Title 14--Aeronautics and Space
   CHAPTER I--FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION
     SUBCHAPTER C--AIRCRAFT
         PART 23--AIRWORTHINESS STANDARDS: NORMAL, UTILITY, ACROBATIC, AND
          COMMUTER CATEGORY AIRPLANES
           Special Federal Aviation Regulations
               SFAR No. 23
               SFAR No. 41
           Subpart A--General
               Sec. 23.1 Applicability.
               Sec. 23.2 Special retroactive requirements.
               Sec. 23.3 Airplane categories.
           Subpart B--Flight
             General
               Sec. 23.21 Proof of compliance.
               Sec. 23.23 Load distribution limits.
               Sec. 23.25 Weight limits.
               Sec. 23.29 Empty weight and corresponding center of gravity.
               Sec. 23.31 Removable ballast.
               Sec. 23.33 Propeller speed and pitch limits.
             Performance
               Sec. 23.45 General.
               Sec. 23.49 Stalling speed.
               Sec. 23.51 Takeoff.
               Sec. 23.53 Takeoff speeds.
               Sec. 23.55 Accelerate-stop distance.
               Sec. 23.57 Takeoff path.
               Sec. 23.59 Takeoff distance and takeoff run.
               Sec. 23.61 Takeoff flight path.
               Sec. 23.65 Climb: All engines operating.
               Sec. 23.67 Climb: one engine inoperative.
               Sec. 23.75 Landing.
               Sec. 23.77 Balked landing.
             Flight Characteristics
               Sec. 23.141 General.
             Controllability and Maneuverability
               Sec. 23.143 General.
               Sec. 23.145 Longitudinal control.
               Sec. 23.147 Directional and lateral control.
               Sec. 23.149 Minimum control speed.
               Sec. 23.151 Acrobatic maneuvers.
               Sec. 23.153 Control during landings.
               Sec. 23.155 Elevator control force in maneuvers.
               Sec. 23.157 Rate of roll.
             Trim
               Sec. 23.161 Trim.
             Stability
               Sec. 23.171 General.
               Sec. 23.173 Static longitudinal stability.
               Sec. 23.175 Demonstration of static longitudinal stability.
               Sec. 23.177 Static directional and lateral stability.
               Sec. 23.179 [Removed. Amdt. No. 23-45, 58 FR 42158, Aug. 6,
                1993]
               Sec. 23.181 Dynamic stability.
             Stalls
               Sec. 23.201 Wings level stall.
               Sec. 23.203 Turning flight and accelerated stalls.
               Sec. 23.205 Critical engine inoperative stalls.
               Sec. 23.207 Stall warning.
             Spinning
               Sec. 23.221 Spinning.
             Ground and Water Handling Characteristics
               Sec. 23.231 Longitudinal stability and control.
               Sec. 23.233 Directional stability and control.
               Sec. 23.235 Taxiing, takeoff, and landing condition.
               Sec. 23.239 Spray characteristics.
             Miscellaneous Flight Requirements
               Sec. 23.251 Vibration and buffeting.
               Sec. 23.253 High speed characteristics.
           Subpart C--Structure
             General
               Sec. 23.301 Loads.
               Sec. 23.302 Canard or tandem wing configurations.
               Sec. 23.303 Factor of safety.
               Sec. 23.305 Strength and deformation.
               Sec. 23.307 Proof of structure.
             Flight Loads
               Sec. 23.321 General.
               Sec. 23.331 Symmetrical flight conditions.
               Sec. 23.333 Flight envelope.
               Sec. 23.335 Design airspeeds.
               Sec. 23.337 Limit maneuvering load factors.
               Sec. 23.341 Gust loads factors.
               Sec. 23.345 High lift devices.
               Sec. 23.347 Unsymmetrical flight conditions.
               Sec. 23.349 Rolling conditions.
               Sec. 23.351 Yawing conditions.
               Sec. 23.361 Engine torque.
               Sec. 23.363 Side load on engine mount.
               Sec. 23.365 Pressurized cabin loads.
               Sec. 23.367 Unsymmetrical loads due to engine failure.
               Sec. 23.369 Rear lift truss.
               Sec. 23.371 Gyroscopic and areodynamic loads.
               Sec. 23.373 Speed control devices.
             Control Surface and System Loads
               Sec. 23.391 Control surface loads.
               Sec. 23.395 Control system loads.
               Sec. 23.397 Limit control forces and torques.
               Sec. 23.399 Dual control system.
               Sec. 23.405 Secondary control system.
               Sec. 23.407 Trim tab effects.
               Sec. 23.409 Tabs.
               Sec. 23.415 Ground gust conditions.
             Horizontal Stabilizing and Balancing Surfaces
               Sec. 23.421 Balancing loads.
               Sec. 23.423 Maneuvering loads.
               Sec. 23.425 Gust loads.
               Sec. 23.427 Unsymmetrical loads.
             Vertical Surfaces
               Sec. 23.441 Maneuvering loads.
               Sec. 23.443 Gust loads.
               Sec. 23.445 Outboard fins or winglets.
             Ailerons, Wing Flaps, and Special Devices
               Sec. 23.455 Ailerons.
               Sec. 23.457 Wing flaps.
               Sec. 23.459 Special devices.
             Ground Loads
               Sec. 23.471 General.
               Sec. 23.473 Ground load conditions and assumptions.
               Sec. 23.477 Landing gear arrangement.
               Sec. 23.479 Level landing conditions.
               Sec. 23.481 Tail down landing conditions.
               Sec. 23.483 One-wheel landing conditions.
               Sec. 23.485 Side load conditions.
               Sec. 23.493 Braked roll conditions.
               Sec. 23.497 Supplementary conditions for tail wheels.
               Sec. 23.499 Supplementary conditions for nose wheels.
               Sec. 23.505 Supplementary conditions for skiplanes.
               Sec. 23.507 Jacking loads.
               Sec. 23.509 Towing loads.
               Sec. 23.511 Ground load; unsymmetrical loads on multiple-wheel
                units.
             Water Loads
               Sec. 23.521 Water load conditions.
               Sec. 23.523 Design weights and center of gravity positions.
               Sec. 23.525 Application of loads.
               Sec. 23.527 Hull and main float load factors.
               Sec. 23.529 Hull and main float landing conditions.
               Sec. 23.531 Hull and main float takeoff condition.
               Sec. 23.533 Hull and main float bottom pressures.
               Sec. 23.535 Auxiliary float loads.
               Sec. 23.537 Seawing loads.
             Emergency Landing Conditions
               Sec. 23.561 General.
               Sec. 23.562 Emergency landing dynamic conditions.
             Fatigue Evaluation
               Sec. 23.571 Pressurized cabin.
               Sec. 23.572 Wing, empennage, and associated structures.
               Sec. 23.573 Damage tolerance and fatigue evaluation of
                structure.
           Subpart D--Design and Construction
               Sec. 23.601 General.
               Sec. 23.603 Materials and workmanship.
               Sec. 23.605 Fabrication methods.
               Sec. 23.607 Self-locking nuts.
               Sec. 23.609 Protection of structure.
               Sec. 23.611 Accessibility.
               Sec. 23.613 Material strength properties and design values.
               Sec. 23.615 [Removed. Amdt. No. 23-45, 58 FR 42164, Aug. 6,
                1993]
               Sec. 23.619 Special factors.
               Sec. 23.621 Casting factors.
               Sec. 23.623 Bearing factors.
               Sec. 23.625 Fitting factors.
               Sec. 23.627 Fatigue strength.
               Sec. 23.629 Flutter.
             Wings
               Sec. 23.641 Proof of strength.
             Control Surfaces
               Sec. 23.651 Proof of strength.
               Sec. 23.655 Installation.
               Sec. 23.657 Hinges.
               Sec. 23.659 Mass balance.
             Control Systems
               Sec. 23.671 General.
               Sec. 23.672 Stability augmentation and automatic and power-
                operated systems.
               Sec. 23.673 Primary flight controls.
               Sec. 23.675 Stops.
               Sec. 23.677 Trim systems.
               Sec. 23.679 Control system locks.
               Sec. 23.681 Limit load static tests.
               Sec. 23.683 Operation tests.
               Sec. 23.685 Control system details.
               Sec. 23.687 Spring devices.
               Sec. 23.689 Cable systems.
               Sec. 23.693 Joints.
               Sec. 23.697 Wing flap controls.
               Sec. 23.699 Wing flap position indicator.
               Sec. 23.701 Flap interconnection.
             Landing Gear
               Sec. 23.721 General.
               Sec. 23.723 Shock absorption tests.
               Sec. 23.725 Limit drop tests.
               Sec. 23.726 Ground load dynamic tests.
               Sec. 23.727 Reserve energy absorption drop test.
               Sec. 23.729 Landing gear extension and retraction system.
               Sec. 23.731 Wheels.
               Sec. 23.733 Tires.
               Sec. 23.735 Brakes.
               Sec. 23.737 Skis.
             Floats and Hulls
               Sec. 23.751 Main float buoyancy.
               Sec. 23.753 Main float design.
               Sec. 23.755 Hulls.
               Sec. 23.757 Auxiliary floats.
             Personnel and Cargo Accommodations
               Sec. 23.771 Pilot compartment.
               Sec. 23.773 Pilot compartment view.
               Sec. 23.775 Windshields and windows.
               Sec. 23.777 Cockpit controls.
               Sec. 23.779 Motion and effect of cockpit controls.
               Sec. 23.781 Cockpit control knob shape.
               Sec. 23.783 Doors.
               Sec. 23.785 Seats, berths, litters, safety belts, and shoulder
                harnesses.
               Sec. 23.787 Baggage and cargo compartments.
               Sec. 23.803 Emergency evacuation.
               Sec. 23.807 Emergency exits.
               Sec. 23.811 Emergency exit marking.
               Sec. 23.813 Emergency exit access.
               Sec. 23.815 Width of aisle.
               Sec. 23.831 Ventilation.
             Pressurization
               Sec. 23.841 Pressurized cabins.
               Sec. 23.843 Pressurization tests.
             Fire Protection
               Sec. 23.851 Fire extinguishers.
               Sec. 23.853 Compartment interiors.
               Sec. 23.859 Combustion heater fire protection.
               Sec. 23.863 Flammable fluid fire protection.
               Sec. 23.865 Fire protection of flight controls, engine mounts,
                and other flight structure.
             Lightning Evaluation
               Sec. 23.867 Lightning protection of structure.
             Miscellaneous
               Sec. 23.871 Leveling means.
           Subpart E--Powerplant
             General
               Sec. 23.901 Installation.
               Sec. 23.903 Engines.
               Sec. 23.904 Automatic power reserve system.
               Sec. 23.905 Propellers.
               Sec. 23.907 Propeller vibration.
               Sec. 23.909 Turbocharger systems.
               Sec. 23.925 Propeller clearance.
               Sec. 23.929 Engine installation ice protection.
               Sec. 23.933 Reversing systems.
               Sec. 23.934 Turbojet and turbofan engine thrust reverser
                systems tests.
               Sec. 23.937 Turbopropeller-drag limiting systems.
               Sec. 23.939 Powerplant operating characteristics.
               Sec. 23.943 Negative acceleration.
             Fuel System
               Sec. 23.951 General.
               Sec. 23.953 Fuel system independence.
               Sec. 23.954 Fuel system lightning protection.
               Sec. 23.955 Fuel flow.
               Sec. 23.957 Flow between interconnected tanks.
               Sec. 23.959 Unusable fuel supply.
               Sec. 23.961 Fuel system hot weather operation.
               Sec. 23.963 Fuel tanks: general.
               Sec. 23.965 Fuel tank tests.
               Sec. 23.967 Fuel tank installation.
               Sec. 23.969 Fuel tank expansion space
               Sec. 23.971 Fuel tank sump.
               Sec. 23.973 Fuel tank filler connection.
               Sec. 23.975 Fuel tank vents and carburetor vapor vents.
               Sec. 23.977 Fuel tank outlet.
               Sec. 23.979 Pressure fueling systems.
             Fuel System Components
               Sec. 23.991 Fuel pumps.
               Sec. 23.993 Fuel system lines and fittings.
               Sec. 23.994 Fuel system components.
               Sec. 23.995 Fuel valves and controls.
               Sec. 23.997 Fuel strainer or filter.
               Sec. 23.999 Fuel system drains.
               Sec. 23.1001 Fuel jettisoning system.
             Oil System
               Sec. 23.1011 General.
               Sec. 23.1013 Oil tanks.
               Sec. 23.1015 Oil tank tests.
               Sec. 23.1017 Oil lines and fittings.
               Sec. 23.1019 Oil strainer or filter.
               Sec. 23.1021 Oil system drains.
               Sec. 23.1023 Oil radiators.
               Sec. 23.1027 Propeller feathering system.
             Cooling
               Sec. 23.1041 General.
               Sec. 23.1043 Cooling tests.
               Sec. 23.1045 Cooling test procedures for turbine engine powered
                airplanes.
               Sec. 23.1047 Cooling test procedures for reciprocating engine-
                powered airplanes.
             Liquid Cooling
               Sec. 23.1061 Installation.
               Sec. 23.1063 Coolant tank tests.
             Induction System
               Sec. 23.1091 Air induction system.
               Sec. 23.1093 Induction system icing protection.
               Sec. 23.1095 Carburetor deicing fluid flow rate.
               Sec. 23.1097 Carburetor deicing fluid system capacity.
               Sec. 23.1099 Carburetor deicing fluid system detail design.
               Sec. 23.1101 Induction air preheater design.
               Sec. 23.1103 Induction system ducts.
               Sec. 23.1105 Induction system screens.
               Sec. 23.1107 Induction system filters.
               Sec. 23.1109 Turbocharger bleed air system.
               Sec. 23.1111 Turbine engine bleed air system.
             Exhaust System
               Sec. 23.1121 General.
               Sec. 23.1123 Exhaust system.
               Sec. 23.1125 Exhaust heat exchangers.
             Powerplant Controls and Accessories
               Sec. 23.1141 Powerplant controls: general.
               Sec. 23.1142 Auxiliary power unit controls.
               Sec. 23.1143 Engine controls.
               Sec. 23.1145 Ignition switches.
               Sec. 23.1147 Mixture controls.
               Sec. 23.1149 Propeller speed and pitch controls.
               Sec. 23.1153 Propeller feathering controls.
               Sec. 23.1155 Turbine engine reverse thrust and propeller pitch
                settings below the flight regime.
               Sec. 23.1157 Carburetor air temperature controls.
               Sec. 23.1163 Powerplant accessories.
               Sec. 23.1165 Engine ignition systems.
               Sec. 23.1181 Designated fire zones; regions included.
             Powerplant Fire Protection
               Sec. 23.1182 Nacelle areas behind firewalls.
               Sec. 23.1183 Lines, fittings, and components.
               Sec. 23.1189 Shutoff means.
               Sec. 23.1191 Firewalls.
               Sec. 23.1192 Engine accessory compartment diaphragm.
               Sec. 23.1193 Cowling and nacelle.
               Sec. 23.1195 Fire extinguishing systems.
               Sec. 23.1197 Fire extinguishing agents.
               Sec. 23.1199 Extinguishing agent containers.
               Sec. 23.1201 Fire extinguishing system materials.
               Sec. 23.1203 Fire detector system.
           Subpart F--Equipment
             General
               Sec. 23.1301 Function and installation.
               Sec. 23.1303 Flight and navigation instruments.
               Sec. 23.1305 Powerplant instruments.
               Sec. 23.1307 Miscellaneous equipment.
               Sec. 23.1309 Equipment, systems, and installations.
             Instruments: Installation
               Sec. 23.1311 Electronic display instrument systems.
               Sec. 23.1321 Arrangement and visibility.
               Sec. 23.1322 Warning, caution, and advisory lights.
               Sec. 23.1323 Airspeed indicating system.
               Sec. 23.1325 Static pressure system.
               Sec. 23.1327 Magnetic direction indicator.
               Sec. 23.1329 Automatic pilot system.
               Sec. 23.1331 Instruments using a power source.
               Sec. 23.1335 Flight director systems.
               Sec. 23.1337 Powerplant instruments.
             Electrical Systems and Equipment
               Sec. 23.1351 General.
               Sec. 23.1353 Storage battery design and installation.
               Sec. 23.1357 Circuit protective devices.
               Sec. 23.1361 Master switch arrangement.
               Sec. 23.1365 Electric cables and equipment.
               Sec. 23.1367 Switches.
             Lights
               Sec. 23.1381 Instrument lights.
               Sec. 23.1383 Landing lights.
               Sec. 23.1385 Position light system installation.
               Sec. 23.1387 Position light system dihedral angles.
               Sec. 23.1389 Position light distribution and intensities.
               Sec. 23.1391 Minimum intensities in the horizontal plane of
                position lights.
               Sec. 23.1393 Minimum intensities in any vertical plane of
                position lights.
               Sec. 23.1395 Maximum intensities in overlapping beams of
                position lights.
               Sec. 23.1397 Color specifications.
               Sec. 23.1399 Riding light.
               Sec. 23.1401 Anticollision light system.
             Safety Equipment
               Sec. 23.1411 General.
               Sec. 23.1413 Safety belts and harnesses.
               Sec. 23.1415 Ditching equipment.
               Sec. 23.1416 Pneumatic de-icer boot system.
               Sec. 23.1419 Ice protection.
             Miscellaneous Equipment
               Sec. 23.1431 Electronic equipment.
               Sec. 23.1435 Hydraulic systems.
               Sec. 23.1437 Accessories for multiengine airplanes.
               Sec. 23.1438 Pressurization and pneumatic systems.
               Sec. 23.1441 Oxygen equipment and supply.
               Sec. 23.1443 Minimum mass flow of supplemental oxygen.
               Sec. 23.1445 Oxygen distribution system.
               Sec. 23.1447 Equipment standards for oxygen dispensing units.
               Sec. 23.1449 Means for determining use of oxygen.
               Sec. 23.1450 Chemical oxygen generators.
               Sec. 23.1457 Cockpit voice recorders.
               Sec. 23.1459 Flight recorders.
               Sec. 23.1461 Equipment containing high energy rotors.
           Subpart G--Operating Limitations and Information
               Sec. 23.1501 General.
               Sec. 23.1505 Airspeed limitations.
               Sec. 23.1507 Operating maneuvering speed.
               Sec. 23.1511 Flap extended speed.
               Sec. 23.1513 Minimum control speed.
               Sec. 23.1519 Weight and center of gravity.
               Sec. 23.1521 Powerplant limitations.
               Sec. 23.1522 Auxiliary power unit limitations.
               Sec. 23.1523 Minimum flight crew.
               Sec. 23.1524 Maximum passenger seating configuration.
               Sec. 23.1525 Kinds of operation.
               Sec. 23.1527 Maximum operating altitude.
               Sec. 23.1529 Instructions for Continued Airworthiness.
             Markings And Placards
               Sec. 23.1541 General.
               Sec. 23.1543 Instrument markings: general.
               Sec. 23.1545 Airspeed indicator.
               Sec. 23.1547 Magnetic direction indicator.
               Sec. 23.1549 Powerplant and auxiliary power unit instruments.
               Sec. 23.1551 Oil quantity indicator.
               Sec. 23.1553 Fuel quantity indicator.
               Sec. 23.1555 Control markings.
               Sec. 23.1557 Miscellaneous markings and placards.
               Sec. 23.1559 Operating limitations placard.
               Sec. 23.1561 Safety equipment.
               Sec. 23.1563 Airspeed placards.
               Sec. 23.1567 Flight maneuver placard.
             Airplane Flight Manual and Approved Manual Material
               Sec. 23.1581 General.
               Sec. 23.1583 Operating limitations.
               Sec. 23.1585 Operating procedures.
               Sec. 23.1587 Performance information.
               Sec. 23.1589 Loading information.
           Appendix A to Part 23--Simplified Design Load Criteria for
            Conventional, Single-Engine Airplanes of 6,000 Pounds or Less
            Maximum Weight
           Appendix B to Part 23--[Reserved]
           Appendix C to Part 23--Basic Landing Conditions
           Appendix D to Part 23--Wheel Spin-Up and Spring-Back Loads
           Appendix E to Part 23--Limited Weight Credit for Airplanes Equipped
            With Standby Power
           Appendix F to Part 23--Test Procedure
           Appendix G to Part 23--Instructions for Continued Airworthiness
             Appendix H to Part 23--Installation of An Automatic Power Reserve
              (APR) System
           Appendix I to Part 23--Seaplane Loads




                              SFAR No. 23

   1. Applicability. An applicant is entitled to a type certificate in the
 normal category for a reciprocating or turbopropeller multiengine powered
 small airplane that is to be certificated to carry more than 10 occupants and
 that is intended for use in operations under Part 135 of the Federal Aviation
 Regulations if he shows compliance with the applicable requirements of Part
 23 of the Federal Aviation Regulations, as supplemented or modified by the
 additional airworthiness requirements of this regulation.
   2. References. Unless otherwise provided, all references in this regulation
 to specific sections of Part 23 of the Federal Aviation Regulations are those
 sections of Part 23 in effect on March 30, 1967.

                              Flight Requirements

   3. General. Compliance must be shown with the applicable requirements of
 Subpart B of Part 23 of the Federal Aviation Regulations in effect on March
 30, 1967, as supplemented or modified in sections 4 through 10 of this
 regulation.

                                  Performance

   4. General. (a) Unless otherwise prescribed in this regulation, compliance
 with each applicable performance requirement in sections 4 through 7 of this
 regulation must be shown for ambient atmospheric conditions and still air.
   (b) The performance must correspond to the propulsive thrust available
 under the particular ambient atmospheric conditions and the particular flight
 condition. The available propulsive thrust must correspond to engine power or
 thrust, not exceeding the approved power or thrust less--
   (1) Installation losses; and
   (2) The power or equivalent thrust absorbed by the accessories and services
 appropriate to the particular ambient atmospheric conditions and the
 particular flight condition.
   (c) Unless otherwise prescribed in this regulation, the applicant must
 select the take-off, en route, and landing configurations for the airplane.
   (d) The airplane configuration may vary with weight, altitude, and
 temperature, to the extent they are compatible with the operating procedures
 required by paragraph (e) of this section.
   (e) Unless otherwise prescribed in this regulation, in determining the
 critical engine inoperative takeoff performance, the accelerate-stop
 distance, takeoff distance, changes in the airplane's configuration, speed,
 power, and thrust, must be made in accordance with procedures established by
 the applicant for operation in service.
   (f) Procedures for the execution of balked landings must be established by
 the applicant and included in the Airplane Flight Manual.
   (g) The procedures established under paragraphs (e) and (f) of this section
 must--
   (1) Be able to be consistently executed in service by a crew of average
 skill;
   (2) Use methods or devices that are safe and reliable; and
   (3) Include allowance for any time delays, in the execution of the
 procedures, that may reasonably be expected in service.
   5. Takeoff--(a) General. The takeoff speeds described in paragraph (b), the
 accelerate-stop distance described in paragraph (c), and the takeoff distance
 described in paragraph (d), must be determined for--
   (1) Each weight, altitude, and ambient temperature within the operational
 limits selected by the applicant;
   (2) The selected configuration for takeoff;
   (3) The center of gravity in the most unfavorable position;
   (4) The operating engine within approved operating limitation; and
   (5) Takeoff data based on smooth, dry, hard-surface runway.
   (b) Takeoff speeds. (1) The decision speed V1is the calibrated airspeed on
 the ground at which, as a result of engine failure or other reasons, the
 pilot is assumed to have made a decision to continue or discontinue the
 takeoff. The speed V1 must be selected by the applicant but may not be less
 than--
   (i) 1.10 Vs1;
   (ii) 1.10 VMC;
   (iii) A speed that permits acceleration to V1 and stop in accordance with
 paragraph (c) allowing credit for an overrun distance equal to that required
 to stop the airplane from a ground speed of 35 knots utilizing maximum
 braking; or
   (iv) A speed at which the airplane can be rotated for takeoff and shown to
 be adequate to safely continue the takeoff, using normal piloting skill, when
 the critical engine is suddenly made inoperative.
   (2) Other essential takeoff speeds necessary for safe operation of the
 airplane must be determined and shown in the Airplane Flight Manual.
   (c) Accelerate-stop distance. (1) The accelerate-stop distance is the sum
 of the distances necessary to--
   (i) Accelerate the airplane from a standing start to V1; and
   (ii) Decelerate the airplane from V1 to a speed not greater than 35 knots,
 assuming that in the case of engine failure, failure of the critical engine
 is recognized by the pilot at the speed V1. The landing gear must remain in
 the extended position and maximum braking may be utilized during
 deceleration.
   (2) Means other than wheel brakes may be used to determine the accelerate-
 stop distance if that means is available with the critical engine inoperative
 and--
   (i) Is safe and reliable;
   (ii) Is used so that consistent results can be expected under normal
 operating conditions; and
   (iii) Is such that exceptional skill is not required to control the
 airplane.
   (d) All engines operating takeoff distance.  The all engine operating
 takeoff distance is the horizontal distance required to takeoff and climb to
 a height of 50 feet above the takeoff surface according to procedures in FAR
 23.51(a).
   (e) One-engine-inoperative takeoff. The maximum weight must be determined
 for each altitude and temperature within the operational limits established
 for the airplane, at which the airplane has takeoff capability after failure
 of the critical engine at or above V1 determined in accordance with paragraph
 (b) of this section. This capability may be established--
   (1) By demonstrating a measurably positive rate of climb with the airplane
 in the takeoff configuration, landing gear extended; or
   (2) By demonstrating the capability of maintaining flight after engine
 failure utilizing procedures prescribed by the applicant.
   6. Climb--(a) Landing climb: All-engines-operating.  The maximum weight
 must be determined with the airplane in the landing configuration, for each
 altitude, and ambient temperature within the operational limits established
 for the airplane and with the most unfavorable center of gravity and out-of-
 ground effect in free air, at which the steady gradient of climb will not be
 less than 3.3 percent, with:
   (1) The engines at the power that is available 8 seconds after initiation
 of movement of the power or thrust controls from the mimimum flight idle to
 the takeoff position.
   (2) A climb speed not greater than the approach speed established under
 section 7 of this regulation and not less than the greater of 1.05MC or
 1.10VS1.
   (b) En route climb, one-engine-inoperative.  (1) the maximum weight must be
 determined with the airplane in the en route configuration, the critical
 engine inoperative, the remaining engine at not more than maximum continuous
 power or thrust, and the most unfavorable center of gravity, at which the
 gradient at climb will be not less than--
   (i) 1.2 percent (or a gradient equivalent to 0.20 Vso 2 , if greater) at
 5,000 feet and an ambient temperature of 41 deg. F. or
   (ii) 0.6 percent (or a gradient equivalent to 0.01 Vso 2 , if greater) at
 5,000 feet and ambient temperature of 81 deg. F.
   (2) The minimum climb gradient specified in subdivisions (i) and (ii) of
 subparagraph (1) of this paragraph must vary linearly between 41 deg. F. and
 81 deg. F. and must change at the same rate up to the maximum operational
 temperature approved for the airplane.
   7. Landing. The landing distance must be determined for standard atmosphere
 at each weight and altitude in accordance with FAR 23.75(a), except that
 instead of the gliding approach specified in FAR 23.75(a)(1), the landing may
 be preceded by a steady approach down to the 50-foot height at a gradient of
 descent not greater than 5.2 percent (3 deg.) at a calibrated airspeed not
 less than 1.3s1.

                                     Trim

   8. Trim--(a) Lateral and directional trim.  The airplane must maintain
 lateral and directional trim in level flight at a speed of Vh or VMO/MMO,
 whichever is lower, with landing gear and wing flaps retracted.
   (b) Longitudinal trim. The airplane must maintain longitudinal trim during
 the following conditions, except that it need not maintain trim at a speed
 greater than VMO/MMO:
   (1) In the approach conditions specified in FAR 23.161(c) (3) through (5),
 except that instead of the speeds specified therein, trim must be maintained
 with a stick force of not more than 10 pounds down to a speed used in showing
 compliance with section 7 of this regulation or 1.4 Vs1 whichever is lower.
   (2) In level flight at any speed from VH or VMO/MMO, whichever is lower, to
 either Vx or 1.4 Vs1, with the landing gear and wing flaps retracted.

                                   Stability

   9. Static longitudinal stability. (a) In showing compliance with the
 provisions of FAR 23.175(b) and with paragraph (b) of this section, the
 airspeed must return to within +/-7 1/2  percent of the trim speed.
   (b) Cruise stability. The stick force curve must have a stable slope for a
 speed range of +/-50 knots from the trim speed except that the speeds need
 not exceed VFC/MFC or be less than 1.4 Vs1. This speed range will be
 considered to begin at the outer extremes of the friction band and the stick
 force may not exceed 50 pounds with--
   (i) Landing gear retracted;
   (ii) Wing flaps retracted;
   (iii) The maximum cruising power as selected by the applicant as an
 operating limitation for turbine engines or 75 percent of maximum continuous
 power for reciprocating engines except that the power need not exceed that
 required at VMO/MMO:
   (iv) Maximum takeoff weight; and
   (v) The airplane trimmed for level flight with the power specified in
 subparagraph (iii) of this paragraph.
   VFC/MFC may not be less than a speed midway between VMO/MMO and VDF/MDF,
 except that, for altitudes where Mach number is the limiting factor, MFC need
 not exceed the Mach number at which effective speed warning occurs.
   (c) Climb stability. For turbopropeller powered airplanes only. In showing
 compliance with FAR 23.175(a), an applicant must in lieu of the power
 specified in FAR 23.175(a)(4), use the maximum power or thrust selected by
 the applicant as an operating limitation for use during climb at the best
 rate of climb speed except that the speed need not be less than 1.4 Vs1.

                                    Stalls

   10. Stall warning. If artificial stall warning is required to comply with
 the requirements of FAR 23.207, the warning device must give clearly
 distinguishable indications under expected conditions of flight. The use of a
 visual warning device that requires the attention of the crew within the
 cockpit is not acceptable by itself.

                                Control Systems

   11. Electric trim tabs. The airplane must meet the requirements of FAR
 23.677 and in addition it must be shown that the airplane is safely
 controllable and that a pilot can perform all the maneuvers and operations
 necessary to effect a safe landing following any probable electric trim tab
 runaway which might be reasonably expected in service allowing for
 appropriate time delay after pilot recognition of the runaway. This
 demonstration must be conducted at the critical airplane weights and center
 of gravity positions.

                           Instruments: Installation

   12. Arrangement and visibility. Each instrument must meet the requirements
 of FAR 23.1321 and in addition--
   (a) Each flight, navigation, and powerplant instrument for use by any pilot
 must be plainly visible to him from his station with the minimum practicable
 deviation from his normal position and line of vision when he is looking
 forward along the flight path.
   (b) The flight instruments required by FAR 23.1303 and by the applicable
 operating rules must be grouped on the instrument panel and centered as
 nearly as practicable about the vertical plane of each pilot's forward
 vision. In addition--
   (1) The instrument that most effectively indicates the attitude must be on
 the panel in the top center position;
   (2) The instrument that most effectively indicates airspeed must be
 adjacent to and directly to the left of the instrument in the top center
 position;
   (3) The instrument that most effectively indicates altitude must be
 adjacent to and directly to the right of the instrument in the top center
 position; and
   (4) The instrument that most effectively indicates direction of flight must
 be adjacent to and directly below the instrument in the top center position.
   13. Airspeed indicating system. Each airspeed indicating system must meet
 the requirements of FAR 23.1323 and in addition--
   (a) Airspeed indicating instruments must be of an approved type and must be
 calibrated to indicate true airspeed at sea level in the standard atmosphere
 with a mimimum practicable instrument calibration error when the
 corresponding pilot and static pressures are supplied to the instruments.
   (b) The airspeed indicating system must be calibrated to determine the
 system error, i.e., the relation between IAS and CAS, in flight and during
 the accelerate takeoff ground run. The ground run calibration must be
 obtained between 0.8 of the mimimum value of V1 and 1.2 times the maximum
 value of V1, considering the approved ranges of altitude and weight. The
 ground run calibration will be determined assuming an engine failure at the
 mimimum value of V1.
   (c) The airspeed error of the installation excluding the instrument
 calibration error, must not exceed 3 percent or 5 knots whichever is greater,
 throughout the speed range from VMO to 1.3S1 with flaps retracted and from
 1.3 VSO to VFE with flaps in the landing position.
   (d) Information showing the relationship between IAS and CAS must be shown
 in the Airplane Flight Manual.
   14. Static air vent system. The static air vent system must meet the
 requirements of FAR 23.1325. The altimeter system calibration must be
 determined and shown in the Airplane Flight Manual.

                     Operating Limitations and Information

   15. Maximum operating limit speed VMO/MMO. Instead of establishing
 operating limitations based on VME and VNO, the applicant must establish a
 maximum operating limit speed VMO/MMO in accordance with the following:
   (a) The maximum operating limit speed must not exceed the design cruising
 speed Vc and must be sufficiently below VD/MD or VDF/MDF to make it highly
 improbable that the latter speeds will be inadvertently exceeded in flight.
   (b) The speed Vmo must not exceed 0.8 VD/MD or 0.8 VDF/MDF unless flight
 demonstrations involving upsets as specified by the Administrator indicates a
 lower speed margin will not result in speeds exceeding VD/MD or VDF.
 Atmospheric variations, horizontal gusts, and equipment errors, and airframe
 production variations will be taken into account.
   16. Minimum flight crew. In addition to meeting the requirements of FAR
 23.1523, the applicant must establish the minimum number and type of
 qualified flight crew personnel sufficient for safe operation of the airplane
 considering--
   (a) Each kind of operation for which the applicant desires approval;
   (b) The workload on each crewmember considering the following:
   (1) Flight path control.
   (2) Collision avoidance.
   (3) Navigation.
   (4) Communications.
   (5) Operation and monitoring of all essential aircraft systems.
   (6) Command decisions; and
   (c) The accessibility and ease of operation of necessary controls by the
 appropriate crewmember during all normal and emergency operations when at his
 flight station.
   17. Airspeed indicator. The airspeed indicator must meet the requirements
 of FAR 23.1545 except that, the airspeed notations and markings in terms of
 VNO and VNE must be replaced by the VMO/MMO notations. The airspeed indicator
 markings must be easily read and understood by the pilot. A placard adjacent
 to the airspeed indicator is an acceptable means of showing compliance with
 the requirements of FAR 23.1545(c).

                            Airplane Flight Manual

   18. General. The Airplane Flight Manual must be prepared in accordance with
 the requirements of FARs 23.1583 and 23.1587, and in addition the operating
 limitations and performance information set forth in sections 19 and 20 must
 be included.
   19. Operating limitations. The Airplane Flight Manual must include the
 following limitations--
   (a) Airspeed limitations. (1) The maximum operating limit speed VMO/MMO and
 a statement that this speed limit may not be deliberately exceeded in any
 regime of flight (climb, cruise, or descent) unless a higher speed is
 authorized for flight test or pilot training;
   (2) If an airspeed limitation is based upon compressibility effects, a
 statement to this effect and information as to any symptoms, the probable
 behavior of the airplane, and the recommended recovery procedures; and
   (3) The airspeed limits, shown in terms of VMO/MMO instead of VNO and VNE.
   (b) Takeoff weight limitations. The maximum takeoff weight for each airport
 elevation, ambient temperature, and available takeoff runway length within
 the range selected by the applicant. This weight may not exceed the weight at
 which:
   (1) The all-engine operating takeoff distance determined in accordance with
 section 5(d) or the accelerate-stop distance determined in accordance with
 section 5(c), which ever is greater, is equal to the available runway length;
   (2) The airplane complies with the one-engine-inoperative takeoff
 requirements specified in section 5(e); and
   (3) The airplane complies with the one-engine-inoperative en route climb
 requirements specified in section 6(b), assuming that a standard temperature
 lapse rate exists from the airport elevation to the altitude of 5,000 feet,
 except that the weight may not exceed that corresponding to a temperature of
 41 deg. F at 5,000 feet.
   20. Performance information. The Airplane Flight Manual must contain the
 performance information determined in accordance with the provisions of the
 performance requirements of this regulation. The information must include the
 following:
   (a) Sufficient information so that the take-off weight limits specified in
 section 19(b) can be determined for all temperatures and altitudes within the
 operation limitations selected by the applicant.
   (b) The conditions under which the performance information was obtained,
 including the airspeed at the 50-foot height used to determine landing
 distances.
   (c) The performance information (determined by extrapolation and computed
 for the range of weights between the maximum landing and takeoff weights)
 for--
   (1) Climb in the landing configuration; and
   (2) Landing distance.
   (d) Procedure established under section 4 of this regulation related to the
 limitations and information required by this section in the form of guidance
 material including any relevant limitations or information.
   (e) An explanation of significant or unusual flight or ground handling
 characteristics of the airplane.
   (f) Airspeeds, as indicated airspeeds, corresponding to those determined
 for takeoff in accordance with section 5(b).
   21. Maximum operating altitudes. The maximum operating altitude to which
 operation is permitted, as limited by flight, structural, powerplant,
 functional, or equipment characteristics, must be specified in the Airplane
 Flight Manual.
   22. Stowage provision for Airplane Flight Manual. Provision must be made
 for stowing the Airplane Flight Manual in a suitable fixed container which is
 readily accessible to the pilot.
   23. Operating procedures. Procedures for restarting turbine engines in
 flight (including the effects of altitude) must be set forth in the Airplane
 Flight Manual.

                             Airframe Requirements

                                 FLIGHT LOADS

   24. Engine torque. (a) Each turbopropeller engine mount and its supporting
 structure must be designed for the torque effects of--
   (1) The conditions set forth in FAR 23.361(a).
   (2) The limit engine torque corresponding to takeoff power and propeller
 speed, multiplied by a factor accounting for propeller control system
 malfunction, including quick feathering action, simultaneously with 1 g level
 flight loads. In the absence of a rational analysis, a factor of 1.6 must be
 used.
   (b) The limit torque is obtained by multiplying the mean torque by a factor
 of 1.25.
   25. Turbine engine gyroscopic loads. Each turbopropeller engine mount and
 its supporting structure must be designed for the gyroscopic loads that
 result, with the engines at maximum continuous r.p.m., under either--
   (a) The conditions prescribed in FARs 23.351 and 23.423; or
   (b) All possible combinations of the following:
   (1) A yaw velocity of 2.5 radius per second.
   (2) A pitch velocity of 1.0 radians per second.
   (3) A normal load factor of 2.5.
   (4) Maximum continuous thrust.
   26. Unsymmetrical loads due to engine failure. (a) Turbopropeller powered
 airplanes must be designed for the unsymmetrical loads resulting from the
 failure of the critical engine including the following conditions in
 combination with a single malfunction of the propeller drag limiting system,
 considering the probable pilot corrective action on the flight controls.
   (1) At speeds between VMC and VD, the loads resulting from power failure
 because of fuel flow interruption are considered to be limit loads.
   (2) At speeds between VMC and VC, the loads resulting from the
 disconnection of the engine compressor from the turbine or from loss of the
 turbine blades are considered to be ultimate loads.
   (3) The time history of the thrust decay and drag buildup occurring as a
 result of the prescribed engine failures must be substantiated by test or
 other data applicable to the particular engine-propeller combination.
   (4) The timing and magnitude of the probable pilot corrective action must
 be conservatively estimated, considering the characteristics of the
 particular engine-propeller-airplane combination.
   (b) Pilot corrective action may be assumed to be initiated at the time
 maximum yawing velocity is reached, but not earlier than two seconds after
 the engine failure. The magnitude of the corrective action may be based on
 the control forces specified in FAR 23.397 except that lower forces may be
 assumed where it is shown by analysis or test that these forces can control
 the yaw and roll resulting from the prescribed engine failure conditions.

                                 Ground Loads

   27. Dual wheel landing gear units. Each dual wheel landing gear unit and
 its supporting structure must be shown to comply with the following:
   (a) Pivoting. The airplane must be assumed to pivot about one side of the
 main gear with the brakes on that side locked. The limit vertical load factor
 must be 1.0 and the coefficient of friction 0.8. This condition need apply
 only to the main gear and its supporting structure.
   (b) Unequal tire inflation. A 60-40 percent distribution of the loads
 established in accordance with FAR 23.471 through FAR 23.483 must be applied
 to the dual wheels.
   (c) Flat tire. (1) Sixty percent of the loads specified in FAR 23.471
 through FAR 23.483 must be applied to either wheel in a unit.
   (2) Sixty percent of the limit drag and side loads and 100 percent of the
 limit vertical load established in accordance with FARs 23.493 and 23.485
 must be applied to either wheel in a unit except that the vertical load need
 not exceed the maximum vertical load in paragraph (c)(1) of this section.

                              Fatigue Evaluation

   28. Fatigue evaluation of wing and associated structure. Unless it is shown
 that the structure, operating stress levels, materials, and expected use are
 comparable from a fatigue standpoint to a similar design which has had
 substantial satisfactory service experience, the strength, detail design, and
 the fabrication of those parts of the wing, wing carrythrough, and attaching
 structure whose failure would be catastrophic must be evaluated under
 either--
   (a) A fatigue strength investigation in which the structu@ is @own by
 analysis, tests, or both to be able to withstand the repeated loads of
 variable magnitude expected in service; or
   (b) A fail-safe strength investigation in which it is shown by analysis,
 tests, or both that catastrophic failure of the structure is not probable
 after fatigue, or obvious partial failure, of a principal structural element,
 and that the remaining structure is able to withstand a static ultimate load
 factor of 75 percent of the critical limit load factor at Vc. These loads
 must be multiplied by a factor of 1.15 unless the dynamic effects of failure
 under static load are otherwise considered.

                            Design and Construction

   29. Flutter. For Multiengine turbopropeller powered airplanes, a dynamic
 evaluation must be made and must include--
   (a) The significant elastic, inertia, and aerodynamic forces associated
 with the rotations and displacements of the plane of the propeller; and
   (b) Engine-propeller-nacelle stiffness and damping variations appropriate
 to the particular configuration.

                                 Landing Gear

   30. Flap operated landing gear warning device. Airplanes having retractable
 landing gear and wing flaps must be equipped with a warning device that
 functions continuously when the wing flaps are extended to a flap position
 that activates the warning device to give adequate warning before landing,
 using normal landing procedures, if the landing gear is not fully extended
 and locked. There may not be a manual shut off for this warning device. The
 flap position sensing unit may be installed at any suitable location. The
 system for this device may use any part of the system (including the aural
 warning device) provided for other landing gear warning devices.

                      Personnel and Cargo Accommodations

   31. Cargo and baggage compartments.  Cargo and baggage compartments must be
 designed to meet the requirements of FAR 23.787 (a) and (b), and in addition
 means must be provided to protect passengers from injury by the contents of
 any cargo or baggage compartment when the ultimate forward inertia force is
 9g.
   32. Doors and exits. The airplane must meet the requirements of FAR 23.783
 and FAR 23.807 (a)(3), (b), and (c), and in addition:
   (a) There must be a means to lock and safeguard each external door and exit
 against opening in flight either inadvertently by persons, or as a result of
 mechanical failure. Each external door must be operable from both the inside
 and the outside.
   (b) There must be means for direct visual inspection of the locking
 mechanism by crewmembers to determine whether external doors and exits, for
 which the initial opening movement is outward, are fully locked. In addition,
 there must be a visual means to signal to crewmembers when normally used
 external doors are closed and fully locked.
   (c) The passenger entrance door must qualify as a floor level emergency
 exit. Each additional required emergency exit except floor level exits must
 be located over the wing or must be provided with acceptable means to assist
 the occupants in descending to the ground. In addition to the passenger
 entrance door:
   (1) For a total seating capacity of 15 or less, an emergency exit as
 defined in FAR 23.807(b) is required on each side of the cabin.
   (2) For a total seating capacity of 16 through 23, three emergency exits as
 defined in 23.807(b) are required with one on the same side as the door and
 two on the side opposite the door.
   (d) An evacuation demonstration must be conducted utilizing the maximum
 number of occupants for which certification is desired. It must be conducted
 under simulated night conditions utilizing only the emergency exits on the
 most critical side of the aircraft. The participants must be representative
 of average airline passengers with no prior practice or rehearsal for the
 demonstration. Evacuation must be completed within 90 seconds.
   (e) Each emergency exit must be marked with the word "Exit" by a sign which
 has white letters 1 inch high on a red background 2 inches high, be self-
 illuminated or independently internally electrically illuminated, and have a
 minimum luminescence (brightness) of at least 160 microlamberts. The colors
 may be reversed if the passenger compartment illumination is essentially the
 same.
   (f) Access to window type emergency exits must not be obstructed by seats
 or seat backs.
   (g) The width of the main passenger aisle at any point between seats must
 equal or exceed the values in the following table.

                                         Minimum main
                                       passenger aisle
                                            width

                                       Less
                                     than 25
                                      inches   25 inches
                      Total seating    from     and more
                        capacity      floor    from floor

                      10 through 23  9 inches  15 inches.

                                 Miscellaneous

   33. Lightning strike protection. Parts that are electrically insulated from
 the basic airframe must be connected to it through lightning arrestors unless
 a lightning strike on the insulated part--
   (a) Is improbable because of shielding by other parts; or
   (b) Is not hazardous.
   34. Ice protection. If certification with ice protection provisions is
 desired, compliance with the following requirements must be shown:
   (a) The recommended procedures for the use of the ice protection equipment
 must be set forth in the Airplane Flight Manual.
   (b) An analysis must be performed to establish, on the basis of the
 airplane's operational needs, the adequacy of the ice protection system for
 the various components of the airplane. In addition, tests of the ice
 protection system must be conducted to demonstrate that the airplane is
 capable of operating safely in continuous maximum and intermittent maximum
 icing conditions as described in FAR 25, Appendix C.
   (c) Compliance with all or portions of this section may be accomplished by
 reference, where applicable because of similarity of the designs, to analysis
 and tests performed by the applicant for a type certificated model.
   35. Maintenance information. The applicant must make available to the owner
 at the time of delivery of the airplane the information he considers
 essential for the proper maintenance of the airplane. That information must
 include the following:
   (a) Description of systems, including electrical, hydraulic, and fuel
 controls.
   (b) Lubrication instructions setting forth the frequency and the lubricants
 and fluids which are to be used in the various systems.
   (c) Pressures and electrical loads applicable to the various systems.
   (d) Tolerances and adjustments necessary for proper functioning.
   (e) Methods of leveling, raising, and towing.
   (f) Methods of balancing control surfaces.
   (g) Identification of primary and secondary structures.
   (h) Frequency and extent of inspections necessary to the proper operation
 of the airplane.
   (i) Special repair methods applicable to the airplane.
   (j) Special inspection techniques, including those that require X-ray,
 ultrasonic, and magnetic particle inspection.
   (k) List of special tools.

                                  Propulsion

                                    GENERAL

   36. Vibration characteristics. For turbopropeller powered airplanes, the
 engine installation must not result in vibration characteristics of the
 engine exceeding those established during the type certification of the
 engine.
   37. In-flight restarting of engine. If the engine on turbopropeller powered
 airplanes cannot be restarted at the maximum cruise altitude, a determination
 must be made of the altitude below which restarts can be consistently
 accomplished. Restart information must be provided in the Airplane Flight
 Manual.
   38. Engines--(a) For turbopropeller powered airplanes. The engine
 installation must comply with the following requirements:
   (1) Engine isolation. The powerplants must be arranged and isolated from
 each other to allow operation, in at least one configuration, so that the
 failure or malfunction of any engine, or of any system that can affect the
 engine, will not--
   (i) Prevent the continued safe operation of the remaining engines; or
   (ii) Require immediate action by any crewmember for continued safe
 operation.
   (2) Control of engine rotation. There must be a means to individually stop
 and restart the rotation of any engine in flight except that engine rotation
 need not be stopped if continued rotation could not jeopardize the safety of
 the airplane. Each component of the stopping and restarting system on the
 engine side of the firewall, and that might be exposed to fire, must be at
 least fire resistant. If hydraulic propeller feathering systems are used for
 this purpose, the feathering lines must be at least fire resistant under the
 operating conditions that may be expected to exist during feathering.
   (3) Engine speed and gas temperature control devices. The powerplant
 systems associated with engine control devices, systems, and instrumentation
 must provide reasonable assurance that those engine operating limitations
 that adversely affect turbine rotor structural integrity will not be exceeded
 in service.
   (b) For reciprocating-engine powered airplanes.  To provide engine
 isolation, the powerplants must be arranged and isolated from each other to
 allow operation, in at least one configuration, so that the failure or
 malfunction of any engine, or of any system that can affect that engine, will
 not--
   (1) Prevent the continued safe operation of the remaining engines; or
   (2) Require immediate action by any crewmember for continued safe
 operation.
   39. Turbopropeller reversing systems. (a) Turbopropeller reversing systems
 intended for ground operation must be designed so that no single failure or
 malfunction of the system will result in unwanted reverse thrust under any
 expected operating condition. Failure of structural elements need not be
 considered if the probability of this kind of failure is extremely remote.
   (b) Turbopropeller reversing systems intended for in-flight use must be
 designed so that no unsafe condition will result during normal operation of
 the system, or from any failure (or reasonably likely combination of
 failures) of the reversing system, under any anticipated condition of
 operation of the airplane. Failure of structural elements need not be
 considered if the probability of this kind of failure is extremely remote.
   (c) Compliance with this section may be shown by failure analysis, testing,
 or both for propeller systems that allow propeller blades to move from the
 flight low-pitch position to a position that is substantially less than that
 at the normal flight low-pitch stop position. The analysis may include or be
 supported by the analysis made to show compliance with the type certification
 of the propeller and associated installation components. Credit will be given
 for pertinent analysis and testing completed by the engine and propeller
 manufacturers.
   40. Turbopropeller drag-limiting systems.  Turbopropeller drag-limiting
 systems must be designed so that no single failure or malfunction of any of
 the systems during normal or emergency operation results in propeller drag in
 excess of that for which the airplane was designed. Failure of structural
 elements of the drag-limiting systems need not be considered if the
 probability of this kind of failure is extremely remote.
   41. Turbine engine powerplant operating characteristics. For turbopropeller
 powered airplanes, the turbine engine powerplant operating characteristics
 must be investigated in flight to determine that no adverse characteristics
 (such as stall, surge, or flameout) are present to a hazardous degree, during
 normal and emergency operation within the range of operating limitations of
 the airplane and of the engine.
   42. Fuel flow. (a) For turbopropeller powered airplanes--
   (1) The fuel system must provide for continuous supply of fuel to the
 engines for normal operation without interruption due to depletion of fuel in
 any tank other than the main tank; and
   (2) The fuel flow rate for turbopropeller engine fuel pump systems must not
 be less than 125 percent of the fuel flow required to develop the standard
 sea level atmospheric conditions takeoff power selected and included as an
 operating limitation in the Airplane Flight Manual.
   (b) For reciprocating engine powered airplanes, it is acceptable for the
 fuel flow rate for each pump system (main and reserve supply) to be 125
 percent of the takeoff fuel consumption of the engine.

                            Fuel System Components

   43. Fuel pumps. For turbopropeller powered airplanes, a reliable and
 independent power source must be provided for each pump used with turbine
 engines which do not have provisions for mechanically driving the main pumps.
 It must be demonstrated that the pump installations provide a reliability and
 durability equivalent to that provided by FAR 23.991(a).
   44. Fuel strainer or filter. For turbopropeller powered airplanes, the
 following apply:
   (a) There must be a fuel strainer or filter between the tank outlet and the
 fuel metering device of the engine. In addition, the fuel strainer or filter
 must be--
   (1) Between the tank outlet and the engine-driven positive displacement
 pump inlet, if there is an engine-driven positive displacement pump;
   (2) Accessible for drainage and cleaning and, for the strainer screen,
 easily removable; and
   (3) Mounted so that its weight is not supported by the connecting lines or
 by the inlet or outlet connections of the strainer or filter itself.
   (b) Unless there are means in the fuel system to prevent the accumulation
 of ice on the filter, there must be means to automatically maintain the fuel
 flow if ice-clogging of the filter occurs; and
   (c) The fuel strainer or filter must be of adequate capacity (with respect
 to operating limitations established to insure proper service) and of
 appropriate mesh to insure proper engine operation, with the fuel
 contaminated to a degree (with respect to particle size and density) that can
 be reasonably expected in service. The degree of fuel filtering may not be
 less than that established for the engine type certification.
   45. Lightning strike protection. Protection must be provided against the
 ignition of flammable vapors in the fuel vent system due to lightning
 strikes.

                                    Cooling

   46. Cooling test procedures for turbopropeller powered airplanes. (a)
 Turbopropeller powered airplanes must be shown to comply with the
 requirements of FAR 23.1041 during takeoff, climb en route, and landing
 stages of flight that correspond to the applicable performance requirements.
 The cooling test must be conducted with the airplane in the configuration and
 operating under the conditions that are critical relative to cooling during
 each stage of flight. For the cooling tests a temperature is "stabilized"
 when its rate of change is less than 2 deg. F. per minute.
   (b) Temperatures must be stabilized under the conditions from which entry
 is made into each stage of flight being investigated unless the entry
 condition is not one during which component and engine fluid temperatures
 would stabilize, in which case, operation through the full entry condition
 must be conducted before entry into the stage of flight being investigated in
 order to allow temperatures to reach their natural levels at the time of
 entry. The takeoff cooling test must be preceded by a period during which the
 powerplant component and engine fluid temperatures are stabilized with the
 engines at ground idle.
   (c) Cooling tests for each stage of flight must be continued until--
   (1) The component and engine fluid temperatures stabilize;
   (2) The stage of flight is completed; or
   (3) An operating limitation is reached.

                               Induction System

   47. Air induction. For turbopropeller powered airplanes--
   (a) There must be means to prevent hazardous quantities of fuel leakage or
 overflow from drains, vents, or other components of flammable fluid systems
 from entering the engine intake system; and
   (b) The air inlet ducts must be located or protected so as to minimize the
 ingestion of foreign matter during takeoff, landing, and taxiing.
   48. Induction system icing protection. For turbopropeller powered
 airplanes, each turbine engine must be able to operate throughout its flight
 power range without adverse effect on engine operation or serious loss of
 power or thrust, under the icing conditions specified in Appendix C of FAR
 25. In addition, there must be means to indicate to appropriate flight
 crewmembers the functioning of the powerplant ice protection system.
   49. Turbine engine bleed air systems. Turbine engine bleed air systems of
 turbopropeller powered airplanes must be investigated to determine--
   (a) That no hazard to the airplane will result if a duct rupture occurs.
 This condition must consider that a failure of the duct can occur anywhere
 between the engine port and the airplane bleed service; and
   (b) That if the bleed air system is used for direct cabin pressurization,
 it is not possible for hazardous contamination of the cabin air system to
 occur in event of lubrication system failure.

                                Exhaust System

   50. Exhaust system drains. Turbopropeller engine exhaust systems having low
 spots or pockets must incorporate drains at such locations. These drains must
 discharge clear of the airplane in normal and ground attitudes to prevent the
 accumulation of fuel after the failure of an attempted engine start.

                      Powerplant Controls and Accessories

   51. Engine controls. If throttles or power levers for turbopropeller
 powered airplanes are such that any position of these controls will reduce
 the fuel flow to the engine(s) below that necessary for satisfactory and safe
 idle operation of the engine while the airplane is in flight, a means must be
 provided to prevent inadvertent movement of the control into this position.
 The means provided must incorporate a positive lock or stop at this idle
 position and must require a separate and distinct operation by the crew to
 displace the control from the normal engine operating range.
   52. Reverse thrust controls. For turbopropeller powered airplanes, the
 propeller reverse thrust controls must have a means to prevent their
 inadvertent operation. The means must have a positive lock or stop at the
 idle position and must require a separate and distinct operation by the crew
 to displace the control from the flight regime.
   53. Engine ignition systems. Each turbopropeller airplane ignition system
 must be considered an essential electrical load.
   54. Powerplant accessories. The powerplant accessories must meet the
 requirements of FAR 23.1163, and if the continued rotation of any accessory
 remotely driven by the engine is hazardous when malfunctioning occurs, there
 must be means to prevent rotation without interfering with the continued
 operation of the engine.

                          Powerplant Fire Protection

   55. Fire detector system. For turbopropeller powered airplanes, the
 following apply:
   (a) There must be a means that ensures prompt detection of fire in the
 engine compartment. An overtemperature switch in each engine cooling air exit
 is an acceptable method of meeting this requirement.
   (b) Each fire detector must be constructed and installed to withstand the
 vibration, inertia, and other loads to which it may be subjected in
 operation.
   (c) No fire detector may be affected by any oil, water, other fluids, or
 fumes that might be present.
   (d) There must be means to allow the flight crew to check, in flight, the
 functioning of each fire detector electric circuit.
   (e) Wiring and other components of each fire detector system in a fire zone
 must be at least fire resistant.
   56. Fire protection, cowling and nacelle skin. For reciprocating engine
 powered airplanes, the engine cowling must be designed and constructed so
 that no fire originating in the engine compartment can enter, either through
 openings or by burn through, any other region where it would create
 additional hazards.
   57. Flammable fluid fire protection. If flammable fluids or vapors might be
 liberated by the leakage of fluid systems in areas other than engine
 compartments, there must be means to--
   (a) Prevent the ignition of those fluids or vapors by any other equipment;
 or
   (b) Control any fire resulting from that ignition.

                                   Equipment

   58. Powerplant instruments. (a) The following are required for
 turbopropeller airplanes:
   (1) The instruments required by FAR 23.1305 (a)(1) through (4), (b)(2) and
 (4).
   (2) A gas temperature indicator for each engine.
   (3) Free air temperature indicator.
   (4) A fuel flowmeter indicator for each engine.
   (5) Oil pressure warning means for each engine.
   (6) A torque indicator or adequate means for indicating power output for
 each engine.
   (7) Fire warning indicator for each engine.
   (8) A means to indicate when the propeller blade angle is below the low-
 pitch position corresponding to idle operation in flight.
   (9) A means to indicate the functioning of the ice protection system for
 each engine.
   (b) For turbopropeller powered airplanes, the turbopropeller blade position
 indicator must begin indicating when the blade has moved below the flight
 low-pitch position.
   (c) The following instruments are required for reciprocating-engine powered
 airplanes:
   (1) The instruments required by FAR 23.1305.
   (2) A cylinder head temperature indicator for each engine.
   (3) A manifold pressure indicator for each engine.

                            Systems and Equipments

                                    GENERAL

   59. Function and installation. The systems and equipment of the airplane
 must meet the requirements of FAR 23.1301, and the following:
   (a) Each item of additional installed equipment must--
   (1) Be of a kind and design appropriate to its intended function;
   (2) Be labeled as to its identification, function, or operating
 limitations, or any applicable combination of these factors, unless misuse or
 inadvertent actuation cannot create a hazard;
   (3) Be installed according to limitations specified for that equipment; and
   (4) Function properly when installed.
   (b) Systems and installations must be designed to safeguard against hazards
 to the aircraft in the event of their malfunction or failure.
   (c) Where an installation, the functioning of which is necessary in showing
 compliance with the applicable requirements, requires a power supply, such
 installation must be considered an essential load on the power supply, and
 the power sources and the distribution system must be capable of supplying
 the following power loads in probable operation combinations and for probable
 durations:
   (1) All essential loads after failure of any prime mover, power converter,
 or energy storage device.
   (2) All essential loads after failure of any one engine on two-engine
 airplanes.
   (3) In determining the probable operating combinations and durations of
 essential loads for the power failure conditions described in subparagraphs
 (1) and (2) of this paragraph, it is permissible to assume that the power
 loads are reduced in accordance with a monitoring procedure which is
 consistent with safety in the types of operations authorized.
   60. Ventilation. The ventilation system of the airplane must meet the
 requirements of FAR 23.831, and in addition, for pressurized aircraft the
 ventilating air in flight crew and passenger compartments must be free of
 harmful or hazardous concentrations of gases and vapors in normal operation
 and in the event of reasonably probable failures or malfunctioning of the
 ventilating, heating, pressurization, or other systems, and equipment. If
 accumulation of hazardous quantities of smoke in the cockpit area is
 reasonably probable, smoke evacuation must be readily accomplished.

                       Electrical Systems and Equipment

   61. General. The electrical systems and equipment of the airplane must meet
 the requirements of FAR 23.1351, and the following:
   (a) Electrical system capacity. The required generating capacity, and
 number and kinds of power sources must--
   (1) Be determined by an electrical load analysis, and
   (2) Meet the requirements of FAR 23.1301.
   (b) Generating system. The generating system includes electrical power
 sources, main power busses, transmission cables, and associated control,
 regulation, and protective devices. It must be designed so that--
   (1) The system voltage and frequency (as applicable) at the terminals of
 all essential load equipment can be maintained within the limits for which
 the equipment is designed, during any probable operating conditions;
   (2) System transients due to switching, fault clearing, or other causes do
 not make essential loads inoperative, and do not cause a smoke or fire
 hazard;
   (3) There are means, accessible in flight to appropriate crewmembers, for
 the individual and collective disconnection of the electrical power sources
 from the system; and
   (4) There are means to indicate to appropriate crewmembers the generating
 system quantities essential for the safe operation of the system, including
 the voltage and current supplied by each generator.
   62. Electrical equipment and installation.  Electrical equipment controls,
 and wiring must be installed so that operation of any one unit or system of
 units will not adversely affect the simultaneous operation of to the safe
 operation.
   63. Distribution system. (a) For the purpose of complying with this
 section, the distribution system includes the distribution busses, their
 associated feeders and each control and protective device.
   (b) Each system must be designed so that essential load circuits can be
 supplied in the event of reasonably probable faults or open circuits,
 including faults in heavy current carrying cables.
   (c) If two independent sources of electrical power for particular equipment
 or systems are required by this regulation, their electrical energy supply
 must be insured by means such as duplicate electrical equipment, throwover
 switching, or multichannel or loop circuits separately routed.
   64. Circuit protective devices. The circuit protective devices for the
 electrical circuits of the airplane must meet the requirements of FAR
 23.1357, and in addition circuits for loads which are essential to safe
 operation must have individual and exclusive circuit protection.

 [Doc. No. 8070, 34 FR 189, Jan. 7, 1969, as amended by SFAR 23-1, 34 FR
 20176, Dec. 24, 1969; 35 FR 1102, Jan. 28, 1970]






                                  SFAR No. 41

   Editorial Note: For the text of SFAR No. 41, see Part 21 of this chapter.


Sec. 23.1  Applicability.

   (a) This part prescribes airworthiness standards for the issue of type
 certificates, and changes to those certificates, for airplanes in the normal,
 utility, acrobatic, and commuter categories.
   (b) Each person who applies under Part 21 for such a certificate or change
 must show compliance with the applicable requirements of this part.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-34, 52 FR
 1825, Jan. 15, 1987]






 Sec. 23.2  Special retroactive requirements.

   (a) Notwithstanding Secs. 21.17 and 21.101 of this chapter and irrespective
 of the type certification basis, each normal, utility, and acrobatic category
 airplane having a passenger seating configuration, excluding pilot seats, of
 nine or less, manufactured after December 12, 1986, or any such foreign
 airplane for entry into the United States must provide a safety belt and
 shoulder harness for each forward- or aft-facing seat which will protect the
 occupant from serious head injury when subjected to the inertia loads
 resulting from the ultimate static load factors prescribed in Sec.
 23.561(b)(2) of this part, or which will provide the occupant protection
 specified in Sec. 23.562 of this part when that section is applicable to the
 airplane. For other seat orientations, the seat/restraint system must be
 designed to provide a level of occupant protection equivalent to that
 provided for forward- or aft-facing seats with a safety belt and shoulder
 harness installed.
   (b) Each shoulder harness installed at a flight crewmember station, as
 required by this section, must allow the crewmember, when seated with the
 safety belt and shoulder harness fastened, to perform all functions necessary
 for flight operations.
   (c) For the purpose of this section, the date of manufacture is:
   (1) The date the inspection acceptance records, or equivalent, reflect that
 the airplane is complete and meets the FAA approved type design data; or
   (2) In the case of a foreign manufactured airplane, the date the foreign
 civil airworthiness authority certifies the airplane is complete and issues
 an original standard airworthiness certificate, or the equivalent in that
 country.

 [Amdt. 23-36, 53 FR 30812, Aug. 15, 1988]







 Sec. 23.3  Airplane categories.

   (a) The normal category is limited to airplanes that have a seating
 configuration, excluding pilot seats, of nine or less, a maximum certificated
 takeoff weight of 12,500 pounds or less, and intended for nonacrobatic
 operation. Nonacrobatic operation includes:
   (1) Any maneuver incident to normal flying;
   (2) Stalls (except whip stalls); and
   (3) Lazy eights, chandelles, and steep turns, in which the angle of bank is
 not more than 60 degrees.
   (b) The utility category is limited to airplanes that have a seating
 configuration, excluding pilot seats, of nine or less, a maximum certificated
 takeoff weight of 12,500 pounds or less, and intended for limited acrobatic
 operation. Airplanes certificated in the utility category may be used in any
 of the operations covered under paragraph (a) of this section and in limited
 acrobatic operations. Limited acrobatic operation includes:
   (1) Spins (if approved for the particular type of airplane); and
   (2) Lazy eights, chandelles, and steep turns, in which the angle of bank is
 more than 60 degrees.
   (c) The acrobatic category is limited to airplanes that have a seating
 configuration, excluding pilot seats, of nine or less, a maximum certificated
 takeoff weight of 12,500 pounds or less, and intended for use without
 restrictions, other than those shown to be necessary as a result of required
 flight tests.
   (d) The commuter category is limited to propeller-driven, multiengine
 airplanes that have a seating configuration excluding pilot seats, of 19 or
 less, and a maximum certificated takeoff weight of 19,000 pounds or less,
 intended for nonacrobatic operation as described in paragraph (a) of this
 section.
   (e) Airplanes may be type certificated in more than one category of this
 part if the requirements of each requested category are met.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-4, 32 FR
 5934, Apr. 14, 1967; Amdt. 23-34, 52 FR 1825, Jan. 15, 1987; 52 FR 34745,
 Sept. 14, 1987]






                               Subpart B--Flight








--------[You Cited: 14 CFR Subpart A thru Subpart B as of Mar. 10, 1994]-------




                              Subpart A--General






 Sec. 23.1  Applicability.

   (a) This part prescribes airworthiness standards for the issue of type
 certificates, and changes to those certificates, for airplanes in the normal,
 utility, acrobatic, and commuter categories.
   (b) Each person who applies under Part 21 for such a certificate or change
 must show compliance with the applicable requirements of this part.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-34, 52 FR
 1825, Jan. 15, 1987]






 Sec. 23.2  Special retroactive requirements.

   (a) Notwithstanding Secs. 21.17 and 21.101 of this chapter and irrespective
 of the type certification basis, each normal, utility, and acrobatic category
 airplane having a passenger seating configuration, excluding pilot seats, of
 nine or less, manufactured after December 12, 1986, or any such foreign
 airplane for entry into the United States must provide a safety belt and
 shoulder harness for each forward- or aft-facing seat which will protect the
 occupant from serious head injury when subjected to the inertia loads
 resulting from the ultimate static load factors prescribed in Sec.
 23.561(b)(2) of this part, or which will provide the occupant protection
 specified in Sec. 23.562 of this part when that section is applicable to the
 airplane. For other seat orientations, the seat/restraint system must be
 designed to provide a level of occupant protection equivalent to that
 provided for forward- or aft-facing seats with a safety belt and shoulder
 harness installed.
   (b) Each shoulder harness installed at a flight crewmember station, as
 required by this section, must allow the crewmember, when seated with the
 safety belt and shoulder harness fastened, to perform all functions necessary
 for flight operations.
   (c) For the purpose of this section, the date of manufacture is:
   (1) The date the inspection acceptance records, or equivalent, reflect that
 the airplane is complete and meets the FAA approved type design data; or
   (2) In the case of a foreign manufactured airplane, the date the foreign
 civil airworthiness authority certifies the airplane is complete and issues
 an original standard airworthiness certificate, or the equivalent in that
 country.

 [Amdt. 23-36, 53 FR 30812, Aug. 15, 1988]






 Sec. 23.3  Airplane categories.

   (a) The normal category is limited to airplanes that have a seating
 configuration, excluding pilot seats, of nine or less, a maximum certificated
 takeoff weight of 12,500 pounds or less, and intended for nonacrobatic
 operation. Nonacrobatic operation includes:
   (1) Any maneuver incident to normal flying;
   (2) Stalls (except whip stalls); and
   (3) Lazy eights, chandelles, and steep turns, in which the angle of bank is
 not more than 60 degrees.
   (b) The utility category is limited to airplanes that have a seating
 configuration, excluding pilot seats, of nine or less, a maximum certificated
 takeoff weight of 12,500 pounds or less, and intended for limited acrobatic
 operation. Airplanes certificated in the utility category may be used in any
 of the operations covered under paragraph (a) of this section and in limited
 acrobatic operations. Limited acrobatic operation includes:
   (1) Spins (if approved for the particular type of airplane); and
   (2) Lazy eights, chandelles, and steep turns, in which the angle of bank is
 more than 60 degrees.
   (c) The acrobatic category is limited to airplanes that have a seating
 configuration, excluding pilot seats, of nine or less, a maximum certificated
 takeoff weight of 12,500 pounds or less, and intended for use without
 restrictions, other than those shown to be necessary as a result of required
 flight tests.
   (d) The commuter category is limited to propeller-driven, multiengine
 airplanes that have a seating configuration excluding pilot seats, of 19 or
 less, and a maximum certificated takeoff weight of 19,000 pounds or less,
 intended for nonacrobatic operation as described in paragraph (a) of this
 section.
   (e) Airplanes may be type certificated in more than one category of this
 part if the requirements of each requested category are met.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-4, 32 FR
 5934, Apr. 14, 1967; Amdt. 23-34, 52 FR 1825, Jan. 15, 1987; 52 FR 34745,
 Sept. 14, 1987]






                               Subpart B--Flight






                                    General






 Sec. 23.21  Proof of compliance.

