PART 27--AIRWORTHINESS STANDARDS: NORMAL CATEGORY ROTORCRAFT
Special Federal Aviation Regulation No. 29-4
Subpart A--General
Sec. 27.1 Applicability.
Sec. 27.2 Special retroactive requirements.
Subpart B--Flight
General
Sec. 27.21 Proof of compliance.
Sec. 27.25 Weight limits.
Sec. 27.27 Center of gravity limits.
Sec. 27.29 Empty weight and corresponding center of gravity.
Sec. 27.31 Removable ballast.
Sec. 27.33 Main rotor speed and pitch limits.
Performance
Sec. 27.45 General.
Sec. 27.51 Takeoff.
Sec. 27.65 Climb: all engines operating.
Sec. 27.67 Climb: one engine inoperative.
Sec. 27.71 Glide performance.
Sec. 27.73 Performance at minimum operating speed.
Sec. 27.75 Landing.
Sec. 27.79 Limiting height--speed envelope.
Flight Characteristics
Sec. 27.141 General.
Sec. 27.143 Controllability and maneuverability.
Sec. 27.151 Flight controls.
Sec. 27.161 Trim control.
Sec. 27.171 Stability: general.
Sec. 27.173 Static longitudinal stability.
Sec. 27.175 Demonstration of static longitudinal stability.
Sec. 27.177 Static directional stability.
Ground and Water Handling Characteristics
Sec. 27.231 General.
Sec. 27.235 Taxiing condition.
Sec. 27.239 Spray characteristics.
Sec. 27.241 Ground resonance.
Miscellaneous Flight Requirements
Sec. 27.251 Vibration.
Subpart C--Strength Requirements
General
Sec. 27.301 Loads.
Sec. 27.303 Factor of safety.
Sec. 27.305 Strength and deformation.
Sec. 27.307 Proof of structure.
Sec. 27.309 Design limitations.
Flight Loads
Sec. 27.321 General.
Sec. 27.337 Limit maneuvering load factor.
Sec. 27.339 Resultant limit maneuvering loads.
Sec. 27.341 Gust loads.
Sec. 27.351 Yawing conditions.
Sec. 27.361 Engine torque.
Control Service and System Loads
Sec. 27.391 General.
Sec. 27.395 Control system.
Sec. 27.397 Limit pilot forces and torques.
Sec. 27.399 Dual control system.
Sec. 27.401 [Removed. Amdt. 27-27, 55 FR 38966, Sept. 21, 1990]
Sec. 27.403 [Removed. Amdt. 27-27, 55 FR 38966, Sept. 21, 1990]
Sec. 27.411 Ground clearance: tail rotor guard.
Sec. 27.413 [Removed. Amdt. 27-27, 55 FR 38966, Sept. 21, 1990]
Sec. 27.427 Unsymmetrical loads.
Ground Loads
Sec. 27.471 General.
Sec. 27.473 Ground loading conditions and assumptions.
Sec. 27.475 Tires and shock absorbers.
Sec. 27.477 Landing gear arrangement.
Sec. 27.479 Level landing conditions.
Sec. 27.481 Tail-down landing conditions.
Sec. 27.483 One-wheel landing conditions.
Sec. 27.485 Lateral drift landing conditions.
Sec. 27.493 Braked roll conditions.
Sec. 27.497 Ground loading conditions: landing gear with tail
wheels.
Sec. 27.501 Ground loading conditions: landing gear with skids.
Sec. 27.505 Ski landing conditions.
Water Loads
Sec. 27.521 Float landing conditions.
Main Component Requirements
Sec. 27.547 Main rotor structure.
Sec. 27.549 Fuselage, landing gear, and rotor pylon structures.
Emergency Landing Conditions
Sec. 27.561 General.
Sec. 27.562 Emergency landing dynamic conditions.
Sec. 27.563 Structural ditching provisions.
Fatigue Evaluation
Sec. 27.571 Fatigue evaluation of flight structure.
Subpart D--Design and Construction
General
Sec. 27.601 Design.
Sec. 27.603 Materials.
Sec. 27.605 Fabrication methods.
Sec. 27.607 Fasteners.
Sec. 27.609 Protection of structure.
Sec. 27.610 Lightning protection.
Sec. 27.611 Inspection provisions.
Sec. 27.613 Material strength properties and design values.
Sec. 27.619 Special factors.
Sec. 27.621 Casting factors.
Sec. 27.623 Bearing factors.
Sec. 27.625 Fitting factors.
Sec. 27.629 Flutter.
Rotors
Sec. 27.653 Pressure venting and drainage of rotor blades.
Sec. 27.659 Mass balance.
Sec. 27.661 Rotor blade clearance.
Sec. 27.663 Ground resonance prevention means.
Control Systems
Sec. 27.671 General.
Sec. 27.672 Stability augmentation, automatic, and power-
operated systems.
Sec. 27.673 Primary flight control.
Sec. 27.674 Interconnected controls.
Sec. 27.675 Stops.
Sec. 27.679 Control system locks.
Sec. 27.681 Limit load static tests.
Sec. 27.683 Operation tests.
Sec. 27.685 Control system details.
Sec. 27.687 Spring devices.
Sec. 27.691 Autorotation control mechanism.
Sec. 27.695 Power boost and power-operated control system.
Landing Gear
Sec. 27.723 Shock absorption tests.
Sec. 27.725 Limit drop test.
Sec. 27.727 Reserve energy absorption drop test.
Sec. 27.729 Retracting mechanism.
Sec. 27.731 Wheels.
Sec. 27.733 Tires.
Sec. 27.735 Brakes.
Sec. 27.737 Skis.
Floats and Hulls
Sec. 27.751 Main float buoyancy.
Sec. 27.753 Main float design.
Sec. 27.755 Hulls.
Personnel and Cargo Accommodations
Sec. 27.771 Pilot compartment.
Sec. 27.773 Pilot compartment view.
Sec. 27.775 Windshields and windows.
Sec. 27.777 Cockpit controls.
Sec. 27.779 Motion and effect of cockpit controls.
Sec. 27.783 Doors.
Sec. 27.785 Seats, berths, safety belts, and harnesses.
Sec. 27.787 Cargo and baggage compartments.
Sec. 27.801 Ditching.
Sec. 27.807 Emergency exits.
Sec. 27.831 Ventilation.
Sec. 27.833 Heaters.
Fire Protection
Sec. 27.853 Compartment interiors.
Sec. 27.855 Cargo and baggage compartments.
Sec. 27.859 Heating systems.
Sec. 27.861 Fire protection of structure, controls, and other
parts.
Sec. 27.863 Flammable fluid fire protection.
External Load Attaching Means
Sec. 27.865 External load attaching means.
Miscellaneous
Sec. 27.871 Leveling marks.
Sec. 27.873 Ballast provisions.
Subpart E--Powerplant
General
Sec. 27.901 Installation.
Sec. 27.903 Engines.
Sec. 27.907 Engine vibration.
Rotor Drive System
Sec. 27.917 Design.
Sec. 27.921 Rotor brake.
Sec. 27.923 Rotor drive system and control mechanism tests.
Sec. 27.927 Additional tests.
Sec. 27.931 Shafting critical speed.
Sec. 27.935 Shafting joints.
Sec. 27.939 Turbine engine operating characteristics.
Fuel System
Sec. 27.951 General.
Sec. 27.953 Fuel system independence.
Sec. 27.954 Fuel system lightning protection.
Sec. 27.955 Fuel flow.
Sec. 27.959 Unusable fuel supply.
Sec. 27.961 Fuel system hot weather operation.
Sec. 27.963 Fuel tanks: general.
Sec. 27.965 Fuel tank tests.
Sec. 27.969 Fuel tank expansion space.
Sec. 27.971 Fuel tank sump.
