PART 29--AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY ROTORCRAFT
Special Federal Aviation Regulation No. 29-4
Subpart A--General
Sec. 29.1 Applicability.
Sec. 29.2 Special retroactive requirements.
Subpart B--Flight
General
Sec. 29.21 Proof of compliance.
Sec. 29.25 Weight limits.
Sec. 29.27 Center of gravity limits.
Sec. 29.29 Empty weight and corresponding center of gravity.
Sec. 29.31 Removable ballast.
Sec. 29.33 Main rotor speed and pitch limits.
Performance
Sec. 29.45 General.
Sec. 29.51 Takeoff data: general.
Sec. 29.53 Takeoff: Category A.
Sec. 29.59 Takeoff path: Category A.
Sec. 29.63 Takeoff: Category B.
Sec. 29.65 Climb: all engines operating.
Sec. 29.67 Climb: one engine inoperative.
Sec. 29.71 Helicopter angle of glide: Category B.
Sec. 29.73 Performance at minimum operating speed.
Sec. 29.75 Landing.
Sec. 29.77 Balked landing: category A.
Sec. 29.79 Limiting height-speed envelope.
Flight Characteristics
Sec. 29.141 General.
Sec. 29.143 Controllability and maneuverability.
Sec. 29.151 Flight controls.
Sec. 29.161 Trim control.
Sec. 29.171 Stability: general.
Sec. 29.173 Static longitudinal stability.
Sec. 29.175 Demonstration of static longitudinal stability.
Sec. 29.177 Static directional stability.
Sec. 29.181 Dynamic stability: Category A rotorcraft.
Ground and Water Handling Characteristics
Sec. 29.231 General.
Sec. 29.235 Taxiing condition.
Sec. 29.239 Spray characteristics.
Sec. 29.241 Ground resonance.
Miscellaneous Flight Requirements
Sec. 29.251 Vibration.
Subpart C--Strength Requirements
General
Sec. 29.301 Loads.
Sec. 29.303 Factor of safety.
Sec. 29.305 Strength and deformation.
Sec. 29.307 Proof of structure.
Sec. 29.309 Design limitations.
Flight Loads
Sec. 29.321 General.
Sec. 29.337 Limit maneuvering load factor.
Sec. 29.339 Resultant limit maneuvering loads.
Sec. 29.341 Gust loads.
Sec. 29.351 Yawing conditions.
Sec. 29.361 Engine torque.
Control Surface and System Loads
Sec. 29.391 General.
Sec. 29.395 Control system.
Sec. 29.397 Limit pilot forces and torques.
Sec. 29.399 Dual control system.
Sec. 29.401 [Removed. Amdt. 29-31, 55 FR 38966, Sept. 21, 1990]
Sec. 29.403 [Removed. Amdt. 29-31, 55 FR 38966, Sept. 21, 1990]
Sec. 29.411 Ground clearance: tail rotor guard.
Sec. 29.413 [Removed. Amdt. 29-31, 55 FR 38966, Sept. 21, 1990]
Sec. 29.427 Unsymmetrical loads.
Ground Loads
Sec. 29.471 General.
Sec. 29.473 Ground loading conditions and assumptions.
Sec. 29.475 Tires and shock absorbers.
Sec. 29.477 Landing gear arrangement.
Sec. 29.479 Level landing conditions.
Sec. 29.481 Tail-down landing conditions.
Sec. 29.483 One-wheel landing conditions.
Sec. 29.485 Lateral drift landing conditions.
Sec. 29.493 Braked roll conditions.
Sec. 29.497 Ground loading conditions: landing gear with tail
wheels.
Sec. 29.501 Ground loading conditions: landing gear with skids.
Sec. 29.505 Ski landing conditions.
Sec. 29.511 Ground load: unsymmetrical loads on multiple-wheel
units.
Water Loads
Sec. 29.519 Hull type rotorcraft: Water-based and amphibian.
Sec. 29.521 Float landing conditions.
Main Component Requirements
Sec. 29.547 Main rotor structure.
Sec. 29.549 Fuselage and rotor pylon structures.
Sec. 29.551 Auxiliary lifting surfaces.
Emergency Landing Conditions
Sec. 29.561 General.
Sec. 29.562 Emergency landing dynamic conditions.
Sec. 29.563 Structural ditching provisions.
Fatigue Evaluation
Sec. 29.571 Fatigue evaluation of structure.
Subpart D--Design and Construction
General
Sec. 29.601 Design.
Sec. 29.603 Materials.
Sec. 29.605 Fabrication methods.
Sec. 29.607 Fasteners.
Sec. 29.609 Protection of structure.
Sec. 29.610 Lightning protection.
Sec. 29.611 Inspection provisions.
Sec. 29.613 Material strength properties and design values.
Sec. 29.619 Special factors.
Sec. 29.621 Casting factors.
Sec. 29.623 Bearing factors.
Sec. 29.625 Fitting factors.
Sec. 29.629 Flutter.
Rotors
Sec. 29.653 Pressure venting and drainage of rotor blades.
Sec. 29.659 Mass balance.
Sec. 29.661 Rotor blade clearance.
Sec. 29.663 Ground resonance prevention means.
Control Systems
Sec. 29.671 General.
Sec. 29.672 Stability augmentation, automatic, and power-
operated systems.
Sec. 29.673 Primary flight controls.
Sec. 29.674 Interconnected controls.
Sec. 29.675 Stops.
Sec. 29.679 Control system locks.
Sec. 29.681 Limit load static tests.
Sec. 29.683 Operation tests.
Sec. 29.685 Control system details.
Sec. 29.687 Spring devices.
Sec. 29.691 Autorotation control mechanism.
Sec. 29.695 Power boost and power-operated control system.
Landing Gear
Sec. 29.723 Shock absorption tests.
Sec. 29.725 Limit drop test.
Sec. 29.727 Reserve energy absorption drop test.
Sec. 29.729 Retracting mechanism.
Sec. 29.731 Wheels.
Sec. 29.733 Tires.
Sec. 29.735 Brakes.
Sec. 29.737 Skis.
Floats and Hulls
Sec. 29.751 Main float buoyancy.
Sec. 29.753 Main float design.
Sec. 29.755 Hull buoyancy.
Sec. 29.757 Hull and auxiliary float strength.
Personnel and Cargo Accommodations
Sec. 29.771 Pilot compartment.
Sec. 29.773 Pilot compartment view.
Sec. 29.775 Windshields and windows.
Sec. 29.777 Cockpit controls.
Sec. 29.779 Motion and effect of cockpit controls.
Sec. 29.783 Doors.
Sec. 29.785 Seats, safety belts, and harnesses.
Sec. 29.787 Cargo and baggage compartments.
Sec. 29.801 Ditching.
Sec. 29.803 Emergency evacuation.
Sec. 29.805 Flight crew emergency exits.
Sec. 29.807 Passenger emergency exits.
Sec. 29.809 Emergency exit arrangement.
Sec. 29.811 Emergency exit marking.
Sec. 29.812 Emergency lighting.
Sec. 29.813 Emergency exit access.
Sec. 29.815 Main aisle width.
Sec. 29.831 Ventilation.
Sec. 29.833 Heaters.
Fire Protection
Sec. 29.851 Fire extinguishers.
Sec. 29.853 Compartment interiors.
Sec. 29.855 Cargo and baggage compartmen@.
Sec. 29.859 Combustion heater fire protection.
Sec. 29.861 Fire protection of structure, controls, and other
parts.
Sec. 29.863 Flammable fluid fire protection.
External Load Attaching Means
Sec. 29.865 External load attaching means.
Miscellaneous
Sec. 29.871 Leveling marks.
Sec. 29.873 Ballast provisions.
Subpart E--Powerplant
General
Sec. 29.901 Installation.
Sec. 29.903 Engines.
Sec. 29.907 Engine vibration.
Sec. 29.908 Cooling fans.
Rotor Drive System
Sec. 29.917 Design.
Sec. 29.921 Rotor brake.
Sec. 29.923 Rotor drive system and control mechanism tests.
Sec. 29.927 Additional tests.
Sec. 29.931 Shafting critical speed.
Sec. 29.935 Shafting joints.
Sec. 29.939 Turbine engine operating characteristics.
Fuel System
Sec. 29.951 General.
Sec. 29.953 Fuel system independence.
Sec. 29.954 Fuel system lightning protection.
Sec. 29.955 Fuel flow.
Sec. 29.957 Flow between interconnected tanks.
Sec. 29.959 Unusable fuel supply.
Sec. 29.961 Fuel system hot weather operation.
Sec. 29.963 Fuel tanks: general.
Sec. 29.965 Fuel tank tests.
Sec. 29.967 Fuel tank installation.
Sec. 29.969 Fuel tank expansion space.
Sec. 29.971 Fuel tank sump.
Sec. 29.973 Fuel tank filler connection.
Sec. 29.975 Fuel tank vents and carburetor vapor vents.
Sec. 29.977 Fuel tank outlet.
Sec. 29.979 Pressure refueling and fueling provisions below
fuel level.
Fuel System Components
Sec. 29.991 Fuel pumps.
Sec. 29.993 Fuel system lines and fittings.
Sec. 29.995 Fuel valves.
Sec. 29.997 Fuel strainer or filter.
Sec. 29.999 Fuel system drains.
Sec. 29.1001 Fuel jettisoning.
Oil System
Sec. 29.1011 Engines: General.
Sec. 29.1013 Oil tanks.
Sec. 29.1015 Oil tank tests.
Sec. 29.1017 Oil lines and fittings.
Sec. 29.1019 Oil strainer or filter.
Sec. 29.1021 Oil system drains.
Sec. 29.1023 Oil radiators.
Sec. 29.1025 Oil valves.
Sec. 29.1027 Transmission and gearboxes: General.
Cooling
Sec. 29.1041 General.
Sec. 29.1043 Cooling tests.
Sec. 29.1045 Climb cooling test procedures.
Sec. 29.1047 Takeoff cooling test procedures.
Sec. 29.1049 Hovering cooling test procedures.
Induction System
Sec. 29.1091 Air induction.
Sec. 29.1093 Induction system icing protection.
Sec. 29.1101 Carburetor air preheater design.
Sec. 29.1103 Induction systems ducts and air duct systems.
Sec. 29.1105 Induction system screens.
Sec. 29.1107 Inter-coolers and after-coolers.
Sec. 29.1109 Carburetor air cooling.
Exhaust System
Sec. 29.1121 General.
Sec. 29.1123 Exhaust piping.
Sec. 29.1125 Exhaust heat exchangers.
Powerplant Controls and Accessories
Sec. 29.1141 Powerplant controls: general.
Sec. 29.1142 Auxiliary power unit controls.
Sec. 29.1143 Engine controls.
Sec. 29.1145 Ignition switches.
Sec. 29.1147 Mixture controls.
Sec. 29.1151 Rotor brake controls.
Sec. 29.1157 Carburetor air temperature controls.
Sec. 29.1159 Supercharger controls.
Sec. 29.1163 Powerplant accessories.
Sec. 29.1165 Engine ignition systems.
Powerplant Fire Protection
Sec. 29.1181 Designated fire zones: regions included.
Sec. 29.1183 Lines, fittings, and components.
Sec. 29.1185 Flammable fluids.
Sec. 29.1187 Drainage and ventilation of fire zones.
Sec. 29.1189 Shutoff means.
Sec. 29.1191 Firewalls.
Sec. 29.1193 Cowling and engine compartment covering.
Sec. 29.1194 Other surfaces.
Sec. 29.1195 Fire extinguishing systems.
Sec. 29.1197 Fire extinguishing agents.
Sec. 29.1199 Extinguishing agent containers.
Sec. 29.1201 Fire extinguishing system materials.
Sec. 29.1203 Fire detector systems.
Subpart F--Equipment
General
Sec. 29.1301 Function and installation.
Sec. 29.1303 Flight and navigation instruments.
Sec. 29.1305 Powerplant instruments.
Sec. 29.1307 Miscellaneous equipment.
Sec. 29.1309 Equipment, systems, and installations.
Instruments: Installation
Sec. 29.1321 Arrangement and visibility.
Sec. 29.1322 Warning, caution, and advisory lights.
Sec. 29.1323 Airspeed indicating system.
Sec. 29.1325 Static pressure and pressure altimeter systems.
Sec. 29.1327 Magnetic direction indicator.
Sec. 29.1329 Automatic pilot system.
Sec. 29.1331 Instruments using a power supply.
Sec. 29.1333 Instrument systems.
Sec. 29.1335 Flight director systems.
Sec. 29.1337 Powerplant instruments.
Electrical Systems and Equipment
Sec. 29.1351 General.
Sec. 29.1353 Electrical equipment and installations.
Sec. 29.1355 Distribution system.
Sec. 29.1357 Circuit protective devices.
Sec. 29.1359 Electrical system fire and smoke protection.
Sec. 29.1363 Electrical system tests.
Lights
Sec. 29.1381 Instrument lights.
Sec. 29.1383 Landing lights.
Sec. 29.1385 Position light system installation.
Sec. 29.1387 Position light system dihedral angles.
Sec. 29.1389 Position light distribution and intensities.
Sec. 29.1391 Minimum intensities in the horizontal plane of
forward and rear position lights.
Sec. 29.1393 Minimum intensities in any vertical plane of
forward and rear position lights.
Sec. 29.1395 Maximum intensities in overlapping beams of
forward and rear position lights.
Sec. 29.1397 Color specifications.
Sec. 29.1399 Riding light.
Sec. 29.1401 Anticollision light system.
Safety Equipment
Sec. 29.1411 General.
Sec. 29.1413 Safety belts: passenger warning device.
Sec. 29.1415 Ditching equipment.
Sec. 29.1419 Ice protection.
Miscellaneous Equipment
Sec. 29.1431 Electronic equipment.
Sec. 29.1433 Vacuum systems.
Sec. 29.1435 Hydraulic systems.
Sec. 29.1439 Protective breathing equipment.
Sec. 29.1457 Cockpit voice recorders.
Sec. 29.1459 Flight recorders.
Sec. 29.1461 Equipment containing high energy rotors.
Subpart G--Operating Limitations and Information
Sec. 29.1501 General.
Operating Limitations
Sec. 29.1503 Airspeed limitations: general.
Sec. 29.1505 Never-exceed speed.
Sec. 29.1509 Rotor speed.
Sec. 29.1517 Limiting height-speed envelope.
Sec. 29.1519 Weight and center of gravity.
Sec. 29.1521 Powerplant limitations.
Sec. 29.1522 Auxiliary power unit limitations.
Sec. 29.1523 Minimum flight crew.
Sec. 29.1525 Kinds of operations.
Sec. 29.1527 Maximum operating altitude.
Sec. 29.1529 Instructions for continued airworthiness.
Markings and Placards
Sec. 29.1541 General.
Sec. 29.1543 Instrument markings: general.
Sec. 29.1545 Airspeed indicator.
Sec. 29.1547 Magnetic direction indicator.
Sec. 29.1549 Powerplant instruments.
Sec. 29.1551 Oil quantity indicator.
Sec. 29.1553 Fuel quantity indicator.
Sec. 29.1555 Control markings.
Sec. 29.1557 Miscellaneous markings and placards.
Sec. 29.1559 Limitations placard.
Sec. 29.1561 Safety equipment.
Sec. 29.1565 Tail rotor.
Rotorcraft Flight Manual
Sec. 29.1581 General.
Sec. 29.1583 Operating limitations.
Sec. 29.1585 Operating procedures.
Sec. 29.1587 Performance information.
Sec. 29.1589 Loading information.
Appendix A to Part 29--Instructions for Continued Airworthiness
Appendix B to Part 29--Airworthiness Criteria for Helicopter
Instrument Flight
Appendix C to Part 29--Icing Certification
Appendix D--Criteria for Demonstration of Emergency Evacuation
Procedures Under Sec. 29.803
Special Federal Aviation Regulation No. 29-4
Editorial Note: For the text of SFAR No. 29-4, see Part 21 of this chapter.
Subpart A--General
Sec. 29.1 Applicability.
(a) This part prescribes airworthiness standards for the issue of type
certificates, and changes to those certificates, for transport category
rotorcraft.
(b) Transport category rotorcraft must be certificated in accordance with
either the Category A or Category B requirements of this part. A multiengine
rotorcraft may be type certificated as both Category A and Category B with
appropriate and different operating limitations for each category.
(c) Rotorcraft with a maximum weight greater than 20,000 pounds and 10 or
more passenger seats must be type certificated as Category A rotorcraft.
(d) Rotorcraft with a maximum weight greater than 20,000 pounds and nine or
less passenger seats may be type certificated as Category B rotorcraft
provided the Category A requirements of Subparts C, D, E, and F of this part
are met.
(e) Rotorcraft with a maximum weight of 20,000 pounds or less but with 10
or more passenger seats may be type certificated as Category B rotorcraft
provided the Category A requirements of Secs. 29.67(a)(2), 29.79, 29.1517,
and of Subparts C, D, E, and F of this part are met.
(f) Rotorcraft with a maximum weight of 20,000 pounds or less and nine or
less passenger seats may be type certificated as Category B rotorcraft.
(g) Each person who applies under Part 21 for a certificate or change
described in paragraphs (a) through (f) of this section must show compliance
with the applicable requirements of this part.
[Amdt. 29-21, 48 FR 4391, Jan. 31, 1983]
Sec. 29.2 Special retroactive requirements.
For each rotorcraft manufactured after September 16, 1992, each applicant
must show that each occupant's seat is equipped with a safety belt and
shoulder harness that meets the requirements of paragraphs (a), (b), and (c)
of this section.
(a) Each occupant's seat must have a combined safety belt and shoulder
harness with a single-point release. Each pilot's combined safety belt and
shoulder harness must allow each pilot, when seated with safety belt and
shoulder harness fastened, to perform all functions necessary for flight
operations. There must be a means to secure belts and harnesses, when not in
use, to prevent interference with the operation of the rotorcraft and with
rapid egress in an emergency.
(b) Each occupant must be protected from serious head injury by a safety
belt plus a shoulder harness that will prevent the head from contacting any
injurious object.
(c) The safety belt and shoulder harness must meet the static and dynamic
strength requirements, if applicable, specified by the rotorcraft type
certification basis.
(d) For purposes of this section, the date of manufacture is either--
(1) The date the inspection acceptance records, or equivalent, reflect that
the rotorcraft is complete and meets the FAA-Approved Type Design Data; or
(2) The date that the foreign civil airworthiness authority certifies the
rotorcraft is complete and issues an original standard airworthiness
certificate, or equivalent, in that country.
SUMMARY: This final rule amends the airworthiness and operating
regulations to require installation and use of shoulder harnesses at all
seats of rotorcraft manufactured after September 16, 1992. These amendments
respond to a safety recommendation from the National Transportation Safety
Board and are intended to enhance protection of occupants in rotorcraft.
DATES: Effective date: September 16, 1991.
Compliance date: September 16, 1992.
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--
(a) By tests upon a rotorcraft of the type for which certification is
requested, or by calculations based on, and equal in accuracy to, the results
of testing; and
(b) By systematic investigation of each required combination of weight and
center of gravity, if compliance cannot be reasonably inferred from
combinations investigated.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-24, 49 FR
44435, Nov. 6, 1984]
Sec. 29.25 Weight limits.
(a) Maximum weight. The maximum weight (the highest weight at which
compliance with each applicable requirement of this part is shown) or, at the
option of the applicant, the highest weight for each altitude and for each
practicably separable operating condition, such as takeoff, enroute
operation, and landing, must be established so that it is not more than--
(1) The highest weight selected by the applicant;
(2) The design maximum weight (the highest weight at which compliance with
each applicable structural loading condition of this part is shown); or
(3) The highest weight at which compliance with each applicable flight
requirement of this part is shown.
(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 less than--
(1) The lowest weight selected by the applicant;
(2) The design minimum weight (the lowest weight at which compliance with
each structural loading condition of this part is shown); or
(3) The lowest weight at which compliance with each applicable flight
requirement of this part is shown.
(c) Total weight with jettisonable external load. A total weight for the
rotorcraft with jettisonable external load attached that is greater than the
maximum weight established under paragraph (a) of this section may be
established if structural component approval for external load operations
under Part 133 of this chapter is requested and the following conditions are
met:
(1) The portion of the total weight that is greater than the maximum weight
established under paragraph (a) of this section is made up only of the weight
of all or part of the jettisonable external load.
(2) Structural components of the rotorcraft are shown to comply with the
applicable structural requirements of this part under the increased loads and
stresses caused by the weight increase over that established under paragraph
(a) of this section.
(3) Operation of the rotorcraft at a total weight greater than the maximum
certificated weight established under paragraph (a) of this section is
limited by appropriate operating limitations to rotorcraft external load
operations under Part 133 of this chapter.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 FR
55471, Dec. 20, 1976]
Sec. 29.27 Center of gravity limits.
The extreme forward and aft centers of gravity and, where critical, the
extreme lateral centers of gravity must be established for each weight
established under Sec. 29.25. Such an extreme may not lie beyond--
(a) The extremes selected by the applicant;
(b) The extremes within which the structure is proven; or
(c) The extremes within which compliance with the applicable flight
requirements is shown.
[Amdt. 29-3, 33 FR 965, Jan. 26, 1968]
Sec. 29.29 Empty weight and corresponding center of gravity.
(a) The empty weight and corresponding center of gravity must be determined
by weighing the rotorcraft without the crew and payload, but with--
(1) Fixed ballast;
(2) Unusable fuel; and
(3) Full operating fluids, including--
(i) Oil;
(ii) Hydraulic fluid; and
(iii) Other fluids required for normal operation of rotorcraft systems,
except water intended for injection in the engines.
(b) The condition of the rotorcraft at the time of determining empty weight
must be one that is well defined and can be easily repeated, particularly
with respect to the weights of fuel, oil, coolant, and installed equipment.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49
U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of the Dept. of
Transportation Act (49 U.S.C. 1655(c)))
[Doc. No. 5084, 29 FR 16150. Dec. 3, 1964, as amended by Amdt. 29-15, 43 FR
2326, Jan. 16, 1978]
Sec. 29.31 Removable ballast.
Removable ballast may be used in showing compliance with the flight
requirements of this subpart.
Sec. 29.33 Main rotor speed and pitch limits.
(a) Main rotor speed limits. A range of main rotor speeds must be
established that--
(1) With power on, provides adequate margin to accommodate the variations
in rotor speed occurring in any appropriate maneuver, and is consistent with
the kind of governor or synchronizer used; and
(2) With power off, allows each appropriate autorotative maneuver to be
performed throughout the ranges of airspeed and weight for which
certification is requested.
(b) Normal main rotor high pitch limit (power on). For rotorcraft, except
helicopters required to have a main rotor low speed warning under paragraph
(e) of this section, it must be shown, with power on and without exceeding
approved engine maximum limitations, that main rotor speeds substantially
less than the minimum approved main rotor speed will not occur under any
sustained flight condition. This must be met by--
(1) Appropriate setting of the main rotor high pitch stop;
(2) Inherent rotorcraft characteristics that make unsafe low main rotor
speeds unlikely; or
(3) Adequate means to warn the pilot of unsafe main rotor speeds.
(c) Normal main rotor low pitch limit (power off). It must be shown, with
power off, that--
(1) The normal main rotor low pitch limit provides sufficient rotor speed,
in any autorotative condition, under the most critical combinations of weight
and airspeed; and
(2) It is possible to prevent overspeeding of the rotor without exceptional
piloting skill.
(d) Emergency high pitch. If the main rotor high pitch stop is set to meet
paragraph (b)(1) of this section, and if that stop cannot be exceeded
inadvertently, additional pitch may be made available for emergency use.
(e) Main rotor low speed warning for helicopters. For each single engine
helicopter, and each multiengine helicopter that does not have an approved
device that automatically increases power on the operating engines when one
engine fails, there must be a main rotor low speed warning which meets the
following requirements:
(1) The warning must be furnished to the pilot in all flight conditions,
including power-on and power-off flight, when the speed of a main rotor
approaches a value that can jeopardize safe flight.
(2) The warning may be furnished either through the inherent aerodynamic
qualities of the helicopter or by a device.
(3) The warning must be clear and distinct under all conditions, and must
be clearly distinguishable from all other warnings. A visual device that
requires the attention of the crew within the cockpit is not acceptable by
itself.
(4) If a warning device is used, the device must automatically deactivate
and reset when the low-speed condition is corrected. If the device has an
audible warning, it must also be equipped with a means for the pilot to
manually silence the audible warning before the low-speed condition is
corrected.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49
U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of the Dept. of
Transportation Act (49 U.S.C. 1655(c)))
(a) The performance prescribed in this subpart must be determined--
(1) With normal piloting skill and;
(2) Without exceptionally favorable conditions.
(b) Compliance with the performance requirements of this subpart must be
shown--
(1) For still air at sea level with a standard atmosphere and;
(2) For the approved range of atmospheric variables.
(c) The available power must correspond to engine power, not exceeding the
approved power, less--
(1) Installation losses; and
(2) The power absorbed by the accessories and services at the values for
which certification is requested and approved.
(d) For reciprocating engine-powered rotorcraft, the performance, as
affected by engine power, must be based on a relative humidity of 80 percent
in a standard atmosphere.
(e) For turbine engine-powered rotorcraft, the performance, as affected by
engine power, 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.
Between these two temperatures, the relative humidity must vary linearly.
(f) For turbine-engine-power rotorcraft, a means must be provided to permit
the pilot to detemine prior to takeoff that each engine is capable of
developing the power necessary to achieve the applicable rotorcraft
performance prescribed in this subpart.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49
U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. of
Transportation Act (49 U.S.C. 1655(c)))
(a) The takeoff data required by Secs. 29.53(b), 29.59, 29.63, and
29.67(a)(1) and (2) must be determined--
(1) At each weight, altitude, and temperature selected by the applicant;
and
(2) With the operating engines within approved operating limitations.
(b) Takeoff data must--
(1) Be determined on a smooth, dry, hard surface; and
(2) Be corrected to assume a level takeoff surface.
(c) No takeoff made to determine the data required by this section may
require exceptional piloting skill or alertness, or exceptionally favorable
conditions.
Sec. 29.53 Takeoff: Category A.
(a) General. The takeoff performance must be determined and scheduled so
that, if one engine fails at any time after the start of takeoff, the
rotorcraft can--
(1) Return to, and stop safely on, the takeoff area; or
(2) Continue the takeoff and climb-out, and attain a configuration and
airspeed allowing compliance with Sec. 29.67(a)(2).
(b) Critical decision point. The critical decision point must be a
combination of height and speed selected by the applicant in establishing the
flight paths under Sec. 29.59. The critical decision point must be obtained
so as to avoid the critical areas of the limiting height-speed envelope
established under Sec. 29.79.
Sec. 29.59 Takeoff path: Category A.
(a) The takeoff climb-out path, and the rejected takeoff path must be
established so that the takeoff, climb-out, and rejected takeoff are
accomplished with a safe, smooth transition between each stage of the
maneuver. The takeoff may be begun in any manner if--
(1) The takeoff surface is defined; and
(2) Adequate safeguards are maintained to ensure proper center of gravity
and control positions.
(b) The rejected takeoff path must be established with not more than
takeoff power on each engine from the start of takeoff to the critical
decision point, at which point it is assumed that the critical engine becomes
inoperative and that the rotorcraft is brought to a safe stop.
(c) The takeoff climbout path must be established with not more than
takeoff power on each engine from the start of takeoff to the critical
decision point, at which point it is assumed that the critical engine becomes
inoperative and remains inoperative for the rest of the takeoff. The
rotorcraft must be accelerated to achieve the takeoff safety speed and a
height of 35 feet above the ground or greater and the climbout must be made--
(1) At not less than the takeoff safety speed used in meeting the rate of
climb requirements of Sec. 29.67(a)(1); and
(2) So that the airspeed and configuration used in meeting the climb
requirement of Sec. 29.67(a)(2) are attained.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-24, 49 FR
44436, Nov. 6, 1984]
Sec. 29.63 Takeoff: Category B.
The horizontal distance required to take off and climb over a 50-foot
obstacle must be established with the most unfavorable center of gravity. The
takeoff may be begun in any manner if--
(a) The takeoff surface is defined;
(b) Adequate safeguards are maintained to ensure proper center of gravity
and control positions; and
(c) A landing can be made safely at any point along the flight path if an
engine fails.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 FR
55471, Dec. 20, 1976]
Sec. 29.65 Climb: all engines operating.
(a) The steady rate of climb must be determined for each Category B
rotorcraft--
(1) With maximum continuous power on each engine;
(2) With the landing gear retracted;
(3) For the weights, altitudes, and temperatures for which certification is
requested; and
(4) At VY for standard sea level conditions at maximum weight and at speeds
selected by the applicant at or below VNE for other conditions.
(b) For each Category B rotorcraft except helicopters, the rate of climb
determined under paragraph (a) of this section must provide a steady climb
gradient of at least 1:6 under standard sea level conditions.
(c) For Category A helicopters, if VNE at any altitude within the range for
which certification is requested is less than VY at sea level standard
conditions, with maximum weight and maximum continuous power, the steady rate
of climb must be determined--
(1) At the climb speed selected by the applicant at or below VNE;
(2) Within the range from 2,000 feet below the altitude at which VNE is
equal to VY up to the maximum altitude for which certification is requested;
(3) For the weights and temperatures that correspond to the altitude range
set forth in paragraph (c)(2) of this section and for which certification is
requested;
(4) With maximum continuous power on each engine; and
(5) With the landing gear retracted.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49
U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. of
Transportation Act (49 U.S.C. 1655(c)))
[Doc. No. 5084, 29 FR 16150. Dec. 3, 1964, as amended by Amdt. 29-15, 43 FR
2326, Jan. 16, 1978]
Sec. 29.67 Climb: one engine inoperative.
(a) For Category A rotorcraft, the following apply:
(1) The steady rate of climb without ground effect must be at least 100
feet per minute for each weight, altitude, and temperature for which takeoff
and landing data are to be scheduled, with--
(i) The critical engine inoperative and the remaining engines within
approved operating limitations, except that for rotorcraft for which the use
of 30-second/2-minute OEI power is requested, only the 2-minute OEI power may
be used in showing compliance with this paragraph;
(ii) The most unfavorable center of gravity;
(iii) The landing gear extended;
(iv) The takeoff safety speed selected by the applicant; and
(v) Cowl flaps or other means of controlling the engine-cooling air supply
in the position that provides adequate cooling at the temperatures and
altitudes for which certification is requested.
(2) The steady rate of climb without ground effect must be at least 150
feet per minute 1,000 feet above the takeoff and landing surfaces for each
weight, altitude, and temperature for which takeoff and landing data are to
be scheduled, with--
(i) The critical engine inoperative and the remaining engines at--
(A) Maximum continuous power;
(B) Thirty-minute OEI power (for helicopters for which certification for
the use of 30-minute OEI power is requested); or
(C) Continuous OEI power (for helicopters for which certification for the
use of continuous OEI power is requested);
(ii) The most unfavorable center of gravity;
(iii) The landing gear retracted;
(iv) A speed selected by the applicant; and
(v) Cowl flaps, or other means of controlling the engine-cooling air
supply, in the position that provides adequate cooling at the temperatures
and altitudes for which certification is requested.
(3) The steady rate of climb, in feet per minute, at any altitude at which
the rotorcraft is expected to operate, and at any weight within the range of
weights for which certification is requested, must be determined with--
(i) The critical engine inoperative and the remaining engines at--
(A) Maximum continuous power and at 30-minute OEI power (for helicopters
for which certification for use of 30-minute OEI power is requested); or
(B) Continuous OEI power (for helicopters for which certification for the
use of continuous OEI power is requested);
(ii) The most unfavorable center of gravity;
(iii) The landing gear retracted;
(iv) The speed selected by the applicant; and
(v) Cowl flaps or other means of controlling the engine-cooling air supply
in the position that provides adequate cooling at the temperatures and
altitudes for which certification is requested.
(b) For multiengine Category B helicopters meeting the requirements for
Category A in Sec. 29.79, the steady rate of climb (or descent) must be
determined at the speed for the best rate of climb (or minimum rate of
descent) with the critical engine inoperative and the remaining engines at
either--
(1) Maximum continuous power and at 30-minute OEI power (for helicopters
for which certification for the use of 30-minute OEI power is requested); or
(2) Continuous OEI power (for helicopters for which certification for the
use of continuous OEI power is requested).
SUMMARY: This rule adopts new and revised airworthiness standards by
incorporating optional one-engine-inoperative (OEI) power ratings for
multiengine, turbine-powered rotorcraft. These amendments result from a
petition for rulemaking from Aerospace Industries Association of America
(AIA) and the recognition by both government and industry that additional OEI
power rating standards are needed. These amendments enhance rotorcraft safety
after an engine failure or precautionary shutdown by providing higher OEI
power, when necessary. These amendments also assure that the drive system
will maintain its structural integrity and allow continued safe flight while
operating at the new OEI power ratings with the operable engine(s).
For each category B helicopter, except multiengine helicopters meeting the
requirements of Sec. 29.67(b) and the powerplant installation requirements of
category A, the steady angle of glide must be determined in autorotation--
(a) At the forward speed for minimum rate of descent as selected by the
applicant;
(b) At the forward speed for best glide angle;
(c) At maximum weight; and
(d) At the rotor speed or speeds selected by the applicant.
[Amdt. 29-12, 41 FR 55471, Dec. 20, 1976]
Sec. 29.73 Performance at minimum operating speed.
(a) For each category A rotorcraft, the hovering performance must be
determined over the ranges of weight, altitude, and temperature for which
takeoff data are scheduled--
(1) With not more than takeoff power on each engine;
(2) With the landing gear extended; and
(3) At a height consistent with the procedure used in establishing the
takeoff climbout and rejected takeoff paths.
(b) For each category B helicopter--
(1) The hovering performance must be determined over the ranges of weight,
altitude, and temperature for which certification is requested, with--
(i) Takeoff power on each engine;
(ii) The landing gear extended; and
(iii) The helicopter in ground effect at a height consistent with normal
takeoff procedures; and
(2) The hovering ceiling determined under paragraph (b)(1) of this
section--
(i) For reciprocating engine powered helicopters, must be at least 4,000
feet in standard atmosphere at maximum weight;
(ii) For single engine, turbine engine powered helicopters, must be at
least 2,500 feet, in standard atmosphere plus 40 deg. F., at maximum weight;
and
(iii) For multiengine, turbine engine power helicopters, must be available
at each altitude, temperature, and weight for which takeoff data are to be
scheduled.
(c) For rotorcraft other than helicopters, the steady rate of climb at the
minimum operating speed must be determined, over the ranges of weight,
altitude, and temperature for which certification is requested, with--
(1) Takeoff power; and
(2) The landing gear extended.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 FR
965, Jan. 26, 1968]
Sec. 29.75 Landing.
(a) General. For each rotorcraft--
(1) The corrected landing data must--
(i) Be determined on a smooth, dry, hard surface; and
(ii) Assume a level landing surface;
(2) The approach and landing may not require exceptional piloting skill or
exceptionally favorable conditions;
(3) The landing must be made without excessive vertical acceleration or
tendency to bounce, nose over, ground loop, porpoise, or water loop; and
(4) The landing data required by paragraphs (b) and (c) of this section and
by Sec. 29.77 must be determined--
(i) At each weight, altitude, and temperature selected by the applicant;
and
(ii) With each operating engine within approved operating limitations.
(b) Category A. For category A rotorcraft--
(1) The landing performance must be determined and scheduled so that, if
one engine fails at any point in the approach path, the rotorcraft can either
land and stop safely or climb out from a point in the approach path and
attain a rotorcraft configuration and speed allowing compliance with the
climb requirement of Sec. 29.67(a)(2);
(2) The approach and landing paths must be established, with one engine
inoperative, so that the transition between each stage can be made smoothly
and safely;
(3) The approach and landing speeds must be selected by the applicant and
must be appropriate to the type of rotorcraft;
(4) The approach and landing path must be established to avoid the critical
areas of a limiting height-speed envelope established--
(i) Under Sec. 29.79; or
(ii) For the landing condition with one engine inoperative;
(5) It must be possible to make a safe landing on a prepared landing
surface after complete power failure occurring during normal cruise; and
(6) The horizontal distance required to land and come to a complete stop
(or to a speed of approximately three knots for water landings), from a point
50 feet above the landing surface, must be determined from the approach and
landing paths established in accordance with paragraphs (b)(2) through (b)(4)
of this section.
(c) Category B. For category B rotorcraft--
(1) The horizontal distance required to land and come to a complete stop
(or to a speed of approximately three knots for water landings), from a point
50 feet above the landing surface, must be determined with--
(i) Glide speeds appropriate to the type of rotorcraft and chosen by the
applicant; and
(ii) The approach and landing made with power off and entered from steady
autorotation; and
(2) Each multiengine category B rotorcraft that meets the powerplant
installation requirements for category A must meet the requirements of--
(i) Paragraph (c)(1) of this section; or
(ii) Paragraphs (b)(2) through (b)(6) of this section.
(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C.
1354(a), 1421, 1423, 1424), and sec. 6(c), Dept. of Transportation Act (49
U.S.C. 1655(c)))
For category A rotorcraft, the balked landing path must be established so
that--
(a) With one engine inoperative, the transition from each stage of the
maneuver to the next stage can be made smoothly and safely;
(b) From a combination of height and speed in the approach path selected by
the applicant, a safe climbout can be made at speeds allowing compliance with
the climb requirements of Sec. 29.67(a)(1) and (2); andturn off (c) The
rotorcraft does not descend below 35 feet above the landing surface
in the maneuver described in paragraph (b) of this section.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-24, 49 FR
44436, Nov. 6, 1984]
Sec. 29.79 Limiting height-speed envelope.
(a) If there is any combination of height and forward speed (including
hover) under which a safe landing cannot be made under the applicable power
failure condition in paragraph (b) of this section, a limiting height-speed
envelope must be established for--
(1) Category A. Combinations of weight, pressure altitude, and ambient
temperature for which takeoff and landing are approved; and
(2) Category B. (i) Altitude, from standard sea level conditions to the
maximum altitude for which takeoff and landing are approved; and
(ii) Weight, from the maximum weight (at sea level) to the highest weight
approved for takeoff and landing at each altitude. For helicopters, this
weight need not exceed the highest weight allowing hovering out-of-ground-
effect at each altitude.
(b) The applicable power failure conditions are--
(1) For category A rotorcraft, sudden failure of the critical engine with
the remaining engines at the greatest power for which certification is
requested;
(2) For category B rotorcraft, complete power failure, and
(3) For multiengine, category B rotorcraft for which certification under
the powerplant installation requirements of category A is requested, the
condition specified in either paragraph (b)(1) or (2) of this section.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-1, 30 FR
8778, July 13, 1965; Amdt. 29-21, 48 FR 4391, Jan. 31, 1983]
Flight Characteristics
Sec. 29.141 General.