   (a) Each requirement of this subpart must be met at each appropriate
 combination of weight and center of gravity within the range of loading
 conditions for which certification is requested. This must be shown--
   (1) By tests upon an airplane of the type for which certification is
 requested, or by calculations based on, and equal in accuracy to, the results
 of testing; and
   (2) By systematic investigation of each probable combination of weight and
 center of gravity, if compliance cannot be reasonably inferred from
 combinations investigated.
   (b) The following general tolerances are allowed during flight testing.
 However, greater tolerances may be allowed in particular tests:

                           Item                      Tolerance

             Weight                             +5%, -10%.
             Critical items affected by weight  +5%, -1%.
             C.G                                +/-7% total travel.






 Sec. 23.23   Load distribution limits.

   (a) Ranges of weights and centers of gravity within which the airplane may
 be safely operated must be established. If a weight and center of gravity
 combination is allowable only within certain lateral load distribution limits
 that could be inadvertently exceeded, these limits must be established for
 the corresponding weight and center of gravity combinations.
   (b) The load distribution limits may not exceed any of the following:
   (1) The selected limits;
   (2) The limits at which the structure is proven; or
   (3) The limits at which compliance with each applicable flight requirement
 of this subpart is shown.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-17, 41 FR
 55463, Dec. 20, 1976; Amdt. 23-45, 58 FR 42156, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.25  Weight limits.

   (a) Maximum weight. The maximum weight is the highest weight at which
 compliance with each applicable requirement of this Part (other than those
 complied with at the design landing weight) is shown. In addition, for
 commuter category airplanes, the applicant must establish a maximum zero fuel
 weight. The maximum weight must be established so that it is--
   (1) Not more than--
   (i) The highest weight selected by the applicant;
   (ii) The design maximum weight, which is the highest weight at which
 compliance with each applicable structural loading condition of this part
 (other than those complied with at the design landing weight) is shown; or
   (iii) The highest weight at which compliance with each applicable flight
 requirement is shown, except for airplanes equipped with standby power rocket
 engines, in which case it is the highest weight established in accordance
 with Appendix E of this part; or
   (2) Not less than the weight with--
   (i) Each seat occupied, assuming a weight of 170 pounds for each occupant
 for normal and commuter category airplanes, and 190 pounds for utility and
 acrobatic category airplanes, except that seats other than pilot seats may be
 placarded for a lesser weight; and
   (A) Oil at full capacity, and
   (B) At least enough fuel for maximum continuous power operation of at least
 30 minutes for day-VFR approved airplanes and at least 45 minutes for night-
 VFR and IFR approved airplanes; or
   (ii) The required minimum crew, and fuel and oil to full tank capacity.
   (b) Minimum weight. The minimum weight (the lowest weight at which
 compliance with each applicable requirement of this part is shown) must be
 established so that it is not more than the sum of--
   (1) The empty weight determined under Sec. 23.29;
   (2) The weight of the required minimum crew (assuming a weight of 170
 pounds for each crewmember); and
   (3) The weight of--
   (i) For turbojet powered airplanes, 5 percent of the total fuel capacity of
 that particular fuel tank arrangement under investigation, and
   (ii) For other airplanes, the fuel necessary for one-half hour of operation
 at maximum continuous power.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13086, Aug. 13, 1969; Amdt. 23-21, 43 FR 2317, Jan. 16, 1978; Amdt. 23-34, 52
 FR 1825, Jan. 15, 1987; Amdt. No. 23-45, 58 FR 42156, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.29  Empty weight and corresponding center of gravity.

   (a) The empty weight and corresponding center of gravity must be determined
 by weighing the airplane with--
   (1) Fixed ballast;
   (2) Unusable fuel determined under Sec. 23.959; and
   (3) Full operating fluids, including--
   (i) Oil;
   (ii) Hydraulic fluid; and
   (iii) Other fluids required for normal operation of airplane systems,
 except potable water, lavatory precharge water, and water intended for
 injection in the engines.
   (b) The condition of the airplane at the time of determining empty weight
 must be one that is well defined and can be easily repeated.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. 23-21, 43 FR 2317, Jan. 16, 1978]






 Sec. 23.31  Removable ballast.

   Removable ballast may be used in showing compliance with the flight
 requirements of this subpart, if--
   (a) The place for carrying ballast is properly designed and installed, and
 is marked under Sec. 23.1557; and
   (b) Instructions are included in the airplane flight manual, approved
 manual material, or markings and placards, for the proper placement of the
 removable ballast under each loading condition for which removable ballast is
 necessary.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. 23-13, 37 FR 20023, Sept. 23, 1972]






 Sec. 23.33  Propeller speed and pitch limits.

   (a) General. The propeller speed and pitch must be limited to values that
 will assure safe operation under normal operating conditions.
   (b) Propellers not controllable in flight. For each propeller whose pitch
 cannot be controlled in flight--
   (1) During takeoff and initial climb at Vy, the propeller must limit the
 engine r.p.m., at full throttle or at maximum allowable takeoff manifold
 pressure, to a speed not greater than the maximum allowable takeoff r.p.m.;
 and
   (2) During a closed throttle glide at the placarded "never-exceed speed",
 the propeller may not cause an engine speed above 110 percent of maximum
 continuous speed.
   (c) Controllable pitch propellers without constant speed controls. Each
 propeller that can be controlled in flight, but that does not have constant
 speed controls, must have a means to limit the pitch range so that--
   (1) The lowest possible pitch allows compliance with paragraph (b)(1) of
 this section; and
   (2) The highest possible pitch allows compliance with paragraph (b)(2) of
 this section.
   (d) Controllable pitch propellers with constant speed controls. Each
 controllable pitch propeller with constant speed controls must have--
   (1) With the governor in operation, a means at the governor to limit the
 maximum engine speed to the maximum allowable takeoff r.p.m.; and
   (2) With the governor inoperative, the propeller blades at the lowest
 possible pitch, with takeoff power, the airplane stationary, and no wind,
 either--
   (i) A means to limit the maximum engine speed to 103 percent of the maximum
 allowable takeoff r.p.m., or
   (ii) For an engine with an approved overspeed, a means to limit the maximum
 engine and propeller speed to not more than the maximum approved overspeed.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965; Amdt.
 No. 23-45, 58 FR 42156, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






                                  Performance






 Sec. 23.45  General.

   (a) Unless otherwise prescribed, the performance requirements of this
 subpart must be met for still air; and
   (1) Standard atmospheric conditions for normal, utility, and acrobatic
 category airplanes; or
   (2) Ambient atmospheric conditions for commuter category airplanes.
   (b) The performance data must correspond to the propulsive power or thrust
 available under the particular ambient atmospheric conditions, the particular
 flight condition, and the relative humidity specified in paragraph (d) of
 this section.
   (c) The available propulsive thrust must correspond to engine power or
 thrust, not exceeding the approved power or thrust, less--
   (1) Installation losses; and
   (2) The power or equivalent thrust absorbed by the accessories and services
 appropriate to the particular ambient atmospheric conditions and the
 particular flight condition.
   (d) The performance, as affected by engine power or thrust, must be based
 on a relative humidity of--
   (1) 80 percent, at and below standard temperature; and
   (2) 34 percent, at and above standard temperature, plus 50 deg.F.
   (3) Between the two temperatures listed in paragraphs (d)(1) and (d)(2) of
 this section, the relative humidity must vary linearly.
   (e) For commuter category airplanes, the following also apply:
   (1) Unless otherwise prescribed, the applicant must select the takeoff, en
 route, approach, and landing configurations for the airplane;
   (2) The airplane configuration may vary with weight, altitude, and
 temperature, to the extent they are compatible with the operating procedures
 required by paragraph (e)(3) of this section;
   (3) Unless otherwise prescribed, in determining the critical-engine-
 inoperative takeoff performance, takeoff flight path, the accelerate-stop
 distance, takeoff distance, and landing distance, changes in the airplane's
 configuration, speed, power, and thrust must be made in accordance with
 procedures established by the applicant for operation in service;
   (4) Procedures for the execution of missed approaches and balked landings
 associated with the conditions prescribed in Secs. 23.67(e)(3) and 23.77(c)
 must be established; and
   (5) The procedures established under paragraphs (e)(3) and (e)(4) of this
 section must--
   (i) Be able to be consistently executed by a crew of average skill;
   (ii) Use methods or devices that are safe and reliable; and
   (iii) Include allowance for any reasonably expected time delays in the
 execution of the procedures.

 [Amdt. 23-21, 43 FR 2317, Jan. 16, 1978, as amended by Amdt. 23-34, 52 FR
 1826, Jan. 15, 1987; Amdt. No. 23-45, 58 FR 42156, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.49  Stalling speed.

   (a) Vs 0 is the stalling speed, if obtainable, or the minimum steady speed,
 in knots (CAS), at which the airplane is controllable, with the--
   (1) Applicable power or thrust condition set forth in paragraph (e) of this
 section;
   (2) Propellers in the takeoff position;
   (3) Landing gear extended;
   (4) Wing flaps in the landing position;
   (5) Cowl flaps closed;
   (6) Center of gravity in the most unfavorable position within the allowable
 landing range; and
   (7) Weight used when VS0 is being used as a factor to determine compliance
 with a required performance standard.
   (b) Except as provided in Sec. 23.49(c), VS0 at maximum weight may not
 exceed 61 knots for--
   (1) Single-engine airplanes; and
   (2) Multiengine airplanes of 6,000 pounds or less maximum weight that
 cannot meet the minimum rate of climb specified in Sec. 23.67(b) with the
 critical engine inoperative.
   (c) All single-engine airplanes, and those multiengine airplanes of 6,000
 pounds or less maximum weight with a VS0 of more than 61 knots that do not
 meet the requirements of Sec. 23.67(b)(2)(i), must comply with Sec.
 23.562(d).
   (d) VS1 is the calibrated stalling speed, if obtainable, or the minimum
 steady speed, in knots, at which the airplane is controllable, with the--
   (1) Applicable power or thrust condition set forth in paragraph (e) of this
 section;
   (2) Propellers in the takeoff position;
   (3) Airplane in the condition existing in the test in which VS1 is being
 used; and
   (4) Weight used when VS1 is being used as a factor to determine compliance
 with a required performance standard.
   (e) VS0 and VS1 must be determined by flight tests, using the procedure
 specified in Sec. 23.201.
   (f) The following power or thrust conditions must be used to meet the
 requirements of this section:
   (1) For reciprocating engine-powered airplanes, engines idling, throttles
 closed or at not more than the power necessary for zero thrust at a speed not
 more than 110 percent of the stalling speed.
   (2) For turbine engine-powered airplanes, the propulsive thrust may not be
 greater than zero at the stalling speed, or, if the resultant thrust has no
 appreciable effect on the stalling speed, with engines idling and throttles
 closed.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13086, Aug. 13, 1969; Amdt. 23-21, 43 FR 2317, Jan. 16, 1978; Amdt. 23-44,
 58 FR 38639, July 19, 1993]

 *****************************************************************************


 58 FR 38634, No. 136, July 19, 1993

 SUMMARY: This final rule amends the stalling speed requirements applicable to
 single-engine airplanes and to certain multiengine small airplanes of less
 than 6,000 pounds maximum weight. The rule permits those airplanes to have a
 stall speed greater than 61 knots, provided they meet certain additional
 occupant protection standards. These changes are needed to permit the design
 and type certification of higher performance airplanes with increased cruise
 speeds and better specific fuel consumption. The amendments are intended to
 achieve the benefits of certificating higher performance airplanes while
 affording their occupants the same level of protection in an emergency
 landing that is presently provided by airplanes with a 61-knot stall speed.

 EFFECTIVE DATE: August 18, 1993.

 *****************************************************************************






 Sec. 23.51  Takeoff.

   (a) For each airplane (except a skiplane for which landplane takeoff data
 has been determined under this paragraph and furnished in the Airplane Flight
 Manual) the distance required to takeoff and climb over a 50-foot obstacle
 must be determined with--
   (1) The engines operating within approved operating limitations; and
   (2) The cowl flaps in the normal takeoff position.
   (b) The starting point for measuring seaplane and amphibian takeoff
 distance may be the point at which a speed of not more than three knots is
 reached.
   (c) Takeoffs made to determine the data required by this section may not
 require exceptional piloting skill or exceptionally favorable conditions.
   (d) For commuter category airplanes, takeoff performance and data as
 required by Secs. 23.53 through 23.59 must be determined and included in the
 Airplane Flight Manual--
   (1) For each weight, altitude, and ambient temperature within the
 operational limits selected by the applicant;
   (2) For the selected configuration for takeoff;
   (3) For the most unfavorable center of gravity position;
   (4) With the operating engine within approved operating limitations;
   (5) On a smooth, dry, hard surface runway; and
   (6) Corrected for the following operational correction factors:
   (i) Not more than 50 percent of nominal wind components along the takeoff
 path opposite to the direction of takeoff and not less than 150 percent of
 nominal wind components along the takeoff path in the direction of takeoff;
 and
   (ii) Effective runway gradients.

 [Amdt. 23-21, 43 FR 2317, Jan. 16, 1978, as amended by Amdt. 23-34, 52 FR
 1826, Jan. 15, 1987]






 Sec. 23.53   Takeoff speeds.

   (a) For multiengine normal, utility, and acrobatic category airplanes, the
 rotation speed, VR, may not be less than VMC determined in accordance with
 Sec. 23.149.
   (b) Each normal, utility, and acrobatic category airplane, upon reaching a
 height of 50 feet above the takeoff surface level, must have reached a speed
 of not less than the following:
   (1) For multiengine airplanes, the higher of--
   (i) 1.1 VMC; or
   (ii) Any lesser speed, not less than 1.2 VS1, that is shown to be safe for
 continued flight or land-back, if applicable, under all conditions, including
 turbulence and complete failure of the critical engine.
   (2) For single-engine airplanes, any speed, not less than 1.2 VS1, that is
 shown to be safe under all conditions, including turbulence and complete
 engine failure.
   (c) For commuter category airplanes, the following apply:
   (1) The takeoff decision speed, V1, is the calibrated airspeed on the
 ground at which, as a result of engine failure or other reasons, the pilot is
 assumed to have made a decision to continue or discontinue the takeoff. The
 takeoff decision speed, V1, must be selected by the applicant but may not be
 less than the greater of the following:
   (i) 1.10 VS1;
   (ii) 1.10 VMC established in accordance with Sec. 23.149;
   (iii) A speed at which the airplane can be rotated for takeoff and shown to
 be adequate to safely continue the takeoff, using normal piloting skill, when
 the critical engine is suddenly made inoperative; or
   (iv) VEF plus the speed gained with the critcial engine inoperative during
 the time interval between the instant that the critical engine is failed and
 the instant at which the pilot recognizes and reacts to the engine failure as
 indicated by the pilot's application of the first retarding means during the
 accelerate-stop determination of Sec. 23.55.
   (2) The takeoff safety speed, V2, in terms of calibrated airspeed, must be
 selected by the applicant so as to allow the gradient of climb required in
 Sec. 23.67 but must not be less than V1 or less than 1.2VS1.
   (3) The critical engine failure speed, VEF, is the calibrated airspeed at
 which the critical engine is assumed to fail. VEF must be selected by the
 applicant but not less than VMC determined in accordance with Sec. 23.149.
   (4) The rotation speed, VR in terms of calibrated airspeed, must be
 selected by the applicant and may not be less than the greater of the
 following:
   (i) V1; or
   (ii) The speed determined in accordance with Sec. 23.57(c) that allows
 attaining the initial climb out speed, V2, before reaching a height of 35
 feet above the takeoff surface.
   (5) For any given set of conditions, such as weight, altitude,
 configuration, and temperature, a single value of VR must be used to show
 compliance with both the one-engine-inoperative takeoff and all-engines-
 operating takeoff requirements:
   (i) One-engine-inoperative takeoff determined in accordance with Sec.
 23.57; and
   (ii) All-engines-operating takeoff determined in accordance with Sec.
 23.59.
   (6) The one-engine-inoperative takeoff distance, using a normal rotation
 rate at a speed of 5 knots less than VR established in accordance with
 paragraphs (c)(4) and (5) of this section, must be shown not to exceed the
 corresponding one-engine-inoperative takeoff distance determined in
 accordance with Secs. 23.57 and 23.59 using the established VR. The take off
 distance determined in accordance with Sec. 23.59 and the takeoff must be
 safely continued from the point at which the airplane is 35 feet above the
 takeoff surface at a speed not less than 5 knots less than the established V2
 speed.
   (7) The applicant must show, with all engines operating, that marked
 increases in the scheduled takeoff distances determined in accordance with
 Sec. 23.59 do not result from over-rotation of the airplane and out-of-trim
 conditions.

 [Amdt. 23-34, 52 FR 1826, Jan. 15, 1987; 52 FR 34745, Sept. 14, 1987, as
 amended by Amdt. 23-45, 58 FR 42156, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.55  Accelerate-stop distance.

   For each commuter category airplane, the accelerate-stop distance must be
 determined as follows:
   (a) The accelerate-stop distance is the sum of the distances necessary to--
   (1) Accelerate the airplane from a standing start to V1; and
   (2) Come to a full stop from the point at which V1 is reached assuming that
 in the case of engine failure, the pilot has decided to stop as indicated by
 application of the first retarding means at the speed V1.
   (b) Means other than wheel brakes may be used to determine the accelerate-
 stop distance if that means is available with the critical engine inoperative
 and if that means--
   (1) Is safe and reliable;
   (2) Is used so that consistent results can be expected under normal
 operating conditions; and
   (3) Is such that exceptional skill is not required to control the airplane.

 [Amdt. 23-34, 52 FR 1826, Jan. 15, 1987]






 Sec. 23.57  Takeoff path.

   For each commuter category airplane, the takeoff path is as follows:
   (a) The takeoff path extends from a standing start to a point in the
 takeoff at which the airplane is 1,500 feet above the takeoff surface or at
 which the transition from the takeoff to the en route configuration is
 completed, whichever point is higher; and
   (1) The takeoff path must be based on the procedures prescribed in Sec.
 23.45;
   (2) The airplane must be accelerated on the ground to VEF at which point
 the critical engine must be made inoperative and remain inoperative for the
 rest of the takeoff; and
   (3) After reaching VEF, the airplane must be accelerated to V2.
   (b) During the acceleration to speed V2, the nose gear may be raised off
 the ground at a speed not less than VR. However, landing gear retraction may
 not be initiated until the airplane is airborne.
   (c) During the takeoff path determination, in accordance with paragraphs
 (a) and (b) of this section--
   (1) The slope of the airborne part of the takeoff path must be positive at
 each point;
   (2) The airplane must reach V2 before it is 35 feet above the takeoff
 surface, and must continue at a speed as close as practical to, but not less
 than V2, until it is 400 feet above the takeoff surface;
   (3) At each point along the takeoff path, starting at the point at which
 the airplane reaches 400 feet above the takeoff surface, the available
 gradient of climb may not be less than--
   (i) 1.2 percent for two-engine airplanes;
   (ii) 1.5 percent for three-engine airplanes;
   (iii) 1.7 percent for four-engine airplanes; and
   (4) Except for gear retraction and automatic propeller feathering, the
 airplane configuration may not be changed, and no change in power or thrust
 that requires action by the pilot may be made, until the airplane is 400 feet
 above the takeoff surface.
   (d) The takeoff path must be determined by a continuous demonstrated
 takeoff or by synthesis from segments. If the takeoff path is determined by
 the segmental method--
   (1) The segments must be clearly defined and must be related to the
 distinct changes in the configuration, power or thrust, and speed;
   (2) The weight of the airplane, the configuration, and the power or thrust
 must be constant throughout each segment and must correspond to the most
 critical condition prevailing in the segment;
   (3) The flight path must be based on the airplane's performance without
 ground effect;
   (4) The takeoff path data must be checked by continuous demonstrated
 takeoffs up to the point at which the airplane is out of ground effect and
 its speed is stabilized to ensure that the path is conservative relative to
 the continuous path; and
   (5) The airplane is considered to be out of the ground effect when it
 reaches a height equal to its wing span.

 [Amdt. 23-34, 52 FR 1827, Jan. 15, 1987]






 Sec. 23.59  Takeoff distance and takeoff run.

   For each commuter category airplane--
   (a) Takeoff distance is the greater of--
   (1) The horizontal distance along the takeoff path from the start of the
 takeoff to the point at which the airplane is 35 feet above the takeoff
 surface as determined under Sec. 23.57; or
   (2) With all engines operating, 115 percent of the horizontal distance
 along the takeoff path, with all engines operating, from the start of the
 takeoff to the point at which the airplane is 35 feet above the takeoff
 surface, as determined by a procedure consistent with Sec. 23.57.
   (b) If the takeoff distance includes a clearway, the takeoff run is the
 greater of--
   (1) The horizontal distance along the takeoff path from the start of the
 takeoff to a point equidistant between the point at which VLOF is reached and
 the point at which the airplane is 35 feet above the takeoff surface as
 determined under Sec. 23.57; or
   (2) With all engines operating, 115 percent of the horizontal distance
 along the takeoff path, with all engines operating, from the start of the
 takeoff to a point equidistant between the point at which VLOF is reached and
 the point at which the airplane is 35 feet above the takeoff surface
 determined by a procedure consistent with Sec. 23.57.

 [Amdt. 23-34, 52 FR 1827, Jan. 15, 1987]






 Sec. 23.61  Takeoff flight path.

   For each commuter category airplane, the takeoff flight path must be
 determined as follows:
   (a) The takeoff flight path begins 35 feet above the takeoff surface at the
 end of the takeoff distance determined in accordance with Sec. 23.59.
   (b) The net takeoff flight path data must be determined so that they
 represent the actual takeoff flight paths, as determined in accordance with
 Sec. 23.57 and with paragraph (a) of this section, reduced at each point by a
 gradient of climb equal to--
   (1) 0.8 percent for two-engine airplanes;
   (2) 0.9 percent for three-engine airplanes; and
   (3) 1.0 percent for four-engine airplanes.
   (c) The prescribed reduction in climb gradient may be applied as an
 equivalent reduction in acceleration along that part of the takeoff flight
 path at which the airplane is accelerated in level flight.

 [Amdt. 23-34, 52 FR 1827, Jan. 15, 1987]






 Sec. 23.65  Climb: All engines operating.

 Sec. 23.65  Climb: All engines operating.

   (a) Each airplane must have a steady angle of climb at sea level of at
 least 1:12 for landplanes or 1:15 for seaplanes and amphibians with--
   (1) A speed not less than 1.2 VS1;
   (2) Not more than maximum continuous power on each engine;
   (3) The landing gear retracted;
   (4) The wing flaps in the takeoff position; and
   (5) The cowl flaps or other means for controlling the engine cooling air
 supply in the position used in the cooling tests required by Secs. 23.1041
 through 23.1047.
   (b) Each airplane with engines for which the takeoff and maximum continuous
 power ratings are identical and that has fixed-pitch, two-position, or
 similar propellers, may use a lower propeller pitch setting than that allowed
 by Sec. 23.33 to obtain rated engine r.p.m. at Vx, if--
   (1) The airplane shows marginal performance (such as when it can meet the
 rate of climb requirements of paragraph (a) of this section but has
 difficulty in meeting the angle of climb requirements of paragraph (a) of
 this section or of Sec. 23.77); and
   (2) Acceptable engine cooling is shown at the lower speed associated with
 the best angle of climb.
   (c) Each turbine engine-powered airplane must be able to maintain a steady
 gradient of climb of at least 4 percent at a pressure altitude of 5,000 feet
 and a temperature of 81 degrees F (standard temperature plus 40 degree F)
 with the airplane in the configuration prescribed in paragraph (a) of this
 section.
   (d) In addition for commuter category airplanes, performance data must be
 determined for variations in weight, altitude, and temperature at the most
 critical center of gravity for which approval is requested.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13086, Aug. 13, 1969; Amdt. 23-21, 43 FR 2317, Jan. 16, 1978; Amdt. 23-34, 52
 FR 1827, Jan. 15, 1987; Amdt. 23-24, 52 FR 34745, Sept. 14, 1987; Amdt. No.
 23-45, 58 FR 42156, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.67  Climb: one engine inoperative.

   (a) For normal, utility, and acrobatic category, reciprocating engine-
 powered multiengine airplanes, one-engine-inoperative climb gradients must be
 determined with the--
   (1) Critical engine inoperative, and its propeller in the minimum drag
 position;
   (2) Remaining engines at not more than maximum continuous power or thrust;
   (3) Landing gear retracted;
   (4) Wing flaps in the most favorable position; and
   (5) Means for controlling the engine cooling air supply in the position
 used in the engine cooling tests required by Secs. 23.1041 through 23.1047.
   (b) For normal, utility, and acrobatic category reciprocating engine-
 powered multiengine airplanes, the following apply:
   (1) Each airplane of more than 6,000 pounds maximum weight must be able to
 maintain a steady climb gradient of at least 1.5 percent at a pressure
 altitude of 5,000 feet at a speed not less than 1.2 VS1 and at standard
 temperature (41 deg.F) with the airplane in the configuration prescribed in
 paragraph (a) of this section.
   (2) For each airplane of 6,000 pounds or less maximum weight, the following
 apply:
   (i) Each airplane that meets the requirements of Sec. 23.562(d), or that
 has a VS0 of 61 knots or less, must have its steady climb gradient determined
 at a pressure altitude of 5,000 feet at a speed of not less than 1.2 VS1, and
 at standard temperature (41 deg.F), with the airplane in the configuration
 prescribed in paragraph (a) of this section.
   (ii) Except for those airplanes that meet the requirements prescribed in
 Sec. 23.562(d), each airplane with a VS0 of more than 61 knots must be able
 to maintain the steady climb gradient prescribed in paragraph (b)(1) of this
 section.
   (c) For normal, utility, and acrobatic category turbine engine-powered
 multiengine airplanes the following apply:
   (1) The steady climb gradient must be determined at each weight, altitude,
 and ambient temperature within the operational limits established by the
 applicant, with the airplane in the configuration prescribed in paragraph (a)
 of this section.
   (2) Each airplane must be able to maintain at least the following climb
 gradients with the airplane in the configuration prescribed in paragraph (a)
 of this section:
   (i) 1.5 percent at a pressure altitude of 5,000 feet at a speed not less
 than 1.2 VS1, and at standard temperature (41 deg.F); and
   (ii) 0.75 percent at a pressure altitude of 5,000 feet at a speed not less
 than 1.2 VS1 and 81 deg.F (standard temperature plus 40 deg.F).
   (3) The minimum climb gradient specified in paragraphs (c)(2) (i) and (ii)
 of this section must vary linearly between 41 deg.F and 81 deg.F and must
 change at the same rate up to the maximum operating temperature approved for
 the airplane.
   (d) For all multiengine airplanes, the speed for best rate of climb with
 one engine inoperative must be determined.
   (e) For commuter category airplanes, the following apply:
   (1) Takeoff climb: The maximum weight at which the airplane meets the
 minimum climb performance specified in paragraphs (e)(1) (i) and (ii) of this
 section must be determined for each altitude and ambient temperature within
 the operating limitations established for the airplane, out of ground effect
 in free air, with the airplane in the takeoff configuration, with the most
 critical center of gravity, the critical engine inoperative, the remaining
 engines at the maximum takeoff power or thrust, and the propeller of the
 inoperative engine windmilling with the propeller controls in the normal
 position, except that, if an approved automatic propeller feathering system
 is installed, the propeller may be in the feathered position:
   (i) Takeoff, landing gear extended. The minimum steady gradient of climb
 between the lift-off speed, VLOF, and until the landing gear is retracted
 must be measurably positive for two-engine airplanes, not less than 0.3
 percent for three-engine airplanes, or 0.5 percent for four-engine airplanes
 at all points along the flight path; and
   (ii) Takeoff, landing gear retracted. The minimum steady gradient of climb
 must not be less than 2 percent for two-engine airplanes, 2.3 percent for
 three-engine airplanes, and 2.6 percent for four-engine airplanes at the
 speed V2, until the airplane is 400 feet above the takeoff surface. For
 airplanes with fixed landing gear, this requirement must be met with the
 landing gear extended.
   (2) En route climb: The maximum weight must be determined for each altitude
 and ambient temperature within the operational limits established for the
 airplane, at which the steady gradient of climb is not less than 1.2 percent
 for two-engine airplanes, 1.5 percent for three-engine airplanes, and 1.7
 percent for four-engine airplanes at an height of 1,500 feet above the
 takeoff surface, with the airplane in the en route configuration, the
 critical engine inoperative, the remaining engine at the maximum continuous
 power or thrust, and the most unfavorable center of gravity.
   (3) Approach: In the approach configuration corresponding to the normal
 all-engines-operating procedure in which VS1 for this configuration does not
 exceed 110 percent of the VS1 for the related landing configuration, the
 steady gradient of climb may not be less than 2.1 percent for two-engine
 airplanes, 2.4 percent for three-engine airplanes, and 2.7 percent for four-
 engine airplanes, with--
   (i) The critical engine inoperative and the remaining engines at the
 available takeoff power or thrust;
   (ii) The maximum landing weight; and
   (iii) A climb speed established in connection with the normal landing
 procedures but not exceeding 1.5 VS1.

 [Amdt. 23-21, 43 FR 2317, Jan. 16, 1978, as amended by Amdt. 23-34, 52 FR
 1827, Jan. 15, 1987; Amdt. 23-24, 52 FR 34745, Sept. 14, 1987; Amdt. 23-39,
 55 FR 18575, May 2, 1990; Amdt. 23-42, 56 FR 351, Jan. 3, 1991; Amdt. 23-44,
 58 FR 38639, July 19, 1993]

 *****************************************************************************


 58 FR 38634, No. 136, July 19, 1993

 SUMMARY: This final rule amends the stalling speed requirements applicable to
 single-engine airplanes and to certain multiengine small airplanes of less
 than 6,000 pounds maximum weight. The rule permits those airplanes to have a
 stall speed greater than 61 knots, provided they meet certain additional
 occupant protection standards. These changes are needed to permit the design
 and type certification of higher performance airplanes with increased cruise
 speeds and better specific fuel consumption. The amendments are intended to
 achieve the benefits of certificating higher performance airplanes while
 affording their occupants the same level of protection in an emergency
 landing that is presently provided by airplanes with a 61-knot stall speed.

 EFFECTIVE DATE: August 18, 1993.

 *****************************************************************************






 Sec. 23.75  Landing.

   For airplanes (except skiplanes for which landplane landing data have been
 determined under this section and furnished in the Airplane Flight Manual),
 the horizontal distance necessary to land and come to a complete stop (or to
 a speed of approximately 3 knots for water landings of seaplanes and
 amphibians) from a point 50 feet above the landing surface must be determined
 as follows:
   (a) A steady approach with a calibrated airspeed of not less than 1.3 VS1
 must be maintained down to the 50-foot height and--
   (1) The steady approach must be at a gradient of descent not greater than
 5.2 percent (3 degrees) down to the 50-foot height.
   (2) In addition, an applicant may demonstrate by tests that a maximum
 steady approach gradient steeper than 5.2 percent, down to the 50-foot
 height, is safe. The gradient must be established as an operating limitation
 and the information necessary to display the gradient must be available to
 the pilot by an appropriate instrument.
   (b) The landing may not require more than average piloting skill when
 landing during the atmospheric conditions expected to be encountered in
 service, including crosswinds and turbulence.
   (c) The landing must be made without excessive vertical acceleration or
 tendency to bounce, nose over, ground loop, porpoise, or water loop.
   (d) It must be shown that a safe transition to the balked landing
 conditions of Sec. 23.77 can be made from the conditions that exist at the
 50-foot height.
   (e) The pressures on the wheel braking system may not exceed those
 specified by the brake manufacturer.
   (f) Means other than wheel brakes may be used if that means--
   (1) Is safe and reliable;
   (2) Is used so that consistent results can be expected in service; and
   (3) Is such that no more than average skill is required to control the
 airplane.
   (g) If any device is used that depends on the operation of any engine, and
 the landing distance would be increased when a landing is made with that
 engine inoperative, the landing distance must be determined with that engine
 inoperative unless the use of other compensating means will result in a
 landing distance not more than that with each engine operating.
   (h) In addition, for commuter category airplanes, the following apply:
   (1) The landing distance must be determined for standard temperatures at
 each weight, altitude, and wind condition within the operational limits
 established by the applicant;
   (2) A steady gliding approach, or a steady approach at a gradient of
 descent not greater than 5.2 percent (3 deg.), at a calibrated airspeed not
 less than 1.3VS1 must be maintained down to the 50-foot height; and
   (3) The landing distance data must include correction factors for not more
 than 50 percent of the nominal wind components along the landing path
 opposite to the direction of landing and not less than 150 percent of the
 nominal wind components along the landing path in the direction of landing.

 [Amdt. 23-21, 43 FR 2318, Jan. 16, 1978, as amended by Amdt. 23-34, 52 FR
 1828, Jan. 15, 1987; Amdt. 23-42, 56 FR 351, Jan. 3, 1991]

 *****************************************************************************


 56 FR 344, No. 2, Jan. 3, 1991

   SUMMARY: This final rule upgrades the airworthiness standards for normal,
 utility, acrobatic, and commuter category airplanes. This amendment provides
 airworthiness standards for advancements in technology being incorporated in
 current designs, permits type certification of spin resistant airplanes, and
 reduces the regulatory burden in showing compliance with some of the
 requirements for the design and type certification of small airplanes. These
 new and amended airworthiness standards also result in the need for new
 definitions. As a result, new definitions are added.

   DATES: February 4, 1991.

 *****************************************************************************






 Sec. 23.77  Balked landing.

   (a) For balked landings, each normal, utility, and acrobatic category
 airplane must be able to maintain a steady angle of climb at sea level of at
 least 1:30 with--
   (1) Takeoff power on each engine;
   (2) The landing gear extended; and
   (3) The wing flaps in the landing position, except that if the flaps may
 safely be retracted in two seconds or less without loss of altitude and
 without sudden changes of angle of attack or exceptional piloting skill, they
 may be retracted.
   (b) Each normal, utility, and acrobatic category turbine engine-powered
 airplane must be able to maintain a steady rate of climb of at least zero at
 a pressure altitude of 5,000 feet at 81 degrees F (standard temperature plus
 40 degrees F), with the airplanes in the configuration prescribed in
 paragraph (a) of this section.
   (c) For each commuter category airplane, with all engines operating, the
 maximum weight must be determined with the airplane in the landing
 configuration for each altitude and ambient temperature within the
 operational limits established for the airplane, with the most unfavorable
 center of gravity and out-of-ground effect in free air, at which the steady
 gradient of climb will not be less than 3.3 percent with--
   (1) The engines at the power or thrust that is available 8 seconds after
 initiation of movement of the power or thrust controls from the minimum
 flight-idle position to the takeoff position.
   (2) A climb speed not greater than the approach speed established under
 Sec. 23.75 and not less than the greater of 1.05 VMC or 1.10VS1.

 [Amdt. 23-21, 43 FR 2318, Jan. 16, 1978, as amended by Amdt. 23-34, 52 FR
 1828, Jan. 15, 1987; Amdt. 23-24, 52 FR 34745, Sept. 14, 1987]






                            Flight Characteristics






 Sec. 23.141  General.

   The airplane must meet the requirements of Secs. 23.143 through 23.253 at
 all practical loading conditions and operating altitudes for which
 certification has been requested, not exceeding the maximum operating
 altitude established under Sec. 23.1527, and without requiring exceptional
 piloting skill, alertness, or strength.

 [Amdt. 23-45, 58 FR 42156, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






                      Controllability and Maneuverability






 Sec. 23.143  General.

   (a) The airplane must be safely controllable and maneuverable during--
   (1) Takeoff;
   (2) Climb;
   (3) Level flight;
   (4) Descent; and
   (5) Landing (power on and power off with the wing flaps extended and
 retracted).
   (b) It must be possible to make a smooth transition from one flight
 condition to another (including turns and slips) without danger of exceeding
 the limit load factor, under any probable operating condition (including, for
 multiengine airplanes, those conditions normally encountered in the sudden
 failure of any engine).
   (c) If marginal conditions exist with regard to required pilot strength,
 the "strength of pilots" limits must be shown by quantitative tests. In no
 case may the limits exceed those prescribed in the following table:

                 Values in pounds of force as
                applied to the stick, control
                   wheel, or rudder pedals      Pitch  Roll  Yaw

                (a) For temporary application:
                 Stick                             60    30
                 Wheel (Two hands on rim)          75    60
                 Wheel (One hand on rim)           50
                 Rudder Pedal                                150
                (b) For prolonged application      10     5   20

 [Doc. No, 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-14, 38 FR
 31819, Nov. 19, 1973; Amdt. 23-17, 41 FR 55464, Dec. 20, 1976; Amdt. 23-45,
 58 FR 42157, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.145  Longitudinal control.

   (a) With the airplane as nearly as possible in trim at 1.3 VS1, it must be
 possible, at speeds below the trim speed, to pitch the nose downward so that
 the rate of increase in airspeed allows prompt acceleration to the trim speed
 with--
   (1) Maximum continuous power on each engine;
   (2) Power off; and
   (3) Wing flap and landing gear--
   (i) retracted, and
   (ii) extended.
   (b) No change in trim or exertion of more control force, as specified in
 Sec. 23.143(c), than can be readily applied with one hand for a short period
 of time may be required for the following maneuvers:
   (1) With the landing gear extended, the flaps retracted, and the airplanes
 as nearly as possible in trim at 1.4 VS1, extend the flaps as rapidly as
 possible and allow the airspeed to transition from 1.4 VS1 to 1.4 VSO:
   (i) With power off; and
   (ii) With the power necessary to maintain level flight in the initial
 condition.
   (2) With the landing gear and flaps extended--
   (i) With power off and the airplane as nearly as possible in trim at 1.3
 VSO, quickly apply takeoff power or thrust and retract the flaps as rapidly
 as possible to the recommended go-around setting while attaining and
 maintaining, as a minimum, the speed used to show compliance with Sec. 23.77.
 Retract the gear when positive rate of climb is established; and
   (ii) With power off and in level flight at 1.1VSO, and the airplane as
 nearly as possible in trim at 1.2 VSO, it must be possible to maintain
 approximately level flight while retracting the flaps as rapidly as possible
 with simultaneous application of not more than maximum continuous power. If
 gated flap positions are provided, the airplane may be retrimmed between each
 stage of retraction, and the airplane may accelerate to a speed that is 1.1
 times the minimum steady flight speed obtained for the flap gate position.
   (3) With maximum takeoff power, landing gear retracted, flaps in the
 takeoff position, and the airplane as nearly as possible in trim at VFE,
 appropriate to the takeoff flap position, retract the flaps as rapidly as
 possible while maintaining constant speed.
   (4) With power off, flaps and landing gear retracted, and the airplane as
 nearly as possible trim at 1.4 VS, apply takeoff power rapidly while
 maintaining the same airspeed.
   (5) With power off, landing gear and flaps extended, and the airplane as
 nearly as possible in trim at 1.4 VSO, obtain and maintain airspeeds between
 1.1 VSO and either 1.7 VSO or VFE, whichever is lower.
   (c) At speeds above VMO/MMO and up to VD/MD, a maneuvering capability of
 1.5 g must be demonstrated to provide a margin to recover from upset or
 inadvertent speed increase.
   (d) It must be possible, with a pilot control force of not more than 10
 pounds, to maintain a speed of not more than 1.3 VSO, during a power-off
 glide with landing gear and wing flaps extended, for any weight of the
 airplane, up to and including the maximum weight.
   (e) By using normal flight and power controls, except as otherwise noted in
 paragraphs (e)(1) and (e)(2) of this section, it must be possible to
 establish a zero rate of descent at an attitude suitable for a controlled
 landing without exceeding the operational and structural limitations of the
 airplane, as follows:
   (1) For single-engine and multiengine airplanes, without the use of the
 primary longitudinal control system.
   (2) For multiengine airplanes--
   (i) Without the use of the primary directional control; and
   (ii) If a single failure of any one connecting or transmitting link would
 affect both the longitudinal and directional primary control system, without
 the primary longitudinal and directional control system.

 [Amdt. 23-45, 58 FR 42157, Aug. 6, 1993; 58 FR 51970, Oct. 5, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.147   Directional and lateral control.

   For each multiengine airplane, it must be possible, while holding the wings
 level within 5 degrees, to make sudden changes in heading safely in both
 directions. This must be shown at 1.4 VS1 with heading changes up to 15
 degrees (except that the heading change at which the rudder force corresponds
 to the limits specified in Sec. 23.143 need not be exceeded), with the--
   (a) Critical engine inoperative and its propeller in the minimum drag
 position;
   (b) Remaining engines at maximum continuous power;
   (c) Landing gear--
   (1) Retracted; and
   (2) Extended; and
   (d) Flaps in the most favorable climb position.

 [Amdt. 23-45, 58 FR 42157, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.149   Minimum control speed.

   (a) VMC is the calibrated airspeed at which, when the critical engine is
 suddenly made inoperative, it is possible to maintain control of the airplane
 with that engine still inoperative and then maintain straight flight at the
 same speed with an angle of bank of not more than 5 degrees. The ability to
 maintain straight and level flight at VMC in a static condition with a bank
 angle of not more than 5 degrees must also be demonstrated. The method used
 to simulate critical engine failure must represent the most critical mode of
 powerplant failure, with respect to controllability expected in service.
   (b) VMC may not exceed 1.2 VS1, where VS1 is determined at the maximum
 takeoff weight, with--
   (1) Maximum available takeoff power or thrust on the engines;
   (2) The most unfavorable center of gravity;
   (3) The airplane trimmed for takeoff;
   (4) The maximum sea level takeoff weight, or any lesser weight necessary to
 show VMC;
   (5) The airplane in the most critical takeoff configuration, with the
 propeller controls in the recommended takeoff position and the landing gear
 retracted; and
   (6) The airplane airborne and the ground effect negligible.
   (c) A minimum speed to intentionally render the critical engine inoperative
 must be established and designated as the safe, intentional, one-engine-
 inoperative speed, VSSE.
   (d) At Vmc, the rudder pedal force required to maintain control may not
 exceed 150 pounds, and it may not be necessary to reduce power or thrust of
 the operative engines. During the maneuver, the airplane may not assume any
 dangerous attitude and it must be possible to prevent a heading change of
 more than 20 degrees.

 [Amdt. 23-21, 43 FR 2318, Jan. 16, 1978, as amended by Amdt. No. 23-45, 58
 FR 42157, Aug. 6, 1993; 58 FR 51970, Oct. 5, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.151  Acrobatic maneuvers.

   Each acrobatic and utility category airplane must be able to perform safely
 the acrobatic maneuvers for which certification is requested. Safe entry
 speeds for these maneuvers must be determined.






 Sec. 23.153   Control during landings.

   It must be possible, while in the landing configuration, to safely complete
 a landing without exceeding the one hand control force specified in Sec.
 23.143(c) following an approach to land--
   (a) At a speed 5 knots less than the speed used in complying with the
 requirements of Sec. 23.75 and with the airplane in trim, or as nearly as
 possible in trim, and without the trimming control being moved throughout the
 maneuver;
   (b) At an approach gradient equal to the steepest recommended for
 operational use; and
   (c) With only those power or thrust changes that would be made when landing
 normally from an approach at 1.3 VS1.

 [Amdt. 23-45, 58 FR 42157, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.155  Elevator control force in maneuvers.

   (a) The elevator control force needed to achieve the positive limit
 maneuvering load factor may not be less than:
   (1) For wheel controls, W/100 (where W is the maximum weight) or 20 pounds,
 whichever is greater, except that it need not be greater than 50 pounds; or
   (2) For stick controls, W/140 (where W is the maximum weight) or 15 pounds,
 whichever is greater, except that it need not be greater than 35 pounds.
   (b) The requirement of paragraph (a) of this section must be met at 75
 percent of maximum continuous power for reciprocating engines, or the maximum
 power or thrust selected by the applicant as an operating limitation for use
 during cruise for reciprocating or turbine engines, and with the wing flaps
 and landing gear retracted--
   (1) In a turn, with the trim setting used for wings level flight at VA; and
   (2) In a turn with the trim setting used for the maximum wings level flight
 speed, except that the speed may not exceed VNE or VMO/MMO, whichever is
 appropriate.
   (c) Compliance with the requirements of this section may be demonstrated by
 measuring the normal acceleration that is achieved with the limiting stick
 force or by establishing the stick force per g gradient and extrapolating to
 the appropriate limit.

 [Amdt. 23-14, 38 FR 31819, Nov. 19, 1973; 38 FR 32784, Nov. 28, 1973, as
 amended by Amdt. 23-45, 58 FR 42158, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.157  Rate of roll.

   (a) Takeoff. It must be possible, using a favorable combination of
 controls, to roll the airplane from a steady 30-degree banked turn through an
 angle of 60 degrees, so as to reverse the direction of the turn within:
   (1) For an airplane of 6,000 pounds or less maximum weight, 5 seconds from
 initiation of roll; and
   (2) For an airplane of over 6,000 pounds maximum weight,

                                 (W+500)/1,300

 seconds, but not more than 10 seconds, where W is the weight in pounds.
   (b) The requirement of paragraph (a) of this section must be met when
 rolling the airplane in each direction with--
   (1) Flaps in the takeoff position;
   (2) Landing gear retracted;
   (3) For a single-engine airplane, at maximum takeoff power; and for a
 multiengine airplane with the critical engine inoperative and the propeller
 in the minimum drag position, and the other engines at maximum takeoff power;
 and
   (4) The airplane trimmed at a speed equal to the greater of 1.2 VS1 or 1.1
 VMC, or as nearly as possible in trim for straight flight.
   (c) Approach. It must be possible, using a favorable combination of
 controls, to roll the airplane from a steady 30-degree banked turn through an
 angle of 60 degrees, so as to reverse the direction of the turn within:
   (1) For an airplane of 6,000 pounds or less maximum weight, 4 seconds from
 initiation of roll; and
   (2) For an airplane of over 6,000 pounds maximum weight,

                                (W+2,800)/2,200

 seconds, but not more than 7 seconds, where W is the weight in pounds.
   (d) The requirement of paragraph (c) must be met when rolling the airplane
 in either direction in the following conditions:
   (1) Flaps extended;
   (2) Landing gear extended;
   (3) All engines operating at idle power or thrust and with all engines
 operating at the power or thrust for level flight; and
   (4) The airplane trimmed at the speed that is used in determining
 compliance with Sec. 23.75.

 [Amdt. 23-14, 38 FR 31819, Nov. 19, 1973, as amended by Amdt. No. 23-45, 58
 FR 42158, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






                                     Trim






 Sec. 23.161  Trim.

   (a) General. Each airplane must meet the trim requirements of this section
 after being trimmed, and without further pressure upon, or movement of, the
 primary controls or their corresponding trim controls by the pilot or the
 automatic pilot.
   (b) Lateral and directional trim. The airplane must maintain lateral and
 directional trim in level flight with the landing gear and wing flaps
 retracted as follows:
   (1) For normal, utility, and acrobatic category airplanes at a speed of 0.9
 VH, VC, or VM0, whichever is the lower; and
   (2) For commuter category airplanes, at a speed of VH or VMO/MMO, whichever
 is lower.
   (c) Longitudinal trim. The airplane must maintain longitudinal trim under
 each of the following conditions, except that it need not maintain trim at a
 speed greater than VMO/MMO:
   (1) A climb with maximum continuous power at--
   (i) The speed used in determining the climb performance required by Sec.
 23.65 of this part with the landing gear retracted, and the flaps in the
 takeoff position; and
   (ii) The recommended all-engines-operating climb speed specified in Sec.
 23.1585(a)(2)(i) of this part.
   (2) An approach at a gradient of descent of 5.2 percent (3 degrees) with
 the landing gear extended, and with--
   (i) Flaps retracted and at a speed of 1.4 VS1; and
   (ii) The applicable airspeed and flap position used in showing compliance
 with Sec. 23.75.
   (3) Level flight at any speed with the landing gear and wing flaps
 retracted as follows:
   (i) For normal, utility, and acrobatic category airplanes, at any speeds
 from the lesser of VH and VN0 or VM0, as applicable, to 1.4 VS1; and
   (ii) For commuter category airplanes, at a speed of VH or VMO/MMO,
 whichever is lower, to either VX or 1.4VS1.
   (4) A descent at 0.9 VNO or 0.9 VMO, whichever is applicable, with power
 off and with the landing gear and flaps retracted.
   (d) In addition, each multiengine airplane must maintain longitudinal and
 directional trim, and the lateral control force must not exceed 5 pounds, at
 the speed used in complying with Sec. 23.67 for normal, utility, and
 acrobatic categories and at a speed between VY and 1.4 VS1 for commuter
 category with--
   (1) The critical engine inoperative, and if applicable, its propeller in
 the minimum drag position;
   (2) The remaining engines at maximum continuous power;
   (3) The landing gear retracted;
   (4) Wing flaps in the position selected for showing compliance with Sec.
 23.67 for normal, utility, and acrobatic category airplanes and wing flaps
 retracted for commuter category airplanes.
   (5) An angle of bank of not more than five degrees.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-21, 43 FR
 2318, Jan. 16, 1978; Amdt. 23-34, 52 FR 1828, Jan. 15, 1987; Amdt. 23-42, 56
 FR 351, Jan. 3, 1991; Amdt. 23-42, 56 FR 5455, Feb. 11, 1991]

 *****************************************************************************


 56 FR 344, No. 2, Jan. 3, 1991

   SUMMARY: This final rule upgrades the airworthiness standards for normal,
 utility, acrobatic, and commuter category airplanes. This amendment provides
 airworthiness standards for advancements in technology being incorporated in
 current designs, permits type certification of spin resistant airplanes, and
 reduces the regulatory burden in showing compliance with some of the
 requirements for the design and type certification of small airplanes. These
 new and amended airworthiness standards also result in the need for new
 definitions. As a result, new definitions are added.

   DATES: February 4, 1991.

 *****************************************************************************






                                   Stability






 Sec. 23.171  General.

   The airplane must be longitudinally, directionally, and laterally stable
 under Secs. 23.173 through 23.181. In addition, the airplane must show
 suitable stability and control "feel" (static stability) in any condition
 normally encountered in service, if flight tests show it is necessary for
 safe operation.






 Sec. 23.173  Static longitudinal stability.

   Under the conditions specified in Sec. 23.175 and with the airplane trimmed
 as indicated, the characteristics of the elevator control forces and the
 friction within the control system must be as follows:
   (a) A pull must be required to obtain and maintain speeds below the
 specified trim speed and a push required to obtain and maintain speeds above
 the specified trim speed. This must be shown at any speed that can be
 obtained, except that speeds requiring a control force in excess of 40 pounds
 or speeds above the maximum allowable speed or below the minimum speed for
 steady unstalled flight, need not be considered.
   (b) The airspeed must return to within the tolerances specified for
 applicable categories of airplanes when the control force is slowly released
 at any speed within the speed range specified in paragraph (a) of this
 section. The applicable tolerances are--
   (1) The airspeed must return to within plus or minus 10 percent of the
 original trim airspeed; and
   (2) For commuter category airplanes, the airspeed must return to within
 plus or minus 7.5 percent of the original trim airspeed for the cruising
 condition specified in Sec. 23.175(b).
   (c) The stick force must vary with speed so that any substantial speed
 change results in a stick force clearly perceptible to the pilot.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-14, 38 FR
 31820 Nov. 19, 1973; Amdt. 23-34, 52 FR 1828, Jan. 15, 1987]






 Sec. 23.175  Demonstration of static longitudinal stability.

   Static longitudinal stability must be shown as follows:
   (a) Climb. The stick force curve must have a stable slope, at speeds
 between 85 and 115 percent of the trim speed, with--
   (1) Flaps in the climb position;
   (2) Landing gear retracted;
   (3) All reciprocating engines operating at maximum continuous power, or
 turbine engines operating at the maximum power selected by the applicant as
 an operating limitation for use during climb; and
   (4) The airplane trimmed for VY, except that the speed need not be less
 than 1.4 VS1.
   (b) Cruise--Landing gear retracted (or fixed gear). (1) For the cruise
 conditions specified in paragraphs (b) (2) and (3) of this section, the
 following apply:
   (i) The speed need not be less than 1.3 VS1.
   (ii) For airplanes with VNE established under Sec. 23.1505(a), the speed
 need not be greater than VNE.
   (iii) For airplanes with VMO/MMO established under Sec. 23.1505(c), the
 speed need not be greater than a speed midway between VMO/MMO and the lesser
 of VD/MD or the speed demonstrated under Sec. 23.251, except that for
 altitudes where Mach number in the limiting factor, the speed need not exceed
 that corresponding to the Mach number at which effective speed warning
 occurs.
   (2) High speed cruise. The stick force curve must have a stable slope at
 all speeds within a range that is the greater of 15 percent of the trim speed
 plus the resulting free return speed range or 40 knots plus the resulting
 free return speed range for normal, utility, and acrobatic category
 airplanes, above and below the trim speed. For commuter category airplanes,
 the stick force curve must have a stable slope for a speed range of 50 knots
 from the trim speed, except that the speeds need not exceed VFC/MFC or be
 less than 1.4 VS1 and this speed range is considered to begin at the outer
 extremes of the friction band with a stick force not to exceed 50 pounds. In
 addition, for commuter category airplanes, VFC/MFC may not be less than a
 speed midway between VMO/MMO and VDF/MDF, except that, for altitudes where
 Mach number is the limiting factor, MFC need not exceed the Mach number at
 which effective speed warning occurs. These requirements for all categories
 of airplane must be met with--
   (i) Flaps retracted.
   (ii) Seventy-five percent of maximum continuous power for reciprocating
 engines or, for turbine engines, the maximum cruising power or thrust
 selected by the applicant as an operating limitation, except that the power
 need not exceed that required at VNE for airplanes with VNE established under
 Sec. 23.1505(a), or that required at VMO/MMO for airplanes with VMO/MMO
 established under Sec. 23.1505(c).
   (iii) The airplane trimmed for level flight.
   (3) Low speed cruise. The stick force curve must have a stable slope under
 all the conditions prescribed in paragraph (b)(2) of this section, except
 that the power is that required for level flight at a speed midway between
 1.3 VS1 and the trim speed obtained in the high speed cruise condition under
 paragraph (b)(2) of this section.
   (c) Landing gear extended (airplanes with retractable gear). The stick
 force curve must have a stable slope at all speeds within a range from 15
 percent of the trim speed plus the resulting free return speed range below
 the trim speed, to the trim speed (except that the speed range need not
 include speeds less than 1.4 VS1 nor speeds greater than VLE, with--
   (1) Landing gear extended;
   (2) Flaps retracted;
   (3) 75 percent of maximum continuous power for reciprocating engines, or
 for turbine engines, the maximum cruising power or thrust selected by the
 applicant as an operating limitation, except that the power need not exceed
 that required for level flight at VLE; and
   (4) The airplane trimmed for level flight.
   (d) Approach and landing. The stick force curve must have a stable slope at
 speeds between 1.1 VS1 and 1.8 VS1 with--
   (1) Wing flaps in the landing position;
   (2) Landing gear extended;
   (3) The airplane trimmed at a speed in compliance with Sec. 23.161(c)(2).
   (4) Both power off and enough power to maintain a 3 deg. angle of descent.

 [Amdt. 23-7, 34 FR 13087, Aug. 13, 1969, as amended by Amdt. 23-14, 38 FR
 31820, Nov. 19, 1973; Amdt. 23-17, 41 FR 55464, Dec. 20, 1976; Amdt. 23-34,
 52 FR 1828, Jan. 15, 1987; Amdt. 23-45, 58 FR 42158, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.177  Static directional and lateral stability.

   (a) Three-control airplanes. The stability requirements for three-control
 airplanes are as follows:
   (1) The static directional stability, as shown by the tendency to recover
 from a skid with the rudder free, must be positive for any landing gear and
 flap position appropriate to the takeoff, climb, cruise, approach, and
 landing configurations. This must be shown with symmetrical power up to
 maximum continuous power, and at speeds from 1.2 VS1 up to the maximum
 allowable speed for the condition being investigated in the takeoff, climb,
 cruise, and approach configurations. For the landing configuration, the power
 must be up to that necessary to maintain a three degree angle of descent in
 coordinated flight. The angle of sideslip for these tests must be appropriate
 to the type of airplane. At larger angles of sideslip, up to that at which
 full rudder is used or a control force limit in Sec. 23.143 is reached,
 whichever occurs first, and at speeds from 1.2 VS1 to VA, the rudder pedal
 force must not reverse.
   (2) The static lateral stability, as shown by the tendency to raise the low
 wing in a sideslip, must be positive for any landing gear and flap position.
 This must be shown with symmetrical power, up to 75 percent of maximum
 continuous power, at speeds above 1.2 VS1 in the takeoff configuration and
 1.3 VS1 in other configurations, up to the maximum allowable speed for the
 configuration being investigated in the takeoff, climb, approach, and cruise
 configurations. For the landing configuration, the power must be up to that
 necessary to maintain a three degree angle of descent in coordinated flight.
 The angle of bank for these tests must be appropriate to the type of airplane
 but in no case may the constant heading sideslip angle be less than that
 obtainable with 10 deg. bank, or, if less, the maximum bank angle obtainable
 with full rudder deflection or 150 pounds rudder force. The static lateral
 stability must not be negative at 1.2 VS1.
   (3) In straight, steady slips at 1.2 VS1 for any landing gear and flap
 positions, and for any symmetrical power conditions up to 50 percent of
 maximum continuous power, the aileron and rudder control movements and forces
 must increase steadily, but not necessarily in constant proportion, as the
 angle of slip is increased up to the maximum appropriate to the type of
 airplane. At larger slip angles, up to the angle at which full rudder or
 aileron control is used or a control force limit contained in Sec. 23.143 is
 obtained, the aileron and rudder control movements and forces must not
 reverse as the angle of sideslip is increased. Enough bank must accompany the
 sideslip to hold a constant heading. Rapid entry into, and recovery from, a
 maximum sideslip considered appropriate for the airplane must not result in
 uncontrollable flight characteristics.
   (b) Two-control (or simplified control) airplanes. The stability
 requirements for two-control airplanes are as follows:
   (1) The directional stability of the airplane must be shown by showing
 that, in each configuration, it can be rapidly rolled from a 45 degree bank
 in one direction to a 45 degree bank in the opposite direction without
 showing dangerous skid characteristics.
   (2) The lateral stability of the airplane must be shown by showing that it
 will not assume a dangerous attitude or speed when the controls are abandoned
 for two minutes. This must be done in moderately smooth air with the airplane
 trimmed for straight level flight at 0.9 VH or VC, whichever is lower, with
 flaps and landing gear retracted, and with a rearward center of gravity.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan, 9, 1965, as
 amended by Amdt. 23-21, 43 FR 2318, Jan. 16, 1978; Amdt. No. 23-45, 58 FR
 42158, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.179  [Removed. Amdt. No. 23-45, 58 FR 42158, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.181   Dynamic stability.

   (a) Any short period oscillation not including combined lateral-directional
 oscillations occurring between the stalling speed and the maximum allowable
 speed appropriate to the configuration of the airplane must be heavily damped
 with the primary controls--
   (1) Free; and
   (2) In a fixed position.
   (b) Any combined lateral-directional oscillations ("Dutch roll") occurring
 between the stalling speed and the maximum allowable speed appropriate to the
 configuration of the airplane must be damped to 1/10 amplitude in 7 cycles
 with the primary controls--
   (1) Free; and
   (2) In a fixed position.
   (c) If it is determined that the function of a stability augmentation
 system, reference Sec. 23.672, is needed to meet the flight characteristic
 requirements of this part, the primary control requirements of paragraphs
 (a)(2) and (b)(2) of this section are not applicable to the tests needed to
 verify the acceptability of that system.
   (d) During the conditions as specified in Sec. 23.175, when the
 longitudinal control force required to maintain speeds differing from the
 trim speed by at least plus and minus 15 percent is suddenly released, the
 response of the airplane must not exhibit any dangerous characteristics nor
 be excessive in relation to the magnitude of the control force released. Any
 long-period oscillation of flight path, phugoid oscillation, that results
 must not be so unstable as to increase the pilot's workload or otherwise
 endanger the airplane.

 [Amdt. 23-21, 43 FR 2318, Jan. 16, 1978, as amended by Amdt. No. 23-45, 58
 FR 42158, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






                                    Stalls






 Sec. 23.201  Wings level stall.

   (a) For an airplane with independently controlled roll and directional
 controls, it must be possible to produce and to correct roll by unreversed
 use of the rolling control and to produce and to correct yaw by unreversed
 use of the directional control, up to the time the airplane pitches.
   (b) For an airplane with interconnected lateral and directional controls (2
 controls) and for an airplane with only one of these controls, it must be
 possible to produce and correct roll by unreversed use of the rolling control
 without producing excessive yaw, up to the time the airplane pitches.
   (c) The wings level stall characteristics must be demonstrated in flight as
 follows: Starting from a speed above the stall warning speed, the elevator
 control must be pulled back so that the rate of speed reduction will not
 exceed one knot per second until a stall is produced, as shown by an
 uncontrollable downward pitching motion of the airplane, until the control
 reaches the stop or until the activation of an artificial stall barrier, for
 example, stick pusher. Normal use of the elevator control for recovery is
 allowed after the pitching motion has unmistakably developed or after the
 control has been held against the stop for not less than two seconds. In
 addition, engine power may not be increased for recovery until the speed has
 increased to approximately 1.2 VS1.
   (d) Except where made inapplicable by the special features of a particular
 type of airplane, the following apply to the measurement of loss of altitude
 during a stall:
   (1) The loss of altitude encountered in the stall (power on or power off)
 is the change in altitude (as observed on the sensitive altimeter testing
 installation) between the altitude at which the airplane pitches and the
 altitude at which horizontal flight is regained.
   (2) If power is required during stall recovery, the power used must be that
 used under the normal operating procedures selected by the applicant for this
 maneuver; however, the power used to regain level flight may not be increased
 until the speed has increased to approximately 1.2 VS1.
   (e) During the recovery part of the maneuver, it must be possible to
 prevent more than 15 degrees of roll or yaw by the normal use of controls.
   (f) Compliance with the requirements of this section must be shown under
 the following conditions:
   (1) Wing flaps: Full up, full down, and intermediate, if appropriate.
   (2) Landing gear: Retracted and extended.
   (3) Cowl flaps: Appropriate to configuration.
   (4) Power: Power off, and 75 percent maximum continuous power. If the
 power-to-weight ratio at 75 percent continuous power provides undesirable
 stall characteristics at extremely nose-high attitudes, the test may be
 accomplished with the power or thrust required for level flight in the
 landing configuration at maximum landing weight and a speed of 1.4 VSO, but
 the power may not be less than 50 percent of maximum continuous power.
   (5) Trim:  The airplane trimmed at a speed as near 1.5 VS1 as practicable.
   (6) Propeller: Full increase rpm position for the power off condition.

 [Amdt. 23-14, 38 FR 31820, Nov. 19, 1973, as amended by Amdt. 23-45, 58 FR
 42158, Aug. 6, 1993; 58 FR 51970, Oct. 5, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.203  Turning flight and accelerated stalls.

   Turning flight and accelerated stalls must be demonstrated in flight tests
 as follows:
   (a) Establish and maintain a coordinated turn in a 30 degree bank. Reduce
 speed by steadily and progressively tightening the turn with the elevator
 until the airplane is stalled or until the elevator has reached its stop. The
 rate of speed reduction must be constant, and:
   (1) For a turning flight stall, may not exceed one knot per second; and
   (2) For an accelerated stall, be 3 to 5 knots per second with steadily
 increasing normal acceleration.
   (b) When the stall has fully developed or the elevator has reached its
 stop, it must be possible to regain wings level flight by normal use of the
 flight controls but without increasing power, and without--
   (1) Excessive loss of altitude;
   (2) Undue pitchup;
   (3) Uncontrollable tendency to spin;
   (4) Exceeding a bank angle of 60 degrees in the original direction of the
 turn or 30 degrees in the opposite direction in the case of turning flight
 stalls, and without exceeding a bank angle of 90 degrees in the original
 direction of the turn or 60 degrees in the opposite direction in the case of
 accelerated stalls; and
   (5) Exceeding the maximum permissible speed or allowable load factor.
   (c) Compliance with the requirements of this section must be shown with:
   (1) Wing Flaps:  Retracted, fully extended, and in each intermediate
 position, as appropriate.
   (2) Landing gear: Retracted and extended;
   (3) Cowl flaps: Appropriate to configuration;
   (4) Power: Power or thrust off, and 75 percent maximum continuous power or
 thrust. If the power-to-weight ratio at 75 percent continuous power or thrust
 provides undesirable stall characteristics at extremely nose-high attitudes,
 the test may be accomplished with the power or thrust required for level
 flight in the landing configuration at maximum landing weight and a speed of
 1.4 VS0, but the power may not be less than 50 percent of maximum continuous
 power.
   (5) Trim: The airplane trimmed at a speed as near 1.5 VS1 as practicable.

 [Amdt. 23-14, 38 FR 31820, Nov. 19, 1973, as amended by Amdt. No. 23-45, 58
 FR 42159, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.205  Critical engine inoperative stalls.

   (a) A multiengine airplane may not display any undue spinning tendency and
 must be safely recoverable without applying power to the inoperative engine
 when stalled. The operating engines may be throttled back during the recovery
 from stall.
   (b) Compliance with paragraph (a) of the section must be shown with:
   (1) Wing flaps: Retracted and set to the position used to show compliance
 with Sec. 23.67.
   (2) Landing gear: Retracted.
   (3) Cowl flaps: Appropriate to level flight critical engine inoperative.
   (4) Power: Critical engine inoperative and the remaining engine(s) at 75
 percent maximum continuous power or thrust or the power or thrust at which
 the use of maximum control travel just holds the wings laterally level in the
 approach to stall, whichever is lesser.
   (5) Propeller: Normal inoperative position for the inoperative engine.
   (6) Trim: Level flight, critical engine inoperative, except that for an
 airplane of 6,000 pounds or less maximum weight that has a stalling speed of
 61 knots or less and cannot maintain level flight with the critical engine
 inoperative, the airplane must be trimmed for straight flight, critical
 engine inoperative, at a speed as near 1.5 VS1 as practicable.

 [Amdt. 23-14, 38 FR 31820, Nov. 19, 1973, as amended by Amdt. No. 23-45, 58
 FR 42159, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.207  Stall warning.

   (a) There must be a clear and distinctive stall warning, with the flaps and
 landing gear in any normal position, in straight and turning flight.
   (b) The stall warning may be furnished either through the inherent
 aerodynamic qualities of the airplane or by a device that will give clearly
 distinguishable indications under expected conditions of flight. However, a
 visual stall warning device that requires the attention of the crew within
 the cockpit is not acceptable by itself.
   (c) For the stall tests required by Sec. 23.201(c), the stall warning must
 begin at a speed exceeding the stalling speed by a margin of not less than 5
 knots, but not more than the greater of 10 knots or 15 percent of the
 stalling speed, and must continue until the stall occurs.
   (d) For all other stall tests, the stall warning must begin at not less
 than 5 knots above the stall speed and be sufficiently in advance of the
 stall for the stall to be averted by action after the stall warning first
 occurs. In addition, when following the procedures of Sec. 23.1585, the stall
 warning must not operate during a normal takeoff, a takeoff continued with
 one engine inoperative or approach to landing.

 [Amdt. 23-7, 34 FR 13087, Aug. 13, 1969, as amended by Amdt. 23-45, 58 FR
 42159, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight airworthiness standards
 for normal, utility, acrobatic, and commuter category airplanes. The changes
 are based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, in St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






                                   Spinning






 Sec. 23.221  Spinning.