Sec. 27.973 Fuel tank filler connection.
Sec. 27.975 Fuel tank vents.
Sec. 27.977 Fuel tank outlet.
Fuel System Components
Sec. 27.991 Fuel pumps.
Sec. 27.993 Fuel system lines and fittings.
Sec. 27.995 Fuel valves.
Sec. 27.997 Fuel strainer or filter.
Sec. 27.999 Fuel system drains.
Oil System
Sec. 27.1011 Engines: General.
Sec. 27.1013 Oil tanks.
Sec. 27.1015 Oil tank tests.
Sec. 27.1017 Oil lines and fittings.
Sec. 27.1019 Oil strainer or filter.
Sec. 27.1021 Oil system drains.
Sec. 27.1027 Transmissions and gearboxes: General.
Cooling
Sec. 27.1041 General.
Sec. 27.1043 Cooling tests.
Sec. 27.1045 Cooling test procedures.
Induction System
Sec. 27.1091 Air induction.
Sec. 27.1093 Induction system icing protection.
Exhaust System
Sec. 27.1121 General.
Sec. 27.1123 Exhaust piping.
Powerplant Controls and Accessories
Sec. 27.1141 Powerplant controls: general.
Sec. 27.1143 Engine controls.
Sec. 27.1145 Ignition switches.
Sec. 27.1147 Mixture controls.
Sec. 27.1163 Powerplant accessories.
Powerplant Fire Protection
Sec. 27.1183 Lines, fittings, and components.
Sec. 27.1185 Flammable fluids.
Sec. 27.1187 Ventilation.
Sec. 27.1189 Shutoff means.
Sec. 27.1191 Firewalls.
Sec. 27.1193 Cowling and engine compartment covering.
Sec. 27.1194 Other surfaces.
Sec. 27.1195 Fire detector systems.
Subpart F--Equipment
General
Sec. 27.1301 Function and installation.
Sec. 27.1303 Flight and navigation instruments.
Sec. 27.1305 Powerplant instruments.
Sec. 27.1307 Miscellaneous equipment.
Sec. 27.1309 Equipment, systems, and installations.
Instruments: Installation
Sec. 27.1321 Arrangement and visibility.
Sec. 27.1322 Warning, caution, and advisory lights.
Sec. 27.1323 Airspeed indicating system.
Sec. 27.1325 Static pressure systems.
Sec. 27.1327 Magnetic direction indicator.
Sec. 27.1329 Automatic pilot system.
Sec. 27.1335 Flight director systems.
Sec. 27.1337 Powerplant instruments.
Electrical Systems and Equipment
Sec. 27.1351 General.
Sec. 27.1353 Storage battery design and installation.
Sec. 27.1357 Circuit protective devices.
Sec. 27.1361 Master switch.
Sec. 27.1365 Electric cables.
Sec. 27.1367 Switches.
Lights
Sec. 27.1381 Instrument lights.
Sec. 27.1383 Landing lights.
Sec. 27.1385 Position light system installation.
Sec. 27.1387 Position light system dihedral angles.
Sec. 27.1389 Position light distribution and intensities.
Sec. 27.1391 Minimum intensities in the horizontal plane of
forward and rear position lights.
Sec. 27.1393 Minimum intensities in any vertical plane of
forward and rear position lights.
Sec. 27.1395 Maximum intensities in overlapping beams of
forward and rear position lights.
Sec. 27.1397 Color specifications.
Sec. 27.1399 Riding light.
Sec. 27.1401 Anticollision light system.
Safety Equipment
Sec. 27.1411 General.
Sec. 27.1413 Safety belts.
Sec. 27.1415 Ditching equipment.
Sec. 27.1419 Ice protection.
Sec. 27.1435 Hydraulic systems.
Sec. 27.1457 Cockpit voice recorders.
Sec. 27.1459 Flight recorders.
Sec. 27.1461 Equipment containing high energy rotors.
Subpart G--Operating Limitations and Information
Sec. 27.1501 General.
Operating Limitations
Sec. 27.1503 Airspeed limitations: general.
Sec. 27.1505 Never-exceed speed.
Sec. 27.1509 Rotor speed.
Sec. 27.1519 Weight and center of gravity.
Sec. 27.1521 Powerplant limitations.
Sec. 27.1523 Minimum flight crew.
Sec. 27.1525 Kinds of operations.
Sec. 27.1527 Maximum operating altitude.
Sec. 27.1529 Instructions for Continued Airworthiness.
Markings and Placards
Sec. 27.1541 General.
Sec. 27.1543 Instrument markings: general.
Sec. 27.1545 Airspeed indicator.
Sec. 27.1547 Magnetic direction indicator.
Sec. 27.1549 Powerplant instruments.
Sec. 27.1551 Oil quantity indicator.
Sec. 27.1553 Fuel quantity indicator.
Sec. 27.1555 Control markings.
Sec. 27.1557 Miscellaneous markings and placards.
Sec. 27.1559 Limitations placard.
Sec. 27.1561 Safety equipment.
Sec. 27.1565 Tail rotor.
Rotorcraft Flight Manual and Approved Manual Material
Sec. 27.1581 General.
Sec. 27.1583 Operating limitations.
Sec. 27.1585 Operating procedures.
Sec. 27.1587 Performance information.
Sec. 27.1589 Loading information.
Appendix A Part 27--Instructions for Continued Airworthiness
Appendix B to Part 27--Airworthiness Criteria for Helicopter
Instrument Flight
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. 27.1 Applicability.
(a) This part prescribes airworthiness standards for the issue of type
certificates, and changes to those certificates, for normal category
rotorcraft with maximum weights of 6,000 pounds or less.
(b) Each person who applies under Part 21 for such a certificate or change
must show compliance with the applicable requirements of this part.
Sec. 27.2 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 the foreign civil airworthiness authority certifies that 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. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-21, 49 FR
44432, Nov. 6, 1984]
Sec. 27.25 Weight limits.
(a) Maximum weight. The maximum weight (the highest weight at which
compliance with each applicable requirement of this part is shown) must be
established so that it is--
(1) Not more than--
(i) The highest weight selected by the applicant;
(ii) The design maximum (the highest weight at which compliance with each
applicable structural loading condition of this part is shown); or
(iii) The highest weight at which compliance with each applicable flight
requirement of this part is shown; and
(2) Not less than the sum of--
(i) The empty weight determined under Sec. 27.29; and
(ii) The weight of usable fuel appropriate to the intended operation with
full payload;
(iii) The weight of full oil capacity; and
(iv) For each seat, an occupant weight of 170 pounds or any lower weight
for which certification is requested.
(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--
(1) Not more than the sum of--
(i) The empty weight determined under Sec. 27.29; and
(ii) The weight of the minimum crew necessary to operate the rotorcraft,
assuming for each crewmember a weight no more than 170 pounds, or any lower
weight selected by the applicant or included in the loading instructions; and
(2) Not less than--
(i) The lowest weight selected by the applicant;
(ii) The design minimum weight (the lowest weight at which compliance with
each applicable structural loading condition of this part is shown); or
(iii) 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.
(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)))
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. 27.25. 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. 27-2, 33 FR 962, Jan. 26, 1968]
Sec. 27.29 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 roto-craft 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. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-14, 43 FR
2324, Jan. 16, 1978]
Sec. 27.31 Removable ballast.
Removable ballast may be used in showing compliance with the flight
requirements of this subpart.
Sec. 27.33 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 limits (power on). For rotocraft, 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 limits (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 conditons, 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) Unless otherwise prescribed, the performance requirements of this
subpart must be met for still air and a standard atmosphere.
(b) The performance must correspond to the engine power available under the
particular ambient atmospheric conditions, the particular flight condition,
and the relative humidity specified in paragraphs (d) or (e) of this section,
as appropriate.