The rotorcraft must--
(a) Except as specifically required in the applicable section, meet the
flight characteristics requirements of this subpart--
(1) At the approved operating altitudes and temperatures;
(2) Under any critical loading condition within the range of weights and
centers of gravity for which certification is requested; and
(3) For power-on operations, under any condition of speed, power, and rotor
r.p.m. for which certification is requested; and
(4) For power-off operations, under any condition of speed, and rotor
r.p.m. for which certification is requested that is attainable with the
controls rigged in accordance with the approved rigging instructions and
tolerances;
(b) Be able to maintain any required flight condition and make a smooth
transition from any flight condition to any other flight condition without
exceptional piloting skill, alertness, or strength, and without danger of
exceeding the limit load factor under any operating condition probable for
the type, including--
(1) Sudden failure of one engine, for multiengine rotorcraft meeting
Transport Category A engine isolation requirements;
(2) Sudden, complete power failure, for other rotorcraft; and
(3) Sudden, complete control system failures specified in Sec. 29.695 of
this part; and
(c) Have any additional characteristics required for night or instrument
operation, if certification for those kinds of operation is requested.
Requirements for helicopter instrument flight are contained in Appendix B of
this part.
(a) The rotorcraft must be safely controllable and maneuverable--
(1) During steady flight; and
(2) During any maneuver appropriate to the type, including--
(i) Takeoff;
(ii) Climb;
(iii) Level flight;
(iv) Turning flight;
(v) Glide; and
(vi) Landing (power on and power off).
(b) The margin of cyclic control must allow satisfactory roll and pitch
control at VNE w@h--
(1) Critical weight;
(2) Critical center of gravity;
(3) Critical rotor r.p.m.; and
(4) Power off (except for helicopters demonstrating compliance with
paragraph (e) of this section) and power on.
(c) A wind velocity of not less than 17 knots must be established in which
the rotorcraft can be operated without loss of control on or near the ground
in any maneuver appropriate to the type (such as crosswind takeoffs, sideward
flight, and rearward flight), with--
(1) Critical weight;
(2) Critical center of gravity; and
(3) Critical rotor r.p.m.
(d) The rotorcraft, after (1) failure of one engine, in the case of
multiengine rotorcraft that meet Transport Category A engine isolation
requirements, or (2) complete power failure in the case of other rotorcraft,
must be controllable over the range of speeds and altitudes for which
certification is requested when such power failure occurs with maximum
continuous power and critical weight. No corrective action time delay for any
condition following power failure may be less than--
(i) For the cruise condition, one second, or normal pilot reaction time
(whichever is greater); and
(ii) For any other condition, normal pilot reaction time.
(e) For helicopters for which a VNE (power-off) is established under Sec.
29.1505(c), compliance must be demonstrated with the following requirements
with critical weight, critical center of gravity, and critical rotor r.p.m.:
(1) The helicopter must be safely slowed to VNE (power-off), without
exceptional pilot skill after the last operating engine is made inoperative
at power-on VNE.
(2) At a speed of 1.1 VNE (power-off), the margin of cyclic control must
allow satisfactory roll and pitch control with power off.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49
U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of the Dept. of
Transportation Act (49 U.S.C. 1655(c)))
(a) Longitudinal, lateral, directional, and collective controls may not
exhibit excessive breakout force, friction, or preload.
(b) Control system forces and free play may not inhibit a smooth, direct
rotorcraft response to control system input.
[Amdt. 29-24, 49 FR 44436, Nov. 6, 1984]
Sec. 29.161 Trim control.
The trim control--
(a) Must trim any steady longitudinal, lateral, and collective control
forces to zero in level flight at any appropriate speed; and
(b) May not introduce any undesirable discontinuities in control force
gradients.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-24, 49 FR
44436, Nov. 6, 1984]
Sec. 29.171 Stability: general.
The rotorcraft must be able to be flown, without undue pilot fatigue or
strain, in any normal maneuver for a period of time as long as that expected
in normal operation. At least three landings and takeoffs must be made during
this demonstration.
Sec. 29.173 Static longitudinal stability.
(a) The longitudinal control must be designed so that a rearward movement
of the control is necessary to obtain a speed less than the trim speed, and a
forward movement of the control is necessary to obtain a speed more than the
trim speed.
(b) With the throttle and collective pitch held constant during the
maneuvers specified in Sec. 29.175 (a) through (c), the slope of the control
position versus speed curve must be positive throughout the full range of
altitude for which certification is requested.
(c) During the maneuver specified in Sec. 29.175(d), the longitudinal
control position versus speed curve may have a negative slope within the
specified speed range if the negative motion is not greater than 10 percent
of total control travel.
[Amdt. 29-24, 49 FR 44436, Nov. 6, 1984]
Sec. 29.175 Demonstration of static longitudinal stability.
(a) Climb. Static longitudinal stability must be shown in the climb
condition at speeds from 0.85 VY, or 15 knots below VY, whichever is less, to
1.2 VY or 15 knots above VY, whichever is greater, with--
(1) Critical weight;
(2) Critical center of gravity;
(3) Maximum continuous power;
(4) The landing gear retracted; and
(5) The rotorcraft trimmed at VY.
(b) Cruise. Static longitudinal stability must be shown in the cruise
condition at speeds from 0.7 VH or 0.7 VNE, whichever is less, to 1.1 VH or
1.1 VNE, whichever is less, with--
(1) Critical weight;
(2) Critical center of gravity;
(3) Power for level flight at 0.9 VH or 0.9 VNE, whichever is less;
(4) The landing gear retracted, and
(5) The rotorcraft trimmed at 0.9 VH or 0.9 VNE, whichever is less.
(c) Autorotation. Static longitudinal stability must be shown in
autorotation at airspeeds from 0.5 times the speed for minimum rate of
descent, or 0.5 times the maximum range glide speed for Category A
rotorcraft, to VNE or to 1.1 VNE (power-off) if VNE (power-off) is
established under Sec. 29.1505(c), and with--
(1) Critical weight;
(2) Critical center of gravity;
(3) Power off;
(4) The landing gear----
(i) Retracted; and
(ii) Extended; and
(5) The rotorcraft trimmed at appropriate speeds found necessary by the
Administrator to demonstrate stability throughout the prescribed speed range.
(d) Hovering. For helicopters, the longitudinal cyclic control must operate
with the sense, direction of motion, and position as prescribed in Sec.
29.173 between the maximum approved rearward speed and a forward speed of 17
knots with--
(1) Critical weight;
(2) Critical center of gravity;
(3) Power required to maintain an approximate constant height in ground
effect;
(4) The landing gear extended; and
(5) The helicopter trimmed for hovering.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49
U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. of
Transportation Act (49 U.S.C. 1655(c)))
Static directional stability must be positive with throttle and collective
controls held constant at the trim conditions specified in Sec. 29.175 (a),
(b), and (c). Sideslip angle must increase steadily with directional control
deflection for sideslip angles up to +/-10 deg. from trim. Sufficient cues
must accompany sideslip to alert the pilot when approaching sideslip limits.
[Amdt. 29-24, 49 FR 44436, Nov. 6, 1984]
Sec. 29.181 Dynamic stability: Category A rotorcraft.
Any short-period oscillation occurring at any speed from VY to VNE must be
positively damped with the primary flight controls free and in a fixed
position.
[Amdt. 29-24, 49 FR 44437, Nov. 6, 1984]
Ground and Water Handling Characteristics
Sec. 29.231 General.
The rotorcraft must have satisfactory ground and water handling
characteristics, including freedom from uncontrollable tendencies in any
condition expected in operation.
Sec. 29.235 Taxiing condition.
The rotorcraft must be designed to withstand the loads that would occur
when the rotorcraft is taxied over the roughest ground that may reasonably be
expected in normal operation.
Sec. 29.239 Spray characteristics.
If certification for water operation is requested, no spray characteristics
during taxiing, takeoff, or landing may obscure the vision of the pilot or
damage the rotors, propellers, or other parts of the rotorcraft.
Sec. 29.241 Ground resonance.
The rotorcraft may have no dangerous tendency to oscillate on the ground
with the rotor turning.
Miscellaneous Flight Requirements
Sec. 29.251 Vibration.
Each part of the rotorcraft must be free from excessive vibration under
each appropriate speed and power condition.
Subpart C--Strength Requirements
General
Sec. 29.301 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 specified air, ground, and water loads
must be placed in equilibrium with inertia forces, considering each item of
mass in the rotorcraft. These loads must be distributed to closely
approximate or conservatively represent actual conditions.
(c) If deflections under load would significantly change the distribution
of external or internal loads, this redistribution must be taken into
account.
Sec. 29.303 Factor of safety.
Unless otherwise provided, a factor of safety of 1.5 must be used. This
factor applies to external and inertia loads unless its application to the
resulting internal stresses is more conservative.
Sec. 29.305 Strength and deformation.
(a) The structure must be able to support limit loads without detrimental
or 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.
This must be shown by--
(1) Applying ultimate loads to the structure in a static test for at least
three seconds; or
(2) Dynamic tests simulating actual load application.
Sec. 29.307 Proof of structure.
(a) Compliance with the strength and deformation requirements of this
subpart must be shown for each critical loading condition accounting for the
environment to which the structure will be exposed in operation. Structural
analysis (static or fatigue) may be used only if the structure conforms to
those structures for which experience has shown this method to be reliable.
In other cases, substantiating load tests must be made.
(b) Proof of compliance with the strength requirements of this subpart must
include--
(1) Dynamic and endurance tests of rotors, rotor drives, and rotor
controls;
(2) Limit load tests of the control system, including control surfaces;
(3) Operation tests of the control system;
(4) Flight stress measurement tests;
(5) Landing gear drop tests; and
(6) Any additional tests required for new or unusual design features.
SUMMARY: This rule adopts new and revised airworthiness standards for
certification of airframe and related equipment on both normal and transport
category rotorcraft. In addition, one amendment changes an operating rule
affecting external load operators. These amendments grew out of a rotorcraft
regulatory review program and the recognition by both government and industry
that updated safety standards are needed. These amendments provide a high
level of safety in design requirements, while removing certain unnecessary
existing burdens and better utilizing the unique characteristics and
capabilities of rotorcraft.
The following values and limitations must be established to show compliance
with the structural requirements of this subpart:
(a) The design maximum and design minimum weights.
(b) The main rotor r.p.m. ranges, power on and power off.
(c) The maximum forward speeds for each main rotor r.p.m. within the ranges
determined under paragraph (b) of this section.
(d) The maximum rearward and sideward flight speeds.
(e) The center of gravity limits corresponding to the limitations
determined under paragraphs (b), (c), and (d) of this section.
(f) The rotational speed ratios between each powerplant and each connected
rotating component.
(g) The positive and negative limit maneuvering load factors.
Flight Loads
Sec. 29.321 General.
(a) The flight load factor must be assumed to act normal to the
longitudinal axis of the rotorcraft, and to be equal in magnitude and
opposite in direction to the rotorcraft inertia load factor at the center of
gravity.
(b) Compliance with the flight load requirements of this subpart must be
shown--
(1) At each weight from the design minimum weight to the design maximum
weight; and
(2) With any practical distribution of disposable load within the operating
limitations in the Rotorcraft Flight Manual.
Sec. 29.337 Limit maneuvering load factor.
The rotorcraft must be designed for--
(a) A limit maneuvering load factor ranging from a positive limit of 3.5 to
a negative limit of -1.0; or
(b) Any positive limit maneuvering load factor not less than 2.0 and any
negative limit maneuvering load factor of not less than -0.5 for which--
(1) The probability of being exceeded is shown by analysis and flight tests
to be extremely remote; and
(2) The selected values are appropriate to each weight condition between
the design maximum and design minimum weights.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended at Amdt. 29-30, 55 FR
8002, Mar. 6, 1990]
SUMMARY: This rule adopts new and revised airworthiness standards for
certification of airframe and related equipment on both normal and transport
category rotorcraft. In addition, one amendment changes an operating rule
affecting external load operators. These amendments grew out of a rotorcraft
regulatory review program and the recognition by both government and industry
that updated safety standards are needed. These amendments provide a high
level of safety in design requirements, while removing certain unnecessary
existing burdens and better utilizing the unique characteristics and
capabilities of rotorcraft.
The loads resulting from the application of limit maneuvering load factors
are assumed to act at the center of each rotor hub and at each auxiliary
lifting surface, and to act in directions and with distributions of load
among the rotors and auxiliary lifting surfaces, so as to represent each
critical maneuvering condition, including power-on and power-off flight with
the maximum design rotor tip speed ratio. The rotor tip speed ratio is the
ratio of the rotorcraft flight velocity component in the plane of the rotor
disc to the rotational tip speed of the rotor blades, and is expressed as
follows:
V cos a
<mu> = --------
VR
where--
V=The airspeed along the flight path (f.p.s.);
a=The angle between the projection, in the plane of symmetry, of the axis of
no feathering and a line perpendicular to the flight path (radians,
positive when axis is pointing aft);
V=The angular velocity of rotor (radians per second); and
R =The rotor radius (ft.).
Sec. 29.341 Gust loads.
Each rotorcraft must be designed to withstand, at each critical airspeed
including hovering, the loads resulting from vertical and horizontal gusts of
30 feet per second.
Sec. 29.351 Yawing conditions.
(a) Each rotorcraft must be designed for the loads resulting from the
maneuvers specified in paragraphs (b) and (c) of this section, with--
(1) Unbalanced aerodynamic moments about the center of gravity which the
aircraft reacts to in a rational or conservative manner considering the
principal masses furnishing the reacting inertia forces; and
(2) Maximum main rotor speed.
(b) To produce the load required in paragraph (a) of this section, in
unaccelerated flight with zero yaw, at forward speeds from zero up to 0.6
VNE--
(1) Displace the cockpit control suddenly to the maximum deflection limited
by the control stops or by the maximum pilot force specified in Sec.
29.395(a);
(2) Attain a resulting sideslip angle or 90 deg., whichever is less; and
(3) Return the directional control suddenly to neutral.
(c) To produce the load required in paragraph (a) of the section, in
unaccelerated flight with zero yaw, at forward speeds from 0.6 VNE up to VNE
or VH, whichever is less--
(1) Displace the cockpit directional control suddenly to the maximum
deflection limited by the control stops or by the pilot force specified in
Sec. 29.395(a);
(2) Attain a resulting sideslip angle or 15 deg., whichever is less, at the
lesser speed of VNE or VH;
(3) Vary the sideslip angles of paragraphs (b)(2) and (c)(2) of this
section directly with speed; and
(4) Return the directional control suddenly to neutral.
SUMMARY: This rule adopts new and revised airworthiness standards for
certification of airframe and related equipment on both normal and transport
category rotorcraft. In addition, one amendment changes an operating rule
affecting external load operators. These amendments grew out of a rotorcraft
regulatory review program and the recognition by both government and industry
that updated safety standards are needed. These amendments provide a high
level of safety in design requirements, while removing certain unnecessary
existing burdens and better utilizing the unique characteristics and
capabilities of rotorcraft.
The limit engine torque may not be less than the following:
(a) For turbine engines, the highest of--
(1) The mean torque for maximum continuous power multiplied by 1.25;
(2) The torque required by Sec. 29.923;
(3) The torque required by Sec. 29.927; or
(4) The torque imposed by sudden engine stoppage due to malfunction or
structural failure (such as compressor jamming).
(b) For reciprocating engines, the mean torque for maximum continuous power
multiplied by--
(1) 1.33, for engines with five or more cylinders; and
(2) Two, three, and four, for engines with four, three, and two cylinders,
respectively.
[Amdt. 29-26, 53 FR 34215, Sept. 2, 1988]
Control Surface and System Loads
Sec. 29.391 General.
Each auxiliary rotor, each fixed or movable stabilizing or control surface,
and each system operating any flight control must meet the requirements of
Secs. 29.395 through 29.403, 29.411, 29.413, and 29.427.
SUMMARY: This rule adopts new and revised airworthiness standards for
certification of airframe and related equipment on both normal and transport
category rotorcraft. In addition, one amendment changes an operating rule
affecting external load operators. These amendments grew out of a rotorcraft
regulatory review program and the recognition by both government and industry
that updated safety standards are needed. These amendments provide a high
level of safety in design requirements, while removing certain unnecessary
existing burdens and better utilizing the unique characteristics and
capabilities of rotorcraft.
(a) The reaction to the loads prescribed in Sec. 29.397 must be provided
by--
(1) The control stops only;
(2) The control locks only;
(3) The irreversible mechanism only (with the mechanism locked and with the
control surface in the critical positions for the effective parts of the
system within its limit of motion);
(4) The attachment of the control system to the rotor blade pitch control
horn only (with the control in the critical positions for the affected parts
of the system within the limits of its motion); and
(5) The attachment of the control system to the control surface horn (with
the control in the critical positions for the affected parts of the system
within the limits of its motion).
(b) Each primary control system, including its supporting structure, must
be designed as follows:
(1) The system must withstand loads resulting from the limit pilot forces
prescribed in Sec. 29.397;
(2) Notwithstanding paragraph (b)(3) of this section, when power-operated
actuator controls or power boost controls are used, the system must also
withstand the loads resulting from the limit pilot forces prescribed in Sec.
29.397 in conjunction with the forces output of each normally energized power
device, including any single power boost or actuator system failure;
(3) If the system design or the normal operating loads are such that a part
of the system cannot react to the limit pilot forces prescribed in Sec.
29.397, that part of the system must be designed to withstand the maximum
loads that can be obtained in normal operation. The minimum design loads
must, in any case, provide a rugged system for service use, including
consideration of fatigue, jamming, ground gusts, control inertia, and
friction loads. In the absence of a rational analysis, the design loads
resulting from 0.60 of the specified limit pilot forces are acceptable
minimum design loads; and
(4) If operational loads may be exceeded through jamming, ground gusts,
control inertia, or friction, the system must withstand the limit pilot
forces specified in Sec. 29.397, without yielding.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended at Amdt. 29-30, 55 FR
8002, Mar. 6, 1990]
SUMMARY: This rule adopts new and revised airworthiness standards for
certification of airframe and related equipment on both normal and transport
category rotorcraft. In addition, one amendment changes an operating rule
affecting external load operators. These amendments grew out of a rotorcraft
regulatory review program and the recognition by both government and industry
that updated safety standards are needed. These amendments provide a high
level of safety in design requirements, while removing certain unnecessary
existing burdens and better utilizing the unique characteristics and
capabilities of rotorcraft.
(a) Except as provided in paragraph (b) of this section, the limit pilot
forces are as follows:
(1) For foot controls, 130 pounds.
(2) For stick controls, 100 pounds fore and aft, and 67 pounds laterally.
(b) For flap, tab, stabilizer, rotor brake, and landing gear operating
controls, the following apply (R=radius in inches):
(1) Crank wheel, and lever controls, [1 + R]/3 x 50 pounds, but not less
than 50 pounds nor more than 100 pounds for hand operated controls or 130
pounds for foot operated controls, applied at any angle within 20 degrees of
the plane of motion of the control.
(2) Twist controls, 80R pounds.
[Amdt. 29-12, 41 FR 55471, Dec. 20, 1976]
Sec. 29.399 Dual control system.
Each dual primary flight control system must be able to withstand the loads
that result when pilot forces not less than 0.75 times those obtained under
Sec. 29.395 are applied--
(a) In opposition; and
(b) In the same direction.
SUMMARY: This rule amends the airworthiness standards for systems,
propulsion, and airframe for both normal and transport category rotorcraft.
In addition, these amendments introduce safety improvements, clarifying
existing regulations, and standardize terminology. The changes are based on
some of the proposals that were submitted to the FAA by the European
Airworthiness Authorities. These amendments are also intended to encourage
the European community's acceptance of the Federal Aviation Regulations for
rotorcraft type certification, obviate development of different European
standards, and achieve increased commonality of airworthiness standards
among the respective countries.
SUMMARY: This rule amends the airworthiness standards for systems,
propulsion, and airframe for both normal and transport category rotorcraft.
In addition, these amendments introduce safety improvements, clarifying
existing regulations, and standardize terminology. The changes are based on
some of the proposals that were submitted to the FAA by the European
Airworthiness Authorities. These amendments are also intended to encourage
the European community's acceptance of the Federal Aviation Regulations for
rotorcraft type certification, obviate development of different European
standards, and achieve increased commonality of airworthiness standards
among the respective countries.
(a) It must be impossible for the tail rotor to contact the landing surface
during a normal landing.
(b) If a tail rotor guard is required to show compliance with paragraph (a)
of this section--
(1) Suitable design loads must be established for the guard: and
(2) The guard and its supporting structure must be designed to withstand
those loads.
SUMMARY: This rule amends the airworthiness standards for systems,
propulsion, and airframe for both normal and transport category rotorcraft.
In addition, these amendments introduce safety improvements, clarifying
existing regulations, and standardize terminology. The changes are based on
some of the proposals that were submitted to the FAA by the European
Airworthiness Authorities. These amendments are also intended to encourage
the European community's acceptance of the Federal Aviation Regulations for
rotorcraft type certification, obviate development of different European
standards, and achieve increased commonality of airworthiness standards
among the respective countries.
(a) Horizontal tail surfaces and their supporting structure must be
designed for unsymmetrical loads arising from yawing and rotor wake effects
in combination with the prescribed flight conditions.
(b) To meet the design criteria of paragraph (a) of this section, in the
absence of more rational data, both of the following must be met:
(1) One hundred percent of the maximum loading from the symmetrical flight
conditions acts on the surface on one side of the plane of symmetry, and no
loading acts on the other side.
(2) Fifty percent of the maximum loading from the symmetrical flight
conditions acts on the surface on each side of the plane of symmetry, in
opposite directions.
(c) For empennage arrangements where the horizontal tail surfaces are
supported by the vertical tail surfaces, the vertical tail surfaces and
supporting structure must be designed for the combined vertical and
horizontal surface loads resulting from each prescribed flight condition,
considered separately. The flight conditions must be selected so that the
maximum design loads are obtained on each surface. In the absence of more
rational data, the unsymmetrical horizontal tail surface loading
distributions described in this section must be assumed.
[Doc. No. 25570, Amdt. 29-30, 55 FR 8002, Mar. 6, 1990, as amended by Amdt.
29-31, 55 FR 38966, Sept. 21, 1990]
SUMMARY: This rule amends the airworthiness standards for systems,
propulsion, and airframe for both normal and transport category rotorcraft.
In addition, these amendments introduce safety improvements, clarifying
existing regulations, and standardize terminology. The changes are based on
some of the proposals that were submitted to the FAA by the European
Airworthiness Authorities. These amendments are also intended to encourage
the European community's acceptance of the Federal Aviation Regulations for
rotorcraft type certification, obviate development of different European
standards, and achieve increased commonality of airworthiness standards
among the respective countries.
(a) Loads and equilibrium. For limit ground loads--
(1) The limit ground loads obtained in the landing conditions in this part
must be considered to be external loads that would occur in the rotorcraft
structure if it were acting as a rigid body; and
(2) In each specified landing condition, the external loads must be placed
in equilibrium with linear and angular inertia loads in a rational or
conservative manner.
(b) Critical centers of gravity. The critical centers of gravity within the
range for which certification is requested must be selected so that the
maximum design loads are obtained in each landing gear element.
Sec. 29.473 Ground loading conditions and assumptions.
(a) For specified landing conditions, a design maximum weight must be used
that is not less than the maximum weight. A rotor lift may be assumed to act
through the center of gravity throughout the landing impact. This lift may
not exceed two-thirds of the design maximum weight.
(b) Unless otherwise prescribed, for each specified landing condition, the
rotorcraft must be designed for a limit load factor of not less than the
limit inertia load factor substantiated under Sec. 29.725.
(c) Triggering or actuating devices for additional or supplementary energy
absorption may not fail under loads established in the tests prescribed in
Secs. 29.725 and 29.727, but the factor of safety prescribed in Sec. 29.303
need not be used.
[Amdt. 29-3, 33 FR 966, Jan. 26, 1968]
Sec. 29.475 Tires and shock absorbers.
Unless otherwise prescribed, for each specified landing condition, the
tires must be assumed to be in their static position and the shock absorbers
to be in their most critical position.
Sec. 29.477 Landing gear arrangement.
Sections 29.235, 29.479 through 29.485, and 29.493 apply to landing gear
with two wheels aft, and one or more wheels forward, of the center of
gravity.
Sec. 29.479 Level landing conditions.
(a) Attitudes. Under each of the loading conditions prescribed in paragraph
(b) of this section, the rotorcraft is assumed to be in each of the following
level landing attitudes:
(1) An attitude in which each wheel contacts the ground simultaneously.
(2) An attitude in which the aft wheels contact the ground with the forward
wheels just clear of the ground.
(b) Loading conditions. The rotorcraft must be designed for the following
landing loading conditions:
(1) Vertical loads applied under Sec. 29.471.
(2) The loads resulting from a combination of the loads applied under
paragraph (b)(1) of this section with drag loads at each wheel of not less
than 25 percent of the vertical load at that wheel.
(3) The vertical load at the instant of peak drag load combined with a drag
component simulating the forces required to accelerate the wheel rolling
assembly up to the specified ground speed, with--
(i) The ground speed for determination of the spin-up loads being at least
75 percent of the optimum forward flight speed for minimum rate of descent in
autorotation; and
(ii) The loading conditions of paragraph (b) applied to the landing gear
and its attaching structure only.
(4) If there are two wheels forward, a distribution of the loads applied to
those wheels under paragraphs (b)(1) and (2) of this section in a ratio of
40:60.
(c) Pitching moments. Pitching moments are assumed to be resisted by--
(1) In the case of the attitude in paragraph (a)(1) of this section, the
forward landing gear; and
(2) In the case of the attitude in paragraph (a)(2) of this section, the
angular inertia forces.
Sec. 29.481 Tail-down landing conditions.
(a) The rotorcraft is assumed to be in the maximum nose-up attitude
allowing ground clearance by each part of the rotorcraft.
(b) In this attitude, ground loads are assumed to act perpendicular to the
ground.
Sec. 29.483 One-wheel landing conditions.
For the one-wheel landing condition, the rotorcraft is assumed to be in the
level attitude and to contact the ground on one aft wheel. In this attitude--
(a) The vertical load must be the same as that obtained on that side under
Sec. 29.479(b)(1); and
(b) The unbalanced external loads must be reacted by rotorcraft inertia.
Sec. 29.485 Lateral drift landing conditions.
(a) The rotorcraft is assumed to be in the level landing attitude, with--
(1) Side loads combined with one-half of the maximum ground reactions
obtained in the level landing conditions of Sec. 29.479(b)(1); and
(2) The loads obtained under paragraph (a)(1) of this section applied--
(i) At the ground contact point; or
(ii) For full-swiveling gear, at the center of the axle.
(b) The rotorcraft must be designed to withstand, at ground contact--
(1) When only the aft wheels contact the ground, side loads of 0.8 times
the vertical reaction acting inward on one side and 0.6 times the vertical
reaction acting outward on the other side, all combined with the vertical
loads specified in paragraph (a) of this section; and
(2) When the wheels contact the ground simultaneously--
(i) For the aft wheels, the side loads specified in paragraph (b)(1) of
this section; and
(ii) For the forward wheels, a side load of 0.8 times the vertical reaction
combined with the vertical load specified in paragraph (a) of this section.
Sec. 29.493 Braked roll conditions.
Under braked roll conditions with the shock absorbers in their static
positions--
(a) The limit vertical load must be based on a load factor of at least--
(1) 1.33, for the attitude specified in Sec. 29.479(a)(1); and
(2) 1.0, for the attitude specified in Sec. 29.479(a)(2); and
(b) The structure must be designed to withstand, at the ground contact
point of each wheel with brakes, a drag load of at least the lesser of--
(1) The vertical load multiplied by a coefficient of friction of 0.8; and
(2) The maximum value based on limiting brake torque.
Sec. 29.497 Ground loading conditions: landing gear with tail wheels.
(a) General. Rotorcraft with landing gear with two wheels forward and one
wheel aft of the center of gravity must be designed for loading conditions as
prescribed in this section.
(b) Level landing attitude with only the forward wheels contacting the
ground. In this attitude--
(1) The vertical loads must be applied under Secs. 29.471 through 29.475;
(2) The vertical load at each axle must be combined with a drag load at
that axle of not less than 25 percent of that vertical load; and
(3) Unbalanced pitching moments are assumed to be resisted by angular
inertia forces.
(c) Level landing attitude with all wheels contacting the ground
simultaneously. In this attitude, the rotorcraft must be designed for
landing loading conditions as prescribed in paragraph (b) of this section.
(d) Maximum nose-up attitude with only the rear wheel contacting the
ground. The attitude for this condition must be the maximum nose-up attitude
expected in normal operation, including autorotative landings. In this
attitude--
(1) The appropriate ground loads specified in paragraph (b)(1) and (2) of
this section must be determined and applied, using a rational method to
account for the moment arm between the rear wheel ground reaction and the
rotorcraft center of gravity; or
(2) The probability of landing with initial contact on the rear wheel must
be shown to be extremely remote.
(e) Level landing attitude with only one forward wheel contacting the
ground. In this attitude, the rotorcraft must be designed for ground loads
as specified in paragraph (b)(1) and (3) of this section.
(f) Side loads in the level landing attitude. In the attitudes specified
in paragraphs (b) and (c) of this section, the following apply:
(1) The side loads must be combined at each wheel with one-half of the
maximum vertical ground reactions obtained for that wheel under paragraphs
(b) and (c) of this section. In this condition, the side loads must be--
(i) For the forward wheels, 0.8 times the the vertical reaction (on one
side) acting inward, and 0.6 times the vertical reaction (on the other side)
acting outward; and
(ii) For the rear wheel, 0.8 times the vertical reaction.
(2) The loads specified in paragraph (f)(1) of this section must be
applied--
(i) At the ground contact point with the wheel in the trailing position
(for non-full swiveling landing gear or for full swiveling landing gear with
a lock, steering device, or shimmy damper to keep the wheel in the trailing
position); or
(ii) At the center of the axle (for full swiveling landing gear without a
lock, steering device, or shimmy damper).
(g) Braked roll conditions in the level landing attitude. In the attitudes
specified in paragraphs (b) and (c) of this section, and with the shock
absorbers in their static positions, the rotorcraft must be designed for
braked roll loads as follows:
(1) The limit vertical load must be based on a limit vertical load factor
of not less than--
(i) 1.0, for the attitude specified in paragraph (b) of this section; and
(ii) 1.33, for the attitude specified in paragraph (c) of this section.
(2) For each wheel with brakes, a drag load must be applied, at the ground
contact point, of not less than the lesser of--
(i) 0.8 times the vertical load; and
(ii) The maximum based on limiting brake torque.
(h) Rear wheel turning loads in the static ground attitude. In the static
ground attitude, and with the shock absorbers and tires in their static
positions, the rotorcraft must be designed for rear wheel turning loads as
follows:
(1) A vertical ground reaction equal to the static load on the rear wheel
must be combined with an equal side load.
(2) The load specified in paragraph (h)(1) of this section must be applied
to the rear landing gear--
(i) Through the axle, if there is a swivel (the rear wheel being assumed to
be swiveled 90 degrees to the longitudinal axis of the rotorcraft); or
(ii) At the ground contact point if there is a lock, steering device or
shimmy damper (the rear wheel being assumed to be in the trailing position).
(i) Taxiing condition. The rotorcraft and its landing gear must be designed
for the loads that would occur when the rotorcraft is taxied over the
roughest ground that may reasonably be expected in normal operation.
Sec. 29.501 Ground loading conditions: landing gear with skids.
(a) General. Rotorcraft with landing gear with skids must be designed for
the loading conditions specified in this section. In showing compliance with
this section, the following apply:
(1) The design maximum weight, center of gravity, and load factor must be
determined under Secs. 29.471 through 29.475.
(2) Structural yielding of elastic spring members under limit loads is
acceptable.
(3) Design ultimate loads for elastic spring members need not exceed those
obtained in a drop test of the gear with--
(i) A drop height of 1.5 times that specified in Sec. 29.725; and
(ii) An assumed rotor lift of not more than 1.5 times that used in the
limit drop tests prescribed in Sec. 29.725.
(4) Compliance with paragraph (b) through (e) of this section must be shown
with--
(i) The gear in its most critically deflected position for the landing
condition being considered; and
(ii) The ground reactions rationally distributed along the bottom of the
skid tube.
(b) Vertical reactions in the level landing attitude. In the level
attitude, and with the rotorcraft contacting the ground along the bottom of
both skids, the vertical reactions must be applied as prescribed in paragraph
(a) of this section.
(c) Drag reactions in the level landing attitude. In the level attitude,
and with the rotorcraft contacting the ground along the bottom of both skids,
the following apply:
(1) The vertical reactions must be combined with horizontal drag reactions
of 50 percent of the vertical reaction applied at the ground.
(2) The resultant ground loads must equal the vertical load specified in
paragraph (b) of this section.
(d) Sideloads in the level landing attitude. In the level attitude, and
with the rotorcraft contacting the ground along the bottom of both skids, the
following apply:
(1) The vertical ground reaction must be--
(i) Equal to the vertical loads obtained in the condition specified in
paragraph (b) of this section; and
(ii) Divided equally among the skids.
(2) The vertical ground reactions must be combined with a horizontal
sideload of 25 percent of their value.
(3) The total sideload must be applied equally between skids and along the
length of the skids.
(4) The unbalanced moments are assumed to be resisted by angular inertia.
(5) The skid gear must be investigated for--
(i) Inward acting sideloads; and
(ii) Outward acting sideloads.
(e) One-skid landing loads in the level attitude. In the level attitude,
and with the rotorcraft contacting the ground along the bottom of one skid
only, the following apply:
(1) The vertical load on the ground contact side must be the same as that
obtained on that side in the condition specified in paragraph (b) of this
section.
(2) The unbalanced moments are assumed to be resisted by angular inertia.
(f) Special conditions. In addition to the conditions specified in
paragraphs (b) and (c) of this section, the rotorcraft must be designed for
the following ground reactions:
(1) A ground reaction load acting up and aft at an angle of 45 degrees to
the longitudinal axis of the rotorcraft. This load must be--
(i) Equal to 1.33 times the maximum weight;
(ii) Distributed symmetrically among the skids;
(iii) Concentrated at the forward end of the straight part of the skid
tube; and
(iv) Applied only to the forward end of the skid tube and its attachment to
the rotorcraft.
(2) With the rotorcraft in the level landing attitude, a vertical ground
reaction load equal to one-half of the vertical load determined under
paragraph (b) of this section. This load must be--
(i) Applied only to the skid tube and its attachment to the rotorcraft; and
(ii) Distributed equally over 33.3 percent of the length between the skid
tube attachments and centrally located midway between the skid tube
attachments.
[Amdt. 29-3, 33 FR 966, Jan. 26, 1968, as amended at Amdt. 29-30, 55 FR 8002,
Mar. 6, 1990]
SUMMARY: This rule adopts new and revised airworthiness standards for
certification of airframe and related equipment on both normal and transport
category rotorcraft. In addition, one amendment changes an operating rule
affecting external load operators. These amendments grew out of a rotorcraft
regulatory review program and the recognition by both government and industry
that updated safety standards are needed. These amendments provide a high
level of safety in design requirements, while removing certain unnecessary
existing burdens and better utilizing the unique characteristics and
capabilities of rotorcraft.
If certification for ski operation is requested, the rotorcraft, with skis,
must be designed to withstand the following loading conditions (where P is
the maximum static weight on each ski with the rotorcraft at design maximum
weight, and n is the limit load factor determined under Sec. 29.473(b)):
(a) Up-load conditions in which--
(1) A vertical load of Pn and a horizontal load of Pn/4 are simultaneously
applied at the pedestal bearings; and
(2) A vertical load of 1.33 P is applied at the pedestal bearings.
(b) A side load condition in which a side load of 0.35 Pn is applied at the
pedestal bearings in a horizontal plane perpendicular to the centerline of
the rotorcraft.
(c) A torque-load condition in which a torque load of 1.33 P (in foot-
pounds) is applied to the ski about the vertical axis through the centerline
of the pedestal bearings.
Sec. 29.511 Ground load: unsymmetrical loads on multiple-wheel units.
(a) In dual-wheel gear units, 60 percent of the total ground reaction for
the gear unit must be applied to one wheel and 40 percent to the other.
(b) To provide for the case of one deflated tire, 60 percent of the
specified load for the gear unit must be applied to either wheel except that
the vertical ground reaction may not be less than the full static value.
(c) In determining the total load on a gear unit, the transverse shift in
the load centroid, due to unsymmetrical load distribution on the wheels, may
be neglected.
[Amdt. 29-3, 33 FR 966, Jan. 26, 1968]
Water Loads
Sec. 29.519 Hull type rotorcraft: Water-based and amphibian.
(a) General. For hull type rotorcraft, the structure must be designed to
withstand the water loading set forth in paragraphs (b), (c), and (d) of this
section considering the most severe wave heights and profiles for which
approval is desired. The loads for the landing conditions of paragraphs (b)
and (c) of this section must be developed and distributed along and among the
hull and auxiliary floats, if used, in a rational and conservative manner,
assuming a rotor lift not exceeding two-thirds of the rotorcraft weight to
act throughout the landing impact.
(b) Vertical landing conditions. The rotorcraft must initially contact the
most critical wave surface at zero forward speed in likely pitch and roll
attitudes which result in critical design loadings. The vertical descent
velocity may not be less than 6.5 feet per second relative to the mean water
surface.
(c) Forward speed landing conditions. The rotorcraft must contact the most
critical wave at forward velocities from zero up to 30 knots in likely pitch,
roll, and yaw attitudes and with a vertical descent velocity of not less than
6.5 feet per second relative to the mean water surface. A maximum forward
velocity of less than 30 knots may be used in design if it can be
demonstrated that the forward velocity selected would not be exceeded in a
normal one-engine-out landing.
(d) Auxiliary float immersion condition. In addition to the loads from the
landing conditions, the auxiliary float, and its support and attaching
structure in the hull, must be designed for the load developed by a fully
immersed float unless it can be shown that full immersion of the float is
unlikely, in which case the highest likely float buoyancy load must be
applied that considers loading of the float immersed to create restoring
moments compensating for upsetting moments caused by side wind, asymmetrical
rotorcraft loading, water wave action, and rotorcraft inertia.
[Amdt. 29-3, 33 FR 966, Jan. 26, 1968, as amended at Amdt. 29-30, 55 FR 8002,
Mar. 6, 1990]
SUMMARY: This rule adopts new and revised airworthiness standards for
certification of airframe and related equipment on both normal and transport
category rotorcraft. In addition, one amendment changes an operating rule
affecting external load operators. These amendments grew out of a rotorcraft
regulatory review program and the recognition by both government and industry
that updated safety standards are needed. These amendments provide a high
level of safety in design requirements, while removing certain unnecessary
existing burdens and better utilizing the unique characteristics and
capabilities of rotorcraft.