   (a) Normal category. Except as provided in paragraph (d) of this section, a
 single-engine, normal category airplane must demonstrate compliance with
 either the one-turn spin or the spin-resistant requirements of this
 paragraph.
   (1) One-turn spin. The airplane must recover from a one-turn spin or a
 three-second spin, whichever takes longer, in not more than one additional
 turn after the controls have been applied for recovery. In addition--
   (i) For both the flaps-retracted and flaps-extended conditions, the
 applicable airspeed limit and positive limit maneuvering load factor must not
 be exceeded;
   (ii) There must be no excessive back pressure during the spin or recovery;
   (iii) It must be impossible to obtain unrecoverable spins with any use of
 the flight or engine power controls either at the entry into or during the
 spin; and
   (iv) For the flaps-extended condition, the flaps may be retracted during
 the recovery, but not before rotation has ceased.
   (2) Spin resistant. The airplane must be demonstrated to be spin resistant
 by the following:
   (i) During the stall maneuvers contained in Sec. 23.201, the pitch control
 must be pulled back and held against the stop. Then, using ailerons and
 rudders in the proper direction, it must be possible to maintain wings-level
 flight within 15 degrees of bank and to roll the airplane from a 30-degree
 bank in one direction to a 30-degree bank in the other direction;
   (ii) Reduce the airplane speed using pitch control at a rate of
 approximately 1 knot per second until the pitch control reaches the stop;
 then with the pitch control pulled back and held against the stop, apply full
 rudder control in a manner to promote spin entry, for a period of 7 seconds
 or through a 360-degree heading change, whichever occurs first. If the 360-
 degree heading change is reached first, it must have taken no fewer than 4
 seconds. This maneuver must be performed first with the ailerons in the
 neutral position, and then with the ailerons deflected opposite the direction
 of turn in the most adverse manner. Power or thrust and airplane
 configuration must be set in accordance with Sec. 23.201(f) without change
 during the maneuver. At the end of 7 seconds or a 360 degree heading change,
 the airplane must respond immediately and normally to primary flight controls
 applied to regain coordinated, unstalled flight without reversal of control
 effect and without exceeding the temporary control forces specified by Sec.
 23.143(c); and
   (iii) Compliance with Secs. 23.201 and 23.203 must be demonstrated with the
 airplane in uncoordinated flight, corresponding to one ball width
 displacement on a slip-skid indicator, unless one ball width displacement
 cannot be obtained with full rudder, in which case the demonstration must be
 with full rudder applied.
   (b) Utility category. A utility category airplane must meet the
 requirements of paragraph (a) of this section or the requirements of
 paragraph (c) of this section if approval for spinning is requested.
   (c) Acrobatic category. An acrobatic category airplane must meet the
 following requirements:
   (1) The airplane must recover from any point in a spin, in not more than
 one and one-half additional turns after normal recovery application of the
 controls. Prior to normal recovery application of the controls, the spin test
 must proceed for six turns or 3 seconds, whichever takes longer, with flaps
 retracted, and one turn or 3 seconds, whichever takes longer, with flaps
 extended. However, beyond 3 seconds, the spin may be discontinued when spiral
 characteristics appear with flaps retracted.
   (2) For both the flaps-retracted and flaps-extended conditions, the
 applicable airspeed limit and positive limit maneuvering load factor may not
 be exceeded. For the flaps-extended condition, the flaps may be retracted
 during recovery, if a placard is installed prohibiting intentional spins with
 flaps extended.
   (3) It must be impossible to obtain unrecoverable spins with any use of the
 flight or engine power controls either at the entry into or during the spin.
   (d) Airplanes "characteristically incapable of spinning". If it is desired
 to designate an airplane as "characteristically incapable of spinning", this
 characteristic must be shown with--
   (1) A weight five percent more than the highest weight for which approval
 is requested;
   (2) A center of gravity at least three percent aft of the rearmost position
 for which approval is requested;
   (3) An available elevator up-travel four degrees in excess of that to which
 the elevator travel is to be limited for approval; and
   (4) An available rudder travel seven degrees, in both directions, in excess
 of that to which the rudder travel is to be limited for approval.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13087, Aug. 13, 1969; Amdt. 23-42, 56 FR 352, Jan. 3, 1991; 56 FR 12584,
 Mar. 26, 1991]

 *****************************************************************************


 56 FR 344, No. 2, Jan. 3, 1991

   SUMMARY: This final rule upgrades the airworthiness standards for normal,
 utility, acrobatic, and commuter category airplanes. This amendment provides
 airworthiness standards for advancements in technology being incorporated in
 current designs, permits type certification of spin resistant airplanes, and
 reduces the regulatory burden in showing compliance with some of the
 requirements for the design and type certification of small airplanes. These
 new and amended airworthiness standards also result in the need for new
 definitions. As a result, new definitions are added.

   DATES: February 4, 1991.

 *****************************************************************************






                   Ground and Water Handling Characteristics






 Sec. 23.231  Longitudinal stability and control.

   (a) A landplane may have no uncontrollable tendency to nose over in any
 reasonably expected operating condition, including rebound during landing or
 takeoff. Wheel brakes must operate smoothly and may not induce any undue
 tendency to nose over.
   (b) A seaplane or amphibian may not have dangerous or uncontrollable
 porpoising characteristics at any normal operating speed on the water.






 Sec. 23.233  Directional stability and control.

   (a) A 90 degree cross-component of wind velocity, demonstrated to be safe
 for taxiing, takeoff and landing must be established and must not be less
 than 0.2 VSO.
   (b) The airplane must be satisfactorily controllable in power-off landings
 at normal landing speed, without using brakes or engine power to maintain a
 straight path until the speed has decreased to at least 50 percent of the
 speed at touchdown.
   (c) The airplane must have adequate directional control during taxiing.
   (d) Seaplanes must demonstrate satisfactory directional stability and
 control for water operations up to the maximum wind velocity specified in
 paragraph (a) of this section.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. No. 23-45, 58 FR 42159, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.235  Taxiing, takeoff, and landing condition.

   (a) The airplane must be demonstrated to have satisfactory characteristics
 and the shock-absorbing mechanism must not damage the structure of the
 airplane when the airplane is taxied on the roughest ground that may be
 reasonably expected in normal operation, and when takeoffs and landings are
 performed on unpaved runways having the roughest surface that may reasonably
 be expected in normal operation.
   (b) A wave height, demonstrated to be safe for operation, and any necessary
 water handling procedures for seaplanes and amphibians, must be established.

 [Amdt. No. 23-45, 58 FR 42159, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.239  Spray characteristics.

   Spray may not dangerously obscure the vision of the pilots or damage the
 propellers or other parts of a seaplane or amphibian at any time during
 taxiing, takeoff, and landing.






                       Miscellaneous Flight Requirements






 Sec. 23.251  Vibration and buffeting.

   There must be no vibration or buffeting severe enough to result in
 structural damage, and each part of the airplane must be free from excessive
 vibration, under any appropriate speed and power conditions up to VD/MD. In
 addition, there must be no buffeting in any normal flight condition severe
 enough to interfere with the satisfactory control of the airplane or cause
 excessive fatigue to the flight crew. Stall warning buffeting within these
 limits is allowable.

 [Amdt. No. 23-45, 58 FR 42159, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.253  High speed characteristics.

   If a maximum operating speed VMO/MMO is established under Sec. 23.1505(c),
 the following speed increase and recovery characteristics must be met:
   (a) Operating conditions and characteristics likely to cause inadvertent
 speed increases (including upsets in pitch and roll) must be simulated with
 the airplane trimmed at any likely speed up to VMO/MMO. These conditions and
 characteristics include gust upsets, inadvertent control movements, low stick
 force gradients in relation to control friction, passenger movement, leveling
 off from climb, and descent from Mach to airspeed limit altitude.
   (b) Allowing for pilot reaction time after occurrence of the effective
 inherent or artificial speed warning specified in Sec. 23.1303, it must be
 shown that the airplane can be recovered to a normal attitude and its speed
 reduced to VMO/MMO, without--
   (1) Exceptional piloting strength or skill;
   (2) Exceeding VD/MD, the maximum speed shown under Sec. 23.251, or the
 structural limitations; or
   (3) Buffeting that would impair the pilot's ability to read the instruments
 or to control the airplane for recovery.
   (c) There may be no control reversal about any axis at any speed up to the
 maximum speed shown under Sec. 23.251. Any reversal of elevator control force
 or tendency of the airplane to pitch, roll, or yaw must be mild and readily
 controllable, using normal piloting techniques.

 [Amdt. 23-7, 34 FR 13087, Aug. 13, 1969; as amended by Amdt. 23-26, 45 FR
 60170, Sept. 11, 1980; Amdt. No. 23-45, 58 FR 42160, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************


                             Subpart C--Structure



                                    General





 Sec. 23.301  Loads.

   (a) Strength requirements are specified in terms of limit loads (the
 maximum loads to be expected in service) and ultimate loads (limit loads
 multiplied by prescribed factors of safety). Unless otherwise provided,
 prescribed loads are limit loads.
   (b) Unless otherwise provided, the air, ground, and water loads must be
 placed in equilibrium with inertia forces, considering each item of mass in
 the airplane. These loads must be distributed to conservatively approximate
 or closely represent actual conditions. Methods used to determine load
 intensities and distribution on canard and tandem wing configurations must be
 validated by flight test measurement unless the methods used for determining
 those loading conditions are shown to be reliable or conservative on the
 configuration under consideration.
   (c) If deflections under load would significantly change the distribution
 of external or internal loads, this redistribution must be taken into
 account.
   (d) Simplified structural design criteria may be used if they result in
 design loads not less than those prescribed in Secs. 23.331 through 23.521.
 For conventional, single-engine airplanes with design weights of 6,000 pounds
 or less, the design criteria of Appendix A of this part are an approved
 equivalent of Secs. 23.321 through 23.459. If Appendix A is used, the entire
 Appendix must be substituted for the corresponding sections of this part.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. 23-28, 47 FR 13315, Mar. 29, 1982; Amdt. 23-42, 56 FR 352,
 Jan. 3, 1991]

 *****************************************************************************


 56 FR 344, No. 2, Jan. 3, 1991

   SUMMARY: This final rule upgrades the airworthiness standards for normal,
 utility, acrobatic, and commuter category airplanes. This amendment provides
 airworthiness standards for advancements in technology being incorporated in
 current designs, permits type certification of spin resistant airplanes, and
 reduces the regulatory burden in showing compliance with some of the
 requirements for the design and type certification of small airplanes. These
 new and amended airworthiness standards also result in the need for new
 definitions. As a result, new definitions are added.

   DATES: February 4, 1991.

 *****************************************************************************






 Sec. 23.302   Canard or tandem wing configurations.

   The forward structure of a canard or tandem wing configuration must:
   (a) Meet all requirements of subpart C and subpart D of this part
 applicable to a wing; and
   (b) Meet all requirements applicable to the function performed by these
 surfaces.

 [Doc. No. 25811, 56 FR 352, Jan. 3, 1991]

 *****************************************************************************


 56 FR 344, No. 2, Jan. 3, 1991

   SUMMARY: This final rule upgrades the airworthiness standards for normal,
 utility, acrobatic, and commuter category airplanes. This amendment provides
 airworthiness standards for advancements in technology being incorporated in
 current designs, permits type certification of spin resistant airplanes, and
 reduces the regulatory burden in showing compliance with some of the
 requirements for the design and type certification of small airplanes. These
 new and amended airworthiness standards also result in the need for new
 definitions. As a result, new definitions are added.

   DATES: February 4, 1991.

 *****************************************************************************






 Sec. 23.303  Factor of safety.

   Unless otherwise provided, a factor of safety of 1.5 must be used.






 Sec. 23.305  Strength and deformation.

   (a) The structure must be able to support limit loads without detrimental,
 permanent deformation. At any load up to limit loads, the deformation may not
 interfere with safe operation.
   (b) The structure must be able to support ultimate loads without failure
 for at least three seconds, except local failures or structural instabilities
 between limit and ultimate load are acceptable only if the structure can
 sustain the required ultimate load for at least three seconds. However when
 proof of strength is shown by dynamic tests simulating actual load
 conditions, the three second limit does not apply.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. No. 23-45, 58 FR 42160, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.307  Proof of structure.

   (a) Compliance with the strength and deformation requirements of Sec.
 23.305 must be shown for each critical load condition. Structural analysis
 may be used only if the structure conforms to those for which experience has
 shown this method to be reliable. In other cases, substantiating load tests
 must be made. Dynamic tests, including structural flight tests, are
 acceptable if the design load conditions have been simulated.
   (b) Certain parts of the structure must be tested as specified in Subpart D
 of this part.






                                 Flight Loads






 Sec. 23.321  General.

   (a) Flight load factors represent the ratio of the aerodynamic force
 component (acting normal to the assumed longitudinal axis of the airplane) to
 the weight of the airplane. A positive flight load factor is one in which the
 aerodynamic force acts upward, with respect to the airplane.
   (b) Compliance with the flight load requirements of this subpart must be
 shown--
   (1) At each critical altitude within the range in which the airplane may be
 expected to operate;
   (2) At each weight from the design minimum weight to the design maximum
 weight; and
   (3) For each required altitude and weight, for any practicable distribution
 of disposable load within the operating limitations specified in Secs.
 23.1583 through 23.1589.
   (c) When significant, the effects of compressibility must be taken into
 account.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended Amdt. No. 23-45, 58 FR 42160, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.331  Symmetrical flight conditions.

   (a) The appropriate balancing horizontal tail load must be accounted for in
 a rational or conservative manner when determining the wing loads and linear
 inertia loads corresponding to any of the symmetrical flight conditions
 specified in Secs. 23.333 through 23.341.
   (b) The incremental horizontal tail loads due to maneuvering and gusts must
 be reacted by the angular inertia of the airplane in a rational or
 conservative manner.
   (c) Mutual influence of the aerodynamic surfaces must be taken into account
 when determining flight loads.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-42, 56
 FR 352, Jan. 3, 1991]

 *****************************************************************************


 56 FR 344, No. 2, Jan. 3, 1991

   SUMMARY: This final rule upgrades the airworthiness standards for normal,
 utility, acrobatic, and commuter category airplanes. This amendment provides
 airworthiness standards for advancements in technology being incorporated in
 current designs, permits type certification of spin resistant airplanes, and
 reduces the regulatory burden in showing compliance with some of the
 requirements for the design and type certification of small airplanes. These
 new and amended airworthiness standards also result in the need for new
 definitions. As a result, new definitions are added.

   DATES: February 4, 1991.

 *****************************************************************************






 Sec. 23.333  Flight envelope.

   (a) General. Compliance with the strength requirements of this subpart must
 be shown at any combination of airspeed and load factor on and within the
 boundaries of a flight envelope (similar to the one in paragraph (d) of this
 section) that represents the envelope of the flight loading conditions
 specified by the maneuvering and gust criteria of paragraphs (b) and (c) of
 this section respectively.
   (b) Maneuvering envelope. Except where limited by maximum (static) lift
 coefficients, the airplane is assumed to be subjected to symmetrical
 maneuvers resulting in the following limit load factors:
   (1) The positive maneuvering load factor specified in Sec. 23.337 at speeds
 up to VD;
   (2) The negative maneuvering load factor specified in Sec. 23.337 at VC;
 and
   (3) Factors varying linearly with speed from the specified value at VC to
 0.0 at VD for the normal and commuter category, and --1.0 at VD for the
 acrobatic and utility categories.
   (c) Gust envelope. (1) The airplane is assumed to be subjected to
 symmetrical vertical gusts in level flight. The resulting limit load factors
 must correspond to the conditions determined as follows:
   (i) Positive (up) and negative (down) gusts of 50 f.p.s. at VC must be
 considered at altitudes between sea level and 20,000 feet. The gust velocity
 may be reduced linearly from 50 f.p.s. at 20,000 feet to 25 f.p.s. at 50,000
 feet.
   (ii) Positive and negative gusts of 25 f.p.s. at VD must be considered at
 altitudes between sea level and 20,000 feet. The gust velocity may be reduced
 linearly from 25 f.p.s. at 20,000 feet to 12.5 f.p.s. at 50,000 feet.
   (iii) In addition, for commuter category airplanes, positive (up) and
 negative (down) rough air gusts of 66 f.p.s. at VB must be considered at
 altitudes between sea level and 20,000 feet. The gust velocity may be reduced
 linearly from 66 f.p.s. at 20,000 feet to 38 f.p.s. at 50,000 feet.
   (2) The following assumptions must be made:
   (i) The shape of the gust is--

                             Ude              2(Pi)s
                        U =  ----   (1 - cos  ----     )
                             2                25C

 Where--

 s =Distance penetrated into gust (ft.);
 C =Mean geometric chord of wing (ft.); and
 Ude =Derived gust velocity referred to in subparagraph (1) of this section.

   (ii) Gust load factors vary linearly with speed between VC and VD .
   (d) Flight envelope.

                      [ ...Illustration appears here... ]

                               Flight Envelope

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13087, Aug. 13, 1969; Amdt. 23-34, 52 FR 1829, Jan. 15, 1987]






 Sec. 23.335  Design airspeeds.

   Except as provided in paragraph (a) (4) of this section, the selected
 design airspeeds are equivalent airspeeds (EAS).
   (a) Design cruising speed, VC. For VC the following apply:
   (1) VC (in knots) may not be less than--
   (i) 33 W/S (for normal, utility, and commuter category airplanes); and
   (ii) 36<radical>W/S (for acrobatic category airplanes).
   (2) For values of W/S more than 20, the multiplying factors may be
 decreased linearly with W/S to a value of 28.6 where W/S =100.
   (3) VC need not be more than 0.9 VH at sea level.
   (4) At altitudes where an MD is established, a cruising speed MC limited by
 compressibility may be selected.
   (b) Design dive speed VD. For VD, the following apply:
   (1) VD/MD may not be less than 1.25 VC/MC; and
   (2) With VC min, the required minimum design cruising speed, VD (in knots)
 may not be less than--
   (i) 1.40 Vc min (for normal and commuter category airplanes);
   (ii) 1.50 VC min (for utility category airplanes); and
   (iii) 1.55 VC min (for acrobatic category airplanes).
   (3) For values of W/S more than 20, the multiplying factors in paragraph
 (b)(2) of this section may be decreased linearly with W/S to a value of 1.35
 where W/S=100.
   (4) Compliance with paragraphs (b) (1) and (2) of this section need not be
 shown if VD/MD is selected so that the minimum speed margin between VC/MC and
 VD/MD is the greater of the following:
   (i) The speed increase resulting when, from the initial condition of
 stabilized flight at VC/MC,  the airplane is assumed to be upset, flown for
 20 seconds along a flight path 7.5 deg. below the initial path, and then
 pulled up with a load factor of 1.5 (0.5 g. acceleration increment). At least
 75 percent maximum continuous power for reciprocating engines, and maximum
 cruising power for turbines, or, if less, the power required for VC/MC for
 both kinds of engines, must be assumed until the pullup is initiated, at
 which point power reduction and pilot-controlled drag devices may be used.
   (ii) Mach 0.05 (at altitudes where an MD is established).
   (c) Design maneuvering speed VA. For VA,  the following applies:
   (1) VA may not be less than VS<radical>n where--
   (i) VS is a computed stalling speed with flaps retracted at the design
 weight, normally based on the maximum airplane normal force coefficients,
 CNA; and
   (ii) n is the limit maneuvering load factor used in design
   (2) The value of VA need not exceed the value of VC used in design.
   (d) Design speed for maximum gust intensity, VB. For VB, the following
 apply:
   (1) VB may not be less than the speed determined by the intersection of the
 line representing the maximum positive lift Cn max and the line representing
 the rough air gust velocity on the gust V-n diagram, or <radical>(ng) VS1,
 whichever is less, where:
   (i) ng the positive airplane gust load factor due to gust, at speed VC (in
 accordance with Sec. 23.341), and at the particular weight under
 consideration; and
   (ii) VS1 is the stalling speed with the flaps retracted at the particular
 weight under consideration.
   (2) VB need not be greater than VC.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13088, Aug. 13, 1969; Amdt. 23-16, 40 FR 2577, Jan. 14, 1975; Amdt. 23-34, 52
 FR 1829, Jan. 15, 1987; Amdt. 23-24, 52 FR 34745, Sept. 14, 1987]






 Sec. 23.337  Limit maneuvering load factors.

   (a) The positive limit maneuvering load factor n may not be less than--
   (1)

                            2.1+[24,000/(W+10,000)]

 for normal and commuter category airplanes, except that n need not be more
 than 3.8
   (2) 4.4 for utility category airplanes; or
   (3) 6.0 for acrobatic category airplanes.
   (b) The negative limit maneuvering load factor may not be less than--
   (1) 0.4 times the positive load factor for the normal utility and commuter
 categories; or
   (2) 0.5 times the positive load factor for the acrobatic category.
   (c) Maneuvering load factors lower than those specified in this section may
 be used if the airplane has design features that make it impossible to exceed
 these values in flight.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13088, Aug. 13, 1969; Amdt. 23-34, 52 FR 1829, Jan. 15, 1987]






 Sec. 23.341  Gust loads factors.

   (a) The gust load for a canard or tandem wing configuration must be
 computed using a rational analysis, considering the criteria of Sec.
 23.333(c), to develop the gust loading on each lifting surface or may be
 computed in accordance with paragraph (b) of this section provided that the
 resulting net loads are shown to be conservative with respect to the gust
 criteria of Sec. 23.333(c).
   (b) In the absence of a more rational analysis for conventional
 configurations, the gust load factors must be computed as follows:

                                      KgUdeVa
                             n = 1 +  ------------
                                      498(W/S)

 Where--

 Kg=0.88micro-g/5.3+micro-g=gust alleviation factor;
 micro-g=2(W/S)/<rho>Cag=airplane mass ratio;
 Ude=Derived gust velocities referred to in Sec. 23.333(c) (f.p.s.);
 <rho>=Density of air (slugs/cu.ft.);
 W/S =Wing loading (p.s.f.);
 C =Mean geometric chord (ft.);
 g =Acceleration due to gravity (ft./sec.**2)
 V =Airplane equivalent speed (knots); and
 a =Slope of the airplane normal force coefficient curve CNA per radian if the
     gust loads are applied to the wings and horizontal tail surfaces
     simultaneously by a rational method. The wing lift curve slope CL per
     radian may be used when the gust load is applied to the wings only and
     the horizontal tail gust loads are treated as a separate condition.

 [Amdt. 23-7, 34 FR 13088, Aug. 13, 1969, as amended by Amdt. 23-42, 56 FR
 352, Jan. 3, 1991]

 *****************************************************************************


 56 FR 344, No. 2, Jan. 3, 1991

   SUMMARY: This final rule upgrades the airworthiness standards for normal,
 utility, acrobatic, and commuter category airplanes. This amendment provides
 airworthiness standards for advancements in technology being incorporated in
 current designs, permits type certification of spin resistant airplanes, and
 reduces the regulatory burden in showing compliance with some of the
 requirements for the design and type certification of small airplanes. These
 new and amended airworthiness standards also result in the need for new
 definitions. As a result, new definitions are added.

   DATES: February 4, 1991.

 *****************************************************************************






 Sec. 23.345  High lift devices.

   (a) If flaps or similar high lift devices to be used for takeoff, approach,
 or landing are installed, the airplane, with the flaps fully deflected at VF,
 is assumed to be subjected to symmetrical maneuvers and gusts resulting in
 limit load factors within the range determined by--
   (1) Maneuvering, to a positive limit load factor of 2.0; and
   (2) Positive and negative gust of 25 feet per second acting normal to the
 flight path in level flight.
   (b) VF must be assumed to be not less than 1.4 VS or 1.8 VSF, whichever is
 greater, where--

 VS is the computed stalling speed with flaps retracted at the design weight;
   and
 VSF is the computed stalling speed with flaps fully extended at the design
   weight.

 However, if an automatic flap load limiting device is used, the airplane may
 be designed for the critical combinations of airspeed and flap position
 allowed by that device.
   (c) In designing the flaps and supporting structures, the following must be
 accounted for:
   (1) A head-on gust having a velocity of 25 feet per second (EAS).
   (2) The slipstream effects specified in Sec. 23.457(b).
   (d) In determining external loads on the airplane as a whole, thrust,
 slipstream, and pitching acceleration may be assumed to be zero.
   (e) The requirements of Sec. 23.457, and this section may be complied with
 separately or in combination.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13088, Aug. 13, 1969; Amdt. 23-23, 43 FR 50592, Oct. 30, 1978]






 Sec. 23.347  Unsymmetrical flight conditions.

   The airplane is assumed to be subjected to the unsymmetrical flight
 conditions of Secs. 23.349 and 23.351. Unbalanced aerodynamic moments about
 the center of gravity must be reacted in a rational or conservative manner,
 considering the principal masses furnishing the reacting inertia forces.






 Sec. 23.349  Rolling conditions.

   The wing and wing bracing must be designed for the following loading
 conditions:
   (a) Unsymmetrical wing loads appropriate to the category. Unless the
 following values result in unrealistic loads, the rolling accelerations may
 be obtained by modifying the symmetrical flight conditions in Sec. 23.333(d)
 as follows:
   (1) For the acrobatic category, in conditions A and F, assume that 100
 percent of the semispan wing airload acts on one side of the plane of
 symmetry and 60 percent of this load acts on the other side.
   (2) For normal, utility, and commuter categories, in Condition A, assume
 that 100 percent of the semispan wing airload acts on one side of the
 airplane, and 70 percent of this load acts on the other side. For airplanes
 of more than 1,000 pounds design weight, the latter percentage may be
 increased linearly with weight up through 75 percent at 12,500 pounds to the
 maximum gross weight of the airplane.
   (b) The loads resulting from the aileron deflections and speeds specified
 in Sec. 23.455, in combination with an airplane load factor of at least two
 thirds of the positive maneuvering load factor used for design. Unless the
 following values result in unrealistic loads, the effect of aileron
 displacement on wing torsion may be accounted for by adding the following
 increment to the basic airfoil moment coefficient over the aileron portion of
 the span in the critical condition determined in Sec. 23.333(d):

   <Delta>cm=--0.01<delta>

 where--

 <Delta>cm is the moment coefficient increment; and
 <delta> is the down aileron deflection in degrees in the critical condition.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13088, Aug. 13, 1969; Amdt. 23-34, 52 FR 1829, Jan. 15, 1987]






 Sec. 23.351  Yawing conditions.

   The airplane must be designed for yawing loads on the vertical surfaces
 resulting from the loads specified in Secs. 23.441 through 23.445.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-42, 56
 FR 352, Jan. 3, 1991]

 *****************************************************************************


 56 FR 344, No. 2, Jan. 3, 1991

   SUMMARY: This final rule upgrades the airworthiness standards for normal,
 utility, acrobatic, and commuter category airplanes. This amendment provides
 airworthiness standards for advancements in technology being incorporated in
 current designs, permits type certification of spin resistant airplanes, and
 reduces the regulatory burden in showing compliance with some of the
 requirements for the design and type certification of small airplanes. These
 new and amended airworthiness standards also result in the need for new
 definitions. As a result, new definitions are added.

   DATES: February 4, 1991.

 *****************************************************************************






 Sec. 23.361  Engine torque.

   (a) Each engine mount and its supporting structure must be designed for the
 effects of--
   (1) A limit engine torque corresponding to takeoff power and propeller
 speed acting simultaneously with 75 percent of the limit loads from flight
 condition A of Sec. 23.333(d);
   (2) A limit engine torque corresponding to maximum continuous power and
 propeller spped acting simultaneously with the limit loads
 from flight condition A of Sec. 23.333(d); and
   (3) For turbopropeller installations, in addition to the conditions
 specified in paragraphs (a)(1) and (a)(2) of this section, a limit engine
 torque corresponding to takeoff power and propeller speed, multiplied by a
 factor accounting for propeller control system malfunction, including quick
 feathering, acting simultaneously with lg level flight loads. In the absence
 of a rational analysis, a factor of 1.6 must be used.
   (b) For turbine engine installations, the engine mounts and supporting
 structure must be designed to withstand each of the following:
   (1) A limit engine torque load imposed by sudden engine stoppage due to
 malfunction or structural failure (such as compressor jamming).
   (2) A limit engine torque load imposed by the maximum acceleration of the
 engine.
   (c) The limit engine torque to be considered under paragraph (a) of this
 section must be obtained by multiplying the mean torque by a factor of--
   (1) 1.25 for turbopropeller installations;
   (2) 1.33 for engines with five or more cylinders; and
   (3) Two, three, or four, for engines with four, three, or two cylinders,
 respectively.

 [Amdt. 23-26, 45 FR 60171, Sept. 11, 1980, as amended by Amdt. No. 23-45, 58
 FR 42160, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.363  Side load on engine mount.

   (a) Each engine mount and its supporting structure must be designed for a
 limit load factor in a lateral direction, for the side load on the engine
 mount, of not less than--
   (1) 1.33, or
   (2) One-third of the limit load factor for flight condition A.
   (b) The side load prescribed in paragraph (a) of this section may be
 assumed to be independent of other flight conditions.






 Sec. 23.365  Pressurized cabin loads.

   For each pressurized compartment, the following apply:
   (a) The airplane structure must be strong enough to withstand the flight
 loads combined with pressure differential loads from zero up to the maximum
 relief valve setting.
   (b) The external pressure distribution in flight, and any stress
 concentrations, must be accounted for.
   (c) If landings may be made with the cabin pressurized, landing loads must
 be combined with pressure differential loads from zero up to the maximum
 allowed during landing.
   (d) The airplane structure must be strong enough to withstand the pressure
 differential loads corresponding to the maximum relief valve setting
 multiplied by a factor of 1.33, omitting other loads.
   (e) If a pressurized cabin has two or more compartments separated by
 bulkheads or a floor, the primary structure must be designed for the effects
 of sudden release of pressure in any compartment with external doors or
 windows. This condition must be investigated for the effects of failure of
 the largest opening in the compartment. The effects of intercompartmental
 venting may be considered.






 Sec. 23.367  Unsymmetrical loads due to engine failure.

   (a) Turbopropeller airplanes must be designed for the unsymmetrical loads
 resulting from the failure of the critical engine including the following
 conditions in combination with a single malfunction of the propeller drag
 limiting system, considering the probable pilot corrective action on the
 flight controls:
   (1) At speeds between VMC and VD, the loads resulting from power failure
 because of fuel flow interruption are considered to be limit loads.
   (2) At speeds between VMC and VC, the loads resulting from the
 disconnection of the engine compressor from the turbine or from loss of the
 turbine blades are considered to be ultimate loads.
   (3) The time history of the thrust decay and drag buildup occurring as a
 result of the prescribed engine failures must be substantiated by test or
 other data applicable to the particular engine-propeller combination.
   (4) The timing and magnitude of the probable pilot corrective action must
 be conservatively estimated, considering the characteristics of the
 particular engine-propeller-airplane combination.
   (b) Pilot corrective action may be assumed to be initiated at the time
 maximum yawing velocity is reached, but not earlier than 2 seconds after the
 engine failure. The magnitude of the corrective action may be based on the
 limit pilot forces specified in Sec. 23.397 except that lower forces may be
 assumed where it is shown by analysis or test that these forces can control
 the yaw and roll resulting from the prescribed engine failure conditions.

 [Amdt. 23-7, 34 FR 13089, Aug. 13, 1969]






 Sec. 23.369  Rear lift truss.

   (a) If a rear lift truss is used, it must be designed for conditions of
 reversed airflow at a design speed of--

                       V=8.7 <radical>W/S + 8.7 (knots)

   (b) Either aerodynamic data for the particular wing section used, or a
 value of CL equalling -0.8 with a chordwise distribution that is triangular
 between a peak at the trailing edge and zero at the leading edge, must be
 used.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13089, Aug. 13, 1969; 34 FR 17509, Oct. 30, 1969; Amdt. No. 23-45, 58 FR
 42160, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.371  Gyroscopic and areodynamic loads.

   For turbine-powered airplanes, each engine mount and its supporting
 structure must be designed for the combined gyroscopic and aerodynamic
 loads that result, with the
 engines at maximum continuous r.p.m., under either of the following
 conditions:
   (a) The conditions prescribed in Secs. 23.351 and 23.423.
   (b) All possible combinations of the following:
   (1) A yaw velocity of 2.5 radians per second.
   (2) A pitch velocity of 1 radian per second.
   (3) A normal load factor of 2.5.
   (4) Maximum continuous thrust.

 [Amdt. 23-7, 34 FR 13089, Aug. 13, 1969, as amended by Amdt. 23-26, 45 FR
 60171, Sept. 11, 1980; Amdt. No. 23-45, 58 FR 42160, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.373  Speed control devices.

   If speed control devices (such as spoilers and drag flaps) are incorporated
 for use in enroute conditions--
   (a) The airplane must be designed for the symmetrical maneuvers and gusts
 prescribed in Secs. 23.333, 23.337, and 23.341, and the yawing maneuvers and
 lateral gusts in Secs. 23.441 and 23.443, with the device extended at speeds
 up to the placard device extended speed; and
   (b) If the device has automatic operating or load limiting features, the
 airplane must be designed for the maneuver and gust conditions prescribed in
 paragraph (a) of this section at the speeds and corresponding device
 positions that the mechanism allows.

 [Amdt. 23-7, 34 FR 13089, Aug. 13, 1969]






                       Control Surface and System Loads






 Sec. 23.391  Control surface loads.

   (a) The control surface loads specified in Secs. 23.397 through 23.459 are
 assumed to occur in the conditions described in Secs. 23.331 through 23.351.
   (b) If allowed by the following sections, the values of control surface
 loading in Appendix B of this part may be used, instead of particular control
 surface data, to determine the detailed rational requirements of Secs. 23.397
 through 23.459, unless these values result in unrealistic loads.






 Sec. 23.395  Control system loads.

   (a) Each flight control system and its supporting structure must be
 designed for loads corresponding to at least 125 percent of the computed
 hinge moments of the movable control surface in the conditions prescribed in
 Secs. 23.391 through 23.459. In addition, the following apply:
   (1) The system limit loads need not exceed the higher of the loads that can
 be produced by the pilot and automatic devices operating the controls.
 However, autopilot forces need not be added to pilot forces. The system must
 be designed for the maximum effort of the pilot or autopilot, whichever is
 higher. In addition, if the pilot and the autopilot act in opposition, the
 part of the system between them may be designed for the maximum effort of the
 one that imposes the lesser load. Pilot forces used for design need not
 exceed the maximum forces prescribed in Sec. 23.397(b).
   (2) The design must, in any case, provide a rugged system for service use,
 considering jamming, ground gusts, taxiing downwind, control inertia, and
 friction. Compliance with this subparagraph may be shown by designing for
 loads resulting from application of the minimum forces prescribed in Sec.
 23.397(b).
   (b) A 125 percent factor on computed hinge moments must be used to design
 elevator, aileron, and rudder systems. However, a factor as low as 1.0 may be
 used if hinge moments are based on accurate flight test data, the exact
 reduction depending upon the accuracy and reliability of the data.
   (c) Pilot forces used for design are assumed to act at the appropriate
 control grips or pads as they would in flight, and to react at the
 attachments of the control system to the control surface horns.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13089, Aug. 13, 1969]






 Sec. 23.397  Limit control forces and torques.

   (a) In the control surface flight loading condition, the airloads on
 movable surfaces and the corresponding deflections need not exceed those that
 would result in flight from the application of any pilot force within the
 ranges specified in paragraph (b) of this section. In applying this
 criterion, the effects of control system boost and servo-mechanisms, and the
 effects of tabs must be considered. The automatic pilot effort must be used
 for design if it alone can produce higher control surface loads than the
 human pilot.
   (b) The limit pilot forces and torques are as follows:

                                      Maximum forces
                                      or torques for
                                      design weight,
                                     weight equal to
                                       or less than     Minimum forces
                  Control            5,000 pounds /1/   or torques /2/

         Aileron:
          Stick                      67 lbs            40 lbs.
          Wheel /3/                  50 D in.-lbs /4/  40 D in.-lbs./4/
         Elevator:
          Stick                      167 lbs           100 lbs.
          Wheel (symmetrical)        200 lbs           100 lbs.
          Wheel (unsymmetrical) /5/                    100 lbs.
         Rudder                      200 lbs           150 lbs.

         /1/ For design weight (W) more than 5,000 pounds, the
         specified maximum values must be increased linearly with
         weight to 1.18 times the specified values at a design weight
         of 12,500 pounds and for commuter category airplanes, the
         specified values must be increased linearly with weight to
         1.35 times the specified values at a design weight of 19,000
         pounds.

         /2/ If the design of any individual set of control systems or
         surfaces makes these specified minimum forces or torques
         inapplicable, values corresponding to the present hinge
         moments obtained under Sec. 23.415, but not less than 0.6 of
         the specified minimum forces or torques, may be used.

         /3/ The critical parts of the aileron control system must also
         be designed for a single tangential force with a limit value
         of 1.25 times the couple force determined from the above
         criteria.

         /4/ D=wheel diameter (inches).

         /5/ The unsymmetrical force must be applied at one of the
         normal handgrip points on the control wheel.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13089, Aug. 13, 1969; Amdt. 23-17, 41 FR 55464, Dec. 20, 1976; Amdt. 23-34,
 52 FR 1829, Jan. 15, 1987; Amdt. No. 23-45, 58 FR 42160, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.399  Dual control system.

   Each dual control system must be designed for the pilots operating in
 opposition, using individual pilot forces not less than--
   (a) 0.75 times those obtained under Sec. 23.395; or
   (b) The minimum forces specified in Sec. 23.397(b).






 Sec. 23.405  Secondary control system.

   Secondary controls, such as wheel brakes, spoilers, and tab controls, must
 be designed for the maximum forces that a pilot is likely to apply to those
 controls.






 Sec. 23.407  Trim tab effects.

   The effects of trim tabs on the control surface design conditions must be
 accounted for only where the surface loads are limited by maximum pilot
 effort. In these cases, the tabs are considered to be deflected in the
 direction that would assist the pilot. These deflections must correspond to
 the maximum degree of "out of trim" expected at the speed for the condition
 under consideration.






 Sec. 23.409  Tabs.

   Control surface tabs must be designed for the most severe combination of
 airspeed and tab deflection likely to be obtained within the flight envelope
 for any usable loading condition.






 Sec. 23.415  Ground gust conditions.

   (a) The control system must be investigated as follows for control surface
 loads due to ground gusts and taxiing downwind:
   (1) If an investigation of the control system for ground gust loads is not
 required by paragraph (a)(2) of this section, but the applicant elects to
 design a part of the control system of these loads, these loads need only be
 carried from control surface horns through the nearest stops or gust locks
 and their supporting structures.
   (2) If pilot forces less than the minimums specified in Sec. 23.397(b) are
 used for design, the effects of surface loads due to ground gusts and taxiing
 downwind must be investigated for the entire control system according to the
 formula:

                                    H=KcSq

 where--

 H =limit hinge moment (ft.-lbs.);
 c =means chord of the control surface aft of the hinge line (ft.);
 S =area of control surface aft of the hinge line (sq. ft.);
 q =Dynamic pressure (p.s.f.) based on a design speed not less than
     14.6<radical>W/S+14.6 (f.p.s.) except that the design speed need not
     exceed 88 (f.p.s.); and
 K =limit hinge moment factor for ground gusts derived in paragraph (b) of
     this section. (For ailerons and elevators, a positive value of K
     indicates a moment tending to depress the surface and a negative value of
     K indicates a moment tending to raise the surface).

   (b) The limit hinge moment factor K  for ground gusts must be derived as
 follows:

   Surface        K                      Position of controls

 (a) Aileron      0.75  Control column locked lashed in mid-position.
 (b) Aileron   +/-0.50  Ailerons at full throw; + moment on one aileron, -
                         moment on the other.
 (c) Elevator  +/-0.75  (c) Elevator full up (-).
 (d) Elevator           (d) Elevator full down (+).
 (e) Rudder    +/-0.75  (e) Rudder in neutral.
 (f) Rudder             (f) Rudder at full throw.
   (c) The tie-down attachment fittings and the surrounding structure must be
 designed for limit load conditions resulting from wind speeds up to 65 knots
 horizontally from any direction for the weight determined to be critical for
 tie-down.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13089, Aug. 13, 1969; Amdt. No. 23-45, 58 FR 42160, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






                 Horizontal Stabilizing and Balancing Surfaces






 Sec. 23.421  Balancing loads.

   (a) A horizontal surface balancing load is a load necessary to maintain
 equilibrium in any specified flight condition with no pitching acceleration.
   (b) Horizontal balancing surfaces must be designed for the balancing loads
 occurring at any point on the limit maneuvering envelope and in the flap
 conditions specified in Sec. 23.345.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13089, Aug. 13, 1969; Amdt. 23-42, 56 FR 352, Jan. 3, 1991]

 *****************************************************************************


 56 FR 344, No. 2, Jan. 3, 1991

   SUMMARY: This final rule upgrades the airworthiness standards for normal,
 utility, acrobatic, and commuter category airplanes. This amendment provides
 airworthiness standards for advancements in technology being incorporated in
 current designs, permits type certification of spin resistant airplanes, and
 reduces the regulatory burden in showing compliance with some of the
 requirements for the design and type certification of small airplanes. These
 new and amended airworthiness standards also result in the need for new
 definitions. As a result, new definitions are added.

   DATES: February 4, 1991.

 *****************************************************************************






 Sec. 23.423  Maneuvering loads.

   Each horizontal surface and its supporting structure, and the main wing of
 a canard or tandem wing configuration, if that surface has pitch control,
 must be designed for the maneuvering loads imposed by the following
 conditions:
   (a) A sudden movement of the pitching control, at the speed VA, to the
 maximum aft movement, and the maximum forward movement, as limited by the
 control stops, or pilot effort, whichever is critical.
   (b) A sudden aft movement of the pitching control at speeds above VA,
 followed by a forward movement of the pitching control resulting in the
 following combinations of normal and angular acceleration:

                                      Normal         Angular
                                   acceleration    acceleration
                   Condition           (n)        (radian/sec**2)

               Nose-up pitching    1.0           +39nm.Vx(nm-1.5)
               Nose-down ptiching  nm            -39nm.Vx(nm-1.5)

 where--
   (1) nm=positive limit maneuvering load factor used in the design of the
 airplane; and
   (2) V=initial speed in knots.
   The conditions in this paragraph involve loads corresponding to the loads
 that may occur in a "checked maneuver" (a maneuver in which the pitching
 control is suddenly displaced in one direction and then suddenly moved in the
 opposite direction). The deflections and timing of the "checked maneuver"
 must avoid exceeding the limit maneuvering load factor. The total horizontal
 surface load for both nose-up and nose-down pitching conditions is the sum of
 the balancing loads at V and the specified value of the normal load factor n,
 plus the maneuvering load increment due to the specified value of the angular
 acceleration.

 [Doc. No. 25811, 56 FR 353, Jan. 3, 1991; Amdt. 23-42, 56 FR 5455, Feb. 11,
 1991]

 *****************************************************************************


 56 FR 344, No. 2, Jan. 3, 1991

   SUMMARY: This final rule upgrades the airworthiness standards for normal,
 utility, acrobatic, and commuter category airplanes. This amendment provides
 airworthiness standards for advancements in technology being incorporated in
 current designs, permits type certification of spin resistant airplanes, and
 reduces the regulatory burden in showing compliance with some of the
 requirements for the design and type certification of small airplanes. These
 new and amended airworthiness standards also result in the need for new
 definitions. As a result, new definitions are added.

   DATES: February 4, 1991.

 *****************************************************************************






 Sec. 23.425  Gust loads.

   (a) Each horizontal surface, other than a main wing, must be designed for
 loads resulting from--
   (1) Gust velocities specified in Sec. 23.333(c) with flaps retracted; and
   (2) Positive and negative gusts of 25 f.p.s. nominal intensity at VF
 corresponding to the flight conditions specified in Sec. 23.345(a)(2).
   (b) [Reserved]
   (c) When determining the total load on the horizontal surfaces for the
 conditions specified in paragraph (a) of this section, the initial balancing
 loads for steady unaccelerated flight at the pertinent design speeds VF, VC,
 and VD must first be determined. The incremental load resulting from the
 gusts must be added to the initial balancing load to obtain the total load.
   (d) In the absence of a more rational analysis, the incremental load due to
 the gust must be computed as follows only on airplane configurations with
 aft-mounted, horizontal surfaces, unless its use elsewhere is shown to be
 conservative:

                              Kg Ude V<alpha>ht Sht         de
                <Delta>Lht =  ---------------------    (1-  -- )
                                     498                    da

 where--

 <Delta>Lht=Incremental horizontal tailload (lbs.);
 Kg=Gust alleviation factor defined in Sec. 23.341;
 Ude=Derived gust velocity (f.p.s.);
 V=Airplane equivalent speed (knots);
 <alpha>ht=Slope of aft horizontal lift curve (per radian);
 Sht=Area of aft horizontal lift surface (ft**2); and

                              de
                         (1-  --  ) = Downwash factor
                              da

 [Doc. No. 4080, 20 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13089 Aug. 13, 1969; Amdt. 23-42, 56 FR 353, Jan. 3, 1991]

 *****************************************************************************


 56 FR 344, No. 2, Jan. 3, 1991

   SUMMARY: This final rule upgrades the airworthiness standards for normal,
 utility, acrobatic, and commuter category airplanes. This amendment provides
 airworthiness standards for advancements in technology being incorporated in
 current designs, permits type certification of spin resistant airplanes, and
 reduces the regulatory burden in showing compliance with some of the
 requirements for the design and type certification of small airplanes. These
 new and amended airworthiness standards also result in the need for new
 definitions. As a result, new definitions are added.

   DATES: February 4, 1991.

 *****************************************************************************






 Sec. 23.427  Unsymmetrical loads.

   (a) Horizontal surfaces other than main wing and their supporting
 structure must be designed for unsymmetrical loads arising from yawing and
 slipstream effects, in combination with the loads prescribed for the flight
 conditions set forth in Secs. 23.421 through 23.425.
   (b) In the absence of more rational data for airplanes that are
 conventional in regard to location of engines, wings, horizontal surfaces
 other than main wing, and fuselage shape:
   (1) 100 percent of the maximum loading from the symmetrical flight
 conditions may be assumed on the surface on one side of the plane of
 symmetry; and
   (2) The following percentage of that loading must be applied to the
 opposite side:

   Percent=100-10 (n-1), where n is the specified positive maneuvering load
 factor, but this value may not be more than 80 percent.

   (c) For airplanes that are not conventional (such as airplanes with
 horizontal surfaces other than main wing having appreciable dihedral or
 supported by the vertical tail surfaces) the surfaces and supporting
 structures must be designed for combined vertical and horizontal surface
 loads resulting from each prescribed flight condition taken separately.

 [Amdt. 23-14, 38 FR 31820, Nov. 19, 1973, as amended by Amdt. 23-42, 56 FR
 353, Jan. 3, 1991]

 *****************************************************************************


 56 FR 344, No. 2, Jan. 3, 1991

   SUMMARY: This final rule upgrades the airworthiness standards for normal,
 utility, acrobatic, and commuter category airplanes. This amendment provides
 airworthiness standards for advancements in technology being incorporated in
 current designs, permits type certification of spin resistant airplanes, and
 reduces the regulatory burden in showing compliance with some of the
 requirements for the design and type certification of small airplanes. These
 new and amended airworthiness standards also result in the need for new
 definitions. As a result, new definitions are added.

   DATES: February 4, 1991.

 *****************************************************************************






                            Vertical Surfaces






 Sec. 23.441  Maneuvering loads.

   (a) At speeds up to VA, the vertical surfaces must be designed to
 withstand the following conditions. In computing the loads, the yawing
 velocity may be assumed to be zero:
   (1) With the airplane in unaccelerated flight at zero yaw, it is assumed
 that the rudder control is suddenly displaced to the maximum deflection, as
 limited by the control stops or by limit pilot forces.
   (2) With the rudder deflected as specified in paragraph (a)(1) of this
 section, it is assumed that the airplane yaws to the resulting sideslip
 angle. In lieu of a rational analysis, an overswing angle equal to 1.3 times
 the static sideslip angle of paragraph (a)(3) of this section may be assumed.
   (3) A yaw angle of 15 degrees with the rudder control maintained in the
 neutral position (except as limited by pilot strength).
   (b) [Reserved.]
   (c) The yaw angles specified in paragraph (a)(3) of this section may be
 reduced if the yaw angle chosen for a particular speed cannot be exceeded
 in--
   (1) Steady slip conditions;
   (2) Uncoordinated rolls from steep banks; or
   (3) Sudden failure of the critical engine with delayed corrective action.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13090, Aug. 13, 1969; Amdt. 23-14, 38 FR 31821, Nov. 19, 1973; Amdt. 23-28,
 47 FR 13315, Mar. 29, 1982; Amdt. 23-42, 56 FR 353, Jan. 3, 1991]

 *****************************************************************************


 56 FR 344, No. 2, Jan. 3, 1991

   SUMMARY: This final rule upgrades the airworthiness standards for normal,
 utility, acrobatic, and commuter category airplanes. This amendment provides
 airworthiness standards for advancements in technology being incorporated in
 current designs, permits type certification of spin resistant airplanes, and
 reduces the regulatory burden in showing compliance with some of the
 requirements for the design and type certification of small airplanes. These
 new and amended airworthiness standards also result in the need for new
 definitions. As a result, new definitions are added.

   DATES: February 4, 1991.

 *****************************************************************************






 Sec. 23.443  Gust loads.

   (a) Vertical surfaces must be designed to withstand, in unaccelerated
 flight at speed VC, lateral gusts of the values prescribed for VC in Sec.
 23.333(c).
   (b) In addition, for commuter category airplanes, the airplane is assumed
 to encounter derived gusts normal to the plane of symmetry while in
 unaccelerated flight at VB, VC, VD, and VF. The derived gusts and airplane
 speeds corresponding to these conditions, as determined by Secs. 23.341 and
 23.345, must be investigated. The shape of the gust must be as specified in
 Sec. 23.333(c)(2)(i).
   (c) In the absence of a more rational analysis, the gust load must be
 computed as follows:

                             Lvt=KgtUdeV<alpha>vtSvt
                             -----------------------
                                     498
 where--

 Lvt=Vertical surface load (lbs.);
 Kgt=0.88<micro>gt/5.3+<micro>gt=gust alleviation factor;
 <micro>gt=2W/PCtg<alpha>vtSvt(K/1t)**2=lateral mass ratio;
 Ude=Derived gust velocity (f.p.s.);
 P=Air density (slugs/cu.ft.);
 W =Airplane weight (lbs.);
 Svt=Area of vertical surface (ft.**2);
 Ct=Mean geometric chord of vertical surface (ft.);
 <alpha>vt=Lift curve slope of vertical surface (per radian);
 K =Radius of gyration in yaw (ft.);
 1t=Distance from airplane c.g. to lift center of vertical surface (ft.);
 g =Acceleration due to gravity (ft./sec.**2); and
 V =Airplane equivalent speed (knots).

 [Amdt. 23-7, 34 FR 13090, Aug. 13, 1969, as amended by Amdt. 23-34, 52 FR
 1830, Jan. 15, 1987; 52 FR 7262, Mar. 9, 1987; Amdt. 23-24, 52 FR 34745,
 Sept. 14, 1987; Amdt. 23-42, 56 FR 353, Jan. 3, 1991]

 *****************************************************************************


 56 FR 344, No. 2, Jan. 3, 1991

   SUMMARY: This final rule upgrades the airworthiness standards for normal,
 utility, acrobatic, and commuter category airplanes. This amendment provides
 airworthiness standards for advancements in technology being incorporated in
 current designs, permits type certification of spin resistant airplanes, and
 reduces the regulatory burden in showing compliance with some of the
 requirements for the design and type certification of small airplanes. These
 new and amended airworthiness standards also result in the need for new
 definitions. As a result, new definitions are added.

   DATES: February 4, 1991.

 *****************************************************************************






 Sec. 23.445  Outboard fins or winglets.

   (a) If outboard fins or winglets are included on the horizontal surfaces or
 wings, the horizontal surfaces or wings must be designed for their maximum
 load in combination with loads induced by the fins or winglets and moments or
 forces exerted on the horizontal surfaces or wings by the fins or winglets.
   (b) If outboard fins or winglets extend above and below the horizontal
 surface, the critical vertical surface loading (the load per unit area as
 determined under Secs. 23.441 and 23.443) must be applied to--
   (1) The part of the vertical surfaces above the horizontal surface with 80
 percent of that loading applied to the part below the horizontal surface; and
   (2) The part of the vertical surfaces below the horizontal surface with 80
 percent of that loading applied to the part above the horizontal surface.
   (c) The end plate effects of outboard fins or winglets must be taken
 into account in applying the yawing conditions of Secs. 23.441 and 23.443 to
 the vertical surfaces in paragraph (b) of this section.
   (d) When rational methods are used for computing loads, the maneuvering
 loads of Sec. 23.441 on the vertical surfaces and the one-g horizontal
 surface load, including induced loads on the horizontal surface and moments
 or forces exerted on the horizontal surfaces by the vertical surfaces, must
 be applied simultaneously for the structural loading condition.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-14, 38 FR
 31821, Nov. 19, 1973; Amdt. 23-42, 56 FR 353, Jan. 3, 1991]

 *****************************************************************************


 56 FR 344, No. 2, Jan. 3, 1991

   SUMMARY: This final rule upgrades the airworthiness standards for normal,
 utility, acrobatic, and commuter category airplanes. This amendment provides
 airworthiness standards for advancements in technology being incorporated in
 current designs, permits type certification of spin resistant airplanes, and
 reduces the regulatory burden in showing compliance with some of the
 requirements for the design and type certification of small airplanes. These
 new and amended airworthiness standards also result in the need for new
 definitions. As a result, new definitions are added.

   DATES: February 4, 1991.

 *****************************************************************************






                   Ailerons, Wing Flaps, and Special Devices






 Sec. 23.455  Ailerons.

   (a) The ailerons must be designed for the loads to which they are
 subjected--
   (1) In the neutral position during symmetrical flight conditions; and
   (2) By the following deflections (except as limited by pilot effort),
 during unsymmetrical flight conditions:
   (i) Sudden maximum displacement of the aileron control at VA. Suitable
 allowance may be made for control system deflections.
   (ii) Sufficient deflection at VC, where VC is more than VA, to produce a
 rate of roll not less than obtained in paragraph (a)(2)(i) of this section.
   (iii) Sufficient deflection at VD to produce a rate of roll not less than
 one-third of that obtained in paragraph (a)(2)(i) of this section.
   (b) [Reserved]

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13090, Aug. 13, 1969; Amdt. 23-42, 56 FR 353, Jan. 3, 1991]

 *****************************************************************************


 56 FR 344, No. 2, Jan. 3, 1991

   SUMMARY: This final rule upgrades the airworthiness standards for normal,
 utility, acrobatic, and commuter category airplanes. This amendment provides
 airworthiness standards for advancements in technology being incorporated in
 current designs, permits type certification of spin resistant airplanes, and
 reduces the regulatory burden in showing compliance with some of the
 requirements for the design and type certification of small airplanes. These
 new and amended airworthiness standards also result in the need for new
 definitions. As a result, new definitions are added.

   DATES: February 4, 1991.

 *****************************************************************************






 Sec. 23.457  Wing flaps.

   (a) The wing flaps, their operating mechanisms, and their supporting
 structures must be designed for critical loads occurring in the flaps-
 extended flight conditions with the flaps in any position. However, if an
 automatic flap load limiting device is used, these components may be designed
 for the critical combinations of airspeed and flap position allowed by that
 device.
   (b) The effects of propeller slipstream, corresponding to takeoff power,
 must be taken into account at not less than 1.4 VS, where VS is the computed
 stalling speed with flaps fully retracted at the design weight. For the
 investigation of slipstream effects, the load factor may be assumed to be
 1.0.






 Sec. 23.459  Special devices.

   The loading for special devices using aerodynamic surfaces (such as slots
 and spoilers) must be determined from test data.






                                 Ground Loads






 Sec. 23.471  General.

   The limit ground loads specified in this subpart are considered to be
 external loads and inertia forces that act upon an airplane structure. In
 each specified ground load condition, the external reactions must be placed
 in equilibrium with the linear and angular inertia forces in a rational or
 conservative manner.






 Sec. 23.473  Ground load conditions and assumptions.

   (a) The ground load requirements of this subpart must be complied with at
 the design maximum weight except that Secs. 23.479, 23.481, and 23.483 may be
 complied with at a design landing weight (the highest weight for landing
 conditions at the maximum descent velocity) allowed under paragraphs (b) and
 (c) of this section.
   (b) The design landing weight may be as low as--
   (1) 95 percent of the maximum weight if the minimum fuel capacity is enough
 for at least one-half hour of operation at maximum continuous power plus a
 capacity equal to a fuel weight which is the difference between the design
 maximum weight and the design landing weight; or
   (2) The design maximum weight less the weight of 25 percent of the total
 fuel capacity.
   (c) The design landing weight of a multiengine airplane may be less than
 that allowed under paragraph (b) of this section if--
   (1) The airplane meets the one-engine-inoperative climb requirements of
 Sec. 23.67 (a) or (b)(1); and
   (2) Compliance is shown with the fuel jettisoning system requirements of
 Sec. 23.1001.
   (d) The selected limit vertical inertia load factor at the center of
 gravity of the airplane for the ground load conditions prescribed in this
 subpart may not be less than that which would be obtained when landing with a
 descent velocity (V), in feet per second, equal to 4.4 (W/S) 1/4, except that
 this velocity need not be more than 10 feet per second and may not be less
 than seven feet per second.
   (e) Wing lift not exceeding two-thirds of the weight of the airplane may be
 assumed to exist throughout the landing impact and to act through the center
 of gravity. The ground reaction load factor may be equal to the inertia load
 factor minus the ratio of the above assumed wing lift to the airplane weight.
   (f) Energy absorption tests (to determine the limit load factor
 corresponding to the required limit descent velocities) must be made under
 Sec. 23.723(a) unless specifically exempted by that section.
   (g) No inertia load factor used for design purposes may be less than 2.67,
 nor may the limit ground reaction load factor be less than 2.0 at design
 maximum weight, unless these lower values will not be exceeded in taxiing at
 speeds up to takeoff speed over terrain as rough as that expected in service.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13090, Aug. 13, 1969; Amdt. 23-28, 47 FR 13315, Mar. 29, 1982; Amdt. No.
 23-45, 58 FR 42160, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.477  Landing gear arrangement.

   Sections 23.479 through 23.483, or the conditions in Appendix C, apply to
 airplanes with conventional arrangements of main and nose gear, or main and
 tail gear.






 Sec. 23.479  Level landing conditions.

   (a) For a level landing, the airplane is assumed to be in the following
 attitudes:
   (1) For airplanes with tail wheels, a normal level flight attitude.
   (2) For airplanes with nose wheels, attitudes in which--
   (i) The nose and main wheels contact the ground simultaneously; and
   (ii) The main wheels contact the ground and the nose wheel is just clear of
 the ground.

 The attitude used in paragraph (a)(2)(i) of this section may be used in the
 analysis required under paragraph (a)(2)(ii) of this section.
   (b) When investigating landing conditions, the drag components simulating
 the forces required to accelerate the tires and wheels up to the landing
 speed (spin-up) must be properly combined with the corresponding
 instantaneous vertical ground reactions, and the forward-acting horizontal
 loads resulting from rapid reduction of the spin-up drag loads (spring-back)
 must be combined with vertical ground reactions at the instant of the peak
 forward load, assuming wing lift and a tire-sliding coefficient of friction
 of 0.8. However, the drag loads may not be less than 25 percent of the
 maximum vertical ground reactions (neglecting wing lift).
   (c) In the absence of specific tests or a more rational analysis for
 determining the wheel spin-up and spring-back loads for landing conditions,
 the method set forth in appendix D of this part must be used. If appendix D
 of this part is used, the drag components used for design must not be less
 than those given by appendix C of this part.
   (d) For airplanes with tip tanks or large overhung masses (such as turbo-
 propeller or jet engines) supported by the wing, the tip tanks and the
 structure supporting the tanks or overhung masses must be designed for the
 effects of dynamic responses under the level landing conditions of either
 paragraph (a)(1) or (a)(2)(ii) of this section. In evaluating the effects of
 dynamic response, an airplane lift equal to the weight of the airplane may be
 assumed.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-17, 41 FR
 55464, Dec. 20, 1976; Amdt. No. 23-45, 58 FR 42160, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.481  Tail down landing conditions.

   (a) For a tail down landing, the airplane is assumed to be in the following
 attitudes:
   (1) For airplanes with tail wheels, an attitude in which the main and tail
 wheels contact the ground simultaneously.
   (2) For airplanes with nose wheels, a stalling attitude, or the maximum
 angle allowing ground clearance by each part of the airplane, whichever is
 less.
   (b) For airplanes with either tail or nose wheels, ground reactions are
 assumed to be vertical, with the wheels up to speed before the maximum
 vertical load is attained.






 Sec. 23.483  One-wheel landing conditions.

   For the one-wheel landing condition, the airplane is assumed to be in the
 level attitude and to contact the ground on one side of the main landing
 gear. In this attitude, the ground reactions must be the same as those
 obtained on that side under Sec. 23.479.






 Sec. 23.485  Side load conditions.

   (a) For the side load condition, the airplane is assumed to be in a level
 attitude with only the main wheels contacting the ground and with the shock
 absorbers and tires in their static positions.
   (b) The limit vertical load factor must be 1.33, with the vertical ground
 reaction divided equally between the main wheels.
   (c) The limit side inertia factor must be 0.83, with the side ground
 reaction divided between the main wheels so that--
   (1) 0.5 (W) is acting inboard on one side; and
   (2) 0.33 (W) is acting outboard on the other side.
   (d) The side loads prescribed in paragraph (c) of this section are assumed
 to be applied at the ground contact point and the drag loads may be assumed
 to be zero.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. 23-45, 58 FR 42160, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.493  Braked roll conditions.

   Under braked roll conditions, with the shock absorbers and tires in their
 static positions, the following apply:
   (a) The limit vertical load factor must be 1.33.
   (b) The attitudes and ground contacts must be those described in Sec.
 23.479 for level landings.
   (c) A drag reaction equal to the vertical reaction at the wheel multiplied
 by a coefficient of friction of 0.8 must be applied at the ground contact
 point of each wheel with brakes, except that the drag reaction need not
 exceed the maximum value based on limiting brake torque.






 Sec. 23.497  Supplementary conditions for tail wheels.

   In determining the ground loads on the tail wheel and affected supporting
 structures, the following apply:
   (a) For the obstruction load, the limit ground reaction obtained in the
 tail down landing condition is assumed to act up and aft through the axle at
 45 degrees. The shock absorber and tire may be assumed to be in their static
 positions.
   (b) For the side load, a limit vertical ground reaction equal to the static
 load on the tail wheel, in combination with a side component of equal
 magnitude, is assumed. In addition--
   (1) If a swivel is used, the tail wheel is assumed to be swiveled 90
 degrees to the airplane longitudinal axis with the resultant ground load
 passing through the axle;
   (2) If a lock, steering device, or shimmy damper is used, the tail wheel is
 also assumed to be in the trailing position with the side load acting at the
 ground contact point; and
   (3) The shock absorber and tire are assumed to be in their static
 positions.






 Sec. 23.499  Supplementary conditions for nose wheels.

   In determining the ground loads on nose wheels and affected supporting
 structures, and assuming that the shock absorbers and tires are in their
 static positions, the following conditions must be met:
   (a) For aft loads, the limit force components at the axle must be--
   (1) A vertical component of 2.25 times the static load on the wheel; and
   (2) A drag component of 0.8 times the vertical load.
   (b) For forward loads, the limit force components at the axle must be--
   (1) A vertical component of 2.25 times the static load on the wheel; and
   (2) A forward component of 0.4 times the vertical load.
   (c) For side loads, the limit force components at ground contact must be--
   (1) A vertical component of 2.25 times the static load on the wheel; and
   (2) A side component of 0.7 times the vertical load.






 Sec. 23.505  Supplementary conditions for skiplanes.

   In determining ground loads for skiplanes, and assuming that the airplane
 is resting on the ground with one main ski frozen at rest and the other skis
 free to slide, a limit side force equal to 0.036 times the design maximum
 weight must be applied near the tail assembly, with a factor of safety of 1.

 [Amdt. 23-7, 34 FR 13090, Aug. 13, 1969]






 Sec. 23.507  Jacking loads.

   (a) The airplane must be designed for the loads developed when the aircraft
 is supported on jacks at the design maximum weight assuming the following
 load factors for landing gear jacking points at a three-point attitude and
 for primary flight structure jacking points in the level attitude:
   (1) Vertical-load factor of 1.35 times the static reactions.
   (2) Fore, aft, and lateral load factors of 0.4 times the vertical static
 reactions.
   (b) The horizontal loads at the jack points must be reacted by inertia
 forces so as to result in no change in the direction of the resultant loads
 at the jack points.
   (c) The horizontal loads must be considered in all combinations with the
 vertical load.

 [Amdt. 23-14, 38 FR 31821, Nov. 19, 1973]






 Sec. 23.509  Towing loads.

   The towing loads of this section must be applied to the design of tow
 fittings and their immediate attaching structure.
   (a) The towing loads specified in paragraph (d) of this section must be
 considered separately. These loads must be applied at the towing fittings and
 must act parallel to the ground. In addition:
   (1) A vertical load factor equal to 1.0 must be considered acting at the
 center of gravity; and
   (2) The shock struts and tires must be in there static positions.
   (b) For towing points not on the landing gear but near the plane of
 symmetry of the airplane, the drag and side tow load components specified for
 the auxiliary gear apply. For towing points located outboard of the main
 gear, the drag and side tow load components specified for the main gear
 apply. Where the specified angle of swivel cannot be reached, the maximum
 obtainable angle must be used.
   (c) The towing loads specified in paragraph (d) of this section must be
 reacted as follows:
   (1) The side component of the towing load at the main gear must be reacted
 by a side force at the static ground line of the wheel to which the load is
 applied.
   (2) The towing loads at the auxiliary gear and the drag components of the
 towing loads at the main gear must be reacted as follows:
   (i) A reaction with a maximum value equal to the vertical reaction must be
 applied at the axle of the wheel to which the load is applied. Enough
 airplane inertia to achieve equilibrium must be applied.
   (ii) The loads must be reacted by airplane inertia.
   (d) The prescribed towing loads are as follows, where W is the design
 maximum weight:

                                                      Load

    Tow point        Position      Magnitude        No.           Direction

  Main gear                           0.225W  1                Forward,
                                               2                parallel to
                                               3                drag axis.
                                               4                Forward, at 30
                                                                deg. to drag
                                                                axis.
                                                                Aft, parallel
                                                                to drag axis.
                                                                Aft, at 30
                                                                deg. to drag
                                                                axis.

  Auxiliary gear  Swiveled              0.3W  5                Forward.
                   forward                     6                Aft.
                  Swiveled aft          0.3W  7                Forward.
                                               8                Aft.
                  Swiveled 45          0.15W  9                Forward, in
                   deg. from                   10               plane of
                   forward                                      wheel.
                                                                Aft, in plane
                                                                of wheel.
                  Swiveled 45          0.15W  11               Forward, in
                   deg. from aft               12               plane of
                                                                wheel.
                                                                Aft, in plane
                                                                of wheel.

 [Amdt. 23-14, 38 FR 31821, Nov. 19, 1973]






 Sec. 23.511  Ground load; unsymmetrical loads on multiple-wheel units.

   (a) Pivoting loads. The airplane is assumed to pivot about on side of the
 main gear with--
   (1) The brakes on the pivoting unit locked; and
   (2) Loads corresponding to a limit vertical load factor of 1, and
 coefficient of friction of 0.8 applied to the main gear and its supporting
 structure.
   (b) Unequal tire loads. The loads established under Secs. 23.471 through
 23.483 must be applied in turn, in a 60/40 percent distribution, to the dual
 wheels and tires in each dual wheel landing gear unit.
   (c) Deflated tire loads. For the deflated tire condition--
   (1) 60 percent of the loads established under Secs. 23.471 through 23.483
 must be applied in turn to each wheel in a landing gear unit; and
   (2) 60 percent of the limit drag and side loads, and 100 percent of the
 limit vertical load established under Secs. 23.485 and 23.493 or lesser
 vertical load obtained under paragraph (c)(1) of this section, must be
 applied in turn to each wheel in the dual wheel landing gear unit.

 [Amdt. 23-7, 34 FR 13090, Aug. 13, 1969]






                                  Water Loads






 Sec. 23.521  Water load conditions.

   (a) The structure of seaplanes and amphibians must be designed for water
 loads developed during takeoff and landing with the seaplane in any attitude
 likely to occur in normal operation at appropriate forward and sinking
 velocities under the most severe sea conditions likely to be encountered.
   (b) Unless the applicant makes a rational analysis of the water loads,
 Secs. 23.523 through 23.537 apply.
   (c) Floats previously approved by the FAA may be installed on airplanes
 that are certificated under this part, provided that the floats meet the
 criteria of paragraph (a) of this section.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. No. 23-45, 58 FR 42160, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight airworthiness standards
 for normal, utility, acrobatic, and commuter category airplanes. The changes
 are based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, in St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.523  Design weights and center of gravity positions.

   (a) Design weights. The water load requirements must be met at each
 operating weight up to the design landing weight except that, for the takeoff
 condition prescribed in Sec. 23.531, the design water takeoff weight (the
 maximum weight for water taxi and takeoff run) must be used.
   (b) Center of gravity positions. The critical centers of gravity within the
 limits for which certification is requested must be considered to reach
 maximum design loads for each part of the seaplane structure.

 [Amdt. No. 23-45, 58 FR 42160, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.525   Application of loads.

   (a) Unless otherwise prescribed, the seaplane as a whole is assumed to be
 subjected to the loads corresponding to the load factors specified in Sec.
 23.527.
   (b) In applying the loads resulting from the load factors prescribed in
 Sec. 23.527, the loads may be distributed over the hull or main float bottom
 (in order to avoid excessive local shear loads and bending moments at the
 location of water load application) using pressures not less than those
 prescribed in Sec. 23.533(c).
   (c) For twin float seaplanes, each float must be treated as an equivalent
 hull on a fictitious seaplane with a weight equal to one-half the weight of
 the twin float seaplane.
   (d) Except in the takeoff condition of Sec. 23.531, the aerodynamic lift on
 the seaplane during the impact is assumed to be  2/3  of the weight of the
 seaplane.