(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 appropriate to the
particular ambient atmopheric conditions and the particular flight condition.
(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 degrees F.
Between these two temperatures, the relative humidity must vary linearly.
(f) For turbine-engine-powered rotorcraft, a means must be provided to
permit the pilot to determine 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) of the Dept. of
Transportation Act (49 U.S.C. 1655(c)))
[Amdt. 27-14, 43 FR 2324, Jan. 16, 1978, as amended by Amdt. 27-21, 49 FR
44432, Nov. 6, 1984]
Sec. 27.51 Takeoff.
(a) The takeoff, with takeoff power and r.p.m., and with the extreme
forward center of gravity--
(1) May not require exceptional piloting skill or exceptionally favorable
conditions; and
(2) Must be made in such a manner that a landing can be made safely at any
point along the flight path if an engine fails.
(b) Paragraph (a) of this section must be met throughout the ranges of--
(1) Altitude, from standard sea level conditions to the maximum altitude
capability of the rotorcraft, or 7,000 feet, whichever is less; and
(2) Weight, from the maximum weight (at sea level) to each lesser weight
selected by the applicant for each altitude covered by paragraph (b)(1) of
this section.
Sec. 27.65 Climb: all engines operating.
(a) For rotorcraft other than helicopters--
(1) The steady rate of climb, at VY, must be determined--
(i) With maximum continuous power on each engine;
(ii) With the landing gear retracted; and
(iii) For the weights, altitudes, and temperatures for which certification
is requested; and
(2) The climb gradient, at the rate of climb determined in accordance with
paragraph (a)(1) of this section, must be either--
(i) At least 1:10 if the horizontal distance required to take off and climb
over a 50-foot obstacle is determined for each weight, altitude, and
temperature within the range for which certification is requested; or
(ii) At least 1:6 under standard sea level conditions.
(b) Each helicopter must meet the following requirements:
(1) VY must be determined--
(i) For standard sea level conditions;
(ii) At maximum weight; and
(iii) With maximum continuous power on each engine.
(2) If at any altitude within the range for which certification is
requested, VNE is less than VY the steady rate of climb must be determined--
(i) At the climb speed selected by the applicant at or below VNE;
(ii) 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;
(iii) For the weights and temperatures that correspond to the altitude
range set forth in paragraph (b)(2)(ii) of this section and for which
certification is requested; and
(iv) With maximum continuous power on each engine.
(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. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-14, 43 FR
2324, Jan. 16, 1978]
Sec. 27.67 Climb: one engine inoperative.
For multiengine helicopters, the steady rate of climb (or descent), at Vy
(or at the speed for minimum rate of descent), must be determined with--
(a) Maximum weight;
(b) The critical engine inoperative and the remaining engines at either--
(1) Maximum continuous power and, for helicopters for which certification
for the use of 30-minute OEI power is requested, at 30-minute OEI power; or
(2) Continuous OEI power for helicopters for which certification for the
use of continuous OEI power is requested.
(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 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-23, 53 FR
34210, Sept. 2, 1988]
Sec. 27.71 Glide performance.
For single-engine helicopters and multiengine helicopters that do not meet
the Category A engine isolation requirements of Part 29 of this chapter, the
minimum rate of descent airspeed and the best angle-of-glide airspeed must be
determined in autorotation at--
(a) Maximum weight; and
(b) Rotor speed(s) selected by the applicant.
[Amdt. 27-21, 49 FR 44433, Nov. 6, 1984]
Sec. 27.73 Performance at minimum operating speed.
(a) For helicopters--
(1) The hovering ceiling must be determined over the ranges of weight,
altitude, and temperature for which certification is requested, with--
(i) Takeoff power;
(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 (a)(1) of this section
must be at least--
(i) For reciprocating engine powered helicopters, 4,000 feet at maximum
weight with a standard atmosphere; or
(ii) For turbine engine powered helicopters, 2,500 feet pressure altitude
at maximum weight at a temperature of standard +40 degrees F.
(b) 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.
Sec. 27.75 Landing.
(a) The rotorcraft must be able to be landed with no excessive vertical
acceleration, no tendency to bounce, nose over, ground loop, porpoise, or
water loop, and without exceptional piloting skill or exceptionally favorable
conditions, with--
(1) Approach or glide speeds appropriate to the type of rotorcraft and
selected by the applicant;
(2) The approach and landing made with--
(i) Power off, for single-engine rotorcraft; and
(ii) For multiengine rotocraft, one engine inoperative and with each
operating engine within approved operating limitations; and
(3) The approach and landing entered from steady autorotation.
(b) Multiengine rotorcraft must be able to be landed safely after complete
power failure under normal operating conditions.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-14, 43 FR
2324, Jan. 16, 1978]
Sec. 27.79 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 (including all pertinent information) for that
condition, throughout the ranges of--
(1) Altitude, from standard sea level conditions to the maximum altitude
capability of the rotorcraft, or 7,000 feet, whichever is less; and
(2) Weight, from the maximum weight (at sea level) to the lesser weight
selected by the applicant for each altitude covered by paragraph (a)(1) of
this section. For helicopters, the weight at altitudes above sea level may
not be less than the maximum weight or the highest weight allowing hovering
out of ground effect which is lower.
(b) The applicable power failure conditions are--
(1) For single-engine helicopters, full autorotation;
(2) For multiengine helicopters, one engine inoperative (where engine
isolation features insure continued operation of the remaining engines), and
the remaining engines at the greatest power for which certification is
requested, and
(3) For other rotocraft, conditions appropriate to the type.
(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)))
The rotorcraft must--
(a) Except as specifically required in the applicable section, meet the
flight characteristics requirements of this subpart--
(1) At the altitudes and temperatures expected in operation;
(2) Under any critical loading condition within the range of weights and
centers of gravity for which certification is requested;
(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 of Part 29 of this
chapter;
(2) Sudden, complete power failure for other rotorcraft; and
(3) Sudden, complete control system failures specified in Sec. 27.695 of
this part; and
(c) Have any additional characteristic 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.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-2, 33 FR
962, Jan. 26, 1968; Amdt. 27-11, 41 FR 55468, Dec. 20, 1976; Amdt. 27-19, 48
FR 4389, Jan. 31, 1983; Amdt. 27-21, 49 FR 44433, Nov. 6, 1984]
Sec. 27.143 Controllability and maneuverability.
(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;
(vi) Landing (power on and power off); and
(vii) Recovery to power-on flight from a balked autorotative approach.
(b) The margin of cyclic control must allow satisfactory roll and pitch
control at VNE with--
(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;
(3) Critical rotor r.p.m.; and
(4) Altitude, from standard sea level conditions to the maximum altitude
capability of the rotorcraft or 7,000 feet, whichever is less.
(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 engine 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.
27.1505(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. 27-21, 49 FR 44433, Nov. 6, 1984]
Sec. 27.161 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. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-21, 49 FR
44433, Nov. 6, 1984]
Sec. 27.171 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. 27.173 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. 27.175 (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. 27.175(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. 27-21, 49 FR 44433, Nov. 6, 1984]
Sec. 27.175 Demonstration of static longitudinal stability.
(a) Climb. Static longitudinal stability must be shown in the climb
condition at speeds from 0.85 VY to 1.2 VY, 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 VNH, 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 to VNE, or to 1.1 VNE (power-off) if VNE (power-off) is established
under Sec. 27.1505(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 and direction of motion prescribed in Sec. 27.173 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) of the 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. 27.175 (a)
and (b). This must be shown by steadily increasing 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. 27-21, 49 FR 44433, Nov. 6, 1984]
Ground and Water Handling Characteristics
Sec. 27.231 General.
The rotorcraft must have satisfactory ground and water handling
characteristics, including freedom from uncontrollable tendencies in any
condition expected in operation.
Sec. 27.235 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. 27.239 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. 27.241 Ground resonance.