If certification for float operation (including float amphibian operation)
is requested, the rotorcraft, with floats, must be designed to withstand the
following loading conditions (where the limit load factor is determined under
Sec. 29.473(b) or assumed to be equal to that determined for wheel landing
gear):
(a) Up-load conditions in which--
(1) A load is applied so that, with the rotorcraft in the static level
attitude, the resultant water reaction passes vertically through the center
of gravity; and
(2) The vertical load prescribed in paragraph (a)(1) of this section is
applied simultaneously with an aft component of 0.25 times the vertical
component
(b) A side load condition in which--
(1) A vertical load of 0.75 times the total vertical load specified in
paragraph (a)(1) of this section is divided equally among the floats; and
(2) For each float, the load share determined under paragraph (b)(1) of
this section, combined with a total side load of 0.25 times the total
vertical load specified in paragraph (b)(1) of this section, is applied to
that float only.
[Amdt. 29-3, 33 FR 967, Jan. 26, 1968]
Main Component Requirements
Sec. 29.547 Main rotor structure.
(a) Each main rotor assembly (including rotor hubs and blades) must be
designed as prescribed in this section.
(b) [Reserved]
(c) The main rotor structure must be designed to withstand the following
loads prescribed in Secs. 29.337 through 29.341, and 29.351:
(1) Critical flight loads.
(2) Limit loads occurring under normal conditions of autorotation.
(d) The main rotor structure must be designed to withstand loads
simulating--
(1) For the rotor bl@es, hubs, and flapping hinges, the impact force of
each blade against its stop during ground operation; and
(2) Any other critical condition expected in normal operation.
(e) The main rotor structure must be designed to withstand the limit torque
at any rotational speed, including zero. In addition:
(1) The limit torque need not be greater than the torque defined by a
torque limiting device (where provided), and may not be less than the greater
of--
(i) The maximum torque likely to be transmitted to the rotor structure, in
either direction, by the rotor drive or by sudden application of the rotor
brake; and
(ii) The limit engine torque specified in Sec. 29.361.
(2) The limit torque must be equally and rationally distributed to the
rotor blades.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-4, 33 FR
14106, Sept. 18, 1968]
Sec. 29.549 Fuselage and rotor pylon structures.
(a) Each fuselage and rotor pylon structure must be designed to withstand--
(1) The critical loads prescribed in Secs. 29.337 through 29.341, and
29.351;
(2) The applicable ground loads prescribed in Secs. 29.235, 29.471 through
29.485, 29.493, 29.497, 29.505, and 29.521; and
(3) The loads prescribed in Sec. 29.547 (d)(1) and (e)(1)(i).
(b) Auxiliary rotor thrust, the torque reaction of each rotor drive system,
and the balancing air and inertia loads occurring under accelerated flight
conditions, must be considered.
(c) Each engine mount and adjacent fuselage structure must be designed to
withstand the loads occurring under accelerated flight and landing
conditions, including engine torque.
(d) [Reserved]
(e) If approval for the use of 2 1/2 -minute OEI power is requested, each
engine mount and adjacent structure must be designed to withstand the loads
resulting from a limit torque equal to 1.25 times the mean torque for 2 1/2 -
minute OEI power combined with 1g flight loads.
Each auxiliary lifting surface must be designed to withstand--
(a) The critical flight loads in Secs. 29.337 through 29.341, and 29.351;
(b) the applicable ground loads in Secs. 29.235, 29.471 through 29.485,
29.493, 29.505, and 29.521; and
(c) Any other critical condition expected in normal operation.
Emergency Landing Conditions
Sec. 29.561 General.
(a) The rotorcraft, although it may be damaged in emergency landing
conditions on land or water, must be designed as prescribed in this section
to protect the occupants under those conditions.
(b) The structure must be designed to give each occupant every reasonable
chance of escaping serious injury in a crash landing when--
* * * * *
(b) The structure must be designed to give each occupant every reasonable
chance of escaping serious injury in a minor crash landing when--
(1) Proper use is made of seats, belts, and other safety design provisions;
(2) The wheels are retracted (where applicable); and
(3) Each occupant and each item of mass inside the cabin that could injure
an occupant is restrained when subjected to the following ultimate inertial
load factors relative to the surrounding structure:
(i) Upward--4g.
(ii) Forward--16g.
(iii) Sideward--8g.
(iv) Downward--20g, after the intended displacement of the seat device.
(c) The supporting structure must be designed to restrain under any
ultimate inertial load factor up to those specified in this paragraph, any
item of mass above and/or behind the crew and passenger compartment that
could injure an occupant if it came loose in an emergency landing. Items of
mass to be considered include, but are not limited to, rotors, transmis@on,
and engines. The items of mass must be restrained for the following ultimate
inertial load factors:
(1) Upward--1.5g.
(2) Forward--8g.
(3) Sideward--2g.
(4) Downward--4g.
(d) Any fuselage structure in the area of internal fuel tanks below the
passenger floor level must be designed to resist the following ultimate
inertial factors and loads, and to protect the fuel tanks from rupture, if
rupture is likely when those loads are applied to that area:
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 27-25, 54 FR
47319, Nov. 13, 1989]
Sec. 29.562 Emergency landing dynamic conditions.
(a) The rotorcraft, although it may be damaged in a crash landing, must be
designed to reasonably protect each occupant when--
(1) The occupant properly uses the seats, safety belts, and shoulder
harnesses provided in the design; and
(2) The occupant is exposed to loads equivalent to those resulting from the
conditions prescribed in this section.
(b) Each seat type design or other seating device approved for crew or
passenger occupancy during takeoff and landing must successfully complete
dynamic tests or be demonstrated by rational analysis based on dynamic tests
of a similar type seat in accordance with the following criteria. The tests
must be conducted with an occupant simulated by a 170-pound anthropomorphic
test dummy (ATD), as defined by 49 CFR 572, Subpart B, or its equivalent,
sitting in the normal upright position.
(i) A change in downward velocity of not less than 30 feet per second when
the seat or other seating device is oriented in its nominal position with
respect to the rotorcraft's reference system, the rotorcraft's longitudinal
axis is canted upward 60 deg. with respect to the impact velocity vector, and
the rotorcraft's lateral axis is perpendicular to a vertical plane containing
the impact velocity vector and the rotorcraft's longitudinal axis. Peak floor
deceleration must occur in not more than 0.031 seconds after impact and must
reach a minimum of 30g's.
(2) A change in forward velocity of not less than 42 feet per second when
the seat or other seating device is oriented in its nominal position with
respect to the rotorcraft's reference system, the rotorcraft's longitudinal
axis is yawed 10 deg. either right or left of the impact velocity vector
(whichever would cause the greatest load on the shoulder harness), the
rotorcraft's lateral axis is contained in a horizontal plane containing the
impact velocity vector, and the rotorcraft's vertical axis is perpendicular
to a horizontal plane containing the impact velocity vector. Peak floor
deceleration must occur in not more than 0.071 seconds after impact and must
reach a minimum of 18.4g's.
(3) Where floor rails or floor or sidewall floor attachment devices are
used to attach the seating devices to the airframe structure for the
conditions of this section, the rails or devices must be misaligned with
respect to each other by at least 10 deg. vertically (i.e., pitch out of
parallel) and by at least a 10 deg. lateral roll, with the directions
optional, to account for possible floor warp.
(c) Compliance with the following must be shown:
(1) The seating device system must remain intact although it may experience
separation intended as part of its design.
(2) The attachment between the seating device and the airframe structure
must remain intact although the structure may have exceeded its limit load.
(3) The ATD's shoulder harness strap or straps must remain on or in the
immediate vicinity of the ATD's shoulder during the impact.
(4) The safety belt must remain on the ATD's pelvis during the impact.
(5) The ATD's head either does not contact any portion of the crew or
passenger compartment or, if contact is made, the head impact does not exceed
a head injury criteria (HIC) of 1,000 as determined by this equation.
**2.5
1 t2
HIC = (t2-t1) [------ I a(t)dt ]
(t2-t1) t1
Where: a(t) is the resultant acceleration at the center of gravity of the
head form expressed as a multiple of g (the acceleration of gravity) and t2 -
t1 is the time duration, in seconds, of major head impact, not to exceed 0.05
seconds.
(6) Loads in individual shoulder harness straps must not exceed 1,750
pounds. If dual straps are used for retaining the upper torso, the total
harness strap loads must not exceed 2,000 pounds.
(7) The maximum compressive load measured between the pelvis and the lumbar
column of the ATD must not exceed 1,500 pounds.
(d) An alternate approach that achieves an equivalent or greater level of
occupant protection, as required by this section, must be substantiated on a
rational basis.
[Amdt. 27-25, 54 FR 47320, Nov. 13, 1989]
Sec. 29.563 Structural ditching provisions.
If certification with ditching provisions is requested, structural strength
for ditching must meet the requirements of this section and Sec. 29.801(e).
(a) Forward speed landing conditions. The rotorcraft must initially contact
the most critical wave for reasonably probable water conditions at forward
velocities from zero up to 30 knots in likely pitch, roll, and yaw attitudes.
The rotorcraft limit vertical descent velocity may not be less than 5 feet
per second relative to the mean water surface. Rotor lift may be used to act
through the center of gravity throughout the landing impact. This lift may
not exceed two-thirds of the design maximum weight. A maximum forward
velocity of less than 30 knots may be used in design if it can be
demonstrated that the forward velocity selected would not be exceeded in a
normal one-engine-out touchdown.
(b) Auxiliary or emergency float conditions.--(1) Floats fixed or deployed
before initial water contact. In addition to the landing loads in paragraph
(a) of this section, each auxiliary or emergency float, or its support and
attaching structure in the airframe or fuselage, must be designed for the
load developed by a fully immersed float unless it can be shown that full
immersion is unlikely. If full immersion is unlikely, the highest likely
float buoyancy load must be applied. The highest likely buoyancy load must
include consideration of a partially immersed float creating restoring
moments to compensate the upsetting moments caused by side wind,
unsymmetrical rotorcraft loading, water wave action, rotorcraft inertia, and
probable structural damage and leakage considered under Sec. 29.801(d).
Maximum roll and pitch angles determined from compliance with Sec. 29.801(d)
may be used, if significant, to determine the extent of immersion of each
float. If the floats are deployed in flight, appropriate air loads derived
from the flight limitations with the floats deployed shall be used in
substantiation of the floats and their attachment to the rotorcraft. For this
purpose, the design airspeed for limit load is the float deployed airspeed
operating limit multiplied by 1.11.
(2) Floats deployed after initial water contact. Each float must be
designed for full or partial immersion prescribed in paragraph (b)(1) of this
section. In addition, each float must be designed for combined vertical and
drag loads using a relative limit speed of 20 knots between the rotorcraft
and the water. The vertical load may not be less than the highest likely
buoyancy load determined under paragraph (b)(1) of this section.
SUMMARY: This rule adopts new and revised airworthiness standards for
certification of airframe and related equipment on both normal and transport
category rotorcraft. In addition, one amendment changes an operating rule
affecting external load operators. These amendments grew out of a rotorcraft
regulatory review program and the recognition by both government and industry
that updated safety standards are needed. These amendments provide a high
level of safety in design requirements, while removing certain unnecessary
existing burdens and better utilizing the unique characteristics and
capabilities of rotorcraft.
(a) General. An evaluation of the strength of principal elements, detail
design points, and fabrication techniques must show that catastrophic failure
due to fatigue, considering the effects of environment, intrinsic/discrete
flaws, or accidental damage will be avoided. Parts to be evaluated include,
but are not limited to, rotors, rotor drive systems between the engines and
rotor hubs, controls, fuselage, fixed and movable control surfaces, engine
and transmission mountings, landing gear, and their related primary
attachments. In addition, the following apply:
(1) Each evaluation required by this section must include--
(i) The identification of principal structural elements, the failure of
which could result in catastrophic failure of the rotorcraft;
(ii) In-flight measurement in determining the loads or stresses for items
in paragraph (a)(1)(i) of this section in all critical conditions throughout
the range of limitations in Sec. 29.309 (including altitude effects), except
that maneuvering load factors need not exceed the maximum values expected in
operations; and
(iii) Loading spectra as severe as those expected in operation based on
loads or stresses determined under paragraph (a)(1)(ii) of this section,
including external load operations, if applicable, and other high frequency
power cycle operations.
(2) Based on the evaluations required by this section, inspections,
replacement times, combinations thereof, or other procedures must be
established as necessary to avoid catastrophic failure. These inspections,
replacement times, combinations thereof, or other procedures must be included
in the airworthiness limitations section of the Instructions for Continued
Airworthiness required by Sec. 29.1529 and section A29.4 of Appendix A of
this part.
(b) Fatigue tolerance evaluation (including tolerance to flaws). The
structure must be shown by analysis supported by test evidence and, if
available, service experience to be of fatigue tolerant design. The fatigue
tolerance evaluation must include the requirements of either paragraph (b)
(1), (2), or (3) of this section, or a combination thereof, and also must
include a determination of the probable locations and modes of damage caused
by fatigue, considering environmental effects, intrinsic/discrete flaws, or
accidental damage. Compliance with the flaw tolerance requirements of
paragraph (b) (1) or (2) of this section is required unless the applicant
establishes that these fatigue flaw tolerant methods for a particular
structure cannot be achieved within the limitations of geometry,
inspectability, or good design practice. Under these circumstances, the safe-
life evaluation of paragraph (b)(3) of this section is required.
(1) Flaw tolerant safe-life evaluation. It must be shown that the
structure, with flaws present, is able to withstand repeated loads of
variable magnitude without detectable flaw growth for the following time
intervals--
(i) Life of the rotorcraft; or
(ii) Within a replacement time furnished under section A29.4 of appendix A
to this part.
(2) Fail-safe (residual strength after flaw growth) evaluation. It must be
shown that the structure remaining after a partial failure is able to
withstand design limit loads without failure within an inspection period
furnished under section A29.4 of appendix A to this part. Limit loads are
defined in Sec. 29.301(a).
(i) The residual strength evaluation must show that the remaining structure
after flaw growth is able to withstand design limit loads without failure
within its operational life.
(ii) Inspection intervals and methods must be established as necessary to
ensure that failures are detected prior to residual strength conditions being
reached.
(iii) If significant changes in structural stiffness or geometry, or both,
follow from a structural failure or partial failure, the effect on flaw
tolerance must be further investigated.
(3) Safe-life evaluation. It must be shown that the structure is able to
withstand repeated loads of variable magnitude without detectable cracks for
the following time intervals--
(i) Life of the rotorcraft; or
(ii) Within a replacement time furnished under section A29.4 of appendix A
to this part.
Subpart D--Design and Construction
General
Sec. 29.601 Design.
(a) The rotorcraft may have no design features or details that experience
has shown to be hazardous or unreliable.
(b) The suitability of each questionable design detail and part must be
established by tests.
Sec. 29.603 Materials.
The suitability and durability of materials used for parts, the failure of
which could adversely affect safety, must--
(a) Be established on the basis of experience or tests;
(b) Meet approved specifications that ensure their having the strength and
other properties assumed in the design data; and
(c) Take into account the effects of environmental conditions, such as
temperature and humidity, expected in service.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49
U.S.C. 1354(a), 1421, 1423, 1424), and sec. 6(c), Dept. of Transportation Act
(49 U.S.C. 1655(c)))
(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 according to an approved process specification.
(b) Each new aircraft fabrication method must be substantiated by a test
program.
(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C.
1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 U.S.C.
1655(c)))
[Doc. No. 5084, 29 FR 16150. Dec. 3, 1964, as amended by Amdt. 29-17, 43 FR
50599, Oct. 30, 1978]
Sec. 29.607 Fasteners.
(a) Each removable bolt, screw, nut, pin, or other fastener whose loss
could jeopardize the safe operation of the rotorcraft must incorporate two
separate locking devices. The fastener and its locking devices may not be
adversely affected by the environmental conditions associated with the
particular installation.
(b) 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. 29-5, 33 FR 14533, Sept. 27, 1968]
Sec. 29.609 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 provisions for ventilation and drainage where necessary to prevent
the accumulation of corrosive, flammable, or noxious fluids.
Sec. 29.610 Lightning protection.
(a) The rotorcraft must be protected against catastrophic effects from
lightning.
(b) For metallic components, compliance with paragraph (a) of this section
may be shown by--
(1) Electrically bonding the components properly to the airframe; or
(2) Designing the components so that a strike will not endanger the
rotorcraft.
(c) For nonmetallic components, compliance with paragraph (a) of this
section may be shown by--
(1) Designing the components to minmize the effect of a strike; or
(2) Incorporating acceptable means of diverting the resulting electrical
current to not endanger the rotorcraft.
[Amdt. 29-24, 49 FR 44437, Nov. 6, 1984]
Sec. 29.611 Inspection provisions.
There must be means to allow close examination of each part that requires--
(a) Recurring inspection;
(b) Adjustment for proper alignment and functioning; or
(c) Lubrication.
Sec. 29.613 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 paragraphs (d) and
(e) of this section, compliance with this paragraph must be shown by
selecting design values that assure 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; and
(2) For redundant structures, those 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 strength, detail design, and fabrication of the structure must
minimize the probability of disastrous fatigue failure, particularly at
points of stress concentration.
(d) Design values may be those contained in the following publications
(available from the Naval Publications and Forms Center, 5801 Tabor Avenue,
Philadelphia, PA 19120) or other values approved by the Administrator:
(1) MIL--HDBK-5, "Metallic Materials and Elements for Flight Vehicle
Structure".
(2) MIL--HDBK-17, "Plastics for Flight Vehicles".
(3) ANC-18, "Design of Wood Aircraft Structures".
(4) MIL--HDBK-23, "Composite Construction for Flight Vehicles".
(e) Other design values may be used if a selection of the material is made
in which a specimen of each individual item is tested before use and it is
determined that the actual strength properties of that particular item will
equal or exceed those used in design.
(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C.
1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 U.S.C.
1655(c)))
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-17, 43 FR
50599, Oct. 30, 1978; Amdt. 29-30, 55 FR 8003, Mar. 6, 1990]
SUMMARY: This rule adopts new and revised airworthiness standards for
certification of airframe and related equipment on both normal and transport
category rotorcraft. In addition, one amendment changes an operating rule
affecting external load operators. These amendments grew out of a rotorcraft
regulatory review program and the recognition by both government and industry
that updated safety standards are needed. These amendments provide a high
level of safety in design requirements, while removing certain unnecessary
existing burdens and better utilizing the unique characteristics and
capabilities of rotorcraft.
(a) The special factors prescribed in Secs. 29.621 through 29.625 apply to
each part of the structure whose strength is--
(1) Uncertain;
(2) Likely to deteriorate in service before normal replacement; or
(3) Subject to appreciable variability due to--
(i) Uncertainties in manufacturing processes; or
(ii) Uncertainties in inspection methods.
(b) For each part of the rotorcraft to which Secs. 29.621 through 29.625
apply, the factor of safety prescribed in Sec. 29.303 must be multiplied by a
special factor equal to--
(1) The applicable special factors prescribed in Secs. 29.621 through
29.625; or
(2) Any other factor great enough to ensure that the probability of the
part being understrength because of the uncertainties specified in paragraph
(a) of this section is extremely remote.
Sec. 29.621 Casting factors.
(a) General. The factors, tests, and inspections specified in paragraphs
(b) and (c) 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 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 rotorcraft or result in serious
injury to any occupant, the following apply:
(1) Each critical casting must--
(i) Have a casting factor of not less than 1.25; and
(ii) Receive 100 percent inspection by visual, radiographic, and magnetic
particle (for ferromagnetic materials) or penetrate (for nonferromagnetic
materials) inspection methods or approved equivalent inspection methods.
(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. 29.305 at an ultimate load
corresponding to a casting factor of 1.25; and
(ii) The deformation requirements of Sec. 29.305 at a load of 1.15 times
the limit load.
(d) Noncritical castings. For each casting other than those specified in
paragraph (c) 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 greater 100 percent visual.
Less than 2.0, greater than 1.5 100 percent visual, and magnetic particle
(ferromagnetic materials), penetrant
(nonferromagnetic materials), or approved
equivalent inspection methods.
1.25 through 1.50 100 percent visual, and magnetic particle
(ferromagnetic materials), penetrant
(nonferromagnetic materials), and
radiographic or approved equivalent
inspection methods.
(2) The percentage of castings inspected by nonvisual methods may be
reduced below that specified in paragraph (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.
Sec. 29.623 Bearing factors.
(a) Except as provided in paragraph (b) of this section, 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) No bearing factor need be used on a part for which any larger special
factor is prescribed.
Sec. 29.625 Fitting factors.
For each fitting (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--
(1) For joints made under approved practices and based on comprehensive
test data (such as continuous joints in metal plating, welded joints, and
scarf joints in wood); and
(2) With respect to any bearing surface for which a larger special factor
is used.
(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.
Sec. 29.629 Flutter.
Each aerodynamic surface of the rotorcraft must be free from flutter
under each appropriate speed and power condition.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended at Amdt. 29-30, 55
FR 8003, Mar. 6, 1990]
SUMMARY: This rule adopts new and revised airworthiness standards for
certification of airframe and related equipment on both normal and transport
category rotorcraft. In addition, one amendment changes an operating rule
affecting external load operators. These amendments grew out of a rotorcraft
regulatory review program and the recognition by both government and industry
that updated safety standards are needed. These amendments provide a high
level of safety in design requirements, while removing certain unnecessary
existing burdens and better utilizing the unique characteristics and
capabilities of rotorcraft.
Sec. 29.653 Pressure venting and drainage of rotor blades.
(a) For each rotor blade--
(1) There must be means for venting the internal pressure of the blade;
(2) Drainage holes must be provided for the blade; and
(3) The blade must be designed to prevent water from becoming trapped in
it.
(b) Paragraphs (a)(1) and (2) of this section does not apply to sealed
rotor blades capable of withstanding the maximum pressure differentials
expected in service.
[Amdt. 29-3, 33 FR 967, Jan. 26, 1968]
Sec. 29.659 Mass balance.
(a) The rotor and blades must be mass balanced as necessary to--
(1) Prevent excessive vibration; and
(2) Prevent flutter at any speed up to the maximum forward speed.
(b) The structural integrity of the mass balance installation must be
substantiated.
[Amdt. 29-3, 33 FR 967, Jan. 26, 1968]
Sec. 29.661 Rotor blade clearance.
There must be enough clearance between the rotor blades and other parts of
the structure to prevent the blades from striking any part of the structure
during any operating condition.
[Amdt. 29-3, 33 FR 967, Jan. 26, 1968]
Sec. 29.663 Ground resonance prevention means.
(a) The reliability of the means for preventing ground resonance must be
shown either by analysis and tests, or reliable service experience, or by
showing through analysis or tests that malfunction or failure of a single
means will not cause ground resonance.
(b) The probable range of variations, during service, of the damping action
of the ground resonance prevention means must be established and must be
investigated during the test required by Sec. 29.241.
SUMMARY: This rule adopts new and revised airworthiness standards for
certification of airframe and related equipment on both normal and transport
category rotorcraft. In addition, one amendment changes an operating rule
affecting external load operators. These amendments grew out of a rotorcraft
regulatory review program and the recognition by both government and industry
that updated safety standards are needed. These amendments provide a high
level of safety in design requirements, while removing certain unnecessary
existing burdens and better utilizing the unique characteristics and
capabilities of rotorcraft.
(a) Each control and control system must operate with the ease, smoothness,
and positiveness appropriate to its function.
(b) Each element of each flight control system must be designed, or
distinctively and permanently marked, to minimize the probability of any
incorrect assembly that could result in the malfunction of the system.
(c) A means must be provided to allow full control movement of all primary
flight controls prior to flight, or a means must be provided that will allow
the pilot to determine that full control authority is available prior to
flight.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-24, 49 FR
44437, Nov. 6, 1984]
Sec. 29.672 Stability augmentation, automatic, and power-operated systems.
If the functioning of stability augmentation or other automatic or power-
operated system is necessary to show compliance with the flight
characteristics requirements of this part, the system must comply with Sec.
29.671 of this part 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 which could result in an unsafe condition
if the pilot is unaware of the failure. Warning systems must not activate the
control systems.
(b) The design of the stability augmentation system or of any other
automatic or power-operated system must allow initial counteraction of
failures without requiring exceptional pilot skill or strength, by overriding
the failure by moving the flight controls in the normal sense, and by
deactivating the failed system.
(c) It must be show that after any single failure of the stability
augmentation system or any other automatic or power-operated system--
(1) The rotorcraft is safely controllable when the failure or malfunction
occurs at any speed or altitude within the approved operating limitations;
(2) The controllability and maneuverability requirements of this part are
met within a practical operational flight envelope (for example, speed,
altitude, normal acceleration, and rotorcraft configurations) which is
described in the Rotorcraft Flight Manual; and
(3) The trim and stability characteristics are not impaired below a level
needed to allow continued safe flight and landing.
[Amdt. 29-24, 49 FR 44437, Nov. 6, 1984]
Sec. 29.673 Primary flight controls.
Primary flight controls are those used by the pilot for immediate control
of pitch, roll, yaw, and vertical motion of the rotorcraft.
[Amdt. 29-24, 49 FR 44437, Nov. 6, 1984]
Sec. 29.674 Interconnected controls.
Each primary flight control system must provide for safe flight and landing
and operate independently after a malfunction, failure, or jam of any
auxiliary interconnected control.
SUMMARY: This rule adopts new and revised airworthiness standards for
certification of airframe and related equipment on both normal and transport
category rotorcraft. In addition, one amendment changes an operating rule
affecting external load operators. These amendments grew out of a rotorcraft
regulatory review program and the recognition by both government and industry
that updated safety standards are needed. These amendments provide a high
level of safety in design requirements, while removing certain unnecessary
existing burdens and better utilizing the unique characteristics and
capabilities of rotorcraft.
(a) Each control system must have stops that positively limit the range of
motionof the pilot's controls.
(b) Each stop must be located in the system so that the range of travel of
its control is not appreciably affected by--
(1) Wear;
(2) Slackness; or
(3) Takeup adjustments.
(c) Each stop must be able to withstand the loads corresponding to the
design conditions for the system.
(d) For each main rotor blade--
(1) Stops that are appropriate to the blade design must be provided to
limit travel of the blade about its hinge points; and
(2) There must be means to keep the blade from hitting the droop stops
during any operation other than starting and stopping the rotor.
(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C.
1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 U.S.C.
1655(c)))
[Doc. No. 5084, 29 FR 16150. Dec. 3, 1964, as amended by Amdt. 29-17, 43 FR
50599, Oct. 30, 1978]
Sec. 29.679 Control system locks.
If there is a device to lock the control system with the rotorcraft on the
ground or water, there must be means to--
(a) Automatically disengage the lock when the pilot operates the controls
in a normal manner, or limit the operation of the rotorcraft so as to give
unmistakable warning to the pilot before takeoff; and
(b) Prevent the lock from engaging in flight.
Sec. 29.681 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. 29.683 Operation tests.
It must be shown by operation tests that, when the controls are operated
from the pilot compartment with the control system loaded to correspond with
loads specified for the system, the system is free from--
(a) Jamming;
(b) Excessive friction; and
(c) Excessive deflection.
Sec. 29.685 Control system details.
(a) Each detail of each control system must be designed 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) Cable systems must be designed as follows:
(1) Cables, cable fittings, turnbuckles, splices, and pulleys must be of an
acceptable kind.
(2) The design of cable systems must prevent any hazardous change in cable
tension throughout the range of travel under any operating conditions and
temperature variations.
(3) No cable smaller than 1/8 inch diameter may be used in any primary
control system.
(4) Pulley kinds and sizes must correspond to the cables with which they
are used. The pulley-cable combinations and strength values specified in MIL-
HDBK-5 must be used unless they are inapplicable.
(5) Pulleys must have close fitting guards to prevent the cables from being
displaced or fouled.
(6) Pulleys must lie close enough to the plane passing through the cable to
prevent the cable from rubbing against the pulley flange.
(7) No fairlead may cause a change in cable direction of more than three
degrees.
(8) No clevis pin subject to load or motion and retained only by cotter
pins may be used in the control system.
(9) Turnbuckles attached to parts having angular motion must be installed
to prevent binding throughout the range of travel.
(10) There must be means for visual inspection at each fairlead, pulley,
terminal, and turnbuckle.
(e) Control system joints subject to angular motion must incorporate the
following special factors with respect to the ultimate bearing strength of
the softest material used as a bearing:
(1) 3.33 for push-pull systems other than ball and roller bearing systems.
(2) 2.0 for cable systems.
(f) For control system joints, the manufacturer's static, non-Brinell
rating of ball and roller bearings may not be exceeded.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 FR
55471, Dec. 20, 1976]
Sec. 29.687 Spring devices.
(a) Each control system spring device whose failure could cause flutter or
other unsafe characteristics must be reliable.
(b) Compliance with paragraph (a) of this section must be shown by tests
simulating service conditions.
Sec. 29.691 Autorotation control mechanism.
Each main rotor blade pitch control mechanism must allow rapid entry into
autorotation after power failure.
Sec. 29.695 Power boost and power-operated control system.
(a) If a power boost or power-operated control system is used, an alternate
system must be immediately available that allows continued safe flight and
landing in the event of--
(1) Any single failure in the power portion of the system; or
(2) The failure of all engines.
(b) Each alternate system may be a duplicate power portion or a manually
operated mechanical system. The power portion includes the power source (such
as hydrualic pumps), and such items as valves, lines, and actuators.
(c) The failure of mechanical parts (such as piston rods and links), and
the jamming of power cylinders, must be considered unless they are extremely
improbable.
Landing Gear
Sec. 29.723 Shock absorption tests.
The landing inertia load factor and the reserve energy absorption capacity
of the landing gear must be substantiated by the tests prescribed in Secs.
29.725 and 29.727, respectively. These tests must be conducted on the
complete rotorcraft or on units consisting of wheel, tire, and shock absorber
in their proper relation.
Sec. 29.725 Limit drop test.
The limit drop test must be conducted as follows:
(a) The drop height must be at least 8 inches.
(b) If considered, the rotor lift specified in Sec. 29.473(a) must be
introduced into the drop test by appropriate energy absorbing devices or by
the use of an effective mass.
(c) Each landing gear unit must be tested in the attitude simulating the
landing condition that is most critical from the standpoint of the energy to
be absorbed by it.
(d) When an effective mass is used in showing compliance with paragraph (b)
of this section, the following formulae may be used instead of more rational
computations.
h+(1-L)d
We = W x ------------ ; and
h+d
We
n = nj ---- + L
W
where:
We=the effective weight to be used in the drop test (lbs.).
W=WM for main gear units (lbs.), equal to the static reaction on the
particular unit with the rotorcraft in the most critical attitude. A
rational method may be used in computing a main gear static reaction,
taking into consideration the moment arm between the main wheel reaction
and the rotorcraft center of gravity.
W=WN for nose gear units (lbs.), equal to the vertical component of the
static reaction that would exist at the nose wheel, assuming that the
mass of the rotorcraft acts at the center of gravity and exerts a force
of 1.0g downward and 0.25g forward.
W=Wt for tailwheel units (lbs.) equal to whichever of the following is
critical--
(1) The static weight on the tailwheel with the rotorcraft resting on all
wheels; or
(2) The vertical component of the ground reaction that would occur at the
tailwheel assuming that the mass of the rotorcraft acts at the center of
gravity and exerts a force of 1g downward with the rotorcraft in the maximum
nose-up attitude considered in the nose-up landing conditions.
h =specified free drop height (inches).
L=ratio of assumed rotor lift to the rotorcraft weight.
d=deflection under impact of the tire (at the proper inflation pressure) plus
the vertical component of the axle travel (inches) relative to the drop
mass.
n=limit inertia load factor.
nj=the load factor developed, during impact, on the mass used in the drop
test (i.e., the acceleration dv/dt in g's recorded in the drop test plus
1.0).
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 FR
967, Jan. 26. 1968]
Sec. 29.727 Reserve energy absorption drop test.
The reserve energy absorption drop test must be conducted as follows:
(a) The drop height must be 1.5 times that specified in Sec. 29.725(a).
(b) Rotor lift, where considered in a manner similar to that prescribed in
Sec. 29.725(b), may not exceed 1.5 times the lift allowed under that
paragraph.
(c) The landing gear must withstand this test without collapsing. Collapse
of the landing gear occurs when a member of the nose, tail, or main gear will
not support the rotorcraft in the proper attitude or allows the rotorcraft
structure, other than landing gear and external accessories, to impact the
landing surface.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended at Amdt. 29-30, 55 FR
8003, Mar. 6, 1990]
SUMMARY: This rule adopts new and revised airworthiness standards for
certification of airframe and related equipment on both normal and transport
category rotorcraft. In addition, one amendment changes an operating rule
affecting external load operators. These amendments grew out of a rotorcraft
regulatory review program and the recognition by both government and industry
that updated safety standards are needed. These amendments provide a high
level of safety in design requirements, while removing certain unnecessary
existing burdens and better utilizing the unique characteristics and
capabilities of rotorcraft.
For rotorcraft with retractable landing gear, the following apply:
(a) Loads. The landing gear, retracting mechanism, wheel well doors, and
supporting structure must be designed for--
(1) The loads occurring in any maneuvering condition with the gear
retracted;
(2) The combined friction, inertia, and air loads occurring during
retraction and extension at any airspeed up to the design maximum landing
gear operating speed; and
(3) The flight loads, including those in yawed flight, occurring with the
gear extended at any airspeed up to the design maximum landing gear extended
speed.
(b) Landing gear lock. A positive means must be provided to keep the gear
extended.
(c) Emergency operation. When other than manual power is used to operate
the gear, emergency means must be provided for extending the gear in the
event of--
(1) Any reasonably probable failure in the normal retraction system; or
(2) The failure of any single source of hydraulic, electric, or equivalent
energy.
(d) Operation tests. The proper functioning of the retracting mechanism
must be shown by operation tests.
(e) Position indicator. There must be means to indicate to the pilot when
the gear is secured in the extreme positions.
(f) Control. The location and operation of the retraction control must meet
the requirements of Secs. 29.777 and 29.779.
(g) Landing gear warning. An aural or equally effective landing gear
warning device must be provided that functions continuously when the
rotorcraft is in a normal landing mode and the landing gear is not fully
extended and locked. A manual shutoff capability must be provided for the
warning device and the warning system must automatically reset when the
rotorcraft is no longer in the landing mode.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-24, 49 FR
44437, Nov. 6, 1984]
Sec. 29.731 Wheels.
(a) Each landing gear wheel must be approved.
(b) The maximum static load rating of each wheel may not be less than the
corresponding static ground reaction with--
(1) Maximum weight; and
(2) Critical center of gravity.
(c) 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.
Sec. 29.733 Tires.
Each landing gear wheel must have a tire--
(a) That is a proper fit on the rim of the wheel; and
(b) Of a rating that is not exceeded under--
(1) The design maximum weight;
(2) A load on each main wheel tire equal to the static ground reaction
corresponding to the critical center of gravity; and
(3) A load on nose wheel tires (to be compared with the dynamic rating
established for those tires) equal to the reaction obtained at the nose
wheel, assuming that the mass of the rotorcraft acts as the most critical
center of gravity and exerts a force of 1.0 g downward and 0.25 g forward,
the reactions being distributed to the nose and main wheels according to the
principles of statics with the drag reaction at the ground applied only at
wheels with brakes.
(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 or systems.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 FR
55471, Dec. 20, 1976]
Sec. 29.735 Brakes.
For rotorcraft with wheel-type landing gear, a braking device must be
installed that is--
(a) Controllable by the pilot;
(b) Usable during power-off landings; and
(c) Adequate to--
(1) Counteract any normal unbalanced torque when starting or stopping the
rotor; and
(2) Hold the rotorcraft parked on a 10-degree slope on a dry, smooth
pavement.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-24, 49 FR
44437, Nov. 6, 1984]
Sec. 29.737 Skis.
(a) 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.
(b) There must be a stabilizing means to maintain the ski in an appropriate
position during flight. This means must have enough strength to withstand the
maximum aerodynamic and inertia loads on the ski.
Floats and Hulls
Sec. 29.751 Main float buoyancy.
(a) For main floats, the buoyancy necessary to support the maximum weight
of the rotorcraft in fresh water must be exceeded by--
(1) 50 percent, for single floats; and
(2) 60 percent, for multiple floats.
(b) Each main float must have enough water-tight compartments so that, with
any single main float compartment flooded, the mainfloats will provide a
margin of positive stability great enough to minimize the probability of
capsizing.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 FR
967, Jan. 26, 1968]
Sec. 29.753 Main float design.
(a) Bag floats. Each bag float must be designed to withstand--
(1) The maximum pressure differential that might be developed at the
maximum altitude for which certification with that float is requested; and
(2) The vertical loads prescribed in Sec. 29.521(a), distributed along the
length of the bag over three-quarters of its projected area.
(b) Rigid floats. Each rigid float must be able to withstand the vertical,
horizontal, and side loads prescribed in Sec. 29.521. An appropriate load
distribution under critical conditions must be used.
Sec. 29.755 Hull buoyancy.
Water-based and amphibian rotorcraft. The hull and auxiliary floats,
if used, must have enough watertight compartments so that, with any single
compartment of the hull or auxiliary floats flooded, the buoyancy of the hull
and auxiliary floats, and wheel tires if used, provides a margin of positive
water stability great enough to minimize the probability of capsizing the
rotorcraft for the worst combination of wave heights and surface winds for
which approval is desired.
[Amdt. 29-3, 33 FR 967, Jan. 26, 1968, as amended at Amdt. 29-30, 55 FR 8003,
Mar. 6, 1990]
SUMMARY: This rule adopts new and revised airworthiness standards for
certification of airframe and related equipment on both normal and transport
category rotorcraft. In addition, one amendment changes an operating rule
affecting external load operators. These amendments grew out of a rotorcraft
regulatory review program and the recognition by both government and industry
that updated safety standards are needed. These amendments provide a high
level of safety in design requirements, while removing certain unnecessary
existing burdens and better utilizing the unique characteristics and
capabilities of rotorcraft.
The hull, and auxiliary floats if used, must withstand the water loads
prescribed by Sec. 29.519 with a rational and conservative distribution of
local and distributed water pressures over the hull and float bottom.
[Amdt. 29-3, 33 FR 967, Jan. 26, 1968]
Personnel and Cargo Accommodations
Sec. 29.771 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) If there is provision for a second pilot, the rotorcraft must be
controllable with equal safety from either pilot position. Flight and
powerplant controls must be designed to prevent confusion or inadvertent
operation when the rotorcraft is piloted from either position;
(c) The vibration and noise characteristics of cockpit appurtenances may
not interfere with safe operation;
(d) Inflight leakage of rain or snow that could distract the crew or harm
the structure must be prevented.
(a) Nonprecipitation conditions. For nonprecipitation conditions, the
following apply:
(1) Each pilot compartment must be arranged to give the pilots a
sufficiently extensive, clear, and undistorted view for safe operation.
(2) Each pilot compartment must be free of glare and reflection that could
interfere with the pilot's view. If certification for night operation is
requested, this must be shown by night flight tests.
(b) Precipitation conditions. For precipitation conditions, the following
apply:
(1) Each pilot must have a sufficiently extensive view for safe operation--
(i) In heavy rain at forward speeds up to VH; and
(ii) In the most severe icing condition for which certification is
requested.