 [Amdt. No. 23-45, 58 FR 42161, Aug. 6, 1993; 58 FR 51970, Oct. 5, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.527   Hull and main float load factors.

   (a) Water reaction load factors nw must be computed in the following
 manner:
   (1) For the step landing case

                             C1(VSO)**2
                  nw = -------------------------
                         (Tan**(2/3)b)W**(1/3)

   (2) For the bow and stern landing cases

                             C1(VSO)**2                    K1
                  nw = ------------------------- x ------------------
                         (Tan**(2/3)b)W**(1/3)     (1+(rx)**2)**(2/3)

   (b) The following values are used:
   (1) nw=water reaction load factor (that is, the water reaction divided by
 seaplane weight).
   (2) C1=empirical seaplane operations factor equal to 0.012 (except that
 this factor may not be less than that necessary to obtain the minimum value
 of step load factor of 2.33).
   (3) VSO=seaplane stalling speed in knots with flaps extended in the
 appropriate landing position and with no slipstream effect.
   (4) b=Angle of dead rise at the longitudinal station at which the load
 factor is being determined in accordance with figure 1 of appendix I of this
 part.
   (5) W=seaplane landing weight in pounds.
   (6) K1=empirical hull station weighing factor, in accordance with figure 2
 of appendix I of this part.
   (7) rx=ratio of distance, measured parallel to hull reference axis, from
 the center of gravity of the seaplane to the hull longitudinal station at
 which the load factor is being computed to the radius of gyration in pitch of
 the seaplane, the hull reference axis being a straight line, in the plane of
 symmetry, tangential to the keel at the main step.
   (c) For a twin float seaplane, because of the effect of flexibility of the
 attachment of the floats to the seaplane, the factor K1 may be reduced at the
 bow and stern to 0.8 of the value shown in figure 2 of appendix I of this
 part. This reduction applies only to the design of the carrythrough and
 seaplane structure.

 [Amdt. No. 23-45, 58 FR 42161, Aug. 6, 1993; 58 FR 51970, Oct. 5, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.529   Hull and main float landing conditions.

   (a) Symmetrical step, bow, and stern landing. For symmetrical step, bow,
 and stern landings, the limit water reaction load factors are those computed
 under Sec. 23.527. In addition--
   (1) For symmetrical step landings, the resultant water load must be applied
 at the keel, through the center of gravity, and must be directed
 perpendicularly to the keel line;
   (2) For symmetrical bow landings, the resultant water load must be applied
 at the keel, one-fifth of the longitudinal distance from the bow to the step,
 and must be directed perpendicularly to the keel line; and
   (3) For symmetrical stern landings, the resultant water load must be
 applied at the keel, at a point 85 percent of the longitudinal distance from
 the step to the stern post, and must be directed perpendicularly to the keel
 line.
   (b) Unsymmetrical landing for hull and single float seaplanes.
 Unsymmetrical step, bow, and stern landing conditions must be investigated.
 In addition--
   (1) The loading for each condition consists of an upward component and a
 side component equal, respectively, to 0.75 and 0.25 tan b times the
 resultant load in the corresponding symmetrical landing condition; and
   (2) The point of application and direction of the upward component of the
 load is the same as that in the symmetrical condition, and the point of
 application of the side component is at the same longitudinal station as the
 upward component but is directed inward perpendicularly to the plane of
 symmetry at a point midway between the keel and chine lines.
   (c) Unsymmetrical landing; twin float seaplanes. The unsymmetrical loading
 consists of an upward load at the step of each float of 0.75 and a side load
 of 0.25 tan b at one float times the step landing load reached under Sec.
 23.527. The side load is directed inboard, perpendicularly to the plane of
 symmetry midway between the keel and chine lines of the float, at the same
 longitudinal station as the upward load.

 [Amdt. No. 23-45, 58 FR 42161, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.531   Hull and main float takeoff condition.

   For the wing and its attachment to the hull or main float--
   (a) The aerodynamic wing lift is assumed to be zero; and
   (b) A downward inertia load, corresponding to a load factor computed from
 the following formula, must be applied:

                                   CTO(VSI)**2
                           n = ---------------------
                               (Tan**(2/3)b)W**(1/3)

 Where--
 n=inertia load factor;
 CTO=empirical seaplane operations factor equal to 0.004;
 VS1=seaplane stalling speed (knots) at the design takeoff weight with the
   flaps extended in the appropriate takeoff position;
 b=angle of dead rise at the main step (degrees); and
 W=design water takeoff weight in pounds.

 [Amdt. No. 23-45, 58 FR 42161, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.533  Hull and main float bottom pressures.

   (a) General. The hull and main float structure, including frames and
 bulkheads, stringers, and bottom plating, must be designed under this
 section.
   (b) Local pressures. For the design of the bottom plating and stringers and
 their attachments to the supporting structure, the following pressure
 distributions must be applied:
   (1) For an unflared bottom, the pressure at the chine is 0.75 times the
 pressure at the keel, and the pressures between the keel and chine vary
 linearly, in accordance with figure 3 of appendix I of this part. The
 pressure at the keel (p.s.i.) is computed as follows:

                               C2K2(VSI)**2
                         PK = ----------------
                                 Tan(bk)

 where--
 Pk=pressure (p.s.i.) at the keel;
 C2=0.00213;
 K2=hull station weighing factor, in accordance with figure 2 of appendix I of
   this part;
 VS1=seaplane stalling speed (knots) at the design water takeoff weight with
   flaps extended in the appropriate takeoff position; and
 bK=angle of dead rise at keel, in accordance with figure 1 of appendix I of
   this part.

   (2) For a flared bottom, the pressure at the beginning of the flare is the
 same as that for an unflared bottom, and the pressure between the chine and
 the beginning of the flare varies linearly, in accordance with figure 3 of
 appendix I of this part. The pressure distribution is the same as that
 prescribed in paragraph (b)(1) of this section for an unflared bottom except
 that the pressure at the chine is computed as follows:

                               C3K2(VSI)**2
                        Pch = ----------------
                                 Tan(b)

 where--
 Pch=pressure (p.s.i.) at the chine;
 C3=0.0016;
 K2=hull station weighing factor, in accordance with figure 2 of appendix I of
   this part;
 VS1=seaplane stalling speed (knots) at the design water takeoff weight with
   flaps extended in the appropriate takeoff position; and
 b=angle of dead rise at appropriate station.

   The area over which these pressures are applied must simulate pressures
 occurring during high localized impacts on the hull or float, but need not
 extend over an area that would induce critical stresses in the frames or in
 the overall structure.
   (c) Distributed pressures. For the design of the frames, keel, and chine
 structure, the following pressure distributions apply:
   (1) Symmetrical pressures are computed as follows:

                             C4K2(VSO)**2
                        P = ----------------
                               Tan(b)

 where--
 P=pressure (p.s.i.);
 C4=0.078 C1 (with C1 computed under Sec. 23.527);
 K2=hull station weighing factor, determined in accordance with figure 2 of
   appendix I of this part;
 VS0=seaplane stalling speed (knots) with landing flaps extended in the
   appropriate position and with no slipstream effect; and
 b=angle of dead rise at appropriate station.

   (2) The unsymmetrical pressure distribution consists of the pressures
 prescribed in paragraph (c)(1) of this section on one side of the hull or
 main float centerline and one-half of that pressure on the other side of the
 hull or main float centerline, in accordance with figure 3 of appendix I of
 this part.
   (3) These pressures are uniform and must be applied simultaneously over the
 entire hull or main float bottom. The loads obtained must be carried into the
 sidewall structure of the hull proper, but need not be transmitted in a fore
 and aft direction as shear and bending loads.

 [Amdt. No. 23-45, 58 FR 42161, Aug. 6, 1993; 58 FR 51970, Oct. 5, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.535  Auxiliary float loads.

   (a) General. Auxiliary floats and their attachments and supporting
 structures must be designed for the conditions prescribed in this section. In
 the cases specified in paragraphs (b) through (e) of this section, the
 prescribed water loads may be distributed over the float bottom to avoid
 excessive local loads, using bottom pressures not less than those prescribed
 in paragraph (g) of this section.
   (b) Step loading. The resultant water load must be applied in the plane of
 symmetry of the float at a point three-fourths of the distance from the bow
 to the step and must be perpendicular to the keel. The resultant limit load
 is computed as follows, except that the value of L need not exceed three
 times the weight of the displaced water when the float is completely
 submerged:

                                     C5(VSO)**2(W**(2/3))
                           L = ------------------------------
                               Tan**(2/3)bs(1+(ry)**2)**(2/3)
 where--
 L=limit load (lbs.);
 C5=0.0053;
 VS0=seaplane stalling speed (knots) with landing flaps extended in the
   appropriate position and with no slipstream effect;
 W=seaplane design landing weight in pounds;
 bs=angle of dead rise at a station  3/4  of the distance from the bow to the
   step, but need not be less than 15 degrees; and
 ry=ratio of the lateral distance between the center of gravity and the plane
   of symmetry of the float to the radius of gyration in roll.

   (c) Bow loading. The resultant limit load must be applied in the plane of
 symmetry of the float at a point one-fourth of the distance from the bow to
 the step and must be perpendicular to the tangent to the keel line at that
 point. The magnitude of the resultant load is that specified in paragraph (b)
 of this section.
   (d) Unsymmetrical step loading. The resultant water load consists of a
 component equal to 0.75 times the load specified in paragraph (a) of this
 section and a side component equal to 0.025 tan b times the load specified in
 paragraph (b) of this section. The side load must be applied perpendicularly
 to the plane of symmetry of the float at a point midway between the keel and
 the chine.
   (e) Unsymmetrical bow loading. The resultant water load consists of a
 component equal to 0.75 times the load specified in paragraph (b) of this
 section and a side component equal to 0.25 tan b times the load specified in
 paragraph (c) of this section. The side load must be applied perpendicularly
 to the plane of symmetry at a point midway between the keel and the chine.
   (f) Immersed float condition. The resultant load must be applied at the
 centroid of the cross section of the float at a point one-third of the
 distance from the bow to the step. The limit load components are as follows:

          vertical = PgV

                     CXP(V**(2/3))(K VSO)**2
               aft = -----------------------
                                2

                     CYP(V**(2/3))(K VSO)**2
              side = -----------------------
                                2

 where--
 P=mass density of water (slugs/ft. 3 )
 V=volume of float (ft. 3 );
 CX=coefficient of drag force, equal to 0.133;
 CY=coefficient of side force, equal to 0.106;
 K=0.8, except that lower values may be used if it is shown that the floats
   are incapable of submerging at a speed of 0.8 Vso in normal operations;
 Vso=seaplane stalling speed (knots) with landing flaps extended in the
   appropriate position and with no slipstream effect; and
 g=acceleration due to gravity (ft/sec**2).
   (g) Float bottom pressures. The float bottom pressures must be established
 under Sec. 23.533, except that the value of K2 in the formulae may be taken
 as 1.0. The angle of dead rise to be used in determining the float bottom
 pressures is set forth in paragraph (b) of this section.

 [Amdt. No. 23-45, 58 FR 42162, Aug. 6, 1993; 58 FR 51970, Oct. 5, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.537  Seawing loads.

   Seawing design loads must be based on applicable test data.

 [Amdt. No. 23-45, 58 FR 42163, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






                         Emergency Landing Conditions






 Sec. 23.561  General.

   (a) The airplane, although it may be damaged in emergency landing
 conditions, must be designed as prescribed in this section to protect each
 occupant under those conditions.
   (b) The structure must be designed to protect each occupant during
 emergency landing conditions when--
   (1) Proper use is made of the seats, safety belts, and shoulder harnesses
 provided for in the design;
   (2) The occupant experiences the static inertia loads corresponding to the
 following ultimate load factors--
   (i) Upward, 3.0g for normal, utility, and commuter category airplanes, or
 4.5g for acrobatic category airplanes;
   (ii) Forward, 9.0g;
   (iii) Sideward, 1.5g; and
   (3) The items of mass within the cabin, that could injure an occupant,
 experience the static inertia loads corresponding to the following ultimate
 load factors--
   (i) Upward, 3.0g;
   (ii) Forward, 18.0g; and
   (iii) Sideward, 4.5g.
   (c) Each airplane with retractable landing gear must be designed to protect
 each occupant in a landing--
   (1) With the wheels retracted;
   (2) With moderate descent velocity; and
   (3) Assuming, in the absence of a more rational analysis--
   (i) A downward ultimate inertia force of 3 g; and
   (ii) A coefficient of friction of 0.5 at the ground.
   (d) If it is not established that a turnover is unlikely during an
 emergency landing, the structure must be designed to protect the occupants in
 a complete turnover as follows:
   (1) The likelihood of a turnover may be shown by an analysis assuming the
 following conditions--
   (i) Maximum weight;
   (ii) Most forward center of gravity position;
   (iii) Longitudinal load factor of 9.0g;
   (iv) Vertical load factor of 1.0g; and
   (v) For airplanes with tricycle landing gear, the nose wheel strut failed
 with the nose contacting the ground.
   (2) For determining the loads to be applied to the inverted airplane after
 a turnover, an upward ultimate inertia load factor of 3.0g and a coefficient
 of friction with the ground of 0.5 must be used.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13090, Aug. 13, 1969; Amdt. 23-24, 52 FR 34745, Sept. 14, 1987; Amdt. 23-36,
 53 FR 30812, Aug. 15, 1988]






 Sec. 23.562  Emergency landing dynamic conditions.

   (a) Each seat/restraint system for use in a normal, utility, or acrobatic
 category airplane must be designed to protect each occupant during an
 emergency landing when--
   (1) Proper use is made of seats, safety belts, and shoulder harnesses
 provided for in the design; and
   (2) The occupant is exposed to the loads resulting from the conditions
 prescribed in this section.
   (b) Except for those seat/restraint systems that are required to meet
 paragraph (d) of this section, each seat/restraint system for crew or
 passenger occupancy in a normal, utility, or acrobatic category airplane,
 must successfully complete dynamic tests or be demonstrated by rational
 analysis supported by dynamic tests, in accordance with each of the following
 conditions. These tests must be conducted with an occupant simulated by an
 anthropomorphic test dummy (ATD) defined by 49 CFR Part 572, Subpart B, or an
 FAA-approved equivalent, with a nominal weight of 170 pounds and seated in
 the normal upright position.
   (1) For the first test, the change in velocity may not be less than 31 feet
 per second. The seat/restraint system must be oriented in its nominal
 position with respect to the airplane and with the horizontal plane of the
 airplane pitched up 60 degrees, with no yaw, relative to the impact vector.
 For seat/restraint systems to be installed in the first row of the airplane,
 peak deceleration must occur in not more than 0.05 seconds after impact and
 must reach a minimum of 19g. For all other seat/restraint systems, peak
 deceleration must occur in not more than 0.06 seconds after impact and must
 reach a minimum of 15g.
   (2) For the second test, the change in velocity may not be less than 42
 feet per second. The seat/restraint system must be oriented in its nominal
 position with respect to the airplane and with the vertical plane of the
 airplane yawed 10 degrees, with no pitch, relative to the impact vector in a
 direction that results in the greatest load on the shoulder harness. For
 seat/restraint systems to be installed in the first row of the airplane, peak
 deceleration must occur in not more than 0.05 seconds after impact and must
 reach a minimum of 26g. For all other seat/restraint systems, peak
 deceleration must occur in not more than 0.06 seconds after impact and must
 reach a minimum of 21g.
   (3) To account for floor warpage, the floor rails or attachment devices
 used to attach the seat/restraint system to the airframe structure must be
 preloaded to misalign with respect to each other by at least 10 degrees
 vertically (i.e., pitch out of parallel) and one of the rails or attachment
 devices must be preloaded to misalign by 10 degrees in roll prior to
 conducting the test defined by paragraph (b)(2) of this section.
   (c) Compliance with the following requirements must be shown during the
 dynamic tests conducted in accordance with paragraph (b) of this section:
   (1) The seat/restraint system must restrain the ATD although seat/restraint
 system components may experience deformation, elongation, displacement, or
 crushing intended as part of the design.
   (2) The attachment between the seat/restraint system and the test fixture
 must remain intact, although the seat structure may have deformed.
   (3) Each shoulder harness strap must remain on the ATD's shoulder during
 the impact.
   (4) The safety belt must remain on the ATD's pelvis during the impact.
   (5) The results of the dynamic tests must show that the occupant is
 protected from serious head injury.
   (i) When contact with adjacent seats, structure, or other items in the
 cabin can occur, protection must be provided so that the head impact does not
 exceed a head injury criteria (HIC) of 1,000.
   (ii) The value of HIC is defined as--

                       1          t2          2.5
   HIC= { (t2-t1) [----------     S   a(t)dt ]      }
                    (t2-t1)       t1                  Max

 Where: t1 is the initial integration time, expressed in seconds, t2 is the
     final integration time, expressed in seconds, (t2-t1) is the time
     duration of the major head impact, expressed in seconds, and a(t) is the
     resultant deceleration at the center of gravity of the head form
     expressed as a multiple of g (units of gravity).

   (iii) Compliance with the HIC limit must be demonstrated by measuring the
 head impact during dynamic testing as prescribed in paragraphs (b)(1) and
 (b)(2) of this section or by a separate showing of compliance with the head
 injury criteria using test or analysis procedures.
   (6) Loads in individual shoulder harness straps may not exceed 1,750
 pounds. If dual straps are used for retaining the upper torso, the total
 strap loads may not exceed 2,000 pounds.
   (7) The compression load measured between the pelvis and the lumbar spine
 of the ATD may not exceed 1,500 pounds.
   (d) For all single-engine airplanes with a VS0 of more than 61 knots at
 maximum weight, and those multiengine airplanes of 6,000 pounds or less
 maximum weight with a VS0 of more than 61 knots at maximum weight that do not
 comply with Sec. 23.67(b)(2)(i):
   (1) The ultimate load factors of Sec. 23.561(b) must be increased by
 multiplying the load factors by the square of the ratio of the increased
 stall speed to 61 knots. The increased ultimate load factors need not exceed
 the values reached at a VS0 of 79 knots. The upward ultimate load factor for
 acrobatic category airplanes need not exceed 5.0g.
   (2) The seat/restraint system test required by paragraph (b)(1) of this
 section must be conducted in accordance with the following criteria:
   (i) The change in velocity may not be less than 31 feet per second.
   (ii)(A) The peak deceleration (gp) of 19g and 15g must be increased and
 multiplied by the square of the ratio of the increased stall speed to 61
 knots:

 gp=19.0 (VS0/61)**2 or gp=15.0 (VS0/61)**2

   (B) The peak deceleration need not exceed the value reached at a VS0 of 79
 knots.
   (iii) The peak deceleration must occur in not more than time (tr), which
 must be computed as follows:

                                      31      .96
                              tr = -------- = ---
                                   32.2(gp)   gp

 where--

 gp=The peak deceleration calculated in accordance with paragraph (d)(2)(ii)
     of this section
 tr=The rise time (in seconds) to the peak deceleration.

   (e) An alternate approach that achieves an equivalent, or greater, level of
 occupant protection to that required by this section may be used if
 substantiated on a rational basis.

 [Amdt. 23-36, 53 FR 30812, Aug. 15, 1988, as amended by Amdt. 23-44, 58 FR
 38639, July 19, 1993]

 *****************************************************************************


 58 FR 38634, No. 136, July 19, 1993

 SUMMARY: This final rule amends the stalling speed requirements applicable to
 single-engine airplanes and to certain multiengine small airplanes of less
 than 6,000 pounds maximum weight. The rule permits those airplanes to have a
 stall speed greater than 61 knots, provided they meet certain additional
 occupant protection standards. These changes are needed to permit the design
 and type certification of higher performance airplanes with increased cruise
 speeds and better specific fuel consumption. The amendments are intended to
 achieve the benefits of certificating higher performance airplanes while
 affording their occupants the same level of protection in an emergency
 landing that is presently provided by airplanes with a 61-knot stall speed.

 EFFECTIVE DATE: August 18, 1993.

 *****************************************************************************






                              Fatigue Evaluation






 Sec. 23.571  Pressurized cabin.

   The strength, detail design, and fabrication of the pressure cabin
 structure must be evaluated under one of the following:
   (a) A fatigue strength investigation, in which the structure is shown by
 analysis, tests, or both to be able to withstand the repeated loads of
 variable magnitude expected in service. Analysis alone is considered
 acceptable only when it is conservative and applied to simple structures.
   (b) A fail safe strength investigation, in which it is shown by analysis,
 tests, or both that catastrophic failure of the structure is not probable
 after fatigue failure, or obvious partial failure, of a principal structural
 element, and that the remaining structures are able to withstand a static
 ultimate load factor of 75 percent of the limit load factor at VC,
 considering the combined effects of normal operating pressures, expected
 external aerodynamic pressures, and flight loads. These loads must be
 multiplied by a factor of 1.15 unless the dynamic effects of failure under
 static load are otherwise considered.
   (c) The damage tolerance evaluation of Sec. 23.573(b).

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-14, 38 FR
 31821, Nov. 19, 1973; Amdt. No. 23-45, 58 FR 42163, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.572   Wing, empennage, and associated structures.

   (a) The strength, detail design, and fabrication of those parts of the
 wings (including canards, tandem wings, and winglets/tip fins), empennage,
 their carry-through and attaching structures, whose failure would be
 catastrophic, must be evaluated under one of the following unless it is
 shown that the structure, operating stress level, materials, and expected
 uses are comparable, from a fatigue standpoint, to a similar design that has
 had extensive satisfactory service experience:
   (1) A fatigue strength investigation in which the structure is shown by
 analysis, tests, or both to be able to withstand the repeated loads of
 variable magnitude expected in service. Analysis alone is acceptable only
 when it is conservative and applied to simple structures; or
   (2) A fail-safe strength investigation in which it is shown by analysis,
 tests, or both, that catastrophic failure of the structure is not probably
 after fatigue failure, or obvious partial failure, of a principal structural
 element, and that the remaining structure is able to withstand a static
 ultimate load factor of 75 percent of the critical limit load factor at Vc.
 These loads must be multiplied by a factor of 1.15 unless the dynamic effects
 of failure under static load are otherwise considered.
   (3) The damage tolerance evaluation of Sec. 23.573(b).
   (b) Each evaluation required by this section must--
   (1) Include typical loading spectra (e.g. taxi, ground-air-ground cycles,
 maneuver, gust);
   (2) Account for any significant effects due to the mutual influence of
 aerodynamic surfaces; and
   (3) Consider any significant effects from propeller slipstream loading, and
 buffet from vortex impingements.

 [Amdt. 23-7, 34 FR 13090, Aug. 13, 1969, as amended by Amdt. 23-14, 38 FR
 31821, Nov. 19, 1973; Amdt. 23-34, 52 FR 1830, Jan. 15, 1987; Amdt. 23-38, 54
 FR 39511, Sept. 26, 1989; Amdt. No. 23-45, 58 FR 42163, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.573   Damage tolerance and fatigue evaluation of structure.

   (a) Composite airframe structure. Composite airframe structure must be
 evaluated under this paragraph instead of Secs. 23.571 and 23.572. The
 applicant must evaluate the composite airframe structure, the failure of
 which would result in catastrophic loss of the airplane, in each wing
 (including canards, tandem wings, and winglets), empennage, their
 carrythrough and attaching structure, moveable control surfaces and their
 attaching structure fuselage, and pressure cabin using the damage-tolerance
 criteria prescribed in paragraphs (a)(1) through (a)(4) of this section
 unless shown to be impractical. If the applicant establishes that damage-
 tolerance criteria is impractical for a particular structure, the structure
 must be evaluated in accordance with paragraphs (a)(1) and (a)(6) of this
 section. Where bonded joints are used, the structure must also be evaluated
 in accordance with paragraph (a)(5) of this section. The effects of material
 variability and environmental conditions on the strength and durability
 properties of the composite materials must be accounted for in the
 evaluations required by this section.
   (1) It must be demonstrated by tests, or by analysis supported by tests,
 that the structure is capable of carrying ultimate load with damage up to the
 threshold of detectability considering the inspection procedures employed.
   (2) The growth rate or no-growth of damage that may occur from fatigue,
 corrosion, manufacturing flaws or impact damage, under repeated loads
 expected in service, must be established by tests or analysis supported by
 tests.
   (3) The structure must be shown by residual strength tests, or analysis
 supported by residual strength tests, to be able to withstand critical limit
 flight loads, considered as ultimate loads, with the extent of detectable
 damage consistent with the results of the damage tolerance evaluations. For
 pressurized cabins, the following loads must be withstood:
   (i) Critical limit flight loads with the combined effects of normal
 operating pressure and expected external aerodynamic pressures.
   (ii) The expected external aerodynamic pressures in 1g flight combined with
 a cabin differential pressure equal to 1.1 times the normal operating
 differential pressure without any other load.
   (4) The damage growth, between initial detectability and the value selected
 for residual strength demonstrations, factored to obtain inspection
 intervals, must allow development of an inspection program suitable for
 application by operation and maintenance personnel.
   (5) The limit load capacity of each bonded joint must be substantiated by
 one of the following methods:
   (i) The maximum disbonds of each bonded joint consistent with the
 capability to withstand the loads in paragraph (a)(3) of this section must be
 determined by analysis, tests, or both. Disbonds of each bonded joint greater
 than this must be prevented by design features; or
   (ii) Proof testing must be conducted on each production article that will
 apply the critical limit design load to each critical bonded joint; or
   (iii) Repeatable and reliable non-destructive inspection techniques must be
 established that ensure the strength of each joint.
   (6) Structural components for which the damage tolerance method is shown to
 be impractical must be shown by component fatigue tests, or analysis
 supported by tests, to be able to withstand the repeated loads of variable
 magnitude expected in service. Sufficient component, subcomponent, element,
 or coupon tests must be done to establish the fatigue scatter factor and the
 environmental effects. Damage up to the threshold of detectability and
 ultimate load residual strength capability must be considered in the
 demonstration.
   (b) Metallic airframe structure. If the applicant elects to use Sec.
 23.571(c) or Sec. 23.572(a)(3), then the damage tolerance evaluation must
 include a determination of the probable locations and modes of damage due to
 fatigue, corrosion, or accidental damage. The determination must be by
 analysis supported by test evidence and, if available, service experience.
 Damage at multiple sites due to fatigue must be included where the design is
 such that this type of damage can be expected to occur. The evaluation must
 incorporate repeated load and static analyses supported by test evidence. The
 extent of damage for residual strength evaluation at any time within the
 operational life of the airplane must be consistent with the initial
 detectability and subsequent growth under repeated loads. The residual
 strength evaluation must show that the remaining structure is able to
 withstand critical limit flight loads, considered as ultimate, with the
 extent of detectable damage consistent with the results of the damage
 tolerance evaluations. For pressurized cabins, the following load must be
 withstood:
   (1) The normal operating differential pressure combined with the expected
 external aerodynamic pressures applied simultaneously with the flight loading
 conditions specified in this part, and
   (2) The expected external aerodynamic pressures in 1g flight combined with
 a cabin differential pressure equal to 1.1 times the normal operating
 differential pressure without any other load.
   (c) Inspection. Based on evaluations required by this section, inspections
 or other procedures must be established as necessary to prevent catastrophic
 failure and must be included in the Airworthiness Limitations section of the
 Instructions for Continued Airworthiness required by Sec. 23.1529.

 [Amdt. No. 23-45, 58 FR 42163, Aug. 6, 1993; 58 FR 51970, Oct. 5, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************



                      Subpart D--Design and Construction






 Sec. 23.601  General.

   The suitability of each questionable design detail and part having an
 important bearing on safety in operations, must be established by tests.






 Sec. 23.603  Materials and workmanship.

   (a) The suitability and durability of materials used for parts, the failure
 of which could adversely affect safety, must--
   (1) Be established by experience or tests;
   (2) Meet approved specifications that ensure their having the strength and
 other properties assumed in the design data; and
   (3) Take into account the effects of environmental conditions, such as
 temperature and humidity, expected in service.
   (b) Workmanship must be of a high standard.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-17, 41 FR
 55464, Dec. 20, 1976; Amdt. 23-23, 43 FR 50592, Oct. 10, 1978]






 Sec. 23.605  Fabrication methods.

   (a) The methods of fabrication used must produce consistently sound
 structures. If a fabrication process (such as gluing, spot welding, or heat-
 treating) requires close control to reach this objective, the process must be
 performed under an approved process specification.
   (b) Each new aircraft fabrication method must be substantiated by a test
 program.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. 23-23, 43 FR 50592, Oct. 10, 1978]






 Sec. 23.607  Self-locking nuts.

   No self-locking nut may be used on any bolt subject to rotation in
 operation unless a nonfriction locking device is used in addition to the
 self-locking device.

 [Amdt. 23-17, 41 FR 55464, Dec. 20, 1976]






 Sec. 23.609  Protection of structure.

   Each part of the structure must--
   (a) Be suitably protected against deterioration or loss of strength in
 service due to any cause, including--
   (1) Weathering;
   (2) Corrosion; and
   (3) Abrasion; and
   (b) Have adequate provisions for ventilation and drainage.






 Sec. 23.611  Accessibility.

   Means must be provided to allow inspection (including inspection of
 principal structural elements and control systems), close examination,
 repair, and replacement of each part requiring maintenance, adjustments for
 proper alignment and function, lubrication or servicing.

 [Amdt. 23-7, 34 FR 13090, Aug. 13, 1969]






 Sec. 23.613  Material strength properties and design values.

   (a) Material strength properties must be based on enough tests of material
 meeting specifications to establish design values on a statistical basis.
   (b) Design values must be chosen to minimize the probability of structural
 failure due to material variability. Except as provided in paragraph (e) of
 this section, compliance with this paragraph must be shown by selecting
 design values that ensure material strength with the following probability:
   (1) Where applied loads are eventually distributed through a single member
 within an assembly, the failure of which would result in loss of structural
 integrity of the component; 99 percent probability with 95 percent
 confidence.
   (2) For redundant structure, in which the failure of individual elements
 would result in applied loads being safely distributed to other load carrying
 members; 90 percent probability with 95 percent confidence.
   (c) The effects of temperature on allowable stresses used for design in an
 essential component or structure must be considered where thermal effects are
 significant under normal operating conditions.
   (d) The design of the structure must minimize the probability of
 catastrophic fatigue failure, particularly at points of stress concentration.
   (e) Design values greater than the guaranteed minimums required by this
 section may be used where only guaranteed minimum values are normally allowed
 if a "premium selection" of the material is made in which a specimen of each
 individual item is tested before use to determine that the actual strength
 properties of that particular item will equal or exceed those used in design.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. 23-23, 43 FR 50592, Oct. 30, 1978; Amdt. No. 23-45, 58 FR
 42163, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.615  [Removed. Amdt. No. 23-45, 58 FR 42164, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.619  Special factors.

   The factor of safety prescribed in Sec. 23.303 must be multiplied by the
 highest pertinent special factors of safety prescribed in Secs. 23.621
 through 23.625 for each part of the structure whose strength is--
   (a) Uncertain;
   (b) Likely to deteriorate in service before normal replacement; or
   (c) Subject to appreciable variability because of uncertainties in
 manufacturing processes or inspection methods.

 [Amdt. 23-7, 34 FR 13091, Aug. 13, 1969]






 Sec. 23.621  Casting factors.

   (a) General. The factors, tests, and inspections specified in paragraphs
 (b) through (d) of this section must be applied in addition to those
 necessary to establish foundry quality control. The inspections must meet
 approved specifications. Paragraphs (c) and (d) of this section apply to any
 structural castings except castings that are pressure tested as parts of
 hydraulic or other fluid systems and do not support structural loads.
   (b) Bearing stresses and surfaces.  The casting factors specified in
 paragraphs (c) and (d) of this section--
   (1) Need not exceed 1.25 with respect to bearing stresses regardless of the
 method of inspection used; and
   (2) Need not be used with respect to the bearing surfaces of a part whose
 bearing factor is larger than the applicable casting factor.
   (c) Critical castings. For each casting whose failure would preclude
 continued safe flight and landing of the airplane or result in serious injury
 to occupants, the following apply:
   (1) Each critical casting must either--
   (i) Have a casting factor of not less than 1.25 and receive 100 percent
 inspection by visual, radiographic, and either magnetic particle, penetrant
 or other approved equivalent non-destructive inspection method; or
   (ii) Have a casting factor of not less than 2.0 and receive 100 percent
 visual inspection and 100 percent approved non-destructive inspection. When
 an approved quality control procedure is established and an acceptable
 statistical analysis supports reduction, non-destructive inspection may be
 reduced from 100 percent, and applied on a sampling basis.
   (2) For each critical casting with a casting factor less than 1.50, three
 sample castings must be static tested and shown to meet--
   (i) The strength requirements of Sec. 23.305 at an ultimate load
 corresponding to a casting factor of 1.25; and
   (ii) The deformation requirements of Sec. 23.305 at a load of 1.15 times
 the limit load.
   (3) Examples of these castings are structural attachment fittings, parts of
 flight control systems, control surface hinges and balance weight
 attachments, seat, berth, safety belt, and fuel and oil tank supports and
 attachments, and cabin pressure valves.
   (d) Non-critical castings. For each casting other than those specified in
 paragraph (c) or (e) of this section, the following apply:
   (1) Except as provided in paragraphs (d) (2) and (3) of this section, the
 casting factors and corresponding inspections must meet the following table:

         Casting factor                            Inspection

 2.0 or more                      100 percent visual.
 Less than 2.0 but more than 1.5  100 percent visual, and magnetic particle or
                                   penetrant or equivalent nondestructive
                                   inspection methods.
 1.25 through 1.50                100 percent visual, magnetic particle or
                                   penetrant, and radiographic, or approved
                                   equivalent nondestructive inspection
                                   methods.

   (2) The percentage of castings inspected by nonvisual methods may be
 reduced below that specified in subparagraph (d)(1) of this section when an
 approved quality control procedure is established.
   (3) For castings procured to a specification that guarantees the mechanical
 properties of the material in the casting and provides for demonstration of
 these properties by test of coupons cut from the castings on a sampling
 basis--
   (i) A casting factor of 1.0 may be used; and
   (ii) The castings must be inspected as provided in paragraph (d)(1) of this
 section for casting factors of "1.25 through 1.50" and tested under paragraph
 (c)(2) of this section.
   (e) Non-structural castings. Castings used for non-structural purposes do
 not require evaluation, testing or close inspection.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965; Amdt.
 No. 23-45, 58 FR 42164, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.623  Bearing factors.

   (a) Each part that has clearance (free fit), and that is subject to
 pounding or vibration, must have a bearing factor large enough to provide for
 the effects of normal relative motion.
   (b) For control surface hinges and control system joints, compliance with
 the factors prescribed in Secs. 23.657 and 23.693, respectively, meets
 paragraph (a) of this section.

 [Amdt. 23-7, 34 FR 13091, Aug. 13, 1969]






 Sec. 23.625  Fitting factors.

   For each fitting (a part or terminal used to join one structural member to
 another), the following apply:
   (a) For each fitting whose strength is not proven by limit and ultimate
 load tests in which actual stress conditions are simulated in the fitting and
 surrounding structures, a fitting factor of at least 1.15 must be applied to
 each part of--
   (1) The fitting;
   (2) The means of attachment; and
   (3) The bearing on the joined members.
   (b) No fitting factor need be used for joint designs based on comprehensive
 test data (such as continuous joints in metal plating, welded joints, and
 scarf joints in wood).
   (c) For each integral fitting, the part must be treated as a fitting up to
 the point at which the section properties become typical of the member.
   (d) For each seat, berth, safety belt, and harness, its attachment to the
 structure must be shown, by analysis, tests, or both, to be able to withstand
 the inertia forces prescribed in Sec. 23.561 multiplied by a fitting factor
 of 1.33.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13091, Aug. 13, 1969]






 Sec. 23.627  Fatigue strength.

   The structure must be designed, as far as practicable, to avoid points of
 stress concentration where variable stresses above the fatigue limit are
 likely to occur in normal service.






 Sec. 23.629  Flutter.

   (a) It must be shown by one of the methods specified in paragraph (b), (c),
 or (d) of this section, or a combination of these methods, that the airplane
 is free from flutter, control reversal, and divergence for any condition of
 operation within the limit V-n envelope, and at all speeds up to the speed
 specified for the selected method. In addition--
   (1) Adequate tolerances must be established for quantities which affect
 flutter, including speed, damping, mass balance, and control system
 stiffness; and
   (2) The natural frequencies of main structural components must be
 determined by vibration tests or other approved methods.
   (b) A rational analysis may be used to show that the airplane is free from
 flutter, control reversal, and divergence if the analysis shows freedom from
 flutter for all speeds up to 1.2VD.
   (c) Flight flutter tests may be used to show that the airplane is free from
 flutter, control reversal, and divergence if it is shown by these tests
 that--
   (1) Proper and adequate attempts to induce flutter have been made within
 the speed range up to VD;
   (2) The vibratory response of the structure during the test indicates
 freedom from flutter;
   (3) A proper margin of damping exists at VD; and
   (4) There is no large and rapid reduction in damping as VD is approached.
   (d) Compliance with the rigidity and mass balance criteria (pages 4-12), in
 Airframe and Equipment Engineering Report No. 45 (as corrected) "Simplified
 Flutter Prevention Criteria" (published by the Federal Aviation
 Administration) may be accomplished to show that the airplane is free from
 flutter, control reversal, or divergence if--
   (1) VD/MD for the airplane is less than 260 knots (EAS) and less than Mach
 0.5,
   (2) The wing and aileron flutter prevention criteria, as represented by the
 wing torsional stiffness and aileron balance criteria, are limited in use to
 airplanes without large mass concentrations (such as engines, floats, or fuel
 tanks in outer wing panels) along the wing span, and
   (3) The airplane--
   (i) Does not have a T-tail or boom tail,
   (ii) Does not have unusual mass distributions or other unconventional
 design features that affect the applicability of the criteria, and
   (iii) Has fixed-fin and fixed-stabilizer surfaces.
   (e) For turbopropeller-powered airplanes, the dynamic evaluation must
 include--
   (1) Whirl mode degree of freedom which takes into account the stability of
 the plane of rotation of the propeller and significant elastic, inertial, and
 aerodynamic forces, and
   (2) Propeller, engine, engine mount, and airplane structure stiffness and
 damping variations appropriate to the particular configuration.
   (f) Freedom from flutter, control reversal, and divergence up to VD/MD must
 be shown as follows:
   (1) For airplanes that meet the criteria of paragraphs (d)(1) through
 (d)(3) of this section, after the failure, malfunction, or disconnection of
 any single element in any tab control system.
   (2) For airplanes other than those described in paragraph (f)(1) of this
 section, after the failure, malfunction, or disconnection of any single
 element in the primary flight control system, any tab control system, or any
 flutter damper.
   (g) For airplanes showing compliance with the fail-safe criteria of Secs.
 23.571 and 23.572, the airplane must be shown by analysis or test to be free
 from flutter to VD/MD after fatigue failure, or obvious partial failure of a
 principle structural element.
   (h) For airplanes showing compliance with the damage-tolerance criteria of
 Sec. 23.573, the airplane must be shown by analysis or test to be free from
 flutter to VD/MD with the extent of damage for which residual strength is
 demonstrated.

 [Amdt. 23-23, 43 FR 50592, Oct. 30, 1978, as amended by Amdt. 23-31, 49 FR
 46867, Nov. 28, 1984; Amdt. No. 23-45, 58 FR 42164, Aug. 6, 1993; 58 FR
 51970, Oct. 5, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






                                     Wings






 Sec. 23.641  Proof of strength.

   The strength of stressed-skin wings must be proven by load tests or by
 combined structural analysis and load tests.






                               Control Surfaces






 Sec. 23.651  Proof of strength.

   (a) Limit load tests of control surfaces are required. These tests must
 include the horn or fitting to which the control system is attached.
   (b) In structural analyses, rigging loads due to wire bracing must be
 accounted for in a rational or conservative manner.






 Sec. 23.655  Installation.

   (a) Movable surfaces must be installed so that there is no interference
 between any surfaces, their bracing, or adjacent fixed structure, when one
 surface is held in its most critical clearance positions and the others are
 operated through their full movement.
   (b) If an adjustable stabilizer is used, it must have stops that will limit
 its range of travel to that allowing safe flight and landing.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. No. 23-45, 58 FR 42164, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.657  Hinges.

   (a) Control surface hinges, except ball and roller bearing hinges, must
 have a factor of safety of not less than 6.67 with respect to the ultimate
 bearing strength of the softest material used as a bearing.
   (b) For ball or roller bearing hinges, the approved rating of the bearing
 may not be exceeded.
   (c) Hinges must have enough strength and rigidity for loads parallel to the
 hinge line.






 Sec. 23.659  Mass balance.

   The supporting structure and the attachment of concentrated mass balance
 weights used on control surfaces must be designed for--
   (a) 24 g normal to the plane of the control surface;
   (b) 12 g fore and aft; and
   (c) 12 g parallel to the hinge line.






                                Control Systems






 Sec. 23.671  General.

   (a) Each control must operate easily, smoothly, and positively enough to
 allow proper performance of its functions.
   (b) Controls must be arranged and identified to provide for convenience in
 operation and to prevent the possibility of confusion and subsequent
 inadvertent operation.






 Sec. 23.672  Stability augmentation and automatic and power-operated systems.

   If the functioning of stability augmentation or other automatic or power-
 operated systems is necessary to show compliance with the flight
 characteristics requirements of this part, such systems must comply with Sec.
 23.671 and the following:
   (a) A warning, which is clearly distinguishable to the pilot under expected
 flight conditions without requiring the pilot's attention, must be provided
 for any failure in the stability augmentation system or in any other
 automatic or power-operated system that could result in an unsafe condition
 if the pilot was not aware of the failure. Warning systems must not activate
 the control system.
   (b) The design of the stability augmentation system or of any other
 automatic or power-operated system must permit initial counteraction of
 failures without requiring exceptional pilot skill or strength, by either the
 deactivation of the system or a failed portion thereof, or by overriding the
 failure by movement of the flight controls in the normal sense.
   (c) It must be shown that, after any single failure of the stability
 augmentation system or any other automatic or power-operated system--
   (1) The airplane is safely controllable when the failure or malfunction
 occurs at any speed or altitude within the approved operating limitations
 that is critical for the type of failure being considered;
   (2) The controllability and maneuverability requirements of this part are
 met within a practical operational flight envelope (for example, speed,
 altitude, normal acceleration, and airplane configuration) that is described
 in the Airplane Flight Manual (AFM); and
   (3) The trim, stability, and stall characteristics are not impaired below a
 level needed to permit continued safe flight and landing.

 [Amdt. No. 23-45, 58 FR 42164, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.673  Primary flight controls.

   (a) Primary flight controls are those used by the pilot for the immediate
 control of pitch, roll, and yaw.
   (b) The design of two-control airplanes must minimize the likelihood of
 complete loss of lateral or directional control in the event of failure of
 any connecting or transmitting element in the control system.






 Sec. 23.675  Stops.

   (a) Each control system must have stops that positively limit the range of
 motion of each movable aerodynamic surface controlled by the system.
   (b) Each stop must be located so that wear, slackness, or takeup
 adjustments will not adversely affect the control characteristics of the
 airplane because of a change in the range of surface travel.
   (c) Each stop must be able to withstand any loads corresponding to the
 design conditions for the control system.

 [Amdt. 23-17, 41 FR 55464, Dec. 20, 1976]






 Sec. 23.677  Trim systems.

   (a) Proper precautions must be taken to prevent inadvertent, improper, or
 abrupt trim tab operation. There must be means near the trim control to
 indicate to the pilot the direction of trim control movement relative to
 airplane motion. In addition, there must be means to indicate to the pilot
 the position of the trim device with respect to the range of adjustment. This
 means must be visible to the pilot and must be located and designed to
 prevent confusion.
   (b) Trimming devices must be designed so that, when any one connecting or
 transmitting element in the primary flight control system fails, adequate
 control for safe flight and landing is available with--
   (1) For single-engine airplanes, the longitudinal trimming devices; or
   (2) For multiengine airplanes, the longitudinal and directional trimming
 devices.
   (c) Tab controls must be irreversible unless the tab is properly balanced
 and has no unsafe flutter characteristics. Irreversible tab systems must have
 adequate rigidity and reliability in the portion of the system from the tab
 to the attachment of the irreversible unit to the airplane structure.
   (d) It must be demonstrated that the airplane is safely controllable and
 that the pilot can perform all maneuvers and operations necessary to effect a
 safe landing following any probable powered trim system runaway that
 reasonably might be expected in service, allowing for appropriate time delay
 after pilot recognition of the trim system runaway. The demonstration must be
 conducted at critical airplane weights and center of gravity positions.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13091, Aug. 13, 1969; Amdt. 23-34, 52 FR 1830, Jan. 15, 1987; Amdt. 23-42,
 56 FR 353, Jan. 3, 1991]

 *****************************************************************************


 56 FR 344, No. 2, Jan. 3, 1991

   SUMMARY: This final rule upgrades the airworthiness standards for normal,
 utility, acrobatic, and commuter category airplanes. This amendment provides
 airworthiness standards for advancements in technology being incorporated in
 current designs, permits type certification of spin resistant airplanes, and
 reduces the regulatory burden in showing compliance with some of the
 requirements for the design and type certification of small airplanes. These
 new and amended airworthiness standards also result in the need for new
 definitions. As a result, new definitions are added.

   DATES: February 4, 1991.

 *****************************************************************************






 Sec. 23.679  Control system locks.

   If there is a device to lock the control system on the ground or water:
   (a) There must be a means to--
   (1) Give unmistakable warning to the pilot when lock is engaged; or
   (2) Automatically disengage the device when the pilot operates the primary
 flight controls in a normal manner.
   (b) The device must be installed to limit the operation of the airplane so
 that, when the device is engaged, the pilot receives unmistakable warning at
 the start of the takeoff.
   (c) The device must have a means to preclude the possibility of it becoming
 inadvertently engaged in flight.

 [Amdt. No. 23-45, 58 FR 42164, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.681  Limit load static tests.

   (a) Compliance with the limit load requirements of this part must be shown
 by tests in which--
   (1) The direction of the test loads produces the most severe loading in the
 control system; and
   (2) Each fitting, pulley, and bracket used in attaching the system to the
 main structure is included.
   (b) Compliance must be shown (by analyses or individual load tests) with
 the special factor requirements for control system joints subject to angular
 motion.






 Sec. 23.683  Operation tests.

   (a) It must be shown by operation tests that, when the controls are
 operated from the pilot compartment with the system loaded as prescribed in
 paragraph (b) of this section, the system is free from--
   (1) Jamming;
   (2) Excessive friction; and
   (3) Excessive deflection.
   (b) The prescribed test loads are--
   (1) For the entire system, loads corresponding to the limit airloads on the
 appropriate surface, or the limit pilot forces in Sec. 23.397(b), whichever
 are less; and
   (2) For secondary controls, loads not less than those corresponding to the
 maximum pilot effort established under Sec. 23.405.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13091, Aug. 13, 1969]






 Sec. 23.685  Control system details.

   (a) Each detail of each control system must be designed and installed to
 prevent jamming, chafing, and interference from cargo, passengers, loose
 objects, or the freezing of moisture.
   (b) There must be means in the cockpit to prevent the entry of foreign
 objects into places where they would jam the system.
   (c) There must be means to prevent the slapping of cables or tubes against
 other parts.
   (d) Each element of the flight control system must have design features, or
 must be distinctively and permanently marked, to minimize the possibility of
 incorrect assembly that could result in malfunctioning of the control system.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-17, 41 FR
 55464, Dec. 20, 1976]






 Sec. 23.687  Spring devices.

   The reliability of any spring device used in the control system must be
 established by tests simulating service conditions unless failure of the
 spring will not cause flutter or unsafe flight characteristics.






 Sec. 23.689  Cable systems.

   (a) Each cable, cable fitting, turnbuckle, splice, and pulley used must
 meet approved specifications. In addition--
   (1) No cable smaller than 1/8  inch diameter may be used in primary control
 systems;
   (2) Each cable system must be designed so that there will be no hazardous
 change in cable tension throughout the range of travel under operating
 conditions and temperature variations; and
   (3) There must be means for visual inspection at each fairlead, pulley,
 terminal, and turnbuckle.
   (b) Each kind and size of pulley must correspond to the cable with which it
 is used. Each pulley must have closely fitted guards to prevent the cables
 from being misplaced or fouled, even when slack. Each pulley must lie in the
 plane passing through the cable so that the cable does not rub against the
 pulley flange.
   (c) Fairleads must be installed so that they do not cause a change in cable
 direction of more than three degrees.
   (d) Clevis pins subject to load or motion and retained only by cotter pins
 may not be used in the control system.
   (e) Turnbuckles must be attached to parts having angular motion in a manner
 that will positively prevent binding throughout the range of travel.
   (f) Tab control cables are not part of the primary control system and may
 be less than 1/8  inch diameter in airplanes that are safely controllable
 with the tabs in the most adverse positions.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13091, Aug. 13, 1969]






 Sec. 23.693  Joints.

   Control system joints (in push-pull systems) that are subject to angular
 motion, except those in ball and roller bearing systems, must have a special
 factor of safety of not less than 3.33 with respect to the ultimate bearing
 strength of the softest material used as a bearing. This factor may be
 reduced to 2.0 for joints in cable control systems. For ball or roller
 bearings, the approved ratings may not be exceeded.






 Sec. 23.697  Wing flap controls.

   (a) Each wing flap control must be designed so that, when the flap has been
 placed in any position upon which compliance with the performance
 requirements of this part is based, the flap will not move from that position
 unless the control is adjusted or is moved by the automatic operation of a
 flap load limiting device.
   (b) The rate of movement of the flaps in response to the operation of the
 pilot's control or automatic device must give satisfactory flight and
 performance characteristics under steady or changing conditions of airspeed,
 engine power, and attitude.






 Sec. 23.699  Wing flap position indicator.

   There must be a wing flap position indicator for--
   (a) Flap installations with only the retracted and fully extended position,
 unless--
   (1) A direct operating mechanism provides a sense of "feel" and position
 (such as when a mechanical linkage is employed); or
   (2) The flap position is readily determined without seriously detracting
 from other piloting duties under any flight condition, day or night; and
   (b) Flap installation with intermediate flap positions if--
   (1) Any flap position other than retracted or fully extended is used to
 show compliance with the performance requirements of this part; and
   (2) The flap installation does not meet the requirements of paragraph
 (a)(1) of this section.






 Sec. 23.701  Flap interconnection.

   (a) The main wing flaps and related movable surfaces as a system must--
   (1) Be synchronized by mechanical connection; or
   (2) Maintain synchronization so that the occurrence of an unsafe condition
 has been shown to be extremely improbable; or
   (b) The airplane must be shown to have safe flight characteristics with any
 combination of extreme positions of individual movable surfaces (mechanically
 interconnected surfaces are to be considered as a single surface).
   (c) If an interconnection is used in multiengine airplanes, it must be
 designed to account for the unsummetrical loads resulting from flight with
 the engines on one side of the plane of symmetry inoperative and the
 remaining engines at takeoff power. For single-engine airplanes, and
 multiengine airplanes with no slipstream effects on the flaps, it may be
 assumed that 100 percent of the critical air load acts on one side and 70
 percent on the other.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-14, 38 FR
 31821, Nov. 19, 1973; Amdt. 23-42, 56 FR 353, Jan. 3, 1991; Amdt. 23-42, 56
 FR 5455, Feb. 11, 1991]

 *****************************************************************************


 56 FR 344, No. 2, Jan. 3, 1991

   SUMMARY: This final rule upgrades the airworthiness standards for normal,
 utility, acrobatic, and commuter category airplanes. This amendment provides
 airworthiness standards for advancements in technology being incorporated in
 current designs, permits type certification of spin resistant airplanes, and
 reduces the regulatory burden in showing compliance with some of the
 requirements for the design and type certification of small airplanes. These
 new and amended airworthiness standards also result in the need for new
 definitions. As a result, new definitions are added.

   DATES: February 4, 1991.

 *****************************************************************************






                                 Landing Gear






 Sec. 23.721   General.

   For commuter category airplanes that have a passenger seating
 configuration, excluding pilot seats, of 10 or more, the following general
 requirements for the landing gear apply:
   (a) The main landing-gear system must be designed so that if it fails due
 to overloads during takeoff and landing (assuming the overloads to act in the
 upward and aft directions), the failure mode is not likely to cause the
 spillage of enough fuel from any part of the fuel system to consitute a fire
 hazard.
   (b) Each airplane must be designed so that, with the airplane under
 control, it can be landed on a paved runway with any one or more landing-gear
 legs not extended without sustaining a structural component failure that is
 likely to cause the spillage of enough fuel to consitute a fire hazard.
   (c) Compliance with the provisions of this section may be shown by analysis
 or tests, or both.

 [Amdt. 23-34, 52 FR 1830, Jan. 15, 1987]






 Sec. 23.723  Shock absorption tests.

   (a) It must be shown that the limit load factors selected for design in
 accordance with Sec. 23.473 for takeoff and landing weights, respectively,
 will not be exceeded. This must be shown by energy absorption tests except
 that analysis based on tests conducted on a landing gear system with
 identical energy absorption characteristics may be used for increases in
 previously approved takeoff and landing weights.
   (b) The landing gear may not fail, but may yield, in a test showing its
 reserved energy absorption capacity, simulating a descent velocity of 1.2
 times the limit descent velocity, assuming wing lift equal to the weight of
 the airplane.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. 23-23, 43 FR 50593, Oct. 30, 1978]






 Sec. 23.725  Limit drop tests.

   (a) If compliance with Sec. 23.723(a) is shown by free drop tests, these
 tests must be made on the complete airplane, or on units consisting of wheel,
 tire, and shock absorber, in their proper relation, from free drop heights
 not less than those determined by the following formula:

                          h (inches) = 3.6 (W/S)  1/2

 However, the free drop height may not be less than 9.2 inches and need not be
 more than 18.7 inches.
   (b) If the effect of wing lift is provided for in free drop tests, the
 landing gear must be dropped with an effective weight equal to

                                      h+(1-L)d
                            We = W x  ------------
                                      h+d

 where--

 We =the effective weight to be used in the drop test (lbs.);
 h = specified free drop height (inches);
 d = deflection under impact of the tire (at the approved inflation pressure)
     plus the vertical component of the axle travel relative to the drop mass
     (inches);
 W=WM for main gear units (lbs), equal to the static weight on that unit with
     the airplane in the level attitude (with the nose wheel clear in the case
     of nose wheel type airplanes);
 W=WT for tail gear units (lbs.), equal to the static weight on the tail unit
     with the airplane in the tail-down attitude;
 W=WN for nose wheel units lbs.), equal to the vertical component of the
     static reaction that would exist at the nose wheel, assuming that the
     mass of the airplane acts at the center of gravity and exerts a force of
     1.0 g downward and 0.33 g forward; and
 L= the ratio of the assumed wing lift to the airplane weight, but not more
     than 0.667.

   (c) The limit inertia load factor must be determined in a rational or
 conservative manner, during the drop test, using a landing gear unit
 attitude, and applied drag loads, that represent the landing conditions.
   (d) The value of d used in the computation of We in paragraph (b) of this
 section may not exceed the value actually obtained in the drop test.
   (e) The limit inertia load factor must be determined from the drop test in
 paragraph (b) of this section according to the following formula:

                                       We
                               n = nj  ----  + L
                                       W

 where--

 nj=the load factor developed in the drop test (that is, the acceleration (dv/
     dt) in g's recorded in the drop test) plus 1.0; and
 We, W, and L are the same as in the drop test computation.

   (f) The value of n determined in accordance with paragraph (e) may not be
 more than the limit inertia load factor used in the landing conditions in
 Sec. 23.473.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13091, Aug. 13, 1969]






 Sec. 23.726  Ground load dynamic tests.

   (a) If compliance with the ground load requirements of Secs. 23.479 through
 23.483 is shown dynamically by drop test, one drop test must be conducted
 that meets Sec. 23.725 except that the drop height must be--
   (1) 2.25 times the drop height prescribed in Sec. 23.725(a); or
   (2) Sufficient to develop 1.5 times the limit load factor.
   (b) The critical landing condition for each of the design conditions
 specified in Secs. 23.479 through 23.483 must be used for proof of strength.

 [Amdt. 23-7, 34 FR 13091, Aug. 13, 1969]






 Sec. 23.727  Reserve energy absorption drop test.

   (a) If compliance with the reserve energy absorption requirement in Sec.
 23.723(b) is shown by free drop tests, the drop height may not be less than
 1.44 times that specified in Sec. 23.725.
   (b) If the effect of wing lift is provided for, the units must be dropped
 with an effective mass equal to We=Wh/(h+d), when the symbols and other
 details are the same as in Sec. 23.725.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13091, Aug. 13, 1969]






 Sec. 23.729  Landing gear extension and retraction system.

   (a) General. For airplanes with retractable landing gear, the following
 apply:
   (1) Each landing gear retracting mechanism and its supporting structure
 must be designed for maximum flight load factors with the gear retracted and
 must be designed for the combination of friction, inertia, brake torque, and
 air loads, occurring during retraction at any airspeed up to 1.6 VS1 with
 flaps retracted, and for any load factor up to those specified in Sec. 23.345
 for the flaps-extended condition.
   (2) The landing gear and retracting mechanism, including the wheel well
 doors, must withstand flight loads, including loads resulting from all yawing
 conditions specified in Sec. 23.351, with the landing gear extended at any
 speed up to at least 1.6 VS1 with the flaps retracted.
   (b) Landing gear lock. There must be positive means (other than the use of
 hydraulic pressure) to keep the landing gear extended.
   (c) Emergency operation. For a landplane having retractable landing gear
 that cannot be extended manually, there must be means to extend the landing
 gear in the event of either--
   (1) Any reasonably probable failure in the normal landing gear operation
 system; or
   (2) Any reasonably probable failure in a power source that would prevent
 the operation of the normal landing gear operation system.
   (d) Operation test. The proper functioning of the retracting mechanism must
 be shown by operation tests.
   (e) Position indicator. If a retractable landing gear is used, there must
 be a landing gear position indicator (as well as necessary switches to
 actuate the indicator) or other means to inform the pilot that the gear is
 secured in the extended (or retracted) position. If switches are used, they
 must be located and coupled to the landing gear mechanical system in a manner
 that prevents an erroneous indication of either "down and locked" if the
 landing gear is not in a fully extended position, or of "up and locked" if
 the landing gear is not in the fully retracted position. The switches may be
 located where they are operated by the actual landing gear locking latch or
 device.
   (f) Landing gear warning. For landplanes, the following aural or equally
 effective landing gear warning devices must be provided:
   (1) A device that functions continuously when one or more throttles are
 closed beyond the power settings normally used for landing approach if the
 landing gear is not fully extended and locked. A throttle stop may not be
 used in place of an aural device. If there is a manual shutoff for the
 warning device prescribed in this paragraph, the warning system must be
 designed so that when the warning has been suspended after one or more
 throttles are closed, subsequent retardation of any throttle to, or beyond,
 the position for normal landing approach will activate the warning device.
   (2) A device that functions continuously when the wing flaps are extended
 beyond the maximum approach flap position, using a normal landing procedure,
 if the landing gear is not fully extended and locked. There may not be a
 manual shutoff for this warning device. The flap position sensing unit may be
 installed at any suitable location. The system for this device may use any
 part of the system (including the aural warning device) for the device
 required in paragraph (f)(1) of this section.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13091, Aug. 13, 1969; Amdt. 23-21, 43 FR 2318, Jan. 1978; Amdt. 23-26, 45 FR
 60171, Sept. 11, 1980; Amdt. No. 23-45, 58 FR 42164, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.731  Wheels.

   (a) The maximum static load rating of each wheel may not be less than the
 corresponding static ground reaction with--
   (1) Design maximum weight; and
   (2) Critical center of gravity.
   (b) The maximum limit load rating of each wheel must equal or exceed the
 maximum radial limit load determined under the applicable ground load
 requirements of this part.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. No. 23-45, 58 FR 42165, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.733  Tires.

   (a) Each landing gear wheel must have a tire whose approved tire ratings
 (static and dynamic) are not exceeded--
   (1) By a load on each main wheel tire) to be compared to the static rating
 approved for such tires) equal to the corresponding static ground reaction
 under the design maximum weight and critical center of gravity; and
   (2) By a load on nose wheel tires (to be compared with the dynamic rating
 approved for such tires) equal to the reaction obtained at the nose wheel,
 assuming the mass of the airplane to be concentrated at the most critical
 center of gravity and exerting a force of 1.0 W downward and 0.31 W forward
 (where W is the design maximum weight), with the reactions distributed to the
 nose and main wheels by the principles of statics and with the drag reaction
 at the ground applied only at wheels with brakes.
   (b) If specially constructed tires are used, the wheels must be plainly and
 conspicuously marked to that effect. The markings must include the make,
 size, number of plies, and identification marking of the proper tire.
   (c) Each tire installed on a retractable landing gear system must, at the
 maximum size of the tire type expected in service, have a clearance to
 surrounding structure and systems that is adequate to prevent contact between
 the tire and any part of the structure of systems.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13092, Aug. 13, 1969; Amdt. 23-17, 41 FR 55464, Dec. 20, 1976; Amdt. No.
 23-45, 58 FR 42165, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.735  Brakes.

   (a) Brakes must be provided so that the brake kinetic energy capacity
 rating of each main wheel brake assembly is not less than the kinetic energy
 absorption requirements determined under either of the following methods:
   (1) The brake kinetic energy absorption requirements must be based on a
 conservative rational analysis of the sequence of events expected during
 landing at the design landing weight.
   (2) Instead of a rational analysis, the kinetic energy absorption
 requirements for each main wheel brake assembly may be derived from the
 following formula:

   KE=0.0443 WV**2/N

 where--

 KE=Kinetic energy per wheel (ft.-lb.);
 W=Design landing weight (lb.);
 V=Airplane speed in knots. V must be not less than Vs<radical>, the poweroff
     stalling speed of the airplane at sea level, at the design landing
     weight, and in the landing configuration; and
 N=Number of main wheels with brakes.

   (b) Brakes must be able to prevent the wheels from rolling on a paved
 runway with takeoff power on the critical engine, but need not prevent
 movement of the airplane with wheels locked.
   (c) If antiskid devices are installed, the devices and associated systems
 must be designed so that no single probable malfunction or failure will
 result in a hazardous loss of braking ability or directional control of the
 airplane.

 [Amdt. 23-7, 34 FR 13092, Aug. 13, 1969, as amended by Amdt. 23-24, 44 FR
 68742, Nov. 29, 1979; Amdt. 23-42, 56 FR 354, Jan. 3, 1991]

 *****************************************************************************


 56 FR 344, No. 2, Jan. 3, 1991

   SUMMARY: This final rule upgrades the airworthiness standards for normal,
 utility, acrobatic, and commuter category airplanes. This amendment provides
 airworthiness standards for advancements in technology being incorporated in
 current designs, permits type certification of spin resistant airplanes, and
 reduces the regulatory burden in showing compliance with some of the
 requirements for the design and type certification of small airplanes. These
 new and amended airworthiness standards also result in the need for new
 definitions. As a result, new definitions are added.

   DATES: February 4, 1991.

 *****************************************************************************






 Sec. 23.737  Skis.

   The maximum limit load rating of each ski must equal or exceed the maximum
 limit load determined under the applicable ground load requirements of this
 part.

 [Amdt. No. 73-45, 58 FR 42165, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






                               Floats and Hulls






 Sec. 23.751  Main float buoyancy.

   (a) Each main float must have--
   (1) A buoyancy of 80 percent in excess of the buoyancy required by that
 float to support its portion of the maximum weight of the seaplane or
 amphibian in fresh water; and
   (2) Enough watertight compartments to provide reasonable assurance that the
 seaplane or amphibian will stay afloat without capsizing if any two
 compartments of any main float are flooded.
   (b) Each main float must contain at least four watertight compartments
 approximately equal in volume.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. No. 23-45, 58 FR 42165, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.753  Main float design.

   Each seaplane main float must meet the requirements of Sec. 23.521.

 [Amdt. No. 23-45, 58 FR 42165, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.755  Hulls.

   (a) The hull of a hull seaplane or amphibian of 1,500 pounds or more
 maximum weight must have watertight compartments designed and arranged so
 that the hull auxiliary floats, and tires (if used), will keep the airplane
 afloat without capsizing in fresh water when--
   (1) For airplanes of 5,000 pounds or more maximum weight, any two adjacent
 compartments are flooded; and
   (2) For airplanes of 1,500 pounds up to, but not including, 5,000 pounds
 maximum weight, any single compartment is flooded.
   (b) The hulls of hull seaplanes or amphibians of less than 1,500 pounds
 maximum weight need not be compartmented.
   (c) Bulkheads with watertight doors may be used for communication between
 compartments.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. No. 23-45, 58 FR 42165, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.757  Auxiliary floats.

   Auxiliary floats must be arranged so that, when completely submerged in
 fresh water, they provide a righting moment of at least 1.5 times the
 upsetting moment caused by the seaplane or amphibian being tilted.






                      Personnel and Cargo Accommodations






 Sec. 23.771  Pilot compartment.

   For each pilot compartment--
   (a) The compartment and its equipment must allow each pilot to perform his
 duties without unreasonable concentration or fatigue;
   (b) Where the flight crew are separated from the passengers by a partition,
 an opening or openable window or door must be provided to facilitate
 communication between flight crew and the passengers; and
   (c) The aerodynamic controls listed in Sec. 23.779, excluding cables and
 control rods, must be located with respect to the propellers so that no part
 of the pilot or the controls lies in the region between the plane of rotation
 of any inboard propeller and the surface generated by a line passing through
 the center of the propeller hub making an angle of 5 degrees forward or aft
 of the plane of rotation of the propeller.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-14, 38 FR
 31821, Nov. 19, 1973]






 Sec. 23.773   Pilot compartment view.

   (a) Each pilot compartment must be--
   (1) Arranged with sufficiently extensive, clear and undistorted view to
 enable the pilot to safely taxi, takeoff, approach, land, and perform any
 maneuvers within the operating limitations of the airplane.
   (2) Free from glare and reflections that could interfere with the pilot's
 vision. Compliance must be shown in all operations for which certification is
 requested; and
   (3) Designed so that each pilot is protected from the elements so that
 moderate rain conditions do not unduly impair the pilot's view of the flight
 path in normal flight and while landing.
   (b) Each pilot compartment must have a means to either remove or prevent
 the formation of fog or frost on an area of the internal portion of the
 windshield and side windows sufficiently large to provide the view specified
 in paragraph (a)(1) of this section. Compliance must be shown under all
 expected external and internal ambient operating conditions, unless it can be
 shown that the windshield and side windows can be easily cleared by the pilot
 without interruption of moral pilot duties.

 [Amdt. 23-45, 58 FR 42165, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.775  Windshields and windows.

   (a) Nonsplintering safety glass must be used in internal glass panes.
   (b) The design of windshields, windows, and canopies in pressurized
 airplanes must be based on factors peculiar to high altitude operation,
 including--
   (1) The effects of continuous and cyclic pressurization loadings;
   (2) The inherent characteristics of the material used; and
   (3) The effects of temperatures and temperature gradients.
   (c) On pressurized airplanes that do not comply with the fail-safe
 requirements of paragraph (e) of this section, an enclosure canopy including
 a representative part of the installation must be subjected to special tests
 to account for the combined effects of continuous and cyclic pressurization
 loadings and flight loads.
   (d) The windshield and side windows forward of the pilot's back when he is
 seated in the normal flight position must have a luminous transmittance value
 of not less than 70 percent.
   (e) If certification for operation above 25,000 feet is requested the
 windshields, window panels, and canopies must be strong enough to withstand
 the maximum cabin pressure differential loads combined with critical
 aerodynamic pressure and temperature effects, after failure of any load-
 carrying element of the windshield, window panel, or canopy.
   (f) Unless operation in known or forecast icing conditions is prohibited by
 operating limitations, a means must be provided to prevent or to clear
 accumulations of ice from the windshield so that the pilot has adequate view
 for taxi, takeoff, approach, landing, and to perform any maneuvers within the
 operating limitations of the airplane.
   (g) In the event of any probable single failure, a transparency heating
 system must be incapable of raising the temperature of any windshield or
 window to a point where there would be--
   (1) Structural failure that adversely affects the integrity of the cabin;
 or
   (2) There would be a danger of fire.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13092, Aug. 13, 1969; Amdt. No. 23-45, 58 FR 42165, Aug. 6, 1993; 58 FR
 51970, Oct. 5, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.777  Cockpit controls.