The rotorcraft may have no dangerous tendency to oscillate on the ground
with the rotor turning.
Miscellaneous Flight Requirements
Sec. 27.251 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. 27.301 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. 27.303 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. 27.305 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. 27.307 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 test 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 weight.
(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. 27.321 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.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 41 FR
55468, Dec. 20, 1976]
Sec. 27.337 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.
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 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);
omega=The angular velocity of rotor (radians per second); and
R=The rotor radius (ft).
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 41 FR
55469, Dec. 20, 1976]
Sec. 27.341 Gust loads.
The rotorcraft must be designed to withstand, at each critical airspeed
including hovering, the loads resulting from a vertical gust of 30 feet per
second.
Sec. 27.351 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 directional control suddenly to the maximum
deflection limited by the control stops or by the pilot force specified in
Sec. 27.395(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 this 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. 27.395(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.
(a) For turbine engines, the limit torque may not be less than the highest
of--
(1) The mean torque for maximum continuous power multiplied by 1.25;
(2) The torque required by Sec. 27.923;
(3) The torque required by Sec. 27.927; or
(4) The torque imposed by sudden engine stoppage due to malfunction or
structural failure (such as compressor jamming).
(b) For reciprocating engines, the limit torque may not be less than 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. 27-23, 53 FR 34210, Sept. 2, 1988]
Control Service and System Loads
Sec. 27.391 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. 27.395, 27.397, 27.399, 27.401, 27.403, 27.411, 27.413, and 27.427.
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 part of each control system from the pilot's controls to the
control stops must be designed to withstand pilot forces of not less than--
(1) The forces specified in Sec. 27.397; or
(2) If the system prevents the pilot from applying the limit pilot forces
to the system, the maximum forces that the system allows the pilot to apply,
but not less than 0.60 times the forces specified in Sec. 27.397.
(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. 27.397.
(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 force 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.
27.397, 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 rational analysis, the design loads
resulting from 0.60 of the specified limit pilot forces are acceptable
minimum design loads.
(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. 27.397, without yielding.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended at Amdt. 27-26, 55
FR 7999, 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 follows 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. 27-11, 41 FR 55469, Dec. 20, 1976]
Sec. 27.399 Dual control system.
Each dual primary flight control system must be designed to withstand the
loads that result when pilot forces of 0.75 times those obtained under Sec.
27.395 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 but 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 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. 27-26, 55 FR 7999, Mar. 6, 1990, as amended by Amdt.
27-27, 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. 27.473 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. 27.725.
[Amdt. 27-2, 33 FR 963, Jan. 26, 1968]
Sec. 27.475 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. 27.477 Landing gear arrangement.
Sections 27.235, 27.479 through 27.485, and 27.493 apply to landing gear
with two wheels aft, and one or more wheels forward, of the center of
gravity.
Sec. 27.479 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 all wheels contact 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. 27.471.
(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) 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.
(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. 27.483 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. 27.479(b)(1); and
(b) The unbalanced external loads must be reacted by rotorcraft inertia.
Sec. 27.485 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. 27.479 (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 all 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. 27.493 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. 27.479(a)(1); and
(2) 1.0 for the attitude specified in Sec. 27.479(a)(2); and
(b) The structure must be designed to withstand at the ground contact point
of each wheel with brakes, a drag load 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. 27.497 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. 27.471 through 27.475;
(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 paragraphs (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 paragraphs (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 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 sideload.
(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 loads that would occur when the rotorcraft is taxied over the roughest
ground that may reasonably be expected in normal operation.
Sec. 27.501 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. 27.471 through 27.475.
(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. 27.725; and
(ii) An assumed rotor lift of not more than 1.5 times that used in the
limit drop tests prescribed in Sec. 27.725.
(4) Complianc@with paragraphs (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 the 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.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-2, 33 FR
963, Jan. 26, 1968; Amdt. 27-26, 55 FR 8000, 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.
EFFECTIVE DATE: April 5, 1990.
The incorporation by reference of certain publications listed in the
regulations is approved by the Director of the Federal Register as of April
5, 1990.
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. 27.473(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.
Water Loads
Sec. 27.521 Float landing conditions.
If certification for float 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. 27.473(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.
Main Component Requirements
Sec. 27.547 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. 27.337 through 27.341:
(1) Critical flight loads.
(2) Limit loads occurring under normal conditions of autorotation. For this
condition, the rotor r.p.m. must be selected to include the effects of
altitude.
(d) The main rotor structure must be designed to withstand loads
simulating--
(1) For the rotor blades, 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; and
(ii) The limit engine torque specified in Sec. 27.361.
(2) The limit torque must be distributed to the rotor blades in a rational
manner.
Sec. 27.549 Fuselage, landing gear, and rotor pylon structures.
(a) Each fuselage, landing gear, and rotor pylon structure must be designed
as prescribed in this section. Resultant rotor forces may be represented as a
single force applied at the rotor hub attachment point.
(b) Each structure must be designed to withstand--
(1) The critical loads prescribed in Secs. 27.337 through 27.341;
(2) The applicable ground loads prescribed in Secs. 27.235, 27.471 through
27.485, 27.493, 27.497, 27.501, 27.505, and 27.521; and
(3) The loads prescribed in Sec. 27.547 (d) (2) and (e).
(c) Auxiliary rotor thrust, and the balancing air and inertia loads
occurring under accelerated flight conditions, must be considered.
(d) Each engine mount and adjacent fuselage structure must be designed to
withstand the loads occurring under accelerated flight and landing
conditions, including engine torque.
(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--
(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 intended displacement of the seat device.
(c) The supporting structure must be designed to restrain, under any
ultimate inertial load 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, transmissions, 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 when
those loads are applied to that area:
(i) Upward--1.5g.
(ii) Forward--4.0g.
(iii) Sideward--2.0g.
(iv) Downward--4.0g.
SUMMARY: These amendments add comprehensive crash resistant fuel system
design and test criteria to the airworthiness standards for normal and
transport category rotorcraft. cation 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.
EFFECTIVE DATE: November 2, 1994.
*****************************************************************************
Sec. 27.562 Emergency landing dynamic conditions.
(a) The rotorcraft, although it may be damaged in an emergency 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 the loads 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.
(1) 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 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 upper torso 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.
[54 FR 47318, Nov. 13, 1989]
Sec. 27.563 Structural ditching provisions.
If certification with ditching provisions is requested, structural strength
for ditching must meet the requirements of this section and Sec. 27.801(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, of 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. 27.801(d).
Maximum roll and pitch angles determined from compliance with Sec. 27.801(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 perscribed 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.
Sec. 27.571 Fatigue evaluation of flight structure.
(a) General. Each portion of the flight structure (the flight structure
includes rotors, rotor drive systems between the engines and the rotor hubs,
controls, fuselage, landing gear, and their related primary attachments), the
failure of which could be catastrophic, must be identified and must be
evaluated under paragraph (b), (c), (d), or (e) of this section. The
following apply to each fatigue evaluation:
(1) The procedure for the evaluation must be approved.
(2) The locations of probable failure must be determined.
(3) Inflight measurement must be included in determining the following:
(i) Loads or stresses in all critical conditions throughout the range of
limitations in Sec. 27.309, except that maneuvering load factors need not
exceed the maximum values expected in operation.
(ii) The effect of altitude upon these loads or stresses.
(4) The loading spectra must be as severe as those expected in operation
including, but not limited to, external cargo operations, if applicable, and
ground-air-ground cycles. The loading spectra must be based on loads or
stresses determined under paragraph (a)(3) of this section.
(b) Fatigue tolerance evaluation. It must be shown that the fatigue
tolerance of the structure ensures that the probability of catastrophic
fatigue failure is extremely remote without establishing replacement times,
inspection intervals or other procedures under section A27.4 of Appendix A.