(2) The first pilot must have a window that--
(i) Is openable under the conditions prescribed in paragraph (b) (1) of
this section; and
(ii) Provides the view prescribed in that paragraph.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 FR
967, Jan. 26, 1968]
Sec. 29.775 Windshields and windows.
Windshields and windows must be made of material that will not break into
dangerous fragments.
SUMMARY: This rule amends the airworthiness standards for systems,
propulsion, and airframe for both normal and transport category rotorcraft.
In addition, these amendments introduce safety improvements, clarifying
existing regulations, and standardize terminology. The changes are based on
some of the proposals that were submitted to the FAA by the European
Airworthiness Authorities. These amendments are also intended to encourage
the European community's acceptance of the Federal Aviation Regulations for
rotorcraft type certification, obviate development of different European
standards, and achieve increased commonality of airworthiness standards
among the respective countries.
Cockpit controls must be--
(a) Located to provide convenient operation and to prevent confusion and
inadvertent operation; and
(b) Located and arranged with respect to the pilot's seats so that there is
full and unrestricted movement of each control without interference from the
cockpit structure or the pilot's clothing when pilots from 5'2'' to 6'0'' in
height are seated.
Sec. 29.779 Motion and effect of cockpit controls.
Cockpit controls must be designed so that they operate in accordance with
the following movements and actuation:
(a) Flight controls, including the collective pitch control, must operate
with a sense of motion which corresponds to the effect on the rotorcraft.
(b) Twist-grip engine power controls must be designed so that, for lefthand
operation, the motion of the pilot's hand is clockwise to increase power when
the hand is viewed from the edge containing the index finger. Other engine
power controls, excluding the collective control, must operate with a forward
motion to increase power.
(c) Normal landing gear controls must operate downward to extend the
landing gear.
[Amdt. 29-24, 49 FR 44437, Nov. 6, 1984]
Sec. 29.783 Doors.
(a) Each closed cabin must have at least one adequate and easily accessible
external door.
(b) Each external door must be located, and appropriate operating
procedures must be established, to ensure that persons using the door will
not be endangered by the rotors, propellers, engine intakes, and exhausts
when the operating procedures are used.
(c) There must be means for locking crew and external passenger doors and
for preventing their opening in flight inadvertently or as a result of
mechanical failure. It must be possible to open external doors from inside
and outside the cabin with the rotorcraft on the ground even though persons
may be crowded against the door on the inside of the rotorcraft. The means of
opening must be simple and obvious and so arranged and marked that it can be
readily located and operated.
(d) There must be reasonable provisions to prevent the jamming of any
external doors in a minor crash as a result of fuselage deformation under the
following ultimate inertial forces except for cargo or service doors not
suitable for use as an exit in an emergency:
(1) Upward--1.5g.
(2) Forward--4.0g.
(3) Sideward--2.0g.
(4) Downward--4.0g.
(e) There must be means for direct visual inspection of the locking
mechanism by crewmembers to determine whether the external doors (including
passenger, crew, service, and cargo doors) are fully locked. There must be
visual means to signal to appropriate crewmembers when normally used external
doors are closed and fully locked.
(f) For outward opening external doors usable for entrance or egress, there
must be an auxiliary safety latching device to prevent the door from opening
when the primary latching mechanism fails. If the door does not meet the
requirements of paragraph (c) of this section with this device in place,
suitable operating procedures must be established to prevent the use of the
device during takeoff and landing.
(g) If an integral stair is installed in a passenger entry door that is
qualified as a passenger emergency exit, the stair must be designed so that
under the following conditions the effectiveness of passenger emergency
egress will not be impaired:
(1) The door, integral stair, and operating mechanism have been subjected
to the inertial forces specified in paragraph (d) of this section, acting
separately relative to the surrounding structure.
(2) The rotorcraft is in the normal ground attitude and in each of the
attitudes corresponding to collapse of one or more legs, or primary members,
as applicable, of the landing gear.
(h) Nonjettisonable doors used as ditching emergency exits must have means
to enable them to be secured in the open position and remain secure for
emergency egress in sea state conditions prescribed for ditching.
SUMMARY: This rule amends the airworthiness standards for systems,
propulsion, and airframe for both normal and transport category rotorcraft.
In addition, these amendments introduce safety improvements, clarifying
existing regulations, and standardize terminology. The changes are based on
some of the proposals that were submitted to the FAA by the European
Airworthiness Authorities. These amendments are also intended to encourage
the European community's acceptance of the Federal Aviation Regulations for
rotorcraft type certification, obviate development of different European
standards, and achieve increased commonality of airworthiness standards
among the respective countries.
(a) Each seat, safety belt, harness, and adjacent part of the rotorcraft at
each station designated for occupancy during takeoff and landing must be free
of potentially injurious objects, sharp edges, protuberances, and hard
surfaces and must be designed so that a person making proper use of these
facilities will not suffer serious injury in an emergency landing as a result
of the inertial factors specified in Sec. 29.561(b) and dynamic conditions
specified in Sec. 29.562.
(b) Each occupant must be protected from serious head injury by a safety
belt plus a shoulder harness that will prevent the head from contacting any
injurious object, except as provided for in Sec. 29.562(c)(5). A shoulder
harness (upper torso restraint), in combination with the safety belt,
constitutes a torso restraint system as described in TSO-C114.
(c) Each occupant's seat must have a combined safety belt and shoulder
harness with a single-point release. Each pilot's combined safety belt and
shoulder harness must allow each pilot when seated with safety belt and
shoulder harness fastened to perform all functions necessary for flight
operations. There must be a means to secure belt and harness when not in use
to prevent interference with the operation of the rotorcraft and with rapid
egress in an emergency.
(a) Each seat, berth, safety belt, harness, and adjacent part of the
rotorcraft at each station designated for occupancy during takeoff and
landing must be free of potentially injurious objects, sharp edges,
protuberances, and hard surfaces and must be designed so that a person making
proper use of these facilities will not suffer serious injury in an emergency
landing as a result of the inertia forces specified in Sec. 29.561.
(b) Each occupant must be protected from head injury by--
(1) For each crewmember seat and each seat beside a crewmember front seat,
a safety belt and harness that will prevent the head from contacting any
injurious object; and
(2) For each seat not covered under paragraph (b)(1)--
(i) A safety belt plus the absence of injurious objects within striking
radius of the head;
(ii) A safety belt, plus a shoulder harness that will prevent the head from
contracting any injurious object; or
(iii) A safety belt plus an energy-absorbing rest that will support the
arms, shoulders, head and spine.
(c) Each pilot's seat must have a combined safety belt and shoulder harness
with a single-point release that allows the pilot, when seated with safety
belt and shoulder harness fastened, to perform all of the pilot's necessary
functions. There must be a means to secure belts and harnesses, when not in
use, to prevent interference with the operation of the rotorcraft and with
rapid egress in an emergency.
(d) If seat backs do not have a firm handhold, there must be hand grips or
rails along each aisle to let the occupants steady themselves while using the
aisle in moderately rough air.
(e) Each projecting object that would injure persons seated or moving about
in the rotorcraft in normal flight must be padded.
(f) Each seat and its supporting structure must be designed for an occupant
weight of at least 170 pounds, considering the maximum load factors, inertial
forces, and reactions between the occupant, seat, and safety belt or harness
corresponding with the applicable flight and ground-load conditions,
including the emergency landing conditions of Sec. 29.561(b). In addition--
(1) Each pilot seat must be designed for the reactions resulting from the
application of the pilot forces prescribed in Sec. 29.397; and
(2) The inertial forces prescribed in Sec. 29.561(b) must be multiplied by
a factor of 1.33 in determining the strength of the attachment of--
(i) Each seat to the structure; and
(ii) Each safety belt or harness to the seat or structure.
(g) When the safety belt and shoulder harness are combined, the rated
strength of the safety belt and shoulder harness may not be less than that
corresponding to the inertial forces specified in Sec. 29.561(b), considering
the occupant weight of at least 170 pounds, considering the dimensional
characteristics of the restraint system installation, and using a
distribution of at least a 60-percent load to the safety belt and at least a
40-percent load to the shoulder harness. If the safety belt is capable of
being used without the shoulder harness, the inertial forces specified must
be met by the safety belt alone.
(h) When a headrest is used, the headrest and its supporting structure must
be designed to resist the inertia forces specified in Sec. 29.561, with a
1.33 fitting factor and a head weight of at least 13 pounds.
(i) Each seating device system includes the device such as the seat, the
cushions, the occupant restraint system and attachment devices.
(j) Each seating device system may use design features such as crushing or
separation of certain parts of the seat in the design to reduce occupant
loads for the emergency landing dynamic conditions of Sec. 29.562; otherwise,
the system must remain intact and must not interfere with rapid evacuation of
the rotorcraft.
(k) For purposes of this section, a litter is defined as a device designed
to carry a nonambulatory person, primarily in a recumbent position, into and
on the rotorcraft. Each berth or litter must be designed to withstand the
load reaction of an occupant weight of at least 170 pounds when the occupant
is subjected to the forward inertial factors specified in Sec. 29.561(b). A
berth or litter installed within 15 deg. or less of the longitudinal axis of
the rotorcraft must be provided with a padded end-board, cloth diaphragm, or
equivalent means that can withstand the forward load reaction. A berth or
litter oriented greater than 15 deg. with the longitudinal axis of the
rotorcraft must be equipped with appropriate restraints, such as straps or
safety belts, to withstand the forward reaction. In addition--
(1) The berth or litter must have a restraint system and must not have
corners or other protuberances likely to cause serious injury to a person
occupying it during emergency landing conditions; and
(2) The berth or litter attachment and the occupant restraint system
attachments to the structure must be designed to withstand the critical loads
resulting from flight and ground load conditions and from the conditions
prescribed in Sec. 29.561(b).
(a) Each cargo and baggage 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 specified flight and
ground load conditions, except the emergency landing conditions of Sec.
29.561.
(b) There must be means to prevent the contents of any compartment from
becoming a hazard by shifting under the loads specified in paragraph (a) of
this section.
(c) Under the emergency landing conditions of Sec. 29.561, cargo and
baggage compartments must--
(1) Be positioned so that if the contents break loose they are unlikely to
cause injury to the occupants or restrict any of the escape facilities
provided for use after an emergency landing; or
(2) Have sufficient strength to withstand the conditions specified in Sec.
29.561, including the means of restraint and their attachments required by
paragraph (b) of this section. Sufficient strength must be provided for the
maximum authorized weight of cargo and baggage at the critical loading
distribution.
(d) If cargo compartment lamps are installed, each lamp must be installed
so as to prevent contact between lamp bulb and cargo.
SUMMARY: This rule amends the airworthiness standards for systems,
propulsion, and airframe for both normal and transport category rotorcraft.
In addition, these amendments introduce safety improvements, clarifying
existing regulations, and standardize terminology. The changes are based on
some of the proposals that were submitted to the FAA by the European
Airworthiness Authorities. These amendments are also intended to encourage
the European community's acceptance of the Federal Aviation Regulations for
rotorcraft type certification, obviate development of different European
standards, and achieve increased commonality of airworthiness standards
among the respective countries.
(a) If certification with ditching provisions is requested, the rotorcraft
must meet the requirements of this section and Secs. 29.807(d), 29.1411 and
29.1415.
(b) Each practicable design measure, compatible with the general
characteristics of the rotorcraft, must be taken to minimize the probability
that in an emergency landing on water, the behavior of the rotorcraft would
cause immediate injury to the occupants or would make it impossible for them
to escape.
(c) The probable behavior of the rotorcraft in a water landing must be
investigated by model tests or by comparison with rotorcraft of similar
configuration for which the ditching characteristics are known. Scoops,
flaps, projections, and any other factors likely to affect the hydrodynamic
characteristics of the rotorcraft must be considered.
(d) It must be shown that, under reasonably probable water conditions, the
flotation time and trim of the rotorcraft will allow the occupants to leave
the rotorcraft and enter the liferafts required by Sec. 29.1415. If
compliance with this provision is shown by bouyancy and trim computations,
appropriate allowances must be made for probable structural damage and
leakage. If the rotorcraft has fuel tanks (with fuel jettisoning provisions)
that can reasonably be expected to withstand a ditching without leakage, the
jettisonable volume of fuel may be considered as bouyancy volume.
(e) Unless the effects of the collapse of external doors and windows are
accounted for in the investigation of the probable behavior of the rotorcraft
in a water landing (as prescribed in paragraphs (c) and (d) of this section),
the external doors and windows must be designed to withstand the probable
maximum local pressures.
[Amdt. 29-12, 41 FR 55472, Dec. 20, 1976]
Sec. 29.803 Emergency evacuation.
(a) Each crew and passenger area must have means for rapid evacuation in a
crash landing, with the landing gear (1) extended and (2) retracted,
considering the possibility of fire.
(b) Passenger entrance, crew, and service doors may be considered as
emergency exits if they meet the requirements of this section and of Secs.
29.805 through 29.815.
(c) Limited amphibian rotorcraft must meet paragraphs (a) and (b) of this
section. In addition, the following apply:
(1) Each external door, window, and exit must withstand the probable
maximum local water pressures, unless it can be shown that its failure will
not be hazardous to the passengers and crew or have an adverse effect on the
rotorcraft's water stability that would preclude safe evacuation of the
occupants.
(2) At least two exits, one per side, meeting the miminum dimensions of the
exit specified in Sec. 29.807(a)(4) and located above the water level must be
provided for passenger seating capacities up to 39, inclusive. For passenger
seating capacities from 40 to 59, inclusive, two exits, one per side, above
the water level must be provided meeting the minimum dimensions of the exit
specified in Sec. 29.807(a)(3). In all cases, there must be at least one
emergency exit located above the water level for each 35 passengers.
(d) Except as provided in paragraph (e) of this section, the following
categories of rotorcraft must be tested in accordance with the requirements
of appendix D of this part to demonstrate that the maximum seating capacity,
including the crewmembers required by the operating rules, can be evacuated
from the rotorcraft to the ground within 90 seconds:
(1) Rotorcraft with a seating capacity of more than 44 passengers.
(2) Rotorcraft with all of the following:
(i) Ten or more passengers per passenger exit as determined under Sec.
29.807(b).
(ii) No main aisle, as described in Sec. 29.815, for each row of passenger
seats.
(iii) Access to each passenger exit for each passenger by virtue of design
features of seats, such as folding or break-over seat backs or folding seats.
(e) A combination of analysis and tests may be used to show that the
rotorcraft is capable of being evacuated within 90 seconds under the
conditions specified in Sec. 29.803(d) if the Administrator finds that the
combination of analysis and tests will provide data, with respect to the
emergency evacuation capability of the rotorcraft, equivalent to that which
would be obtained by actual demonstration.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 FR
967, Jan. 26, 1968; Amdt. 29-30, 55 FR 8004, Mar. 6, 1990]
SUMMARY: This rule adopts new and revised airworthiness standards for
certification of airframe and related equipment on both normal and transport
category rotorcraft. In addition, one amendment changes an operating rule
affecting external load operators. These amendments grew out of a rotorcraft
regulatory review program and the recognition by both government and industry
that updated safety standards are needed. These amendments provide a high
level of safety in design requirements, while removing certain unnecessary
existing burdens and better utilizing the unique characteristics and
capabilities of rotorcraft.
(a) For rotorcraft with passenger emergency exits that are not convenient
to the flight crew, there must be flight crew emergency exits, on both sides
of the rotorcraft or as a top hatch, in the flight crew area.
(b) Each flight crew emergency exit must be of sufficient size and must be
located so as to allow rapid evacuation of the flight crew. This must be
shown by test.
(c) Each exit must not be obstructed by water or flotation devices after a
ditching. This must be shown by test, demonstration, or analysis.
[Amdt. 29-3, 33 FR 968, Jan. 26, 1968, as amended at Amdt. 29-30, 55 FR 8004,
Mar. 6, 1990]
SUMMARY: This rule adopts new and revised airworthiness standards for
certification of airframe and related equipment on both normal and transport
category rotorcraft. In addition, one amendment changes an operating rule
affecting external load operators. These amendments grew out of a rotorcraft
regulatory review program and the recognition by both government and industry
that updated safety standards are needed. These amendments provide a high
level of safety in design requirements, while removing certain unnecessary
existing burdens and better utilizing the unique characteristics and
capabilities of rotorcraft.
(a) Type. For the purpose of this part, the types of passenger emergency
exit are as follows:
(1) Type I. This type must have a rectangular opening of not less than 24
inches wide by 48 inches high, with corner radii not greater than one-third
the width of the exit, in the passenger area in the side of the fuselage at
floor level and as far away as practicable from areas that might become
potential fire hazards in a crash.
(2) Type II. This type is the same as Type I, except that the opening must
be at least 20 inches wide by 44 inches high.
(3) Type III. This type is the same as Type I, except that--
(i) The opening must be at least 20 inches wide by 36 inches high; and
(ii) The exits need not be at floor level.
(4) Type IV. This type must have a rectangular opening of not less than 19
inches wide by 26 inches high, with corner radii not greater than one-third
the width of the exit, in the side of the fuselage with a step-up inside the
rotorcraft of not more than 29 inches.
Openings with dimensions larger than those specified in this section may be
used, regardless of shape, if the base of the opening has a flat surface of
not less than the specified width.
(b) Passenger emergency exits; side-of-fuselage. Emergency exits must be
accessible to the passengers and, except as provided in paragraph (d) of this
section, must be provided in accordance with the following table:
Emergency exits for
each side of the
fuselage
Passenger
seating Type Type Type Type
capacity I II III IV
1 through 10 1
11 through 19 1 or 2
20 through 39 1 1
40 through 59 1 1
60 through 79 1 1 or 2
(c) Passenger emergency exits; other than side-of-fuselage. In addition to
the requirements of paragraph (b) of this section--
(1) There must be enough openings in the top, bottom, or ends of the
fuselage to allow evacuation with the rotorcraft on its side; or
(2) The probability of the rotorcraft coming to rest on its side in a crash
landing must be extremely remote.
(d) Ditching emergency exits for passengers. If certification with ditching
provisions is requested, ditching emergency exits must be provided in
accordance with the following requirements and must be proven by test,
demonstration, or analysis unless the emergency exits required by paragraph
(b) of this section already meet these requirements.
(1) For rotorcraft that have a passenger seating configuration, excluding
pilots seats, of nine seats or less, one exit above the waterline in each
side of the rotorcraft, meeting at least the dimensions of a Type IV exit.
(2) For rotorcraft that have a passenger seating configuration, excluding
pilots seats, of 10 seats or more, one exit above the waterline in a side of
the rotorcraft meeting at least the dimensions of a Type III exit, for each
unit (or part of a unit) of 35 passenger seats, but no less than two such
exits in the passenger cabin, with one on each side of the rotorcraft.
However, where it has been shown through analysis, ditching demonstrations,
or any other tests found necessary by the Administrator, that the evacuation
capability of the rotorcraft during ditching is improved by the use of larger
exits, or by other means, the passenger seat to exit ratio may be increased.
(3) Flotation devices, whether stowed or deployed, may not interfere with
or obstruct the exits.
(e) Ramp exits. One Type I exit only, or one Type II exit only, that is
required in the side of the fuselage under paragraph (b) of this section, may
be installed instead in the ramp of floor ramp rotorcraft if--
(1) Its installation in the side of the fuselage is impractical; and
(2) Its installation in the ramp meets Sec. 29.813.
(f) Tests. The proper functioning of each emergency exit must be shown by
test.
[Amdt. 29-3, 33 FR 968, Jan. 26, 1968, as amended by Amdt. 29-12, 41 FR
55472, Dec. 20, 1976; Amdt. 29-30, 55 FR 8004, Mar. 6, 1990]
SUMMARY: This rule adopts new and revised airworthiness standards for
certification of airframe and related equipment on both normal and transport
category rotorcraft. In addition, one amendment changes an operating rule
affecting external load operators. These amendments grew out of a rotorcraft
regulatory review program and the recognition by both government and industry
that updated safety standards are needed. These amendments provide a high
level of safety in design requirements, while removing certain unnecessary
existing burdens and better utilizing the unique characteristics and
capabilities of rotorcraft.
(a) Each emergency exit must consist of a movable door or hatch in the
external walls of the fuselage and must provide an unobstructed opening to
the outside.
(b) Each emergency exit must be openable from the inside and from the
outside.
(c) The means of opening each emergency exit must be simple and obvious and
may not require exceptional effort.
(d) There must be means for locking each emergency exit and for preventing
opening in flight inadvertently or as a result of mechanical failure.
(e) There must be means to minimize the probability of the jamming of any
emergency exit in a minor crash landing as a result of fuselage deformation
under the ultimate inertial forces in Sec. 29.783(d).
(f) Except as provided in paragraph (h) of this section, each land-based
rotorcraft emergency exit must have an approved slide as stated in paragraph
(g) of this section, or its equivalent, to assist occupants in descending to
the ground from each floor level exit and an approved rope, or its
equivalent, for all other exits, if the exit threshold is more that 6 feet
above the ground--
(1) With the rotorcraft on the ground and with the landing gear extended;
(2) With one or more legs or part of the landing gear collapsed, broken, or
not extended; and
(3) With the rotorcraft resting on its side, if required by Sec. 29.803(d).
(g) The slide for each passenger emergency exit must be a self-supporting
slide or equivalent, and must be designed to meet the following requirements:
(1) It must be automatically deployed, and deployment must begin during the
interval between the time the exit opening means is actuated from inside the
rotorcraft and the time the exit is fully opened. However, each passenger
emergency exit which is also a passenger entrance door or a service door must
be provided with means to prevent deployment of the slide when the exit is
opened from either the inside or the outside under nonemergency conditions
for normal use.
(2) It must be automatically erected within 10 seconds after deployment is
begun.
(3) It must be of such length after full deployment that the lower end is
self-supporting on the ground and provides safe evacuation of occupants to
the ground after collapse of one or more legs or part of the landing gear.
(4) It must have the capability, in 25-knot winds directed from the most
critical angle, to deploy and, with the assistance of only one person, to
remain usable after full deployment to evacuate occupants safely to the
ground.
(5) Each slide installation must be qualified by five consecutive
deployment and inflation tests conducted (per exit) without failure, and at
least three tests of each such five-test series must be conducted using a
single representative sample of the device. The sample devices must be
deployed and inflated by the system's primary means after being subjected to
the inertia forces specified in Sec. 29.561(b). If any part of the system
fails or does not function properly during the required tests, the cause of
the failure or malfunction must be corrected by positive means and after
that, the full series of five consecutive deployment and inflation tests must
be conducted without failure.
(h) For rotorcraft having 30 or fewer passenger seats and having an exit
threshold more than 6 feet above the ground, a rope or other assist means may
be used in place of the slide specified in paragraph (f) of this section,
provided an evacuation demonstration is accomplished as prescribed in Sec.
29.803(d) or (e).
(i) If a rope, with its attachment, is used for compliance with paragraph
(f), (g), or (h) of this section, it must--
(1) Withstand a 400-pound static load; and
(2) Attach to the fuselage structure at or above the top of the emergency
exit opening, or at another approved location if the stowed rope would reduce
the pilot's view in flight.
SUMMARY: This rule adopts new and revised airworthiness standards for
certification of airframe and related equipment on both normal and transport
category rotorcraft. In addition, one amendment changes an operating rule
affecting external load operators. These amendments grew out of a rotorcraft
regulatory review program and the recognition by both government and industry
that updated safety standards are needed. These amendments provide a high
level of safety in design requirements, while removing certain unnecessary
existing burdens and better utilizing the unique characteristics and
capabilities of rotorcraft.
(a) Each passenger emergency exit, its means of access, and its means of
opening must be conspicuously marked for the guidance of occupants using the
exits in daylight or in the dark. Such markings must be designed to remain
visible for rotorcraft equipped for overwater flights if the rotorcraft is
capsized and the cabin is submerged.
(b) The identity and location of each passenger emergency exit must be
recognizable from a distance equal to the width of the cabin.
(c) The location of each passenger emergency exit must be indicated by a
sign visible to occupants approaching along the main passenger aisle. There
must be a locating sign--
(1) Next to or above the aisle near each floor emergency exit, except that
one sign may serve two exits if both exists can be seen readily from that
sign; and
(2) On each bulkhead or divider that prevents fore and aft vision along the
passenger cabin, to indicate emergency exits beyond and obscured by it,
except that if this is not possible the sign may be placed at another
appropriate location.
(d) Each passenger emergency exit marking and each locating sign must have
white letters 1 inch high on a red background 2 inches high, be self or
electrically illuminated, and have a minimum luminescence (brightness) of at
least 160 microlamberts. The colors may be reversed if this will increase the
emergency illumination of the passenger compartment.
(e) The location of each passenger emergency exit operating handle and
instructions for opening must be shown--
(1) For each emergency exit, by a marking on or near the exit that is
readable from a distance of 30 inches; and
(2) For each Type I or Type II emergency exit with a locking mechanism
released by rotary motion of the handle, by--
(i) A red arrow, with a shaft at least three-fourths inch wide and a head
twice the width of the shaft, extending along at least 70 degrees of arc at a
radius approximately equal to three-fourths of the handle length; and
(ii) The word "open" in red letters 1 inch high, placed horizontally near
the head of the arrow.
(f) Each emergency exit, and its means of opening, must be marked on the
outside of the rotorcraft. In addition, the following apply:
(1) There must be a 2-inch colored band outlining each passenger emergency
exit, except small rotorcraft with a maximum weight of 12,500 pounds or less
may have a 2-inch colored band outlining each exit release lever or device of
passenger emergency exits which are normally used doors.
(2) Each outside marking, including the band, must have color contrast to
be readily distinguishable from the surrounding fuselage surface. The
contrast must be such that, if the reflectance of the darker color is 15
percent or less, the reflectance of the lighter color must be at least 45
percent. "Reflectance" is the ratio of the luminous flux reflected by a body
to the luminous flux it receives. When the reflectance of the darker color is
greater than 15 percent, at least a 30 percent difference between its
reflectance and the reflectance of the lighter color must be provided.
(g) Exits marked as such, though in excess of the required number of exits,
must meet the requirements for emergency exits of the particular type.
Emergency exits need only be marked with the word "Exit."
SUMMARY: This rule amends the airworthiness standards for systems,
propulsion, and airframe for both normal and transport category rotorcraft.
In addition, these amendments introduce safety improvements, clarifying
existing regulations, and standardize terminology. The changes are based on
some of the proposals that were submitted to the FAA by the European
Airworthiness Authorities. These amendments are also intended to encourage
the European community's acceptance of the Federal Aviation Regulations for
rotorcraft type certification, obviate development of different European
standards, and achieve increased commonality of airworthiness standards
among the respective countries.
For transport Category A rotorcraft, the following apply:
(a) A source of light with its power supply independent of the main
lighting system must be installed to--
(1) Illuminate each passenger emergency exit marking and locating sign; and
(2) Provide enough general lighting in the passenger cabin so that the
average illumination, when measured at 40-inch intervals at seat armrest
height on the center line of the main passenger aisle, is at least 0.05 foot-
candle.
(b) Exterior emergency lighting must be provided at each emergency exit.
The illumination may not be less than 0.05 foot-candle (measured normal to
the direction of incident light) for minimum width on the ground surface,
with landing gear extended, equal to the width of the emergency exit where an
evacuee is likely to make first contact with the ground outside the cabin.
The exterior emergency lighting may be provided by either interior or
exterior sources with light intensity measurements made with the emergency
exits open.
(c) Each light required by paragraph (a) or (b) of this section must be
operable manually from the cockpit station and from a point in the passenger
compartment that is readily accessible. The cockpit control device must have
an "on," "off," and "armed" position so that when turned on at the cockpit or
passenger compartment station or when armed at the cockpit station, the
emergency lights will either illuminate or remain illuminated upon
interruption of the rotorcraft's normal electric power.
(d) Any means required to assist the occupants in descending to the ground
must be illuminated so that the erected assist means is visible from the
rotorcraft.
(1) The assist means must be provided with an illumination of not less than
0.03 foot-candle (measured normal to the direction of the incident light) at
the ground end of the erected assist means where an evacuee using the
established escape route would normally make first contact with the ground,
with the rotorcraft in each of the attitudes corresponding to the collapse of
one or more legs of the landing gear.
(2) If the emergency lighting subsystem illuminating the assist means is
independent of the rotorcraft's main emergency lighting system, it--
(i) Must automatically be activated when the assist means is erected;
(ii) Must provide the illumination required by paragraph (d)(1); and
(iii) May not be adversely affected by stowage.
(e) The energy supply to each emergency lighting unit must provide the
required level of illumination for at least 10 minutes at the critical
ambient conditions after an emergency landing.
(f) If storage batteries are used as the energy supply for the emergency
lighting system, they may be recharged from the rotorcraft's main electrical
power system provided the charging circuit is designed to preclude
inadvertent battery discharge into charging circuit faults.
[Amdt. 29-24, 49 FR 44438, Nov. 6, 1984]
Sec. 29.813 Emergency exit access.
(a) Each passageway between passenger compartments, and each passageway
leading to Type I and Type II emergency exits, must be--
(1) Unobstructed; and
(2) At least 20 inches wide.
(b) For each emergency exit covered by Sec. 29.809(f), there must be enough
space adjacent to that exit to allow a crewmember to assist in the evacuation
of passengers without reducing the unobstructed width of the passageway below
that required for that exit.
(c) There must be access from each aisle to each Type III and Type IV exit,
and
(1) For rotorcraft that have a passenger seating configuration, excluding
pilot seats, of 20 or more, the projected opening of the exit provided must
not be obstructed by seats, berths, or other protrusions (including seatbacks
in any position) for a distance from that exit of not less than the width of
the narrowest passenger seat installed on the rotorcraft;
(2) For rotorcraft that have a passenger seating configuration, excluding
pilot seats, of 19 or less, there may be minor obstructions in the region
described in paragraph (c) (1) of this section, if there are compensating
factors to maintain the effectiveness of the exit.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 FR
55472, Dec. 20, 1976]
Sec. 29.815 Main aisle width.
The main passenger aisle width between seats must equal or exceed the
values in the following table:
Minimum main
passenger aisle
width
Less 25
than 25 Inches
inches and more
Passenger from from
seating floor floor
capacity (inches) (inches)
10 or less 12 15
11 through 19 12 20
20 or more 15 20
/1/ A narrower width not less than 9
inches may be approved when
substantiated by tests found necessary
by the Administrator.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 FR
55472, Dec. 20, 1976]
Sec. 29.831 Ventilation.
(a) Each passenger and crew compartment must be ventilated, and each crew
compartment must have enough fresh air (but not less than 10 cu. ft. per
minute per crewmember) to let crewmembers perform their duties without undue
discomfort or fatigue.
(b) Crew and passenger compartment air must be free from harmful or
hazardous concentrations of gases or vapors.
(c) The concentration of carbon monoxide may not exceed one part in 20,000
parts of air during forward flight. If the concentration exceeds this value
under other conditions, there must be suitable operating restrictions.
(d) There must be means to ensure compliance with paragraphs (b) and (c) of
this section under any reasonably probable failure of any ventilating,
heating, or other system or equipment.
Sec. 29.833 Heaters.
Each combustion heater must be approved.
Fire Protection
Sec. 29.851 Fire extinguishers.
(a) Hand fire extinguishers. For hand fire extinguishers the following
apply:
(1) Each hand fire extinguisher must be approved.
(2) The kinds and quantities of each extinguishing agent used must be
appropriate to the kinds of fires likely to occur where that agent is used.
(3) Each extinguisher for use in a personnel compartment must be designed
to minimize the hazard of toxic gas concentrations.
(b) Built-in fire extinguishers. If a built-in fire extinguishing system is
required--
(1) The capacity of each system, in relation to the volume of the
compartment where used and the ventilation rate, must be adequate for any
fire likely to occur in that compartment.
(2) Each system must be installed so that--
(i) No extinguishing agent likely to enter personnel compartments will be
present in a quantity that is hazardous to the occupants; and
(ii) No discharge of the extinguisher can cause structural damage.
Sec. 29.853 Compartment interiors.
For each compartment to be used by the crew or passengers--
(a) The materials (including finishes or decorative surfaces applied to the
materials) must meet the following test criteria as applicable:
(1) 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 Part 25 of this chapter, or other
approved 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.
(2) 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 (a)(3) of this section, must be self extinguishing when tested
vertically in accordance with the applicable portion of Appendix F of Part 25
of this chapter, 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.
(3) Acrylic windows and signs, parts constructed in whole or in part of
elastometric materials, edge lighted instrument assemblies consisting of two
or more instruments in a common housing, seat belts, 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 Part 25 of this
chapter, or other approved equivalent methods.
(4) Except for electrical wire and 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 (a)(1), (a)(2), or (a)(3) of this section may not
have a burn rate greater than 4 inches per minute when tested horizontally in
accordance with the applicable portions of Appendix F of Part 25 of this
chapter, or other approved equivalent methods.
(b) In addition to meeting the requirements of paragraph (a)(2), seat
cushions, except those on flight crewmember seats, must meet the test
requirements of Part II of Appendix F of Part 25 of this chapter, or
equivalent.
(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.
(d) Each receptacle for towels, paper, or waste must be at least fire-
resistant and must have means for containing possible fires;
(e) There must be a hand fire extinguisher for the flight crewmembers; and
(f) At least the following number of hand fire extinguishers must be
conveniently located in passenger compartments:
Passenger Fire
capacity extinguishers
7 through 30 1
31 through 60 2
61 or more 3
(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C.
1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 U.S.C.
1655(c)))
(a) Each cargo and baggage compartment must be construced of or lined with
materials in accordance with the following:
(1) For accessible and inaccessible compartments not occupied by passengers
or crew, the material must be at least fire resistant.
(2) Materials must meet the requirements in Sec. 29.853(a)(1), (a)(2), and
(a)(3) for cargo or baggage compartments in which--
(i) The presence of a compartment fire would be easily discovered by a
crewmember while at the crewmember's station;
(ii) Each part of the compartment is easily accessible in flight;
(iii) The compartment has a volume of 200 cubic feet or less; and
(iv) Notwithstanding Sec. 29.1439(a), protective breathing equipment is not
required.
(b) No compartment may contain any controls, wiring, lines, equipment, or
accessories whose damage or failure would affect safe operation, unless those
items are protected so that--
(1) They cannot be damaged by the movement of cargo in the compartment; and
(2) Their breakage or failure will not create a fire hazard.
(c) The design and sealing of inaccessible compartments must be adequate to
contain compartment fires until a landing and safe evacuation can be made.
(d) Each cargo and baggage compartment that is not sealed so as to contain
cargo compartment fires completely without endangering the safety of a
rotorcraft or its occupants must be designed, or must have a device, to
ensure detection of fires or smoke by a crewmember while at his station and
to prevent the accumulation of harmful quantities of smoke, flame,
extinguishing agents, and other noxious gases in any crew or passenger
compartment. This must be shown in flight.
(e) For rotorcraft used for the carriage of cargo only, the cabin area may
be considered a cargo compartment and, in addition to paragraphs (a) through
(d) of this section, the following apply:
(1) There must be means to shut off the ventilating airflow to or within
the compartment. Controls for this purpose must be accessible to the flight
crew in the crew compartment.
(2) Required crew emergency exits must be accessible under all cargo
loading conditions.
(3) Sources of heat within each compartment must be shielded and insulated
to prevent igniting the cargo.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 FR
969, Jan 26, 1968; Amdt. 29-24, 49 FR 44438, Nov. 6, 1984; Amdt. 29-30, 55 FR
8004, Mar. 6, 1990]
SUMMARY: This rule adopts new and revised airworthiness standards for
certification of airframe and related equipment on both normal and transport
category rotorcraft. In addition, one amendment changes an operating rule
affecting external load operators. These amendments grew out of a rotorcraft
regulatory review program and the recognition by both government and industry
that updated safety standards are needed. These amendments provide a high
level of safety in design requirements, while removing certain unnecessary
existing burdens and better utilizing the unique characteristics and
capabilities of rotorcraft.
(a) Combustion heater fire zones. The following combustion heater fire
zones must be protected against fire under the applicable provisions of Secs.
29.1181 through 29.1191, and 29.1195 through 29.1203:
(1) The region surrounding any heater, if that region contains any
flammable fluid system components (including 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) Each part of any ventilating air passage that--
(i) Surrounds the combustion chamber; and
(ii) Would not contain (without damage to other rotorcraft components) any
fire that may occur within the passage.
(b) Ventilating air ducts. Each ventilating air duct passing through any
fire zone 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 so constructed or isolated from that system
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 communicate with the ventilating airstream
unless flames from backfires or reverse burning cannot enter the ventilating
airstream under any operating condition, including reverse flow or
malfunction 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. There must be means 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. For each combustion heater, safety control
means must be provided as follows:
(1) Means independent of the components provided for the normal continuous
control of air temperature, airflow, and fuel flow must be provided, for each
heater, to automatically shut off the ignition and fuel supply of that heater
at a point remote from that heater when any of the following occurs:
(i) The heat exchanger temperature exceeds safe limits.
(ii) The ventilating air temperature exceeds safe limits.
(iii) The combustion airflow becomes inadequate for safe operation.
(iv) The ventilating airflow becomes inadequate for safe operation.
(2) The means of complying with paragraph (e)(1) of this section for any
individual heater must--
(i) Be independent of components serving any other heater whose heat output
is essential for safe operation; and
(ii) Keep the heater off until restarted by the crew.
(3) There must be 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) of this section.
(f) Air intakes. Each combustion and ventilating air intake must be where
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 malfunction of any other component.
(g) Heater exhaust. Each heater exhaust system must meet the requirements
of Secs. 29.1121 and 29.1123. In addition--
(1) Each exhaust shroud must be sealed so that no flammable fluids or
hazardous quantities of vapors can reach the exhaust systems through joints;
and
(2) No exhaust system may restrict the prompt relief of any backfire that,
if so restricted, could cause heater failure.
(h) Heater fuel systems. Each heater fuel system must meet the powerplant
fuel system requirements affecting safe heater operation. Each heater fuel
system component in 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 for safe drainage of any fuel that might
accumulate in the combustion chamber or the heat 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 against hazardous ice accumulation under
any operating condition.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-2, 32 FR
6914, May 5, 1967]
Sec. 29.861 Fire protection of structure, controls, and other parts.
Each part of the structure, controls, and the rotor mechanism, and other
parts essential to controlled landing and (for category A) flight that would
be affected by powerplant fires must be isolated under Sec. 29.1191, or must
be--
(a) For category A rotorcraft, fireproof; and
(b) For Category B rotorcraft, fireproof or protected so that they can
perform their essential functions for at least 5 minutes under any
foreseeable powerplant fire conditions.
[Doc. 5084, 29 FR 16150, Dec. 3, 1964, as amended at Amdt. 29-30, 55 FR
8005, Mar. 6, 1990]
SUMMARY: This rule adopts new and revised airworthiness standards for
certification of airframe and related equipment on both normal and transport
category rotorcraft. In addition, one amendment changes an operating rule
affecting external load operators. These amendments grew out of a rotorcraft
regulatory review program and the recognition by both government and industry
that updated safety standards are needed. These amendments provide a high
level of safety in design requirements, while removing certain unnecessary
existing burdens and better utilizing the unique characteristics and
capabilities of rotorcraft.