   (a) Each cockpit control must be located and (except where its function is
 obvious) identified to provide convenient operation and to prevent confusion
 and inadvertent operation.
   (b) The controls must be located and arranged so that the pilot, when
 seated, has full and unrestricted movement of each control without
 interference from either his clothing or the cockpit structure.
   (c) Powerplant controls must be located--
   (1) For multiengine airplanes, on the pedestal or overhead at or near the
 center of the cockpit;
   (2) For tandem seated single-engine airplanes, on the left side console or
 instrument panel;
   (3) For other single-engine airplanes at or near the center of the cockpit,
 on the pedestal, instrument panel, or overhead; and
   (4) For airplanes, with side-by-side pilot seats and with two sets of
 powerplant controls, on left and right consoles.
   (d) The control location order from left to right must be power (thrust)
 lever, propeller (rpm control), and mixture control (condition lever and fuel
 cutoff for turbine-powered airplanes). Power (thrust) levers must be at least
 one inch higher or longer to make them more prominent than propeller (rpm
 control) or mixture controls. Carburetor heat or alternate air control must
 be to the left of the throttle or at least eight inches from the mixture
 control when located other than on a pedestal. Carburetor heat or alternate
 air control, when located on a pedestal must be aft or below the power
 (thrust) lever. Supercharger controls must be located below or aft of the
 propeller controls. Airplanes with tandem seating or single-place airplanes
 may utilize control locations on the left side of the cabin compartment;
 however, location order from left to right must be power (thrust) lever,
 propeller (rpm control) and mixture control.
   (e) Identical powerplant controls for each engine must be located to
 prevent confusion as to the engines they control.
   (1) Conventional multiengine powerplant controls must be located so that
 the left control(s) operates the left engines(s) and the right control(s)
 operates the right engine(s).
   (2) On twin-engine airplanes with front and rear engine locations (tandem),
 the left powerplant controls must operate the front engine and the right
 powerplant controls must operate the rear engine.
   (f) Wing flap and auxiliary lift device controls must be located--
   (1) Centrally, or to the right of the pedestal or powerplant throttle
 control centerline; and
   (2) Far enough away from the landing gear control to avoid confusion.
   (g) The landing gear control must be located to the left of the throttle
 centerline or pedestal centerline.
   (h) Each fuel feed selector control must comply with Sec. 23.995 and be
 located and arranged so that the pilot can see and reach it without moving
 any seat or primary flight control when his seat is at any position in which
 it can be placed.
   (1) For a mechanical fuel selector:
   (i) The indication of the selected fuel valve position must be by means of
 a pointer and must provide positive identification and feel (detent, etc.) of
 the selected position.
   (ii) The position indicator pointer must be located at the part of the
 handle that is the maximum dimension of the handle measured from the center
 of rotation.
   (2) For electrical or electronic fuel selector:
   (i) Digital controls or electrical switches must be properly labelled.
   (ii) Means must be provided to indicate to the flight crew the tank or
 function selected. Selector switch position is not acceptable as a means of
 indication. The "off" or "closed" position must be indicated in red.
   (3) If the fuel valve selector handle or electrical or digital selection is
 also a fuel shut-off selector, the off position marking must be colored red.
 If a separate emergency shut-off means is provided, it also must be colored
 red.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13092, Aug. 13, 1969; Amdt. 23-33, 51 FR 26656, July 24, 1986]






 Sec. 23.779  Motion and effect of cockpit controls.

   Cockpit controls must be designed so that they operate in accordance with
 the following movement and actuation:
   (a) Aerodynamic controls:

                                                 Motion and effect

 (1) Primary controls:
   Aileron                              Right (clockwise) for right wing down.
   Elevator                             Rearward for nose up.
   Rudder                               Right pedal forward for nose right.
 (2) Secondary controls:
   Flaps (or auxiliary lift devices)    Forward or up for flaps up or
                                         auxiliary device stowed; rearward or
                                         down for flaps down or auxiliary
                                         device deployed.
   Trim tabs (or equivalent)            Switch motion or mechanical rotation
                                         of control to produce similar
                                         rotation of the airplane about an
                                         axis parallel to the axis control.
                                         Axis of roll trim control may be
                                         displaced to accommodate comfortable
                                         actuation by the pilot. For single-
                                         engine airplanes, direction of
                                         pilot's hand movement must be in the
                                         same sense as airplane response for
                                         rudder trim if only a portion of a
                                         rotational element is accessible.

   (b) Powerplant and auxiliary controls:

                                                   Motion and effect

  (1) Powerplant controls:
    Power (thrust) lever                 Forward to increase forward thrust
                                          and rearward to increase rearward
                                          thrust.
    Propellers                           Forward to increase rpm.
    Mixture                              Forward or upward for rich.
    Carburetor, air heat or alternate    Forward or upward for cold.
     air
    Supercharger                         Forward or upward for low blower.
    Turbosuperchargers                   Forward, upward, or clockwise to
                                          increase pressure.
    Rotary controls                      Clockwise from off to full on.
  (2) Auxiliary controls:
    Fuel tank selector                   Right for right tanks, left for left
                                          tanks.
    Landing gear                         Down to extend.
    Speed brakes                         Aft to extend.

 [Amdt. 23-33, 51 FR 26656, July 24, 1986]






 Sec. 23.781  Cockpit control knob shape.

   (a) Flap and landing gear control knobs must conform to the general shapes
 (but not necessarily the exact sizes or specific proportions) in the
 following figure:

                      [ ...Illustration appears here... ]

                               Flap Control Knob

                      [ ...Illustration appears here... ]

                           Landing Gear Control Knob

   (b) Powerplant control knobs must conform to the general shapes (but not
 necessarily the exact sizes or specific proportions) in the following figure:

                      [ ...Illustration appears here... ]

                          Power (Thrust) Control Knob

                      [ ...Illustration appears here... ]

                               RPM Control Knob

                      [ ...Illustration appears here... ]

                              Mixture Control Knob

                      [ ...Illustration appears here... ]

                    Carb Heat or Alternate Air Control Knob

                      [ ...Illustration appears here... ]

                           Supercharger Control Knob

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. 23-33, 51 FR 26657, July 24, 1986]






 Sec. 23.783  Doors.

   (a) Each closed cabin with passenger accommodations must have at least one
 adequate and easily accessible external door.
   (b) No passenger door may be located with respect to any propeller disc so
 as to endanger persons using that door.
   (c) Each external passenger or crew door must comply with the following
 requirements:
   (1) There must be a means to lock and safeguard the door against
 inadvertent opening during flight by persons, by cargo, or as a result of
 mechanical failure.
   (2) The door must be openable from the inside and the outside when the
 internal locking mechanism is in the locked position.
   (3) There must be a means of opening which is simple and obvious and is
 arranged and marked inside and outside so that the door can be readily
 located, unlocked, and opened, even in darkness.
   (4) The door must meet the marking requirements of Sec. 23.811 of this
 part.
   (5) The door must be reasonably free from jamming as a result of fuselage
 deformation in an emergency landing.
   (6) Auxiliary locking devices that are actuated externally to the airplane
 may be used but such devices must be overridden by the normal internal
 opening means.
   (d) In addition, each external passenger or crew door, for a commuter
 category airplane, must comply with the following requirements:
   (1) Each door must be openable from both the inside and outside, even
 though persons may be crowded against the door on the inside of the airplane.
   (2) If inward opening doors are used, there must be a means to prevent
 occupants from crowding against the door to the extent that would interfere
 with opening the door.
   (3) Auxiliary locking devices may be used.
   (e) Each external door on a commuter category airplane, each external door
 forward of any engine or propeller on a normal, utility, or acrobatic
 category airplane, and each door of the pressure vessel on a pressurized
 airplane must comply with the following requirements:
   (1) There must be a means to lock and safeguard each external door,
 including cargo and service type doors, against inadvertent opening in
 flight, by persons, by cargo, or as a result of mechanical failure or failure
 of a single structural element, either during or after closure.
   (2) There must be a provision for direct visual inspection of the locking
 mechanism to determine if the external door, for which the initial opening
 movement is not inward, is fully closed and locked. The provisions must be
 discernible, under operating lighting conditions, by a crewmember using a
 flashlight or an equivalent lighting source.
   (3) There must be a visual warning means to signal a flight crewmember if
 the external door is not fully closed and locked. The means must be designed
 so that any failure, or combination of failures, that would result in an
 erroneous closed and locked indication is improbable for doors for which the
 initial opening movement is not inward.

 [Docket No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. 23-36, 53 FR 30813, Aug. 15, 1988]






 Sec. 23.785  Seats, berths, litters, safety belts, and shoulder harnesses.

   (a) Each seat/restraint system and the supporting structure must be
 designed to support occupants weighing at least 215 pounds when subjected to
 the maximum load factors corresponding to the specified flight and ground
 load conditions, as defined in the approved operating envelope of the
 airplane. In addition, these loads must be multiplied by a factor of 1.33 in
 determining the strength of all fittings and the attachment of--
   (1) Each seat to the structure; and
   (2) Each safety belt and shoulder harness to the seat or structure.
   (b) Each forward-facing or aft-facing seat/restraint system in normal,
 utility, or acrobatic category airplanes must consist of a seat, safety belt,
 and shoulder harness that are designed to provide the occupant protection
 provisions required in Sec. 23.562 of this part. Other seat orientations must
 provide the same level of occupant protection as a forward-facing or aft-
 facing seat with a safety belt and shoulder harness, and provide the
 protection provisions of Sec. 23.562 of this part.
   (c) For commuter category airplanes, each seat and the supporting structure
 must be designed for occupants weighing at least 170 pounds when subjected to
 the inertia loads resulting from the ultimate static load factors prescribed
 in Sec. 23.561(b)(2) of this part, and each occupant must be protected from
 serious head injury when subjected to the inertia loads resulting from these
 load factors by a safety belt and shoulder harness for the front seats; and a
 safety belt, or a safety belt and shoulder harness, for each seat other than
 the front seats.
   (d) Each restraint system must have a single-point release for occupant
 evacuation.
   (e) The restraint system for each crewmember must allow the crewmember,
 when seated with the safety belt and shoulder harness fastened, to perform
 all functions necessary for flight operations.
   (f) Each pilot seat must be designed for the reactions resulting from the
 application of pilot forces to the primary flight controls as prescribed in
 Sec. 23.395 of this part.
   (g) There must be a means to secure each safety belt and shoulder harness,
 when not in use, to prevent interference with the operation of the airplane
 and with rapid occupant egress in an emergency.
   (h) Unless otherwise placarded, each seat in a utility or acrobatic
 category airplane must be designed to accommodate an occupant wearing a
 parachute.
   (i) The cabin area surrounding each seat, including the structure, interior
 walls, instrument panel, control wheel, pedals, and seats within striking
 distance of the occupant's head or torso (with the restraint system fastened)
 must be free of potentially injurious objects, sharp edges, protuberances,
 and hard surfaces. If energy absorbing designs or devices are used to meet
 this requirement, they must protect the occupant from serious injury when the
 occupant is subjected to the inertia loads resulting from the ultimate static
 load factors prescribed in Sec. 23.561(b)(2) of this part, or they must
 comply with the occupant protection provisions of Sec. 23.562 of this part,
 as required in paragraphs (b) and (c) of this section.
   (j) Each seat track must be fitted with stops to prevent the seat from
 sliding off the track.
   (k) Each seat/restraint system may use design features, such as crushing or
 separation of certain components, to reduce occupant loads when showing
 compliance with the requirements of Sec. 23.562 of this part; otherwise, the
 system must remain intact.
   (l) For the purposes of this section, a front seat is a seat located at a
 flight crewmember station or any seat located alongside such a seat.
   (m) Each berth, or provisions for a litter, installed parallel to the
 longitudinal axis of the airplane, must be designed so that the forward part
 has a padded end-board, canvas diaphragm, or equivalent means that can
 withstand the load reactions from a 215-pound occupant when subjected to the
 inertia loads resulting from the ultimate static load factors of Sec.
 23.561(b)(2) of this part. In addition--
   (1) Each berth or litter must have an occupant restraint system and may not
 have corners or other parts likely to cause serious injury to a person
 occupying it during emergency landing conditions; and
   (2) Occupant restraint system attachments for the berth or litter must
 withstand the inertia loads resulting from the ultimate static load factors
 of Sec. 23.561(b)(2) of this part.
   (n) Proof of compliance with the static strength requirements of this
 section for seats and berths approved as part of the type design and for seat
 and berth installations may be shown by--
   (1) Structural analysis, if the structure conforms to conventional airplane
 types for which existing methods of analysis are known to be reliable;
   (2) A combination of structural analysis and static load tests to limit
 load; or
   (3) Static load tests to ultimate loads.

 [Amdt. 23-36, 53 FR 30813, Aug. 15, 1988; Amdt. 23-36, 54 FR 50737, Dec. 11,
 1989]






 Sec. 23.787   Baggage and cargo compartments.

   (a) Each cargo compartment must be designed for its placarded maximum
 weight of contents and for the critical load distributions at the appropriate
 maximum load factors corresponding to the flight and ground load conditions
 of this part.
   (b) There must be means to prevent the contents of any cargo compartment
 from becoming a hazard by shifting, and to protect any controls, wiring,
 lines, equipment or accessories whose damage or failure would affect safe
 operations.
   (c) There must be a means to protect occupants from injury by the contents
 of any baggage or cargo compartment, located aft of the occupants and
 separated by structure, when the ultimate forward inertia load factor is 9g
 and assuming the maximun allowed baggage or cargo weight for the compartment.
   (d) Cargo compartments must be constructed of materials which are at least
 flame resistant.
   (e) Designs which provide for baggage or cargo to be carried in the same
 compartment as passengers must have a means to protect the occupants from
 injury when the cargo is subjected to the inertia loads resulting from the
 ultimate static load factors of Sec. 23.561(b)(3) of this part, assuming the
 maximum allowed baggage or cargo weight for the compartment.
   (f) If cargo compartment lamps are installed, each lamp must be installed
 so as to prevent contact between lamp bulb and cargo.
   (g) Baggage compartments used in commuter category airplanes must also meet
 the requirements of paragraphs (a), (b), (d), and (f) of this section.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-14, 38 FR
 31822, Nov. 19, 1973; Amdt. 23-17, 41 FR 55464, Dec. 20, 1976; Amdt. 23-34,
 52 FR 1830, Jan. 15, 1987; Amdt. 23-34, 52 FR 34745, Sept. 14, 1987; Amdt.
 23-36, 53 FR 30814, Aug. 15, 1988; 53 FR 34194, Sept. 2, 1988]






 Sec. 23.803   Emergency evacuation.

   For commuter category airplanes, an evacuation demonstration must be
 conducted utilizing the maximum number of occupants for which certification
 is desired. The demonstration must be conducted under simulated night
 conditions using only the emergency exits on the most critical side of the
 airplane. The participants must be representative of average airline
 passengers with no prior practice or rehearsal for the demonstration.
 Evacuation must be completed within 90 seconds.

 [Amdt. 23-34, 52 FR 1831, Jan. 15, 1987]






 Sec. 23.807  Emergency exits.

   (a) Number and location. Emergency exits must be located to allow escape
 without crowding in any probable crash attitude. The airplane must have at
 least the following emergency exits:
   (1) For all airplanes with a seating capacity of two or more, excluding
 airplanes with canopies, at least one emergency exit on the opposite side of
 the cabin from the main door specified in Sec. 23.783 of this part.
   (2) [Reserved]
   (3) If the pilot compartment is separated from the cabin by a door that is
 likely to block the pilot's escape in a minor crash, there must be an exit in
 the pilot's compartment. The number of exits required by paragraph (a)(1) of
 this section must then be separately determined for the passenger
 compartment, using the seating capacity of that compartment.
   (b) Type and operation. Emergency exits must be movable windows, panels,
 canopies, or external doors, openable from both inside and outside the
 airplane, that provide a clear and unobstructed opening large enough to admit
 a 19-by-26-inch ellipse. Auxiliary locking devices used to secure the
 airplane must be designed to be overridden by the normal internal opening
 means. In addition, each emergency exit must--
   (1) Be readily accessible, requiring no exceptional agility to be used in
 emergencies;
   (2) Have a method of opening that is simple and obvious;
   (3) Be arranged and marked for easy location and operation, even in
 darkness;
   (4) Have reasonable provisions against jamming by fuselage deformation; and
   (5) In the case of acrobatic category airplanes, allow each occupant to
 bail out quickly with parachutes at any speed between VSO and VD.
   (c) Tests. The proper functioning of each emergency exit must be shown by
 tests.
   (d) Doors and exits. In addition, for commuter category airplanes the
 following requirements apply:
   (1) The passenger entrance door must qualify as a floor level emergency
 exit. If an integral stair is installed at such a passenger entry door, the
 stair must be designed so that when subjected to the inertia forces specified
 in Sec. 23.561, and following the collapse of one or more legs of the landing
 gear, it will not interfere to an extent that will reduce the effectiveness
 of emergency egress through the passenger entry door. Each additional
 required emergency exit, except floor level exits, must be located over the
 wing or must be provided with acceptable means to assist the occupants in
 descending to the ground. In addition to the passenger entrance door--
   (i) For a total passenger seating capacity of 15 or less, an emergency exit
 as defined in paragraph (b) of this section is required on each side of the
 cabin; and
   (ii) For a total passenger seating capacity of 16 through 19, three
 emergency exits, as defined in paragraph (b) of this section, are required
 with one on the same side as the door and two on the side opposite the door.
   (2) A means must be provided to lock each emergency exit and to safeguard
 against its opening in flight, either inadvertently by persons or as a result
 of mechanical failure. In addition, a means for direct visual inspection of
 the locking mechanism must be provided to determine that each emergency exit
 for which the initial opening movement is outward is fully locked.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13092, Aug. 13, 1969; Amdt. 23-10, 36 FR 2864, Feb. 11, 1971; Amdt. 23-34, 52
 FR 1831, Jan. 15, 1987; Amdt. 23-36, 53 FR 30814, Aug. 15, 1988; 53 FR 34194,
 Sept. 2, 1988]






 Sec. 23.811   Emergency exit marking.

   (a) Each emergency exit and external door in the passenger compartment must
 be externally marked and readily identifiable from outside the airplane by--
   (1) A conspicuous visual identification scheme; and
   (2) A permanent decal or placard on or adjacent to the emergency exit which
 shows the means of opening the emergency exit, including any special
 instructions, if applicable.
   (b) In addition, for commuter category airplanes, these exits and doors
 must be internally marked with the word "exit" by a sign which has white
 letters 1 inch high on a red background 2 inches high, be self-illuminated or
 independently, internally electrically illuminated, and have a minimum
 brightness of at least 160 microlamberts. The color may be reversed if the
 passenger compartment illumination is essentially the same.

 [Amdt. 23-36, 53 FR 30814, Aug. 15, 1988; 53 FR 34194, Sept. 2, 1988]






 Sec. 23.813   Emergency exit access.

   For commuter category airplanes, access to window-type emergency exits may
 not be obstructed by seats or seat backs.

 [Amdt. 23-36, 53 FR 30815, Aug. 15, 1988]






 Sec. 23.815   Width of aisle.

   For commuter category airplanes, the width of the main passenger aisle at
 any point between seats must equal or exceed the values in the following
 table:

                                         Minimum main
                                       passenger aisle
                                            width

                                       Less
                                     than 25
                        Number of     inches   25 inches
                        passenger      from     and more
                          seats       floor    from floor

                      10 through 19  9 inches  15 inches.

 [Amdt. 23-34, 52 FR 1831, Jan. 15, 1987]






 Sec. 23.831  Ventilation.

   (a) Each passenger and crew compartment must be suitably ventilated. Carbon
 monoxide concentration may not exceed one part in 20,000 parts of air.
   (b) For pressurized airplanes, the ventilating air in the flightcrew and
 passenger compartments must be free of harmful or hazardous concentrations
 of gases and vapors in normal operations and in the event of reasonably
 probable failures or malfunctioning of the ventilating, heating,
 pressurization, or other systems and equipment. If accumulation of hazardous
 quantities of smoke in the cockpit area is reasonably probable, smoke
 evacuation must be readily accomplished starting with full pressurization
 and without depressurizing beyond safe limits.

 [Docket No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. 23-34, 52 FR 1831, Jan. 15, 1987; Amdt. 23-42, 56 FR 354,
 Jan. 3, 1991]

 *****************************************************************************


 56 FR 344, No. 2, Jan. 3, 1991

   SUMMARY: This final rule upgrades the airworthiness standards for normal,
 utility, acrobatic, and commuter category airplanes. This amendment provides
 airworthiness standards for advancements in technology being incorporated in
 current designs, permits type certification of spin resistant airplanes, and
 reduces the regulatory burden in showing compliance with some of the
 requirements for the design and type certification of small airplanes. These
 new and amended airworthiness standards also result in the need for new
 definitions. As a result, new definitions are added.

   DATES: February 4, 1991.

 *****************************************************************************






                                Pressurization






 Sec. 23.841  Pressurized cabins.

   (a) If certification for operation over 31,000 feet is requested, the
 airplane must be able to maintain a cabin pressure altitude of not more than
 15,000 feet in event of any probable failure or malfunction in the
 pressurization system.
   (b) Pressurized cabins must have at least the following valves, controls,
 and indicators, for controlling cabin pressure:
   (1) Two pressure relief valves to automatically limit the positive pressure
 differential to a predetermined value at the maximum rate of flow delivered
 by the pressure source. The combined capacity of the relief valves must be
 large enough so that the failure of any one valve would not cause an
 appreciable rise in the pressure differential. The pressure differential is
 positive when the internal pressure is greater than the external.
   (2) Two reverse pressure differential relief valves (or their equivalent)
 to automatically prevent a negative pressure differential that would damage
 the structure. However, one valve is enough if it is of a design that
 reasonably precludes its malfunctioning.
   (3) A means by which the pressure differential can be rapidly equalized.
   (4) An automatic or manual regulator for controlling the intake or exhaust
 airflow, or both, for maintaining the required internal pressures and airflow
 rates.
   (5) Instruments to indicate to the pilot the pressure differential, the
 cabin pressure altitude, and the rate of change of cabin pressure altitude.
   (6) Warning indication at the pilot station to indicate when the safe or
 preset pressure differential is exceeded and when a cabin pressure altitude
 of 10,000 feet is exceeded.
   (7) A warning placard for the pilot if the structure is not designed for
 pressure differentials up to the maximum relief valve setting in combination
 with landing loads.
   (8) A means to stop rotation of the compressor or to divert airflow from
 the cabin if continued rotation of an engine-driven cabin compressor or
 continued flow of any compressor bleed air will create a hazard if a
 malfunction occurs.

 [Amdt. 23-14, 38 FR 31822, Nov. 19, 1973, as amended by Amdt. 23-17, 41 FR
 55464, Dec. 20, 1976]






 Sec. 23.843  Pressurization tests.

   (a) Strength test. The complete pressurized cabin, including doors,
 windows, canopy, and valves, must be tested as a pressure vessel for the
 pressure differential specified in Sec. 23.365(d).
   (b) Functional tests. The following functional tests must be performed:
   (1) Tests of the functioning and capacity of the positive and negative
 pressure differential valves, and of the emergency release valve, to simulate
 the effects of closed regulator valves.
   (2) Tests of the pressurization system to show proper functioning under
 each possible condition of pressure, temperature, and moisture, up to the
 maximum altitude for which certification is requested.
   (3) Flight tests, to show the performance of the pressure supply, pressure
 and flow regulators, indicators, and warning signals, in steady and stepped
 climbs and descents at rates corresponding to the maximum attainable within
 the operating limitations of the airplane, up to the maximum altitude for
 which certification is requested.
   (4) Tests of each door and emergency exit, to show that they operate
 properly after being subjected to the flight tests prescribed in paragraph
 (b) (3) of this section.






                                Fire Protection






 Sec. 23.851   Fire extinguishers.

   (a) There must be at least one hand fire extinguisher for use in the pilot
 compartment that is located within easy access of the pilot while seated.
   (b) There must be at least one hand fire extinguisher located conveniently
 in the passenger compartment--
   (1) Of each airplane accommodating more than 6 passengers; and
   (2) Of each commuter category airplane.
   (c) For hand fire extinguishers, the following apply:
   (1) The type and quantity of each extinguishing agent used must be
 appropriate to the kinds of fire likely to occur where that agent is to be
 used.
   (2) Each extinguisher for use in a personnel compartment must be designed
 to minimize the hazard of toxic gas concentrations.

 [Amdt. 23-45, 58 FR 42165, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.853  Compartment interiors.

   For each compartment to be used by the crew or passengers:
   (a) The materials must be at least flame-resistant;
   (b) [Reserved]
   (c) If smoking is to be prohibited, there must be a placard so stating, and
 if smoking is to be allowed--
   (1) There must be an adequate number of self-contained, removable ashtrays;
 and
   (2) Where the crew compartment is separated from the passenger compartment,
 there must be at least one illuminated sign (using either letters or symbols)
 notifying all passengers when smoking is prohibited. Signs which notify when
 smoking is prohibited must--
   (i) When illuminated, be legible to each passenger seated in the passenger
 cabin under all probable lighting conditions; and
   (ii) Be so constructed that the crew can turn the illumination on and off;
 and
   (d) In addition, for commuter category airplanes the following requirements
 apply:
   (1) Each disposal receptacle for towels, paper, or waste must be fully
 enclosed and constructed of at least fire resistant materials and must
 contain fires likely to occur in it under normal use. The ability of the
 disposal receptacle to contain those fires under all probable conditions of
 wear, misalignment, and ventilation expected in service must be demonstrated
 by test. A placard containing the legible words "No Cigarette Disposal" must
 be located on or near each disposal receptacle door.
   (2) Lavatories must have "No Smoking" or "No Smoking in Lavatory" placards
 located conspicuously on each side of the entry door and self-contained,
 removable ashtrays located conspicuously on or near the entry side of each
 lavatory door, except that one ashtray may serve more than one lavatory door
 if it can be seen from the cabin side of each lavatory door served. The
 placards must have red letters at least 1/2  inch high on a white background
 at least 1 inch high (a "No Smoking" symbol may be included on the placard).
   (3) Materials (including finishes or decorative surfaces applied to the
 materials) used in each compartment occupied by the crew or passengers must
 meet the following test criteria as applicable:
   (i) Interior ceiling panels, interior wall panels, partitions, galley
 structure, large cabinet walls, structural flooring, and materials used in
 the construction of stowage compartments (other than underseat stowage
 compartments and compartments for stowing small items such as magazines and
 maps) must be self-extinguishing when tested vertically in accordance with
 the applicable portions of Appendix F of this part or by other equivalent
 methods. The average burn length may not exceed 6 inches and the average
 flame time after removal of the flame source may not exceed 15 seconds.
 Drippings from the test specimen may not continue to flame for more than an
 average of 3 seconds after falling.
   (ii) Floor covering, textiles (including draperies and upholstery), seat
 cushions, padding, decorative and nondecorative coated fabrics, leather,
 trays and galley furnishings, electrical conduit, thermal and acoustical
 insulation and insulation covering, air ducting, joint and edge covering,
 cargo compartment liners, insulation blankets, cargo covers and
 transparencies, molded and thermoformed parts, air ducting joints, and trim
 strips (decorative and chafing), that are constructed of materials not
 covered in paragraph (d)(3)(iv) of this section must be self extinguishing
 when tested vertically in accordance with the applicable portions of Appendix
 F of this part or other approved equivalent methods. The average burn length
 may not exceed 8 inches and the average flame time after removal of the flame
 source may not exceed 15 seconds. Drippings from the test specimen may not
 continue to flame for more than an average of 5 seconds after falling.
   (iii) Motion picture film must be safety film meeting the Standard
 Specifications for Safety Photographic Film PH1.25 (available from the
 American National Standards Institute, 1430 Broadway, New York, N.Y. 10018)
 or an FAA approved equivalent. If the film travels through ducts, the ducts
 must meet the requirements of paragraph (d)(3)(ii) of this section.
   (iv) Acrylic windows and signs, parts constructed in whole or in part of
 elastomeric materials, edge-lighted instrument assemblies consisting of two
 or more instruments in a common housing, seatbelts, shoulder harnesses, and
 cargo and baggage tiedown equipment, including containers, bins, pallets,
 etc., used in passenger or crew compartments, may not have an average burn
 rate greater than 2.5 inches per minute when tested horizontally in
 accordance with the applicable portions of Appendix F of this part or by
 other approved equivalent methods.
   (v) Except for electrical wire cable insulation, and for small parts (such
 as knobs, handles, rollers, fasteners, clips, grommets, rub strips, pulleys,
 and small electrical parts) that the Administrator finds would not contribute
 significantly to the propagation of a fire, materials in items not specified
 in paragraphs (d)(3) (i), (ii), (iii), or (iv) of this section may not have a
 burn rate greater than 4.0 inches per minute when tested horizontally in
 accordance with the applicable portions of Appendix F of this part or by
 other approved equivalent methods.
   (e) Lines, tanks, or equipment containing fuel, oil, or other flammable
 fluids may not be installed in such compartments unless adequately shielded,
 isolated, or otherwise protected so that any breakage or failure of such an
 item would not create a hazard.
   (f) Airplane materials located on the cabin side of the firewall must be
 self-extinguishing or be located at such a distance from the firewall, or
 otherwise protected, so that ignition will not occur if the firewall is
 subjected to a flame temperature of not less than 2,000 degrees F for 15
 minutes. For self-extinguishing materials (except electrical wire and cable
 insulation and small parts that the Administrator finds would not contribute
 significantly to the propagation of a fire), a vertifical self-extinguishing
 test must be conducted in accordance with Appendix F of this part or an
 equivalent method approved by the Administrator. The average burn length of
 the material may not exceed 6 inches and the average flame time after removal
 of the flame source may not exceed 15 seconds. Drippings from the material
 test specimen may not continue to flame for more than an average of 3 seconds
 after falling.

 [Amdt. 23-14, 23 FR 31822, Nov. 19, 1973, as amended by Amdt. 23-23, 43 FR
 50593, Oct. 30, 1978; Amdt. 23-25, 45 FR 7755, Feb. 4, 1980; Amdt. 23-34, 52
 FR 1831, Jan. 15, 1987]






 Sec. 23.859  Combustion heater fire protection.

   (a) Combustion heater fire regions. The following combustion heater fire
 regions must be protected from fire in accordance with the applicable
 provisions of Secs.23.1182 through 23.1191 and 23.1203:
   (1) The region surrounding the heater, if this region contains any
 flammable fluid system components (excluding the heater fuel system) that
 could--
   (i) Be damaged by heater malfunctioning; or
   (ii) Allow flammable fluids or vapors to reach the heater in case of
 leakage.
   (2) The region surrounding the heater, if the heater fuel system has
 fittings that, if they leaked, would allow fuel vapor to enter this region.
   (3) The part of the ventilating air passage that surrounds the combustion
 chamber.
   (b) Ventilating air ducts. Each ventilating air duct passing through any
 fire region must be fireproof. In addition--
   (1) Unless isolation is provided by fireproof valves or by equally
 effective means, the ventilating air duct downstream of each heater must be
 fireproof for a distance great enough to ensure that any fire originating in
 the heater can be contained in the duct; and
   (2) Each part of any ventilating duct passing through any region having a
 flammable fluid system must be constructed or isolated from that system so
 that the malfunctioning of any component of that system cannot introduce
 flammable fluids or vapors into the ventilating airstream.
   (c) Combustion air ducts. Each combustion air duct must be fireproof for a
 distance great enough to prevent damage from backfiring or reverse flame
 propagation. In addition--
   (1) No combustion air duct may have a common opening with the ventilating
 airstream unless flames from backfires or reverse burning cannot enter the
 ventilating airstream under any operating condition, including reverse flow
 or malfunctioning of the heater or its associated components; and
   (2) No combustion air duct may restrict the prompt relief of any backfire
 that, if so restricted, could cause heater failure.
   (d) Heater controls: general. Provision must be made to prevent the
 hazardous accumulation of water or ice on or in any heater control component,
 control system tubing, or safety control.
   (e) Heater safety controls. (1) Each combustion heater must have the
 following safety controls:
   (i) Means independent of the components for the normal continuous control
 of air temperature, airflow, and fuel flow must be provided to automatically
 shut off the ignition and fuel supply to that heater at a point remote from
 that heater when any of the following occurs:
   (A) The heater exchanger temperature exceeds safe limits.
   (B) The ventilating air temperature exceeds safe limits.
   (C) The combustion airflow becomes inadequate for safe operation.
   (D) The ventilating airflow becomes inadequate for safe operation.
   (ii) Means to warn the crew when any heater whose heat output is essential
 for safe operation has been shut off by the automatic means prescribed in
 paragraph (e)(1)(i@of this section.
   (2) The means for complying with paragraph (e)(1)(i) of this section for
 any individual heater must--
   (i) Be independent of components serving any other heater whose heat output
 is essential for safe operations; and
   (ii) Keep the heater off until restarted by the crew.
   (f) Air intakes. Each combustion and ventilating air intake must be located
 so that no flammable fluids or vapors can enter the heater system under any
 operating condition--
   (1) During normal operation; or
   (2) As a result of the malfunctioning of any other component.
   (g) Heater exhaust. Heater exhaust systems must meet the provisions of
 Secs. 23.1121 and 23.1123. In addition, there must be provisions in the
 design of the heater exhaust system to safely expel the products of
 combustion to prevent the occurrence of--
   (1) Fuel leakage from the exhaust to surrounding compartments;
   (2) Exhaust gas impingement on surrounding equipment or structure;
   (3) Ignition of flammable fluids by the exhaust, if the exhaust is in a
 compartment containing flammable fluid lines; and
   (4) Restrictions in the exhaust system to relieve backfires that, if so
 restricted, could cause heater failure.
   (h) Heater fuel systems. Each heater fuel system must meet each powerplant
 fuel system requirement affecting safe heater operation. Each heater fuel
 system component within the ventilating airstream must be protected by
 shrouds so that no leakage from those components can enter the ventilating
 airstream.
   (i) Drains. There must be means to safely drain fuel that might accumulate
 within the combustion chamber or the heater exchanger. In addition--
   (1) Each part of any drain that operates at high temperatures must be
 protected in the same manner as heater exhausts; and
   (2) Each drain must be protected from hazardous ice accumulation under any
 operating condition.

 [Amdt. 23--27, 45 FR 70387, Oct. 23, 1980]






 Sec. 23.863   Flammable fluid fire protection.

   (a) In each area where flammable fluids or vapors might escape by leakage
 of a fluid system, there must be means to minimize the probability of
 ignition of the fluids and vapors, and the resultant hazard if ignition does
 occur.
   (b) Compliance with paragraph (a) of this section must be shown by analysis
 or tests, and the following factors must be considered:
   (1) Possible sources and paths of fluid leakage, and means of detecting
 leakage.
   (2) Flammability characteristics of fluids, including effects of any
 combustible or absorbing materials.
   (3) Possible ignition sources, including electrical faults, overheating of
 equipment, and malfunctioning of protective devices.
   (4) Means available for controlling or extinguishing a fire, such as
 stopping flow of fluids, shutting down equipment, fireproof containment, or
 use of extinguishing agents.
   (5) Ability of airplane components that are critical to safety of flight to
 withstand fire and heat.
   (c) If action by the flight crew is required to prevent or counteract a
 fluid fire (e.g. equipment shutdown or actuation of a fire extinguisher),
 quick acting means must be provided to alert the crew.
   (d) Each area where flammable fluids or vapors might escape by leakage of a
 fluid system must be identified and defined.

 [Amdt. 23-23, 43 FR 50593, Oct. 30, 1978]






 Sec. 23.865   Fire protection of flight controls, engine mounts, and other
 flight structure.

   Flight controls, engine mounts, excluding those portions that are
 certificated as part of the engine, and other flight structure located in the
 engine compartment must be constructed of fireproof material or shielded so
 that they are capable of withstanding the effects of a fire. Engine vibration
 isolators must incorporate suitable features to ensure that the engine is
 retained if the non-fireproof portions of the isolators deteriorate from the
 effects of a fire.

 [Amdt. 23-45, 58 FR 42165, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.







                             Lightning Evaluation






 Sec. 23.867  Lightning protection of structure.

   (a) The airplane must be protected against catastrophic effects from
 lightning.
   (b) For metallic components, compliance with paragraph (a) of this section
 may be shown by--
   (1) Bonding the components properly to the airframe; or
   (2) Designing the components so that a strike will not endanger the
 airplane.
   (c) For nonmetallic components, compliance with paragraph (a) of this
 section may be shown by--
   (1) Designing the components to minimize the effect of a strike; or
   (2) Incorporating acceptable means of diverting the resulting electrical
 current so as not to endanger the airplane.

 [Amdt. 23-7, 34 FR 13092, Aug. 13, 1969]






                                 Miscellaneous






 Sec. 23.871  Leveling means.

   There must be means for determining when the airplane is in a level
 position on the ground.

 [Amdt. 23-7, 34 FR 13092, Aug. 13, 1969]


                             Subpart E--Powerplant

                                   General



 Sec. 23.901  Installation.

   (a) For the purpose of this part, the airplane powerplant installation
 includes each component that--
   (1) Is necessary for propulsion; and
   (2) Affects the safety of the major propulsive units.
   (b) Each powerplant installation must be constructed and arranged to--
   (1) Ensure safe operation to the maximum altitude for which approval is
 requested.
   (2) Be accessible for necessary inspections and maintenance.
   (c) Engine cowls and nacelles must be easily removable or openable by the
 pilot to provide adequate access to and exposure of the engine compartment
 for preflight checks.
   (d) Each turbine engine installation must be constructed and arranged to--
   (1) Result in vibration characteristics that do not exceed those
 established during the type certification of the engine.
   (2) Provide continued safe operation without a hazardous loss of power or
 thrust while being operated in rain for at least 3 minutes with the rate of
 water ingestion being not less than 4 percent by weight, of the engine
 induction airflow rate at the maximum installed power or thrust approved for
 takeoff and at flight idle. The engine must accelerate and decelerate safely
 following stabilized operation under these rain conditions.
   (e) The installation must comply with--
   (1) The instructions provided under the engine type certificate and the
 propeller type certificate.
   (2) The applicable provisions of this subpart.
   (f) Each auxiliary power unit installation must meet the applicable
 portions of this part.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13092, Aug. 13, 1969; Amdt. 23-18, 42 FR 15041, Mar. 17, 1977; Amdt. 23-29,
 49 FR 6846, Feb. 23, 1984; Amdt. 23-34, 52 FR 1832, Jan. 15, 1987; Amdt. 23-
 34, 52 FR 34745, Sept. 14, 1987; Amdt. 23-43, 58 FR 18970, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.903   Engines.

   (a) Engine type certificate.
   (1) Each engine must have a type certificate and must meet the applicable
 requirements of part 34 of this chapter.
   (2) Each turbine engine must either--
   (i) Comply with Sec. 33.77 of this chapter in effect on October 31, 1974,
 or as later amended; or
   (ii) Be shown to have a foreign object ingestion service history in similar
 installation locations which has not resulted in any unsafe condition.
   (b) Turbine engine installations. For turbine engine installations--
   (1) Design precautions must be taken to minimize the hazards to the
 airplane in the event of an engine rotor failure or of a fire originating
 inside the engine which burns through the engine case.
   (2) The powerplant systems associated with engine control devices, systems,
 and instrumentation must be designed to give reasonable assurance that those
 operating limitations that adversely affect turbine rotor structural
 integrity will not be exceeded in service.
   (c) The powerplants must be arranged and isolated from each other to allow
 operation, in at least one configuration, so that the failure or malfunction
 of any engine, or the failure or malfunction (including destruction by fire
 in the engine compartment) of any system that can affect an engine (other
 than a fuel tank if only one fuel tank is installed), will not:
   (1) Prevent the continued safe operation of the remaining engines; or
   (2) Require immediate action by any crewmember for continued safe operation
 of the remaining engines.
   (d) Starting and stopping (piston engine).
   (1) The design of the installation must be such that risk of fire or
 mechanical damage to the engine or airplane, as a result of starting the
 engine in any conditions in which starting is to be permitted, is reduced to
 a minimum. Any techniques and associated limitations for engine starting must
 be established and included in the Airplane Flight Manual, approved manual
 material, or applicable operating placards. Means must be provided for--
   (i) Restarting any engine of a multiengine airplane in flight, and
   (ii) Stopping any engine in flight, after engine failure, if continued
 engine rotation would cause a hazard to the airplane.
   (2) In addition, for commuter category airplanes, the following apply:
   (i) Each component of the stopping system on the engine side of the
 firewall that might be exposed to fire must be at least fire resistant.
   (ii) If hydraulic propeller feathering systems are used for this purpose,
 the feathering lines must be at least fire resistant under the operating
 conditions that may be expected to exist during feathering.
   (e) Starting and stopping (turbine engine). Turbine engine installations
 must comply with the following:
   (1) The design of the installation must be such that risk of fire or
 mechanical damage to the engine or the airplane, as a result of starting the
 engine in any conditions in which starting is to be permitted, is reduced to
 a minimum. Any techniques and associated limitations must be established and
 included in the Airplane Flight Manual, approved manual material, or
 applicable operating placards.
   (2) There must be means for stopping combustion within any engine and for
 stopping the rotation of any engine if continued rotation would cause a
 hazard to the airplane. Each component of the engine stopping system located
 in any fire zone must be fire resistant. If hydraulic propeller feathering
 systems are used for stopping the engine, the hydraulic feathering lines or
 hoses must be fire resistant.
   (3) It must be possible to restart an engine in flight. Any techniques and
 associated limitations must be established and included in the Airplane
 Flight Manual, approved manual material, or applicable operating placards.
   (4) It must be demonstrated in flight that when restarting engines
 following a false start, all fuel or vapor is discharged in such a way that
 it does not constitute a fire hazard.
   (f) Restart capability. An altitude and airspeed envelope must be
 established for the airplane for in-flight engine restarting and each
 installed engine must have a restart capability within that envelope.
   (g) For turbine engine powered airplanes, if the minimum windmilling speed
 of the engines, following the in-flight shutdown of all engines, is
 insufficient to provide the necessary electrical power for engine ignition, a
 power source independent of the engine-driven electrical power generating
 system must be provided to permit in-flight engine ignition for restarting.

 [Amdt. 23-14, 38 FR 31822, Nov. 19, 1973, as amended by Amdt. 23-17, 41 FR
 55464, Dec. 20, 1976; Amdt. 23-26, 45 FR 60171, Sept. 11, 1980; Amdt. 23-29,
 49 FR 6847, Feb. 23, 1984; Amdt. 23-34, 52 FR 1832, Jan. 15, 1987; Amdt. 23-
 40, 55 FR 32861, Aug. 10, 1990; Amdt. 23-43, 58 FR 18970, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.904  Automatic power reserve system.

   If installed, an automatic power reserve (APR) system that automatically
 advances the power or thrust on the operating engine(s), when any engine
 fails during takeoff, must comply with appendix H of this part.

 [Amdt. 23-43, 58 FR 18970, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.905   Propellers.

   (a) Each propeller must have a type certificate.
   (b) Engine power and propeller shaft rotational speed may not exceed the
 limits for which the propeller is certificated.
   (c) Each featherable propeller must have a means to unfeather it in flight.
   (d) Each component of the propeller blade pitch control system must meet
 the requirements of Sec. 35.42 of this chapter.
   (e) All areas of the airplane forward of the pusher propeller that are
 likely to accumulate and shed ice into the propeller disc during any
 operating condition must be suitably protected to prevent ice formation, or
 it must be shown that any ice shed into the propeller disc will not create a
 hazardous condition.
   (f) Each pusher propeller must be marked so that the disc is conspicuous
 under normal daylight ground conditions.
   (g) If the engine exhaust gases are discharged into the pusher propeller
 disc, it must be shown by tests, or analysis supported by tests, that the
 propeller is capable of continuous safe operation.
   (h) All engine cowling, access doors, and other removable items must be
 designed to ensure that they will not separate from the airplane and contact
 the pusher propeller.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-26, 45 FR
 60171, Sept. 11, 1980; Amdt. 23-29, 49 FR 6847, Feb. 23, 1984; Amdt. 23-43,
 58 FR 18970, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.907  Propeller vibration.

   (a) Each propeller with metal blades or highly stressed metal components
 must be shown to have vibration stresses, in normal operating conditions,
 that do not exceed values that have been shown by the propeller manufacturer
 to be safe for continuous operation. This must be shown by--
   (1) Measurement of stresses through direct testing of the propeller;
   (2) Comparison with similar installations for which these measurements have
 been made; or
   (3) Any other acceptable test method or service experience that proves the
 safety of the installation.
   (b) Proof of safe vibration characteristics for any type of propeller,
 except for conventional, fixed-pitch, wood propellers must be shown where
 necessary.






 Sec. 23.909  Turbocharger systems.

   (a) Each turbocharger must be approved under the engine type certificate or
 it must be shown that the turbocharger system, while in its normal engine
 installation and operating in the engine environment--
   (1) Can withstand, without defect, an endurance test of 150 hours that
 meets the applicable requirements of Sec. 33.49 of this subchapter; and
   (2) Will have no adverse effect upon the engine.
   (b) Control system malfunctions, vibrations, and abnormal speeds and
 temperatures expected in service may not damage the turbocharger compressor
 or turbine.
   (c) Each turbocharger case must be able to contain fragments of a
 compressor or turbine that fails at the highest speed that is obtainable with
 normal speed control devices inoperative.
   (d) Each intercooler installation, where provided, must comply with the
 following--
   (1) The mounting provisions of the intercooler must be designed to
 withstand the loads imposed on the system;
   (2) It must be shown that, under the installed vibration environment, the
 intercooler will not fail in a manner allowing portions of the intercooler to
 be ingested by the engine; and
   (3) Airflow through the intercooler must not discharge directly on any
 airplane component (e.g., windshield) unless such discharge is shown to cause
 no hazard to the airplane under all operating conditions.
   (e) Engine power, cooling characteristics, operating limits, and procedures
 affected by the turbocharger system installations must be evaluated.
 Turbocharger operating procedures and limitations must be included in the
 Airplane Flight Manual in accordance with Sec. 23.1581.

 [Amdt. 23-7, 34 FR 13092, Aug. 13, 1969, as amended by Amdt. 23-43, 58 FR
 18970, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.925  Propeller clearance.

   Unless smaller clearances are substantiated, propeller clearances with the
 airplane at maximum weight, with the most adverse center of gravity, and with
 the propeller in the most adverse pitch position, may not be less than the
 following:
   (a) Ground clearance. There must be a clearance of at least seven inches
 (for each airplane with nose wheel landing gear) or nine inches (for each
 airplane with tail wheel landing gear) between each propeller and the ground
 with the landing gear statically deflected and in the level, normal takeoff,
 or taxing attitude, whichever is most critical. In addition, for each
 airplane with conventional landing gear struts using fluid or mechanical
 means for absorbing landing shocks, there must be positive clearance between
 the propeller and the ground in the level takeoff attitude with the critical
 tire completely deflated and the corresponding landing gear strut bottomed.
 Positive clearance for airplanes using leaf spring struts is shown with a
 deflection corresponding to 1.5g.
   (b) Aft-mounted propellers. In addition to the clearances specified in
 paragraph (a) of this section, the airplane must be designed such that the
 propeller will not contact the runway surface when the airplane is in the
 maximum pitch attitude attainable during normal takeoff and landings. If a
 tail wheel, bumper, or an energy absorption device is provided to show
 compliance with this paragraph, the following apply:
   (1) Suitable design loads must be established for the tail wheel, bumper,
 or energy absorption device; and
   (2) The supporting structure of the tail wheel, bumper, or energy
 absorption device must be designed to withstand the loads established in
 paragraph (b)(1) of this section and inspection/replacement criteria must be
 established for the tail wheel, bumper, or energy absorption device and
 provided as part of the information required by Sec. 23.1529.
   (c) Water clearance. There must be a clearance of at least 18 inches
 between each propeller and the water, unless compliance with Sec. 23.239 can
 be shown with a lesser clearance.
   (d) Structural clearance. There must be--
   (1) At least one inch radial clearance between the blade tips and the
 airplane structure, plus any additional radial clearance necessary to prevent
 harmful vibration;
   (2) At least one-half inch longitudinal clearance between the propeller
 blades or cuffs and stationary parts of the airplane; and
   (3) Positive clearance between other rotating parts of the propeller or
 spinner and stationary parts of the airplane.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-43, 58 FR
 18971, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.929  Engine installation ice protection.

   Propellers (except wooden propellers) and other components of complete
 engine installations must be protected against the accumulation of ice as
 necessary to enable satisfactory functioning without appreciable loss of
 power when operated in the icing conditions for which certification is
 requested.

 [Amdt. 23-14, 33 FR 31822, Nov. 19, 1973]






 Sec. 23.933  Reversing systems.

   (a) For turbojet and turbofan reversing systems. (1) Each system intended
 for ground operation only must be designed so that no single failure or
 malfunction of the system will result in unwanted reverse thrust under any
 expected operating condition. Failure of structural elements need not be
 considered if the probability of this type of failure is extremely remote.
   (2) Each system intended for in-flight use must be designed so that no
 unsafe condition will result during normal operation of the system, or from
 any failure, or likely combination of failures, of the reversing system under
 any operating condition including ground operation. Failure of structural
 elements need not be considered if the probability of this type of failure is
 extremely remote.
   (3) Each system must have a means to prevent the engine from producing more
 than idle forward thrust when the reversing system malfunctions; except that
 it may produce any greater forward thrust that is shown to allow directional
 control to be maintained, with aerodynamic means alone, under the most
 critical reversing condition expected in operation.
   (b) For propeller reversing systems. (1) Each system must be designed so
 that no single failure, likely combination of failures or malfunction of the
 system will result in unwanted reverse thrust under any operating condition.
 Failure of structural elements need not be considered if the probability of
 this type of failure is extremely remote.
   (2) Compliance with paragraph (a)(1) of this section must be shown by
 failure analysis, or testing, or both, for propeller systems that allow the
 propeller blades to move from the flight low-pitch position to a position
 that is substantially less than the normal flight, low-pitch position. The
 analysis may include or be supported by the analysis made to show compliance
 with Sec. 35.21 for the type certification of the propeller and associated
 installation components. Credit will be given for pertinent analysis and
 testing completed by the engine and propeller manufacturers.

 [Amdt. 23-43, 58 FR 18971, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.934   Turbojet and turbofan engine thrust reverser systems tests.

   Thrust reverser systems of turbojet or turbofan engines must meet the
 requirements of Sec. 33.97 of this chapter or it must be demonstrated by
 tests that engine operation and vibratory levels are not affected.

 [Amdt. 23-43, 58 FR 18971, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.937  Turbopropeller-drag limiting systems.

   (a) Turbopropeller-powered airplane propeller-drag limiting systems must
 be designed so that no single failure or malfunction of any of the systems
 during normal or emergency operation results in propeller drag in excess of
 that for which the airplane was designed under the structural requirements of
 this part. Failure of structural elements of the drag limiting systems need
 not be considered if the probability of this kind of failure is extremely
 remote.
   (b) As used in this section, drag limiting systems include manual or
 automatic devices that, when actuated after engine power loss, can move the
 propeller blades toward the feather position to reduce windmilling drag to a
 safe level.

 [Amdt. 23-7, 34 FR 13093, Aug. 13, 1969, as amended by Amdt. 23-43, 58 FR
 18971, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.939  Powerplant operating characteristics.

   (a) Turbine engine powerplant operating characteristics must be
 investigated in flight to determine that no adverse characteristics (such as
 stall, surge, or flameout) are present, to a hazardous degree, during normal
 and emergency operation within the range of operating limitations of the
 airplane and of the engine.
   (b) Turbocharged reciprocating engine operating characteristics must be
 investigated in flight to assure that no adverse characteristics, as a result
 of an inadvertent overboost, surge, flooding, or vapor lock, are present
 during normal or emergency operation of the engine(s) throughout the range of
 operating limitations of both airplane and engine.
   (c) For turbine engines, the air inlet system must not, as a result of
 airflow distortion during normal operation, cause vibration harmful to the
 engine.

 [Amdt. 23-7, 34 FR 13093 Aug. 13, 1969, as amended by Amdt. 23-14, 38 FR
 31823, Nov. 19, 1973; Amdt. 23-18, 42 FR 15041, Mar. 17, 1977; Amdt. 23-42,
 56 FR 354, Jan. 3, 1991]

 *****************************************************************************


 56 FR 344, No. 2, Jan. 3, 1991

   SUMMARY: This final rule upgrades the airworthiness standards for normal,
 utility, acrobatic, and commuter category airplanes. This amendment provides
 airworthiness standards for advancements in technology being incorporated in
 current designs, permits type certification of spin resistant airplanes, and
 reduces the regulatory burden in showing compliance with some of the
 requirements for the design and type certification of small airplanes. These
 new and amended airworthiness standards also result in the need for new
 definitions. As a result, new definitions are added.

   DATES: February 4, 1991.

 *****************************************************************************






 Sec. 23.943  Negative acceleration.

   No hazardous malfunction of an engine, an auxiliary power unit approved for
 use in flight, or any component or system associated with the powerplant or
 auxiliary power unit may occur when the airplane is operated at the negative
 accelerations within the flight envelopes prescribed in Sec. 23.333. This
 must be shown for the greatest value and duration of the acceleration
 expected in service.

 [Amdt. 23-18, 42 FR 15041, Mar. 17, 1977, as amended by Amdt. 23-43, 58 FR
 18971, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






                                  Fuel System






 Sec. 23.951  General.

   (a) Each fuel system must be constructed and arranged to ensure fuel flow
 at a rate and pressure established for proper engine and auxiliary power unit
 functioning under each likely operating condition, including any maneuver for
 which certification is requested and during which the engine or auxiliary
 power unit is permitted to be in operation.
   (b) Each fuel system must be arranged so that--
   (1) No fuel pump can draw fuel from more than one tank at a time; or
   (2) There are means to prevent introducing air into the system.
   (c) Each fuel system for a turbine engine must be capable of sustained
 operation throughout its flow and pressure range with fuel initially
 saturated with water at 80 deg. F and having 0.75cc of free water per gallon
 added and cooled to the most critical condition for icing likely to be
 encountered in operation.
   (d) Each fuel system for a turbine engine powered airplane must meet the
 applicable fuel venting requirements of part 34 of this chapter.

 [Amdt. 23-15, 39 FR 35459, Oct. 1, 1974, as amended by Amdt. 23-40, 55 FR
 32861, Aug. 10, 1990; Amdt. 23-43, 58 FR 18971, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.953  Fuel system independence.

   (a) Each fuel system for a multiengine airplane must be arranged so that,
 in at least one system configuration, the failure of any one component (other
 than a fuel tank) will not result in the loss of power of more than one
 engine or require immediate action by the pilot to prevent the loss of power
 of more than one engine.
   (b) If a single fuel tank (or series of fuel tanks interconnected to
 function as a single fuel tank) is used on a multiengine airplane, the
 following must be provided:
   (1) Independent tank outlets for each engine, each incorporating a shut-off
 valve at the tank. This shutoff valve may also serve as the fire wall shutoff
 valve required if the line between the valve and the engine compartment does
 not contain more than one quart of fuel (or any greater amount shown to be
 safe) that can escape into the engine compartment.
   (2) At least two vents arranged to minimize the probability of both vents
 becoming obstructed simultaneously.
   (3) Filler caps designed to minimize the probability of incorrect
 installation or inflight loss.
   (4) A fuel system in which those parts of the system from each tank outlet
 to any engine are independent of each part of the system supplying fuel to
 any other engine.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13093 Aug. 13, 1969; Amdt. 23-43, 58 FR 18971, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.954  Fuel system lightning protection.

   The fuel system must be designed and arranged to prevent the ignition of
 fuel vapor within the system by--
   (a) Direct lightning strikes to areas having a high probability of stroke
 attachment;
   (b) Swept lightning strokes on areas where swept strokes are highly
 probable; and
   (c) Corona or streamering at fuel vent outlets.

 [Amdt. 23-7, 34 FR 13093, Aug. 13, 1969]






 Sec. 23.955  Fuel flow.

   (a) General. The ability of the fuel system to provide fuel at the rates
 specified in this section and at a pressure sufficient for proper engine
 operation must be shown in the attitude that is most critical with respect to
 fuel feed and quantity of unusable fuel. These conditions may be simulated in
 a suitable mockup. In addition--
   (1) The quantity of fuel in the tank may not exceed the amount established
 as the unusable fuel supply for that tank under Sec. 23.959 plus that
 necessary to show compliance with this section; and
   (2) If there is a fuel flowmeter, it must be blocked during the flow test
 and the fuel must flow through the meter or its bypass; and
   (3) If there is a flowmeter without a bypass, it must not have any failure
 mode that would restrict fuel flow below the level required in this fuel flow
 demonstration; and
   (4) The fuel flow must include that flow needed for vapor return flow, jet
 pump drive flow, and for all other purposes for which fuel is used.
   (b) Gravity systems. The fuel flow rate for gravity systems (main and
 reserve supply) must be 150 percent of the takeoff fuel consumption of the
 engine.
   (c) Pump systems. The fuel flow rate for each pump system (main and reserve
 supply) for each reciprocating engine must be 125 percent of the fuel flow
 required by the engine at the maximum takeoff power approved under this part.
   (1) This flow rate is required for each main pump and each emergency pump,
 and must be available when the pump is operating as it would during takeoff;
   (2) For each hand-operated pump, this rate must occur at not more than 60
 complete cycles (120 single strokes) per minute.
   (3) The fuel pressure, with main and emergency pumps operating
 simultaneously, must not exceed the fuel inlet pressure limits of the engine
 unless it can be shown that no adverse effect occurs.
   (d) Auxiliary fuel systems and fuel transfer systems. Paragraphs (b), (c),
 and (f) of this section apply to each auxiliary and transfer system, except
 that--
   (1) The required fuel flow rate must be established upon the basis of
 maximum continuous power and engine rotational speed, instead of takeoff
 power and fuel consumption; and
   (2) If there is a placard providing operating instructions, a lesser flow
 rate may be used for transferring fuel from any auxiliary tank into a larger
 main tank. This lesser flow rate must be adequate to maintain engine maximum
 continuous power but the flow rate must not overfill the main tank at lower
 engine powers.
   (e) Multiple fuel tanks. For reciprocating engines that are supplied with
 fuel from more than one tank, if engine power loss becomes apparent due to
 fuel depletion from the tank selected, it must be possible after switching to
 any full tank, in level flight, to obtain 75 percent maximum continuous power
 on that engine in not more than--
   (1) 10 seconds for naturally aspirated single-engine airplanes;
   (2) 20 seconds for turbocharged single-engine airplanes, provided that 75
 percent maximum continuous naturally aspirated power is regained within 10
 seconds; or
   (3) 20 seconds for multiengine airplanes.
   (f) Turbine engine fuel systems. Each turbine engine fuel system must
 provide at least 100 percent of the fuel flow required by the engine under
 each intended operation condition and maneuver. The conditions may be
 simulated in a suitable mockup. This flow must--
   (1) Be shown with the airplane in the most adverse fuel feed condition
 (with respect to altitudes, attitudes, and other conditions) that is expected
 in operation; and
   (2) For multiengine airplanes, notwithstanding the lower flow rate allowed
 by paragraph (d) of this section, be automatically uninterrupted with respect
 to any engine until all the fuel scheduled for use by that engine has been
 consumed. In addition--
   (i) For the purposes of this section, "fuel scheduled for use by that
 engine" means all fuel in any tank intended for use by a specific engine.
   (ii) The fuel system design must clearly indicate the engine for which fuel
 in any tank is scheduled.
   (iii) Compliance with this paragraph must require no pilot action after
 completion of the engine starting phase of operations.
   (3) For single-engine airplanes, require no pilot action after completion
 of the engine starting phase of operations unless means are provided that
 unmistakenly alert the pilot to take any needed action at least five minutes
 prior to the needed action; such pilot action must not cause any change in
 engine operation; and such pilot action must not distract pilot attention
 from essential flight duties during any phase of operations for which the
 airplane is approved.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13093, Aug. 13, 1969; Amdt. 23-43, 58 FR 18971, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.957  Flow between interconnected tanks.

   (a) It must be impossible, in a gravity feed system with interconnected
 tank outlets, for enough fuel to flow between the tanks to cause an overflow
 of fuel from any tank vent under the conditions in Sec. 23.959, except that
 full tanks must be used.
   (b) If fuel can be pumped from one tank to another in flight, the fuel tank
 vents and the fuel transfer system must be designed so that no structural
 damage to any airplane component can occur because of overfilling of any
 tank.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-43, 58 FR
 18972, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.959  Unusable fuel supply.

   The unusable fuel supply for each tank must be established as not less than
 that quantity at which the first evidence of malfunctioning occurs under the
 most adverse fuel feed condition occurring under each intended operation and
 flight maneuver involving that tank. Fuel system component failures need not
 be considered.

 [Amdt. 23-7, 34 FR 13093, Aug. 13, 1969, as amended by Amdt. 23-18, 42 FR
 15041, Mar. 17, 1977]






 Sec. 23.961   Fuel system hot weather operation.

   Each fuel system must be free from vapor lock when using fuel at its
 critical temperature, with respect to vapor formation, when operating the
 airplane in all critical operating and environmental conditions for which
 approval is requested. For turbine fuel, the initial temperature must be 110
 deg.F, -0 deg., +5 deg.F or the maximum outside air temperature for which
 approval is requested, whichever is more critical.

 [Amdt. 23-43, 58 FR 18972, Apr. 9, 1993; 58 FR 27060, May 6, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.963  Fuel tanks: general.

   (a) Each fuel tank must be able to withstand, without failure, the
 vibration, inertia, fluid, and structural loads that it may be subjected to
 in operation.
   (b) Each flexible fuel tank liner must be of an acceptable kind.
   (c) Each integral fuel tank must have adequate facilities for interior
 inspection and repair.
   (d) The total usable capacity of the fuel tanks must be enough for at least
 one-half hour of operation at maximum continuous power.
   (e) Each fuel quantity indicator must be adjusted, as specified in Sec.
 23.1337(b), to account for the unusable fuel supply determined under Sec.
 23.959.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt 23-34, 52 FR 1832, Jan. 15, 1987; Amdt. 23-43, 58 FR 18972,
 Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.965  Fuel tank tests.

   (a) Each fuel tank must be able to withstand the following pressures
 without failure or leakage:
   (1) For each conventional metal tank and nonmetallic tank with walls not
 supported by the airplane structure, a pressure of 3.5 p.s.i., or that
 pressure developed during maximum ultimate acceleration with a full tank,
 whichever is greater.
   (2) For each integral tank, the pressure developed during the maximum limit
 acceleration of the airplane with a full tank, with simultaneous application
 of the critical limit structural loads.
   (3) For each nonmetallic tank with walls supported by the airplane
 structure and constructed in an acceptable manner using acceptable basic tank
 material, and with actual or simulated support conditions, a pressure of 2
 p.s.i. for the first tank of a specific design. The supporting structure must
 be designed for the critical loads occurring in the flight or landing
 strength conditions combined with the fuel pressure loads resulting from the
 corresponding accelerations.
   (b) Each fuel tank with large, unsupported, or unstiffened flat
 surfaces,whose failure or deformation could cause fuel leakage, must be able
 to withstand the following test without leakage, failure, or excessive
 deformation of the tank walls:
   (1) Each complete tank assembly and its support must be vibration tested
 while mounted to simulate the actual installation.
   (2) Except as specified in paragraph (b)(4) of this section, the tank
 assembly must be vibrated for 25 hours at a total displacement of not less
 than 1/32  of an inch (unless another displacement is substantiated) while
 2/3  filled with water or other suitable test fluid.
   (3) The test frequency of vibration must be as follows:
   (i) If no frequency of vibration resulting from any rpm within the normal
 operating range of engine or propeller speeds is critical, the test frequency
 of vibration cycles per minute is obtained by multiplying the maximum
 continuous propeller speed in rpm by 0.9 for propeller-driven airplanes, and
 for non-propeller-driven airplanes, 2,000 cycles per minute.
   (ii) If only one frequency of vibration resulting from any rpm within the
 normal operating range of engine or propeller speeds is critical, that
 frequency of vibration must be the test frequency.
   (iii) If more than one frequency of vibration resulting from any rpm within
 the normal operating range of engine or propeller speeds is critical, the
 most critical of these frequencies must be the test frequency.
   (c) Each integral tank using methods of construction and sealing not
 previously proven to be adequate by test data or service experience must be
 able to withstand the vibration test specified in paragraphs (b) (1) through
 (4) of this section.
   (d) Each tank with a nonmetallic liner must be subjected to the sloshing
 test outlined in paragraph (b)(5) of this section, with the fuel at room
 temperature. In addition, a specimen liner of the same basic construction as
 that to be used in the airplane must, when installed in a suitable test tank,
 withstand the sloshing test with fuel at a temperature of 110 deg. F.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-43, 58 FR
 18972, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.967  Fuel tank installation.

   (a) Each fuel tank must be supported so that tank loads are not
 concentrated. In addition--
   (1) There must be pads, if necessary, to prevent chafing between each tank
 and its supports;
   (2) Padding must be nonabsorbent or treated to prevent the absorption of
 fuel;
   (3) If a flexible tank liner is used, it must be supported so that it is
 not required to withstand fluid loads;
   (4) Interior surfaces adjacent to the liner must be smooth and free from
 projections that could cause wear, unless--
   (i) Provisions are made for protection of the liner at those points; or
   (ii) The construction of the liner itself provides such protection; and
   (5) A positive pressure must be maintained within the vapor space of each
 bladder cell under any condition of operation, except for a particular
 condition for which it is shown that a zero or negative pressure will not
 cause the bladder cell to collapse; and
   (6) Syphoning of fuel (other than minor spillage) or collapse of bladder
 fuel cells may not result from improper securing or loss of the fuel filler
 cap.
   (b) Each tank compartment must be ventilated and drained to prevent the
 accumulation of flammable fluids or vapors. Each compartment adjacent to a
 tank that is an integral part of the airplane structure must also be
 ventilated and drained.
   (c) No fuel tank may be on the engine side of the firewall. There must be
 at least one-half inch of clearance between the fuel tank and the firewall.
 No part of the engine nacelle skin that lies immediately behind a major air
 opening from the engine compartment may act as the wall of an integral tank.
   (d) Each fuel tank must be isolated from personnel compartments by a fume-
 proof and fuel-proof enclosure that is vented and drained to the exterior of
 the airplane. The required enclosure must sustain any personnel compartment
 pressurization loads without permanent deformation or failure under the
 conditions of Secs. 23.365 and 23.843 of this part. A bladder-type fuel cell,
 if used, must have a retaining shell at least equivalent to a metal fuel tank
 in structural integrity.
   (e) Fuel tanks must be designed, located, and installed so as to retain
 fuel:
   (1) When subjected to the inertia loads resulting from the ultimate static
 load factors prescribed in Sec. 23.561(b)(2) of this part; and
   (2) Under conditions likely to occur when the airplane lands on a paved
 runway at a normal landing speed under each of the following conditions:
   (i) The airplane in a normal landing attitude and its landing gear
 retracted.
   (ii) The most critical landing gear leg collapsed and the other landing
 gear legs extended.

 In showing compliance with paragraph (e)(2) of this section, the tearing away
 of an engine mount must be considered unless all the engines are installed
 above the wing or on the tail or fuselage of the airplane.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13903, Aug. 13, 1969; Amdt. 23-14, 38 FR 31823, Nov. 19, 1973; Amdt. 23-18,
 42 FR 15041, Mar. 17, 1977; Amdt. 23-26, 45 FR 60171, Sept. 11, 1980; Amdt.
 23-36, 53 FR 30815, Aug. 15, 1988; Amdt. 23-43, 58 FR 18972, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.969  Fuel tank expansion space

   Each fuel tank must have an expansion space of not less than two percent of
 the tank capacity, unless the tank vent discharges clear of the airplane (in
 which case no expansion space is required). It must be impossible to fill the
 expansion space inadvertently with the airplane in the normal ground
 attitude.






 Sec. 23.971  Fuel tank sump.

   (a) Each fuel tank must have a drainable sump with an effective capacity,
 in the normal ground and flight attitudes, of 0.25 percent of the tank
 capacity, or 1/16  gallon, whichever is greater.
   (b) Each fuel tank must allow drainage of any hazardous quantity of water
 from any part of the tank to its sump with the airplane in the normal ground
 attitude.
   (c) Each reciprocating engine fuel system must have a sediment bowl or
 chamber that is accessible for drainage; has a capacity of 1 ounce for every
 20 gallons of fuel tank capacity; and each fuel tank outlet is located so
 that, in the normal flight attitude, water will drain from all parts of the
 tank except the sump to the sediment bowl or chamber.
   (d) Each sump, sediment bowl, and sediment chamber drain required by
 paragraphs (a), (b), and (c) of this section must comply with the drain
 provisions of Sec. 23.999 (b)(1) and (b)(2).

 [Amdt. 23-43, 58 FR 18972, Apr. 9, 1993; 58 FR 27060, May 6, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.973  Fuel tank filler connection.