(c) Replacement time evaluation. it must be shown that the probability of
catastrophic fatigue failure is extremely remote within a replacement time
furnished under section A27.4 of Appendix A.
(d) Fail-safe evaluation. The following apply to fail-safe evaluation:
(1) It must be shown that all partial failures will become readily
detectable under inspection procedures furnished under section A27.4 of
Appendix A.
(2) The interval between the time when any partial failure becomes readily
detectable under paragraph (d)(1) of this section, and the time when any such
failure is expected to reduce the remaining strength of the structure to
limit or maximum attainable loads (whichever is less), must be determined.
(3) It must be shown that the interval determined under paragraph (d)(2) of
this section is long enough, in relation to the inspection intervals and
related procedures furnished under section A27.4 of Appendix A, to provide a
probability of detection great enough to ensure that the probability of
catastrophic failure is extremely remote.
(e) Combination of replacement time and failsafe evaluations. A component
may be evaluated under a combination of paragraphs (c) and (d) of this
section. For such component it must be shown that the probability of
catastrophic failure is extremely remote with an approved combination of
replacement time, inspection intervals, and related procedures furnished
under section A27.4 of Appendix A.
(Secs. 313(a), 601, 603, 604, and 605, 72 Stat. 752, 775, and 778, (49 U.S.C.
1354(a), 1421, 1423, 1424, and 1425; sec. 6(c), 49 U.S.C. 1655(c)))
[Amdt. 27-3, 33 FR 14106, Sept. 18, 1968, as amended by Amdt. 27-12, 42 FR
15044, Mar. 17, 1977; Amdt. 27-18, 45 FR 60177, Sept. 11 1980; Amdt. 27-26,
55 FR 8000, 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 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. 27.603 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, Federal Aviation Act of 1958 (49 U.S.C.
1354(a), 1421, 1423, 1424); and sec. 6(c) of the 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, and 605 of the Federal Aviation Act of 1958 (49
U.S.C. 1354(a), 1421, 1423, 1424 and 1425); sec. 6(c) of the Dept. of
Transportation Act (49 U.S.C. 1655(c)))
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-16, 43 FR
50599, Oct. 30, 1978]
Sec. 27.607 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. 27-4, 33 FR 14533, Sept. 27, 1968]
Sec. 27.609 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. 27.610 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 minimize the effect of a strike; or
(2) Incorporating acceptable means of diverting the resulting electrical
current so as not to endanger the rotorcraft.
[Amdt. 27-21, 49 FR 44433, Nov. 6, 1984]
Sec. 27.611 Inspection provisions.
There must be means to allow the close examination of each part that
requires--
(a) Recurring inspection;
(b) Adjustment for proper alignment and functioning; or
(c) Lubrication.
Sec. 27.613 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 structure, 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, Pennsylvania 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. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-16, 43 FR
50599, Oct. 30, 1978; Amdt. 27-26, 55 FR 8000, 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. 27.621 through 27.625 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 to which Secs. 27.621 through 27.625 apply, the factor of
safety prescribed in Sec. 27.303 must be multiplied by a special factor equal
to--
(1) The applicable special factors prescribed in Secs. 27.621 through
27.625; 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. 27.621 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. 27.305 at an ultimate load
corresponding to a casting factor of 1.25; and
(ii) The deformation requirements of Sec. 27.305 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. 27.623 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. 27.625 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. 27.629 Flutter.
Each aeordynamic surface of the rotorcraft must be free from flutter
under each appropriate speed and power condition.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended at Amdt. 27-26, 55
FR 8000, 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. 27.653 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. 27-2, 33 FR 963, Jan. 26, 1968]
Sec. 27.659 Mass balance.
(a) The rotors 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. 27-2, 33 FR 963, Jan. 26, 1968]
Sec. 27.661 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. 27-2, 33 FR 963, Jan. 26, 1968]
Sec. 27.663 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. 27.241.
[Amdt. 27-2, 33 FR 963, Jan. 26, 1968, as amended at Amdt. 27-26, 55 FR 8000,
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 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.
Sec. 27.672 Stability augmentation, automatic, and power-operated systems.
If the functioning of stability augmentation or other automatic or power-
operated systems is necessary to show compliance with the flight
characteristics requirements of this part, such systems must comply with Sec.
27.671 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 movement of the flight controls in the normal sense and
deactivating the failed system.
(c) It must be shown that after any single failure of the stability
augmentation system or any other automatic or power-operated system--
(1) The 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 permit continued safe flight and landing.
Primary flight controls are those used by the pilot for immediate control
of pitch, roll, yaw, and vertical motion of the rotorcraft.
[Amdt. 27-21, 49 FR 44434, Nov. 6, 1984]
Sec. 27.674 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
motion of 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. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-16, 43 FR
50599, Oct. 30, 1978]
Sec. 27.679 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) Give unmistakable warning to the pilot when the lock is engaged; and
(b) Prevent the lock from engaging in flight.
Sec. 27.681 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. 27.683 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. 27.685 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 the 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 three thirty-seconds of an 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 which must be
used are specified in Military Handbook MIL-HDBK-5C, Vol. 1 & Vol. 2,
Metallic Materials and Elements for Flight Vehicle Structures, (Sept. 15,
1976, as amended through December 15, 1978). This incorporation by reference
was approved by the Director of the Federal Register in accordance with 5
U.S.C. section 552(a) and 1 CFR part 51. Copies may be obtained from the
Naval Publications and Forms Center, 5801 Tabor Avenue, Philadelphia,
Pennsylvania, 19120. Copies may be inspected at the FAA, Rotorcraft Standards
Staff, 4400 Blue Mount Road, Fort Worth, Texas, or at the Office of the
Federal Register, 1100 L Street NW., Room 8301, Washington, DC.
(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 3 deg..
(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 must not be exceeded.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 41 FR
55469, Dec. 20, 1976; Amdt. 27-26, 55 FR 8001, 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 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. 27.691 Autorotation control mechanism.
Each main rotor blade pitch control mechanism must allow rapid entry into
autorotation after power failure.
Sec. 27.695 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 hydraulic 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. 27.723 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.
27.725 and 27.727, 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. 27.725 Limit drop test.
The limit drop test must be conducted as follows:
(a) The drop height must be--
(1) 13 inches from the lowest point of the landing gear to the ground; or
(2) Any lesser height, not less than eight inches, resulting in a drop
contact velocity equal to the greatest probable sinking speed likely to occur
at ground contact in normal power-off landings.
(b) If considered, the rotor lift specified in Sec. 27.473(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 formula 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 lg downward with the rotorcraft in the maximum
nose-up attitude considered in the nose-up landing conditions.
h =specified free drop height (inches).
L =ration 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 travels (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).
Sec. 27.727 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. 27.725(a).
(b) Rotor lift, where considered in a manner similar to that prescribed in
Sec. 27.725(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 the landing gear and external accessories, to impact
the landing surface.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended at Amdt. 27-26, 55 FR
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 mechansim, 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 a 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. 27.777 and 27.779.
(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.
[Amdt. 27-21, 49 FR 44434, Nov. 6, 1984]
Sec. 27.731 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. 27.733 Tires.
(a) Each landing gear wheel must have a tire--
(1) That is a proper fit on the rim of the wheel; and
(2) Of the proper rating.
(b) The maximum static load rating of each tire must equal or exceed the
static ground reaction obtained at its wheel, assuming--
(1) The design maximum weight; and
(2) The most unfavorable center of gravity.
(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. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 41 FR
55469, Dec. 20, 1976]
Sec. 27.735 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. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-21, 49 FR
44434, Nov. 6, 1984]
Sec. 27.737 Skis.
The maximum limit load rating of each ski must equal or exceed the maximum
limit load determined under the applicable ground load requirements of this
part.