(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 hazards 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 rotorcraft 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.
(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C.
1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 U.S.C.
1655(c)))
[Amdt. 29-17, 43 FR 50600, Oct. 30, 1978]
External Load Attaching Means
Sec. 29.865 External load attaching means.
(a) It must be shown by analysis or test, or both, that the rotorcraft
external load attaching means can withstand a limit static load equal to 2.5,
or some lower factor approved under Secs. 29.337 through 29.341, multiplied
by the maximum external load for which authorization is requested. The load
is applied in the vertical direction and in any direction making an angle of
30 deg. with the vertical, except for those directions having a forward
component. However, the 30 deg. angle may be reduced to a lesser angle if--
(1) An operating limitation is established limiting external load
operations to such angles for which compliance with this paragraph has been
shown; or
(2) It is shown that the lesser angle can not be exceeded in service.
(b) The external load attaching means for Class B and Class C rotorcraft-
load combinations must include a device to enable the pilot to release the
external load quickly during flight. This quick-release device, and the means
by which it is controlled, must comply with the following:
(1) A control for the quick-release device must be installed on one of the
pilot's primary controls and must be designed and located so that it may be
operated by the pilot without hazardously limiting his ability to control the
rotorcraft during an emergency situation.
(2) In addition a manual mechanical control for the quick-release device,
readily accessible either to the pilot or to another crew member, must be
provided.
(3) The quick-release device must function properly with all external loads
up to and including the maximum external load for which authorization is
requested.
(c) A placard or marking must be installed next to the external-load
attaching means stating the maximum authorized external load as demonstrated
under Sec. 29.25 and this section.
(d) The fatigue evaluation of Sec. 29.571(a) does not apply to this section
except for a failure of the cargo attaching means that results in a hazard to
the rotorcraft.
[Amdt. 29-12, 41 FR 55472, Dec. 20, 1976, as amended at Amdt. 29-30, 55 FR
8005, Mar. 6, 1990]
SUMMARY: This rule adopts new and revised airworthiness standards for
certification of airframe and related equipment on both normal and transport
category rotorcraft. In addition, one amendment changes an operating rule
affecting external load operators. These amendments grew out of a rotorcraft
regulatory review program and the recognition by both government and industry
that updated safety standards are needed. These amendments provide a high
level of safety in design requirements, while removing certain unnecessary
existing burdens and better utilizing the unique characteristics and
capabilities of rotorcraft.
There must be reference marks for leveling the rotorcraft on the ground.
Sec. 29.873 Ballast provisions.
Ballast provisions must be designed and constructed to prevent inadvertent
shifting of ballast in flight.
Subpart E--Powerplant
General
Sec. 29.901 Installation.
(a) For the purpose of this part, the powerplant installation includes each
part of the rotorcraft (other than the main and auxiliary rotor structures)
that--
(1) Is necessary for propulsion;
(2) Affects the control of the major propulsive units; or
(3) Affects the safety of the major propulsive units between normal
inspections or overhauls.
(b) For each powerplant installation--
(1) The installation must comply with--
(i) The installation instructions provided under Sec. 33.5 of this chapter;
and
(ii) The applicable provisions of this subpart.
(2) Each component of the installation must be constructed, arranged, and
installed to ensure its continued safe operation between normal inspections
or overhauls for the range of temperature and altitude for which approval is
requested.
(3) Accessibility must be provided to allow any inspection and maintenance
necessary for continued airworthiness; and
(4) Electrical interconnections must be provided to prevent differences of
potential between major components of the installation and the rest of the
rotorcraft.
(5) Axial and radial expansion of turbine engines may not affect the safety
of the installation.
(6) Design precautions must be taken to minimize the possibility of
incorrect assembly of components and equipment essential to safe operation of
the rotorcraft, except where operation with the incorrect assembly can be
shown to be extremely improbable.
(c) For each powerplant and auxiliary power unit installation, it must be
established that no single failure or malfunction or probable combination of
failures will jeopardize the safe operation of the rotorcraft except that--
(1) The failure of structural elements need not be considered if the
probability of such failure is extremely remote; and
(2) The failure of engine rotor discs need not be considered.
(d) Each auxiliary power unit installation must meet the applicable
provisions of this subpart.
(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C.
1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 U.S.C.
1655(c)))
(a) Engine type certification. Each engine must have an approved type
certificate. Reciprocating engines for use in helicopters must be qualified
in accordance with Sec. 33.49(d) of this chapter or be otherwise approved for
the intended usage.
(b) Category A; engine isolation. For each category A rotorcraft, 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 of any system that can affect any engine, will not--
(1) Prevent the continued safe operation of the remaining engines; or
(2) Require immediate action, other than normal pilot action with primary
flight controls, by any crewmember to maintain safe operation.
(c) Category A; control of engine rotation. For each Category A rotorcraft,
there must be a means for stopping the rotation of any engine individually in
flight, except that, for turbine engine installations, the means for stopping
the engine need be provided only where necessary for safety. In addition--
(1) Each component of the engine stopping system that is located on the
engine side of the firewall, and that might be exposed to fire, must be at
least fire resistant; or
(2) Duplicate means must be available for stopping the engine and the
controls must be where all are not likely to be damaged at the same time in
case of fire.
(d) Turbine engine installation. For turbine engine installations, the
powerplant systems associated with engine control devices, systems, and
instrumentation must be designed to give reasonable assurance that those
engine operating limitations that adversely affect turbine rotor structural
integrity will not be exceeded in service.
(e) Restart capability. (1) A means to restart any engine in flight must be
provided.
(2) Except for the in-flight shutdown of all engines, engine restart
capability must be demonstrated throughout a flight envelope for the
rotorcraft.
(3) Following the in-flight shutdown of all engines, in-flight engine
restart capability must be provided.
(Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 U.S.C. 1354(a), 1421, and
1423; sec. 6(c), 49 U.S.C. 1655(c))
SUMMARY: This rule amends the airworthiness standards for systems,
propulsion, and airframe for both normal and transport category rotorcraft.
In addition, these amendments introduce safety improvements, clarifying
existing regulations, and standardize terminology. The changes are based on
some of the proposals that were submitted to the FAA by the European
Airworthiness Authorities. These amendments are also intended to encourage
the European community's acceptance of the Federal Aviation Regulations for
rotorcraft type certification, obviate development of different European
standards, and achieve increased commonality of airworthiness standards
among the respective countries.
(a) Each engine must be installed to prevent the harmful vibration of any
part of the engine or rotorcraft.
(b) The addition of the rotor and the rotor drive system to the engine may
not subject the principal rotating parts of the engine to excessive vibration
stresses. This must be shown by a vibration investigation.
Sec. 29.908 Cooling fans.
For cooling fans that are a part of a powerplant installation the following
apply:
(a) Category A. For cooling fans installed in Category A rotorcraft, it
must be shown that a fan blade failure will not prevent continued safe flight
either because of damage caused by the failed blade or loss of cooling air.
(b) Category B. For cooling fans installed in category B rotorcraft, there
must be means to protect the rotorcraft and allow a safe landing if a fan
blade fails. It must be shown that--
(1) The fan blade would be contained in the case of a failure;
(2) Each fan is located so that a fan blade failure will not jeopardize
safety; or
(3) Each fan blade can withstand an ultimate load of 1.5 times the
centrifugal force expected in service, limited by either--
(i) The highest rotational speeds achievable under uncontrolled conditions;
or
(ii) An overspeed limiting device.
(c) Fatigue evaluation. Unless a fatigue evaluation under Sec. 29.571 is
conducted, it must be shown that cooling fan blades are not operating at
resonant conditions within the operating limits of the rotorcraft.
(Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 U.S.C. 1354(a), 1421, and
1423; sec. 6(c), 49 U.S.C. 1655 (c))
[Amdt. 29-13, 42 FR 15046, Mar. 17, 1977, as amended by Amdt. 29-26, 53 FR
34215, Sept. 2, 1988]
Rotor Drive System
Sec. 29.917 Design.
(a) General. The rotor drive system includes any part necessary to transmit
power from the engines to the rotor hubs. This includes gear boxes, shafting,
universal joints, couplings, rotor brake assemblies, clutches, supporting
bearings for shafting, any attendant accessory pads or drives, and any
cooling fans that are a part of, attached to, or mounted on the rotor drive
system.
(b) Arrangement. Rotor drive systems must be arranged as follows:
(1) Each rotor drive system of multiengine rotorcraft must be arranged so
that each rotor necessary for operation and control will continue to be
driven by the remaining engines if any engine fails.
(2) For single-engine rotorcraft, each rotor drive system must be so
arranged that each rotor necessary for control in autorotation will continue
to be driven by the main rotors after disengagement of the engine from the
main and auxiliary rotors.
(3) Each rotor drive system must incorporate a unit for each engine to
automatically disengage that engine from the main and auxiliary rotors if
that engine fails.
(4) If a torque limiting device is used in the rotor drive system, it must
be located so as to allow continued control of the rotorcraft when the device
is operating.
(5) If the rotors must be phased for intermeshing, each system must provide
constant and positive phase relationship under any operating condition.
(6) If a rotor dephasing device is incorporated, there must be means to
keep the rotors locked in proper phase before operation.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 FR
55472, Dec. 20, 1976]
Sec. 29.921 Rotor brake.
If there is a means to control the rotation of the rotor drive system
independently of the engine, any limitations on the use of that means must be
specified, and the control for that means must be guarded to prevent
inadvertent operation.
Sec. 29.923 Rotor drive system and control mechanism tests.
(a) Endurance tests, general. Each rotor drive system and rotor control
mechanism must be tested, as prescribed in paragraphs (b) through (n) of this
section, for at least 200 hours plus the time required to meet the
requirements of paragraphs (b)(2), (b)(3), and (k) of this section. These
tests must be conducted as follows:
(1) Ten-hour test cycles must be used, except that the test cycle must be
extended to include the OEI test of paragraphs (b)(2) and (k), of this
section if OEI ratings are requested.
(2) The tests must be conducted on the rotorcraft.
(3) The test torque and rotational speed must be--
(i) Determined by the powerplant limitations; and
(ii) Absorbed by the rotors to be approved for the rotorcraft.
(b) Endurance tests; takeoff run. The takeoff run must be conducted as
follows:
(1) Except as prescribed in paragraphs (b)(2) and (b)(3) of this section,
the takeoff torque run must consist of 1 hour of alternate runs of 5 minutes
at takeoff torque and the maximum speed for use with takeoff torque, and 5
minutes at as low an engine idle speed as practicable. The engine must be
declutched from the rotor drive system, and the rotor brake, if furnished and
so intended, must be applied during the first minute of the idle run. During
the remaining 4 minutes of the idle run, the clutch must be engaged so that
the engine drives the rotors at the minimum practical r.p.m. The engine and
the rotor drive system must be accelerated at the maximum rate. When
declutching the engine, it must be decelerated rapidly enough to allow the
operation of the overrunning clutch.
(2) For helicopters for which the use of a 2 1/2 -minute OEI rating is
requested, the takeoff run must be conducted as prescribed in paragraph
(b)(1) of this section, except for the third and sixth runs for which the
takeoff torque and the maximum speed for use with takeoff torque are
prescribed in that paragraph. For these runs, the following apply:
(i) Each run must consist of at least one period of 2 1/2 minutes with
takeoff torque and the maximum speed for use with takeoff torque on all
engines.
(ii) Each run must consist of at least one period, for each engine in
sequence, during which that engine simulates a power failure and the
remaining engines are run at the 2 1/2 -minute OEI torque and the maximum
speed for use with 2 1/2 -minute OEI torque for 2 1/2 minutes.
(3) For multiengine, turbine-powered rotorcraft for which the use of 30-
second/2-minute OEI power is requested, the takeoff run must be conducted as
prescribed in paragraph (b)(1) of this section except for the following:
(i) Immediately following any one 5-minute power-on run required by
paragraph (b)(1) of this section, each power source must simulate a failure,
in turn, and apply the maximum torque and the maximum speed for use with 30-
second OEI power to the remaining affected drive system power inputs for not
less than 30 seconds, followed by application of the maximum torque and the
maximum speed for use with 2-minute OEI power for not less than 2 minutes. At
least one run sequence must be conducted from a simulated "flight idle"
condition. When conducted on a bench test, the test sequence must be
conducted following stabilization at takeoff power.
(ii) For the purpose of this paragraph, an affected power input includes
all parts of the rotor drive system which can be adversely affected by the
application of higher or asymmetric torque and speed prescribed by the test.
(iii) This test may be conducted on a representative bench test facility
when engine limitations either preclude repeated use of this power or would
result in premature engine removals during the test. The loads, the vibration
frequency, and the methods of application to the affected rotor drive system
components must be representative of rotorcraft conditions. Test components
must be those used to show compliance with the remainder of this section.
(c) Endurance tests; maximum continuous run. Three hours of continuous
operation at maximum continuous torque and the maximum speed for use with
maximum continuous torque must be conducted as follows:
(1) The main rotor controls must be operated at a minimum of 15 times each
hour through the main rotor pitch positions of maximum vertical thrust,
maximum forward thrust component, maximum aft thrust component, maximum left
thrust component, and maximum right thrust component, except that the control
movements need not produce loads or blade flapping motion exceeding the
maximum loads of motions encountered in flight.
(2) The directional controls must be operated at a minimum of 15 times each
hour through the control extremes of maximum right turning torque, neutral
torque as required by the power applied to the main rotor, and maximum left
turning torque.
(3) Each maximum control position must be held for at least 10 seconds, and
the rate of change of control position must be at least as rapid as that for
normal operation.
(d) Endurance tests; 90 percent of maximum continuous run. One hour of
continuous operation at 90 percent of maximum continuous torque and the
maximum speed for use with 90 percent of maximum continuous torque must be
conducted.
(e) Endurance tests; 80 percent of maximum continuous run. One hour of
continuous operation at 80 percent of maximum continuous torque and the
minimum speed for use with 80 percent of maximum continuous torque must be
conducted.
(f) Endurance tests; 60 percent of maximum continuous run. Two hours or,
for helicopters for which the use of either 30-minute OEI power or continuous
OEI power is requested, 1 hour of continuous operation at 60 percent of
maximum continuous torque and the minimum speed for use with 60 percent of
maximum continuous torque must be conducted.
(g) Endurance tests; engine malfunctioning run. It must be determined
whether malfunctioning of components, such as the engine fuel or ignition
systems, or whether unequal engine power can cause dynamic conditions
detrimental to the drive system. If so, a suitable number of hours of
operation must be accomplished under those conditions, 1 hour of which must
be included in each cycle, and the remaining hours of which must be
accomplished at the end of the 20 cycles. If no detrimental condition
results, an additional hour of operation in compliance with paragraph (b) of
this section must be conducted in accordance with the run schedule of
paragraph (b)(1) of this section without consideration of paragraph (b)(2) of
this section.
(h) Endurance tests; overspeed run. One hour of continuous operation must
be conducted at maximum continuous torque and the maximum power-on overspeed
expected in service, assuming that speed and torque limiting devices, if any,
function properly.
(i) Endurance tests; rotor control positions. When the rotor controls are
not being cycled during the tie-down tests, the rotor must be operated, using
the procedures prescribed in paragraph (c) of this section, to produce each
of the maximum thrust positions for the following percentages of test time
(except that the control positions need not produce loads or blade flapping
motion exceeding the maximum loads or motions encountered in flight):
(1) For full vertical thrust, 20 percent.
(2) For the forward thrust component, 50 percent.
(3) For the right thrust component, 10 percent.
(4) For the left thrust component, 10 percent.
(5) For the aft thrust component, 10 percent.
(j) Endurance tests, clutch and brake engagements. A total of at least 400
clutch and brake engagements, including the engagements of paragraph (b) of
this section, must be made during the takeoff torque runs and, if necessary,
at each change of torque and speed throughout the test. In each clutch
engagement, the shaft on the driven side of the clutch must be accelerated
from rest. The clutch engagements must be accomplished at the speed and by
the method prescribed by the applicant. During deceleration after each clutch
engagement, the engines must be stopped rapidly enough to allow the engines
to be automatically disengaged from the rotors and rotor drives. If a rotor
brake is installed for stopping the rotor, the clutch, during brake
engagements, must be disengaged above 40 percent of maximum continuous rotor
speed and the rotors allowed to decelerate to 40 percent of maximum
continuous rotor speed, at which time the rotor brake must be applied. If the
clutch design does not allow stopping the rotors with the engine running, or
if no clutch is provided, the engine must be stopped before each application
of the rotor brake, and then immediately be started after the rotors stop.
(k) Endurance tests; OEI power run--(1) 30-minute OEI power run. For
rotorcraft for which the use of 30-minute OEI power is requested, a run at
30-minute OEI torque and the maximum speed for use with 30-minute OEI torque
must be conducted as follows: For each engine, in sequence, that engine must
be inoperative and the remaining engines must be run for a 30-minute period.
(2) Continuous OEI power run. For rotorcraft for which the use of
continuous OEI power is requested, a run at continuous OEI torque and the
maximum speed for use with continuous OEI torque must be conducted as
follows: For each engine, in sequence, that engine must be inoperative and
the remaining engines must be run for 1 hour.
(3) The number of periods prescribed in paragraph (k)(1) or (k)(2) of this
section may not be less than the number of engines, nor may it be less than
two.
(l) [Reserved]
(m) Any components that are affected by maneuvering and gust loads must be
investigated for the same flight conditions as are the main rotors, and their
service lives must be determined by fatigue tests or by other acceptable
methods. In addition, a level of safety equal to that of the main rotors must
be provided for--
(1) Each component in the rotor drive system whose failure would cause an
uncontrolled landing;
(2) Each component essential to the phasing of rotors on multirotor
rotorcraft, or that furnishes a driving link for the essential control of
rotors in autorotation; and
(3) Each component common to two or more engines on multiengine rotorcraft.
(n) Special tests. Each rotor drive system designed to operate at two or
more gear ratios must be subjected to special testing for durations necessary
to substantiate the safety of the rotor drive system.
(o) Each part tested as prescribed in this section must be in a serviceable
condition at the end of the tests. No intervening disassembly which might
affect test results may be conducted.
(p) Endurance tests; operating lubricants. To be approved for use in rotor
drive and control systems, lubricants must meet the specifications of
lubricants used during the tests prescribed by this section. Additional or
alternate lubricants may be qualified by equivalent testing or by comparative
analysis of lubricant specifications and rotor drive and control system
characteristics. In addition--
(1) At least three 10-hour cycles required by this section must be
conducted with transmission and gearbox lubricant temperatures, at the
location prescribed for measurement, not lower than the maximum operating
temperature for which approval is requested;
(2) For pressure lubricated systems, at least three 10-hour cycles required
by this section must be conducted with the lubricant pressure, at the
location prescribed for measurement, not higher than the minimum operating
pressure for which approval is requested; and
(3) The test conditions of paragraphs (p)(1) and (p)(2) of this section
must be applied simultaneously and must be extended to include operation at
any one-engine-inoperative rating for which approval is requested.
(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C.
1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 U.S.C.
1655(c)))
SUMMARY: This rule adopts new and revised airworthiness standards by
incorporating optional one-engine-inoperative (OEI) power ratings for
multiengine, turbine-powered rotorcraft. These amendments result from a
petition for rulemaking from Aerospace Industries Association of America
(AIA) and the recognition by both government and industry that additional OEI
power rating standards are needed. These amendments enhance rotorcraft safety
after an engine failure or precautionary shutdown by providing higher OEI
power, when necessary. These amendments also assure that the drive system
will maintain its structural integrity and allow continued safe flight while
operating at the new OEI power ratings with the operable engine(s).
(a) Any additional dynamic, endurance, and operational tests, and vibratory
investigations necessary to determine that the rotor drive mechanism is safe,
must be performed.
(b) If turbine engine torque output to the transmission can exceed the
highest engine or transmission torque limit, and that output is not directly
controlled by the pilot under normal operating conditions (such as where the
primary engine power control is accomplished through the flight control), the
following test must be made:
(1) Under conditions associated with all engines operating, make 200
applications, for 10 seconds each, of torque that is at least equal to the
lesser of--
(i) The maximum torque used in meeting Sec. 29.923 plus 10 percent; or
(ii) The maximum torque attainable under probable operating conditions,
assuming that torque limiting devices, if any, function properly.
(2) For multiengine rotorcraft under conditions associated with each
engine, in turn, becoming inoperative, apply to the remaining transmission
torque inputs the maximum torque attainable under probable operating
conditions, assuming that torque limiting devices, if any, function properly.
Each transmission input must be tested at this maximum torque for at least
fifteen minutes.
(c) Lubrication system failure. For lubrication systems required for proper
operation of rotor drive systems, the following apply:
(1) Category A. Unless such failures are extremely remote, it must be shown
by test that any failure which results in loss of lubricant in any normal use
lubrication system will not prevent continued safe operation, although not
necessarily without damage, at a torque and rotational speed prescribed by
the applicant for continued flight, for at least 30 minutes after perception
by the flightcrew of the lubrication system failure or loss of lubricant.
(2) Category B. The requirements of Category A apply except that the rotor
drive system need only be capable of operating under autorotative conditions
for at least 15 minutes.
(d) Overspeed test. The rotor drive system must be subjected to 50
overspeed runs, each 30+/-3 seconds in duration, at not less than either the
higher of the rotational speed to be expected from an engine control device
failure or 105 percent of the maximum rotational speed, including transients,
to be expected in service. If speed and torque limiting devices are
installed, are independent of the normal engine control, and are shown to be
reliable, their rotational speed limits need not be exceeded. These runs must
be conducted as follows:
(1) Overspeed runs must be alternated with stabilizing runs of from 1 to 5
minutes duration each at 60 to 80 percent of maximum continuous speed.
(2) Acceleration and deceleration must be accomplished in a period not
longer than 10 seconds (except where maximum engine acceleration rate will
require more than 10 seconds), and the time for changing speeds may not be
deducted from the specified time for the overspeed runs.
(3) Overspeed runs must be made with the rotors in the flattest pitch for
smooth operation.
(e) The tests prescribed in paragraphs (b) and (d) of this section must be
conducted on the rotorcraft and the torque must be absorbed by the rotors to
be installed, except that other ground or flight test facilities with other
appropriate methods of torque absorption may be used if the conditions of
support and vibration closely simulate the conditions that would exist during
a test on the rotorcraft.
(f) Each test prescribed by this section must be conducted without
intervening disassembly and, except for the lubrication system failure test
required by paragraph (c) of this section, each part tested must be in a
serviceable condition at the conclusion of the test.
(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C.
1354(a), 1421, 1423 1424), sec. 6(c), Dept. of Transportation Act (49 U.S.C.
1655(c)))
(a) The critical speeds of any shafting must be determined by demonstration
except that analytical methods may be used if reliable methods of analysis
are available for the particular design.
(b) If any critical speed lies within, or close to, the operating ranges
for idling, power-on, and autorotative conditions, the stresses occurring at
that speed must be within safe limits. This must be shown by tests.
(c) If analytical methods are used and show that no critical speed lies
within the permissible operating ranges, the margins between the calculated
critical speeds and the limits of the allowable operating ranges must be
adequate to allow for possible variations between the computed and actual
values.
[Amdt. 29-12, 41 FR 55472, Dec. 20, 1976]
Sec. 29.935 Shafting joints.
Each universal joint, slip joint, and other shafting joints whose
lubrication is necessary for operation must have provision for lubrication.
Sec. 29.939 Turbine engine operating characteristics.
(a) Turbine engine operating characteristics must be investigated in flight
to determine that no adverse characteristics (such as stall, surge, of
flameout) are present, to a hazardous degree, during normal and emergency
operation within the range of operating limitations of the rotorcraft and of
the engine.
(b) The turbine engine air inlet system may not, as a result of airflow
distortion during normal operation, cause vibration harmful to the engine.
(c) For governor-controlled engines, it must be shown that there exists no
hazardous torsional instability of the drive system associated with critical
combinations of power, rotational speed, and control displacement.
[Amdt. 29-2, 32 FR 6914, May 5, 1967, as amended by Amdt. 29-12, 41 FR 55473,
Dec. 20, 1976]
Fuel System
Sec. 29.951 General.
(a) Each fuel system must be constructed and arranged to ensure a flow of
fuel at a rate and pressure established for proper engine and auxiliary power
unit functioning under any likely operating conditions, including the
maneuvers 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 engine or 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 degrees 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.
Unless other means acceptable to the Administrator are employed to minimize
the hazard of fuel fires to occupants following an otherwise survivable
impact (crash landing), the fuel systems must incorporate the design features
of this section. These systems must be shown to be capable of sustaining the
static and dynamic deceleration loads of this section, considered as ultimate
loads acting alone, measured at the system component's center of gravity
without structural damage to the system components, fuel tanks, or their
attachments that would leak fuel to an ignition source.
(a) Drop test requirements. Each tank, or the most critical tank, must be
drop-tested as follows:
(1) The drop height must be at least 50 feet.
(2) The drop impact surface must be nondeforming.
(3) The tanks must be filled with water to 80 percent of the normal, full
capacity.
(4) The tank must be enclosed in a surrounding structure representative of
the installation unless it can be established that the surrounding structure
is free of projections or other design features likely to contribute to
rupture of the tank.
(5) The tank must drop freely and impact in a horizontal position +/-10
deg..
(6) After the drop test, there must be no leakage.
(b) Fuel tank load factors. Except for fuel tanks located so that tank
rupture with fuel release to either significant ignition sources, such as
engines, heaters, and auxiliary power units, or occupants is extremely
remote, each fuel tank must be designed and installed to retain its contents
under the following ultimate inertial load factors, acting alone.
(1) For fuel tanks in the cabin:
(i) Upward--4g.
(ii) Forward--16g.
(iii) Sideward--8g.
(iv) Downward--20g.
(2) For fuel tanks located above or behind the crew or passenger
compartment that, if loosened, could injure an occupant in an emergency
landing:
(i) Upward--1.5g.
(ii) Forward--8g.
(iii) Sideward--2g.
(iv) Downward--4g.
(3) For fuel tanks in other areas:
(i) Upward--1.5g.
(ii) Forward--4g.
(iii) Sideward--2g.
(iv) Downward--4g.
(c) Fuel line self-sealing breakaway couplings. Self-sealing breakaway
couplings must be installed unless hazardous relative motion of fuel system
components to each other or to local rotorcraft structure is demonstrated to
be extremely improbable or unless other means are provided. The couplings or
equivalent devices must be installed at all fuel tank-to-fuel line
connections, tank-to-tank interconnects, and at other points in the fuel
system where local structural deformation could lead to the release of fuel.
(1) The design and construction of self-sealing breakaway couplings must
incorporate the following design features:
(i) The load necessary to separate a breakaway coupling must be between 25
to 50 percent of the minimum ultimate failure load (ultimate strength) of the
weakest component in the fluid-carrying line. The separation load must in no
case be less than 300 pounds, regardless of the size of the fluid line.
(ii) A breakaway coupling must separate whenever its ultimate load (as
defined in paragraph (c)(1)(i) of this section) is applied in the failure
modes most likely to occur.
(iii) All breakaway couplings must incorporate design provisions to
visually ascertain that the coupling is locked together (leak-free) and is
open during normal installation and service.
(iv) All breakaway couplings must incorporate design provisions to prevent
uncoupling or unintended closing due to operational shocks, vibrations, or
accelerations.
(v) No breakaway coupling design may allow the release of fuel once the
coupling has performed its intended function.
(2) All individual breakaway couplings, coupling fuel feed systems, or
equivalent means must be designed, tested, installed, and maintained so
inadvertent fuel shutoff in flight is improbable in accordance with Sec.
29.955(a) and must comply with the fatigue evaluation requirements of Sec.
29.571 without leaking.
(3) Alternate, equivalent means to the use of breakaway couplings must not
create a survivable impact-induced load on the fuel line to which it is
installed greater than 25 to 50 percent of the ultimate load (strength) of
the weakest component in the line and must comply with the fatigue
requirements of Sec. 29.571 without leaking.
(d) Frangible or deformable structural attachments. Unless hazardous
relative motion of fuel tanks and fuel system components to local rotorcraft
structure is demonstrated to be extremely improbable in an otherwise
survivable impact, frangible or locally deformable attachments of fuel tanks
and fuel system components to local rotorcraft structure must be used. The
attachment of fuel tanks and fuel system components to local rotorcraft
structure, whether frangible or locally deformable, must be designed such
that its separation or relative local deformation will occur without rupture
or local tear-out of the fuel tank or fuel system component that will cause
fuel leakage. The ultimate strength of frangible or deformable attachments
must be as follows:
(1) The load required to separate a frangible attachment from its support
structure, or deform a locally deformable attachment relative to its support
structure, must be between 25 and 50 percent of the minimum ultimate load
(ultimate strength) of the weakest component in the attached system. In no
case may the load be less than 300 pounds.
(2) A frangible or locally deformable attachment must separate or locally
deform as intended whenever its ultimate load (as defined in paragraph (d)(1)
of this section) is applied in the modes most likely to occur.
(3) All frangible or locally deformable attachments must comply with the
fatigue requirements of Sec. 29.571.
(e) Separation of fuel and ignition sources. To provide maximum crash
resistance, fuel must be located as far as practicable from all occupiable
areas and from all potential ignition sources.
(f) Other basic mechanical design criteria. Fuel tanks, fuel lines,
electrical wires, and electrical devices must be designed, constructed, and
installed, as far as practicable, to be crash resistant.
(g) Rigid or semirigid fuel tanks. Rigid or semirigid fuel tank or bladder
walls must be impact and tear resistant.
SUMMARY: These amendments add comprehensive crash resistant fuel system
design and test criteria to the airworthiness standards for normal and
transport category rotorcraft. Application of these standards will minimize
fuel spillage near ignition sources and potential ignition sources and,
therefore, will improve the evacuation time needed for crew and passengers to
escape a post-crash fire (PCF). Implementation of these amendments will
minimize the PCF hazard saving lives and substantially reducing the severity
of physiological injuries sustained from PCF's in otherwise survivable
accidents.
(a) For category A rotorcraft--
(1) The fuel system must meet the requirements of Sec. 29.903(b); and
(2) Unless other provisions are made to meet paragraph (a)(1) of this
section, the fuel system must allow fuel to be supplied to each engine
through a system independent of those parts of each system supplying fuel to
other engines.
(b) Each fuel system for a multiengine category B rotorcraft must meet the
requirements of paragraph (a)(2) of this section. However, separate fuel
tanks need not be provided for each engine.
Sec. 29.954 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 to areas where swept strokes are highly
probable; and
(c) Corona and streamering at fuel vent outlets.
[Amdt. 29-26, 53 FR 34217, Sept. 2, 1988]
Sec. 29.955 Fuel flow.
(a) General. The fuel system for each engine must provide the engine with
at least 100 percent of the fuel required under all operating and maneuvering
conditions to be approved for the rotorcraft, including, as applicable, the
fuel required to operate the engines under the test conditions required by
Sec. 29.927. Unless equivalent methods are used, compliance must be shown by
test during which the following provisions are met, except that combinations
of conditions which are shown to be improbable need not be considered.
(1) The fuel pressure, corrected for accelerations (load factors), must be
within the limits specified by the engine type certificate data sheet.
(2) The fuel level in the tank may not exceed that established as the
unusable fuel supply for that tank under Sec. 29.959, plus that necessary to
conduct the test.
(3) The fuel head between the tank and the engine must be critical with
respect to rotorcraft flight attitudes.
(4) The fuel flow transmitter, if installed, and the critical fuel pump
(for pump-fed systems) must be installed to produce (by actual or simulated
failure) the critical restriction to fuel flow to be expected from component
failure.
(5) Critical values of engine rotational speed, electrical power, or other
sources of fuel pump motive power must be applied.
(6) Critical values of fuel properties which adversely affect fuel flow are
applied during demonstrations of fuel flow capability.
(7) The fuel filter required by Sec. 29.997 is blocked to the degree
necessary to simulate the accumulation of fuel contamination required to
activate the indicator required by Sec. 29.1305(a)(17).
(b) Fuel transfer system. If normal operation of the fuel system requires
fuel to be transferred to another tank, the transfer must occur automatically
via a system which has been shown to maintain the fuel level in the receiving
tank within acceptable limits during flight or surface operation of the
rotorcraft.
(c) Multiple fuel tanks. If an engine can be supplied with fuel from more
than one tank, the fuel system, in addition to having appropriate manual
switching capability, must be designed to prevent interruption of fuel flow
to that engine, without attention by the flightcrew, when any tank supplying
fuel to that engine is depleted of usable fuel during normal operation and
any other tank that normally supplies fuel to that engine alone contains
usable fuel.
[Amdt. 29-26, 53 FR 34217, Sept. 2, 1988]
Sec. 29.957 Flow between interconnected tanks.
(a) Where tank outlets are interconnected and allow fuel to flow between
them due to gravity or flight accelerations, it must be impossible for fuel
to flow between tanks in quantities great enough to cause overflow from the
tank vent in any sustained flight condition.
(b) If fuel can be pumped from one tank to another in flight--
(1) The design of the vents and the fuel transfer system must prevent
structural damage to tanks from overfilling; and
(2) There must be means to warn the crew before overflow through the vents
occurs.
Sec. 29.959 Unusable fuel supply.
The unusable fuel supply for each tank must be established as not less than
the quantity at which the first evidence of malfunction occurs under the most
adverse fuel feed condition occurring under any intended operations and
flight maneuvers involving that tank.
Sec. 29.961 Fuel system hot weather operation.
Each suction lift fuel system and other fuel systems conducive to vapor
formation must be shown to operate satisfactorily (within certification
limits) when using fuel at the most critical temperature for vapor formation
under critical operating conditions including, if applicable, the engine
operating conditions defined by Sec. 29.927(b)(1) and (b)(2).
[Amdt. 29-26, 53 FR 34217, Sept. 2, 1988]
Sec. 29.963 Fuel tanks: general.
(a) Each fuel tank must be able to withstand, without failure, the
vibration, inertia, fluid, and structural loads to which it may be subjected
in operation.
(b) Each flexible fuel tank bladder or liner must be approved or shown to
be suitable for the particular application and must be puncture resistant.
Puncture resistance must be shown by meeting the TSO-C80, paragraph 16.0,
requirements using a minimum puncture force of 370 pounds.
(c) Each integral fuel tank must have facilities for inspection and repair
of its interior.
(d) The maximum exposed surface temperature of all components in the fuel
tank must be less by a safe margin than the lowest expected autoignition
temperature of the fuel or fuel vapor in the tank. Compliance with this
requirement must be shown under all operating conditions and under all normal
or malfunction conditions of all components inside the tank.
(e) Each fuel tank installed in personnel compartments must be isolated by
fume-proof and fuel-proof enclosures that are drained and vented to the
exterior of the rotorcraft. The design and construction of the enclosures
must provide necessary protection for the tank, must be crash resistant
during a survivable impact in accordance with Sec. 29.952, and must be
adequate to withstand loads and abrasions to be expected in personnel
compartments.
SUMMARY: These amendments add comprehensive crash resistant fuel system
design and test criteria to the airworthiness standards for normal and
transport category rotorcraft. Application of these standards will minimize
fuel spillage near ignition sources and potential ignition sources and,
therefore, will improve the evacuation time needed for crew and passengers to
escape a post-crash fire (PCF). Implementation of these amendments will
minimize the PCF hazard saving lives and substantially reducing the severity
of physiological injuries sustained from PCF's in otherwise survivable
accidents.
(a) Each fuel tank must be able to withstand the applicable pressure tests
in this section without failure or leakage. If practicable, test pressures
may be applied in a manner simulating the pressure distribution in service.
(b) Each conventional metal tank, each nonmetallic tank with walls that are
not supported by the rotorcraft structure, and each integral tank must be
subjected to a pressure of 3.5 p.s.i. unless the pressure developed during
maximum limit acceleration or emergency deceleration with a full tank exceeds
this value, in which case a hydrostatic head, or equivalent test, must be
applied to duplicate the acceleration loads as far as possible. However, the
pressure need not exceed 3.5 p.s.i. on surfaces not exposed to the
acceleration loading.
(c) Each nonmetallic tank with walls supported by the rotorcraft structure
must be subjected to the following tests:
(1) A pressure test of at least 2.0 p.s.i. This test may be conducted on
the tank alone in conjunction with the test specified in paragraph (c)(2) of
this section.
(2) A pressure test, with the tank mounted in the rotorcraft structure,
equal to the load developed by the reaction of the contents, with the tank
full, during maximum limit acceleration or emergency deceleration. However,
the pressure need not exceed 2.0 p.s.i. on surfaces faces not exposed to the
acceleration loading.
(d) Each tank with large unsupported or unstiffened flat areas, or with
other features whose failure or deformation could cause leakage, must be
subjected to the following test or its equivalent:
(1) Each complete tank assembly and its supprots must be vibration tested
while mounted to simulate the actual installation.
(2) The tank assembly must be vibrated for 25 hours while two-thirds full
of any suitable fluid. The amplitude of vibration may not be less than one
thirty-second of an inch, unless otherwise substantiated.
(3) The test frequency of vibration must be as follows:
(i) If no frequency of vibration resulting from any r.p.m. within the
normal operating range of engine or rotor system speeds is critical, the test
frequency of vibration, in number of cycles per minute, must, unless a
frequency based on a more rational analysis is used, be the number obtained
by averaging the maximum and minimum power-on engine speeds (r.p.m.) for
reciprocating engine powered rotorcraft or 2,000 c.p.m. for turbine engine
powered rotorcraft.
(ii) If only one frequency of vibration resulting from any r.p.m. within
the normal operating range of engine or rotor system speeds is critical, that
frequency of vibration must be the test frequency.
(iii) If more than one frequency of vibration resulting from any r.p.m.
within the normal operating range of engine or rotor system speeds is
critical, the most critical of these frequencies must be the test frequency.
(4) Under paragraph (d)(3)(ii) and (iii), the time of test must be adjusted
to accomplish the same number of vibration cycles as would be accomplished in
25 hours at the frequency specified in paragraph (d)(3)(i) of this section.
(5) During the test, the tank assembly must be rocked at the rate of 16 to
20 complete cycles per minute through an angle of 15 degrees on both sides of
the horizontal (30 degrees total), about the most critical axis, for 25
hours. If motion about more than one axis is likely to be critical, the tank
must be rocked about each critical axis for 12 1/2 hours.
(Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 U.S.C. 1354(a), 1421, and
1423; sec. 6(c), 49 U.S.C. 1655 (c))
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-13, 42 FR
15046, Mar. 17, 1977]
Sec. 29.967 Fuel tank installation.
(a) Each fuel tank must be supported so that tank loads are not
concentrated on unsupported tank surfaces. In addition--
(1) There must be pads, if necessary, to prevent chafing between each tank
and its supports;
(2) The padding must be nonabsorbent or treated to prevent the absorption
of fuel;
(3) If flexible tank liners are used, they must be supported so that they
are not required to withstand fluid loads; and
(4) Each interior surface of tank compartments must be smooth and free of
projections that could cause wear of the liner, unless--
(i) There are means for protection of the liner at those points; or
(ii) The construction of the liner itself provides such protection.