   (a) Each fuel tank filler connection must be marked as prescribed in Sec.
 23.1557(c).
   (b) Spilled fuel must be prevented from entering the fuel tank compartment
 or any part of the airplane other than the tank itself.
   (c) Each filler cap must provide a fuel-tight seal for the main filler
 opening. However, there may be small openings in the fuel tank cap for
 venting purposes or for the purpose of allowing passage of a fuel gauge
 through the cap provided such openings comply with the requirements of Sec.
 23.975(a).
   (d) Each fuel filling point, except pressure fueling connection points,
 must have a provision for electrically bonding the airplane to ground fueling
 equipment.
   (e) For airplanes with engines requiring gasoline as the only permissible
 fuel, the inside diameter of the fuel filler opening must be no larger than
 2.36 inches.
   (f) For airplanes with turbine engines, and not equipped with pressure
 fueling provisions, the inside diameter of the fuel filler opening must be no
 smaller than 2.95 inches.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. 23-18, 42 FR 15041, Mar. 17, 1977; Amdt. 23-43, 58 FR 18972,
 Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.975  Fuel tank vents and carburetor vapor vents.

   (a) Each fuel tank must be vented from the top part of the expansion space.
 In addition--
   (1) Each vent outlet must be located and constructed in a manner that
 minimizes the possibility of its being obstructed by ice or other foreign
 matter;
   (2) Each vent must be constructed to prevent siphoning of fuel during
 normal operation;
   (3) The venting capacity must allow the rapid relief of excessive
 differences of pressure between the interior and exterior of the tank;
   (4) Airspaces of tanks with interconnected outlets must be interconnected;
   (5) There may be no undrainable points in any vent line where moisture can
 accumulate with the airplane in either the ground or level flight attitudes.
 Any drain valves installed in the vent lines must discharge clear of the
 airplane and be accessible for drainage;
   (6) No vent may terminate at a point where the discharge of fuel from the
 vent outlet will constitute a fire hazard or from which fumes may enter
 personnel compartments; and
   (7) Vents must be arranged to prevent the loss of fuel, except fuel
 discharged because of thermal expansion, when the airplane is parked in any
 direction on a ramp having a one-percent slope.
   (b) Each carburetor with vapor elimination connections and each fuel
 injection engine employing vapor return provisions must have a separate vent
 line to lead vapors back to the top of one of the fuel tanks. If there is
 more than one tank and it is necessary to use these tanks in a definite
 sequence for any reason, the vapor vent line must lead back to the fuel tank
 to be used first, unless the relative capacities of the tanks are such that
 return to another tank is preferable.
   (c) For acrobatic category airplanes, excessive loss of fuel during
 acrobatic maneuvers, including short periods of inverted flight, must be
 prevented. It must be impossible for fuel to siphon from the vent when normal
 flight has been resumed after any acrobatic maneuver for which certification
 is requested.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. 23-18, 42 FR 15041, Mar. 17, 1977; Amdt. 23-29, 49 FR 6847,
 Feb. 23, 1984; Amdt. 23-43, 58 FR 18973, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.977  Fuel tank outlet.

   (a) There must be a fuel strainer for the fuel tank outlet or for the
 booster pump. This strainer must--
   (1) For reciprocating engine powered airplanes, have 8 to 16 meshes per
 inch; and
   (2) For turbine engine powered airplanes, prevent the passage of any object
 that could restrict fuel flow or damage any fuel system component.
   (b) The clear area of each fuel tank outlet strainer must be at least five
 times the area of the outlet line.
   (c) The diameter of each strainer must be at least that of the fuel tank
 outlet.
   (d) Each strainer must be accessible for inspection and cleaning.

 [Amdt. 23-17, 41 FR 55465, Dec. 20, 1976, as amended by Amdt. 23-43, 58 FR
 18973, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.979  Pressure fueling systems.

   For pressure fueling systems, the following apply:
   (a) Each pressure fueling system fuel manifold connection must have means
 to prevent the escape of hazardous quantities of fuel from the system if the
 fuel entry valve fails.
   (b) An automatic shutoff means must be provided to prevent the quantity of
 fuel in each tank from exceeding the maximum quantity approved for that tank.
 This means must allow checking for proper shutoff operation before each
 fueling of the tank.
   (c) A means must be provided to prevent damage to the fuel system in the
 event of failure of the automatic shutoff means prescribed in paragraph (b)
 of this section.
   (d) All parts of the fuel system up to the tank which are subjected to
 fueling pressures must have a proof pressure of 1.33 times, and an ultimate
 pressure of at least 2.0 times, the surge pressure likely to occur during
 fueling.

 [Amdt. 23-14, 38 FR 31823, Nov. 19, 1973]






                            Fuel System Components






 Sec. 23.991  Fuel pumps.

   (a) Main pumps. For main pumps, the following apply:
   (1) For reciprocating engine installations having fuel pumps to supply fuel
 to the engine, at least one pump for each engine must be directly driven by
 the engine and must meet Sec. 23.955. This pump is a main pump.
   (2) For turbine engine installations, each fuel pump required for proper
 engine operation, or required to meet the fuel system requirements of this
 subpart (other than those in paragraph (b) of this section), is a main pump.
 In addition--
   (i) There must be at least one main pump for each turbine engine;
   (ii) The power supply for the main pump for each engine must be independent
 of the power supply for each main pump for any other engine; and
   (iii) For each main pump, provision must be made to allow the bypass of
 each positive displacement fuel pump other than a fuel injection pump
 approved as part of the engine.
   (b) Emergency pumps. There must be an emergency pump immediately available
 to supply fuel to the engine if any main pump (other than a fuel injection
 pump approved as part of an engine) fails. The power supply for each
 emergency pump must be independent of the power supply for each corresponding
 main pump.
   (c) Warning means. If both the main pump and emergency pump operate
 continuously, there must be a means to indicate to the appropriate flight
 crewmembers a malfunction of either pump.
   (d) Operation of any fuel pump may not affect engine operation so as to
 create a hazard, regardless of the engine power or thrust setting or the
 functional status of any other fuel pump.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13093, Aug. 13, 1969; Amdt. 23-26, 45 FR 60171, Sept. 11, 1980; Amdt. 23-43,
 58 FR 18973, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.993  Fuel system lines and fittings.

   (a) Each fuel line must be installed and supported to prevent excessive
 vibration and to withstand loads due to fuel pressure and accelerated flight
 conditions.
   (b) Each fuel line connected to components of the airplane between which
 relative motion could exist must have provisions for flexibility.
   (c) Each flexible connection in fuel lines that may be under pressure and
 subjected to axial loading must use flexible hose assemblies.
   (d) Each flexible hose must be shown to be suitable for the particular
 application.
   (e) No flexible hose that might be adversely affected by exposure to high
 temperatures may be used where excessive temperatures will exist during
 operation or after engine shutdown.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-43, 58 FR
 18973, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.994   Fuel system components.

   Fuel system components in an engine nacelle or in the fuselage must be
 protected from damage which could result in spillage of enough fuel to
 constitute a fire hazard as a result of a wheels-up landing on a paved
 runway.

 [Amdt. 23-29, 49 FR 6847, Feb. 23, 1984]






 Sec. 23.995  Fuel valves and controls.

   (a) There must be a means to allow appropriate flight crew members to
 rapidly shut off, in flight, the fuel to each engine individually.
   (b) No shutoff valve may be on the engine side of any firewall. In
 addition, there must be means to--
   (1) Guard against inadvertent operation of each shutoff valve; and
   (2) Allow appropriate flight crew members to reopen each valve rapidly
 after it has been closed.
   (c) Each valve and fuel system control must be supported so that loads
 resulting from its operation or from accelerated flight conditions are not
 transmitted to the lines connected to the valve.
   (d) Each valve and fuel system control must be installed so that gravity
 and vibration will not affect the selected position.
   (e) Each fuel valve handle and its connections to the valve mechanism must
 have design features that minimize the possibility of incorrect installation.
   (f) Each check valve must be constructed, or otherwise incorporate
 provisions, to preclude incorrect assembly or connection of the valve.
   (g) Fuel tank selector valves must--
   (1) Require a separate and distinct action to place the selector in the
 "OFF" position; and
   (2) Have the tank selector positions located in such a manner that it is
 impossible for the selector to pass through the "OFF" position when changing
 from one tank to another.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-14, 38 FR
 31823, Nov. 19, 1973; Amdt. 23-17, 41 FR 55465, Dec. 20, 1976; Amdt. 23-18,
 42 FR 15041, Mar. 17, 1977; Amdt. 23-29, 49 FR 6847, Feb. 23, 1984]






 Sec. 23.997  Fuel strainer or filter.

   There must be a fuel strainer or filter between the fuel tank outlet and
 the inlet of either the fuel metering device or an engine driven positive
 displacement pump, whichever is nearer the fuel tank outlet. This fuel
 strainer or filter must--
   (a) Be accessible for draining and cleaning and must incorporate a screen
 or element which is easily removable;
   (b) Have a sediment trap and drain except that it need not have a drain if
 the strainer or filter is easily removable for drain purposes;
   (c) Be mounted so that its weight is not supported by the connecting lines
 or by the inlet or outlet connections of the strainer or filter itself,
 unless adequate strength margins under all loading conditions are provided in
 the lines and connections; and
   (d) Have the capacity (with respect to operating limitations established
 for the engine) to ensure that engine fuel system functioning is not
 impaired, with the fuel contaminated to a degree (with respect to particle
 size and density) that is greater than that established for the engine during
 its type certification.
   (e) In addition, for commuter category airplanes, unless means are provided
 in the fuel system to prevent the accumulation of ice on the filter, a means
 must be provided to automatically maintain the fuel flow if ice clogging of
 the filter occurs.

 [Amdt. 23-15, 39 FR 35459, Oct. 1, 1974, as amended by Amdt. 23-29, 49 FR
 6847, Feb. 23, 1984; Amdt. 23-34, 52 FR 1832, Jan. 15, 1987; Amdt. 23-43, 58
 FR 18973, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.999  Fuel system drains.

   (a) There must be at least one drain to allow safe drainage of the entire
 fuel system with the airplane in its normal ground attitude.
   (b) Each drain required by paragraph (a) of this section and Sec. 23.971
 must--
   (1) Discharge clear of all parts of the airplane;
   (2) Have a drain valve--
   (i) That has manual or automatic means for positive locking in the closed
 position;
   (ii) That is readily accessible;
   (iii) That can be easily opened and closed;
   (iv) That allows the fuel to be caught for examination;
   (v) That can be observed for proper closing; and
   (vi) That is either located or protected to prevent fuel spillage in the
 event of a landing with landing gear retracted.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-17, 41 FR
 55465, Dec. 20, 1976; Amdt. 23-43, 58 FR 18973, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1001  Fuel jettisoning system.

   (a) If the design landing weight is less than that permitted under the
 requirements of Sec. 23.473(b), the airplane must have a fuel jettisoning
 system installed that is able to jettison enough fuel to bring the maximum
 weight down to the design landing weight. The average rate of fuel
 jettisoning must be at least 1 percent of the maximum weight per minute,
 except that the time required to jettison the fuel need not be less than 10
 minutes.
   (b) Fuel jettisoning must be demonstrated at maximum weight with flaps and
 landing gear up and in--
   (1) A power-off glide at 1.4 VS1;
   (2) A climb at the one-engine-inoperative best rate-of-climb speed, with
 the critical engine inoperative and the remaining engines at maximum
 continuous power; and
   (3) Level flight at 1.4 VS1, if the results of the tests in the conditions
 specified in paragraphs (b)(1) and (2) of this section show that this
 condition could be critical.
   (c) During the flight tests prescribed in paragraph (b) of this section, it
 must be shown that--
   (1) The fuel jettisoning system and its operation are free from fire
 hazard;
   (2) The fuel discharges clear of any part of the airplane;
   (3) Fuel or fumes do not enter any parts of the airplane; and
   (4) The jettisoning operation does not adversely affect the controllability
 of the airplane.
   (d) For reciprocating engine powered airplanes, the jettisoning system must
 be designed so that it is not possible to jettison the fuel in the tanks used
 for takeoff and landing below the level allowing 45 minutes flight at 75
 percent maximum continuous power. However, if there is an auxiliary control
 independent of the main jettisoning control, the system may be designed to
 jettison all the fuel.
   (e) For turbine engine powered airplanes, the jettisoning system must be
 designed so that it is not possible to jettison fuel in the tanks used for
 takeoff and landing below the level allowing climb from sea level to 10,000
 feet and thereafter allowing 45 minutes cruise at a speed for maximum range.
   (f) The fuel jettisoning valve must be designed to allow flight crewmembers
 to close the valve during any part of the jettisoning operation.
   (g) Unless it is shown that using any means (including flaps, slots, and
 slats) for changing the airflow across or around the wings does not adversely
 affect fuel jettisoning, there must be a placard, adjacent to the jettisoning
 control, to warn flight crewmembers against jettisoning fuel while the means
 that change the airflow are being used.
   (h) The fuel jettisoning system must be designed so that any reasonably
 probable single malfunction in the system will not result in a hazardous
 condition due to unsymmetrical jettisoning of, or inability to jettison,
 fuel.

 [Amdt. 23-7, 34 FR 13094, Aug. 13, 1969, as amended by Amdt. 23-43, 58 FR
 18973, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






                                  Oil System






 Sec. 23.1011  General.

   (a) For oil systems and components that have been approved under the engine
 airworthiness requirements and where those requirements are equal to or more
 severe than the corresponding requirements of subpart E of this part, that
 approval need not be duplicated. Where the requirements of subpart E of this
 part are more severe, substantiation must be shown to the requirements of
 subpart E of this part.
   (b) Each engine must have an independent oil system that can supply it with
 an appropriate quantity of oil at a temperature not above that safe for
 continuous operation.
   (c) The usable oil tank capacity may not be less than the product of the
 endurance of the airplane under critical operating conditions and the maximum
 oil consumption of the engine under the same conditions, plus a suitable
 margin to ensure adequate circulation and cooling.
   (d) For an oil system without an oil transfer system, only the usable oil
 tank capacity may be considered. The amount of oil in the engine oil lines,
 the oil radiator, and the feathering reserve, may not be considered.
   (e) If an oil transfer system is used, and the transfer pump can pump some
 of the oil in the transfer lines into the main engine oil tanks, the amount
 of oil in these lines that can be pumped by the transfer pump may be included
 in the oil capacity.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-43, 58 FR
 18973, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1013  Oil tanks.

   (a) Installation. Each oil tank must be installed to--
   (1) Meet the requirements of Sec. 23.967 (a) and (b); and
   (2) Withstand any vibration, inertia, and fluid loads expected in
 operation.
   (b) Expansion space. Oil tank expansion space must be provided so that--
   (1) Each oil tank used with a reciprocating engine has an expansion space
 of not less than the greater of 10 percent of the tank capacity or 0.5
 gallon, and each oil tank used with a turbine engine has an expansion space
 of not less than 10 percent of the tank capacity; and
   (2) It is impossible to fill the expansion space inadvertently with the
 airplane in the normal ground attitude.
   (c) Filler connection. Each oil tank filler connection must be marked as
 specified in Sec. 23.1557(c). Each recessed oil tank filler connection of an
 oil tank used with a turbine engine, that can retain any appreciable quantity
 of oil, must have provisions for fitting a drain.
   (d) Vent. Oil tanks must be vented as follows:
   (1) Each oil tank must be vented to the engine crankcase from the top part
 of the expansion space so that the vent connection is not covered by oil
 under any normal flight condition.
   (2) Oil tank vents must be arranged so that condensed water vapor that
 might freeze and obstruct the line cannot accumulate at any point.
   (3) For acrobatic category airplanes, there must be means to prevent
 hazardous loss of oil during acrobatic maneuvers, including short periods of
 inverted flight.
   (e) Outlet. No oil tank outlet may be enclosed by any screen or guard that
 would reduce the flow of oil below a safe value at any operating temperature.
 No oil tank outlet diameter may be less than the diameter of the engine oil
 pump inlet. Each oil tank used with a turbine engine must have means to
 prevent entrance into the tank itself, or into the tank outlet, of any object
 that might obstruct the flow of oil through the system. There must be a
 shutoff valve at the outlet of each oil tank used with a turbine engine,
 unless the external portion of the oil system (including oil tank supports)
 is fireproof.
   (f) Flexible liners. Each flexible oil tank liner must be of an acceptable
 kind.
   (g) Each oil tank filler cap of an oil tank that is used with an engine
 must provide an oiltight seal.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-15, 39 FR
 35459 Oct. 1, 1974; Amdt. 23-43, 58 FR 18973, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1015  Oil tank tests.

   Each oil tank must be tested under Sec. 23.965, except that--
   (a) The applied pressure must be five p.s.i. for the tank construction
 instead of the pressures specified in Sec. 23.965(a);
   (b) For a tank with a nonmetallic liner the test fluid must be oil rather
 than fuel as specified in Sec. 23.965(d), and the slosh test on a specimen
 liner must be conducted with the oil at 250 deg. F.; and
   (c) For pressurized tanks used with a turbine engine, the test pressure may
 not be less than 5 p.s.i. plus the maximum operating pressure of the tank.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-15, 39 FR
 35460, Oct. 1, 1974]






 Sec. 23.1017  Oil lines and fittings.

   (a) Oil lines. Oil lines must meet Sec. 23.993 and must accommodate a flow
 of oil at a rate and pressure adequate for proper engine functioning under
 any normal operating condition.
   (b) Breather lines. Breather lines must be arranged so that--
   (1) Condensed water vapor or oil that might freeze and obstruct the line
 cannot accumulate at any point;
   (2) The breather discharge will not constitute a fire hazard if foaming
 occurs, or cause emitted oil to strike the pilot's windshield;
   (3) The breather does not discharge into the engine air induction system;
 and
   (4) For acrobatic category airplanes, there is no excessive loss of oil
 from the breather during acrobatic maneuvers, including short periods of
 inverted flight.
   (5) The breather outlet is protected against blockage by ice or foreign
 matter.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13094, Aug. 13, 1969; Amdt. 23-14, 38 FR 31823, Nov. 19, 1973]






 Sec. 23.1019  Oil strainer or filter.

   (a) Each turbine engine installation must incorporate an oil strainer or
 filter through which all of the engine oil flows and which meets the
 following requirements:
   (1) Each oil strainer or filter that has a bypass, must be constructed and
 installed so that oil will flow at the normal rate through the rest of the
 system with the strainer or filter completely blocked.
   (2) The oil strainer or filter must have the capacity (with respect to
 operating limitations established for the engine) to ensure that engine oil
 system functioning is not impaired when the oil is contaminated to a degree
 (with respect to particle size and density) that is greater than that
 established for the engine for its type certification.
   (3) The oil strainer or filter, unless it is installed at an oil tank
 outlet, must incorporate a means to indicate contamination before it reaches
 the capacity established in accordance with paragraph (a)(2) of this section.
   (4) The bypass of a strainer or filter must be constructed and installed so
 that the release of collected contaminants is minimized by appropriate
 location of the bypass to ensure that collected contaminants are not in the
 bypass flow path.
   (5) An oil strainer or filter that has no bypass, except one that is
 installed at an oil tank outlet, must have a means to connect it to the
 warning system required in Sec. 23.1305(c)(9).
   (b) Each oil strainer or filter in a powerplant installation using
 reciprocating engines must be constructed and installed so that oil will flow
 at the normal rate through the rest of the system with the strainer or filter
 element completely blocked.

 [Amdt. 23-15, 39 FR 35460, Oct. 1, 1974, as amended by Amdt. 23-29, 49 FR
 6847, Feb. 23, 1984; Amdt. 23-43, 58 FR 18973, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1021   Oil system drains.

   A drain (or drains) must be provided to allow safe drainage of the oil
 system. Each drain must--
   (a) Be accessible;
   (b) Have drain valves, or other closures, employing manual or automatic
 shut-off means for positive locking in the closed position; and
   (c) Be located or protected to prevent inadvertent operation.

 [Amdt. 23-29, 49 FR 6847, Feb. 23, 1984, as amended by Amdt. 23-43, 58 FR
 18973, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1023  Oil radiators.

   Each oil radiator and its supporting structures must be able to withstand
 the vibration, inertia, and oil pressure loads to which it would be subjected
 in operation.






 Sec. 23.1027  Propeller feathering system.

   (a) If the propeller feathering system uses engine oil and that oil supply
 can become depleted due to failure of any part of the oil system, a means
 must be incorporated to reserve enough oil to operate the feathering system.
   (b) The amount of reserved oil must be enough to accomplish feathering and
 must be available only to the feathering pump.
   (c) The ability of the system to accomplish feathering with the reserved
 oil must be shown.
   (d) Provision must be made to prevent sludge or other foreign matter from
 affecting the safe operation of the propeller feathering system.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-14, 38 FR
 31823, Nov. 19, 1973; Amdt. 23-43, 58 FR 18973, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






                                    Cooling






 Sec. 23.1041  General.

   The powerplant and auxiliary power unit cooling provisions must maintain
 the temperatures of powerplant components and engine fluids, and auxiliary
 power unit components and fluids within the limits established for those
 components and fluids under the most adverse ground, water, and flight
 operations to the maximum altitude for which approval is requested, and after
 normal engine and auxiliary power unit shutdown.

 [Amdt. 23-43, 58 FR 18973, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1043  Cooling tests.

   (a) General. Compliance with Sec. 23.1041 must be shown under critical
 ground, water, and flight operating conditions to the maximum altitude for
 which approval is requested. For turbosupercharged engines, each
 turbosupercharger must be operated through that part of the climb profile for
 which operation with the turbosupercharger is requested and in a manner
 consistent with its intended operation. For these tests, the following apply:
   (1) If the tests are conducted under conditions deviating from the maximum
 ambient atmospheric temperatures specified in paragraph (b) of this section,
 the recorded powerplant temperatures must be corrected under paragraphs (c)
 and (d) of this section, unless a more rational correction method is
 applicable.
   (2) No corrected temperature determined under paragraph (a)(1) of this
 section may exceed established limits.
   (3) The fuel uses during the cooling tests must be of the minimum grade
 approved for the engines, and the mixture settings must be those used in
 normal operation.
   (4) [Reserved]
   (5) Water taxing tests must be conducted on each hull seaplane that may
 reasonably be expected to be taxied for extended periods.
   (b) Maximum ambient atmospheric temperature. A maximum ambient atmospheric
 temperature corresponding to sea level conditions of at least 100 degrees F
 must be established. The assumed temperature lapse rate is 3.6 degrees F per
 thousand feet of altitude above sea level until a temperature of -69.7
 degrees F is reached, above which altitude the temperature is considered
 constant at -69.7 degrees F. However, for winterization installations, the
 applicant may select a maximum ambient atmospheric temperature corresponding
 to sea level conditions of less than 100 degrees F.
   (c) Correction factor (except cylinder barrels). Unless a more rational
 correction applies, temperatures of engine fluids and powerplant components
 (except cylinder barrels) for which temperature limits are established, must
 be corrected by adding to them the difference between the maximum ambient
 atmospheric temperature and the temperature of the ambient air at the time of
 the first occurrence of the maximum component or fluid temperature recorded
 during the cooling test.
   (d) Correction factor for cylinder barrel temperatures. Cylinder barrel
 temperatures must be corrected by adding to them 0.7 times the difference
 between the maximum ambient atmospheric temperature and the temperature of
 the ambient air at the time of the first occurrence of the maximum cylinder
 barrel temperature recorded during the cooling test.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13094, Aug. 13, 1969; Amdt. 23-21, 43 FR 2319, Jan. 16, 1978]






 Sec. 23.1045  Cooling test procedures for turbine engine powered airplanes.

   (a) Compliance with Sec. 23.1041 must be shown for the takeoff, climb, en
 route, and landing stages of flight that correspond to the applicable
 performance requirements. The cooling tests must be conducted with the
 airplane in the configuration, and operating under the conditions, that are
 critical relative to cooling during each stage of flight. For the cooling
 tests, a temperature is "stabilized" when its rate of change is less than 2
 deg. F. per minute.
   (b) Temperatures must be stabilized under the conditions from which entry
 is made into each stage of flight being investigated, unless the entry
 condition normally is not one during which component and engine fluid
 temperatures would stabilize (in which case, operation through the full entry
 condition must be conducted before entry into the stage of flight being
 investigated in order to allow temperatures to reach their natural levels at
 the time of entry). The takeoff cooling test must be preceded by a period
 during which the powerplant component and engine fluid temperatures are
 stabilized with the engines at ground idle.
   (c) Cooling tests for each stage of flight must be continued until--
   (1) The component and engine fluid temperatures stabilize;
   (2) The stage of flight is completed; or
   (3) An operating limitation is reached.

 [Amdt. 23-7, 34 FR 13094, Aug. 13, 1969]






 Sec. 23.1047  Cooling test procedures for reciprocating engine-powered
     airplanes.

   (a) For each single-engine airplane powered with a reciprocating engine,
 engine cooling tests must be conducted as follows:
   (1) Engine temperatures must be stabilized in flight with the engines at
 not less than 75 percent of maximum continuous power.
   (2) After temperatures have stabilized, a climb must be begun at the lowest
 pra@icable altitude and continued for 1 minute with the engine at takeoff
 power.
   (3) At the end of 1 minute, the climb must be continued at maximum
 continuous power for at least 5 minutes after the occurrence of the highest
 temperature recorded.
   (b) The climb required in paragraph (a) of this section must be conducted
 at a speed not more than the best rate-of-climb speed with maximum continuous
 power unless--
   (1) The slope of the flight path at the speed chosen for the cooling test
 is equal to or greater than the minimum required angle of climb determined
 under Sec. 23.65; and
   (2) The airplane has a cylinder head temperature indicator as specified in
 Sec. 23.1305(b)(3).
   (c) The stabilizing and climb parts of the test must be conducted with cowl
 flap settings selected by the applicant.
   (d) For each multiengine airplane powered with reciprocating engines, that
 meets the minimum one-engine-inoperative climb performance specified in
 Sec. 23.67(b)(1), engine cooling tests must be conducted as follows:
   (1) The airplane must be in the configuration specified in Sec. 23.67(a),
 except that, when above the critical altitude, the operating engines must be
 at maximum continuous power or at full throttle.
   (2) The stabilizing and climb parts of the tests must be conducted with
 cowl flap settings selected by the applicant.
   (3) The temperatures of the operating engines must be stabilized in flight,
 with the engines at not less than 75 percent of the maximum continuous power.
   (4) After engine temperatures have stabilized, a climb must be--
   (i) Begun from 1,000 feet below the critical altitude (or, if this is
 impracticable, at the lowest altitude that the terrain will allow) or 1,000
 feet below the altitude at which the single-engine-inoperative rate of climb
 is 0.02 Vso 2  whichever is lower; and
   (ii) Continued for at least 5 minutes after the highest temperature has
 been recorded.
   (5) The climb must be conducted at a speed not more than the highest speed
 at which compliance with the climb requirement of Sec. 23.67(b)(1) can be
 shown. If the speed used exceeds the speed for best rate of climb with one
 engine inoperative, the airplane must have a cylinder head temperature
 indicator as specified in Sec. 23.1337(e).
   (e) For each multiengine airplane powered with reciprocating engines that
 cannot meet the minimum one-engine-inoperative climb performance specified in
 Sec. 23.67(b)(1), engine cooling tests must be conducted as prescribed in
 paragraph (d) of this section, except that, after stabilizing temperatures
 in flight, the climb (or descent, for airplanes with zero or negative
 one-engine-inoperative rates of climb) must be--
   (1) Begun as close to sea level as is practicable; and
   (2) Conducted at the best rate-of-climb speed (or the speed of minimum rate
 of descent, for airplanes with zero or negative one-engine-inoperative rates
 of climb).

 [Amdt. 23-7, 34 FR 13094, Aug. 13, 1969, as amended by Amdt. 23-21, 43 FR
 2319, Jan. 16, 1978; Amdt. 23-42, 56 FR 354, Jan. 3, 1991; Amdt. 23-43, 58 FR
 18973, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






                                Liquid Cooling






 Sec. 23.1061  Installation.

   (a) General. Each liquid-cooled engine must have an independent cooling
 system (including coolant tank) installed so that--
   (1) Each coolant tank is supported so that tank loads are distributed over
 a large part of the tank surface;
   (2) There are pads or other isolation means between the tank and its
 supports to prevent chafing.
   (3) Pads or any other isolation means that is used must be nonabsorbent or
 must be treated to prevent absorption of flammable fluids; and
   (4) No air or vapor can be trapped in any part of the system, except the
 coolant tank expansion space, during filling or during operation.
   (b) Coolant tank. The tank capacity must be at least one gallon, plus 10
 percent of the cooling system capacity. In addition--
   (1) Each coolant tank must be able to withstand the vibration, inertia, and
 fluid loads to which it may be subjected in operation;
   (2) Each coolant tank must have an expansion space of at least 10 percent
 of the total cooling system capacity; and
   (3) It must be impossible to fill the expansion space inadvertently with
 the airplane in the normal ground attitude.
   (c) Filler connection. Each coolant tank filler connection must be marked
 as specified in Sec. 23.1557(c). In addition--
   (1) Spilled coolant must be prevented from entering the coolant tank
 compartment or any part of the airplane other than the tank itself; and
   (2) Each recessed coolant filler connection must have a drain that
 discharges clear of the entire airplane.
   (d) Lines and fittings. Each coolant system line and fitting must meet the
 requirements of Sec. 23.993, except that the inside diameter of the engine
 coolant inlet and outlet lines may not be less than the diameter of the
 corresponding engine inlet and outlet connections.
   (e) Radiators. Each coolant radiator must be able to withstand any
 vibration, inertia, and coolant pressure load to which it may normally be
 subjected. In addition--
   (1) Each radiator must be supported to allow expansion due to operating
 temperatures and prevent the transmittal of harmful vibration to the
 radiator; and
   (2) If flammable coolant is used, the air intake duct to the coolant
 radiator must be located so that (in case of fire) flames from the nacelle
 cannot strike the radiator.
   (f) Drains. There must be an accessible drain that--
   (1) Drains the entire cooling system (including the coolant tank, radiator,
 and the engine) when the airplane is in the normal ground altitude;
   (2) Discharges clear of the entire airplane; and
   (3) Has means to positively lock it closed.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-43, 58 FR
 18973, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1063  Coolant tank tests.

   Each coolant tank must be tested under Sec. 23.965, except that--
   (a) The test required by Sec. 23.965(a) (1) must be replaced with a similar
 test using the sum of the pressure developed during the maximum ultimate
 acceleration with a full tank or a pressure of 3.5 pounds per square inch,
 whichever is greater, plus the maximum working pressure of the system; and
   (b) For a tank with a nonmetallic liner the test fluid must be coolant
 rather than fuel as specified in Sec. 23.965(d), and the slosh test on a
 specimen liner must be conducted with the coolant at operating temperature.






                               Induction System






 Sec. 23.1091  Air induction system.

   (a) The air induction system for each engine and auxiliary power unit and
 their accessories must supply the air required by that engine and auxiliary
 power unit and their accessories under the operating conditions for which
 certification is requested.
   (b) Each reciprocating engine installation must have at least two separate
 air intake sources and must meet the following:
   (1) Primary air intakes may open within the cowling if that part of the
 cowling is isolated from the engine accessory section by a fire-resistant
 diaphragm or if there are means to prevent the emergence of backfire flames.
   (2) Each alternate air intake must be located in a sheltered position and
 may not open within the cowling if the emergence of backfire flames will
 result in a hazard.
   (3) The supplying of air to the engine through the alternate air intake
 system may not result in a loss of excessive power in addition to the power
 loss due to the rise in air temperature.
   (4) Each automatic alternate air door must have an override means
 accessible to the flight crew.
   (5) Each automatic alternate air door must have a means to indicate to the
 flight crew when it is not closed.
   (c) For turbine engine powered airplanes--
   (1) There must be means to prevent hazardous quantities of fuel leakage or
 overflow from drains, vents, or other components of flammable fluid systems
 from entering the engine or auxiliary power unit and their accessories intake
 system; and
   (2) The airplane must be designed to prevent water, slush or other foreign
 material on the runway, taxiway, or other airport operating surface from
 being directed into the engine or auxiliary power unit air inlet ducts in
 hazardous quantities during takeoff, landing, and taxiing.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13095, Aug. 13, 1969; Amdt. 23-43, 58 FR 18973, Apr. 9, 1993; 58 FR 27060,
 May 6, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1093  Induction system icing protection.

   (a) Reciprocating engines. Each reciprocating engine air induction system
 must have means to prevent and eliminate icing. Unless this is done by other
 means, it must be shown that, in air free of visible moisture at a
 temperature of 30 deg. F.--
   (1) Each airplane with sea level engines using conventional venturi
 carburetors has a preheater that can provide a heat rise of 90 deg. F. with
 the engines at 75 percent of maximum continuous power;
   (2) Each airplane with altitude engines using conventional venturi
 carburetors has a preheater that can provide a heat rise of 120 deg. F. with
 the engines at 75 percent of maximum continuous power;
   (3) Each airplane with altitude engines using fuel metering devices
 tending to prevent icing has a preheater that, with the engines at 60 percent
 of maximum continuous power, can provide a heat rise of--
   (i) 100 deg. F.; or
   (ii) 40 deg. F., if a fluid deicing system meeting the requirements of
 Secs. 23.1095 through 23.1099 is installed;
   (4) Each airplane with sea level engine(s) using fuel metering device
 tending to prevent icing has a sheltered alternate source of air with a
 preheat of not less than 60 deg.F with the engines at 75 percent of maximum
 continuous power;
   (5) Each airplane with sea level or altitude engine(s) using fuel injection
 systems having metering components on which impact ice may accumulate has a
 preheater capable of providing a heat rise of 75 deg.F when the engine is
 operating at 75 percent of its maximum continuous power; and
   (6) Each airplane with sea level or altitude engine(s) using fuel injection
 systems not having fuel metering components projecting into the airstream on
 which ice may form, and introducing fuel into the air induction system
 downstream of any components or other obstruction on which ice produced by
 fuel evaporation may form, has a sheltered alternate source of air with a
 preheat of not less than 60 deg.F with the engines at 75 percent of its
 maximum continuous power.
   (b) Turbine engines.
   (1) Each turbine engine and its air inlet system must operate throughout
 the flight power range of the engine (including idling), without the
 accumulation of ice on engine or inlet system components that would adversely
 affect engine operation or cause a serious loss of power or thrust--
   (i) Under the icing conditions specified in appendix C of part 25 of this
 chapter; and
   (ii) In snow, both falling and blowing, within the limitations established
 for the airplane for such operation.
   (2) Each turbine engine must idle for 30 minutes on the ground, with the
 air bleed available for engine icing protection at its critical condition,
 without adverse effect, in an atmosphere that is at a temperature between 15
 deg. and 30 deg.F (between -9 deg. and -1 deg.C) and has a liquid water
 content not less than 0.3 grams per cubic meter in the form of drops having a
 mean effective diameter not less than 20 microns, followed by momentary
 operation at takeoff power or thrust. During the 30 minutes of idle
 operation, the engine may be run up periodically to a moderate power or
 thrust setting in a manner acceptable to the Administrator.
   (c) For airplanes with reciprocating engines having superchargers to
 pressurize the air before it enters the fuel metering device, the heat rise
 in the air caused by that supercharging at any altitude may be utilized in
 determining compliance with paragraph (a) of this section if the heat rise
 utilized is that which will be available, automatically, for the applicable
 altitudes and operating condition because of supercharging.

 [Amdt. 23-7, 34 FR 13095, Aug. 13, 1969, as amended by Amdt. 23-15, 39 FR
 35460, Oct. 1, 1974; Amdt. 23-17, 41 FR 55465, Dec. 20, 1976; Amdt. 23-18, 42
 FR 15041, Mar. 17, 1977; Amdt. 23-29, 49 FR 6847, Feb. 23, 1984; Amdt. 23-43,
 58 FR 18974, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1095  Carburetor deicing fluid flow rate.

   (a) If a carburetor deicing fluid system is used, it must be able to
 simultaneously supply each engine with a rate of fluid flow, expressed in
 pounds per hour, of not less than 2.5 times the square root of the maximum
 continuous power of the engine.
   (b) The fluid must be introduced into the air induction system--
   (1) Close to, and upstream of, the carburetor; and
   (2) So that it is equally distributed over the entire cross section of the
 induction system air passages.






 Sec. 23.1097  Carburetor deicing fluid system capacity.

   (a) The capacity of each carburetor deicing fluid system--
   (1) May not be less than the greater of--
   (i) That required to provide fluid at the rate specified in Sec. 23.1095
 for a time equal to three percent of the maximum endurance of the airplane;
 or
   (ii) 20 minutes at that flow rate; and
   (2) Need not exceed that required for two hours of operation.
   (b) If the available preheat exceeds 50 deg. F. but is less than 100 deg.
 F., the capacity of the system may be decreased in proportion to the heat
 rise available in excess of 50 deg. F.






 Sec. 23.1099  Carburetor deicing fluid system detail design.

   Each carburetor deicing fluid system must meet the applicable requirements
 for the design of a fuel system, except as specified in Secs. 23.1095 and
 23.1097.






 Sec. 23.1101   Induction air preheater design.

   Each exhaust-heated, induction air preheater must be designed and
 constructed to--
   (a) Ensure ventilation of the preheater when the induction air preheater is
 not being used during engine operation;
   (b) Allow inspection of the exhaust manifold parts that it surrounds; and
   (c) Allow inspection of critical parts of the preheater itself.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-43, 58 FR
 18974, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1103  Induction system ducts.

   (a) Each induction system duct must have a drain to prevent the
 accumulation of fuel or moisture in the normal ground and flight attitudes.
 No drain may discharge where it will cause a fire hazard.
   (b) Each duct connected to components between which relative motion could
 exist must have means for flexibility.
   (c) Each flexible induction system duct must be capable of withstanding the
 effects of temperature extremes, fuel, oil, water, and solvents to which it
 is expected to be exposed in service and maintenance without hazardous
 deterioration or delamination.
   (d) For reciprocating engine installations, each induction system duct must
 be--
   (1) Strong enough to prevent induction system failures resulting from
 normal backfire conditions; and
   (2) Fire resistant in any compartment for which a fire extinguishing system
 is required.
   (e) Each inlet system duct for an auxiliary power unit must be--
   (1) Fireproof within the auxiliary power unit compartment;
   (2) Fireproof for a sufficient distance upstream of the auxiliary power
 unit compartment to prevent hot gas reverse flow from burning through the
 duct and entering any other compartment of the airplane in which a hazard
 would be created by the entry of the hot gases;
   (3) Constructed of materials suitable to the environmental conditions
 expected in service, except in those areas requiring fireproof or fire
 resistant materials; and
   (4) Constructed of materials that will not absorb or trap hazardous
 quantities of flammable fluids that could be ignited by a surge or reverse-
 flow condition.
   (f) Induction system ducts that supply air to a cabin pressurization system
 must be suitably constructed of material that will not produce hazardous
 quantities of toxic gases or isolated to prevent hazardous quantities of
 toxic gases from entering the cabin during a powerplant fire.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13095, Aug. 13, 1969; Amdt. 23-43, 58 FR 18974, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1105  Induction system screens.

   If induction system screens are used--
   (a) Each screen must be upstream of the carburetor;
   (b) No screen may be in any part of the induction system that is the only
 passage through which air can reach the engine, unless--
   (1) The available preheat is at least 100 deg. F.; and
   (2) The screen can be deiced by heated air;
   (c) No screen may be deiced by alcohol alone; and
   (d) It must be impossible for fuel to strike any screen.






 Sec. 23.1107   Induction system filters.

   On reciprocating-engine installations, if an air filter is used to protect
 the engine against foreign material particles in the induction air supply--
   (a) Each air filter must be capable of withstanding the effects of
 temperature extremes, rain, fuel, oil, and solvents to which it is expected
 to be exposed in service and maintenance; and
   (b) Each air filter shall have a design feature to prevent material
 separated from the filter media from interfering with proper fuel metering
 operation.

 [Amdt. 23-43, 58 FR 18974, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1109  Turbocharger bleed air system.

   The following applies to turbocharged bleed air systems used for cabin
 pressurization:
   (a) The cabin air system may not be subject to hazardous contamination
 following any probable failure of the turbocharger or its lubrication system.
   (b) The turbocharger supply air must be taken from a source where it cannot
 be contaminated by harmful or hazardous gases or vapors following any
 probable failure or malfunction of the engine exhaust, hydraulic, fuel, or
 oil system.

 [Doc. No. 25811, 56 FR 354, Jan. 3, 1991]

 *****************************************************************************


 56 FR 344, No. 2, Jan. 3, 1991

   SUMMARY: This final rule upgrades the airworthiness standards for normal,
 utility, acrobatic, and commuter category airplanes. This amendment provides
 airworthiness standards for advancements in technology being incorporated in
 current designs, permits type certification of spin resistant airplanes, and
 reduces the regulatory burden in showing compliance with some of the
 requirements for the design and type certification of small airplanes. These
 new and amended airworthiness standards also result in the need for new
 definitions. As a result, new definitions are added.

   DATES: February 4, 1991.

 *****************************************************************************






 Sec. 23.1111  Turbine engine bleed air system.

   For turbine engine bleed air systems, the following apply:
   (a) No hazard may result if duct rupture or failure occurs anywhere between
 the engine port and the airplane unit served by the bleed air.
   (b) The effect on airplane and engine performance of using maximum bleed
 air must be established.
   (c) Hazardous contamination of cabin air systems may not result from
 failures of the engine lubricating system.

 [Amdt. 23-7, 34 FR 13095, Aug. 13, 1969, as amended by Amdt. 23-17, 41 FR
 55465, Dec. 20, 1976]






                                Exhaust System






 Sec. 23.1121  General.

   For powerplant and auxiliary power unit installations, the following
 apply--
   (a) Each exhaust system must ensure safe disposal of exhaust gases without
 fire hazard or carbon monoxide contamination in any personnel compartment.
   (b) Each exhaust system part with a surface hot enough to ignite flammable
 fluids or vapors must be located or shielded so that leakage from any system
 carrying flammable fluids or vapors will not result in a fire caused by
 impingement of the fluids or vapors on any part of the exhaust system
 including shields for the exhaust system.
   (c) Each exhaust system must be separated by fireproof shields from
 adjacent flammable parts of the airplane that are outside of the engine and
 auxiliary power unit compartments.
   (d) No exhaust gases may discharge dangerously near any fuel or oil system
 drain.
   (e) No exhaust gases may be discharged where they will cause a glare
 seriously affecting pilot vision at night.
   (f) Each exhaust system component must be ventilated to prevent points of
 excessively high temperature.
   (g) If significant traps exists, each turbine engine exhaust system must
 have drains discharging clear of the airplane, in any normal ground and
 flight attitude, to prevent fuel accumulation after the failure of an
 attempted engine start.
   (h) Each exhaust heat exchanger must incorporate means to prevent blockage
 of the exhaust port after any internal heat exchanger failure.
   (i) For the purpose of compliance with Sec. 23.603, the failure of any part
 of the exhaust system will be considered to adversely affect safety.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13095, Aug. 13, 1969; Amdt. 23-18, 42 FR 15042, Mar. 17, 1977; Amdt. 23-43,
 18974, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1123  Exhaust system.

   (a) Each exhaust system must be fireproof and corrosion-resistant, and
 must have means to prevent failure due to expansion by operating
 temperatures.
   (b) Each exhaust system must be supported to withstand the vibration and
 inertia loads to which it may be subjected in operation.
   (c) Parts of the system connected to components between which relative
 motion could exist must have means for flexibility.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-43, 58 FR
 18974, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1125  Exhaust heat exchangers.

   For reciprocating engine powered airplanes the following apply:
   (a) Each exhaust heat exchanger must be constructed and installed to
 withstand the vibration, inertia, and other loads that it may be subjected to
 in normal operation. In addition--
   (1) Each exchanger must be suitable for continued operation at high
 temperatures and resistant to corrosion from exhaust gases;
   (2) There must be means for inspection of critical parts of each exchanger;
 and
   (3) Each exchanger must have cooling provisions wherever it is subject to
 contact with exhaust gases.
   (b) Each heat exchanger used for heating ventilating air must be
 constructed so that exhaust gases may not enter the ventilating air.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-17, 41 FR
 55465, Dec. 20, 1976]






                      Powerplant Controls and Accessories






 Sec. 23.1141  Powerplant controls: general.

   (a) Powerplant controls must be located and arranged under Sec. 23.777 and
 marked under Sec. 23.1555(a).
   (b) Each flexible control must be of an acceptable kind.
   (c) Each control must be able to maintain any necessary position without--
   (1) Constant attention by flight crew members; or
   (2) Tendency to creep due to control loads or vibration.
   (d) Each control must be able to withstand operating loads without failure
 or excessive deflection.
   (e) For turbine engine powered airplanes, no single failure or malfunction,
 or probable combination thereof, in any powerplant control system may cause
 the failure of any powerplant function necessary for safety.
   (f) The portion of each powerplant control located in the engine
 compartment that is required to be operated in the event of fire must be at
 least fire resistant.
   (g) Powerplant valve controls located in the cockpit must have--
   (1) For manual valves, positive stops or in the case of fuel valves
 suitable index provisions, in the open and closed position; and
   (2) For power-assisted valves, a means to indicate to the flight crew when
 the valve--
   (i) Is in the fully open or fully closed position; or
   (ii) Is moving between the fully open and fully closed position.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13095, Aug. 13, 1969; Amdt. 23-14, 38 FR 31823, Nov. 19, 1973; Amdt. 23-18,
 42 FR 15042, Mar. 17, 1977]






 Sec. 23.1142   Auxiliary power unit controls.

   Means must be provided on the flight deck for the starting, stopping,
 monitoring, and emergency shutdown of each installed auxiliary power unit.

 [Amdt. 23-43, 58 FR 18974, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1143  Engine controls.

   (a) There must be a separate power or thrust control for each engine and a
 separate control for each supercharger that requires a control.
   (b) Power, thrust, and supercharger controls must be arranged to allow--
   (1) Separate control of each engine and each supercharger; and
   (2) Simultaneous control of all engines and all superchargers.
   (c) Each power, thrust, or supercharger control must give a positive and
 immediate responsive means of controlling its engine or supercharger.
   (d) The power, thrust, or supercharger controls for each engine or
 supercharger must be independent of those for every other engine or
 supercharger.
   (e) For each fluid injection (other than fuel) system and its controls not
 provided and approved as part of the engine, the applicant must show that the
 flow of the injection fluid is adequately controlled.
   (f) If a power or thrust control incorporates a fuel shutoff feature, the
 control must have a means to prevent the inadvertent movement of the control
 into the shutoff position. The means must--
   (1) Have a positive lock or stop at the idle position; and
   (2) Require a separate and distinct operation to place the control in the
 shutoff position.
   (g) For reciprocating single-engine airplanes, each power or thrust control
 must be designed so that if the control separates at the engine fuel metering
 device, the airplane is capable of continued safe flight and landing.

 [Amdt. 23-7, 34 FR 13095, Aug. 13, 1969, as amended by Amdt. 23-17, 41 FR
 55465, Dec. 20, 1976; Amdt. 23-29, 49 FR 6847, Feb. 23, 1984; Amdt. 23-43, 58
 FR 18974, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1145  Ignition switches.

   (a) Ignition switches must control and shut off each ignition circuit on
 each engine.
   (b) There must be means to quickly shut off all ignition on multiengine
 airplanes by the grouping of switches or by a master ignition control.
   (c) Each group of ignition switches, except ignition switches for turbine
 engines for which continuous ignition is not required, and each master
 ignition control must have a means to prevent its inadvertent operation.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. 23-18, 42 FR 15042, Mar. 17, 1977; Amdt. 23-43, 58 FR 18974,
 Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1147  Mixture controls.

   (a) If there are mixture controls, each engine must have a separate
 control, and each mixture control must have guards or must be shaped or
 arranged to prevent confusion by feel with other controls.
   (1) The controls must be grouped and arranged to allow--
   (i) Separate control of each engine; and
   (ii) Simultaneous control of all engines.
   (2) The controls must require a separate and distinct operation to move the
 control toward lean or shut-off position.
   (b) For reciprocating single-engine airplanes, each manual engine mixture
 control must be designed so that, if the control separates at the engine fuel
 metering device, the airplane is capable of continued safe flight and
 landing.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13096, Aug. 13, 1969; Amdt. 23-33, 51 FR 26657, July 24, 1986; Amdt. 23-43,
 58 FR 18974, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1149  Propeller speed and pitch controls.

   (a) If there are propeller speed or pitch controls, they must be grouped
 and arranged to allow--
   (1) Separate control of each propeller; and
   (2) Simultaneous control of all propellers.
   (b) The controls must allow ready synchronization of all propellers on
 multiengine airplanes.






 Sec. 23.1153  Propeller feathering controls.

   If there are propeller feathering controls, each propeller must have a
 separate control. Each control must have means to prevent inadvertent
 operation.






 Sec. 23.1155  Turbine engine reverse thrust and propeller pitch settings
     below the flight regime.

   For turbine engine installations, each control for reverse thrust and for
 propeller pitch settings below the flight regime must have means to prevent
 its inadvertent operation. The means must have a positive lock or stop at the
 flight idle position and must require a separate and distinct operation by
 the crew to displace the control from the flight regime (forward thrust
 regime for turbojet powered airplanes).

 [Amdt. 23-7, 34 FR 13096, Aug. 13, 1969]






 Sec. 23.1157  Carburetor air temperature controls.

   There must be a separate carburetor air temperature control for each
 engine.






 Sec. 23.1163   Powerplant accessories.

   (a) Each engine mounted accessory must--
   (1) Be approved for mounting on the engine involved and use the provisions
 on the engines for mounting; or
   (2) Have torque limiting means on all accessory drives in order to prevent
 the torque limits established for those drives from being exceeded; and
   (3) In addition to paragraphs (a)(1) or (a)(2) of this section, be sealed
 to prevent contamination of the engine oil system and the accessory system.
   (b) Electrical equipment subject to arcing or sparking must be installed to
 minimize the probability of contact with any flammable fluids or vapors that
 might be present in a free state.
   (c) Each generator rated at or more than 6 kilowatts must be designed and
 installed to minimize the probability of a fire hazard in the event it
 malfunctions.
   (d) If the continued rotation of any accessory remotely driven by the
 engine is hazardous when malfunctioning occurs, a means to prevent rotation
 without interfering with the continued operation of the engine must be
 provided.
   (e) Each accessory driven by a gearbox that is not approved as part of the
 powerplant driving the gearbox must--
   (1) Have torque limiting means to prevent the torque limits established for
 the affected drive from being exceeded;
   (2) Use the provisions on the gearbox for mounting; and
   (3) Be sealed to prevent contamination of the gearbox oil system and the
 accessory system.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-14, 38 FR
 31823, Nov. 19, 1973; Amdt. 23-29, 49 FR 6847, Feb. 23, 1984; Amdt. 23-34, 52
 FR 1832, Jan. 15, 1987; Amdt. 23-42, 56 FR 354, Jan. 3, 1991]

 *****************************************************************************


 56 FR 344, No. 2, Jan. 3, 1991

   SUMMARY: This final rule upgrades the airworthiness standards for normal,
 utility, acrobatic, and commuter category airplanes. This amendment provides
 airworthiness standards for advancements in technology being incorporated in
 current designs, permits type certification of spin resistant airplanes, and
 reduces the regulatory burden in showing compliance with some of the
 requirements for the design and type certification of small airplanes. These
 new and amended airworthiness standards also result in the need for new
 definitions. As a result, new definitions are added.

   DATES: February 4, 1991.

 *****************************************************************************






 Sec. 23.1165   Engine ignition systems.

   (a) Each battery ignition system must be supplemented by a generator that
 is automatically available as an alternate source of electrical energy to
 allow continued engine operation if any battery becomes depleted.
   (b) The capacity of batteries and generators must be large enough to meet
 the simultaneous demands of the engine ignition system and the greatest
 demands of any electrical system components that draw from the same source.
   (c) The design of the engine ignition system must account for--
   (1) The condition of an inoperative generator;
   (2) The condition of a completely depleted battery with the generator
 running at its normal operating speed; and
   (3) The condition of a completely depleted battery with the generator
 operating at idling speed, if there is only one battery.
   (d) There must be means to warn appropriate crewmembers if malfunctioning
 of any part of the electrical system is causing the continuous discharge of
 any battery used for engine ignition.
   (e) Each turbine engine ignition system must be independent of any
 electrical circuit that is not used for assisting, controlling, or analyzing
 the operation of that system.
   (f) In addition, for commuter category airplanes, each turbopropeller
 ignition system must be an essential electrical load.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-17, 41 FR
 55465 Dec. 20, 1976; Amdt. 23-34, 52 FR 1833, Jan. 15, 1987]






 Sec. 23.1181   Designated fire zones; regions included.

   Designated fire zones are--
   (a) For reciprocating engines--
   (1) The power section;
   (2) The accessory section;
   (3) Any complete powerplant compartment in which there is no isolation
 between the power section and the accessory section.
   (b) For turbine engines--
   (1) The compressor and accessory sections;
   (2) The combustor, turbine and tailpipe sections that contain lines or
 components carrying flammable fluids or gases.
   (c) Any auxiliary power unit compartment; and
   (d) Any fuel-burning heater, and other combustion equipment installation
 described in Sec. 23.859;

 [Amdt. 23-43, 58 FR 18975, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






                          Powerplant Fire Protection






 Sec. 23.1182   Nacelle areas behind firewalls.

   Components, lines, and fittings, except those subject to the provisions of
 Sec. 23.1351(e), located behind the engine-compartment firewall must be
 constructed of such materials and located at such distances from the firewall
 that they will not suffer damage sufficient to endanger the airplane if a
 portion of the engine side of the firewall is subjected to a flame
 temperature of not less than 2000 deg. F for 15 minutes.

 [Amdt. 23-14, 38 FR 31816, Nov. 19, 1973]






 Sec. 23.1183   Lines, fittings, and components.

   (a) Except as provided in paragraph (b) of this section, each component,
 line, and fitting carrying flammable fluids, gas, or air in any area subject
 to engine fire conditions must be at least fire resistant, except that
 flammable fluid tanks and supports which are part of and attached to the
 engine must be fireproof or be enclosed by a fireproof shield unless damage
 by fire to any non-fireproof part will not cause leakage or spillage of
 flammable fluid. Components must be shielded or located so as to safeguard
 against the ignition of leaking flammable fluid. Flexible hose assemblies
 (hose and end fittings) must be approved. An integral oil sump of less than
 25-quart capacity on a reciprocating engine need not be fireproof nor be
 enclosed by a fireproof shield.
   (b) Paragraph (a) of this section does not apply to--
   (1) Lines, fittings, and components which are already approved as part of a
 type certificated engine; and
   (2) Vent and drain lines, and their fittings, whose failure will not result
 in, or add to, a fire hazard.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-5, 32 FR
 6912, May 5, 1967; Amdt. 23-15, 39 FR 35460, Oct. 1, 1974; Amdt. 23-29, 49 FR
 6847, Feb. 23, 1984]






 Sec. 23.1189   Shutoff means.

   (a) For each multiengine airplane, the following apply:
   (1) Each engine installation must have means to shut off or otherwise
 prevent hazardous quantities of fuel, oil, deicing fluid, and other flammable
 liquids from flowing into, within, or through any engine compartment, except
 in lines, fittings, and components forming an integral part of an engine.
   (2) The closing of the fuel shutoff valve for any engine may not make any
 fuel unavailable to the remaining engines that would be available to those
 engines with that valve open.
   (3) Operation of any shutoff means may not interfere with the later
 emergency operation of other equipment such as propeller feathering devices.
   (4) Each shutoff must be outside of the engine compartment unless an equal
 degree of safety is provided with the shutoff inside the compartment.
   (5) Not more than one quart of flammable fluid may escape into the engine
 compartment after engine shutoff. For those installations where the flammable
 fluid that escapes after shutdown cannot be limited to one quart, it must be
 demonstrated that this greater amount can be safely contained or drained
 overboard.
   (6) There must be means to guard against inadvertent operation of each
 shutoff means, and to make it possible for the crew to reopen the shutoff
 means in flight after it has been closed.
   (b) Turbine engine installations need not have an engine oil system shutoff
 if--
   (1) The oil tank is integral with, or mounted on, the engine; and
   (2) All oil system components external to the engine are fireproof or
 located in areas not subject to engine fire conditions.
   (c) Power operated valves must have means to indicate to the flight crew
 when the valve has reached the selected position and must be designed so that
 the valve will not move from the selected position under vibration conditions
 likely to exist at the valve location.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13096, Aug. 13, 1969; Amdt. 23-14, 38 FR 31823, Nov. 19, 1973; Amdt. 23-29,
 49 FR 6847, Feb. 23, 1984; Amdt. 23-43, 58 FR 18975, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1191   Firewalls.

   (a) Each engine, auxiliary power unit, fuel burning heater, and other
 combustion equipment, must be isolated from the rest of the airplane by
 firewalls, shrouds, or equivalent means.
   (b) Each firewall or shroud must be constructed so that no hazardous
 quantity of liquid, gas, or flame can pass from the isolated compartment to
 other parts of the airplane.
   (c) Each opening in the firewall or shroud must be sealed with close
 fitting, fireproof grommets, bushings, or firewall fittings.
   (d) [Reserved]
   (e) Each firewall and shroud must be fireproof and protected against
 corrosion.
   (f) Compliance with the criteria for fireproof materials or components must
 be shown as follows:
   (1) The flame to which the materials or components are subjected must be
 2000 +/- 150 deg.F.
   (2) Sheet materials approximately 10 inches square must be subjected to the
 flame from a suitable burner.
   (3) The flame must be large enough to maintain the required test
 temperature over an area approximately five inches square.
   (g) Firewall materials and fittings must resist flame penetration for at
 least 15 minutes.
   (h) The following materials may be used in firewalls or shrouds without
 being tested as required by this section:
   (1) Stainless steel sheet, 0.015 inch thick.
   (2) Mild steel sheet (coated with aluminum or otherwise protected against
 corrosion) 0.018 inch thick.
   (3) Terne plate, 0.018 inch thick.
   (4) Monel metal, 0.018 inch thick.
   (5) Steel or copper base alloy firewall fittings.
   (6) Titanium sheet, 0.016 inch thick.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-43, 58 FR
 18975, Apr. 9, 1993; 58 FR 27060, May 6, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1192   Engine accessory compartment diaphragm.

   For aircooled radial engines, the engine power section and all portions of
 the exhaust sytem must be isolated from the engine accessory compartment by a
 diaphragm that meets the firewall requirements of Sec. 23.1191.

 [Amdt. 23-14, 38 FR 31823, Nov. 19, 1973]






 Sec. 23.1193   Cowling and nacelle.

   (a) Each cowling must be constructed and supported so that it can resist
 any vibration, inertia, and air loads to which it may be subjected in
 operation.
   (b) There must be means for rapid and complete drainage of each part of the
 cowling in the normal ground and flight attitudes. Drain operation may be
 shown by test, analysis, or both, to ensure that under normal aerodynamic
 pressure distribution expected in service each drain will operate as
 designed. No drain may discharge where it will cause a fire hazard.
   (c) Cowling must be at least fire resistant.
   (d) Each part behind an opening in the engine compartment cowling must be
 at least fire resistant for a distance of at least 24 inches aft of the
 opening.
   (e) Each part of the cowling subjected to high temperatures due to its
 nearness to exhaust sytem ports or exhaust gas impingement, must be fire
 proof.
   (f) Each nacelle of a multiengine airplane with supercharged engines must
 be designed and constructed so that with the landing gear retracted, a fire
 in the engine compartment will not burn through a cowling or nacelle and
 enter a nacelle area other than the engine compartment.
   (g) In addition, for commuter category airplanes, the airplane must be
 designed so that no fire originating in any engine compartment can enter,
 either through openings or by burn-through, any other region where it would
 create additional hazards.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. 23-18, 42 FR 15042, Mar. 17, 1977; Amdt. 23-34, 52 FR 1833,
 Jan. 15, 1987; Amdt. 23-43, 58 FR 18975, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1195   Fire extinguishing systems.

   (a) For commuter category airplanes, fire extinguishing systems must be
 installed and compliance shown with the following:
   (1) Except for combustor, turbine, and tailpipe sections of turbine-engine
 installations that contain lines or components carrying flammable fluids or
 gases for which a fire originating in these sections is shown to be
 controllable, a fire extinguisher system must serve each engine compartment;
   (2) The fire extinguishing system, the quantity of the extinguishing agent,
 the rate of discharge, and the discharge distribution must be adequate to
 extinguish fires. An individual "one shot" system may be used.
   (3) The fire extinguishing system for a nacelle must be able to
 simultaneously protect each compartment of the nacelle for which protection
 is provided.
   (b) If an auxiliary power unit is installed in any airplane certificated to
 this part, that auxiliary power unit compartment must be served by a fire
 extinguishing system meeting the requirements of paragraph (a)(2) of this
 section.

 [Amdt. 23-34, 52 FR 1833, Jan. 15, 1987, as amended by Amdt. 23-43, 58 FR
 18975, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1197   Fire extinguishing agents.

   For commuter category airplanes, the following applies:
   (a) Fire extinguishing agents must--
   (1) Be capable of extinguishing flames emanating from any burning of fluids
 or other combustible materials in the area protected by the fire
 extinguishing system; and
   (2) Have thermal stability over the temperature range likely to be
 experienced in the compartment in which they are stored.
   (b) If any toxic extinguishing agent is used, provisions must be made to
 prevent harmful concentrations of fluid or fluid vapors (from leakage during
 normal operation of the airplane or as a result of discharging the fire
 extinguisher on the ground or in flight) from entering any personnel
 compartment, even though a defect may exist in the extinguishing system. This
 must be shown by test except for built-in carbon dioxide fuselage compartment
 fire extinguishing systems for which--
   (1) Five pounds or less of carbon dioxide will be discharged, under
 established fire control procedures, into any fuselage compartment; or
   (2) Protective breathing equipment is available for each flight crewmember
 on flight deck duty.

 [Amdt. 23-34, 52 FR 1833, Jan. 15, 1987]






 Sec. 23.1199  Extinguishing agent containers.

   For commuter category airplanes, the following applies:
   (a) Each extinguishing agent container must have a pressure relief to
 prevent bursting of the container by excessive internal pressures.
   (b) The discharge end of each discharge line from a pressure relief
 connection must be located so that discharge of the fire extinguishing agent
 would not damage the airplane. The line must also be located or protected to
 prevent clogging caused by ice or other foreign matter.
   (c) A means must be provided for each fire extinguishing agent container to
 indicate that the container has discharged or that the charging pressure is
 below the established minimum necessary for proper functioning.
   (d) The temperature of each container must be maintained, under intended
 operating conditions, to prevent the pressure in the container from--
   (1) Falling below that necessary to provide an adequate rate of discharge;
 or
   (2) Rising high enough to cause premature discharge.
   (e) If a pyrotechnic capsule is used to discharge the extinguishing agent,
 each container must be installed so that temperature conditions will not
 cause hazardous deterioration of the pyrotechnic capsule.

 [Amdt. 23-34, 52 FR 1833, Jan. 15, 1987; 52 FR 34745, Sept. 14, 1987]






 Sec. 23.1201  Fire extinguishing system materials.

   For commuter category airplanes, the following apply:
   (a) No material in any fire extinguishing system may react chemically with
 any extinguishing agent so as to create a hazard.
   (b) Each system component in an engine compartment must be fireproof.

 [Amdt. 23-34, 52 FR 1833, Jan. 15, 1987; 52 FR 7262, Mar. 9, 1987]






 Sec. 23.1203   Fire detector system.

   (a) There must be means that ensure the prompt detection of a fire in--
   (1) An engine compartment of--
   (i) Multiengine turbine powered airplanes;
   (ii) Multiengine reciprocating engine powered airplanes incorporating
 turbochargers;
   (iii) Airplanes with engine(s) located where they are not readily visible
 from the cockpit; and
   (iv) All commuter category airplanes.
   (2) The auxiliary power unit compartment of any airplane incorporating an
 auxiliary power unit.
   (b) Each fire detector must be constructed and installed to withstand the
 vibration, inertia, and other loads to which it may be subjected in
 operation.
   (c) No fire detector may be affected by any oil, water, other fluids, or
 fumes that might be present.
   (d) There must be means to allow the crew to check, in flight, the
 functioning of each fire detector electric circuit.
   (e) Wiring and other components of each fire detector system in a fire
 zone must be at least fire resistant.

 [Amdt. 23-18, 42 FR 15042, Mar. 17, 1977, as amended by Amdt. 23-34, 52 FR
 1833, Jan. 15, 1987; Amdt. 23-43, 58 FR 18975, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************



                             Subpart F--Equipment






                                    General




 Sec. 23.1301  Function and installation.

   Each item of installed equipment must--
   (a) Be of a kind and design appropriate to its intended function.
   (b) Be labeled as to its identification, function, or operating
 limitations, or any applicable combination of these factors;
   (c) Be installed according to limitations specified for that equipment; and
   (d) Function properly when installed.

 [Amdt. 23-20, 42 FR 36968, July 18, 1977]






 Sec. 23.1303  Flight and navigation instruments.

   The following are required flight and navigational instruments:
   (a) An airspeed indicator.
   (b) An altimeter.
   (c) A direction indicator (nonstabilized magnetic compass).
   (d) For turbine engine powered airplanes, a free air temperature indicator
 or an air-temperature indicator which provides indications that are
 convertible to free-air.
   (e) A speed warning device for--
   (1) Turbine engine powered airplanes; and
   (2) Other airplanes for which Vmo/Mmo and Vd/Md are established under Secs.
 23.335(b)(4) and 23.1505(c) if Vmo/Mmo is greater than 0.8 Vd/Md.
   The speed warning device must give effective aural warning (differing
 distinctively from aural warnings used for other purposes) to the pilots
 whenever the speed exceeds Vmo plus 6 knots or Mmo+0.01. The upper limit of
 the production tolerance for the warning device may not exceed the prescribed
 warning speed.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-17, 41 FR
 55465, Dec. 20, 1976; Amdt. 23-43, 58 FR 18975, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1305  Powerplant instruments.

   The following are required powerplant instruments:
   (a) For all airplanes.
   (1) A fuel quantity indicator for each fuel tank, installed in accordance
 with Sec. 23.1337(b).
   (2) An oil pressure indicator for each engine.
   (3) An oil temperature indicator for each engine.
   (4) An oil quantity measuring device for each oil tank which meets the
 requirements of Sec. 23.1337(d).
   (5) A fire warning means for those airplanes required to comply with Sec.
 23.1203.
   (b) For reciprocating engine-powered airplanes. In addition to the
 powerplant instruments required by paragraph (a) of this section, the
 following powerplant instruments are required:
   (1) An induction system air temperature indicator for each engine equipped
 with a preheater and having induction air temperature limitations that can be
 exceeded with preheat.
   (2) A tachometer indicator for each engine.
   (3) A cylinder head temperature indicator for--
   (i) Each air-cooled engine with cowl flaps;
   (ii) Each airplane for which compliance with Sec. 23.1041 is shown at a
 speed higher than Vy; and
   (iii) Each commuter category airplane.
   (4) A fuel pressure indicator for each pump fed engine.
   (5) A manifold pressure indicator for each altitude engine and for each
 engine with a controllable propeller.
   (6) For each turbocharger installation:
   (i) If limitations are established for either carburetor (or manifold) air
 inlet temperature or exhaust gas or turbocharger turbine inlet temperature,
 indicators must be furnished for each temperature for which the limitation is
 established unless it is shown that the limitation will not be exceeded in
 all intended operations.
   (ii) If its oil system is separate from the engine oil system, oil pressure
 and oil temperature indicators must be provided.
   (7) A coolant temperature indicator for each liquid-cooled engine.
   (c) For turbine engine-powered airplanes. In addition to the powerplant
 instruments required by paragraph (a) of this section, the following
 powerplant instruments are required:
   (1) A gas temperature indicator for each engine.
   (2) A fuel flowmeter indicator for each engine.
   (3) A fuel low pressure warning means for each engine.
   (4) A fuel low level warning means for any fuel tank that should not be
 depleted of fuel in normal operations.
   (5) A tachometer indicator (to indicate the speed of the rotors with
 established limiting speeds) for each engine.
   (6) An oil low pressure warning means for each engine.
   (7) An indicating means to indicate the functioning of the powerplant ice
 protection system for each engine.
   (8) For each engine, an indicating means for the fuel strainer or filter
 required by Sec. 23.997 to indicate the occurrence of contamination of the
 strainer or filter before it reaches the capacity established in accordance
 with Sec. 23.997(d).
   (9) For each engine, a warning means for the oil strainer or filter
 required by Sec. 23.1019, if it has no bypass, to warn the pilot of the
 occurrence of contamination of the strainer or filter screen before it
 reaches the capacity established in accordance with Sec. 23.1019(a)(5).
   (10) An indicating means to indicate the functioning of any heater used to
 prevent ice clogging of fuel system components.
   (d) For turbojet/turbofan engine-powered airplanes. In addition to the
 powerplant instruments required by paragraphs (a) and (c) of this section,
 the following powerplant instruments are required:
   (1) For each engine, an indicator to indicate thrust or to indicate a
 parameter that can be related to thrust, including a free air temperature
 indicator if needed for this purpose.
   (2) For each engine, a position indicating means to indicate to the flight
 crew when the thrust reverser, if installed, is in the reverse thrust
 position.
   (e) For turbopropeller-powered airplanes. In addition to the powerplant
 instruments required by paragraphs (a) and (c) of this section, the following
 powerplant instruments are required:
   (1) A torque indicator for each engine.
   (2) A position indicating means to indicate to the flight crew when the
 propeller blade angle is below the flight low pitch position, for each
 propeller, unless it can be shown that such occurrence is highly improbable.