Floats and Hulls
Sec. 27.751 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 main floats will provide a
margin of positive stability great enough to minimize the probability of
capsizing.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-2, 33 FR
963, Jan. 26, 1968]
Sec. 27.753 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. 27.521(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. 27.521. These loads may be
distributed along the length of the float.
Sec. 27.755 Hulls.
For each rotorcraft, with a hull and auxiliary floats, that is to be
approved for both taking off from and landing on water, the hull and
auxiliary floats must have enough watertight compartments so that, with any
single compartment flooded, the buoyancy of the hull and auxiliary floats
(and wheel tires if used) provides a margin of positive stability great
enough to minimize the probability of capsizing.
Personnel and Cargo Accommodations
Sec. 27.771 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 seat; and
(c) The vibration and noise characteristics of cockpit appurtenances may
not interfere with safe operation.
Sec. 27.773 Pilot compartment view.
(a) Each pilot compartment must be free from glare and reflections that
could interfere with the pilot's view, and designed so that--
(1) Each pilot's view is sufficiently extensive, clear, and undistorted for
safe operation; and
(2) Each pilot is protected from the elements so that moderate rain
conditions do not unduly impair his view of the flight path in normal flight
and while landing.
(b) If certification for night operation is requested, compliance with
paragraph (a) of this section must be shown in night flight tests.
Sec. 27.775 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 pilots' 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. 27.779 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. 27-21, 49 FR 44434, Nov. 6, 1984]
Sec. 27.783 Doors.
(a) Each closed cabin must have at least one adequate and easily accessible
external door.
(b) Each external door must be located where persons using it will not be
endangered by the rotors, propellers, engine intakes, and exhausts when
appropriate operating procedures are used. If opening procedures are
required, they must be marked inside, on or adjacent to the door opening
device.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended at Amdt. 27-26, 55 FR
8001, 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. 27.785 Seats, berths, safety belts, and harnesses.
(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 static inertial load factors specified in Sec. 27.561(b) and dynamic
conditions specified in Sec. 27.562.
(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. 27.562(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 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 enable the occupants to steady themselves while
using the aisle in moderately rough air.
(e) Each projecting object that could 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 occupant, seat, and safety belt or harness
corresponding with the applicable flight and ground load conditions,
including the emergency landing conditions of Sec. 27.561(b). In addition--
(1) Each pilot seat must be designed for the reactions resulting from the
application of the pilot forces prescribed in Sec. 27.397; and
(2) The inertial forces prescribed in Sec. 27.561(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. 27.561(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. 27.561, 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 seats to reduce occupant loads for the
emergency landing dynamic conditions of Sec. 27.562; otherwise, the system
must remain intact and must not interfere with rapid evacuation of the
rotorcraft.
(k) For the 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.
27.561(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 diaphram, 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 load
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. 27.561(b).
[Amdt. 27-21, 49 FR 44434, Nov. 6, 1984, as amended by Amdt. 27-25, 54 FR
47319, Nov. 13, 1989]
Sec. 27.787 Cargo and baggage compartments.
(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.
27.561.
(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. 27.561, 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.
27.561 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. 27.807(d), 27.1411 and
27.1415.
(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 factor 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 life rafts required by Sec. 27.1415. If
compliance with this provision is shown by buoyancy 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 buoyancy 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. 27-11, 41 FR 55469, Dec. 20, 1976]
Sec. 27.807 Emergency exits.
(a) Number and location. Rotorcraft with closed cabins must have at least
one emergency exit on the opposite side of the cabin from the main door.
(b) Type and operation. Each emergency exit prescribed in paragraph (a) of
this section must--
(1) Consist of a movable window or panel, or additional external door,
providing an unobstructed opening that will admit a 19-by 26-inch ellipse;
(2) Be readily accessible, require no exceptional agility of a person using
it, and be located so as to allow ready use, without crowding, in any
probable attitudes that may result from a crash;
(3) Have a simple and obvious method of opening and be arranged and marked
so as to be readily located and operated, even in darkness; and
(4) Be reasonably protected from jamming by fuselage deformation.
(c) Tests. The proper functioning of each emergency exit must be shown by
test.
(d) Ditching emergency exits for passengers. If certification with ditching
provisions is requested, one emergency exit on each side of the fuselage must
be proven by test, demonstration, or analysis to--
(1) Be above the waterline;
(2) Have at least the dimensions specified in paragraph (b) of this
section; and
(3) Open without interference from flotation devices whether stowed or
deployed.
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 ventilating system for the pilot and passenger compartments must be
designed to prevent the presence of excessive quantities of fuel fumes and
carbon monoxide.
(b) The concentration of carbon monoxide may not exceed one part in 20,000
parts of air during forward flight or hovering in still air. If the
concentration exceeds this value under other conditions, there must be
suitable operating restrictions.
Sec. 27.833 Heaters.
Each combustion heater must be approved.
[Amdt. 27-23, 53 FR 34210, Sept. 2, 1988]
Fire Protection
Sec. 27.853 Compartment interiors.
For each compartment to be used by the crew or passengers--
(a) The materials must be at least flash-resistant;
(b) The wall and ceiling linings, and the covering of upholstery, floors,
and furnishings must be at least flame resistant; and
(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.
[Amdt. 27-17, 45 FR 7755, Feb. 4, 1980]
Sec. 27.855 Cargo and baggage compartments.
(a) Each cargo and baggage compartment must be constructed of, or lined
with, materials that are at least--
(1) Flame resistant, in the case of compartments that are readily
accessible to a crewmember in flight; and
(2) Fire resistant, in the case of other compartments.
(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.
Sec. 27.859 Heating systems.
(a) General. For each heating system that involves the passage of cabin air
over, or close to, the exhaust manifold, there must be means to prevent
carbon monoxide from entering any cabin or pilot compartment.
(b) Heat exchangers. Each heat exchanger must be--
(1) Of suitable materials;
(2) Adequately cooled under all conditions; and
(3) Easily disassembled for inspection.
(c) Combustion heater fire protection. Except for heaters which incorporate
designs to prevent hazards in the event of fuel leakage in the heater fuel
system, fire within the ventilating air passage, or any other heater
malfunction, each heater zone must incorporate the fire protection features
of the applicable requirements of Secs. 27.1183, 27.1185, 27.1189, 27.1191,
and be provided with--
(1) Approved, quick-acting fire detectors in numbers and locations ensuring
prompt detection of fire in the heater region.
(2) Fire extinguisher systems that provide at least one adequate discharge
to all areas of the heater region.
(3) Complete drainage of each part of each zone to minimize the hazards
resulting from failure or malfunction of any component containing flammable
fluids. The drainage means must be--
(i) Effective under conditions expected to prevail when drainage is needed;
and
(ii) Arranged so that no discharged fluid will cause an additional fire
hazard.
(4) Ventilation, arranged so that no discharged vapors will cause an
additional fire hazard.
(d) Ventilating air ducts. Each ventilating air duct passing through any
heater region must be fireproof.
(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.
(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.
(e) Combustion air ducts. Each combustion air duct must be fireproof for a
distance great enough to prevent damage from backfiring or reverse flame
propagation.
(1) No combustion air duct may connect 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.
(2) No combustion air duct may restrict the prompt relief of any backfire
that, if so restricted, could cause heater failure.
(f) Heater control: 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.
(g) 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 (g)(1) of this section for any
individual heater must--
(i) Be independent of components serving any other heater, the heat output
of which 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, the heat output
of which is essential for safe operation, has been shut off by the automatic
means prescribed in paragraph (g)(1) of this section.
(h) Air intakes. Each combustion and ventilating air intake must be located
so that no flammable fluids or vapors can enter the heater system--
(1) During normal operation; or
(2) As a result of the malfunction of any other component.