(b) Any spaces adjacent to tank surfaces must be adequately ventilated to
avoid accumulation of fuel or fumes in those spaces due to minor leakage. If
the tank is in a sealed compartment, ventilation may be limited to drain
holes that prevent clogging and that prevent excessive pressure resulting
from altitude changes. If flexible tank liners are installed, the venting
arrangement for the spaces between the liner and its container must maintain
the proper relationship to tank vent pressures for any expected flight
condition.
(c) The location of each tank must meet the requirements of Sec. 29.1185(b)
and (c).
(d) No rotorcraft skin immediately adjacent to a major air outlet from the
engine compartment may act as the wall of an integral tank.
SUMMARY: These amendments add comprehensive crash resistant fuel system
design and test criteria to the airworthiness standards for normal and
transport category rotorcraft. Application of these standards will minimize
fuel spillage near ignition sources and potential ignition sources and,
therefore, will improve the evacuation time needed for crew and passengers to
escape a post-crash fire (PCF). Implementation of these amendments will
minimize the PCF hazard saving lives and substantially reducing the severity
of physiological injuries sustained from PCF's in otherwise survivable
accidents.
Each fuel tank or each group of fuel tanks with interconnected vent systems
must have an expansion space of not less than 2 percent of the combined tank
capacity. It must be impossible to fill the fuel tank expansion space
inadvertently with the rotorcraft in the normal ground attitude.
[Amdt. 29-26, 53 FR 34217, Sept. 2, 1988]
Sec. 29.971 Fuel tank sump.
(a) Each fuel tank must have a sump with a capacity of not less than the
greater of--
(1) 0.10 per cent of the tank capacity; or
(2) 1/16 gallon.
(b) The capacity prescribed in paragraph (a) of this section must be
effective with the rotorcraft in any normal attitude, and must be located so
that the sump contents cannot escape through the tank outlet opening.
(c) Each fuel tank must allow drainage of hazardous quantities of water
from each part of the tank to the sump with the rotorcraft in any ground
attitude to be expected in service.
(d) Each fuel tank sump must have a drain that allows complete drainage of
the sump on the ground.
(a) Each fuel tank filler connection must prevent the entrance of fuel into
any part of the rotorcraft other than the tank itself during normal
operations and must be crash resistant during a survivable impact in
accordance with Sec. 29.952(c). In addition--
(1) Each filler must be marked as prescribed in Sec. 29.1557(c)(1);
(2) Each recessed filler connection that can retain any appreciable
quantity of fuel must have a drain that discharges clear of the entire
rotorcraft; and
(3) Each filler cap must provide a fuel-tight seal under the fluid pressure
expected in normal operation and in a survivable impact.
(b) Each filler cap or filler cap cover must warn when the cap is not fully
locked or seated on the filler connection.
SUMMARY: These amendments add comprehensive crash resistant fuel system
design and test criteria to the airworthiness standards for normal and
transport category rotorcraft. Application of these standards will minimize
fuel spillage near ignition sources and potential ignition sources and,
therefore, will improve the evacuation time needed for crew and passengers to
escape a post-crash fire (PCF). Implementation of these amendments will
minimize the PCF hazard saving lives and substantially reducing the severity
of physiological injuries sustained from PCF's in otherwise survivable
accidents.
Sec. 29.975 Fuel tank vents and carburetor vapor vents.
(a) Fuel tank vents. Each fuel tank must be vented from the top part of the
expansion space so that venting is effective under normal flight conditions.
In addition--
(1) The vents must be arranged to avoid stoppage by dirt or ice formation;
(2) The vent arrangement must prevent siphoning of fuel during normal
operation;
(3) The venting capacity and vent pressure levels must maintain acceptable
differences of pressure between the interior and exterior of the tank,
during--
(i) Normal flight operation;
(ii) Maximum rate of ascent and descent; and
(iii) Refueling and defueling (where applicable);
(4) Airspaces of tanks with interconnected outlets must be interconnected;
(5) There may be no point in any vent line where moisture can accumulate
with the rotorcraft in the ground attitude or the level flight attitude,
unless drainage is provided;
(6) No vent or drainage provision may end at any point--
(i) Where the discharge of fuel from the vent outlet would constitute a
fire hazard; or
(ii) From which fumes could enter personnel compartments; and
(7) The venting system must be designed to minimize spillage of fuel
through the vents to an ignition source in the event of a rollover during
landing, ground operations, or a survivable impact, unless a rollover is
shown to be extremely remote.
(b) Carburetor vapor vents. Each carburetor with vapor elimination
connections must have a vent line to lead vapors back to one of the fuel
tanks. In addition--
(1) Each vent system must have means to avoid stoppage by ice; and
(2) If there is more than one fuel tank, and it is necessary to use the
tanks in a definite sequence, each vapor vent return line must lead back to
the fuel tank used for takeoff and landing.
SUMMARY: These amendments add comprehensive crash resistant fuel system
design and test criteria to the airworthiness standards for normal and
transport category rotorcraft. Application of these standards will minimize
fuel spillage near ignition sources and potential ignition sources and,
therefore, will improve the evacuation time needed for crew and passengers to
escape a post-crash fire (PCF). Implementation of these amendments will
minimize the PCF hazard saving lives and substantially reducing the severity
of physiological injuries sustained from PCF's in otherwise survivable
accidents.
(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 finger strainer must be accessible for inspection and cleaning.
[Amdt. 29-12, 41 FR 55473, Dec. 20, 1976]
Sec. 29.979 Pressure refueling and fueling provisions below fuel level.
(a) Each fueling connection below the fuel level in each tank must have
means to prevent the escape of hazardous quantities of fuel from that tank in
case of malfunction of the fuel entry valve.
(b) For systems intended for pressure refueling, a means in addition to the
normal means for limiting the tank content must be installed to prevent
damage to the tank in case of failure of the normal means.
(c) The rotorcraft pressure fueling system (not fuel tanks and fuel tank
vents) must withstand an ultimate load that is 2.0 times the load arising
from the maximum pressure, including surge, that is likely to occur during
fueling. The maximum surge pressure must be established with any combination
of tank valves being either intentionally or inadvertently closed.
(d) The rotorcraft defueling system (not including fuel tanks and fuel tank
vents) must withstand an ultimate load that is 2.0 times the load arising
from the maximum permissible defueling pressure (positive or negative) at the
rotorcraft fueling connection.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 FR
55473, Dec. 20, 1976]
Fuel System Components
Sec. 29.991 Fuel pumps.
(a) Compliance with Sec. 29.955 must not be jeopardized by failure of--
(1) Any one pump except pumps that are approved and installed as parts of a
type certificated engine; or
(2) Any component required for pump operation except the engine served by
that pump.
(b) The following fuel pump installation requirements apply:
(1) When necessary to maintain the proper fuel pressure--
(i) A connection must be provided to transmit the carburetor air intake
static pressure to the proper fuel pump relief valve connection; and
(ii) The gauge balance lines must be independently connected to the
carburetor inlet pressure to avoid incorrect fuel pressure readings.
(2) The installation of fuel pumps having seals or diaphragms that may leak
must have means for draining leaking fuel.
(3) Each drain line must discharge where it will not create a fire hazard.
[Amdt. 29-26, 53 FR 34217, Sept. 2, 1988]
Sec. 29.993 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, valve actuation, and
accelerated flight conditions.
(b) Each fuel line connected to components of the rotorcraft between which
relative motion could exist must have provisions for flexibility.
(c) Each flexible connection in fuel lines that may be under pressure or
subjected to axial loading must use flexible hose assemblies.
(d) Flexible hose must be approved.
(e) No flexible hose that might be adversely affected by high temperatures
may be used where excessive temperatures will exist during operation or after
engine shutdown.
Sec. 29.995 Fuel valves.
In addition to meeting the requirements of Sec. 29.1189, each fuel valve
must--
(a) [Reserved]
(b) Be supported so that no loads resulting from their operation or from
accelerated flight conditions are transmitted to the lines attached to the
valve.
(Secs. 313(a), 601, and 603, 72 Stat. 759, 775, 49 U.S.C. 1354(a), 1421, and
1423; sec. 6(c), 49 U.S.C. 1655 (c))
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-13, 42 FR
15046, Mar. 17, 1977]
Sec. 29.997 Fuel strainer or filter.
There must be a fuel strainer or filter between the fuel tank outlet and
the inlet of the first fuel system component which is susceptible to fuel
contamination, including but not limited to the fuel metering device or an
engine 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 inself,
unless adequate strengh margins under all loading conditions are provided in
the lines and connections; and
(d) Provide a means to remove from the fuel any contaminant which would
jeopardize the flow of fuel through rotorcraft or engine fuel system
components required for proper rotorcraft or engine fuel system operation.
(a) There must be at least one accessible drain at the lowest point in each
fuel system to completely drain the system with the rotorcraft in any ground
attitude to be expected in service.
(b) Each drain required by paragraph (a) of this section including the
drains prescribed in Sec. 29.971 must--
(1) Discharge clear of all parts of the rotorcraft;
(2) Have manual or automatic means to ensure positive closure in the off
position; and
(3) Have a drain valve--
(i) That is readily accessible and which can be easily opened and closed;
and
(ii) That is either located or protected to prevent fuel spillage in the
event of a landing with landing gear retracted.
If a fuel jettisoning system is installed, the following apply:
(a) Fuel jettisoning must be safe during all flight regimes for which
jettisoning is to be authorized.
(b) In showing compliance with paragraph (a) of this section, it must be
shown that--
(1) The fuel jettisoning system and its operation are free from fire
hazard;
(2) No hazard results from fuel or fuel vapors which impinge on any part of
the rotorcraft during fuel jettisoning; and
(3) Controllability of the rotorcraft remains satisfactory throughout the
fuel jettisoning operation.
(c) Means must be provided to automatically prevent jettisoning fuel below
the level required for an all-engine climb at maximum continuous power from
sea level to 5,000 feet altitude and cruise thereafter for 30 minutes at
maximum range engine power.
(d) The controls for any fuel jettisoning system must be designed to allow
flight personnel (minimum crew) to safely interrupt fuel jettisoning during
any part of the jettisoning operation.
(e) The fuel jettisoning system must be designed to comply with the
powerplant installation requirements of Sec. 29.901(c).
(f) An auxiliary fuel jettisoning system which meets the requirements of
paragraphs (a), (b), (d), and (e) of this section may be installed to
jettison additional fuel provided it has separate and independent controls.
[Amdt. 29-26, 53 FR 34218, Sept. 2, 1988]
Oil System
Sec. 29.1011 Engines: General.
(a) 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.
(b) The usable oil capacity of each system may not be less than the product
of the endurance of the rotorcraft under critical operating conditions and
the maximum allowable oil consumption of the engine under the same
conditions, plus a suitable margin to ensure adequate circulation and
cooling. Instead of a rational analysis of endurance and consumption, a
usable oil capacity of one gallon for each 40 gallons of usable fuel may be
used for reciprocating engine installations.
(c) Oil-fuel ratios lower than those prescribed in paragraph (c) of this
section may be used if they are substantiated by data on the oil consumption
of the engine.
(d) The ability of the engine and oil cooling provisions to maintain the
oil temperature at or below the maximum established value must be shown under
the applicable requirements of Secs. 29.1041 through 29.1049.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-26, 53 FR
34218, Sept. 2, 1988]
Sec. 29.1013 Oil tanks.
(a) Installation. Each oil tank installation must meet the requirements of
Sec. 29.967.
(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;
(2) Each reserve oil tank not directly connected to any engine has an
expansion space of not less than two percent of the tank capacity; and
(3) It is impossible to fill the expansion space inadvertently with the
rotorcraft in the normal ground attitude.
(c) Filler connections. Each recessed oil tank filler connection that can
retain any appreciable quantity of oil must have a drain that discharges
clear of the entire rotorcraft. In addition--
(1) Each oil tank filler cap must provide an oil-tight seal under the
pressure expected in operation;
(2) For category A rotorcraft, each oil tank filler cap or filler cap cover
must incorporate features that provide a warning when caps are not fully
locked or seated on the filler connection; and
(3) Each oil filler must be marked under Sec. 29.1557(c)(2).
(d) Vent. Oil tanks must be vented as follows:
(1) Each oil tank must be vented from the top part of the expansion space
so that venting is effective under all normal flight conditions.
(2) Oil tank vents must be arranged so that condensed water vapor that
might freeze and obstruct the line cannot accumulate at any point;
(e) Outlet. There must be 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. No oil tank outlet may be enclosed by a screen or guard
that would reduce the flow of oil below a safe value at any operating
temperature. 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 approved or shown
to be suitable for the particular installation.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-10, 39 FR
35462, Oct. 1, 1974]
Sec. 29.1015 Oil tank tests.
Each oil tank must be designed and installed so that--
(a) It can withstand, without failure, any vibration, inertia, and fluid
loads to which it may be subjected in operation; and
(b) It meets the requirements of Sec. 29.965, except that instead of the
pressure specified in Sec. 29.965(b)--
(1) 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;
and
(2) For all other tanks, the test pressure may not be less than 5 p.s.i.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-10, 39 FR
35462, Oct. 1, 1974]
Sec. 29.1017 Oil lines and fittings.
(a) Each oil line must meet the requirements of Sec. 29.993.
(b) Breather lines must be arranged so that--
(1) Condensed water vapor 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; and
(3) The breather does not discharge into the engine air induction system.
Sec. 29.1019 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 under Part 33 of this chapter.
(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. 29.1305(a)(18).
(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.
A drain (or drains) must be provided to allow safe drainage of the oil
system. Each drain must--
(a) Be accessible; and
(b) Have manual or automatic means for positive locking in the closed
position.
[Amdt. 29-22, 49 FR 6850, Feb. 23, 1984]
Sec. 29.1023 Oil radiators.
(a) Each oil radiator must be able to withstand any vibration, inertia, and
oil pressure loads to which it would be subjected in operation.
(b) Each oil radiator air duct must be located, or equipped, so that, in
case of fire, and with the airflow as it would be with and without the engine
operating, flames cannot directly strike the radiator.
Sec. 29.1025 Oil valves.
(a) Each oil shutoff must meet the requirements of Sec. 29.1189.
(b) The closing of oil shutoffs may not prevent autorotation.
(c) Each oil valve must have positive stops or suitable index provisions in
the "on" and "off" positions and must be supported so that no loads resulting
from its operation or from accelerated flight conditions are transmitted to
the lines attached to the valve.
Sec. 29.1027 Transmission and gearboxes: General.
(a) The oil system for components of the rotor drive system that require
continuous lubrication must be sufficiently independent of the lubrication
systems of the engine(s) to ensure--
(1) Operation with any engine inoperative; and
(2) Safe autorotation.
(b) Pressure lubrication systems for transmissions and gearboxes must
comply with the requirements of Secs. 29.1013, paragraphs (c), (d), and (f)
only, 29.1015, 29.1017, 29.1021, 29.1023, and 29.1337(d). In addition, the
system must have--
(1) An oil strainer or filter through which all the lubricant flows, and
must--
(i) Be designed to remove from the lubricant any contaminant which may
damage transmission and drive system components or impede the flow of
lubricant to a hazardous degree; and
(ii) Be equipped with a bypass constructed and installed so that--
(A) The lubricant will flow at the normal rate through the rest of the
system with the strainer or filter completely blocked; and
(B) The release of collected contaminants is minimized by appropriate
location of the bypass to ensure that collected contaminants are not in the
bypass flowpath;
(iii) Be equipped with a means to indicate collection of contaminants on
the filter or strainer at or before opening of the bypass;
(2) For each lubricant tank or sump outlet supplying lubrication to rotor
drive systems and rotor drive system components, a screen to prevent entrance
into the lubrication system of any object that might obstruct the flow of
lubricant from the outlet to the filter required by paragraph (b)(1) of this
section. The requirements of paragraph (b)(1) of this section do not apply to
screens installed at lubricant tank or sump outlets.
(c) Splash type lubrication systems for rotor drive system gearboxes must
comply with Secs. 29.1021 and 29.1337(d).
[Amdt. 29-26, 53 FR 34218, Sept. 2, 1988]
Cooling
Sec. 29.1041 General.
(a) The powerplant and auxiliary power unit cooling provisions must be able
to maintain the temperatures of powerplant components, engine fluids, and
auxiliary power unit components and fluids within the temperature limits
established for these components and fluids, under ground, water, and flight
operating conditions for which certification is requested, and after normal
engine or auxiliary power unit shutdown, or both.
(b) There must be cooling provisions to maintain the fluid temperatures in
any power transmission within safe values under any critical surface (ground
or water) and flight operating conditions.
(c) Except for ground-use-only auxiliary power units, compliance with
paragraphs (a) and (b) of this section must be shown by flight tests in which
the temperatures of selected powerplant component and auxiliary power unit
component, engine, and transmission fluids are obtained under the conditions
prescribed in those paragraphs.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-26, 53 FR
34218, Sept. 2, 1988]
Sec. 29.1043 Cooling tests.
(a) General. For the tests prescribed in Sec. 29.1041(c), the following
apply:
(1) If the tests are conducted under conditions deviating from the maximum
ambient atmospheric temperature 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 used 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) The test procedures must be as prescribed in Secs. 29.1045 through
29.1049.
(5) For the purposes of the cooling tests, a temperature is "stabilized"
when its rate of change is less than 2 deg.F per minute.
(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.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49
U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c) of the Dept. of
Transportation Act (49 U.S.C. 1655(c)))
(a) Climb cooling tests must be conducted under this section for--
(1) Category A rotorcraft; and
(2) Multiengine category B rotorcraft for which certification is requested
under the category A powerplant installation requirements, and under the
requirements of Sec. 29.861(a) at the steady rate of climb or descent
established under Sec. 29.67(b).
(b) The climb or descent cooling tests must be conducted with the engine
inoperative that produces the most adverse cooling conditions for the
remaining engines and powerplant components.
(c) Each operating engine must--
(1) For helicopters for which the use of 30-minute OEI power is requested,
be at 30-minute OEI power for 30 minutes, and then at maximum continuous
power (or at full throttle when above the critical altitude);
(2) For helicopters for which the use of continuous OEI power is requested,
be at continuous OEI power (or at full throttle when above the critical
altitude); and
(3) For other rotorcraft, be at maximum continuous power (or at full
throttle when above the critical altitude).
(d) After temperatures have stabilized in flight, the climb must be--
(1) Begun from an altitude not greater than the lower of--
(i) 1,000 feet below the engine critcal altitude; and
(ii) 1,000 feet below the maximum altitude at which the rate of climb is
150 f.p.m; and
(2) Continued for at least five minutes after the occurrence of the highest
temperature recorded, or until the rotorcraft reaches the maximum altitude
for which certification is requested.
(e) For category B rotorcraft without a positive rate of climb, the descent
must begin at the all-engine-critical altitude and end at the higher of--
(1) The maximum altitude at which level flight can be maintained with one
engine operative; and
(2) Sea level.
(f) The climb or descent must be conducted at an airspeed representing a
normal operational practice for the configuration being tested. However, if
the cooling provisions are sensitive to rotorcraft speed, the most critical
airspeed must be used, but need not exceed the speeds established under Sec.
29.67(a)(2) or Sec. 29.67(b). The climb cooling test may be conducted in
conjunction with the takeoff cooling test of Sec. 29.1047.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-26, 53 FR
34218, Sept. 2, 1988]
Sec. 29.1047 Takeoff cooling test procedures.
(a) Category A. For each category A rotorcraft, cooling must be shown
during takeoff and subsequent climb as follows:
(1) Each temperature must be stabilized while hovering in ground effect
with--
(i) The power necessary for hovering;
(ii) The appropriate cowl flap and shutter settings; and
(iii) The maximum weight.
(2) After the temperatures have stabilized, a climb must be started at the
lowest practicable altitude and must be conducted with one engine
inoperative.
(3) The operating engines must be at the greatest power for which approval
is sought (or at full throttle when above the critical altitude) for the same
period as this power is used in determining the takeoff climbout path under
Sec. 29.59.
(4) At the end of the time interval prescribed in paragraph (b)(3) of this
section, the power must be changed to that used in meeting Sec. 29.67(a)(2)
and the climb must be continued for--
(i) Thirty minutes, if 30-minute OEI power is used; or
(ii) At least 5 minutes after the occurrence of the highest temperature
recorded, if continuous OEI power or maximum continuous power is used.
(5) The speeds must be those used in determining the takeoff flight path
under Sec. 29.59.
(b) Category B. For each category B rotorcraft, cooling must be shown
during takeoff and subsequent climb as follows:
(1) Each temperature must be stabilized while hovering in ground effect
with--
(i) The power necessary for hovering;
(ii) The appropriate cowl flap and shutter settings; and
(iii) The maximum weight.
(2) After the temperatures have stabilized, a climb must be started at the
lowest practicable altitude with takeoff power.
(3) Takeoff power must be used for the same time interval as takeoff power
is used in determining the takeoff flight path under Sec. 29.63.
(4) At the end of the time interval prescribed in paragraph (a)(3) of this
section, the power must be reduced to maximum continuous power and the climb
must be continued for at least five minutes after the occurence of the
highest temperature recorded.
(5) The cooling test must be conducted at an airspeed corresponding to
normal operating practice for the configuration being tested. However, if the
cooling provisions are sensitive to rotorcraft speed, the most critical
airspeed must be used, but need not exceed the speed for best rate of climb
with maximum continuous power.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-1, 30 FR
8778, July 13, 1965; Amdt. 29-26, 53 FR 34219, Sept. 2, 1988]
Sec. 29.1049 Hovering cooling test procedures.
The hovering cooling provisions must be shown--
(a) At maximum weight or at the greatest weight at which the rotorcraft can
hover (if less), at sea level, with the power required to hover but not more
than maximum continuous power, in the ground effect in still air, until at
least five minutes after the occurrence of the highest temperature recorded;
and
(b) With maximum continuous power, maximum weight, and at the altitude
resulting in zero rate of climb for this configuration, until at least five
minutes after the occurrence of the highest temperature recorded.
Induction System
Sec. 29.1091 Air induction.
(a) The air induction system for each engine and auxiliary power unit must
supply the air required by that engine and auxiliary power unit under the
operating conditions for which certification is requested.
(b) Each engine and auxiliary power unit air induction system must provide
air for proper fuel metering and mixture distribution with the induction
system valves in any position.
(c) No air intake may open within the engine accessory section or within
other areas of any powerplant compartment where emergence of backfire flame
would constitute a fire hazard.
(d) Each reciprocating engine must have an alternate air source.
(e) Each alternate air intake must be located to prevent the entrance of
rain, ice, or other foreign matter.
(f) For turbine engine powered rotorcraft and rotorcraft incorporating
auxiliary power units--
(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 intake system; and
(2) The air inlet ducts must be located or protected so as to minimize the
ingestion of foreign matter during takeoff, landing, and taxiing.
(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C.
1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 U.S.C.
1655(c)))
(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., and with the engines at 60 percent of maximum
continuous power--
(1) Each rotorcraft with sea level engines using conventional venturi
carburetors has a preheater that can provide a heat rise of 90 deg. F.;
(2) Each rotorcraft with sea level engines using carburetors tending to
prevent icing has a preheater that can provide a heat rise of 70 deg. F.;
(3) Each rotorcraft with altitude engines using conventional venturi
carburetors has a preheater that can provide a heat rise of 120 deg. F.; and
(4) Each rotorcraft with altitude engines using carburetors tending to
prevent icing has a preheater that can provide a heat rise of 100 deg. F.
(b) Turbine engines. (1) It must be shown that each turbine engine and its
air inlet system can operate throughout the flight power range of the engine
(including idling)--
(i) Without accumulating ice on engine or inlet system components that
would adversely affect engine operation or cause a serious loss of power
under the icing conditions specified in Appendix C of this Part; and
(ii) In snow, both falling and blowing, without adverse effect on engine
operation, within the limitations established for the rotorcraft.
(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) Supercharged reciprocating engines. For each engine having a
supercharger to pressurize the air before it enters the carburetor, 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
altitude and operation condition because of supercharging.
(Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 U.S.C. 1354(a), 1421, and
1423; sec. 6(c), 49 U.S.C. 1655 (c))
Each carburetor air preheater must be designed and constructed to--
(a) Ensure ventilation of the preheater when the engine is operated in cold
air;
(b) Allow inspection of the exhaust manifold parts that it surrounds; and
(c) Allow inspection of critical parts of the preheater itself.
Sec. 29.1103 Induction systems ducts and air duct systems.
(a) Each induction system duct upstream of the first stage of the engine
supercharger and of the auxiliary power unit compressor must have a drain to
prevent the hazardous accumulation of fuel and moisture in the ground
attitude. No drain may discharge where it might cause a fire hazard.
(b) Each duct must be strong enough to prevent induction system failure
from normal backfire conditions.
(c) Each duct connected to components between which relative motion could
exist must have means for flexibility.
(d) Each duct within any fire zone for which a fire-extinguishing system is
required must be at least--
(1) Fireproof, if it passes through any firewall; or
(2) Fire resistant, for other ducts, except that ducts for auxiliary power
units must be fireproof within the auxiliary power unit fire zone.
(e) Each auxiliary power unit induction system duct must be fireproof for a
sufficient distance upstream of the auxiliary power unit compartment to
prevent hot gas reverse flow from burning through auxiliary power unit ducts
and entering any other compartment or area of the rotorcraft in which a
hazard would be created resulting from the entry of hot gases. The materials
used to form the remainder of the induction system duct and plenum chamber of
the auxiliary power unit must be capable of resisting the maximum heat
conditions likely to occur.
(f) Each auxiliary power unit induction system duct must be constructed of
materials that will not absorb or trap hazardous quantities of flammable
fluids that could be ignited in the event of a surge or reverse flow
condition.
(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C.
1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 U.S.C.
1655(c)))
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-17, 43 FR
50602, Oct. 30, 1978]
Sec. 29.1105 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 it 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. 29.1107 Inter-coolers and after-coolers.
Each inter-cooler and after-cooler must be able to withstand the vibration,
inertia, and air pressure loads to which it would be subjected in operation.
Sec. 29.1109 Carburetor air cooling.
It must be shown under Sec. 29.1043 that each installation using two-stage
superchargers has means to maintain the air temperature, at the carburetor
inlet, at or below the maximum established value.
Exhaust System
Sec. 29.1121 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 component upon which hot exhaust gases could impinge, or that
could be subjected to high temperatures from exhaust system parts, must be
fireproof. Each exhaust system component must be separated by a fireproof
shield from adjacent parts of the rotorcraft that are outside the engine and
auxiliary power unit compartments.
(d) No exhaust gases may discharge so as to cause a fire hazard with
respect to any flammable fluid vent or drain.
(e) No exhaust gases may discharge 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) Each exhaust shroud must be ventilated or insulated to avoid, during
normal operation, a temperature high enough to ignite any flammable fluids or
vapors outside the shroud.
(h) If significant traps exist, each turbine engine exhaust system must
have drains discharging clear of the rotorcraft, in any normal ground and
flight attitudes, to prevent fuel accumulation after the failure of an
attempted engine start.
(Secs. 313(a), 601, and 603, 72 Stat. 752, 755, 49 U.S.C. 1354(a), 1421, and
1423; sec. 6(c), 49 U.S.C. 1655 (c))
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 FR
970, Jan. 26, 1968; Amdt. 29-13, 42 FR 15046, Mar. 17, 1977]
Sec. 29.1123 Exhaust piping.
(a) Exhaust piping must be heat and corrosion resistant, and must have
provisions to prevent failure due to expansion by operating temperatures.
(b) Exhaust piping must be supported to withstand any vibration and inertia
loads to which it would be subjected in operation.
(c) Exhaust piping connected to components between which relative motion
could exist must have provisions for flexibility.
Sec. 29.1125 Exhaust heat exchangers.
For reciprocating engine powered rotorcraft the following apply:
(a) Each exhaust heat exchanger must be constructed and installed to
withstand the vibration, inertia, and other loads to which it would be
subjected in 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 inspecting the critical parts of each
exchanger;
(3) Each exchanger must have cooling provisions wherever it is subject to
contact with exhaust gases; and
(4) Each exhaust heat exchanger muff may have stagnant areas or liquid
traps that would increase the probability of ignition of flammable fluids or
vapors that might be present in case of the failure or malfunction of
components carrying flammable fluids.
(b) If an exhaust heat exchanger is used for heating ventilating air used
by personnel--
(1) There must be a secondary heat exchanger between the primary exhaust
gas heat exchanger and the ventilating air system; or
(2) Other means must be used to prevent harmful contamination of the
ventilating air.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 FR
55473, Dec. 20, 1976]
Powerplant Controls and Accessories
Sec. 29.1141 Powerplant controls: general.
(a) Powerplant controls must be located and arranged under Sec. 29.777 and
marked under Sec. 29.1555.
(b) Each control must be located so that it cannot be inadvertently
operated by persons entering, leaving, or moving normally in the cockpit.
(c) Each flexible powerplant control must be approved.
(d) Each control must be able to maintain any set position without--
(1) Constant attention; or
(2) Tendency to creep due to control loads or vibration.
(e) Each control must be able to withstand operating loads without
excessive deflection.
(f) Controls of powerplant valves required for safety 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.
(Secs. 313(a), 601, and 603, 72 Stat. 752, 775, 49 U.S.C. 1354(a), 1421, and
1423; sec. 6(c), 49 U.S.C. 1655(c))
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-13, 42 FR
15046, Mar. 17, 1977; Amdt. 29-26, 53 FR 34219, Sept. 2, 1988]
Sec. 29.1142 Auxiliary power unit controls.
Means must be provided on the flight deck for starting, stopping, and
emergency shutdown of each installed auxiliary power unit.
(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C.
1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 U.S.C.
1655(c)))
[Amdt. 29-17, 43 FR 50602, Oct. 30, 1978]
Sec. 29.1143 Engine controls.
(a) There must be a separate power control for each engine.
(b) Power controls must be arranged to allow ready synchronization of all
engines by--
(1) Separate control of each engine; and
(2) Simultaneous control of all engines.
(c) Each power control must provide a positive and immediately responsive
means of controlling its engine.
(d) Each fluid injection control other than fuel system control must be in
the corresponding power control. However, the injection system pump may have
a separate control.
(e) If a power 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.
(f) For rotorcraft to be certificated for a 30-second OEI power rating, a
means must be provided to automatically activate and control the 30-second
OEI power and prevent any engine from exceeding the installed engine limits
associated with the 30-second OEI power rating approved for the rotorcraft.
[Amdt. 29-26, 53 FR 34219, Sept. 2, 1988, as amended by Amdt. 29-34, 59 FR
47768, Sept. 16, 1994]
SUMMARY: This rule adopts new and revised airworthiness standards by
incorporating optional one-engine-inoperative (OEI) power ratings for
multiengine, turbine-powered rotorcraft. These amendments result from a
petition for rulemaking from Aerospace Industries Association of America
(AIA) and the recognition by both government and industry that additional OEI
power rating standards are needed. These amendments enhance rotorcraft safety
after an engine failure or precautionary shutdown by providing higher OEI
power, when necessary. These amendments also assure that the drive system
will maintain its structural integrity and allow continued safe flight while
operating at the new OEI power ratings with the operable engine(s).
(a) Ignition switches must control each ignition circuit on each engine.
(b) There must be means to quickly shut off all ignition 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.
(Secs. 313(a), 601, and 603, 72 Stat. 759, 775, 49 U.S.C. 1354(a), 1421, and
1423; sec. 6(c), 49 U.S.C. 1655 (c))
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-13, 42 FR
15046, Mar. 17, 1977]
Sec. 29.1147 Mixture controls.
(a) If there are mixture controls, each engine must have a separate
control, and the controls must be arranged to allow--
(1) Separate control of each engine; and
(2) Simultaneous control of all engines.
(b) Each intermediate position of the mixture controls that corresponds to
a normal operating setting must be identifiable by feel and sight.
Sec. 29.1151 Rotor brake controls.
(a) It must be impossible to apply the rotor brake inadvertently in flight.
(b) There must be means to warn the crew if the rotor brake has not been
completely released before takeoff.
Sec. 29.1157 Carburetor air temperature controls.
There must be a separate carburetor air temperature control for each
engine.
Sec. 29.1159 Supercharger controls.
Each supercharger control must be accessible to--
(a) The pilots; or
(b) (If there is a separate flight engineer station with a control panel)
the flight engineer.
Sec. 29.1163 Powerplant accessories.
(a) Each engine mounted accessory must--
(1) Be approved for mounting on the engine involved;
(2) Use the provisions on the engine for mounting; and
(3) Be sealed in such a way as 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 igniting flammable fluids or vapors.
(c) If continued rotation of an engine-driven cabin supercharger or any
remote accessory driven by the engine will be a hazard if they malfunction,
there must be means to prevent their hazardous rotation without interfering
with the continued operation of the engine.
(d) Unless other means are provided, torque limiting means must be provided
for accessory drives located on any component of the transmission and rotor
drive system to prevent damage to these components from excessive accessory
load.
(a) Each battery ignition system must be supplemented with 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) Magneto ground wiring (for separate ignition circuits) that lies on the
engine side of any firewall must be installed, located, or protected, to
minimize the probability of the simultaneous failure of two or more wires as
a result of mechanical damage, electrical fault, or other cause.
(e) No ground wire for any engine may be routed through a fire zone of
another engine unless each part of that wire within that zone is fireproof.
(f) Each ignition system must be independent of any electrical circuit that
is not used for assisting, controlling, or analyzing the operation of that
system.
(g) There must be means to warn appropriate crewmembers if the
malfunctioning of any part of the electrical system is causing the continuous
discharge of any battery necessary for engine ignition.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 FR
55473, Dec. 20, 1976]
Powerplant Fire Protection
Sec. 29.1181 Designated fire zones: regions included.
(a) Designated fire zones are--
(1) The engine power section of reciprocating engines;
(2) The engine accessory section of reciprocating engines;
(3) Any complete powerplant compartment in which there is no isolation
between the engine power section and the engine accessory section, for
reciprocating engines;
(4) Any auxiliary power unit compartment;
(5) Any fuel-burning heater and other combustion equipment installation
described in Sec. 29.859;
(6) The compressor and accessory sections of turbine engines; and
(7) The combustor, turbine, and tailpipe sections of turbine engine
installations except sections that do not contain lines and components
carrying flammable fluids or gases and are isolated from the designated fire
zone prescribed in paragraph (a)(6) of this section by a firewall that meets
Sec. 29.1191.
(b) Each designated fire zone must meet the requirements of Secs. 29.1183
through 29.1203.
[Amdt. 29-3, 33 FR 970, Jan. 26, 1968, as amended by Amdt. 29-26, 53 FR
34219, Sept. 2, 1988]
Sec. 29.1183 Lines, fittings, and components.
(a) Except as provided in paragraph (b) of this section, each line,
fitting, and other component carrying flammable fluid in any area subject to
engine fire conditions and each component which conveys or contains flammable
fluid in a designated fire zone must be fire resistant, except that flammable
fluid tanks and supports in a designated fire zone 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. 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.
(a) No tank or reservoir that is part of a system containing flammable
fluids or gases may be in a designated fire zone unless the fluid contained,
the design of the system, the materials used in the tank and its supports,
the shutoff means, and the connections, lines, and controls provide a degree
of safety equal to that which would exist if the tank or reservoir were
outside such a zone.
(b) Each fuel tank must be isolated from the engines by a firewall or
shroud.
(c) There must be at least one-half inch of clear airspace between each
tank or reservoir and each firewall or shroud isolating a designated fire
zone, unless equivalent means are used to prevent heat transfer from the fire
zone to the flammable fluid.
(d) Absorbent material close to flammable fluid system components that
might leak must be covered or treated to prevent the absorption of hazardous
quantities of fluids.
Sec. 29.1187 Drainage and ventilation of fire zones.
(a) There must be complete drainage of each part of each designated fire
zone to minimize the hazards resulting from failure or malfunction of any
component containing flammable fluids. The drainage means must be--
(1) Effective under conditions expected to prevail when drainage is needed;
and
(2) Arranged so that no discharged fluid will cause an additional fire
hazard.
(b) Each designated fire zone must be ventilated to prevent the
accumulation of flammable vapors.
(c) No ventilation opening may be where it would allow the entry of
flammable fluids, vapors, or flame from other zones.
(d) Ventilation means must be arranged so that no discharged vapors will
cause an additional fire hazard.
(e) For category A rotorcraft, there must be means to allow the crew to
shut off the sources of forced ventilation in any fire zone (other than the
engine power section of the powerplant compartment) unless the amount of
extinguishing agent and the rate of discharge are based on the maximum
airflow through that zone.
Sec. 29.1189 Shutoff means.
(a) There must be means to shut off or otherwise prevent hazardous
quantities of fuel, oil, de-icing fluid, and other flammable fluids from
flowing into, within, or through any designated fire zone, except that this
means need not be provided--
(1) For lines, fittings, and components forming an integral part of an
engine;
(2) For oil systems for turbine engine installations in which all
components of the system, including oil tanks, are fireproof or located in
areas not subject to engine fire conditions; or
(3) For engine oil systems in category B rotorcraft using reciprocating
engines of less than 500 cubic inches displacement.
(b) The closing of any fuel shutoff valve for any engine may not make fuel
unavailable to the remaining engines.
(c) For category A rotorcraft, no hazardous quantity of flammable fluid may
drain into any designated fire zone after shutoff has been accomplished, nor
may the closing of any fuel shutoff valve for an engine make fuel unavailable
to the remaining engines.
(d) The operation of any shutoff may not interfere with the later emergency
operation of any other equipment, such as the means for declutching the
engine from the rotor drive.
(e) Each shutoff valve and its control must be designed, located, and
protected to function properly under any condition likely to result from fire
in a designated fire zone.
(f) Except for ground-use-only auxiliary power unit installations, there
must be means to prevent inadvertent operation of each shutoff and to make it
possible to reopen it in flight after it has been closed.
(a) Each engine, including the combustor, turbine, and tailpipe sections of
turbine engine installations, must be isolated by a firewall, shroud, or
equivalent means, from personnel compartments, structures, controls, rotor
mechanisms, and other parts that are--
(1) Essential to controlled flight and landing; and
(2) Not protected under Sec. 29.861.
(b) Each auxiliary power unit, combustion heater, and other combustion
equipment to be used in flight, must be isolated from the rest of the
rotorcraft by firewalls, shrouds, or equivalent means.
(c) Each firewall or shroud must be constructed so that no hazardous
quantity of air, fluid, or flame can pass from any engine compartment to
other parts of the rotorcraft.
(d) Each opening in the firewall or shroud must be sealed with close-
fitting fireproof grommets, bushings, or firewall fittings.
(e) Each firewall and shroud must be fireproof and protected against
corrosion.
(f) In meeting this section, account must be taken of the probable path of
a fire as affected by the airflow in normal flight and in autorotation.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 FR
970, Jan. 26, 1968]
Sec. 29.1193 Cowling and engine compartment covering.
(a) Each cowling and engine compartment covering must be constructed and
supported so that it can resist the vibration, inertia, and air loads to
which it may be subjected in operation.