 [Amdt. 23-43, 58 FR 18975, Apr. 9, 1993; 58 FR 27060, May 6, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1307  Miscellaneous equipment.

   (a) There must be a seat or berth for each occupant.
   (b) The following miscellaneous equipment is required as prescribed in this
 subpart:
   (1) A master switch arrangement.
   (2) An adequate source of electrical energy.
   (3) Electrical protective devices.
   (c) The equipment necessary for an airplane to operate at the maximum
 operating altitude and in the kinds of operations and meteorological
 conditions for which certification is requested and is approved in accordance
 with Sec. 23.1559 must be included in the type design.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. 23-23, 43 FR 50593, Oct. 30, 1978; Amdt. 23-43, 58 FR 18976,
 Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1309  Equipment, systems, and installations.

   (a) Each item of equipment, each system, and each installation:
   (1) When performing its intended function, may not adversely affect the
 response, operation, or accuracy of any--
   (i) Equipment essential to safe operation; or
   (ii) Other equipment unless there is a means to inform the pilot of the
 effect.
   (2) In a single-engine airplane, must be designed to minimize hazards to
 the airplane in the event of a probable malfunction or failure.
   (3) In a multiengine airplane, must be designed to prevent hazards to the
 airplane in the event of a probable malfunction or failure.
   (b) The design of each item of equipment, each system, and each
 installation must be examined separately and in relationship to other
 airplane systems and installations to determine if the airplane is dependent
 upon its function for continued safe flight and landing and, for airplanes
 not limited to VFR conditions, if failure of a system would significantly
 reduce the capability of the airplane or the ability of the crew to cope with
 adverse operating conditions. Each item of equipment, each system, and each
 installation identified by this examination as one upon which the airplane is
 dependent for proper functioning to ensure continued safe flight and landing,
 or whose failure would significantly reduce the capability of the airplane or
 the ability of the crew to cope with adverse operating conditions, must be
 designed to comply with the following additional requirements:
   (1) It must perform its intended function under any foreseeable operating
 condition.
   (2) When systems and associated components are considered separately and in
 relation to other systems--
   (i) The occurrence of any failure condition that would prevent the
 continued safe flight and landing of the airplane must be extremely
 improbable; and
   (ii) The occurrence of any other failure condition that would significantly
 reduce the capability of the airplane or the ability of the crew to cope with
 adverse operating conditions must be improbable.
   (3) Warning information must be provided to alert the crew to unsafe system
 operating conditions and to enable them to take appropriate corrective
 action. Systems, controls, and associated monitoring and warning means must
 be designed to minimize crew errors that could create additional hazards.
   (4) Compliance with the requirements of paragraph (b)(2) of this section
 may be shown by analysis and, where necessary, by appropriate ground, flight,
 or simulator tests. The analysis must consider--
   (i) Possible modes of failure, including malfunctions and damage from
 external sources;
   (ii) The probability of multiple failures, and the probability of
 undetected faults.;
   (iii) The resulting effects on the airplane and occupants, considering the
 stage of flight and operating conditions; and
   (iv) The crew warning cues, corrective action required, and the crew's
 capability of determining faults.
   (c) Each item of equipment, each system, and each installation whose
 functioning is required by this chapter and that requires a power supply is
 an "essential load" on the power supply. The power sources and the system
 must be able to supply the following power loads in probable operating
 combinations and for probable durations:
   (1) Loads connected to the power distribution system with the system
 functioning normally.
   (2) Essential loads after failure of--
   (i) Any one engine on two-engine airplanes; or
   (ii) Any two engines on an airplane with three or more engines; or
   (iii) Any power converter or energy storage device.
   (3) Essential loads for which an alternate source of power is required, as
 applicable, by the operating rules of this chapter, after any failure or
 malfunction in any one power supply system, distribution system, or other
 utilization system.
   (d) In determining compliance with paragraph (c)(2) of this section, the
 power loads may be assumed to be reduced under a monitoring procedure
 consistent with safety in the kinds of operations authorized. Loads not
 required in controlled flight need not be considered for the two-engine-
 inoperative condition on airplanes with three or more engines.
   (e) In showing compliance with this section with regard to the electrical
 power system and to equipment design and installation, critical environmental
 and atmospheric conditions, including radio frequency energy and the effects
 (both direct and indirect) of lightning strikes, must be considered. For
 electrical generation, distribution, and utilization equipment required by or
 used in complying with this chapter, the ability to provide continuous, safe
 service under forseeable environmental conditions may be shown by
 environmental tests, design analysis, or reference to previous comparable
 service experience on other airplanes.
   (f) As used in this section, "system" refers to all pneumatic systems,
 fluid systems, electrical systems, mechanical systems, and powerplant systems
 included in the airplane design, except for the following:
   (1) Powerplant systems provided as part of the certificated engine.
   (2) The flight structure (such a wing, empennage, control surfaces and
 their systems, the fuselage, engine mounting, and landing gear and their
 related primary attachments) whose requirements are specific in subparts C
 and D of this part.

 [Doc. No. 25812, Amdt. 23-41, 55 FR 43309, Oct. 26, 1990; 55 FR 47028,
 Nov. 8, 1990]

 *****************************************************************************


 55 FR 43306, No. 208, Oct. 26, 1990

   SUMMARY: This final rule amends the airworthiness standards for equipment,
 systems, and installations and establishes airworthiness standards for the
 installation of electronic display instrument systems in normal, utility,
 acrobatic, and commuter category airplanes. It also provides alternative
 airworthiness standards for the instrument configuration for general, air
 taxi and commercial operations. This amendment updates the airworthiness and
 operating requirements to reflect advanced technology being incorporated in
 current designs while maintaining an acceptable level of safety.

   EFFECTIVE DATE: November 26, 1990.

 *****************************************************************************






                           Instruments: Installation






 Sec. 23.1311  Electronic display instrument systems.

   (a) Electonic display indicator requirements in this section are
 independent to each pilot station required by the airworthiness standards or
 by the applicable operating rules for each airplane that is to be approved
 for operation in IFR conditions.
   (b) Electronic display indicators required by Sec. 23.1303(a), (b), and (c)
 must be independent of the airplane's electrical power system.
   (c) Electronic display indicators, including those with features that make
 isolation and independence between powerplant instrument systems impractical
 must--
   (1) Be easily legible under all lighting conditions encountered in the
 cockpit, including direct sunlight, considering the expected electronic
 display brightness level at the end of an electronic display indicator's
 useful life. Specific limitations on display system useful life must be
 addressed in the Instructions for Continued Airworthiness requirements of
 Sec. 23.1529;
   (2) Not inhibit the primary display of attitude, airspeed, altitude, or
 powerplant parameters needed by any pilot to set power within established
 limitations, in any normal mode of operation;
   (3) Not inhibit the primary display of engine parameters needed by any
 pilot to properly set or monitor powerplant limitations during the engine
 starting mode of operation;
   (4) Have independent secondary attitude and rate-of-turn instruments that
 comply with Sec. 23.1321(a) if the primary electronic display instrument
 system for a pilot presents this information. Instrument displays that are
 located in accordance with Sec. 23.1321(d) are considered the primary
 displays. A rate-of-turn instrument is not required if a third attitude
 instrument system is installed in accordance with the instrument requirements
 prescribed in Sec. 121.305(j) of this chapter.
   (5) Incorporate sensory cues for the pilot that are equivalent to those in
 the instrument being replaced by the electronic display indicators; and
   (6) Incorporate visual displays of instrument markings, required by Secs.
 23.1541 through 23.1553, or visual displays that alert the pilot to abnormal
 operational values or approaches to established limitation values, for each
 parameter required to be displayed by this part.
   (d) The electronic display indicators, including their systems and
 installations, and considering other airplane systems, must be designed so
 that one display of information essential for continued safe flight and
 landing will remain available to the crew, without need for immediate action
 by any pilot for continued safe operation, after any single failure or
 probable combination of failures.
   (e) As used in this section, "instrument" includes devices that are
 physically contained in one unit, and devices that are composed of two or
 more physically separate units or components connected together (such as a
 remote indicating gyroscopic direction indicator that includes a magnetic
 sensing element, a gyroscopic unit, an amplifier, and an indicator connected
 together). As used in this section, "primary" display refers to the display
 of a parameter that is located in the instrument panel such that the pilot
 looks at it first when wanting to view that parameter.

 [Doc. No. 25812, Amdt. 23-41, 55 FR 43310, Oct. 26, 1990; 55 FR 47028,
 Nov. 8, 1990]

 *****************************************************************************


 55 FR 43306, No. 208, Oct. 26, 1990

   SUMMARY: This final rule amends the airworthiness standards for equipment,
 systems, and installations and establishes airworthiness standards for the
 installation of electronic display instrument systems in normal, utility,
 acrobatic, and commuter category airplanes. It also provides alternative
 airworthiness standards for the instrument configuration for general, air
 taxi and commercial operations. This amendment updates the airworthiness and
 operating requirements to reflect advanced technology being incorporated in
 current designs while maintaining an acceptable level of safety.

   EFFECTIVE DATE: November 26, 1990.

 *****************************************************************************






 Sec. 23.1321   Arrangement and visibility.

   (a) Each flight, navigation, and powerplant instrument for use by any
 required pilot during takeoff, initial climb, final approach, and landing
 must be located so that any pilot seated at the controls can monitor the
 airplane's flight path and these instruments with minimum head and eye
 movement. The powerplant instruments for these flight conditions are those
 needed to set power within powerplant limitations.
   (b) For each multiengine airplane, identical powerplant instruments must be
 located so as to prevent confusion as to which engine each instrument
 relates.
   (c) Instrument panel vibration may not damage, or impair the accuracy of,
 any instrument.
   (d) For each airplane certificated for flight under instrument flight rules
 or of more than 6,000 pounds maximum weight, the flight instruments required
 by Sec. 23.1303, and, as applicable, by the operating rules of this chapter,
 must be grouped on the instrument panel and centered as nearly as practicable
 about the vertical plane of each required pilot's forward vision. In
 addition:
   (1) The instrument that most effectively indicates the attitude must be on
 the panel in the top center position;
   (2) The instrument that most effectively indicates airspeed must be
 adjacent to and directly to the left of the instrument in the top center
 position;
   (3) The instrument that most effectively indicates altitude must be
 adjacent to and directly to the right of the instrument in the top center
 position;
   (4) The instrument that most effectively indicates direction of flight,
 other than the magnetic direction indicator required by Sec. 23.1303(c), must
 be adjacent to and directly below the instrument in the top center position;
 and
   (5) Electronic display indicators may be used for compliance with
 paragraphs (d)(1) through (d)(4) of this section when such displays comply
 with requirements in Sec. 23.1311.
   (e) If a visual indicator is provided to indicate malfunction of an
 instrument, it must be effective under all probable cockpit lighting
 conditions.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-14, 38 FR
 31824, Nov. 19, 1973; Amdt. 23-20, 42 FR 36968, July 18, 1977; Amdt. 23-41,
 55 FR 43310, Oct. 26, 1990; 55 FR 46888, Nov. 7, 1990]

 *****************************************************************************


 55 FR 43306, No. 208, Oct. 26, 1990

   SUMMARY: This final rule amends the airworthiness standards for equipment,
 systems, and installations and establishes airworthiness standards for the
 installation of electronic display instrument systems in normal, utility,
 acrobatic, and commuter category airplanes. It also provides alternative
 airworthiness standards for the instrument configuration for general, air
 taxi and commercial operations. This amendment updates the airworthiness and
 operating requirements to reflect advanced technology being incorporated in
 current designs while maintaining an acceptable level of safety.

   EFFECTIVE DATE: November 26, 1990.

 *****************************************************************************






 Sec. 23.1322   Warning, caution, and advisory lights.

   If warning, caution, or advisory lights are installed in the cockpit, they
 must, unless otherwise approved by the Administrator, be--
   (a) Red, for warning lights (lights indicating a hazard which may require
 immediate corrective action);
   (b) Amber, for caution lights (lights indicating the possible need for
 future corrective action);
   (c) Green, for safe operation lights; and
   (d) Any other color, including white, for lights not described in
 paragraphs (a) through (c) of this section, provided the color differs
 sufficiently from the colors prescribed in paragraphs (a) through (c) of this
 section to avoid possible confusion.
   (e) Effective under all probable cockpit lighting conditions.

 [Amdt. 23-17, 41 FR 55465, Dec. 20, 1976, as amended by Amdt. 23-43, 58 FR
 18976, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1323   Airspeed indicating system.

   (a) Each airspeed indicating instrucment must be calibrated to indicate
 true airspeed (at sea level with a standard atmosphere) with a minimum
 practicable instrument calibration error when the corresponding pitot and
 static pressures are applied.
   (b) Each airspeed system must be calibrated in flight to determine the
 system error. The system error, including position error, but excluding the
 airspeed indicator instrument calibration error, may not exceed three percent
 of the calibrated airspeed or five knots, whichever is greater, throughout
 the following speed ranges:
   (1) 1.3 VS1 to VMO/MMO or VNE, whichever is appropriate with flaps
 retracted.
   (2) 1.3 VS1 to VFE with flaps extended.
   (c) In addition, for commuter category airplanes, the airspeed indicating
 system must be calibrated to determine the system error in flight and during
 the accelerate-takeoff ground run. The ground run calibration must be
 obtained between 0.8 of the minimum value of V1, and 1.2 times the maximum
 value of V1 considering the approved ranges of altitude and weight. The
 ground run calibration must be determined assuming an engine failure at the
 minimum value of V1.
   (d) For commuter category airplanes, the information showing the
 relationship between IAS and CAS determined in accordance with paragraph (c)
 of this section must be shown in the Airplane Flight Manual.
   (e) If certification for instrument flight rules or flight in icing
 conditions is requested, each airspeed system must have a heated pitot tube
 or an equivalent means of preventing malfunction due to icing.

 [Amdt. 23-20, 42 FR 36968, July 18, 1977, as amended by Amdt. 23-34, 52 FR
 1834, Jan. 15, 1987; 52 FR 34745, Sept. 14, 1987; Amdt. 23-42, 56 FR 354,
 Jan. 3, 1991]

 *****************************************************************************


 56 FR 344, No. 2, Jan. 3, 1991

   SUMMARY: This final rule upgrades the airworthiness standards for normal,
 utility, acrobatic, and commuter category airplanes. This amendment provides
 airworthiness standards for advancements in technology being incorporated in
 current designs, permits type certification of spin resistant airplanes, and
 reduces the regulatory burden in showing compliance with some of the
 requirements for the design and type certification of small airplanes. These
 new and amended airworthiness standards also result in the need for new
 definitions. As a result, new definitions are added.

   DATES: February 4, 1991.

 *****************************************************************************






 Sec. 23.1325   Static pressure system.

   (a) Each instrument provided with static pressure case connections must be
 so vented that the influence of airplane speed, the opening and closing of
 windows, airflow variations, moisture, or other foreign matter will least
 affect the accuracy of the instruments except as noted in paragraph (b)(3) of
 this section.
   (b) If a static pressure system is necessary for the functioning of
 instruments, systems, or devices, it must comply with the provisions of
 paragraphs (b) (1) through (3) of this section.
   (1) The design and installation of a static pressure system must be such
 that--
   (i) Positive drainage of moisture is provided;
   (ii) Chafing of the tubing, and excessive distortion or restriction at
 bends in the tubing, is avoided; and
   (iii) The materials used are durable, suitable for the purpose intended,
 and protected against corrosion.
   (2) A proof test must be conducted to demonstrate the integrity of the
 static pressure system in the following manner:
   (i) Unpressurized airplanes. Evacuate the static pressure system to a
 pressure differential of approximately 1 inch of mercury or to a reading on
 the altimeter, 1,000 feet above the aircraft elevation at the time of the
 test. Without additional pumping for a period of 1 minute, the loss of
 indicated altitude must not exceed 100 feet on the altimeter.
   (ii) Pressurized airplanes. Evacuate the static pressure system until a
 pressure differential equivalent to the maximum cabin pressure differential
 for which the airplane is type certificated is achieved. Without additional
 pumping for a period of 1 minute, the loss of indicated altitude must not
 exceed 2 percent of the equivalent altitude of the maximum cabin differential
 pressure or 100 feet, whichever is greater.
   (3) If a static pressure system is provided for any instrument, device, or
 system required by the operating rules of this chapter, each static pressure
 port must be designed or located in such a manner that the correlation
 between air pressure in the static pressure system and true ambient
 atmospheric static pressure is not altered when the airplane encounters icing
 conditions. An antiicing means or an alternate source of static pressure may
 be used in showing compliance with this requirement. If the reading of the
 altimeter, when on the alternate static pressure system differs from the
 reading of the altimeter when on the primary static system by more than 50
 feet, a correction card must be provided for the alternate static system.
   (c) Except as provided in paragraph (d) of this section, if the static
 pressure system incorporates both a primary and an alternate static pressure
 source, the means for selecting one or the other source must be designed so
 that--
   (1) When either source is selected, the other is blocked off; and
   (2) Both sources cannot be blocked off simultaneously.
   (d) For unpressurized airplanes, paragraph (c)(1) of this section does not
 apply if it can be demonstrated that the static pressure system calibration,
 when either static pressure source is selected, is not changed by the other
 static pressure source being open or blocked.
   (e) Each system must be designed and installed so that the error in
 indicated pressure altitude, at sea level, with a standard atmosphere,
 excluding instrument calibration error, does not result in an error of more
 than +30 feet per 100 knots speed for the appropriate configuration in the
 speed range between 1.3 Vs0 with flaps extended and 1.8 Vs1 with flaps
 retracted. However, the error need not be less than +30 feet.
   (f) For commuter category airplanes, the altimeter system calibration,
 required by paragraph (e) of this section, must be shown in the Airplane
 Flight Manual.
   (g) For airplanes prohibited from flight in instrument meteorological
 conditions, in accordance with Sec. 23.1559(b) of this part, paragraph (b)(3)
 of this section does not apply.

 [Amdt. 23-1, 30 FR 8261, June 29, 1965, as amended by Amdt. 23-6, 32 FR 7586,
 May 24, 1967; 32 FR 13505, Sept. 27, 1967; 32 FR 13714, Sept. 30, 1967; Amdt.
 23-20, 42 FR 36968, July 18, 1977; Amdt. 23-34, 52 FR 1834, Jan. 15, 1987;
 Amdt. 23-42, 56 FR 354, Jan. 3, 1991]

 *****************************************************************************


 56 FR 344, No. 2, Jan. 3, 1991

   SUMMARY: This final rule upgrades the airworthiness standards for normal,
 utility, acrobatic, and commuter category airplanes. This amendment provides
 airworthiness standards for advancements in technology being incorporated in
 current designs, permits type certification of spin resistant airplanes, and
 reduces the regulatory burden in showing compliance with some of the
 requirements for the design and type certification of small airplanes. These
 new and amended airworthiness standards also result in the need for new
 definitions. As a result, new definitions are added.

   DATES: February 4, 1991.

 *****************************************************************************






 Sec. 23.1327   Magnetic direction indicator.

   (a) Except as provided in paragraph (b) of this section--
   (1) Each magnetic direction indicator must be installed so that its
 accuracy is not excessively affected by the airplane's vibration or magnetic
 fields; and
   (2) The compensated installation may not have a deviation in level flight,
 greater than ten degrees on any heading.
   (b) A magnetic nonstabilized direction indicator may deviate more than ten
 degrees due to the operation of electrically powered systems such as
 electrically heated windshields if either a magnetic stabilized direction
 indicator, which does not have a deviation in level flight greater than ten
 degrees on any heading, or a gyroscopic direction indicator, is installed.
 Deviations of a magnetic nonstabilized direction indicator of more than 10
 degrees must be placarded in accordance with Sec. 23.1547(e).

 [Amdt. 23-20, 42 FR 36969, July 18, 1977]






 Sec. 23.1329   Automatic pilot system.

   If an automatic pilot system is installed, it must meet the following:
   (a) Each system must be designed so that the automatic pilot can--
   (1) Be quickly and positively disengaged by the pilots to prevent it from
 interfering with their control of the airplane; or
   (2) Be sufficiently overpowered by one pilot to let him control the
 airplane.
   (b) If the provisions of paragraph (a)(1) of this section are applied, the
 quick release (emergency) control must be located on the control wheel (both
 control wheels if the airplane can be operated from either pilot seat) on the
 side opposite the throttles, or on the stick control, such that it can be
 operated without moving the hand from its normal position on the control.
   (c) Unless there is automatic synchronization, each system must have a
 means to readily indicate to the pilot the alignment of the actuating device
 in relation to the control system it operates.
   (d) Each manually operated control for the system operation must be readily
 accessible to the pilot. Each control must operate in the same plane and
 sense of motion as specified in Sec. 23.779 for cockpit controls. The
 direction of motion must be plainly indicated on or near each control.
   (e) Each system must be designed and adjusted so that, within the range of
 adjustment available to the pilot, it cannot produce hazardous loads on the
 airplane or create hazardous deviations in the flight path, under any flight
 condition appropriate to its use, either during normal operation or in the
 event of a malfunction, assuming that corrective action begins within a
 reasonable period of time.
   (f) Each system must be designed so that a single malfunction will not
 produce a hardover signal in more than one control axis. If the automatic
 pilot integrates signals from auxiliary controls or furnishes signals for
 operation of other equipment, positive interlocks and sequencing of
 engagement to prevent improper operation are required.
   (g) There must be protection against adverse interaction of integrated
 components, resulting from a malfunction.
   (h) If the automatic pilot system can be coupled to airborne navigation
 equipment, means must be provided to indicate to the flight crew the current
 mode of operation. Selector switch position is not acceptable as a means of
 indication.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. 23-23, 43 FR 50593, Oct. 30, 1978; Amdt. 23-43, 58 FR 18976,
 Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1331   Instruments using a power source.

   For each instrument that uses a power source, the following apply:
   (a) Each instrument must have an integral visual power annunciator or
 separate power indicator to indicate when power is not adequate to sustain
 proper instrument performance. If a separate indicator is used, it must be
 located so that the pilot using the instruments can monitor the indicator
 with minimum head and eye movement. The power must be sensed at or near the
 point where it enters the instrument. For electric and vacuum/pressure
 instruments, the power is considered to be adequate when the voltage or the
 vacuum/pressure, respectively, is within approved limits.
   (b) The installation and power supply systems must be designed so that--
   (1) The failure of one instrument will not interfere with the proper supply
 of energy to the remaining instrument; and
   (2) The failure of the energy supply from one source will not interfere
 with the proper supply of energy from any other source.
   (c) There must be at least two independent sources of power (not driven by
 the same engine on multiengine airplanes), and a manual or an automatic means
 to select each power source.

 [Amdt. 23-43, 58 FR 18976, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1335   Flight director systems.

   If a flight director system is installed, means must be provided to
 indicate to the flight crew its current mode of operation. Selector switch
 position is not acceptable as a means of indication.

 [Amdt. 23-20, 42 FR 36969, July 18, 1977]






 Sec. 23.1337   Powerplant instruments.

   (a) Instruments and instrument lines.
   (1) Each powerplant and auxiliary power unit instrument line must meet the
 requirements of Sec. 23.993.
   (2) Each line carrying flammable fluids under pressure must--
   (i) Have restricting orifices or other safety devices at the source of
 pressure to prevent the escape of excessive fluid if the line fails; and
   (ii) Be installed and located so that the escape of fluids would not create
 a hazard.
   (3) Each powerplant and auxiliary power unit instrument that utilizes
 flammable fluids must be installed and located so that the escape of fluid
 would not create a hazard.
   (b) Fuel quantity indicator. There must be a means to indicate to the
 flight crewmembers the quantity of fuel in each tank during flight. An
 indicator, calibrated in either gallons or pounds, and clearly marked to
 indicate which scale is being used, may be used. In addition--
   (1) Each fuel quantity indicator must be calibrated to read "zero" during
 level flight when the quantity of fuel remaining in the tank is equal to the
 unusable fuel supply determined under Sec. 23.959;
   (2) Each exposed sight gauge used as a fuel quantity indicator must be
 protected against damage;
   (3) Each sight gauge that forms a trap in which water can collect and
 freeze must have means to allow drainage on the ground;
   (4) Tanks with interconnected outlets and airspaces may be considered as
 one tank and need not have separate indicators; and
   (5) No fuel quantity indicator is required for an auxiliary tank that is
 used only to transfer fuel to other tanks if the relative size of the tank,
 the rate of fuel transfer, and operating instructions are adequate to--
   (i) Guard against overflow; and
   (ii) Give the flight crewmembers prompt warning if transfer is not
 proceeding as planned.
   (c) Fuel flowmeter system. If a fuel flowmeter system is installed, each
 metering component must have a means to by-pass the fuel supply if
 malfunctioning of that component severely restricts fuel flow.
   (d) Oil quantity indicator. There must be a means to indicate the quantity
 of oil in each tank--
   (1) On the ground (such as by a stick gauge); and
   (2) In flight, to the flight crew members, if there is an oil transfer
 system or a reserve oil supply system.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13096, Aug. 13, 1969; Amdt. 23-18, 42 FR 15042, Mar. 17, 1977; Amdt. 23-43,
 58 FR 18976, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






                       Electrical Systems and Equipment






 Sec. 23.1351   General.

   (a) Electrical system capacity. Each electrical system must be adequate for
 the intended use. In addition--
   (1) Electric power sources, their transmission cables, and their associated
 control and protective devices, must be able to furnish the required power at
 the proper voltage to each load circuit essential for safe operation; and
   (2) Compliance with paragraph (a)(1) of this section must be shown as
 follows--
   (i) For normal, utility, and acrobatic category airplanes, by an electrical
 load analysis or by electrical measurements that account for the electrical
 loads applied to the electrical system in probable combinations and for
 probable durations; and
   (ii) For commuter category airplanes, by an electrical load analysis that
 accounts for the electrical loads applied to the electrical system in
 probable combinations and for probable durations.
   (b) Function. For each electrical system, the following apply:
   (1) Each system, when installed, must be--
   (i) Free from hazards in itself, in its method of operation, and in its
 effects on other parts of the airplane;
   (ii) Protected from fuel, oil, water, other detrimental substances, and
 mechanical damage; and
   (iii) So designed that the risk of electrical shock to crew, passengers,
 and ground personnel is reduced to a minimum.
   (2) Electric power sources must function properly when connected in
 combination or independently, except alternators installed in normal,
 utility, and acrobatic category airplanes, may depend on a battery for
 initial excitation or for stabilization.
   (3) No failure or malfunction of any electric power source may impair the
 ability of any remaining source to supply load circuits essential for safe
 operation, except the operation of an alternator that depends on a battery
 for initial excitation or for stabilization may be stopped by failure of that
 battery in normal, utility, and acrobatic category airplanes.
   (4) Each electric power source control must allow the independent operation
 of each source, except in normal, utility and acrobatic category airplanes,
 controls associated with alternators which depend on a battery for initial
 excitation or for stabilization need not break the connection between the
 alternator and its battery.
   (5) In addition, for commuter category airplanes, the following apply:
   (i) Each system must be designed so that essential load circuits can be
 supplied in the event of reasonably probable faults or open circuits
 including faults in heavy current carrying cables;
   (ii) A means must be accessible in flight to the flight crewmembers for the
 individual and collective disconnection of the electrical power sources from
 the system;
   (iii) The system must be designed so that voltage and frequency, if
 applicable, at the terminals of all essential load equipment can be
 maintained within the limits for which the equipment is designed during any
 probable operating conditions;
   (iv) If two independent sources of electrical power for particular
 equipment or systems are required, their electrical energy supply must be
 ensured by means such as duplicate electrical equipment, throwover switching,
 or multichannel or loop circuits separately routed; and
   (v) For the purpose of complying with paragraph (b)(5) of this section, the
 distribution system includes the distribution busses, their associated
 feeders, and each control and protective device.
   (c) Generating System. There must be at least one generator/alternator if
 the electrical system supplies power to load circuits essential for safe
 operation. In addition--
   (1) Each generator/alternator must be able to deliver its continuous rated
 power, or such power as is limited by its regulation system.
   (2) Generator/alternator voltage control equipment must be able to
 dependably regulate the generator/alternator output within rated limits.
   (3) Means must be provided to disconnect each generator/alternator from the
 battery and other generators/alternators when enough reverse current exists
 that might damage the generator/alternator, or will adversely affect the
 airplane electrical system.
   (4) There must be a means to give immediate warning to the flight crew of a
 failure of any generator/alternator.
   (5) Each generator/alternator must have an overvoltage control designed and
 installed to prevent damage to the electrical system, or to equipment
 supplied by the electrical system that could result if that generator/
 alternator were to develop an overvoltage condition.
   (d) Instruments. A means must exist to indicate to appropriate flight
 crewmembers the electric power system quantities essential for safe
 operation.
   (1) For normal, utility, and acrobatic category airplanes with direct
 current systems, an ammeter that can be switched into each generator feeder
 may be used and, if only one generator exists, the ammeter may be in the
 battery feeder.
   (2) For commuter category airplanes, the essential electric power system
 quantities include the voltage and current supplied by each generator.
   (e) Fire resistance. Electrical equipment must be so designed and installed
 that in the event of a fire in the engine compartment, during which the
 surface of the firewall adjacent to the fire is heated to 2,000 deg. F for 5
 minutes or to a lesser temperature substantiated by the applicant, the
 equipment essential to continued safe operation and located behind the
 firewall will function satisfactorily and will not create an additional fire
 hazard.
   (f) External power. If provisions are made for connecting external power to
 the airplane, and that external power can be electrically connected to
 equipment other than that used for engine starting, means must be provided to
 ensure that no external power supply having a reverse polarity, or a reverse
 phase sequence, can supply power to the airplane's electrical system.
   (g) It must be shown by analysis, tests, or both, that the airplane can be
 operated safely in VFR conditions, for a period of not less than five
 minutes, with the normal electrical power (electrical power sources excluding
 the battery and any other standby electrical sources) inoperative, with
 critical type fuel (from the standpoint of flameout and restart capability),
 and with the airplane initially at the maximum certificated altitude. Parts
 of the electrical system may remain on if--
   (1) A single malfunction, including a wire bundle or junction box fire,
 cannot result in loss of the part turned off and the part turned on; and
   (2) The parts turned on are electrically and mechanically isolated from the
 parts turned off.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13096, Aug. 13, 1969; Amdt. 23-14, 38 FR 31824, Nov. 19, 1973; Amdt. 23-17,
 41 FR 55465, Dec. 20, 1976; Amdt. 23-20, 42 FR 36969, July 18, 1977; Amdt.
 23-34, 52 FR 1834, Jan. 15, 1987; 52 FR 34745, Sept. 14, 1987; Amdt. 23-43,
 58 FR 18976, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1353   Storage battery design and installation.

   (a) Each storage battery must be designed and installed as prescribed in
 this section.
   (b) Safe cell temperatures and pressures must be maintained during any
 probable charging and discharging condition. No uncontrolled increase in cell
 temperature may result when the battery is recharged (after previous complete
 discharge)--
   (1) At maximum regulated voltage or power;
   (2) During a flight of maximum duration; and
   (3) Under the most adverse cooling condition likely to occur in service.
   (c) Compliance with paragraph (b) of this section must be shown by tests
 unless experience with similar batteries and installations has shown that
 maintaining safe cell temperatures and pressures presents no problem.
   (d) No explosive or toxic gases emitted by any battery in normal operation,
 or as the result of any probable malfunction in the charging system or
 battery installation, may accumulate in hazardous quantities within the
 airplane.
   (e) No corrosive fluids or gases that may escape from the battery may
 damage surrounding structures or adjacent essential equipment.
   (f) Each nickel cadmium battery installation capable of being used to start
 an engine or auxiliary power unit must have provisions to prevent any
 hazardous effect on structure or essential systems that may be caused by the
 maximum amount of heat the battery can generate during a short circuit of the
 battery or of its individual cells.
   (g) Nickel cadmium battery installations capable of being used to start an
 engine or auxiliary power unit must have--
   (1) A system to control the charging rate of the battery automatically so
 as to prevent battery overheating;
   (2) A battery temperature sensing and over-temperature warning system with
 a means for disconnecting the battery from its charging source in the event
 of an over-temperature condition; or
   (3) A battery failure sensing and warning system with a means for
 disconnecting the battery from its charging source in the event of battery
 failure.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. 23-20, 42 FR 36969, July 18, 1977; Amdt. 23-21, 43 FR 2319,
 Jan. 16, 1978]






 Sec. 23.1357   Circuit protective devices.

   (a) Protective devices, such as fuses or circuit breakers, must be
 installed in all electrical circuits other than--
   (1) Main circuits of starter motors used during starting only; and
   (2) Circuits in which no hazard is presented by their omission.
   (b) A protective device for a circuit essential to flight safety may not be
 used to protect any other circuit.
   (c) Each resettable circuit protective device ("trip free" device in which
 the tripping mechanism cannot be overridden by the operating control) must be
 designed so that--
   (1) A manual operation is required to restore service after tripping; and
   (2) If an overload or circuit fault exists, the device will open the
 circuit regardless of the position of the operating control.
   (d) If the ability to reset a circuit breaker or replace a fuse is
 essential to safety in flight, that circuit breaker or fuse must be so
 located and identified that it can be readily reset or replaced in flight.
   (e) For fuses identified as replaceable in flight--
   (1) There must be one spare of each rating or 50 percent spare fuses of
 each rating, whichever is greater; and
   (2) The spare fuse(s) must be readily accessible to any required pilot.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. 23-20, 42 FR 36969, July 18, 1977; Amdt. 23-43, 58 FR
 18976, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1361   Master switch arrangement.

   (a) There must be a master switch arrangement to allow ready disconnection
 of each electric power source from power distribution systems, except as
 provided in paragraph (b) of this section. The point of disconnection must be
 adjacent to the sources controlled by the switch arrangement. If separate
 switches are incorporated into the master switch arrangement, a means must be
 provided for the switch arrangement to be operated by one hand with a single
 movement.
   (b) Load circuits may be connected so that they remain energized when the
 master switch is open, if the circuits are isolated, or physically shielded,
 to prevent their igniting flammable fluids or vapors that might be liberated
 by the leakage or rupture of any flammable fluid system; and
   (1) The circuits are required for continued operation of the engine; or
   (2) The circuits are protected by circuit protective devices with a rating
 of five amperes or less adjacent to the electric power source.
   (3) In addition, two or more circuits installed in accordance with the
 requirements of paragraph (b)(2) of this section must not be used to supply a
 load of more than five amperes.
   (c) The master switch or its controls must be so installed that the switch
 is easily discernible and accessible to a crewmember in flight.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. 23-20, 42 FR 36969, July 18, 1977; Amdt. 23-43, 58 FR 18977,
 Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1365   Electric cables and equipment.

   (a) Each electric connecting cable must be of adequate capacity.
   (b) Each cable and associated equipment that would overheat in the event of
 circuit overload or fault must be at least flame resistant and may not emit
 dangerous quantities of toxic fumes.
   (c) Main power cables (including generator cables) in the fuselage must be
 designed to allow a reasonable degree of deformation and stretching without
 failure and must--
   (1) Be separated from flammable fluid lines; or
   (2) Be shrouded by means of electrically insulated flexible conduit, or
 equivalent, which is in addition to the normal cable insulation.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-14, 38 FR
 31824, Nov. 19, 1973; Amdt. 23-43, 58 FR 18977, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1367   Switches.

   Each switch must be--
   (a) Able to carry its rated current;
   (b) Constructed with enough distance or insulating material between current
 carrying parts and the housing so that vibration in flight will not cause
 shorting;
   (c) Accessible to appropriate flight crewmembers; and
   (d) Labeled as to operation and the circuit controlled.






                                    Lights






 Sec. 23.1381   Instrument lights.

   The instrument lights must--
   (a) Make each instrument and control easily readable and discernible;
   (b) Be installed so that their direct rays, and rays reflected from the
 windshield or other surface, are shielded from the pilot's eyes; and
   (c) Have enough distance or insulating material between current carrying
 parts and the housing so that vibration in flight will not cause shorting.

 A cabin dome light is not an instrument light.






 Sec. 23.1383   Landing lights.

   (a) Each installed landing light must be acceptable.
   (b) Each landing light must be installed so that--
   (1) No dangerous glare is visible to the pilot;
   (2) The pilot is not seriously affected by halation; and
   (3) It provides enough light for night landing.






 Sec. 23.1385   Position light system installation.

   (a) General. Each part of each position light system must meet the
 applicable requirements of this section and each system as a whole must meet
 the requirements of Secs. 23.1387 through 23.1397.
   (b) Left and right position lights. Left and right position lights must
 consist of a red and a green light spaced laterally as far apart as
 practicable and installed on the airplane such that, with the airplane in the
 normal flying position, the red light is on the left side and the green light
 is on the right side.
   (c) Rear position light. The rear position light must be a white light
 mounted as far aft as practicable on the tail or on each wing tip.
   (d) Light covers and color filters.  Each light cover or color filter must
 be at least flame resistant and may not change color or shape or lose any
 appreciable light transmission during normal use.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-17, 41 FR
 55465, Dec. 20, 1976; Amdt. 23-43, 58 FR 18977, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1387   Position light system dihedral angles.

   (a) Except as provided in paragraph (e) of this section, each position
 light must, as installed, show unbroken light within the dihedral angles
 described in this section.
   (b) Dihedral angle L (left) is formed by two intersecting vertical planes,
 the first parallel to the longitudinal axis of the airplane, and the other at
 110 degrees to the left of the first, as viewed when looking forward along
 the longitudinal axis.
   (c) Dihedral angle R (right) is formed by two intersecting vertical planes,
 the first parallel to the longitudinal axis of the airplane, and the other at
 110 degrees to the right of the first, as viewed when looking forward along
 the longitudinal axis.
   (d) Dihedral angle A (aft) is formed by two intersecting vertical planes
 making angles of 70 degrees to the right and to the left, respectively, to a
 vertical plane passing through the longitudinal axis, as viewed when looking
 aft along the longitudinal axis.
   (e) If the rear position light, when mounted as far aft as practicable in
 accordance with Sec. 23.1385(c), cannot show unbroken light within dihedral
 angle A (as defined in paragraph (d) of this section), a solid angle or
 angles of obstructed visibility totaling not more than 0.04 steradians is
 allowable within that dihedral angle, if such solid angle is within a cone
 whose apex is at the rear position light and whose elements make an angle of
 30 deg. with a vertical line passing through the rear position light.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. 23-12, 36 FR 21278, Nov. 5, 1971; Amdt. 23-43, 58 FR 18977,
 Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1389   Position light distribution and intensities.

   (a) General. The intensities prescribed in this section must be provided by
 new equipment with each light cover and color filter in place. Intensities
 must be determined with the light source operating at a steady value equal to
 the average luminous output of the source at the normal operating voltage of
 the airplane. The light distribution and intensity of each position light
 must meet the requirements of paragraph (b) of this section.
   (b) Position lights.  The light distribution and intensities of position
 lights must be expressed in terms of minimum intensities in the horizontal
 plane, minimum intensities in any vertical plane, and maximum intensities in
 overlapping beams, within dihedral angles L, R, and A, and must meet the
 following requirements:
   (1) Intensities in the horizontal plane.  Each intensity in the horizontal
 plane (the plane containing the longitudinal axis of the airplane and
 perpendicular to the plane of symmetry of the airplane) must equal or exceed
 the values in Sec. 23.1391.
   (2) Intensities in any vertical plane.  Each intensity in any vertical
 plane (the plane perpendicular to the horizontal plane) must equal or exceed
 the appropriate value in Sec. 23.1393, where I is the minimum intensity
 prescribed in Sec. 23.1391 for the corresponding angles in the horizontal
 plane.
   (3) Intensities in overlaps between adjacent signals. No intensity in any
 overlap between adjacent signals may exceed the values in Sec. 23.1395,
 except that higher intensities in overlaps may be used with main beam
 intensities substantially greater than the minima specified in Secs. 23.1391
 and 23.1393, if the overlap intensities in relation to the main beam
 intensities do not adversely affect signal clarity. When the peak intensity
 of the left and right position lights is more than 100 candles, the maximum
 overlap intensities between them may exceed the values in Sec. 23.1395 if the
 overlap intensity in Area A is not more than 10 percent of peak position
 light intensity and the overlap intensity in Area B is not more than 2.5
 percent of peak position light intensity.
   (c) Rear position light installation. A single rear position light may be
 installed in a position displaced laterally from the plane of symmetry of an
 airplane if--
   (1) The axis of the maximum cone of illumination is parallel to the flight
 path in level flight; and
   (2) There is no obstruction aft of the light and between planes 70 degrees
 to the right and left of the axis of maximum illumination.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-43, 58 FR
 18977, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1391   Minimum intensities in the horizontal plane of position
     lights.

   Each position light intensity must equal or exceed the applicable values in
 the following table:

                            Angle from right or left
   Dihedral angle (light     of longitudinal axis,
         included)          measured from dead ahead    Intensity (candles)

  L and R (red and green)   0 deg. to 10 deg.         40
                             10 deg. to 20 deg.        30
                             20 deg. to 110 deg.       5
  A (rear white)            110 deg. to 180 deg.      20

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-43, 58 FR
 18977, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1393   Minimum intensities in any vertical plane of position lights.

   Each position light intensity must equal or exceed the applicable values in
 the following table:

                          Angle above or
                            below the       Intensity,
                         horizontal plane       l

                        0 deg.                    1.00
                        0 deg. to 5 deg.          0.90
                        5 deg. to 10 deg.         0.80
                        10 deg. to 15 deg.        0.70
                        15 deg. to 20 deg.        0.50
                        20 deg. to 30 deg.        0.30
                        30 deg. to 40 deg.        0.10
                        40 deg. to 90 deg.        0.05

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-43, 58 FR
 18977, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1395   Maximum intensities in overlapping beams of position lights.

   No position light intensity may exceed the applicable values in the
 following equal or exceed the applicable values in Sec. 23.1389(b)(3):

                                               Maximum intensity

                                               Area A     Area B
                         Overlaps             (candles)  (candles)

              Green in dihedral angle L              10          1
              Red in dihedral angle R                10          1
              Green in dihedral angle A               5          1
              Red in dihedral angle A                 5          1
              Rear white in dihedral angle L          5          1
              Rear white in dihedral angle R          5          1

 Where--
   (a) Area A includes all directions in the adjacent dihedral angle that pass
 through the light source and intersect the common boundary plane at more than
 10 degrees but less than 20 degrees; and
   (b) Area B includes all directions in the adjacent dihedral angle that pass
 through the light source and intersect the common boundary plane at more than
 20 degrees.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-43, 58 FR
 18977, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1397   Color specifications.

   Each position light color must have the applicable International Commission
 on Illumination chromaticity coordinates as follows:
   (a) Aviation red--

   "y" is not greater than 0.335; and
   "z" is not greater than 0.002.

   (b) Aviation green--

   "x" is not greater than 0.440-0.320 y;
   "x" is not greater than y -0.170; and
   "y" is not less than 0.390-0.170 x.

   (c) Aviation white--

   "x" is not less than 0.300 and not greater than 0.540;
   "y" is not less than "x -0.040" or "y0 -0.010," whichever is the smaller;
 and
   "y" is not greater than "x+0.020" nor "0.636-0.400 x ";

   Where "y0" is the "y" coordinate of the Planckian radiator for the value of
 "x"  considered.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, amended by Amdt. 23-11, 36 FR
 12971, July 10, 1971]






 Sec. 23.1399   Riding light.

   (a) Each riding (anchor) light required for a seaplane or amphibian, must
 be installed so that it can--
   (1) Show a white light for at least two miles at night under clear
 atmospheric conditions; and
   (2) Show the maximum unbroken light practicable when the airplane is moored
 or drifting on the water.
   (b) Externally hung lights may be used.






 Sec. 23.1401   Anticollision light system.

   (a) General. If certification for night operation is requested, the
 airplane must have an anticollision light system that--
   (1) Consists of one or more approved anticollision lights located so that
 their light will not impair the flight crewmembers' vision or detract from
 the conspicuity of the position lights; and
   (2) Meets the requirements of paragraphs (b) through (f) of this section.
   (b) Field of coverage. The system must consist of enough lights to
 illuminate the vital areas around the airplane, considering the physical
 configuration and flight characteristics of the airplane. The field of
 coverage must extend in each direction within at least 75 degrees above and
 75 degrees below the horizontal plane of the airplane, except that there may
 be solid angles of obstructed visibility totaling not more than 0.5
 steradians.
   (c) Flashing characteristics. The arrangement of the system, that is, the
 number of light sources, beam width, speed of rotation, and other
 characteristics, must give an effective flash frequency of not less than 40,
 nor more than 100, cycles per minute. The effective flash frequency is the
 frequency at which the airplane's complete anticollision light system is
 observed from a distance, and applies to each sector of light including any
 overlaps that exist when the system consists of more than one light source.
 In overlaps, flash frequencies may exceed 100, but not 180, cycles per
 minute.
   (d) Color. Each anticollision light must be either aviation red or aviation
 white and must meet the applicable requirements of Sec. 23.1397.
   (e) Light intensity. The minimum light intensities in any vertical plane,
 measured with the red filter (if used) and expressed in terms of "effective"
 intensities, must meet the requirements of paragraph (f) of this section. The
 following relation must be assumed:

                             t2
                           INTEGRAL       I(t)dt
                             t1
                          Ie =
                          --------------
                          0.2+(t2-t1)

 where:

 Ie =effective intensity (candles).
 I(t) =instantaneous intensity as a function of time.
 t2-t1 =flash time interval (seconds).

 Normally, the maximum value of effective intensity is obtained when t2 and t1
 are chosen so that the effective intensity is equal to the instantaneous
 intensity at t2 and t1.
   (f) Minimum effective intensities for anticollision lights. Each
 anticollision light effective intensity must equal or exceed the applicable
 values in the following table.

                           Angle above or    Effective
                             below the       intensity
                          horizontal plane   (candles)

                         0 deg. to 5 deg.          400
                         5 deg. to 10 deg.         240
                         10 deg. to 20 deg.         80
                         20 deg. to 30 deg.         40
                         30 deg. to 75 deg.         20

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-11, 36 FR
 12972, July 10, 1971; Amdt. 23-20, 42 FR 36969, July 18, 1977]






                               Safety Equipment






 Sec. 23.1411   General.

   (a) Required safety equipment to be used by the flight crew in an
 emergency, such as automatic liferaft releases, must be readily accessible.
   (b) Stowage provisions for required safety equipment must be furnished and
 must--
   (1) Be arranged so that the equipment is directly accessible and its
 location is obvious; and
   (2) Protect the safety equipment from damage caused by being subjected to
 the inertia loads resulting from the ultimate static load factors specified
 in Sec. 23.561(b)(3) of this part.

 [Amdt. 23-17, 41 FR 55465, Dec. 20, 1976, as amended by Amdt. 23-36, 53 FR
 30815, Aug. 15, 1988]






 Sec. 23.1413   Safety belts and harnesses.

   Each safety belt and shoulder harness must be equipped with a metal-to-
 metal latching device.

 [Amdt. 23-36, 53 FR 30815, Aug. 15, 1988]






 Sec. 23.1415   Ditching equipment.

   (a) Emergency flotation and signaling equipment required by any operating
 rule in this chapter must be installed so that it is readily available to the
 crew and passengers.
   (b) Each raft and each life preserver must be approved.
   (c) Each raft released automatically or by the pilot must be attached to
 the airplane by a line to keep it alongside the airplane. This line must be
 weak enough to break before submerging the empty raft to which it is
 attached.
   (d) Each signaling device required by any operating rule in this chapter,
 must be accessible, function satisfactorily, and must be free of any hazard
 in its operation.






 Sec. 23.1416   Pneumatic de-icer boot system.

   If certification with ice protection provisions is desired and a pneumatic
 de-icer boot system is installed--
   (a) The system must meet the requirements specified in Sec. 23.1419.
   (b) The system and its components must be designed to perform their
 intended function under any normal system operating temperature or pressure,
 and
   (c) Means to indicate to the flight crew that the pneumatic de-icer boot
 system is receiving adequate pressure and is functioning normally must be
 provided.

 [Amdt. 23-23, 43 FR 50593, Oct. 30, 1978]






 Sec. 23.1419   Ice protection.

   If certification with ice protection provisions is desired, compliance with
 the requirements of this section and other applicable sections of this part
 must be shown:
   (a) An analysis must be performed to establish, on the basis of the
 airplane's operational needs, the adequacy of the ice protection system for
 the various components of the airplane. In addition, tests of the ice
 protection system must be conducted to demonstrate that the airplane is
 capable of operating safely in continuous maximum and intermittent maximum
 icing conditions, as described in appendix C of part 25 of this chapter. As
 used in this section, "Capable of operating safely," means that airplane
 performance, controllability, maneuverability, and stability must not be less
 than that required in part 23, subpart B.
   (b) Except as provided by paragraph (c) of this section, in addition to the
 analysis and physical evaluation prescribed in paragraph (a) of this section,
 the effectiveness of the ice protection system and its components must be
 shown by flight tests of the airplane or its components in measured natural
 atmospheric icing conditions and by one or more of the following tests, as
 found necessary to determine the adequacy of the ice protection system--
   (1) Laboratory dry air or simulated icing tests, or a combination of both,
 of the components or models of the components.
   (2) Flight dry air tests of the ice protection system as a whole, or its
 individual components.
   (3) Flight test of the airplane or its components in measured simulated
 icing conditions.
   (c) If certification with ice protection has been accomplished on prior
 type certificated airplanes whose designs include components that are
 thermodynamically and aerodynamically equivalent to those used on a new
 airplane design, certification of these equivalent components may be
 accomplished by reference to previously accomplished tests, required in Sec.
 23.1419 (a) and (b), provided that the applicant accounts for any differences
 in installation of these components.
   (d) A means must be identified or provided for determining the formation of
 ice on the critical parts of the airplane. Adequate lighting must be provided
 for the use of this means during night operation. Also, when monitoring of
 the external surfaces of the airplane by the flight crew is required for
 operation of the ice protection equipment, external lighting must be provided
 that is adequate to enable the monitoring to be done at night. Any
 illumination that is used must be of a type that will not cause glare or
 reflection that would handicap crewmembers in the performance of their
 duties. The Airplane Flight Manual or other approved manual material must
 describe the means of determining ice formation and must contain information
 for the safe operation of the airplane in icing conditions.

 [Amdt. 23-43, 58 FR 18977, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






                            Miscellaneous Equipment






 Sec. 23.1431   Electronic equipment.

   (a) In showing compliance with Sec. 23.1309(b) (1) and (2) with respect to
 radio and electronic equipment and their installations, critical
 environmental conditions must be considered.
   (b) Radio and electronic equipment, controls, and wiring must be installed
 so that operation of any unit or system of units will not adversely affect
 the simultaneous operation of any other radio or electronic unit, or system
 of units, required by this chapter.

 [Amdt. 23-43, 58 FR 18977, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1435   Hydraulic systems.

   (a) Design. Each hydraulic system must be designed as follows:
   (1) Each hydraulic system and its elements must withstand, without
 yielding, the structural loads expected in addition to hydraulic loads.
   (2) A means to indicate the pressure in each hydraulic system which
 supplies two or more primary functions must be provided to the flight crew.
   (3) There must be means to ensure that the pressure, including transient
 (surge) pressure, in any part of the system will not exceed the safe limit
 above design operating pressure and to prevent excessive pressure resulting
 from fluid volumetric changes in all lines which are likely to remain closed
 long enough for such changes to occur.
   (4) The minimum design burst pressure must be 2.5 times the operating
 pressure.
   (b) Tests. Each system must be substantiated by proof pressure tests. When
 proof tested, no part of any system may fail, malfunction, or experience a
 permanent set. The proof load of each system must be at least 1.5 times the
 maximum operating pressure of that system.
   (c) Accumulators. A hydraulic accumulator or pressurized reservoir must not
 be installed on the engine side of any firewall unless--
   (1) It is an integral part of an engine or propeller, or
   (2) It is a nonpressurized reservoir and the total capacity of all such
 nonpressurized reservoirs is one quart or less.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13096, Aug. 13, 1969; Amdt. 23-14, 38 FR 31824, Nov. 19, 1973; Amdt. 23-43,
 58 FR 18978, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1437   Accessories for multiengine airplanes.

   For multiengine airplanes, engine-driven accessories essential to safe
 operation must be distributed among two or more engines so that the failure
 of any one engine will not impair safe operation through the malfunctioning
 of these accessories.






 Sec. 23.1438   Pressurization and pneumatic systems.

   (a) Pressurization system elements must be burst pressure tested to 2.0
 times, and proof pressure tested to 1.5 times, the maximum normal operating
 pressure.
   (b) Pneumatic system elements must be burst pressure tested to 3.0 times,
 and proof pressure tested to 1.5 times, the maximum normal operating
 pressure.
   (c) An analysis, or a combination of analysis and test, may be substituted
 for any test required by paragraph (a) or (b) of this section if the
 Administrator finds it equivalent to the required test.

 [Amdt. 23-20, 42 FR 36969, July 18, 1977]






 Sec. 23.1441   Oxygen equipment and supply.

   (a) If certification with supplemental oxygen equipment is requested, or
 the airplane is approved for operations at or above altitudes where oxygen is
 required to be used by the operating rules, oxygen equipment must be provided
 that meets the requirements of this section and Secs. 23.1443 through
 23.1449. Portable oxygen equipment may be used to meet the requirements of
 this part if the portable equipment is shown to comply with the applicable
 requirements, is identified in the airplane type design, and its stowage
 provisions are found to be in compliance with the requirements of Sec.
 23.561.
   (b) The oxygen system must be free from hazards in itself, in its method of
 operation, and its effect upon other components.
   (c) There must be a means to allow the crew to readily determine, during
 the flight, the quantity of oxygen available in each source of supply.
   (d) Each required flight crewmember must be provided with--
   (1) Demand oxygen equipment if the airplane is to be certificated for
 operation above 25,000 feet.
   (2) Pressure demand oxygen equipment if the airplane is to be certificated
 for operation above 40,000 feet.
   (e) There must be a means, readily available to the crew in flight, to turn
 on and to shut off the oxygen supply at the high pressure source. This
 shutoff requirement does not apply to chemical oxygen generators.

 [Amdt. 23-9, 35 FR 6386, Apr. 21, 1970, as amended by Amdt. 23-43, 58 FR
 18978, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1443   Minimum mass flow of supplemental oxygen.

   (a) If continuous flow oxygen equipment is installed, an applicant must
 show compliance with the requirements of either paragraphs (a)(1) and (a)(2)
 or paragraph (a)(3) of this section:
   (1) For each passenger, the minimum mass flow of supplemental oxygen
 required at various cabin pressure altitudes may not be less than the flow
 required to maintain, during inspiration and while using the oxygen equipment
 (including masks) provided, the following mean tracheal oxygen partial
 pressures;
   (i) At cabin pressure altitudes above 10,000 feet up to and including
 18,500 feet, a mean tracheal oxygen partial pressure of 100 mm. Hg when
 breathing 15 liters per minute, Body Temperature, Pressure, Saturated (BTPS)
 and with a tidal volume of 700 cc. with a constant time interval between
 respirations.
   (ii) At cabin pressure altitudes above 18,500 feet up to and including
 40,000 feet, a mean tracheal oxygen partial pressure of 83.8 mm. Hg when
 breathing 30 liters per minute, BTPS, and with a tidal volume of 1,100 cc.
 with a constant time interval between respirations.
   (2) For each flight crewmember, the minimum mass flow may not be less than
 the flow required to maintain, during inspiration, a mean tracheal oxygen
 partial pressure of 149 mm. Hg when breathing 15 liters per minute, BTPS, and
 with a maximum tidal volume of 700 cc. with a constant time interval between
 respirations.
   (3) The minimum mass flow of supplemental oxygen supplied for each user
 must be at a rate not less than that shown in the following figure for each
 altitude up to and including the maximum operating altitude of the airplane.

       [INSERT: Line graph plotting oxygen mass flow in liters per minute
             against cabin pressure altitude in thousands of feet]

   (b) If demand equipment is installed for use by flight crewmembers, the
 minimum mass flow of supplemental oxygen required for each flight crewmember
 may not be less than the flow required to maintain, during inspiration, a
 mean tracheal oxygen partial pressure of 122 mm. Hg up to and including a
 cabin pressure altitude of 35,000 feet, and 95 percent oxygen between cabin
 pressure altitudes of 35,000 and 40,000 feet, when breathing 20 liters per
 minute BTPS. In addition, there must be means to allow the crew to use
 undiluted oxygen at their discretion.
   (c) If first-aid oxygen equipment is installed, the minimum mass flow of
 oxygen to each user may not be less than 4 liters per minute, STPD. However,
 there may be a means to decrease this flow to not less than 2 liters per
 minute, STPD, at any cabin altitude. The quantity of oxygen required is based
 upon an average flow rate of 3 liters per minute per person for whom first-
 aid oxygen is required.
   (d) As used in this section:
   (1) BTPS means Body Temperature, and Pressure, Saturated (which is, 37
 deg.C, and the ambient pressure to which the body is exposed, minus 47 mm.
 Hg, which is the tracheal pressure displaced by water vapor pressure when the
 breathed air becomes saturated with water vapor at 37 deg.C).
   (2) STPD means Standard, Temperature, and Pressure, Dry (which is, 0 deg.C
 at 760 mm. Hg with no water vapor).

 [Amdt. 23-43, 58 FR 18978, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1445   Oxygen distribution system.

   (a) Except for flexible lines from oxygen outlets to the dispensing units,
 or where shown to be otherwise suitable to the installation, nonmetallic
 tubing must not be used for any oxygen line that is normally pressurized
 during flight.
   (b) Nonmetallic oxygen distribution lines must not be routed where they may
 be subjected to elevated temperatures, electrical arcing, and released
 flammable fluids that might result from any probable failure.

 [Amdt. 23-43, 58 FR 18978, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1447   Equipment standards for oxygen dispensing units.

   If oxygen dispensing units are installed, the following apply:
   (a) There must be an individual dispensing unit for each occupant for whom
 supplemental oxygen is to be supplied. Each dispensing unit must:
   (1) Provide for effective utilization of the oxygen being delivered to the
 unit.
   (2) Be capable of being readily placed into position on the face of the
 user.
   (3) Be equipped with a suitable means to retain the unit in position on the
 face.
   (b) If certification for operation up to and including 18,000 feet (MSL) is
 requested, each oxygen dispensing unit must:
   (1) Cover the nose and mouth of the user; or
   (2) Be a nasal cannula, in which case one oxygen dispensing unit covering
 both the nose and mouth of the user must be available. In addition, each
 nasal cannula or its connecting tubing must have permanently affixed--
   (i) A visible warning against smoking while in use;
   (ii) An illustration of the correct method of donning; and
   (iii) A visible warning against use with nasal obstructions or head colds
 with resultant nasal congestion.
   (c) If certification for operation above 18,000 feet (MSL) is requested,
 each oxygen dispensing unit must cover the nose and mouth of the user.
   (d) For a pressurized airplane designed to operate at flight altitudes
 above 25,000 feet (MSL), an oxygen dispensing unit connected to an oxygen
 supply terminal must be immediately available to each occupant, wherever
 seated.
   (e) If certification for operation above 30,000 feet is requested, the
 dispensing units must meet the following requirements:
   (1) The dispensing units for passengers must be automatically presented to
 each occupant before the cabin pressure altitude exceeds 15,000 feet.
   (2) The dispensing units for flight crewmembers must be automatically
 presented to each flight crewmember before the cabin pressure altitude
 exceeds 15,000 feet, or the units must be of the quick-donning type,
 connected to an oxygen supply terminal that is immediately available to
 flight crewmembers at their station.
   (f) If an automatic dispensing unit (hose and mask, or other unit) system
 is installed, the crew must be provided with a manual means to make the
 dispensing units immediately available in the event of failure of the
 automatic system.

 [Amdt. 23-9, 35 FR 6387, Apr. 21, 1970, as amended by Amdt. 23-20, 42 FR
 36969, July 18, 1977; Amdt. 23-30, 49 FR 7340, Feb. 28, 1984; Amdt. 23-43,
 58 FR 18978, Apr. 9, 1993]

 *****************************************************************************


 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






 Sec. 23.1449   Means for determining use of oxygen.

   There must be a means to allow the crew to determine whether oxygen is
 being delivered to the dispensing equipment.

 [Amdt. 23-9, 35 FR 6387, Apr. 21, 1970]






 Sec. 23.1450   Chemical oxygen generators.

   (a) For the purpose of this section, a chemical oxygen generator is defined
 as a device which produces oxygen by chemical reaction.
   (b) Each chemical oxygen generator must be designed and installed in
 accordance with the following requirements:
   (1) Surface temperature developed by the generator during operation may not
 create a hazard to the airplane or to its occupants.
   (2) Means must be provided to relieve any internal pressure that may be
 hazardous.
   (c) In addition to meeting the requirements in paragraph (b) of this
 section, each portable chemical oxygen generator that is capable of sustained
 operation by successive replacement of a generator element must be placarded
 to show--
   (1) The rate of oxygen flow, in liters per minute;
   (2) The duration of oxygen flow, in minutes, for the replaceable generator
 element; and
   (3) A warning that the replaceable generator element may be hot, unless the
 element construction is such that the surface temperature cannot exceed 100
 deg. F.

 [Amdt. 23-20, 42 FR 36969, July 18, 1977]






 Sec. 23.1457  Cockpit voice recorders.

   (a) Each cockpit voice recorder required by the operating rules of this
 chapter must be approved and must be installed so that it will record the
 following:
   (1) Voice communications transmitted from or received in the airplane by
 radio.
   (2) Voice communications of flight crewmembers on the flight deck.
   (3) Voice communications of flight crewmembers on the flight deck, using
 the airplane's interphone system.
   (4) Voice or audio signals identifying navigation or approach aids
 introduced into a headset or speaker.
   (5) Voice communications of flight crewmembers using the passenger
 loudspeaker system, if there is such a system and if the fourth channel is
 available in accordance with the requirements of paragraph (c)(4)(ii) of this
 section.
   (b) The recording requirements of paragraph (a)(2) of this section must be
 met by installing a cockpit-mounted area microphone, located in the best
 position for recording voice communications originating at the first and
 second pilot stations and voice communications of other crewmembers on the
 flight deck when directed to those stations. The microphone must be so
 located and, if necessary, the preamplifiers and filters of the recorder must
 be so adjusted or supplemented, so that the intelligibility of the recorded
 communications is as high as practicable when recorded under flight cockpit
 noise conditions and played back. Repeated aural or visual playback of the
 record may be used in evaluating intelligibility.
   (c) Each cockpit voice recorder must be installed so that the part of the
 communication or audio signals specified in paragraph (a) of this section
 obtained from each of the following sources is recorded on a separate
 channel:
   (1) For the first channel, from each boom, mask, or handheld microphone,
 headset, or speaker used at the first pilot station.
   (2) For the second channel from each boom, mask, or handheld microphone,
 headset, or speaker used at the second pilot station.
   (3) For the third channel--from the cockpit-mounted area microphone.
   (4) For the fourth channel from:
   (i) Each boom, mask, or handheld microphone, headset, or speaker used at
 the station for the third and fourth crewmembers.
   (ii) If the stations specified in paragraph (c)(4)(i) of this section are
 not required or if the signal at such a station is picked up by another
 channel, each microphone on the flight deck that is used with the passenger
 loudspeaker system, if its signals are not picked up by another channel.
   (5) And that as far as is practicable all sounds received by the microphone
 listed in paragraphs (c) (1), (2), and (4) of this section must be recorded
 without interruption irrespective of the position of the interphone-
 transmitter key switch. The design shall ensure that sidetone for the flight
 crew is produced only when the interphone, public address system, or radio
 transmitters are in use.
   (d) Each cockpit voice recorder must be installed so that:
   (1) It receives its electric power from the bus that provides the maximum
 reliability for operation of the cockpit voice recorder without jeopardizing
 service to essential or emergency loads.
   (2) There is an automatic means to simultaneously stop the recorder and
 prevent each erasure feature from functioning, within 10 minutes after crash
 impact; and
   (3) There is an aural or visual means for preflight checking of the
 recorder for proper operation.
   (e) The record container must be located and mounted to minimize the
 probability of rupture of the container as a result of crash impact and
 consequent heat damage to the record from fire. In meeting this requirement,
 the record container must be as far aft as practicable, but may not be where
 aft mounted engines may crush the container during impact. However, it need
 not be outside of the pressurized compartment.
   (f) If the cockpit voice recorder has a bulk erasure device, the
 installation must be designed to minimize the probability of inadvertent
 operation and actuation of the device during crash impact.
   (g) Each recorder container must:
   (1) Be either bright orange or bright yellow;
   (2) Have reflective tape affixed to its external surface to facilitate its
 location under water; and
   (3) Have an underwater locating device, when required by the operating
 rules of this chapter, on or adjacent to the container which is secured in
 such manner that they are not likely to be separated during crash impact.

 [Amdt. 23-35, 53 FR 26142, July 11, 1988]






 Sec. 23.1459   Flight recorders.

   (a) Each flight recorder required by the operating rules of this chapter
 must be installed so that:
   (1) It is supplied with airspeed, altitude, and directional data obtained
 from sources that meet the accuracy requirements of Secs. 23.1323, 23.1325,
 and 23.1327, as appropriate;
   (2) The vertical acceleration sensor is rigidly attached, and located
 longitudinally either within the approved center of gravity limits of the
 airplane, or at a distance forward or aft of these limits that does not
 exceed 25 percent of the airplane's mean aerodynamic chord;
   (3) It receives its electrical power power from the bus that provides the
 maximum reliability for operation of the flight recorder without jeopardizing
 service to essential or emergency loads;
   (4) There is an aural or visual means for preflight checking of the
 recorder for proper recording of data in the storage medium.
   (5) Except for recorders powered solely by the engine-driven electrical
 generator system, there is an automatic means to simultaneously stop a
 recorder that has a data erasure feature and prevent each erasure feature
 from functioning, within 10 minutes after crash impact; and
   (b) Each nonejectable record container must be located and mounted so as to
 minimize the probability of container rupture resulting from crash impact and
 subsequent damage to the record from fire. In meeting this requirement the
 record container must be located as far aft as practicable, but need not be
 aft of the pressurized compartment, and may not be where aft-mounted engines
 may crush the container upon impact.
   (c) A correlation must be established between the flight recorder readings
 of airspeed, altitude, and heading and the corresponding readings (taking
 into account correction factors) of the first pilot's instruments. The
 correlation must cover the airspeed range over which the airplane is to be
 operated, the range of altitude to which the airplane is limited, and 360
 degrees of heading. Correlation may be established on the ground as
 appropriate.
   (d) Each recorder container must:
   (1) Be either bright orange or bright yellow;
   (2) Have reflective tape affixed to its external surface to facilitate its
 location under water; and
   (3) Have an underwater locating device, when required by the operating
 rules of this chapter, on or adjacent to the container which is secured in
 such a manner that they are not likely to be separated during crash impact.
   (e) Any novel or unique design or operational characteristics of the
 aircraft shall be evaluated to determine if any dedicated parameters must be
 recorded on flight recorders in addition to or in place of existing
 requirements.

 [Amdt. 23-35, 53 FR 26143, July 11, 1988]






 Sec. 23.1461   Equipment containing high energy rotors.

   (a) Equipment containing high energy rotors must meet paragraph (b), (c),
 or (d) of this section.
   (b) High energy rotors contained in equipment must be able to withstand
 damage caused by malfunctions, vibration, abnormal speeds, and abnormal
 temperatures. In addition--
   (1) Auxiliary rotor cases must be able to contain damage caused by the
 failure of high energy rotor blades; and
   (2) Equipment control devices, systems, and instrumentation must reasonably
 ensure that no operating limitations affecting the integrity of high energy
 rotors will be exceeded in service.
   (c) It must be shown by test that equipment containing high energy rotors
 can contain any failure of a high energy rotor that occurs at the highest
 speed obtainable with the normal speed control devices inoperative.
   (d) Equipment containing high energy rotors must be located where rotor
 failure will neither endanger the occupants nor adversely affect continued
 safe flight.

 [Amdt. 23-20, 42 FR 36969, July 18, 1977]






               Subpart G--Operating Limitations and Information






 Sec. 23.1501  General.

   (a) Each operating limitation specified in Secs. 23.1505 through 23.1527
 and other limitations and information necessary for safe operation must be
 established.
   (b) The operating limitations and other information necessary for safe
 operation must be made available to the crewmembers as prescribed in Secs.
 23.1541 through 23.1589.

 [Amdt. 23-21, 43 FR 2319, Jan. 16, 1978]






 Sec. 23.1505  Airspeed limitations.