(i) Heater exhaust. Each heater exhaust system must meet the requirements
of Secs. 27.1121 and 27.1123.
(1) Each exhaust shroud must be sealed so that no flammable fluids or
hazardous quantities of vapors can reach the exhaust system through joints.
(2) No exhaust system may restrict the prompt relief of any backfire that,
if so restricted, could cause heater failure.
(j) 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.
(k) Drains. There must be means for safe drainage of any fuel that might
accumulate in the combustion chamber or the heat exchanger.
(1) Each part of any drain that operates at high temperatures must be
protected in the same manner as heater exhausts.
(2) Each drain must be protected against hazardous ice accumulation under
any operating condition.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-23, 53 FR
34211, Sept. 2, 1988]
Sec. 27.861 Fire protection of structure, controls, and other parts.
Each part of the structure, controls, rotor mechanism, and other parts
essential to a controlled landing that would be affected by powerplant fires
must be fireproof or protected so they can perform their essential functions
for at least 5 minutes under any foreseeable powerplant fire conditions.
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. 27-16, 43 FR 50599, Oct. 30, 1978]
External Load Attaching Means
Sec. 27.865 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. 27.337 through 27.341, 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 crewmember, 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. 27.25 and this section.
(d) The fatigue evaluation of Sec. 27.571(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. 27-11, 41 FR 55469, Dec. 20, 1976, as amended at Amdt. 27-26, 55 FR
8001, 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. 27.873 Ballast provisions.
Ballast provisions must be designed and constructed to prevent inadvertent
shifting of ballast in flight.
Subpart E--Powerplant
General
Sec. 27.901 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) 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;
(2) Accessibility must be provided to allow any inspection and maintenance
necessary for continued airworthiness;
(3) Electrical interconnections must be provided to prevent differences of
potential between major components of the installation and the rest of the
rotorcraft;
(4) Axial and radial expansion of turbine engines may not affect the safety
of the installation; and
(5) 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) The installation must comply with--
(1) The installation instructions provided under Sec. 33.5 of this chapter;
and
(2) The applicable provisions of this subpart.
(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))
(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) Engine or drive system cooling fan blade protection. (1) If an engine
or rotor drive system cooling fan is installed, there must be a means to
protect the rotorcraft and allow a safe landing if a fan blade fails. This
must be shown by showing that--
(i) The fan blades are contained in case of failure;
(ii) Each fan is located so that a failure will not jeopardize safety; or
(iii) Each fan blade can withstand an ultimate load of 1.5 times the
centrifugal force resulting from operation limited by the following:
(A) For fans driven directly by the engine--
(1) The terminal engine r.p.m. under uncontrolled conditions; or
(2) An overspeed limiting device.
(B) For fans driven by the rotor drive system, the maximum rotor drive
system rotational speed to be expected in service, including transients.
(2) Unless a fatigue evaluation under Sec. 27.571 is conducted, it must be
shown that cooling fan blades are not operating at resonant conditions within
the operating limits of the rotorcraft.
(c) 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.
(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.
(c) No part of the rotor drive system may be subjected to excessive
vibration stresses.
Rotor Drive System
Sec. 27.917 Design.
(a) 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.
(b) Each rotor drive system must be arranged so 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.
(c) 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.
(d) 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.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 41 FR
55469, Dec. 20, 1976]
Sec. 27.923 Rotor drive system and control mechanism tests.
(a) 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.
(b) Each rotor drive system and control mechanism must be tested for not
less than 100 hours. The test 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.
(c) A 60-hour part of the test prescribed in paragraph (b) of this section
must be run at not less than maximum continuous torque and the maximum speed
for use with maximum continuous torque. In this test, the main rotor controls
must be set in the position that will give maximum longitudinal cyclic pitch
change to simulate forward flight. The auxiliary rotor controls must be in
the position for normal operation under the conditions of the test.
(d) A 30-hour or, for rotorcraft for which the use of either 30-minute OEI
power or continuous OEI power is requested, a 25-hour part of the test
prescribed in paragraph (b) of this section must be run at not less than 75
percent of maximum continuous torque and the minimum speed for use with 75
percent of maximum continuous torque. The main and auxiliary rotor controls
must be in the position for normal operation under the conditions of the
test.
(e) A 10-hour part of the test prescribed in paragraph (b) of this section
must be run at not less than takeoff torque and the maximum speed for use
with takeoff torque. The main and auxiliary rotor controls must be in the
normal position for vertical ascent.
(1) For multiengine rotorcraft for which the use of 2 1/2 minute OEI power
is requested, 12 runs during the 10-hour test must be conducted as follows:
(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 2 1/2 minute OEI torque and the maximum speed
for use with 2 1/2 minute OEI torque for 2 1/2 minutes.
(2) For multiengine turbine-powered rotorcraft for which the use of 30-
second and 2-minute OEI power is requested, 10 runs must be conducted as
follows:
(i) Immediately following a takeoff run of at least 5 minutes, 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 removal 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.
(f) The parts of the test prescribed in paragraphs (c) and (d) of this
section must be conducted in intervals of not less than 30 minutes and may be
accomplished either on the ground or in flight. The part of the test
prescribed in paragraph (e) of this section must be conducted in intervals of
not less than five minutes.
(g) At intervals of not more than five hours during the tests prescribed in
paragraphs (c), (d), and (e) of this section, the engine must be stopped
rapidly enough to allow the engine and rotor drive to be automatically
disengaged from the rotors.
(h) Under the operating conditions specified in paragraph (c) of this
section, 500 complete cycles of lateral control, 500 complete cycles of
longitudinal control of the main rotors, and 500 complete cycles of control
of each auxiliary rotor must be accomplished. A "complete cycle" involves
movement of the controls from the neutral position, through both extreme
positions, and back to the neutral position, except that control movements
need not produce loads or flapping motions exceeding the maximum loads or
motions encountered in flight. The cycling may be accomplished during the
testing prescribed in paragraph (c) of this section.
(i) At least 200 start-up clutch engagements must be accomplished--
(1) So that the shaft on the driven side of the clutch is accelerated; and
(2) Using a speed and method selected by the applicant.
(j) For multiengine rotorcraft for which the use of 30-minute OEI power is
requested, five runs must be made at 30-minute OEI torque and the maximum
speed for use with 30-minute OEI torque, in which each engine, in sequence,
is made inoperative and the remaining engine(s) is run for a 30-minute
period.
(k) For multiengine rotorcraft for which the use of continuous OEI power is
requested, five runs must be made at continuous OEI torque and the maximum
speed for use with continuous OEI torque, in which each engine, in sequence,
is made inoperative and the remaining engine(s) is run for a 1-hour period.
(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 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 rating 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, or torque that is at least equal to the
lesser of--
(i) The maximum torque used in meeting Sec. 27.923 plus 10 percent; or
(ii) The maximum attainable torque output of the engines, 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 15
minutes.
(3) The tests prescribed in this paragraph must be conducted on the
rotorcraft at the maximum rotational speed intended for the power condition
of the test 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.
(c) It must be shown by tests that the rotor drive system is capable of
operating under autorotative conditions for 15 minutes after the loss of
pressure in the rotor drive primary oil system.
(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. 27-2, 33 FR 963, Jan. 26, 1968, as amended by Amdt. 27-12, 42 FR
15045, Mar. 17, 1977; Amdt. 27-23, 53 FR 34212, Sept. 2, 1988]
Sec. 27.931 Shafting critical speed.
(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.
Sec. 27.935 Shafting joints.
Each universal joint, slip joint, and other shafting joints whose
lubrication is necessary for operation must have provision for lubrication.
Sec. 27.939 Turbine engine operating characteristics.