(b) Cowling must meet the drainage and ventilation requirements of Sec.
29.1187.
(c) On rotorcraft with a diaphragm isolating the engine power section from
the engine accessory section, each part of the accessory section cowling
subject to flame in case of fire in the engine power section of the
powerplant must--
(1) Be fireproof; and
(2) Meet the requirements of Sec. 29.1191.
(d) Each part of the cowling or engine compartment covering subject to high
temperatures due to its nearness to exhaust system parts or exhaust gas
impingement must be fireproof.
(e) Each rotorcraft must--
(1) Be designated and constructed so that no fire originating in any fire
zone can enter, either through openings or by burning through external skin,
any other zone or region where it would create additional hazards;
(2) Meet the requirements of paragraph (e)(1) of this section with the
landing gear retracted (if applicable); and
(3) Have fireproof skin in areas subject to flame if a fire starts in or
burns out of any designated fire zone.
(f) A means of retention for each openable or readily removable panel,
cowling, or engine or rotor drive system covering must be provided to
preclude hazardous damage to rotors or critical control components in the
event of--
(1) Structural or mechanical failure of the normal retention means, unless
such failure is extremely improbable; or
(2) Fire in a fire zone, if such fire could adversely affect the normal
means of retention.
(Secs. 313(a), 601, and 603, 72 Stat. 759, 775, 49 U.S.C. 1354(a), 1421, and
1423; sec. 6(c), 49 U.S.C. 1655(c))
All surfaces aft of, and near, engine compartments and designated fire
zones, other than tail surfaces not subject to heat, flames, or sparks
emanating from a designated fire zone or engine compartment, must be at least
fire resistant.
[Amdt. 29-3, 33 FR 970, Jan. 26, 1968]
Sec. 29.1195 Fire extinguishing systems.
(a) Each turbine engine powered rotorcraft and Category A reciprocating
engine powered rotorcraft, and each Category B reciprocating engine powered
rotorcraft with engines of more than 1,500 cubic inches must have a fire
extinguishing system for the designated fire zones. The fire extinguishing
system for a powerplant must be able to simultaneously protect all zones of
the powerplant compartment for which protection is provided.
(b) For multiengine powered rotorcraft, the fire extinguishing system, the
quantity of extinguishing agent, and the rate of discharge must--
(1) For each auxiliary power unit and combustion equipment, provide at
least one adequate discharge; and
(2) For each other designated fire zone, provide two adequate discharges.
(c) For single engine rotorcraft, the quantity of extinguishing agent and
the rate of discharge must provide at least one adequate discharge for the
engine compartment.
(d) It must be shown by either actual or simulated flight tests that under
critical airflow conditions in flight the discharge of the extinguishing
agent in each designated fire zone will provide an agent concentration
capable of extinguishing fires in that zone and of minimizing the probability
of reignition.
(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C.
1354(a), 1421, 1423, 1424), sec. 6(c), Dept. of Transportation Act (49 U.S.C.
1655(c)))
(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, it must be shown by test that
entry of harmful concentrations of fluid or fluid vapors into any personnel
compartment (due to leakage during normal operation of the rotorcraft, or
discharge on the ground or in flight) is prevented, even though a defect may
exist in the extinguishing system.
(Secs. 313(a), 601, and 603, 72 Stat. 759, 775, 49 U.S.C. 1354(a), 1421, and
1423; sec. 6(c), 49 U.S.C. 1655(c))
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 FR
55473, Dec. 20, 1976; Amdt. 29-13, 42 FR 15047, Mar. 17, 1977]
Sec. 29.1199 Extinguishing agent containers.
(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 rotorcraft. The line must also be located or protected
to prevent clogging caused by ice or other foreign matter.
(c) There must be a means 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.
(Secs. 313(a), 601, and 603, 72 Stat. 759, 775, 49 U.S.C. 1354(a), 1421, and
1423; sec. 6(c), 49 U.S.C. 1655 (c))
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-13, 42 FR
15047, Mar. 17, 1977]
Sec. 29.1201 Fire extinguishing system materials.
(a) No materials 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.
Sec. 29.1203 Fire detector systems.
(a) For each turbine engine powered rotorcraft and Category A reciprocating
engine powered rotorcraft, and for each Category B reciprocating engine
powered rotorcraft with engines of more than 900 cubic inches displacement,
there must be approved, quick-acting fire detectors in designated fire zones
and in the combustor, turbine, and tailpipe sections of turbine installations
(whether or not such sections are designated fire zones) in numbers and
locations ensuring prompt detection of fire in those zones.
(b) Each fire detector must be constructed and installed to withstand any
vibration, inertia, and other loads to which it would 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 crewmembers to check, in flight, the
functioning of each fire detector system electrical circuit.
(e) The writing and other components of each fire detector system in an
engine compartment must be at least fire resistant.
(f) No fire detector system component for any fire zone may pass through
another fire zone, unless--
(1) It is protected against the possibility of false warnings resulting
from fires in zones through which it passes; or
(2) The zones involved are simultaneously protected by the same detector
and extinguishing systems.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 FR
970, Jan. 26, 1968]
Subpart F--Equipment
General
Sec. 29.1301 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.
Sec. 29.1303 Flight and navigation instruments.
The following are required flight and navigational instruments:
(a) An airspeed indicator. For Category A rotorcraft with VNE less than a
speed at which unmistakable pilot cues provide overspeed warning, a maximum
allowable airspeed indicator must be provided. If maximum allowable airspeed
varies with weight, altitude, temperature, or r.p.m., the indicator must show
that variation.
(b) A sensitive altimeter.
(c) A magnetic direction indicator.
(d) A clock displaying hours, minutes, and seconds with a sweep-second
pointer or digital presentation.
(e) A free-air temperature indicator.
(f) A non-tumbling gyroscopic bank and pitch indicator.
(g) A gyroscopic rate-of-turn indicator combined with an integral slip-skid
indicator (turn-and-bank indicator) except that only a slip-skid indicator is
required on rotorcraft with a third altitude instrument system that--
(1) Is useable through flight altitudes of +/- 80 degrees of pitch and +/-
120 degrees of roll;
(2) Is powered from a source independent of the electrical generating
system;
(3) Continues reliable operation for a minimum of 30 minutes after total
failure of the electrical generating system;
(4) Operates independently of any other altitude indicating system;
(5) Is operative without selection after total failure of the electrical
generating system;
(6) Is located on the instrument panel in a position acceptable to the
Administrator that will make it plainly visible to and useable by any pilot
at his station; and
(7) Is appropriately lighted during all phases of operation.
(h) A gyroscopic direction indicator.
(i) A rate-of-climb (vertical speed) indicator.
(j) For Category A rotorcraft, a speed warning device when VNE is less than
the speed at which unmistakable overspeed warning is provided by other pilot
cues. The speed warning device must give effective aural warning (differing
distinctively from aural warnings used for other purposes) to the pilots
whenever the indicated speed exceeds VNE plus 3 knots and must operate
satisfactorily throughout the approved range of altitudes and temperatures.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49
U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. of
Transportation Act (49 U.S.C. 1655(c)))
The following are required powerplant instruments:
(a) For each rotorcraft--
(1) A carburetor air temperature indicator for each reciprocating engine;
(2) A cylinder head temperature indicator for each air-cooled reciprocating
engine, and a coolant temperature indicator for each liquid-cooled
reciprocating engine;
(3) A fuel quantity indicator for each fuel tank;
(4) A low fuel warning device for each fuel tank which feeds an engine.
This device must--
(i) Provide a warning to the crew when approximately 10 minutes of usable
fuel remains in the tank; and
(ii) Be independent of the normal fuel quantity indicating system.
(5) A manifold pressure indicator, for each reciprocating engine of the
altitude type;
(6) An oil pressure warning device for each pressure-lubricated gearbox to
indicate when the oil pressure falls below a safe value;
(7) An oil quantity indicator for each oil tank and each rotor drive
gearbox, if lubricant is self-contained;
(8) An oil temperature indicator for each engine;
(9) An oil temperature warning device to indicate unsafe oil temperatures
in each main rotor drive gearbox, including gearboxes necessary for rotor
phasing;
(10) A gas temperature indicator for each turbine engine;
(11) A gas producer rotor tachometer for each turbine engine;
(12) A tachometer for each engine that, if combined with the applicable
instrument required by paragraph (a)(13) of this section, indicates rotor
r.p.m. during autorotation.
(13) At least one tachometer to indicate, as applicable--
(i) The r.p.m. of the single main rotor;
(ii) The common r.p.m. of any main rotors whose speeds cannot vary
appreciably with respect to each other; and
(iii) The r.p.m. of each main rotor whose speed can vary appreciably with
respect to that of another main rotor;
(14) A free power turbine tachometer for each turbine engine;
(15) A means, for each turbine engine, to indicate power for that engine;
(16) For each turbine engine, an indicator to indicate the functioning of
the powerplant ice protection system;
(17) An indicator for the filter required by Sec. 29.997 to indicate the
occurrence of contamination of the filter to the degree established in
compliance with Sec. 29.955;
(18) For each turbine engine, a warning means for the oil strainer or
filter required by Sec. 29.1019, if it has no bypass, to warn the pilot of
the occurrence of contamination of the strainer or filter before it reaches
the capacity established in accordance with Sec. 29.1019(a)(2);
(19) An indicator to indicate the functioning of any selectable or
controllable heater used to prevent ice clogging of fuel system components;
(20) An individual fuel pressure indicator for each engine, unless the fuel
system which supplies that engine does not employ any pumps, filters, or
other components subject to degradation or failure which may adversely affect
fuel pressure at the engine;
(21) A means to indicate to the flightcrew the failure of any fuel pump
installed to show compliance with Sec. 29.955;
(22) Warning or caution devices to signal to the flightcrew when
ferromagnetic particles are detected by the chip detector required by Sec.
29.1337(e); and
(23) For auxiliary power units, an individual indicator, warning or caution
device, or other means to advise the flightcrew that limits are being
exceeded, if exceeding these limits can be hazardous, for--
(i) Gas temperature;
(ii) Oil pressure; and
(iii) Rotor speed.
(b) For category A rotorcraft--
(1) An individual oil pressure indicator for each engine, and either an
independent warning device for each engine or a master warning device for the
engines with means for isolating the individual warning circuit from the
master warning device;
(2) An independent fuel pressure warning device for each engine or a master
warning device for all engines with provision for isolating the individual
warning device from the master warning device; and
(3) Fire warning indicators.
(c) For category B rotorcraft--
(1) An individual oil pressure indicator for each engine; and
(2) Fire warning indicators, when fire detection is required.
The following is required miscellaneous equipment:
(a) An approved seat for each occupant.
(b) A master switch arrangement for electrical circuits other than
ignition.
(c) Hand fire extinguishers.
(d) A windshield wiper or equivalent device for each pilot station.
(e) A two-way radio communication system.
[Amdt. 29-12, 41 FR 55473, Dec. 20, 1976]
Sec. 29.1309 Equipment, systems, and installations.
(a) The equipment, systems, and installations whose functioning is required
by this subchapter must be designed and installed to ensure that they perform
their intended functions under any foreseeable operating condition.
(b) The rotorcraft systems and associated components, considered separately
and in relation to other systems, must be designed so that--
(1) For Category B rotorcraft, the equipment, systems, and installations
must be designed to prevent hazards to the rotorcraft if they malfunction or
fail; or
(2) For Category A rotorcraft--
(i) The occurrence of any failure condition which would prevent the
continued safe flight and landing of the rotorcraft is extremely improbable;
and
(ii) The occurrence of any other failure conditions which would reduce the
capability of the rotorcraft or the ability of the crew to cope with adverse
operating conditions is improbable.
(c) 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 which could create additional hazards.
(d) Compliance with the requirements of paragraph (b)(2) of this section
must be shown by analysis and, where necessary, by appropriate ground,
flight, or simulator tests. The analysis must consider--
(1) Possible modes of failure, including malfunctions and damage from
external sources;
(2) The probability of multiple failures and undetected failures;
(3) The resulting effects on the rotorcraft and occupants, considering the
stage of flight and operating conditions; and
(4) The crew warning cues, corrective action required, and the capability
of detecting faults.
(e) For Category A rotorcraft, each installation whose functioning is
required by this subchapter and which 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 system with the system functioning normally.
(2) Essential loads, after failure of any one prime mover, power converter,
or energy storage device.
(3) Essential loads, after failure of--
(i) Any one engine, on rotorcraft with two engines; and
(ii) Any two engines, on rotorcraft with three or more engines.
(f) In determining compliance with paragraphs (e) (2) and (3) 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 for controlled flight need not be considered for the two-engine-
inoperative condition on rotorcraft with three or more engines.
(g) In showing compliance with paragraphs (a) and (b) of this section with
regard to the electrical system and to equipment design and installation,
critical environmental conditions must be considered. For electrical
generation, distribution, and utilization equipment required by or used in
complying with this subchapter, except equipment covered by Technical
Standard Orders containing environmental test procedures, the ability to
provide continuous, safe service under foreseeable environmental conditions
may be shown by environmental tests, design analysis, or reference to
previous comparable service experience on other aircraft.
(h) In showing compliance with paragraphs (a) and (b) of this section, the
effects of lightning strikes on the rotorcraft must be considered in
accordance with Sec. 29.610.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49
U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. of
Transportation Act (49 U.S.C. 1655(c)))
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-14, 42 FR
36972, July 18, 1977; Amdt. 29-24, 49 FR 44438, Nov. 6, 1984]
Instruments: Installation
Sec. 29.1321 Arrangement and visibility.
(a) Each flight, navigation, and powerplant instrument for use by any pilot
must be easily 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) Each instrument necessary for safe operation, including the airspeed
indicator, gyroscopic direction indicator, gyroscopic bank-and-pitch
indicator, slip-skid indicator, altimeter, rate-of-climb indicator, rotor
tachometers, and the indicator most representative of engine power, must be
grouped and centered as nearly as practicable about the vertical plane of the
pilot's forward vision. In addition, for rotorcraft approved for IFR flight--
(1) The instrument that most effectively indicates attitude must be on the
panel in the top center position;
(2) The instrument that most effectively indicates direction of flight must
be adjacent to and directly below the attitude instrument;
(3) The instrument that most effectively indicates airspeed must be
adjacent to and to the left of the attitude instrument; and
(4) The instrument that most effectively indicates altitude or is most
frequently utilized in control of altitude must be adjacent to and to the
right of the attitude instrument.
(c) Other required powerplant instruments must be closely grouped on the
instrument panel.
(d) Identical powerplant instruments for the engines must be located so as
to prevent any confusion as to which engine each instrument relates.
(e) Each powerplant instrument vital to safe operation must be plainly
visible to appropriate crewmembers.
(f) Instrument panel vibration may not damage, or impair the readability or
accuracy of, any instrument.
(g) If a visual indicator is provided to indicate malfunction of an
instrument, it must be effective under all probable cockpit lighting
conditions.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49
U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. of
Transportation Act (49 U.S.C. 1655(c)))
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-14, 42 FR
36972, July 18, 1977; Amdt. 29-21, 48 FR 4391, Jan. 31, 1983]
Sec. 29.1322 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.
[Amdt. 29-12, 41 FR 55474, Dec. 20, 1976]
Sec. 29.1323 Airspeed indicating system.
For each airspeed indicating system, the following apply:
(a) Each airspeed indicating instrument 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 system must be calibrated to determine system error excluding
airspeed instrument error. This calibration must be determined--
(1) In level flight at speeds of 20 knots and greater, and over an
appropriate range of speeds for flight conditions of climb and autorotation;
and
(2) During takeoff, with repeatable and readable indications that ensure--
(i) Consistent realization of the field lengths specified in the Rotorcraft
Flight Manual; and
(ii) Avoidance of the critical areas of the limiting height-speed envelope
established under Sec. 29.79.
(c) For Category A rotorcraft--
(1) The indication must allow consistent definition of the critical
decision point; and
(2) The system error, excluding the airspeed instrument calibration error,
may not exceed--
(i) Three percent or 5 knots, whichever is greater, in level flight at
speeds above 80 percent of takeoff safety speed; and
(ii) Ten knots in climb at speeds from 10 knots below takeoff safety speed
to 10 knots above VY.
(d) For Category B rotorcraft, the system error, excluding the airspeed
instrument calibration error, may not exceed 3 percent or 5 knots, whichever
is greater, in level flight at speeds above 80 percent of the climbout speed
attained at 50 feet when complying with Sec. 29.63.
(e) Each system must be arranged, so far as practicable, to prevent
malfunction or serious error due to the entry of moisture, dirt, or other
substances.
(f) Each system must have a heated pitot tube or an equivalent means of
preventing malfunction due to icing.
Sec. 29.1325 Static pressure and pressure altimeter systems.
(a) Each instrument with static air case connections must be vented to the
outside atmosphere through an appropriate piping system.
(b) Each vent must be located where its orifices are least affected by
airflow variation, moisture, or foreign matter.
(c) Each static pressure port must be designed and located in such manner
that the correlation between air pressure in the static pressure system and
true ambient atmospheric static pressure is not altered when the rotorcraft
encounters icing conditions. An anti-icing 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 altimeter when on the primary static system by
more than 50 feet, a correction card must be provided for the alternate
static system.
(d) Except for the vent into the atmosphere, each system must be airtight.
(e) Each pressure altimeter must be approved and calibrated to indicate
pressure altitude in a standard atmosphere with a minimum practicable
calibration error when the corresponding static pressures are applied.
(f) Each system must be designed and installed so that an 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. However, the error need not be less than
+/-30 feet.
(g) Except as provided in paragraph (h) 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.
(h) For unpressurized rotorcraft, paragraph (g)(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.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49
U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. of
Transportation Act (49 U.S.C. 1655(c)))
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-14, 42 FR
36972, July 18, 1977; Amdt. 29-24, 49 FR 44439, Nov. 6, 1984]
Sec. 29.1327 Magnetic direction indicator.
(a) Each magnetic direction indicator must be installed so that its
accuracy is not excessively affected by the rotorcraft's vibration or
magnetic fields.
(b) The compensated installation may not have a deviation, in level flight,
greater than 10 degrees on any heading.
Sec. 29.1329 Automatic pilot system.
(a) Each automatic pilot system must be designed so that the automatic
pilot can--
(1) Be sufficiently overpowered by one pilot to allow control of the
rotorcraft; and
(2) Be readily and positively disengaged by each pilot to prevent it from
interfering with the control of the rotorcraft.
(b) 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.
(c) Each manually operated control for the system's operation must be
readily accessible to the pilots.
(d) The system must be designed and adjusted so that, within the range of
adjustment available to the pilot, it cannot produce hazardous loads on the
rotorcraft, 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.
(e) If the automatic pilot integrates signals from auxiliary controls or
furnishes signals for operation of other equipment, there must be positive
interlocks and sequencing of engagement to prevent improper operation.
For category A rotorcraft--
(a) Each required flight instrument using a power supply must have--
(1) Two independent sources of power;
(2) A means of selecting either power source; and
(3) A visual means integral with each instrument to indicate when the power
adequate to sustain proper instrument performance is not being supplied. The
power must be measured at or near the point where it enters the instrument.
For electrical instruments, the power is considered to be adequate when the
voltage is within the approved limits; and
(b) The installation and power supply system must be such that failure of
any flight instrument connected to one source, or of the energy supply from
one source, or a fault in any part of the power distribution system does not
interfere with the proper supply of energy from any other source.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-24, 49 FR
44439, Nov. 6, 1984]
Sec. 29.1333 Instrument systems.
For systems that operate the required flight instruments which are located
at each pilot's station, the following apply:
(a) Only the required flight instruments for the first pilot may be
connected to that operating system.
(b) The equipment, systems, and installations must be designed so that one
display of the information essential to the safety of flight which is
provided by the flight instruments remains available to a pilot, without
additional crewmember action, after any single failure or combination of
failures that are not shown to be extremely improbable.
(c) Additional instruments, systems, or equipment may not be connected to
the operating system for a second pilot unless provisions are made to ensure
the continued normal functioning of the required flight instruments in the
event of any malfunction of the additional instruments, systems, or equipment
which is not shown to be extremely improbable.
[Amdt. 29-24, 49 FR 44439, Nov. 6, 1984]
Sec. 29.1335 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.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49
U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. of
Transportation Act (49 U.S.C. 1655(c)))
[Amdt. 29-14, 42 FR 36973, July 18, 1977]
Sec. 29.1337 Powerplant instruments.
(a) Instruments and instrument lines. (1) Each powerplant and auxiliary
power unit instrument line must meet the requirements of Secs. 29.993 and
29.1183.
(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 means to indicate to the flight
crew members the quantity, in gallons or equivalent units, of usable fuel in
each tank during flight. 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. 29.959;
(2) When two or more tanks are closely interconnected by a gravity feed
system and vented, and when it is impossible to feed from each tank
separately, at least one fuel quantity indicator must be installed;
(3) Tanks with interconnected outlets and airspaces may be treated as one
tank and need not have separate indicators; and
(4) Each exposed sight gauge used as a fuel quantity indicator must be
protected against damage.
(c) Fuel flowmeter system. If a fuel flowmeter system is installed, each
metering component must have a means for bypassing the fuel supply if
malfunction of that component severely restricts fuel flow.
(d) Oil quantity indicator. There must be a stick gauge or equivalent means
to indicate the quantity of oil--
(1) In each tank; and
(2) In each transmission gearbox.
(e) Rotor drive system transmissions and gearboxes utilizing ferromagnetic
materials must be equipped with chip detectors designed to indicate the
presence of ferromagnetic particles resulting from damage or excessive wear
within the transmission or gearbox. Each chip detector must--
(1) Be designed to provide a signal to the indicator required by Sec.
29.1305(a)(22); and
(2) Be provided with a means to allow crewmembers to check, in flight, the
function of each detector electrical circuit and signal.
(Secs. 313(a), 601, and 603, 72 Stat. 759, 775, 49 U.S.C. 1354(a), 1421, and
1423; sec. 6(c), 49 U.S.C. 1655(c))
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-13, 42 FR
15047, Mar. 17, 1977; Amdt. 29-26, 53 FR 34219, Sept. 2, 1988]
Electrical Systems and Equipment
Sec. 29.1351 General.
(a) Electrical system capacity. The required generating capacity and the
number and kind of power sources must--
(1) Be determined by an electrical load analysis; and
(2) Meet the requirements of Sec. 29.1309.
(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) Power sources function properly when independent and when connected in
combination;
(2) No failure or malfunction of any power source can create a hazard or
impair the ability of remaining sources to supply essential loads;
(3) The system voltage and frequency (as applicable) at the terminals of
essential load equipment can be maintained within the limits for which the
equipment is designed, during any probable operating condition;
(4) 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;
(5) There are means accessible in flight to appropriate crewmembers for the
individual and collective disconnection of the electrical power sources from
the main bus; and
(6) There are means to indicate to appropriate crewmembers the generating
system quantities essential for the safe operation of the system, such as the
voltage and current supplied by each generator.
(c) External power. If provisions are made for connecting external power to
the rotorcraft, 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 rotorcraft's electrical system.
(d) Operation without normal electrical power. It must be shown by
analysis, tests, or both, that the rotorcraft 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)
inoperative, with critical type fuel (from the standpoint of flameout and
restart capability, and with the rotorcraft 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;
(2) The parts turned on are electrically and mechanically isolated from the
parts turned off; and
(3) The electrical wire and cable insulation, and other materials, of the
parts turned on are self-extinguishing when tested in accordance with Sec.
25.1359(d) in effect on September 1, 1977.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49
U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. of
Transportation Act (49 U.S.C. 1655(c)))
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-14, 42 FR
36973, July 18, 1977]
Sec. 29.1353 Electrical equipment and installations.
(a) 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 any other electrical unit or system essential to
safe operation.
(b) Cables must be grouped, routed, and spaced so that damage to essential
circuits will be minimized if there are faults in heavy current-carrying
cables.
(c) Storage batteries must be designed and installed as follows:
(1) 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)--
(i) At maximum regulated voltage or power;
(ii) During a flight of maximum duration; and
(iii) Under the most adverse cooling condition likely in service.
(2) Compliance with paragraph (a)(1) of this section must be shown by test
unless experience with similar batteries and installations has shown that
maintaining safe cell temperatures and pressures presents no problem.
(3) 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
rotorcraft.
(4) No corrosive fluids or gases that may escape from the battery may
damage surrounding structures or adjacent essential equipment.
(5) 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.
(6) Nickel cadmium battery installations capable of being used to start an
engine or auxiliary power unit must have--
(i) A system to control the charging rate of the battery automatically so
as to prevent battery overheating;
(ii) 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
(iii) A battery failure sensing and warning system with a means for
disconnecting the battery from its charging source in the event of battery
failure.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49
U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. of
Transportation Act (49 U.S.C. 1655(c)))
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-14, 42 FR
36973, July 18, 1977; Amdt. 29-15, 43 FR 2327, Jan. 16, 1978]
Sec. 29.1355 Distribution system.
(a) The distribution system includes the distribution busses, their
associated feeders, and each control and protective device.
(b) If two independent sources of electrical power for particular equipment
or systems are required by this chapter, in the event of the failure of one
power source for such equipment or system, another power source (including
its separate feeder) must be provided automatically or be manually selectable
to maintain equipment or system operation.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49
U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. of
Transportation Act (49 U.S.C. 1655(c)))
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-14, 42 FR
36973, July 18, 1977; Amdt. 29-24, 49 FR 44439, Nov. 6, 1984]
Sec. 29.1357 Circuit protective devices.
(a) Automatic protective devices must be used to minimize distress to the
electrical system and hazard to the rotorcraft system and hazard to the
rotorcraft in the event of wiring faults or serious malfunction of the system
or connected equipment.
(b) The protective and control devices in the generating system must be
designed to de-energize and disconnect faulty power sources and power
transmission equipment from their associated buses with sufficient rapidity
to provide protection from hazardous overvoltage and other malfunctioning.
(c) Each resettable circuit protective device must be designed so that,
when an overload or circuit fault exists, it 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 located
and identified so that it can be readily reset or replaced in flight.
(e) Each essential load must have individual circuit protection. However,
individual protection for each circuit in an essential load system (such as
each position light circuit in a system) is not required.
(f) If fuses are used, there must be spare fuses for use in flight equal to
at least 50 percent of the number of fuses of each rating required for
complete circuit protection.
(g) Automatic reset circuit breakers may be used as integral protectors for
electrical equipment provided there is circuit protection for the cable
supplying power to the equipment.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-24, 49 FR
44440, Nov. 6, 1984]
Sec. 29.1359 Electrical system fire and smoke protection.
(a) Components of the electrical system must meet the applicable fire and
smoke protection provisions of Secs. 29.831 and 29.863.
(b) Electrical cables, terminals, and equipment, in designated fire zones,
and that are used in emergency procedures, must be at least fire resistant.
Sec. 29.1363 Electrical system tests.
(a) When laboratory tests of the electrical system are conducted--
(1) The tests must be performed on a mock-up using the same generating
equipment used in the rotorcraft;
(2) The equipment must simulate the electrical characteristics of the
distribution wiring and connected loads to the extent necessary for valid
test results; and
(3) Laboratory generator drives must simulate the prime movers on the
rotorcraft with respect to their reaction to generator loading, including
loading due to faults.
(b) For each flight condition that cannot be simulated adequately in the
laboratory or by ground tests on the rotorcraft, flight tests must be made.
Lights
Sec. 29.1381 Instrument lights.
The instrument lights must--
(a) Make each instrument, switch, and other device for which they are
provided easily readable; and
(b) Be installed so that--
(1) Their direct rays are shielded from the pilot's eyes; and
(2) No objectionable reflections are visible to the pilot.
Sec. 29.1383 Landing lights.
(a) Each required landing or hovering light must be approved.
(b) Each landing light must be installed so that--
(1) No objectionable glare is visible to the pilot;
(2) The pilot is not adversely affected by halation; and
(3) It provides enough light for night operation, including hovering and
landing.
(c) At least one separate switch must be provided, as applicable--
(1) For each separately installed landing light; and
(2) For each group of landing lights installed at a common location.
Sec. 29.1385 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. 29.1387 through 29.1397.
(b) Forward position lights. Forward position lights must consist of a red
and a green light spaced laterally as far apart as practicable and installed
forward on the rotorcraft so that, with the rotorcraft in the normal flying
position, the red light is on the left side, and the green light is on the
right side. Each light must be approved.
(c) Rear position light. The rear position light must be a white light
mounted as far aft as practicable, and must be approved.
(d) Circuit. The two forward position lights and the rear position light
must make a single circuit.
(e) 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.
Sec. 29.1387 Position light system dihedral angles.
(a) Except as provided in paragraph (e) of this section, each forward and
rear 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 rotorcraft, 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 rotorcraft, 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. 29.1385(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.
(49 U.S.C. 1655(c))
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-9, 36 FR
21279, Nov. 5, 1971]
Sec. 29.1389 Position light distribution and intensities.
(a) General. The intensities prescribed in this section must be provided by
new equipment with light covers and color filters 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
rotorcraft. The light distribution and intensity of each position light must
meet the requirements of paragraph (b) of this section.
(b) Forward and rear position lights. The light distribution and
intensities of forward and rear 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 rotorcraft and
perpendicular to the plane of symmetry of the rotorcraft), must equal or
exceed the values in Sec. 29.1391.
(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. 29.1393 where I is the minimum intensity prescribed
in Sec. 29.1391 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. 29.1395,
except that higher intensities in overlaps may be used with the use of main
beam intensities substantially greater than the minima specified in Secs.
29.1391 and 29.1393 if the overlap intensities in relation to the main beam
intensities do not adversely affect signal clarity.
Sec. 29.1391 Minimum intensities in the horizontal plane of forward and rear
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 (forward red and 0 deg. to 10 deg. 40
green) 10 deg. to 20 deg. 30
20 deg. to 110 deg. 5
A (rear white) 110 deg. to 180 deg. 20
Sec. 29.1393 Minimum intensities in any vertical plane of forward and rear
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 I
0 deg. 1.00
0 deg. to 5 deg. .90
5 deg. to 10 deg. .80
10 deg. to 15 deg. .70
15 deg. to 20 deg. .50
20 deg. to 30 deg. .30
30 deg. to 40 deg. .10
40 deg. to 90 deg. .05
Sec. 29.1395 Maximum intensities in overlapping beams of forward and rear
position lights.
No position light intensity may exceed the applicable values in the
following table, except as provided in Sec. 29.1389(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.
Sec. 29.1397 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--0.320y;
"x" is not greater than y--0.170; and
"y" is not less than 0.390--0.170x.
(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 "yc--0.010," whichever is the smaller;
and
"y" is not greater than "x+0.020" nor "0.636--0.400 x";
Where "Ye" is the "y" coordinate of the Planckian radiator for the value of
"x" considered.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-7, 36 FR
12972, July 10, 1971]
Sec. 29.1399 Riding light.
(a) Each riding light required for water operation 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 a maximum practicable unbroken light with the rotorcraft on the
water.
(b) Externally hung lights may be used.
Sec. 29.1401 Anticollision light system.
(a) General. If certification for night operation is requested, the
rotorcraft must have an anticollision light system that--
(1) Consists of one or more approved anticollision lights located so that
their emitted light will not impair the crew's 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 rotorcraft, considering the physical
configuration and flight characteristics of the rotorcraft. The field of
coverage must extend in each direction within at least 30 degrees above and
30 degrees below the horizontal plane of the rotorcraft, 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 rotorcraft'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 aviation red and must meet the
applicable requirements of Sec. 29.1397.
(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-tl=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 light. 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. 150
5 deg. to 10 deg. 90
10 deg. to 20 deg. 30
20 deg. to 30 deg. 15
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-7, 36 FR
12972, July 10, 1971; Amdt. 29-11, 41 FR 5290, Feb. 5, 1976]
Safety Equipment
Sec. 29.1411 General.
(a) Accessibility. Required safety equipment to be used by the crew in an
emergency, such as automatic liferaft releases, must be readily accessible.
(b) Stowage provisions. Stowage provisions for required emergency 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 inadvertent damage.
(c) Emergency exit descent device. The stowage provisions for the emergency
exit descent device required by Sec. 29.809(f) must be at the exits for which
they are intended.
(d) Liferafts. Liferafts must be stowed near exits through which the rafts
can be launched during an unplanned ditching. Rafts automatically or remotely
released outside the rotorcraft must be attached to the rotorcraft by the
static line prescribed in Sec. 29.1415.
(e) Long-range signaling device. The stowage provisions for the long-range
signaling device required by Sec. 29.1415 must be near an exit available
during an unplanned ditching.
(f) Life preservers. Each life preserver must be within easy reach of each
occupant while seated.
Sec. 29.1413 Safety belts: passenger warning device.
(a) If there are means to indicate to the passengers when safety belts
should be fastened, they must be installed to be operated from either pilot
seat.
(b) Each safety belt must be equipped with a metal to metal latching
device.
(Secs. 313, 314, and 601 through 610 of the Federal Aviation Act of 1958 (49
U.S.C. 1354, 1355, and 1421 through 1430) and sec. 6(c), Dept. of
Transportation Act (49 U.S.C. 1655(c)))
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-16 43 FR
46233, Oct. 5, 1978]
Sec. 29.1415 Ditching equipment.
(a) Emergency flotation and signaling equipment required by any operating
rule of this chapter must meet the requirements of this section.
(b) Each liferaft and each life preserver must be approved. In addition--
(1) Provide not less than two rafts, of an approximately equal rated
capacity and buoyancy to accommodate the occupants of the rotorcraft; and
(2) Each raft must have a trailing line, and must have a static line
designed to hold the raft near the rotorcraft but to release it if the
rotorcraft becomes totally submerged.
(c) Approved survival equipment must be attached to each liferaft.
(d) There must be an approved survival type emergency locator transmitter
for use in one life raft.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-8, 36 FR
18722, Sept. 21, 1971; Amdt. 29-19, 45 FR 38348, June 9, 1980; Amdt. 29-30,
55 FR 8005, Mar. 6, 1990; Amdt. 29-33, 59 FR 32057, June 21, 1994]
*****************************************************************************
59 FR 32050, No. 118, June 21, 1994
SUMMARY: This rule requires that newly installed emergency locator
transmitters (ELT's) on U.S.-registered aircraft be of an improved design
that meets the requirements of a revised Technical Standard Order (TSO) or
later TSO's issued for ELT's. This rule is prompted by unsatisfactory
performance experienced with automatic ELT's manufactured under the original
TSO. Further, it addresses certain safety recommendations made by the
National Transportation Safety Board (NTSB) and the search and rescue (SAR)
community. The FAA is also adopting improved standards for survival ELT's.
The rule is expected to have a dramatic effect on reducing activation
failures and would increase the likelihood of locating airplanes after
accidents. In addition, publication of this document coincides with notice of
the FAA's withdrawal of manufacturing authority for ELT's produced under TSO-
C91.
EFFECTIVE DATE: This document is effective June 21, 1994.
(a) To obtain certification for flight into icing conditions, compliance
with this section must be shown.
(b) It must be demonstrated that the rotorcraft can be safely operated in
the continuous maximum and intermittent maximum icing conditions determined
under Appendix C of this part within the rotorcraft altitude envelope. An
analysis must be performed to establish, on the basis of the rotorcraft's
operational needs, the adequacy of the ice protection system for the various
components of the rotorcraft.
(c) In addition to the analysis and physical evaluation prescribed in
paragraph (b) of this section, the effectiveness of the ice protection system
and its components must be shown by flight tests of the rotorcraft 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 tests of the rotorcraft or its components in measured simulated
icing conditions.
(d) The ice protection provisions of this section are considered to be
applicable primarily to the airframe. Powerplant installation requirements
are contained in Subpart E of this part.
(e) A means must be identified or provided for determining the formation of
ice on critical parts of the rotorcraft. Unless otherwise restricted, the
means must be available for nighttime as well as daytime operation. The
rotorcraft flight manual must describe the means of determining ice formation
and must contain information necessary for safe operation of the rotorcraft
in icing conditions.
[Amdt. 29-21, 48 FR 4391, Jan. 31, 1983]
Miscellaneous Equipment
Sec. 29.1431 Electronic equipment.
(a) Radio communication and navigation equipment installations must be free
from hazards in themselves, in their method of operation, and in their
effects on other components, under any critical environmental conditions.
(b) Radio communication and navigation 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 any other radio or electronic
unit, or system of units, required by this chapter.
Sec. 29.1433 Vacuum systems.
(a) There must be means, in addition to the normal pressure relief, to
automatically relieve the pressure in the discharge lines from the vacuum air
pump when the delivery temperature of the air becomes unsafe.
(b) Each vacuum air system line and fitting on the discharge side of the
pump that might contain flammable vapors or fluids must meet the requirements
of Sec. 29.1183 if they are in a designated fire zone.
(c) Other vacuum air system components in designated fire zones must be at
least fire resistant.
Sec. 29.1435 Hydraulic systems.
(a) Design. Each hydraulic system must be designed as follows:
(1) Each element of the hydraulic system must be designed to withstand,
without detrimental, permanent deformation, any structural loads that may be
imposed simultaneously with the maximum operating hydraulic loads.
(2) Each element of the hydraulic system must be designed to withstand
pressures sufficiently greater than those prescribed in paragraph (b) of this
section to show that the system will not rupture under service conditions.
(3) There must be means to indicate the pressure in each main hydraulic
power system.
(4) There must be means to ensure that no pressure in any part of the
system will exceed a safe limit above the maximum operating pressure of the
system, and to prevent excessive pressures resulting from any fluid
volumetric change in lines likely to remain closed long enough for such a
change to take place. The possibility of detrimental transient (surge)
pressures during operation must be considered.
(5) Each hydraulic line, fitting, and component must be installed and
supported to prevent excessive vibration and to withstand inertia loads. Each
element of the installation must be protected from abrasion, corrosion, and
mechanical damage.
(6) Means for providing flexibility must be used to connect points, in a
hydraulic fluid line, between which relative motion or differential vibration
exists.
(b) Tests. Each element of the system must be tested to a proof pressure of
1.5 times the maximum pressure to which that element will be subjected in
normal operation, without failure, malfunction, or detrimental deformation of
any part of the system.
(c) Fire protection. Each hydraulic system using flammable hydraulic fluid
must meet the applicable requirements of Secs. 29.861, 29.1183, 29.1185, and
29.1189.
Sec. 29.1439 Protective breathing equipment.
(a) If one or more cargo or baggage compartments are to be accessible in
flight, protective breathing equipment must be available for an appropriate
crewmember.
(b) For protective breathing equipment required by paragraph (a) of this
section or by any operating rule of this chapter--
(1) That equipment must be designed to protect the crew from smoke, carbon
dioxide, and other harmful gases while on flight deck duty;
(2) That equipment must include--
(i) Masks covering the eyes, nose, and mouth; or
(ii) Masks covering the nose and mouth, plus accessory equipment to protect
the eyes; and
(3) That equipment must supply protective oxygen of 10 minutes duration per
crewmember at a pressure altitude of 8,000 feet with a respiratory minute
volume of 30 liters per minute BTPD.