   (a) The never-exceed speed VNE must be established so that it is--
   (1) Not less than 0.9 times the minimum value of VD allowed under Sec.
 23.335; and
   (2) Not more than the lesser of--
   (i) 0.9 VD established under Sec. 23.335; or
   (ii) 0.9 times the maximum speed shown under Sec. 23.251.
   (b) The maximum structural cruising speed VNO must be established so that
 it is--
   (1) Not less than the minimum value of VC allowed under Sec. 23.335; and
   (2) Not more than the lesser of--
   (i) VC established under Sec. 23.335; or
   (ii) 0.89 VNE established under paragraph (a) of this section.
   (c) Paragraphs (a) and (b) of this section do not apply to turbine
 airplanes or to airplanes for which a design diving speed VD/MD is
 established under Sec. 23.335(b)(4). For those airplanes, a maximum operating
 limit speed (VMO/MMO-airspeed or Mach number, whichever is critical at a
 particular altitude) must be established as a speed that may not be
 deliberately exceeded in any regime of flight (climb, cruise, or descent)
 unless a higher speed is authorized for flight test or pilot training
 operations. VMO/MMO must be established so that it is not greater than the
 design cruising speed VC/MC and so that it is sufficiently below VD/MD and
 the maximum speed shown under Sec. 23.251 to make it highly improbable that
 the latter speeds will be inadvertently exceeded in operations. The speed
 margin between VMO/MMO and VD/MD or the maximum speed shown under Sec. 23.251
 may not be less than the speed margin established between VC/MC  and VD/MD
 under Sec. 23.335(b), or the speed margin found necessary in the flight test
 conducted under Sec. 23.253.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13096, Aug. 13, 1969]






 Sec. 23.1507   Operating maneuvering speed.

   The maximum maneuvering speed Vo, must be established as an operating
 limitation. Vo is a selected speed that is not greater than Vs[square root]n
 established in Sec. 23.335(c).

 [Amdt. No. 23-45, 58 FR 42165, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.1511   Flap extended speed.

   (a) The flap extended speed VFE must be established so that it is--
   (1) Not less than the minimum value of VF allowed in Secs. 23.345 and
 23.457; and
   (2) Not more than the lesser of--
   (i) VF established under Sec. 23.345; or
   (ii) VF established under Sec. 23.457.
   (b) Additional combinations of flap setting, airspeed, and engine power may
 be established if the structure has been proven for the corresponding design
 conditions.






 Sec. 23.1513   Minimum control speed.

   The minimum control speed VMC, determined under Sec. 23.149, must be
 established as an operating limitation.






 Sec. 23.1519   Weight and center of gravity.

   The weight and center of gravity limitations determined under Sec. 23.23
 must be established as operating limitations.






 Sec. 23.1521   Powerplant limitations.

   (a) General. The powerplant limitations prescribed in this section must be
 established so that they do not exceed the corresponding limits for which the
 engines or propellers are type certificated. In addition, other powerplant
 limitations used in determining compliance with this part must be
 established.
   (b) Takeoff operation. The powerplant takeoff operation must be limited
 by--
   (1) The maximum rotational speed (rpm);
   (2) The maximum allowable manifold pressure (for reciprocating engines);
   (3) The maximum allowable gas temperature (for turbine engines);
   (4) The time limit for the use of the power or thrust corresponding to the
 limitations established in paragraphs (b) (1) through (3) of this section;
 and
   (5) If the time limit in paragraph (b) (4) of this section exceeds two
 minutes, the maximum allowable cylinder head (as applicable), liquid coolant,
 and oil temperatures.
   (c) Continuous operation. The continuous operation must be limited by--
   (1) The maximum rotational speed;
   (2) The maximum allowable manifold pressure (for reciprocating engines);
   (3) The maximum allowable gas temperature (for turbine engines); and
   (4) The maximum allowable cylinder head, oil, and liquid coolant
 temperatures.
   (d) Fuel grade or designation. The minimum fuel grade (for reciprocating
 engines), or fuel designation (for turbine engines), must be established so
 that it is not less than that required for the operation of the engines
 within the limitations in paragraphs (b) and (c) of this section.
   (e) Ambient temperature. For turbine engines, ambient temperature
 limitations (including limitations for winterization installations if
 applicable) must be established as the maximum ambient atmospheric
 temperature at which compliance with the cooling provisions of Secs. 23.1041
 through 23.1047 is shown.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. 23-21, 43 FR 2319, Jan. 16, 1978; Amdt. 23-45, 58 FR 42165,
 Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.1522   Auxiliary power unit limitations.

   If an auxiliary power unit is installed, the limitations established for
 the auxiliary power must be specified in the operating limitations for the
 airplane.

 [Amdt. No. 23-45, 58 FR 42166, Aug. 6, 1993; 58 FR 51970, Oct. 5, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.1523  Minimum flight crew.

   The minimum flight crew must be established so that it is sufficient for
 safe operation considering--
   (a) The workload on individual crewmembers and, in addition for commuter
 category airplanes, each crewmember workload determination must consider the
 following:
   (1) Flight path control,
   (2) Collision avoidance,
   (3) Navigation,
   (4) Communications,
   (5) Operation and monitoring of all essential airplane systems,
   (6) Command decisions, and
   (7) The accessibility and ease of operation of necessary controls by the
 appropriate crewmember during all normal and emergency operations when at the
 crewmember flight station;
   (b) The accessibility and ease of operation of necessary controls by the
 appropriate crewmember; and
   (c) The kinds of operation authorized under Sec. 23.1525.

 [Amdt. 23-21, 43 FR 2319, Jan. 16, 1978, as amended by Amdt. 23-34, 52 FR
 1834, Jan. 15, 1987]






 Sec. 23.1524   Maximum passenger seating configuration.

   The maximum passenger seating configuration must be established.

 [Amdt. 23-10, 36 FR 2864, Feb. 11, 1971]






 Sec. 23.1525   Kinds of operation.

   The kinds of operation authorized (e.g. VFR, IFR, day or night) and the
 meteorological conditions (e.g. icing) to which the operation of the airplane
 is limited of from which it is prohibited, mut be established appropriate to
 the installed equipment.

 [Amdt. No. 23-45, 58 FR 42166, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.1527   Maximum operating altitude.

   (a) The maximum altitude up to which operation is allowed, as limited by
 flight, structural, powerplant, functional or equipment characteristics, must
 be established.
   (b) A maximum operating altitude limitation of not more than 25,000 feet
 must be established for pressurized airplanes unless compliance with Sec.
 23.775(e) is shown.

 [Amdt. 23-45, 45 FR 42166, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.1529   Instructions for Continued Airworthiness.

   The applicant must prepare Instructions for Continued Airworthiness in
 accordance with Appendix G to this part that are acceptable to the
 Administrator. The instructions may be incomplete at type certification if a
 program exists to ensure their completion prior to delivery of the first
 airplane or issuance of a standard certificate of airworthiness, whichever
 occurs later.

 [Amdt. 23-26, 45 FR 60171, Sept. 11, 1980]






                             Markings And Placards






 Sec. 23.1541  General.

   (a) The airplane must contain--
   (1) The markings and placards specified in Secs. 23.1545 through 23.1567;
 and
   (2) Any additional information, instrument markings, and placards required
 for the safe operation if it has unusual design, operating, or handling
 characteristics.
   (b) Each marking and placard prescribed in paragraph (a) of this section--
   (1) Must be displayed in a conspicuous place; and
   (2) May not be easily erased, disfigured, or obscured.
   (c) For airplanes which are to be certificated in more than one category--
   (1) The applicant must select one category upon which the placards and
 markings are to be based; and
   (2) The placards and marking information for all categories in which the
 airplane is to be certificated must be furnished in the Airplane Flight
 Manual.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. 23-21, 43 FR 2319, Jan. 16, 1978]






 Sec. 23.1543   Instrument markings: general.

   For each instrument--
   (a) When markings are on the cover glass of the instrument, there must be
 means to maintain the correct alignment of the glass cover with the face of
 the dial; and
   (b) Each arc and line must be wide enough and located to be clearly visible
 to the pilot.






 Sec. 23.1545   Airspeed indicator.

   (a) Each airspeed indicator must be marked as specified in paragraph (b) of
 this section, with the marks located at the corresponding indicated
 airspeeds.
   (b) The following markings must be made:
   (1) For the never-exceed speed VNE, a radial red line.
   (2) For the caution range, a yellow arc extending from the red line
 specified in paragraph (b)(1) of this section to the upper limit of the green
 arc specified in paragraph (b)(3) of this section.
   (3) For the normal operating range, a green arc with the lower limit at VS1
 with maximum weight and with landing gear and wing flaps retracted, and the
 upper limit at the maximum structural cruising speed VNO established under
 Sec. 23.1505(b).
   (4) For the flap operating range, a white arc with the lower limit at VS0
 at the maximum weight, and the upper limit at the flaps-extended speed VFE
 established under Sec. 23.1511.
   (5) For the one-engine-inoperative best rate of climb speed, Vy, a blue
 sector extending from the Vy speed at sea level to the Vy speed at--
   (i) An altitude of 5,000 feet, if the one-engine-inoperative best rate of
 climb at that altitude is less than 100 feet per minute, or
   (ii) The highest 1,000-foot altitude (at or above 5,000 feet) at which the
 one-engine-inoperative best rate of climb is 100 feet per minute or more.

 Each side of the sector must be labeled to show the altitude for the
 corresponding Vy.
   (6) For the minimum control speed (one-engine-inoperative), Vmc', a red
 radial line.
   (c) If VNE or VNO vary with altitude, there must be means to indicate to
 the pilot the appropriate limitations throughout the operating altitude
 range.
   (d) Paragraphs (b)(1) through (b)(3) and paragraph (c) of this section do
 not apply to aircraft for which a maximum operating speed VMO/MMO is
 established under Sec. 23.1505(c). For those aircraft there must either be a
 maximum allowable airspeed indication showing the variation of VMO/MMO with
 altitude or compressibility limitations (as appropriate), or a radial red
 line marking for VMO/MMO must be made at lowest value of VMO/MMO established
 for any altitude up to the maximum operating altitude for the airplane.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-3, 30 FR
 14240, Nov. 13, 1965; Amdt. 23-7, 34 FR 13097, Aug. 13, 1969; Amdt. 23-23, 43
 FR 50593, Oct. 30, 1978]






 Sec. 23.1547   Magnetic direction indicator.

   (a) A placard meeting the requirements of this section must be installed on
 or near the magnetic direction indicator.
   (b) The placard must show the calibration of the instrument in level flight
 with the engines operating.
   (c) The placard must state whether the calibration was made with radio
 receivers on or off.
   (d) Each calibration reading must be in terms of magnetic headings in not
 more than 30 degree increments.
   (e) If a magnetic nonstabilized direction indicator can have a deviation of
 more than 10 degrees caused by the operation of electrical equipment, the
 placard must state which electrical loads, or combination of loads, would
 cause a deviation of more than 10 degrees when turned on.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. 23-20, 42 FR 36969, July 18, 1977]






 Sec. 23.1549   Powerplant and auxiliary power unit instruments.

   For each required powerplant and auxiliary power unit instrument, as
 appropriate to the type of instruments--
   (a) Each maximum and, if applicable, minimum safe operating limit must be
 marked with a red radial or a red line;
   (b) Each normal operating range must be marked with a green arc or green
 line, not extending beyond the maximum and minimum safe limits;
   (c) Each takeoff and precautionary range must be marked with a yellow arc
 or a yellow line; and
   (d) Each engine, auxiliary power unit, or propeller range that is
 restricted because of excessive vibration stresses must be marked with red
 arcs or red lines.

 [Amdt. 23-12, 41 FR 55466, Dec. 20, 1976, as amended by Amdt. 23-28, 47 FR
 13315, Mar. 29, 1982; Amdt. No. 23-45, 58 FR 42166, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.1551   Oil quantity indicator.

   Each oil quantity indicator must be marked in sufficient increments to
 indicate readily and accurately the quantity of oil.






 Sec. 23.1553   Fuel quantity indicator.

   If the unusable fuel supply for any tank exceeds one gallon, or five
 percent of the tank capacity, whichever is greater, a red arc must be marked
 on its indicator extending from the calibrated zero reading to the lowest
 reading obtainable in level flight.






 Sec. 23.1555   Control markings.

   (a) Each cockpit control, other than primary flight controls and simple
 push button type starter switches, must be plainly marked as to its function
 and method of operation.
   (b) Each secondary control must be suitably marked.
   (c) For powerplant fuel controls--
   (1) Each fuel tank selector control must be marked to indicate the position
 corresponding to each tank and to each existing cross feed position;
   (2) If safe operation requires the use of any tanks in a specific sequence,
 that sequence must be marked on or near the selector for those tanks;
   (3) The conditions under which the full amount of usable fuel in any
 restricted usage fuel tank can safely be used must be stated on a placard
 adjacent to the selector valve for that tank; and
   (4) Each valve control for any engine of a multiengine airplane must be
 marked to indicate the position corresponding to each engine controlled.
   (d) Usable fuel capacity must be marked as follows:
   (1) For fuel systems having no selector controls, the usable fuel capacity
 of the system must be indicated at the fuel quantity indicator.
   (2) For fuel systems having selector controls, the usable fuel capacity
 available at each selector control position must be indicated near the
 selector control.
   (e) For accessory, auxiliary, and emergency controls--
   (1) If retractable landing gear is used, the indicator required by Sec.
 23.729 must be marked so that the pilot can, at any time, ascertain that the
 wheels are secured in the extreme positions; and
   (2) Each emergency control must be red and must be marked as to method of
 operation.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. 23-21, 43 FR 2319, Jan. 16, 1978]






 Sec. 23.1557   Miscellaneous markings and placards.

   (a) Baggage and cargo compartments, and ballast location. Each baggage and
 cargo compartment, and each ballast location, must have a placard stating any
 limitations on contents, including weight, that are necessary under the
 loading requirements.
   (b) Seats. If the maximum allowable weight to be carried in a seat is less
 than 170 pounds, a placard stating the lesser weight must be permanently
 attached to the seat structure.
   (c) Fuel, oil, and coolant filler openings. The following apply:
   (1) Fuel filter openings must be marked at or near the filler cover with--
   (i) For reciprocating engine-powered airplanes--
   (A) The word "Avgas"; and
   (B) The minimum fuel grade.
   (ii) For turbine engine-powered airplanes--
   (A) The words "Jet Fuel"; and
   (B) The permissible fuel designations, or references to the Airplane Flight
 Manual (AFM) for permissible fuel designations.
   (iii) For pressure fueling systems, the maximum permissible fueling supply
 pressure and the maximum permissible defueling pressure.
   (2) Oil filler openings must be marked at or near the filler cover with the
 word "Oil" and the permissible oil designations, or references to the
 Airplane Flight Manual (AFM) for permissible oil designations.
   (3) Coolant filler openings must be marked at or near the filler cover with
 the word "Coolant".
   (d) Emergency exit placards. Each placard and operating control for each
 emergency exit must be red. A placard must be near each emergency exit
 control and must clearly indicate the location of that exit and its method of
 operation.
   (e) The system voltage of each direct current installation must be clearly
 marked adjacent to its exernal power connection.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; as amended by Amdt. 23-21, 42 FR
 15042, Mar. 17, 1977; Amdt. 23-23, 43 FR 50594, Oct. 30, 1978; Amdt. No.
 23-45, 58 FR 42166, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.1559   Operating limitations placard.

   (a) There must be a placard in clear view of the pilot stating--
   (1) For airplanes certificated in one category:

   The markings and placards installed in this airplane contain operating
 limitations which must be complied with when operating this airplane in the
 -------------------- category. (Insert category.) Other operating limitations
 which must be complied with when operating this airplane in this category are
 contained in the Airplane Flight Manual.

   (2) For airplanes certificated in more than one category:

   The markings and placards installed in this airplane contain operating
 limitations which must be complied with when operating this airplane in the
 ---------- category. (Insert category.) Other operating limitations which
 must be complied with when operating this airplane in this category or in the
 ---------- category are contained in the Airplane Flight Manual. (Insert
 category or categories.)

   (b) There must be a placard in clear view of the pilot that specifies the
 kind of operations (such as VFR, IFR, day, or night) and the meteorological
 conditions (such as icing conditions) to which the operation of the airplane
 is limited, or from which it is prohibited, by the equipment installed.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. 23-13, 37 FR 20023, Sept. 23, 1972; 37 FR 21320, Oct. 7,
 1972; Amdt. 23-21, 43 FR 2319, Jan. 16, 1978]






 Sec. 23.1561   Safety equipment.

   (a) Safety equipment must be plainly marked as to method of operation.
   (b) Stowage provisions for required safety equipment must be marked for the
 benefit of occupants.






 Sec. 23.1563   Airspeed placards.

   There must be an airspeed placard in clear view of the pilot and as close
 as practicable to the airspeed indicator. This placard must list--
   (a) The operating maneuvering speed, Vo; and
   (b) The maximum landing gear operating speed VLO.

 [Amdt. 23-7, 34 FR 13097, Aug. 13, 1969, as amended by Amdt. No. 23-45, 58
 FR 42166, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.1567   Flight maneuver placard.

   (a) For normal category airplanes, there must be a placard in front of and
 in clear view of the pilot stating: "No acrobatic maneuvers, including spins,
 approved."
   (b) For utility category airplanes, there must be--
   (1) A placard in clear view of the pilot stating: "Acrobatic maneuvers are
 limited to the following ------" (list approved maneuvers and the recommended
 entry speed for each); and
   (2) For those airplanes that do not meet the spin requirements for
 acrobatic category airplanes, an additional placard in clear view of the
 pilot stating: "Spins Prohibited."
   (c) For acrobatic category airplanes, there must be a placard in clear view
 of the pilot listing the approved acrobatic maneuvers and the recommended
 entry airspeed for each. If inverted flight maneuvers are not approved, the
 placard must bear a notation to this effect.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. 23-13, 37 FR 20023, Sept. 23, 1972; Amdt. 23-21, 43 FR 2319,
 Jan. 16, 1978]






              Airplane Flight Manual and Approved Manual Material






 Sec. 23.1581   General.

   (a) Furnishing information. An Airplane Flight Manual must be furnished
 with each airplane, and it must contain the following:
   (1) Information required by Secs. 23.1583 through 23.1589.
   (2) Other information that is necessary for safe operation because of
 design, operating, or handling characteristics.
   (b) Approved information. (1) Except as provided in paragraph (b)(2) of
 this section, each part of the Airplane Flight Manual containing information
 prescribed in Secs. 23.1583 through 23.1589 must be approved, segregated,
 identified and clearly distinguished from each unapproved part of that
 Airplane Flight Manual.
   (2) The requirements of paragraph (b)(1) of this section do not apply if
 the following is met:
   (i) Each part of the Airplane Flight Manual containing information
 prescribed in Sec. 23.1583 must be limited to such information, and must be
 approved, identified, and clearly distinguished from each other part of the
 Airplane Flight Manual.
   (ii) The information prescribed in Secs. 23.1585 through 23.1589 must be
 determined in accordance with the applicable requirements of this part and
 presented in its entirety in a manner acceptable to the Administrator.
   (3) Each page of the Airplane Flight Manual containing information
 prescribed in this section must be of a type that is not easily erased,
 disfigured, or misplaced, and is capable of being inserted in a manual
 provided by the applicant, or in a folder, or in any other permanent binder.
   (c) [Reserved]
   (d) Table of contents. Each Airplane Flight Manual must include a table of
 contents if the complexity of the manual indicates a need for it.
   (e) Provision must be made for stowing the Airplane Flight Manual in a
 suitable fixed container which is readily accessible to the pilot.
   (f) Revisions and amendments. Each Airplane Flight Manual (AFM) must
 contain a means for recording the incorporation of revisions and amendments.

 [Amdt. 23-21, 43 FR 2319, Jan. 16, 1978, as amended by Amdt. 23-34, 52 FR
 1834, Jan. 15, 1987; Amdt. No. 23-45, 58 FR 42166, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.1583  Operating limitations.

   Operating limitations determined during type certification must be stated,
 including the following:
   (a) Airspeed limitations. The following information must be furnished:
   (1) Information necessary for the marking of the airspeed limits on the
 indicator as required in Sec. 23.1545, and the significance of each of those
 limits and of the color coding used on the indicator.
   (2) The speeds VMC, VO, VLE, and VLO, if established, and their
 significance.
   (3) In addition, for commuter category airplanes--
   (i) The maximum operating limit speed, VMO/MMO and a statement that this
 speed may not be deliberately exceeded in any regime of flight (climb,
 cruise, or descent) unless a higher speed is authorized for flight test or
 pilot training;
   (ii) If an airspeed limitation is based upon compressibility effects, a
 statement to this effect and information as to any symptoms, the probable
 behavior of the airplane, and the recommended recovery procedures; and
   (iii) The airspeed limits must be shown in terms of VMO/MMO instead of VNO
 and VNE.
   (b) Powerplant limitations. The following information must be furnished:
   (1) Limitations required by Sec. 23.1521.
   (2) Explanation of the limitations, when appropriate.
   (3) Information necessary for marking the instruments required by Sec.
 23.1549 through Sec. 23.1553.
   (c) Weight. The airplane flight manual must include--
   (1) The maximum weight; and
   (2) The maximum landing weight, if the design landing weight selected by
 the applicant is less than the maximum weight.
   (3) In addition, for commuter category airplanes, the maximum takeoff
 weight for each altitude, ambient temperature, and required takeoff runway
 length within the range selected by the applicant may not exceed the weight
 at which--
   (i) The all-engine-operating distance determined under Sec. 23.59 or the
 accelerate-stop distance determined under Sec. 23.55, whichever is greater,
 is equal to the available runway length;
   (ii) The airplane complies with the one-engine-inoperative takeoff distance
 requirements of Sec. 23.59; and
   (iii) The airplane complies with the one-engine-inoperative takeoff and en
 route climb requirements of Secs. 23.57 and 23.67.
   (4) In addition, for commuter category airplanes, the maximum landing
 weight for each altitude, ambient temperature, and required landing runway
 length, within the range selected by the applicant. The maximum landing
 weights may not exceed:
   (i) The weight at which the landing distance is determined under Sec.
 23.75; or
   (ii) The weight at which compliance with Sec. 23.77 is shown.
   (d) Center of gravity. The established center of gravity limits must be
 furnished.
   (e) Maneuvers. The following authorized maneuvers, appropriate airspeed
 limitations, and unauthorized maneuvers must be furnished as prescribed in
 this section.
   (1) Normal category airplanes. For normal category airplanes, acrobatic
 maneuvers, including spins, are unauthorized. If the airplane has been shown
 to be "characteristically incapable of spinning" under Sec. 23.221(d), a
 statement to this effect must be entered. Other normal category airplanes
 must be placarded against spins.
   (2) Utility category airplanes. For utility category airplanes, authorized
 maneuvers shown in the type flight tests must be furnished, together with
 recommended entry speeds. No other maneuver is authorized. If the airplane
 has been shown to be "characteristically incapable of spinning" under Sec.
 23.221(d), a statement to this effect must be entered.
   (3) Acrobatic category airplanes. For acrobatic category airplanes, the
 approved flight maneuvers shown in the type flight tests must be included,
 together with recommended entry speeds. A placard listing the use of the
 controls required to recover from spinning maneuvers must be in the cockpit.
   (4) Commuter category airplanes. For commuter category airplanes, acrobatic
 maneuvers, including spins, are unauthorized.
   (f) Flight load factor. The positive limit load factors, in g's, must be
 furnished.
   (g) Flight crew. If a flight crew of more than one is required for safety,
 the number and functions of the minimum flight crew must be furnished.
   (h) Kinds of operation. A list of the kinds of operation to which the
 airplane is limited or from which it is prohibited under Sec. 23.1525, and
 also a list of installed equipment that affects any operating limitation and
 identification as to the equipment's required operational status for the
 kinds of operation for which approval has been given.
   (i)--(j) [Reserved]
   (k) Maximum operating altitude. The maximum altitude established under Sec.
 23.1527 must be furnished.
   (l) Maximum passenger seating configuration.  The maximum passenger seating
 configuration must be furnished.
   (m) Allowable lateral fuel loading. The maximum allowable lateral fuel
 loading differential must be furnished if less than the maximum possible.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13097, Aug. 13, 1969; Amdt. 23-10, 36 FR 2864, Feb. 11, 1971; Amdt. 23-21, 43
 FR 2320, Jan. 16, 1978; Amdt. 23-23, 43 FR 50594, Oct. 30, 1978; Amdt. 23-34,
 52 FR 1834, Jan. 15, 1987; Amdt. No. 23-45, 58 FR 42166, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.1585  Operating procedures.

   (a) For each airplane, information concerning normal, abnormal, and
 emergency procedures and other pertinent information necessary for safe
 operation and the achievement of the scheduled performance must be identified
 and segregated, including--
   (1) The maximum demonstrated values of crosswind velocity for takeoff and
 landing and procedures and information pertinent to operations in crosswinds;
   (2) The speeds, configurations, and procedures for making a normal takeoff
 and the subsequent climb;
   (3) Procedure for abandoning a takeoff due to engine failure or other
 cause;
   (4) The recommended climb speeds, and any variation with altitude;
   (5) An explanation of significant or unusual flight or ground handling
 characteristics of the airplane;
   (6) A recommended speed for flight in rough air. This speed must be chosen
 to protect against the occurrence, as a result of gusts, of structural damage
 to the airplane and loss of control (for example, stalling); and
   (7) For seaplanes and amphibians, water handling procedures and the
 demonstrated wave height.
   (b) For single-engine airplanes, the procedures, speeds, and configurations
 for a glide following an engine failure and subsequent forced landing.
   (c) For multiengine airplanes, the information must include--
   (1) Procedures and speeds for continuing a takeoff following failure of the
 critical engine and the conditions under which takeoff can be safely
 continued, or a warning against attempting to continue the takeoff;
   (2) Procedures, speeds, and configurations for continuing a climb following
 engine failure after takeoff or en route;
   (3) Procedures, speeds, and configurations for making an approach and
 landing with one engine inoperative;
   (4) Procedures, speeds, and configurations for making a go-around with one
 engine inoperative and the conditions under which the go-around can safely be
 executed, or a warning against attempting the go-around maneuver;
   (5) Procedures for restarting engines in flight, including the effects of
 altitude, must be set forth in the Airplane Flight Manual (AFM); and
   (6) The VSSE determined in Sec. 23.149.
   (d) For multiengine airplanes, information identifying each operating
 condition in which the fuel system independence prescribed in Sec. 23.953 is
 necessary for safety must be furnished, together with instructions for
 placing the fuel system in a configuration used to show compliance with that
 section.
   (e) For each airplane showing compliance with Sec. 23.1353 (g)(2) or
 (g)(3), the operating procedures for disconnecting the battery from its
 charging source must be furnished.
   (f) If the unusable fuel supply in any tank exceeds 5 percent of the tank
 capacity, or 1 gallon, whichever is greater, information must be furnished
 which indicates that when the fuel quantity indicator reads "zero" in level
 flight, any fuel remaining in the fuel tank cannot be used safely in flight.
   (g) Information on the total quantity of usable fuel for each fuel tank
 must be furnished.
   (h) In addition, for commuter category airplanes, the procedures for
 restarting turbine engines in flight, including the effects of altitude, must
 be set forth in the Airplane Flight Manual.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-3, 30 FR
 14240, Nov. 13, 1965; Amdt. 23-5, 32 FR 6912, May 5, 1967; Amdt. 23-7, 34 FR
 13097, Aug. 13, 1969; Amdt. 23-21, 43 FR 2320, Jan. 16, 1978; Amdt. 23-23, 43
 FR 50594, Oct. 30, 1978; Amdt. 23-34, 52 FR 1835, Jan. 15, 1987; Amdt. No.
 23-45, 58 FR 42166, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.1587  Performance information.

   The following information must be furnished:
   (a) For normal, utility, and acrobatic category airplanes:
   (1) The takeoff distance determined under Sec. 23.51 and the kind of runway
 surface used in the tests.
   (2) The climb gradient determined under Secs. 23.65 and 23.77, with the
 associated airspeed, power, and the airplane configuration.
   (3) The landing distance determined under Sec. 23.75.
   (4) The one engine inoperative en route climb/descent gradients determined
 under Sec. 23.67 for multiengine airplanes.
   (5) The calculated approximate effect on takeoff distance, landing
 distance, and climb performance for variations in--
   (i) Altitude from sea level to 10,000 feet in a standard atmosphere and
 cruise configuration; and
   (ii) Temperature, at those altitudes from 60 deg.F below standard to 40
 deg.F above standard.
   (b) For skiplanes, a statement of the approximate reduction in climb
 performance may be used instead of complete new data for the skiplane
 configuration if--
   (1) The landing gear is fixed in both the landplane and skiplane
 configurations;
   (2) The climb performance is not critical; and
   (3) The climb reduction in the skiplane configuration does not exceed 50
 feet per minute.
   (c) For each airplane:
   (1) Any loss of altitude more than 100 feet, or any pitch more than 30
 degrees below level flight attitude, occurring during the recovery part of
 maneuvers prescribed in Secs. 23.201(c) and 23.205, if applicable.
   (2) The stalling speed, VSO, at maximum weight.
   (3) The stalling speed, VS1, at maximum weight and with the landing gear
 and wing flaps retracted and the effect upon this stalling speed of angles of
 bank up to 60 degrees.
   (4) The speed used in showing compliance with the cooling and climb
 requirements of Secs. 23.1041 through 23.1047 if this speed is greater than
 the best rate of climb with one engine inoperative for multiengine airplanes
 and the maximum atmospheric temperature at which compliance with the cooling
 requirements has been shown.
   (d) Commuter category airplanes. In addition, for commuter category
 airplanes, the Airplane Flight Manual must contain at least the following
 performance information:
   (1) Sufficient information so that the takeoff weight limits specified in
 Sec. 23.1583 can be determined for all temperatures and altitudes within the
 operational limitations selected by the applicant;
   (2) The conditions under which the performance information was obtained
 including the airspeed at the 50-foot height used to determine the landing
 distance as required by Sec. 23.75;
   (3) The performance information (determined by extrapolation and computed
 for the range of weights between the maximum landing and maximum takeoff
 weights) for--
   (i) Climb in the landing configuration as determined by Sec. 23.77; and
   (ii) Landing distance as determined by Sec. 23.75;
   (4) Procedures information established in accordance with the limitations
 and other information for safe operation of the airplane in the form of
 recommended procedures;
   (5) An explanation of significant or unusual flight and ground handling
 characteristics of the airplane; and
   (6) Airspeed, as calibrated airspeed, corresponding to those established
 while showing compliance to Sec. 23.53, Takeoff speeds. The calibrated
 airspeed may be shown in units of indicated airspeed, and identified as
 indicated airspeed, provided that all pressure sensing and instrumentation
 errors, including the indicator, are accounted for in the flight manual
 data.

 [Amdt. 23-21, 43 FR 2320, Jan. 16, 1978, as amended by Amdt. 23-28, 47 FR
 13315, Mar. 29, 1982; Amdt. 23-34, 52 FR 1835, Jan. 15, 1987; Amdt. 23-39,
 55 FR 18575, May 2, 1990; Amdt. No. 23-45, 58 FR 42167, Aug. 6, 1993; 58 FR
 51970, Oct. 5, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************






 Sec. 23.1589  Loading information.

   The following loading information must be furnished:
   (a) The weight and location of each item of equipment that can be easily
 removed, relocated, or replaced and that is installed when the airplane was
 weighed under the requirement of Sec. 23.25.
   (b) Appropriate loading instr@ctions for each possible loading condition
 between the maximum and minimum weights determined under Sec. 23.25 that can
 result in a center of gravity beyond--
   (1) The extremes selected by the applicant;
   (2) The extremes within which the structure is proven; or
   (3) The extremes within which compliance with each functional requirement
 is shown.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. No. 23-45, 58 FR 42167, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************



 Appendix A to Part 23--Simplified Design Load Criteria for Conventional,
     Single-Engine Airplanes of 6,000 Pounds or Less Maximum Weight

 A23.1  General.

   (a) The design load criteria in this appendix are an approved equivalent of
 those in Secs. 23.321 through 23.459 of this subchapter for the certification
 of conventional, single-engine airplanes of 6,000 pounds or less maximum
 weight.
   (b) Unless otherwise stated, the nomenclature and symbols in this Appendix
 are the same as the corresponding nomenclature and symbols in Part 23.

 A23.3  Special symbols.

 n1 =Airplane Positive Maneuvering Limit Load Factor.
 n2 =Airplane Negative Maneuvering Limit Load Factor.
 n3 =Airplane Positive Gust Limit Load Factor at VC.
 n4 =Airplane Negative Gust Limit Load Factor at VC.
 nflap =Airplane Positive Limit Load Factor With Flaps Fully Extended at VF.

 * V     = Minimum Design Flap Speed =
    F min   11.0 <radical> n1 W/S   [kts]

 * V     = Mimimum Design Manuevering
    A min   Speed = 15.0 <radical> n1 W/S  [kts]

 * V     = Minimum Design Cruising Speed
    C min  = 17.0 <radical> n1 W/S    [kts]

 * V     = Minimum Design Dive Speed =
    D min   24.0 <radical> n1 W/S    [kts]

 A23.5  Certification in more than one category.

   The criteria in this appendix may be used for certification in the normal,
 utility, and acrobatic categories, or in any combination of these categories.
 If certification in more than one category is desired, the design category
 weights must be selected to make the term " n1W " constant for all categories
 or greater for one desired category than for others. The wings and control
 surfaces (including wing flaps and tabs) need only be investigated for the
 maximum value of " n1W ", or for the category corresponding to the maximum
 design weight, where " n1W " is constant. If the acrobatic category is
 selected, a special unsymmetrical flight load investigation in accordance
 with paragraphs A23.9(c)(2) and A23.11(c)(2) of this appendix must be
 completed. The wing, wing carrythrough, and the horizontal tail structures
 must be checked for this condition. The basic fuselage structure need only be
 investigated for the highest load factor design category selected. The local
 supporting structure for dead weight items need only be designed for the
 highest load factor imposed when the particular items are installed in the
 airplane. The engine mount, however, must be designed for a higher side load
 factor, if certification in the acrobatic category is desired, than that
 required for certification in the normal and utility categories. When
 designing for landing loads, the landing gear and the airplane as a whole
 need only be investigated for the category corresponding to the maximum
 design weight. These simplifications apply to single-engine aircraft of
 conventional types for which experience is available, and the Administrator
 may require additional investigations for aircraft with unusual design
 features.

 A23.7  Flight loads.

   (a) Each flight load may be considered independent of altitude and, except
 for the local supporting structure for dead weight items, only the maximum
 design weight conditions must be investigated.
   (b) Table 1 and figures 3 and 4 of this Appendix must be used to determine
 values of n1, n2, n3, and n4, corresponding to the maximum design weights in
 the desired categories.
   (c) Figures 1 and 2 of this Appendix must be used to determine values of n3
 and n4  corresponding to the minimum flying weights in the desired
 categories, and, if these load factors are greater than the load factors at
 the design weight, the supporting structure for dead weight items must be
 substantiated for the resulting higher load factors.
   (d) Each specified wing and tail loading is independent of the center of
 gravity range. The applicant, however, must select a c.g. range, and the
 basic fuselage structure must be investigated for the most adverse dead
 weight loading conditions for the c.g. range selected.
   (e) The following loads and loading conditions are the minimums for which
 strength must be provided in the structure:
   (1) Airplane equilibrium. The aerodynamic wing loads may be considered to
 act normal to the relative wind, and to have a magnitude of 1.05 times the
 airplane normal loads (as determined from paragraphs A23.9 (b) and (c) of
 this appendix) for the positive flight conditions and a magnitude equal to
 the airplane normal loads for the negative conditions. Each chordwise and
 normal component of this wing load must be considered.
   (2) Minimum design airspeeds. The minimum design airspeeds may be chosen by
 the applicant except that they may not be less than the minimum speeds found
 by using figure 3 of this Appendix. In addition, VCmin need not exceed values
 of 0.9 VH actually obtained at sea level for the lowest design weight
 category for which certification is desired. In computing these minimum
 design airspeeds, n1 may not be less than 3.8.
   (3) Flight load factor. The limit flight load factors specified in Table 1
 of this Appendix represent the ratio of the aerodynamic force component
 (acting normal to the assumed longitudinal axis of the airplane) to the
 weight of the airplane. A positive flight load factor is an aerodynamic force
 acting upward, with respect to the airplane.

 A23.9  Flight conditions.

   (a) General. Each design condition in paragraphs (b) and (c) of this
 section must be used to assure sufficient strength for each condition of
 speed and load factor on or within the boundary of a V-n diagram for the
 airplane similar to the diagram in figure 4 of this Appendix. This diagram
 must also be used to determine the airplane structural operating limitations
 as specified in Secs. 23.1501(c) through 23.1513 and Sec. 23.1519.
   (b) Symmetrical flight conditions. The airplane must be designed for
 symmetrical flight conditions as follows:
   (1) The airplane must be designed for at least the four basic flight
 conditions, "A", "D", "E", and "G" as noted on the flight envelope of figure
 4 of this Appendix. In addition, the following requirements apply:
   (i) The design limit flight load factors corresponding to conditions "D"
 and "E" of figure 4 must be at least as great as those specified in Table 1
 and figure 4 of this Appendix, and the design speed for these conditions must
 be at least equal to the value of VD found from figure 3 of this Appendix.
   (ii) For conditions "A" and "G" of figure 4, the load factors must
 correspond to those specified in Table 1 of this Appendix, and the design
 speeds must be computed using these load factors with the maximum static lift
 coefficient CNA determined by the applicant. However, in the absence of more
 precise computations, these latter conditions may be based on a value of CNA
 =+/-1.35 and the design speed for condition "A" may be less than VAmin.
   (iii) Conditions "C" and "F" of figure 4 need only be investigated when
 n3W/S or n4W/S are greater than n1W/S or n2W/S  of this Appendix,
 respectively.
   (2) If flaps or other high lift devices intended for use at the relatively
 low airspeed of approach, landing, and takeoff, are installed, the airplane
 must be designed for the two flight conditions corresponding to the values of
 limit flap-down factors specified in Table 1 of this Appendix with the flaps
 fully extended at not less than the design flap speed VFmin from figure 3 of
 this Appendix.
   (c) Unsymmetrical flight conditions. Each affected structure must be
 designed for unsymmetrical loadings as follows:
   (1) The aft fuselage-to-wing attachment must be designed for the critical
 vertical surface load determined in accordance with paragraph SA23.11(c) (1)
 and (2) of this Appendix.
   (2) The wing and wing carry-through structures must be designed for 100
 percent of condition "A" loading on one side of the plane of symmetry and 70
 percent on the opposite side for certification in the normal and utility
 categories, or 60 percent on the opposite side for certification in the
 acrobatic category.
   (3) The wing and wing carry-through structures must be designed for the
 loads resulting from a combination of 75 percent of the positive maneuvering
 wing loading on both sides of the plane of symmetry and the maximum wing
 torsion resulting from aileron displacement. The effect of aileron
 displacement on wing torsion at VC or VA using the basic airfoil moment
 coefficient modified over the aileron portion of the span, must be computed
 as follows:
   (i) Cm=Cm +0.01<delta><mu> (up aileron side) wing basic airfoil.
   (ii) Cm=Cm -0.01<delta><mu>(down aileron side) wing basic airfoil, where
 <delta><mu> is the up aileron deflection and <delta>d is the down aileron
 deflection.
   (4) <Delta> critical, which is the sum of <delta><mu>+<delta>d must be
 computed as follows:
   (i) Compute <Delta><alpha> and <Delta>b from the formulas:

                                    VA
                             D a =  --  x D p    and
                                    VC

                                        VA
                              Db = 0.5  --  x Dp
                                        VD

 Where Dp = the maximum total deflection (sum of both aileron deflections) at
 VA with VA, VC, and VD described in subparagraph (2) of Sec. 23.7(e) of this
 Appendix.

   (ii) Compute K from the formula:

                                 (Cm-0.01db) VD2
                            K =  ------------------
                                 (Cm-0.01da) VC2

 where <delta><alpha> is the down aileron deflection corresponding to
 <Delta><alpha>, and <delta>b is the down aileron deflection corresponding to
 <Delta>b as computed in step (i).
   (iii) If K is less than 1.0, <Delta><alpha> is <Delta> critical and must be
 used to determine <delta>u and <delta>d. In this case, VC is the critical
 speed which must be used in computing the wing torsion loads over the aileron
 span.
   (iv) If K is equal to or greater than 1.0, <Delta>b is <Delta> critical and
 must be used to determine <delta>u and <delta>d. In this case, Vd is the
 critical speed which must be used in computing the wing torsion loads over
 the aileron span.
   (d) Supplementary conditions; rear lift truss; engine torque; side load on
 engine mount. Each of the following supplementary conditions must be
 investigated:
   (1) In designing the rear lift truss, the special condition specified in
 Sec. 23.369 may be investigated instead of condition "G" of figure 4 of this
 Appendix. If this is done, and if certification in more than one category is
 desired, the value of W/S used in the formula appearing in Sec. 23.369 must
 be that for the category corresponding to the maximum gross weight.
   (2) Each engine mount and its supporting structures must be designed for
 the maximum limit torque corresponding to METO power and propeller speed
 acting simultaneously with the limit loads resulting from the maximum
 positive maneuvering flight load factor n1. The limit torque must be obtained
 by multiplying the mean torque by a factor of 1.33 for engines with five or
 more cylinders. For 4, 3, and 2 cylinder engines, the factor must be 2, 3,
 and 4, respectively.
   (3) Each engine mount and its supporting structure must be designed for the
 loads resulting from a lateral limit load factor of not less than 1.47 for
 the normal and utility categories, or 2.0 for the acrobatic category.

 A23.11  Control surface loads.

   (a) General. Each control surface load must be determined using the
 criteria of paragraph (b) of this section and must lie within the simplified
 loadings of paragraph (c) of this section.
   (b) Limit pilot forces. In each control surface loading condition described
 in paragraphs (c) through (e) of this section, the airloads on the movable
 surfaces and the corresponding deflections need not exceed those which could
 be obtained in flight by employing the maximum limit pilot forces specified
 in the table in Sec. 23.397(b). If the surface loads are limited by these
 maximum limit pilot forces, the tabs must either be considered to be
 deflected to their maximum travel in the direction which would assist the
 pilot or the deflection must correspond to the maximum degree of "out of
 trim" expected at the speed for the condition under consideration. The tab
 load, however, need not exceed the value specified in Table 2 of this
 Appendix.
   (c) Surface loading conditions. Each surface loading condition must be
 investigated as follows:
   (1) Simplified limit surface loadings and distributions for the horizontal
 tail, vertical tail, aileron, wing flaps, and trim tabs are specified in
 Table 2 and figures 5 and 6 of this Appendix. If more than one distribution
 is given, each distribution must be investigated.
   (2) If certification in the acrobatic category is desired, the horizontal
 tail must be investigated for an unsymmetrical load of 100 percent w on one
 side of the airplane centerline and 50 percent on the other side of the
 airplane centerline.
   (d) Outboard fins. Outboard fins must meet the requirements of Sec. 23.455.
   (e) Special devices. Special devices must meet the requirements of Sec.
 23.459.

 A23.13  Control system loads.

   (a) Primary flight controls and systems. Each primary flight control and
 system must be designed as follows:
   (1) The flight control system and its supporting structure must be designed
 for loads corresponding to 125 percent of the computed hinge moments of the
 movable control surface in the conditions prescribed in A23.11 of this
 Appendix. In addition--
   (i) The system limit loads need not exceed those that could be produced by
 the pilot and automatic devices operating the controls; and
   (ii) The design must provide a rugged system for service use, including
 jamming, ground gusts, taxiing downwind, control inertia, and friction.
   (2) Acceptable maximum and minimum limit pilot forces for elevator,
 aileron, and rudder controls are shown in the table in Sec. 23.397(b). These
 pilots loads must be assumed to act at the appropriate control grips or pads
 as they would under flight conditions, and to be reacted at the attachments
 of the control system to the control surface horn.
   (b) Dual controls. If there are dual controls, the systems must be designed
 for pilots operating in opposition, using individual pilot loads equal to 75
 percent of those obtained in accordance with paragraph (a) of this section,
 except that individual pilot loads may not be less than the minimum limit
 pilot forces shown in the table in Sec. 23.397(b).
   (c) Ground gust conditions. Ground gust conditions must meet the
 requirements of Sec. 23.415.
   (d) Secondary controls and systems. Secondary controls and systems must
 meet the requirements of Sec. 23.405.

                       Table 1--Limit Flight Load Factors

                          [Limit flight load factors]

                   Flight load   Normal   Utility   Acrobatic
                     factors    category  category  category

                   Flaps up:
                    n1               3.8       4.4        6.0
                    n2           -0.5 n1
                    n3             (/1/)
                    n4             (/2/)
                   Flaps down:
                    n flap        0.5 n1
                    n flap      /3/ Zero

                   /1/ Find n3 from Fig. 1

                   /2/ Find n4 from Fig. 2

                   /3/ Vertical wing load may be assumed
                   equal to zero and only the flap part of
                   the wing need be checked for this
                   condition.

                      [ ...Illustration appears here... ]

                      [ ...Illustration appears here... ]

                      [ ...Illustration appears here... ]

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13097, Aug. 13, 1969; 34 FR 14727, Sept. 24, 1969; Amdt. 23-16, 40 FR 2577,
 Jan. 14, 1975; Amdt. 23-28, 47 FR 13315, Mar. 29, 1982]






                     Appendix B to Part 23--[Reserved]

 *****************************************************************************


 56 FR 344, No. 2, Jan. 3, 1991

   SUMMARY: This final rule upgrades the airworthiness standards for normal,
 utility, acrobatic, and commuter category airplanes. This amendment provides
 airworthiness standards for advancements in technology being incorporated in
 current designs, permits type certification of spin resistant airplanes, and
 reduces the regulatory burden in showing compliance with some of the
 requirements for the design and type certification of small airplanes. These
 new and amended airworthiness standards also result in the need for new
 definitions. As a result, new definitions are added.

   DATES: February 4, 1991.

 *****************************************************************************






                Appendix C to Part 23--Basic Landing Conditions

                       [C23.1 Basic landing conditions]

                                                       Tail wheel type

                                                                   Tail-down
                  Condition                      Level landing      landing

 Reference section                             23.479(a)(1)       23.481(a)(1)

 Vertical component at c. g                    nW                 nW
 Fore and aft component at c. g                KnW                0
 Lateral component in either direction at c.
  g                                            0                  0
 Shock absorber extension (hydraulic shock
  absorber)                                    Note (2)           Note (2)
 Shock absorber deflection (rubber or spring
  shock absorber), percent                     100                100
 Tire deflection                               Static             Static
 Main wheel loads (both wheels) (Vr)           (n-L)W             (n-L)W b/d
 Main wheel loads (both wheels) (Dr)           KnW                0
 Tail (nose) wheel loads (Vf)                  0                  (n-L)W a/d
 Tail (nose) wheel loads (Df)                  0                  0
 Notes                                         (1), (3), and (4)  (4)

                           [ ...Table continues... ]

                                           Nose wheel type

                                          Level landing
                        Level landing    with nose wheel
                        with inclined     just clear of
       Condition          reactions          ground         Tail-down landing

  Reference section    23.479(a)(2)(i)  23.479(a)(2)(ii)   23.481(a)(2) and
                                                            (b).

  Vertical component   nW               nW                 nW.
   at c. g
  Fore and aft         KnW              KnW                0.
   component at c. g
  Lateral component    0                0                  0.
   in either
   direction at c. g
  Shock absorber       Note (2)         Note (2)           Note (2).
   extension
   (hydraulic shock
   absorber)
  Shock absorber       100              100                100.
   deflection (rubber
   or spring shock
   absorber), percent
  Tire deflection      Static           Static             Static.
  Main wheel loads     (n-L)W a'/d'     (n-L)W             (n-L)W.
   (both wheels) (Vr)
  Main wheel loads     KnW a'/d'        KnW                0.
   (both wheels) (Dr)
  Tail (nose) wheel    (n-L)W b'/d'     0                  0.
   loads (Vf)
  Tail (nose) wheel    KnW b'/d'        0                  0.
   loads (Df)
  Notes                (1)              (1), (3), and (4)  (3) and (4).

  Note (1). K may be determined as follows: K=0.25 for W=3,000 pounds or less;
  K=0.33 for W=6,000 pounds or greater, with linear variation of K between
  these weights.

  Note (2). For the purpose of design, the maximum load factor is assumed to
  occur throughout the shock absorber stroke from 25 percent deflection to 100
  percent deflection unless otherwise shown and the load factor must be used
  with whatever shock absorber extension is most critical for each element of
  the landing gear.

  Note (3). Unbalanced moments must be balanced by a rational or conservative
  method.

  Note (4). L is defined in Sec. 23.735(b).

  Note (5). n is the limit inertia load factor, at the c.g. of the airplane,
  selected under Sec. 23.473 (d), (f), and (g).

                             Basic Landing Conditions

                      [ ...Illustration appears here... ]

                                  Level Landing

                      [ ...Illustration appears here... ]

                                Tail Down Landing

                      [ ...Illustration appears here... ]

                     Level Landing with Inclined Reactions

                      [ ...Illustration appears here... ]

               Level Landing with Nose Wheel Just Clear of Ground

                      [ ...Illustration appears here... ]

                                Tail Down Landing

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR
 13099, Aug. 13, 1969]






 Appendix D to Part 23--Wheel Spin-Up and Spring-Back Loads

 D23.1 Wheel spin-up loads.

   (a) The following method for determining wheel spin-up loads for landing
 conditions is based on NACA T.N. 863. However, the drag component used for
 design may not be less than the drag load prescribed in Sec. 23.479(b).

                  FHmax =1/re <radical> 2Iw(VH--Vc)nFVmax/tS

 where--

 FHmax=maximum rearward horizontal force acting on the wheel (in pounds);
 re=effective rolling radius of wheel under impact based on recommended
     operating tire pressure (which may be assumed to be equal to the rolling
     radius under a static load of njWe) in feet;
 Iw=rotational mass moment of inertia of rolling assembly (in slug feet);
 VH=linear velocity of airplane parallel to ground at instant of contact
     (assumed to be 1.2 VS0,  in feet per second);
 Vc=peripheral speed of tire, if prerotation is used (in feet per second)
     (there must be a positive means of pre-rotation before pre-rotation may
     be considered);
 n=equals effective coefficient of friction (0.80 may be used);
 FVmax=maximum vertical force on wheel (pounds)= njWe, where We and nj are
     defined in Sec. 23.725;
 ts=time interval between ground  contact and attainment of maximum vertical
     force on wheel (seconds). (However, if the value of FVmax, from the above
     equation exceeds 0.8 FVmax, the latter value must be used for FHmax.)

   (b) The equation assumes a linear variation of load factor with time until
 the peak load is reached and under this assumption, the equation determines
 the drag force at the time that the wheel peripheral velocity at radius re
 equals the airplane velocity. Most shock absorbers do not exactly follow a
 linear variation of load factor with time. Therefore, rational or
 conservative allowances must be made to compensate for these variations. On
 most landing gears, the time for wheel spin-up will be less than the time
 required to develop maximum vertical load factor for the specified rate of
 descent and forward velocity. For exceptionally large wheels, a wheel
 peripheral velocity equal to the ground speed may not have been attained at
 the time of maximum vertical gear load. However, as stated above, the drag
 spin-up load need not exceed 0.8 of the maximum vertical loads.
   (c) Dynamic spring-back of the landing gear and adjacent structure at the
 instant just after the wheels come up to speed may result in dynamic forward
 acting loads of considerable magnitude. This effect must be determined, in
 the level landing condition, by assuming that the wheel spin-up loads
 calculated by the methods of this appendix are reversed. Dynamic spring-back
 is likely to become critical for landing gear units having wheels of large
 mass or high landing speeds.

 [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as
 amended by Amdt. No. 23-45, 58 FR 42167, Aug. 6, 1993]

 *****************************************************************************


 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************



 Appendix E to Part 23--Limited Weight Credit for Airplanes Equipped With
     Standby Power

   (a) Each applicant for an increase in the maximum certificated takeoff and
 landing weights of an airplane equipped with a typecertificated standby power
 rocket engine may obtain an increase as specified in paragraph (b) if--
   (1) The installation of the rocket engine has been approved and it has been
 established by flight test that the rocket engine and its controls can be
 operated safely and reliably at the increase in maximum weight; and
   (2) The Airplane Flight Manual, or the placard, markings or manuals
 required in place thereof, set forth in addition to any other operating
 limitations the Administrator may require, the increased weight approved
 under this regulation and a prohibition against the operation of the airplane
 at the approved increased weight when--
   (i) The installed standby power rocket engines have been stored or
 installed in excess of the time limit established by the manufacturer of the
 rocket engine (usually stenciled on the engine casing); or
   (ii) The rocket engine fuel has been expended or discharged.
   (b) The currently approved maximum takeoff and landing weights at which an
 airplane is certificated without a standby power rocket engine installation
 may be increased by an amount which does not exceed any of the following:
   (1) An amount equal in pounds to 0.014 IN, where I is the maximum usable
 impulse in pounds-seconds available from each standby power rocket engine and
 N is the number of rocket engines installed.
   (2) An amount equal to 5 percent of the maximum certificated weight
 approved in accordance with the applicable airworthiness regulations without
 standby power rocket engines installed.
   (3) An amount equal to the weight of the rocket engine installation.
   (4) An amount that, together with the currently approved maximum weight,
 would equal the maximum structural weight established for the airplane
 without standby rocket engines installed.
   (c) For the purposes of this Appendix, "standby power" is power or thrust,
 or both, obtained from rocket engines for a relatively short period and
 actuated only in cases of emergency.
   (d) For the purposes of limited weight credit for airplanes equipped with
 standby power, as set forth in Sec. 23.25(a)(1)(iii) and this Appendix, an
 airplane certificated under Part 4a of the Civil Air Regulations is treated
 as if it had been certificated under Part 3 of the Civil Air Regulations or
 Part 23 of the Federal Aviation Regulations.

 [Amdt. 23-2, 30 FR 8468, July 2, 1965]






                     Appendix F to Part 23--Test Procedure

   An Acceptable Test Procedure for Self-Extinguishing Materials for Showing
 Compliance with Sec. 23.853.
   (a) Conditioning. Specimens must be conditioned to 70 degrees F, plus or
 minus 5 degrees, and at 50 percent plus or minus 5 percent relative humidity
 until moisture equilibrium is reached or for 24 hours. Only one specimen at a
 time may be removed from the conditioning environment immediately before
 subjecting it to the flame.
   (b) Specimen configuration. Materials must be tested either as a section
 cut from a fabricated part as installed in the airplane or as a specimen
 simulating a cut section, such as a specimen cut from a flat sheet of the
 material or a model of the fabricated part. The specimen may be cut from any
 location in a fabricated part; however, fabricated units, such as sandwich
 panels, may not be separated for test. The specimen thickness must be no
 thicker than the minimum thickness to be qualified for use in the airplane,
 except that thick foam parts must be tested in 1/2 -inch thickness. In the
 case of fabrics, both the warp and fill direction of the weave must be tested
 to determine the most critical flammability conditions. When performing the
 test prescribed in paragraphs (d) and (e) of this Appendix, the specimen must
 be mounted in a metal frame so that: (1) The two long edges and the upper
 edge are held securely; (2) the exposed area of the specimen is at least 2
 inches wide and 12 inches long, unless the actual size used in the airplane
 is smaller; and (3) the edge to which the burner flame is applied must not
 consist of the finished or protected edge of the specimen but must be
 representative of the actual cross section of the material or part installed
 in the airplane.
   (c) Apparatus. Except as provided in paragraph (e) of this Appendix, tests
 must be conducted in a draft-free cabinet in accordance with Federal Test
 Method Standard 191 Method 5903 (revised Method 5902) which is available from
 the General Services Administration, Business Service Center, Region 3,
 Seventh and D Streets SW., Washington, D.C. 20407, or with some other
 approved equivalent method. Specimens which are too large for the cabinet
 must be tested in similar draft-free conditions.
   (d) Vertical test. A minimum of three specimens must be tested and the
 results averaged. For fabrics, the direction of weave corresponding to the
 most critical flammability conditions must be parallel to the longest
 dimension. Each specimen must be supported vertically. The specimen must be
 exposed to a Bunsen or Tirrill burner with a nominal 3/8 -inch I.D. tube
 adjusted to give a flame of 1 1/2  inches in height. The minimum flame
 temperature measured by a calibrated thermocouple pryometer in the center of
 the flame must be 1550 deg. F. The lower edge of the specimen must be three-
 fourths inch above the top edge of the burner. The flame must be applied to
 the center line of the lower edge of the specimen. For materials covered by
 Secs. 23.853(d)(3)(i) and 23.853(f), the flame must be applied for 60 seconds
 and then removed. For materials covered by Sec. 23.853(d)(3)(ii), the flame
 must be applied for 12 seconds and then removed. Flame time, burn length, and
 flaming time of drippings, if any, must be recorded. The burn length
 determined in accordance with paragraph (f) of this Appendix must be measured
 to the nearest one-tenth inch.
   (e) Horizontal test. A minimum of three specimens must be tested and the
 results averaged. Each specimen must be supported horizontally. The exposed
 surface when installed in the airplane must be face down for the test. The
 specimen must be exposed to a Bunsen burner or Tirrill burner with a nominal
 3/8 -inch I.D. tube adjusted to give a flame of 1 1/2  inches in height. The
 minimum flame temperature measured by a calibrated thermocouple pyrometer in
 the center of the flame must be 1550 deg. F. The specimen must be positioned
 so that the edge being tested is three-fourths of an inch above the top of,
 and on the center line of, the burner. The flame must be applied for 15
 seconds and then removed. A minimum of 10 inches of the specimen must be used
 for timing purposes, approximately 1 1/2  inches must burn before the burning
 front reaches the timing zone, and the average burn rate must be recorded.
   (f) Burn length. Burn length is the distance from the original edge to the
 farthest evidence of damage to the test specimen due to flame impingement,
 including areas of partial or complete consumption, charring, or
 embrittlement, but not including areas sooted, stained, warped, or
 discolored, nor areas where material has shrunk or melted away from the heat
 source.

 [Amdt. 23-23, 43 FR 50594, Oct. 30, 1978, as amended by Amdt. 23-34, 52 FR
 1835, Jan. 15, 1987; 52 FR 34745, Sept. 14, 1987]






        Appendix G to Part 23--Instructions for Continued Airworthiness

   G23.1  General. (a) This appendix specifies requirements for the
 preparation of Instructions for Continued Airworthiness as required by Sec.
 23.1529.
   (b) The Instructions for Continued Airworthiness for each airplane must
 include the Instructions for Continued Airworthiness for each engine and
 propeller (hereinafter designated 'products'), for each appliance required by
 this chapter, and any required information relating to the interface of those
 appliances and products with the airplane. If Instructions for Continued
 Airworthiness are not supplied by the manufacturer of an appliance or product
 installed in the airplane, the Instructions for Continued Airworthiness for
 the airplane must include the information essential to the continued
 airworthiness of the airplane.
   (c) The applicant must submit to the FAA a program to show how changes to
 the Instructions for Continued Airworthiness made by the applicant or by the
 manufacturers of products and appliances installed in the airplane will be
 distributed.
   G23.2  Format. (a) The Instructions for Continued Airworthiness must be in
 the form of a manual or manuals as appropriate for the quantity of data to be
 provided.
   (b) The format of the manual or manuals must provide for a practical
 arrangement.
   G23.3  Content. The contents of the manual or manuals must be prepared in
 the English language. The Instructions for Continued Airworthiness must
 contain the following manuals or sections, as appropriate, and information:
   (a) Airplane maintenance manual or section. (1) Introduction information
 that includes an explanation of the airplane's features and data to the
 extent necessary for maintenance or preventive maintenance.
   (2) A description of the airplane and its systems and installations
 including its engines, propellers, and appliances.
   (3) Basic control and operation information describing how the airplane
 components and systems are controlled and how they operate, including any
 special procedures and limitations that apply.
   (4) Servicing information that covers details regarding servicing points,
 capacities of tanks, reservoirs, types of fluids to be used, pressures
 applicable to the various systems, location of access panels for inspection
 and servicing, locations of lubrication points, lubricants to be used,
 equipment required for servicing, tow instructions and limitations, mooring,
 jacking, and leveling information.
   (b) Maintenance instructions. (1) Scheduling information for each part of
 the airplane and its engines, auxiliary power units, propellers, accessories,
 instruments, and equipment that provides the recommended periods at which
 they should be cleaned, inspected, adjusted, tested, and lubricated, and the
 degree of inspection, the applicable wear tolerances, and work recommended at
 these periods. However, the applicant may refer to an accessory, instrument,
 or equipment manufacturer as the source of this information if the applicant
 shows that the item has an exceptionally high degree of complexity requiring
 specialized maintenance techniques, test equipment, or expertise. The
 recommended overhaul periods and necessary cross reference to the
 Airworthiness Limitations section of the manual must also be included. In
 addition, the applicant must include an inspection program that includes the
 frequency and extent of the inspections necessary to provide for the
 continued airworthiness of the airplane.
   (2) Troubleshooting information describing probable malfunctions, how to
 recognize those malfunctions, and the remedial action for those malfunctions.
   (3) Information describing the order and method of removing and replacing
 products and parts with any necessary precautions to be taken.
   (4) Other general procedural instructions including procedures for system
 testing during ground running, symmetry checks, weighing and determining the
 center of gravity, lifting and shoring, and storage limitations.
   (c) Diagrams of structural access plates and information needed to gain
 access for inspections when access plates are not provided.
   (d) Details for the application of special inspection techniques including
 radiographic and ultrasonic testing where such processes are specified.
   (e) Information needed to apply protective treatments to the structure
 after inspection.
   (f) All data relative to structural fasteners such as identification,
 discard recommendations, and torque values.
   (g) A list of special tools needed.
   (h) In addition, for commuter category airplanes, the following information
 must be furnished:
   (1) Electrical loads applicable to the various systems;
   (2) Methods of balancing control surfaces;
   (3) Identification of primary and secondary structures; and
   (4) Special repair methods applicable to the airplane.
   G23.4  Airworthiness Limitations section. The Instructions for Continued
 Airworthiness must contain a section titled Airworthiness Limitations that is
 segregated and@learly distinguishable from the rest of the document. This
 section must set forth each mandatory replacement time, structural inspection
 interval, and related structural inspection procedure required for type
 certification. If the Instructions for Continued Airworthiness consist of
 multiple documents, the section required by this paragraph must be included
 in the principal manual. This section must contain a legible statement in a
 prominent location that reads: "The Airworthiness Limitations section is FAA
 approved and specifies maintenance required under Secs. 43.16 and 91.403 of
 the Federal Aviation Regulations unless an alternative program has been FAA
 approved."

 [Amdt. 23-26, 45 FR 60171, Sept. 11, 1980, as amended by Amdt. 23-34, 52 FR
 1835, Jan. 15, 1987; 52 FR 34745, Sept. 14, 1987; Amdt. 23-37, 54 FR 34329,
 Aug. 18, 1989]

   Effective Date Note: At 54 FR 34329, Aug. 18, 1989, Sec. G23.4 in Appendix
 G, Part 23 was amended by changing the cross reference "Sec. 91.163" to "Sec.
 91.403", effective August 18, 1990.






 Appendix H to Part 23--Installation of An Automatic Power Reserve (APR)
 System

   H23.1, General.
   (a) This appendix specifies requirements for installation of an APR engine
 power control system that automatically advances power or thrust on the
 operating engine(s) in the event any engine fails during takeoff.
   (b) With the APR system and associated systems functioning normally, all
 applicable requirements (except as provided in this appendix) must be met
 without requiring any action by the crew to increase power or thrust.
   H23.2, Definitions.
   (a) Automatic power reserve system means the entire automatic system used
 only during takeoff, including all devices both mechanical and electrical
 that sense engine failure, transmit signals, actuate fuel controls or power
 levers on operating engines, including power sources, to achieve the
 scheduled power increase and furnish cockpit information on system operation.
   (b) Selected takeoff power, notwithstanding the definition of "Takeoff
 Power" in part 1 of the Federal Aviation Regulations, means the power
 obtained from each initial power setting approved for takeoff.
   (c) Critical Time Interval, as illustrated in figure H1, means that period
 starting at V1 minus one second and ending at the intersection of the engine
 and APR failure flight path line with the minimum performance all engine
 flight path line. The engine and APR failure flight path line intersects the
 one-engine-inoperative flight path line at 400 feet above the takeoff
 surface. The engine and APR failure flight path is based on the airplane's
 performance and must have a positive gradient of at least 0.5 percent at 400
 feet above the takeoff surface.

                 Figure H1--Critical Time Interval Illustration

        [INSERT: Line graph plotting engine and APR failure flight path
               against minimum performace all engine flight path]

   H23.3, Reliability and performance requirements.
   (a) It must be shown that, during the critical time interval, an APR
 failure that increases or does not affect power on either engine will not
 create a hazard to the airplane, or it must be shown that such failures are
 improbable.
   (b) It must be shown that, during the critical time interval, there are no
 failure modes of the APR system that would result in a failure that will
 decrease the power on either engine or it must be shown that such failures
 are extremely improbable.
   (c) It must be shown that, during the critical time interval, there will be
 no failure of the APR system in combination with an engine failure or it must
 be shown that such failures are extremely improbable.
   (d) All applicable performance requirements must be met with an engine
 failure occurring at the most critical point during takeoff with the APR
 system functioning normally.
   H23.4, Power setting.
   The selected takeoff power set on each engine at the beginning of the
 takeoff roll may not be less than--
   (a) The power necessary to attain, at V1, 90 percent of the maximum takeoff
 power approved for the airplane for the existing conditions;
   (b) That required to permit normal operation of all safety-related systems
 and equipment that are dependent upon engine power or power lever position;
 and
   (c) That shown to be free of hazardous engine response characteristics when
 power is advanced from the selected takeoff power level to the maximum
 approved takeoff power.
   H23.5, Powerplant controls--general.
   (a) In addition to the requirements of Sec. 23.1141, no single failure or
 malfunction (or probable combination thereof) of the APR, including
 associated systems, may cause the failure of any powerplant function
 necessary for safety.
   (b) The APR must be designed to--
   (1) Provide a means to verify to the flight crew before takeoff that the
 APR is in an operating condition to perform its intended function;
   (2) Automatically advance power on the operating engines following an
 engine failure during takeoff to achieve the maximum attainable takeoff power
 without exceeding engine operating limits;
   (3) Prevent deactivation of the APR by manual adjustment of the power
 levers following an engine failure;
   (4) Provide a means for the flight crew to deactivate the automatic
 function. This means must be designed to prevent inadvertent deactivation;
 and
   (5) Allow normal manual decrease or increase in power up to the maximum
 takeoff power approved for the airplane under the existing conditions through
 the use of power levers, as stated in Sec. 23.1141(c), except as provided
 under paragraph (c) of H23.5 of this appendix.
   (c) For airplanes equipped with limiters that automatically prevent engine
 operating limits from being exceeded, other means may be used to increase the
 maximum level of power controlled by the power levers in the event of an APR
 failure. The means must be located on or forward of the power levers, must be
 easily identified and operated under all operating conditions by a single
 action of any pilot with the hand that is normally used to actuate the power
 levers, and must meet the requirements of Sec. 23.777 (a), (b), and (c).
   H23.6, Powerplant instruments.
   In addition to the requirements of Sec. 23.1305:
   (a) A means must be provided to indicate when the APR is in the armed or
 ready condition.
   (b) If the inherent flight characteristics of the airplane do not provide
 warning that an engine has failed, a warning system independent of the APR
 must be provided to give the pilot a clear warning of any engine failure
 during takeoff.
   (c) Following an engine failure at V1 or above, there must be means for the
 crew to readily and quickly verify that the APR has operated satisfactorily.

 [Amdt. 23-43, 58 FR 18979, Apr. 9, 1993]

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 58 FR 18958, No. 67, Apr. 9, 1993

 SUMMARY: This final rule amends the powerplant and equipment airworthiness
 standards for normal, utility, acrobatic, and commuter category airplanes.
 This amendment is based on certain proposals and recommendations discussed at
 the Small Airplane Airworthiness Review Conference held on October 22-26,
 1984, in St. Louis, Missouri, and arises from the recognition by both
 government and industry, that upgraded standards are needed to maintain an
 acceptable level of safety for small airplanes.

 EFFECTIVE DATE: May 10, 1993.

 *****************************************************************************






                    Appendix I to Part 23--Seaplane Loads

                                Appendix I

    Figure 1. Pictorial definition of angles, dimensions, and directions
                               on a seaplane

                             [INSERT: Diagrams]

                   Figure 2. Hull station weighing factor

                             [INSERT: Diagrams]

                 Figure 3. Transverse pressure distributions

                             [INSERT: Diagrams]

 [Amdt. No. 23-45, 58 FR 42167, Aug. 6, 1993; 58 FR 51970, 51971, Oct. 5,
 1993]

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 58 FR 42136, No. 150, Aug. 6, 1993

 SUMMARY: This amendment changes airframe and flight worthiness standards for
 normal, utility, acrobatic, and commuter category airplanes. The changes are
 based on a number of recommendations discussed at the Small Airplane
 Airworthiness Review Conference held on October 22-26, 1984, St. Louis,
 Missouri. These updated safety standards will continue to provide an
 acceptable level of safety in the design requirements for small airplanes
 used in both private and commercial operations. Some of the changes provide
 design requirements applicable to advancements in technology being
 incorporated in current designs. This amendment will also reduce the
 regulatory burden in showing compliance with some requirements while
 maintaining an acceptable level of safety.

 EFFECTIVE DATE: September 7, 1993.

 *****************************************************************************