Sec. 29.1457 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 rotorcraft by
radio.
(2) Voice communications of flight crewmembers on the flight deck.
(3) Voice communications of flight crewmembers on the flight deck, using
the rotorcraft'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 may be
met--
(1) 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; or
(2) By installing a continually energized or voice-actuated lip microphone
at the first and second pilot stations.
The microphone specified in this paragraph must be so located and, if
necessary, the preamplifiers and filters of the recorder must be so adjusted
or supplemented, that the recorded communications are intelligible when
recorded under flight cockpit noise conditions and played back. The level of
intelligibility must be approved by the Administrator. 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 microphone, headset, or speaker used
at the first pilot station.
(2) For the second channel, from each microphone, headset, or speaker used
at the second pilot station.
(3) For the third channel, from the cockpit-mounted area microphone, or the
continually energized or voice-actuated lip microphones at the first and
second pilot stations.
(4) For the fourth channel, from--
(i) Each microphone, headset, or speaker used at the stations for the third
and fourth crewmembers; or
(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.
(iii) Each microphone on the flight deck that is used with the rotorcraft's
loudspeaker system if its signals are not picked up by another channel.
(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.
(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 be either bright orange or bright yellow.
[Amdt. 29-6, 35 FR 7293, May 9, 1970]
Sec. 29.1459 Flight recorders.
(a) Each flight recorder required by the operating rules of Subchapter G 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. 29.1323, 29.1325,
and 29.1327 of this part, as applicable;
(2) The vertical acceleration sensor is rigidly attached, and located
longitudinally within the approved center of gravity limits of the
rotorcraft;
(3) It receives its electrical 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 perflight checking of the
recorder for proper recording of data in the storage medium; and
(5) Except for recorders powered solely by the engine-drive 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 any crash impact.
(b) Each nonejectable recorder 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.
(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. This
correlation must cover the airspeed range over which the aircraft is to be
operated, the range of altitude to which the aircraft 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 a 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 it is not likely to be separated during crash impact.
[Amdt. 29-25, 53 FR 26145, July 11, 1988; 53 FR 26144, July 11, 1988]
Sec. 29.1461 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. 29-3, 33 FR 971, Jan. 26, 1968]
Subpart G--Operating Limitations and Information
Sec. 29.1501 General.
(a) Each operating limitation specified in Secs. 29.1503 through 29.1525
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.
29.1541 through 29.1589.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49
U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. of
Transportation Act (49 U.S.C. 1655(c)))
[Amdt. 29-15, 43 FR 2327, Jan. 16, 1978]
Operating Limitations
Sec. 29.1503 Airspeed limitations: general.
(a) An operating speed range must be established.
(b) When airspeed limitations are a function of weight, weight
distribution, altitude, rotor speed, power, or other factors, airspeed
limitations corresponding with the critical combinations of these factors
must be established.
Sec. 29.1505 Never-exceed speed.
(a) The never-exceed speed, VNE, must be established so that it is--
(1) Not less than 40 knots (CAS); and
(2) Not more than the lesser of--
(i) 0.9 times the maximum forward speeds established under Sec. 29.309;
(ii) 0.9 times the maximum speed shown under Secs. 29.251 and 29.629; or
(iii) 0.9 times the maximum speed substantiated for advancing blade tip
mach number effects under critical altitude conditions.
(b) VNE may vary with altitude, r.p.m., temperature, and weight, if--
(1) No more than two of these variables (or no more than two instruments
integrating more than one of these variables) are used at one time; and
(2) The ranges of these variables (or of the indications on instruments
integrating more than one of these variables) are large enough to allow an
operationally practical and safe variation of VNE.
(c) For helicopters, a stabilized power-off VNE denoted as VNE (power-off)
may be established at a speed less than VNE established pursuant to paragraph
(a) of this section, if the following conditions are met:
(1) VNE (power-off) is not less than a speed midway between the power-on
VNE and the speed used in meeting the requirements of--
(i) Sec. 29.67(a)(3) for Category A helicopters;
(ii) Sec. 29.65(a) for Category B helicopters, except multi-engine
helicopters meeting the requirements of Sec. 29.67(b); and
(iii) Sec. 29.67(b) for multi-engine Category B helicopters meeting the
requirements of Sec. 29.67(b).
(2) VNE (power-off) is--
(i) A constant airspeed;
(ii) A constant amount less than power-on VNE; or
(iii) A constant airspeed for a portion of the altitude range for which
certification is requested, and a constant amount less than power-on VNE for
the remainder of the altitude range.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49
U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. of
Transportation Act (49 U.S.C. 1655(c)))
(a) Maximum power-off (autorotation). The maximum power-off rotor speed
must be established so that it does not exceed 95 percent of the lesser of--
(1) The maximum design r.p.m. determined under Sec. 29.309(b); and
(2) The maximum r.p.m. shown during the type tests.
(b) Minimum power-off. The minimum power-off rotor speed must be
established so that it is not less than 105 percent of the greater of--
(1) The minimum shown during the type tests; and
(2) The minimum determined by design substantiation.
(c) Minimum power-on. The minimum power-on rotor speed must be established
so that it is--
(1) Not less than the greater of--
(i) The minimum shown during the type tests; and
(ii) The minimum determined by design substantiation; and
(2) Not more than a value determined under Sec. 29.33 (a)(1) and (c)(1).
Sec. 29.1517 Limiting height-speed envelope.
For Category A rotorcraft, if a range of heights exists at any speed,
including zero, within which it is not possible to make a safe landing
following power failure, the range of heights and its variation with forward
speed must be established, together with any other pertinent information,
such as the kind of landing surface.
[Amdt. 29-21, 48 FR 4391, Jan. 31, 1983]
Sec. 29.1519 Weight and center of gravity.
The weight and center of gravity limitations determined under Secs. 29.25
and 29.27, respectively, must be established as operating limitations.
Sec. 29.1521 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 are type certificated.
(b) Takeoff operation. The powerplant takeoff operation must be limited
by--
(1) The maximum rotational speed, which may not be greater than--
(i) The maximum value deterimined be the rotor design; or
(ii) The maximum value shown during the type tests;
(2) The maximum allowable manifold pressure (for reciprocating engines);
(3) The maximum allowable turbine inlet or turbine outlet gas temperature
(for turbine engines);
(4) The maximum allowable power or torque for each engine, considering the
power input limitations of the transmission with all engines operating;
(5) The maximum allowable power or torque for each engine considering the
power input limitations of the transmission with one engine inoperative;
(6) The time limit for the use of the power corresponding to the
limitations established in paragraphs (b) (1) through (5) of this section;
and
(7) If the time limit established in paragraph (b)(6) of this section
exceeds 2 minutes--
(i) The maximum allowable cylinder head or coolant outlet temperature (for
reciprocating engines); and
(ii) The maximum allowable engine and transmission oil temperatures.
(c) Continuous operation. The continuous operation must be limited by--
(1) The maximum rotational speed, which may not be greater than--
(i) The maximum value determined by the rotor design; or
(ii) The maximum value shown during the type tests;
(2) The minimum rotational speed shown under the rotor speed requirements
in Sec. 29.1509(c).
(3) The maximum allowable manifold pressure (for reciprocating engines);
(4) The maximum allowable turbine inlet or turbine outlet gas temperature
(for turbine engines);
(5) The maximum allowable power or torque for each engine, considering the
power input limitations of the transmission with all engines operating;
(6) The maximum allowable power or torque for each engine, considering the
power input limitations of the transmission with one engine inoperative; and
(7) The maximum allowable temperatures for--
(i) The cylinder head or coolant outlet (for reciprocating engines);
(ii) The engine oil; and
(iii) The transmission oil.
(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. 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. 29.1041 through 29.1049 is
shown.
(f) Two and one-half minute OEI power operation. Unless otherwise
authorized, the use of 2 1/2 -minute OEI power must be limited to engine
failure operation of multiengine, turbine-powered rotorcraft for not longer
than 2 1/2 minutes for any period in which that power is used. The use of 2
1/2 -minute OEI power must also be limited by--
(1) The maximum rotational speed, which may not be greater than--
(i) The maximum value determined by the rotor design; or
(ii) The maximum value shown during the type tests;
(2) The maximum allowable gas temperature;
(3) The maximum allowable torque; and
(4) The maximum allowable oil temperature.
(g) Thirty-minute OEI power operation. Unless otherwise authorized, the use
of 30-minute OEI power must be limited to multiengine, turbine-powered
rotorcraft for not longer than 30 minutes after failure of an engine. The use
of 30-minute OEI power must also be limited by--
(1) The maximum rotational speed, which may not be greater than--
(i) The maximum value determined by the rotor design; or
(ii) The maximum value shown during the type tests;
(2) The maximum allowable gas temperature;
(3) The maximum allowable torque; and
(4) The maximum allowable oil temperature.
(h) Continuous OEI power operation. Unless otherwise authorized, the use of
continuous OEI power must be limited to multiengine, turbine-powered
rotorcraft for continued flight after failure of an engine. The use of
continuous OEI power must also be limited by--
(1) The maximum rotational speed, which may not be greater than--
(i) The maximum value determined by the rotor design; or
(ii) The maximum value shown during the type tests.
(2) The maximum allowable gas temperature;
(3) The maximum allowable torque; and
(4) The maximum allowable oil temperature.
(i) Rated 30-second OEI power operation. Rated 30-second OEI power is
permitted only on multiengine, turbine-powered rotorcraft, also certificated
for the use of rated 2-minute OEI power, and can only be used for continued
operation of the remaining engine(s) after a failure or precautionary
shutdown of an engine. It must be shown that following application of 30-
second OEI power, any damage will be readily detectable by the applicable
inspections and other related procedures furnished in accordance with Section
A29.4 of Appendix A of this part and Section A33.4 of Appendix A of part 33.
The use of 30-second OEI power must be limited to not more than 30 seconds
for any period in which that power is used, and by--
(1) The maximum rotational speed which may not be greater than--
(i) The maximum value determined by the rotor design; or
(ii) The maximum value demonstrated during the type tests;
(2) The maximum allowable gas temperature; and
(3) The maximum allowable torque.
(j) Rated 2-minute OEI power operation. Rated 2-minute OEI power is
permitted only on multiengine, turbine-powered rotorcraft, also certificated
for the use of rated 30-second OEI power, and can only be used for continued
operation of the remaining engine(s) after a failure or precautionary
shutdown of an engine. It must be shown that following application of 2-
minute OEI power, any damage will be readily detectable by the applicable
inspections and other related procedures furnished in accordance with Section
A29.4 of Appendix a of this part and Section A33.4 of Appendix A of part 33.
The use of 2-minute OEI power must be limited to not more than 2 minutes for
any period in which that power is used, and by--
(1) The maximum rotational speed, which may not be greater than--
(i) The maximum value determined by the rotor design; or
(ii) The maximum value demonstrated during the type tests;
(2) The maximum allowable gas temperature; and
(3) The maximum allowable torque.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49
U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. of
Transportation Act (49 U.S.C. 1655(c)))
SUMMARY: This rule adopts new and revised airworthiness standards by
incorporating optional one-engine-inoperative (OEI) power ratings for
multiengine, turbine-powered rotorcraft. These amendments result from a
petition for rulemaking from Aerospace Industries Association of America
(AIA) and the recognition by both government and industry that additional OEI
power rating standards are needed. These amendments enhance rotorcraft safety
after an engine failure or precautionary shutdown by providing higher OEI
power, when necessary. These amendments also assure that the drive system
will maintain its structural integrity and allow continued safe flight while
operating at the new OEI power ratings with the operable engine(s).
If an auxiliary power unit that meets the requirements of TSO-C77 is
installed in the rotorcraft, the limitations established for that auxiliary
power unit under the TSO including the categories of operation must be
specified as operating limitations for the rotorcraft.
(Secs. 313(a), 601, 603, 604, Federal Aviation Act of 1958 (49 U.S.C.
1354(a), 1421, 1423), sec. 6(c), Dept. of Transportation Act (49 U.S.C.
1655(c)))
[Amdt. 29-17, 43 FR 50602, Oct. 30, 1978]
Sec. 29.1523 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;
(b) The accessibility and ease of operation of necessary controls by the
appropriate crewmember; and
(c) The kinds of operation authorized under Sec. 29.1525.
Sec. 29.1525 Kinds of operations.
The kinds of operations (such as VFR, IFR, day, night, or icing) for which
the rotorcraft is approved are established by demonstrated compliance with
the applicable certification requirements and by the installed equipment.
[Amdt. 29-24, 49 FR 44440, Nov. 6, 1984]
Sec. 29.1527 Maximum operating altitude.
The maximum altitude up to which operation is allowed, as limited by
flight, structural, powerplant, functional, or equipment characteristics,
must be established.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49
U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. of
Transportation Act (49 U.S.C. 1655(c)))
[Amdt. 29-15, 43 FR 2327, Jan. 16, 1978]
Sec. 29.1529 Instructions for continued airworthiness.
The applicant must prepare Instructions for Continued Airworthiness in
accordance with Appendix A 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
rotorcraft or issuance of a standard certificate of airworthiness, whichever
occurs later.
[Amdt. 29-20, 45 FR 60178, Sept. 11, 1980]
Markings and Placards
Sec. 29.1541 General.
(a) The rotorcraft must contain--
(1) The markings and placards specified in Secs. 29.1545 through 29.1565;
and
(2) Any additional information, instrument markings, and placards required
for the safe operation of the rotorcraft 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, disfig-ured, or obscured.
Sec. 29.1543 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. 29.1545 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) A red radial line--
(i) For rotorcraft other than helicopters, at VNE; and
(ii) For helicopters, at a VNE (power-on).
(2) A red, cross-hatched radial line at VNE (power-off) for helicopters, if
VNE (power-off) is less than VNE (power-on).
(3) For the caution range, a yellow arc.
(4) For the safe operating range, a green arc.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49
U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. of
Transportation Act (49 U.S.C. 1655(c)))
(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 heading in not
more than 45 degree increments.
Sec. 29.1549 Powerplant instruments.
For each required powerplant 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 yellow line;
(d) Each engine or propeller range that is restricted because of excessive
vibration stresses must be marked with red arcs or red lines; and
(e) Each OEI limit or approved operating range must be marked to be clearly
differentiated from the markings of paragraphs (a) through (d) of this
section except that no marking is normally required for the 30-second OEI
limit.
SUMMARY: This rule adopts new and revised airworthiness standards by
incorporating optional one-engine-inoperative (OEI) power ratings for
multiengine, turbine-powered rotorcraft. These amendments result from a
petition for rulemaking from Aerospace Industries Association of America
(AIA) and the recognition by both government and industry that additional OEI
power rating standards are needed. These amendments enhance rotorcraft safety
after an engine failure or precautionary shutdown by providing higher OEI
power, when necessary. These amendments also assure that the drive system
will maintain its structural integrity and allow continued safe flight while
operating at the new OEI power ratings with the operable engine(s).
Each oil quantity indicator must be marked with enough increments to
indicate readily and accurately the quantity of oil.
Sec. 29.1553 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. 29.1555 Control markings.
(a) Each cockpit control, other than primary flight controls or control
whose function is obvious, must be plainly marked as to its function and
method of operation.
(b) For powerplant fuel controls--
(1) Each fuel tank selector valve 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 adjacent to, the selector for those
tanks; and
(3) Each valve control for any engine of a multiengine rotorcraft must be
marked to indicate the position corresponding to each engine controlled.
(c) 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.
(d) For accessory, auxiliary, and emergency controls--
(1) Each essential visual position indicator, such as those showing rotor
pitch or landing gear position, must be marked so that each crewmember can
determine at any time the position of the unit to which it relates; and
(2) Each emergency control must be red and must be marked as to method of
operation.
(e) For rotorcraft incorporating retractable landing gear, the maximum
landing gear operating speed must be displayed in clear view of the pilot.
(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 and oil filler openings. The following apply:
(1) Fuel filler openings must be marked at or near the filler cover with--
(i) The word "fuel";
(ii) For reciprocating engine powered rotorcraft, the minimum fuel grade;
(iii) For turbine-engine-powered rotorcraft, the permissible fuel
designations, except that if impractical, this information may be included in
the rotorcraft flight manual, and the fuel filler may be marked with an
appropriate reference to the flight manual; and
(iv) 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".
(d) Emergency exit placards. Each placard and operating control for each
emergency exit must differ in color from the surrounding fuselage surface as
prescribed in Sec. 29.811(h)(2). A placard must be near each emergency exit
control and must clearly indicate the location of that exit and its method of
operation.
There must be a placard in clear view of the pilot that specifies the kinds
of operations (VFR, IFR, day, night, or icing) for which the rotorcraft is
approved.
[Amdt. 29-24, 49 FR 44440, Nov. 6, 1984]
Sec. 29.1561 Safety equipment.
(a) Each safety equipment control to be operated by the crew in emergency,
such as controls for automatic liferaft releases, must be plainly marked as
to its method of operation.
(b) Each location, such as a locker or compartment, that carries any fire
extinguishing, signaling, or other life saving equipment, must be so marked.
(c) Stowage provisions for required emergency equipment must be
conspicuously marked to identify the contents and facilitate removal of the
equipment.
(d) Each liferaft must have obviously marked operating instructions.
(e) Approved survival equipment must be marked for identification and
method of operation.
Sec. 29.1565 Tail rotor.
Each tail rotor must be marked so that its disc is conspicuous under normal
daylight ground conditions.
[Amdt. 29-3, 33 FR 971, Jan. 26, 1968]
Rotorcraft Flight Manual
Sec. 29.1581 General.
(a) Furnishing information. A Rotorcraft Flight Manual must be furnished
with each rotorcraft, and it must contain the following:
(1) Information required by Secs. 29.1583 through 29.1589.
(2) Other information that is necessary for safe operation because of
design, operating, or handling characteristics.
(b) Approved information. Each part of the manual listed in Secs. 29.1583
through 29.1589 that is appropriate to the rotorcraft, must be furnished,
verified, and approved, and must be segregated, indentified, and clearly
distinguished from each unapproved part of that manual.
(c) [Reserved]
(d) Table of contents. Each Rotorcraft Flight Manual must include a table
of contents if the complexity of the manual indicates a need for it.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49
U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. of
Transportation Act (49 U.S.C. 1655(c)))
[Amdt. 29-15, 43 FR 2327, Jan. 16, 1978]
Sec. 29.1583 Operating limitations.
(a) Airspeed and rotor limitations. Information necessary for the marking
of airspeed and rotor limitations on or near their respective indicators must
be furnished. The significance of each limitation and of the color coding
must be explained.
(b) Powerplant limitations. The following information must be furnished:
(1) Limitations required by Sec. 29.1521.
(2) Explanation of the limitations, when appropriate.
(3) Information necessary for marking the instruments required by Secs.
29.1549 through 29.1553.
(c) Weight and loading distribution. The weight and center of gravity
limits required by Secs. 29.25 and 29.27, respectively, must be furnished. If
the variety of possible loading conditions warrants, instructions must be
included to allow ready observance of the limitations.
(d) Flight crew. When a flight crew of more than one is required, the
number and functions of the minimum flight crew determined under Sec. 29.1523
must be furnished.
(e) Kinds of operation. Each kind of operation for which the rotorcraft and
its equipment installations are approved must be listed.
(f) Limiting heights. Enough information must be furnished to allow
compliance with Sec. 29.1517.
(g) Maximum allowable wind. For Category A rotorcraft, the maximum
allowable wind for safe operation near the ground must be furnished.
(h) Altitude. The altitude established under Sec. 29.1527 and an
explanation of the limiting factors must be furnished.
(i) Ambient temperature. Maximum and minimum ambient temperature
limitations must be furnished.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49
U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. of
Transportation Act (49 U.S.C. 1655(c)))
(a) The parts of the manual containing operating procedures must have
information concerning any normal and emergency procedures, and other
information necessary for safe operation, including the applicable
procedures, such as those involving minimum speeds, to be followed if an
engine fails.
(b) For multiengine rotorcraft, information identifying each operating
condition in which the fuel system independence prescribed in Sec. 29.953 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.
(c) For helicopters for which a VNE (power-off) is established under Sec.
29.1505(c), information must be furnished to explain the VNE (power-off) and
the procedures for reducing airspeed to not more than the VNE (power-off)
following failure of all engines.
(d) For each rotorcraft showing compliance with Sec. 29.1353 (c)(6)(ii) or
(c)(6)(iii), the operating procedures for disconnecting the battery from its
charging source must be furnished.
(e) 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.
(f) Information on the total quantity of usable fuel for each fuel tank
must be furnished.
(g) For Category B rotorcraft, the airspeeds and corresponding rotor speeds
for minimum rate of descent and best glide angle as prescribed in Sec. 29.71
must be provided.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49
U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. of
Transportation Act (49 U.S.C. 1655(c)))
Flight manual performance information which exceeds any operating
limitation may be shown only to the extent necessary for presentation clarity
or to determine the effects of approved optional equipment or procedures.
When data beyond operating limits are shown, the limits must be clearly
indicated. The following must be provided:
(a) Category A. For each category A rotorcraft, the Rotorcraft Flight
Manual must contain a summary of the performance data, including data
necessary for the application of any operating rule of this chapter, together
with descriptions of the conditions, such as airspeeds, under which this data
was determined, and must contain--
(1) The indicated airspeeds corresponding with those determined for
takeoff, and the procedures to be followed if the critical engine fails
during takeoff;
(2) The airspeed calibrations;
(3) The techniques, associated airspeeds, and rates of descent for
autorotative landings;
(4) The rejected takeoff distance determined under Sec. 29.59(b) and the
takeoff distance determined under Sec. 29.59(c); and
(5) The landing data determined under Secs. 29.75 and 29.77.
(b) Category B. For each category B rotorcraft, the Rotorcraft Flight
Manual must contain--
(1) The takeoff distance and the climbout speed together with the pertinent
information defining the flight path with respect to autorotative landing if
an engine fails, including the calculated effects of altitude and
temperature;
(2) The steady rates of climb and hovering ceiling, together with the
corresponding airspeeds and other pertinent information, including the
calculated effects of altitude and temperature;
(3) The landing distance, appropriate glide airspeed, and kind of landing
surface, together with any pertinent information that might affect this
distance, including the calculated effects of altitude and temperature;
(4) The maximum safe wind for operation near the ground;
(5) The airspeed calibrations;
(6) The height-speed envelope except for rotorcraft incorporating this as
an operating limitation;
(7) Glide distance as a function of altitude when autorotating at the
speeds and conditions for minimum rate of descent and best glide angle, as
determined in Sec. 29.71;
(8) Maximum safe wind for hover operations out-of-ground effect if hover
performance for that condition is provided; and
(9) Any additional performance data necessary for the application of any
operating rule in this chapter.
There must be loading instructions for each possible loading condition
between the maximum and minimum weights determined under Sec. 29.25 that can
result in a center of gravity beyond any extreme prescribed in Sec. 29.27,
assuming any probable occupant weights.
Appendix A to Part 29--Instructions for Continued Airworthiness
A29.1 GENERAL
(a) This appendix specifies requirements for the preparation of
Instructions for Continued Airworthiness as required by Sec. 29.1529.
(b) The Instructions for Continued Airworthiness for each rotorcraft must
include the Instructions for Continued Airworthiness for each engine and
rotor (hereinafter designated "products" ), for each applicance required by
this chapter, and any required information relating to the interface of those
appliances and products with the rotorcraft. If Instructions for Continued
Airworthiness are not supplied by the manufacturer of an appliance or product
installed in the rotorcraft, the Instructions for Continued Airworthiness for
the rotorcraft must include the information essential to the continued
airworthiness of the rotorcraft.
(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 rotorcraft will be
distributed.
A29.2 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.
A29.3 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) Rotorcraft maintenance manual or section. (1) Introduction information
that includes an explanation of the rotorcraft's features and data to the
extent necessary for maintenance or preventive maintenance.
(2) A description of the rotorcraft and its systems and installations
including its engines, rotors, and appliances.
(3) Basic control and operation information describing how the rotorcraft
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, the 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 rotorcraft and its engines, auxiliary power units, rotors, 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 references 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 rotorcraft.
(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.
A29.4 AIRWORTHINESS LIMITATIONS SECTION
The Instructions for Continued Airworthiness must contain a section titled
Airworthiness Limitations that is segregated and clearly distinguishable from
the rest of the document. This section must set forth each mandatory
replacement time, structural inspection interval, and related structural
inspection procedure approved under Sec. 29.571. 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. 29-20, 45 FR 60178, Sept 11, 1980, as amended by Amdt. 29-27, 54 FR
34330, Aug. 18, 1989]
Effective Date Note: At 54 FR 34330, Aug. 18, 1989, Sec. A29.4 in Appendix
A, Part 29 was amended by changing the cross reference "Sec. 91.163" to "Sec.
91.403", effective August 18, 1990.
Appendix B to Part 29--Airworthiness Criteria for Helicopter Instrument
Flight
I. General. A transport category helicopter may not be type certificated
for operation under the instrument flight rules (IFR) of this chapter unless
it meets the design and installation requirements contained in this appendix.
II. Definitions. (a) VYI means instrument climb speed, utilized instead of
VY for compliance with the climb requirements for instrument flight.
(b) VNEI means instrument flight never exceed speed, utilized instead of
VNE for compliance with maximum limit speed requirements for instrument
flight.
(c) VMINI means instrument flight minimum speed, utilized in complying with
minimum limit speed requirements for instrument flight.
III. Trim. It must be possible to trim the cyclic, collective, and
directional control forces to zero at all approved IFR airspeeds, power
settings, and configurations appropriate to the type.
IV. Static longitudinal stability. (a) General. The helicopter must possess
positive static longitudinal control force stability at critical combinations
of weight and center of gravity at the conditions specified in paragraphs IV
(b) through (f) of this appendix. The stick force must vary with speed so
that any substantial speed change results in a stick force clearly
perceptible to the pilot. The airspeed must return to within 10 percent of
the trim speed when the control force is slowly released for each trim
condition specified in paragraphs IV (b) through (f) of this appendix.
(b) Climb. Stability must be shown in climb thoughout the speed range 20
knots either side of trim with--
(1) The helicopter trimmed at VYI;
(2) Landing gear retracted (if retractable); and
(3) Power required for limit climb rate (at least 1,000 fpm) at VYI or
maximum continuous power, whichever is less.
(c) Cruise. Stability must be shown throughout the speed range from 0.7 to
1.1 VH or VNEI, whichever is lower, not to exceed +/-20 knots from trim
with--
(1) The helicopter trimmed and power adjusted for level flight at 0.9 VH or
0.9 VNEI, whichever is lower; and
(2) Landing gear retracted (if retractable).
(d) Slow cruise. Stability must be shown throughout the speed range from
0.9 VMINI to 1.3 VMINI or 20 knots above trim speed, whichever is greater,
with--
(1) The helicopter trimmed and power adjusted for level flight at 1.1
VMINI; and
(2) Landing gear retracted (if retractable).
(e) Descent. Stability must be shown throughout the speed range 20 knots
either side of trim with--
(1) The helicopter trimmed at 0.8 VH or 0.8 VNEI (or 0.8 VLE for the
landing gear extended case), whichever is lower;
(2) Power required for 1,000 fpm descent at trim speed; and
(3) Landing gear extended and retracted, if applicable.
(f) Approach. Stability must be shown throughout the speed range from 0.7
times the minimum recommended approach speed to 20 knots above the maximum
recommended approach speed with--
(1) The helicopter trimmed at the recommended approach speed or speeds;
(2) Landing gear extended and retracted, if applicable; and
(3) Power required to maintain a 3 deg. glide path and power required to
maintain the steepest approach gradient for which approval is requested.
V. Static lateral-directional stability. (a) Static directional stability
must be positive throughout the approved ranges of airspeed, power, and
vertical speed. In straight, steady sideslips up to +/-10 deg. from trim,
directional control position must increase in approximately constant
proportion to angle of sideslip. At greater angles up to the maximum sideslip
angle appropriate to the type, increased directional control position must
produce increased angle of sideslip.
(b) During sideslips up to +/-10 deg. from trim throughout the approved
ranges of airspeed, power, and vertical speed there must be no negative
dihedral stability perceptible to the pilot through lateral control motion or
force. Longitudinal cycle movement with sideslip must not be excessive.
VI. Dynamic stability. (a) Any oscillation having a period of less than 5
seconds must damp to 1/2 amplitude in not more than one cycle.
(b) Any oscillation having a period of 5 seconds or more but less than 10
seconds must damp to 1/2 amplitude in not more than two cycles.
(c) Any oscillation having a period of 10 seconds or more but less than 20
seconds must be damped.
(d) Any oscillation having a period of 20 seconds or more may not achieve
double amplitude in less than 20 seconds.
(e) Any aperiodic response may not achieve double amplitude in less than 9
seconds.
VII. Stability augmentation system (SAS). (a) If a SAS is used, the
reliability of the SAS must be related to the effects of its failure. The
occurrence of any failure condition which would prevent continued safe flight
and landing must be extremely improbable. For any failure condition of the
SAS which is not shown to be extremely improbable--
(1) The helicopter must be safely controllable and capable of prolonged
instrument flight without undue pilot effort. Additional unrelated probable
failures affecting the control system must be considered; and
(2) The flight characteristics requirements in Subpart B of Part 29 must be
met throughout a practical flight envelope.
(b) The SAS must be designed so that it cannot create a hazardous deviation
in flight path or produce hazardous loads on the helicopter during normal
operation or in the event of malfunction or failure, assuming corrective
action begins within an appropriate period of time. Where multiple systems
are installed, subsequent malfunction conditions must be considered in
sequence unless their occurrence is shown to be improbable.
(c) Thunderstorm lights. In addition to the instrument lights required by
Sec. 29.1381(a), thunderstorm lights which provide high intensity white flood
lighting to the basic flight instruments must be provided. The thunderstorm
lights must be installed to meet the requirements of Sec. 29.1381(b).
VIII. Equipment, systems, and installation. The basic equipment and
installation must comply with Subpart F of Part 29 through Amendment 29-14,
with the following exceptions and additions:
(a) Flight and navigation instruments. (1) A magnetic gyro-stabilized
direction indicator instead of the gyroscopic direction indicator required by
Sec. 29.1303(h); and
(2) A standby attitude indicator which meets the requirements of Secs.
29.1303(g) (1) through (7), instead of a rate-of-turn indicator required by
Sec. 29.1303(g). If standby batteries are provided, they may be charged from
the aircraft electrical system if adequate isolation is incorporated. The
system must be designed so that the standby batteries may not be used for
engine starting.
(b) Miscellaneous requirements. (1) Instrument systems and other systems
essential for IFR flight that could be adversely affected by icing must be
provided with adequate ice protection whether or not the rotorcraft is
certificated for operation in icing conditions.
(2) There must be means in the generating system to automatically de-
energize and disconnect from the main bus any power source developing
hazardous overvoltage.
(3) Each required flight instrument using a power supply (electric, vacuum,
etc.) must have a visual means integral with the instrument to indicate the
adequacy of the power being supplied.
(4) When multiple systems performing like functions are required, each
system must be grouped, routed, and spaced so that physical separation
between systems is provided to ensure that a single malfunction will not
adversely affect more than one system.
(5) For systems that operate the required flight instruments at each
pilot's station--
(i) Only the required flight instruments for the first pilot may be
connected to that operating system;
(ii) Additional instruments, systems, or equipment may not be connected to
an operating system for a second pilot unless provisions are made to ensure
the continued normal functioning of the required instruments in the event of
any malfunction of the additional instruments, systems, or equipment which is
not shown to be extremely improbable;
(iii) The equipment, systems, and installations must be designed so that
one display of the information essential to the safety of flight which is
provided by the instruments will remain available to a pilot, without
additional crew-member action, after any single failure or combination of
failures that is not shown to be extremely improbable; and
(iv) For single-pilot configurations, instruments which require a static
source must be provided with a means of selecting an alternate source and
that source must be calibrated.
IX. Rotorcraft Flight Manual. A Rotorcraft Flight Manual or Rotorcraft
Flight Manual IFR Supplement must be provided and must contain--
(a) Limitations. The approved IFR flight envelope, the IFR flightcrew
composition, the revised kinds of operation, and the steepest IFR precision
approach gradient for which the helicopter is approved;
(b) Procedures. Required information for proper operation of IFR systems
and the recommended procedures in the event of stability augmentation or
electrical system failures; and
(c) Performance. If VYI differs from VY, climb performance at VYI and with
maximum continuous power throughout the ranges of weight, altitude, and
temperature for which approval is requested.
SUMMARY: This rule amends the airworthiness standards for systems,
propulsion, and airframe for both normal and transport category rotorcraft.
In addition, these amendments introduce safety improvements, clarifying
existing regulations, and standardize terminology. The changes are based on
some of the proposals that were submitted to the FAA by the European
Airworthiness Authorities. These amendments are also intended to encourage
the European community's acceptance of the Federal Aviation Regulations for
rotorcraft type certification, obviate development of different European
standards, and achieve increased commonality of airworthiness standards
among the respective countries.
(a) Continuous maximum icing. The maximum continuous intensity of
atmospheric icing conditions (continuous maximum icing) is defined by the
variables of the cloud liquid water content, the mean effective diameter of
the cloud droplets, the ambient air temperature, and the interrelationship of
these three variables as shown in Figure 1 of this appendix. The limiting
icing envelope in terms of altitude and temperature is given in Figure 2 of
this appendix. The interrelationship of cloud liquid water content with drop
diameter and altitude is determined from Figures 1 and 2. The cloud liquid
water content for continuous maximum icing conditions of a horizontal extent,
other than 17.4 nautical miles, is determined by the value of liquid water
content of Figure 1, multiplied by the appropriate factor from Figure 3 of
this appendix.
(b) Intermittent maximum icing. The intermittent maximum intensity of
atmospheric icing conditions (intermittent maximum icing) is defined by the
variables of the cloud liquid water content, the mean effective diameter of
the cloud droplets, the ambient air temperature, and the interrelationship of
these three variables as shown in Figure 4 of this appendix. The limiting
icing envelope in terms of altitude and temperature is given in Figure 5 of
this appendix. The interrelationship of cloud liquid water content with drop
diameter and altitude is determined from Figures 4 and 5. The cloud liquid
water content for intermittent maximum icing conditions of a horizontal
extent, other than 2.6 nautical miles, is determined by the value of cloud
liquid water content of Figure 4 multiplied by the appropriate factor in
Figure 6 of this appendix.
[ ...Illustration appears here... ]
Figures 1-3 -- Continuous Maximum (Stratiform Clouds)
[ ...Illustration appears here... ]
Figures 4-6 -- Intermittent Maximum (Cumuliform Clouds)
[Amdt. 29-21, 48 FR 4393, Jan. 31, 1983]
Appendix D--Criteria for Demonstration of Emergency Evacuation Procedures
Under Sec. 29.803
(a) The demonstration must be conducted either during the dark of the night
or during daylight with the dark of night simulated. If the demonstration is
conducted indoors during daylight hours, it must be conducted inside a
darkened hangar having doors and windows covered. In addition, the doors and
windows of the rotorcraft must be covered if the hangar illumination exceeds
that of a moonless night. Illumination on the floor or ground may be used,
but it must be kept low and shielded against shining into the rotorcraft's
windows or doors.
(b) The rotorcraft must be in a normal attitude with landing gear extended.
(c) Safety equipment such as mats or inverted liferafts may be placed on
the floor or ground to protect participants. No other equipment that is not
part of the rotorcraft's emergency evacuation equipment may be used to aid
the participants in reaching the ground.
(d) Except as provided in paragraph (a) of this appendix, only the
rotorcraft's emergency lighting system may provide illumination.
(e) All emergency equipment required for the planned operation of the
rotorcraft must be installed.
(f) Each external door and exit and each internal door or curtain must be
in the takeoff configuration.
(g) Each crewmember must be seated in the normally assigned seat for
takeoff and must remain in that seat until receiving the signal for
commencement of the demonstration. For compliance with this section, each
crewmember must be--
(1) A member of a regularly scheduled line crew; or
(2) A person having knowledge of the operation of exits and emergency
equipment.
(h) A representative passenger load of persons in normal health must be
used as follows:
(1) At least 25 percent must be over 50 years of age, with at least 40
percent of these being females.
(2) The remaining, 75 percent or less, must be 50 years of age or younger,
with at least 30 percent of these being females.
(3) Three life-size dolls, not included as part of the total passenger
load, must be carried by passengers to simulate live infants 2 years old or
younger, except for a total passenger load of fewer than 44 but more than 19,
one doll must be carried. A doll is not required for a 19 or fewer passenger
load.
(4) Crewmembers, mechanics, and training personnel who maintain or operate
the rotorcraft in the n@mal course of their duties may not be used as
passengers.
(i) No passenger may be assigned a specific seat except as the
Administrator may require. Except as required by paragraph (1) of this
appendix, no employee of the applicant may be seated next to an emergency
exit, except as allowed by the Administrator.
(j) Seat belts and shoulder harnesses (as required) must be fastened.
(k) Before the start of the demonstration, approximately one-half of the
total average amount of carry-on baggage, blankets, pillows, and other
similar articles must be distributed at several locations in the aisles and
emergency exit access ways to create minor obstructions.
(l) No prior indication may be given to any crewmember or passenger of the
particular exits to be used in the demonstration.
(m) The applicant may not practice, rehearse, or describe the demonstration
for the participants nor may any participant have taken part in this type of
demonstration within the preceding 6 months.
(n) A pretakeoff passenger briefing may be given. The passengers may also
be advised to follow directions of crewmembers, but not be instructed on the
procedures to be followed in the demonstration.
(o) If safety equipment, as allowed by paragraph (c) of this appendix, is
provided, either all passenger and cockpit windows must be blacked out or all
emergency exits must have safety equipment to prevent disclosure of the
available emergency exits.
(p) Not more than 50 percent of the emergency exits in the sides of the
fuselage of a rotorcraft that meet all of the requirements applicable to the
required emergency exits for that rotorcraft may be used for demonstration.
Exits that are not to be used for the demonstration must have the exit handle
deactivated or must be indicated by red lights, red tape, or other acceptable
means placed outside the exits to indicate fire or other reasons why they are
unusable. The exits to be used must be representative of all the emergency
exits on the rotorcraft and must be designated by the applicant, subject to
approval by the Administrator. If installed, at least one floor level exit
(Type I; Sec. 29.807(a)(1)) must be used as required by Sec. 29.807(c).
(q) All evacuees must leave the rotorcraft by a means provided as part of
the rotorcraft's equipment.
(r) Approved procedures must be fully utilized during the demonstration.
(s) The evacuation time period is completed when the last occupant has
evacuated the rotorcraft and is on the ground.
SUMMARY: This rule adopts new and revised airworthiness standards for
certification of airframe and related equipment on both normal and transport
category rotorcraft. In addition, one amendment changes an operating rule
affecting external load operators. These amendments grew out of a rotorcraft
regulatory review program and the recognition by both government and industry
that updated safety standards are needed. These amendments provide a high
level of safety in design requirements, while removing certain unnecessary
existing burdens and better utilizing the unique characteristics and
capabilities of rotorcraft.