(a) Turbine engine operating characteristics must be investigated in flight
to determine that no adverse characteristics (such as stall, surge, or
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. 27-1, 32 FR 6914, May 5, 1967, as amended by Amdt. 27-11, 41 FR 55469,
Dec. 20, 1976]
Fuel System
Sec. 27.951 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 functioning under
any likely operating condition, including the maneuvers for which
certification is requested.
(b) Each fuel system must be arranged so that--
(1) No fuel pump can draw fuel from more than one tank at a time; or
(2) There are means to prevent introducing air into the system.
(c) Each fuel system for a turbine engine must be capable of sustained
operation throughout its flow and pressure range with fuel initially
saturated with water at 80 deg. F. and having 0.75cc of free water per gallon
added and cooled to the most critical condition for icing likely to be
encountered in operation.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-9, 39 FR
35461, Oct. 1, 1974]
Sec. 27.952 Fuel system crash resistance.
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 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 tank 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 that
inadvertent fuel shutoff in flight is improbable in accordance with Sec.
27.955(a) and must comply with the fatigue evaluation requirements of Sec.
27.571 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. 27.571 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 components 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. 27.571.
(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) Each fuel system for multiengine rotorcraft must allow fuel to be
supplied to each engine through a system independent of those parts of each
system supplying fuel to other engines. However, separate fuel tanks need not
be provided for each engine.
(b) If a single fuel tank is used on a multiengine rotorcraft, the
following must be provided:
(1) Independent tank outlets for each engine, each incorporating a shutoff
valve at the tank. This shutoff valve may also serve as the firewall shutoff
valve required by Sec. 27.995 if the line between the valve and the engine
compartment does not contain a hazardous amount of fuel that can drain into
the engine compartment.
(2) At least two vents arranged to minimize the probability of both vents
becoming obstructed simultaneously.
(3) Filler caps designed to minimize the probability of incorrect
installation or inflight loss.
(4) A fuel system in which those parts of the system from each tank outlet
to any engine are independent of each part of each system supplying fuel to
other engines.
Sec. 27.954 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; or
(c) Corona and streamering at fuel vent outlets.
[Amdt. 27-23, 53 FR 34212, Sept. 2, 1988]
Sec. 27.955 Fuel flow.
(a) General. The fuel system for each engine must be shown to provide the
engine with at least 100 percent of the fuel required under each operating
and maneuvering condition to be approved for the rotorcraft including, as
applicable, the fuel required to operate the engine(s) under the test
conditions required by Sec. 27.927. 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 critical accelerations, 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. 27.959, plus the minimum
additional fuel necessary to conduct the test.
(3) The fuel head between the tank outlet and the engine inlet must be
critical with respect to rotorcraft flight attitudes.
(4) The critical fuel pump (for pump-fed systems) is installed to produce
(by actual or simulated failure) the critical restriction to fuel flow to be
expected from pump failure.
(5) Critical values of engine rotation 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
must be applied.
(7) The fuel filter required by Sec. 27.997 must be blocked to the degree
necessary to simulate the accumulation of fuel contamination required to
activate the indicator required by Sec. 27.1305(q).
(b) Fuel transfer systems. If normal operation of the fuel system requires
fuel to be transferred to an engine feed tank, the transfer must occur
automatically via a system which has been shown to maintain the fuel level in
the engine feed 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 systems must, in addition to having appropriate
manual switching capability, 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 the engine alone contains
usable fuel.
[Amdt. 27-23, 53 FR 34212, Sept. 2, 1988]
Sec. 27.959 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. 27.961 Fuel system hot weather operation.
Each suction lift fuel system and other fuel systems with features
conducive to vapor formation must be shown by test to operate satisfactorily
(within certification limits) when using fuel at a temperature of 110 deg.F
under critical operating conditions including, if applicable, the engine
operating conditions defined by Sec. 27.927 (b)(1) and (b)(2).
[Amdt. 27-23, 53 FR 34212, Sept. 2, 1988]
Sec. 27.963 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 fuel tank of 10 gallons or greater capacity must have internal
baffles, or must have external support to resist surging.
(c) Each fuel tank must be separated from the engine compartment by a
firewall. At least one-half inch of clear airspace must be provided between
the tank and the firewall.
(d) Spaces adjacent to the surfaces of fuel tanks must be ventilated so
that fumes cannot accumulate in the tank compartment in case of leakage. If
two or more tanks have interconnected outlets, they must be considered as one
tank, and the airspaces in those tanks must be interconnected to prevent the
flow of fuel from one tank to another as a result of a difference in pressure
between those airspaces.
(e) The maximum exposed surface temperature of any component in the fuel
tank must be less, by a safe margin as determined by the Administrator, 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 failure or malfunction conditions of all components
inside the tank.
(f) 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. 27.952, and must be
adequate to withstand loads and abrasions to be expected in personnel
compartments.
(g) 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.
(h) Each integral fuel tank must have provisions for inspection and repair
of its interior.
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, nonmetallic tank with walls that are not
supported by the rotorcraft structure, and 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 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 support 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 calculation 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 paragraphs (d)(3)(ii) and (iii) of this section, 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))
[Amdt. 27-12, 42 FR 15045, Mar. 17, 1977]
Sec. 27.967 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 it is
not necessary for them 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 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. 27.1185
(a) and (c).
(d) No rotorcraft skin immediately adjacent to a major air outlet from the
engine compartment may act as the wall of the 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 tank capacity.
It must be impossible to fill the fuel tank expansion space inadvertently
with the rotorcraft in the normal ground attitude.
[Amdt. 27-23, 53 FR 34213, Sept. 2, 1988]
Sec. 27.971 Fuel tank sump.
(a) Each fuel tank must have a drainable sump with an effective capacity in
any ground attitude to be expected in service of 0.25 percent of the tank
capacity or 1/16 gallon, whichever is greater, unless--
(1) The fuel system has a sediment bowl or chamber that is accessible for
preflight drainage and has a minimum capacity of 1 ounce for every 20 gallons
of fuel tank capacity; and
(2) Each fuel tank drain is located so that in any ground attitude to be
expected in service, water will drain from all parts of the tank to the
sediment bowl or chamber.
(b) Each sump, sediment bowl, and sediment chamber drain required by this
section must comply with the drain provisions of Sec. 27.999(b).
[Amdt. 27-23, 53 FR 34213, Sept. 2, 1988]
Sec. 27.973 Fuel tank filler connection.
(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. 27.952(c). In addition--
(1) Each filler must be marked as prescribed in Sec. 27.1557(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.
(a) Each fuel tank must be vented from the top part of the expansion space
so that venting is effective under all normal flight conditions. Each vent
must minimize the probability of stoppage by dirt or ice.
(b) 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 operation, or a survivable impact, unless a rollover is shown
to be extremely remote.
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 stainer for the fuel tank outlet or for the
booster pump. This strainer must--
(1) For reciprocating engine powered rotorcraft, have 8 to 16 meshes per
inch; and
(2) For turbine engine powered rotorcraft, 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. 27-11, 41 FR 55470, Dec. 20, 1976]
Fuel System Components
Sec. 27.991 Fuel pumps.
Compliance with Sec. 27.955 may not be jeopardized by failure of--
(a) Any one pump except pumps that are approved and installed as parts of a
type certificated engine; or
(b) Any component required for pump operation except, for engine driven
pumps, the engine served by that pump.
[Amdt. 27-23, 53 FR 34213, Sept. 2, 1988]
Sec. 27.993 Fuel system lines and fittings.
(a) Each fuel line must be installed and supported to prevent excessive
vibration and to withstand loads due to fuel pressure and accelerated flight
conditions.
(b) Each fuel line connected to components of the rotorcraft between which
relative motion could exist must have provisions for flexibility.
(c) Flexible hose must be approved.
(d) Each flexible connection in fuel lines that may be under pressure or
subjected to axial loading must use flexible hose assemblies.
(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.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-2, 33 FR
964, Jan. 26, 1968]
Sec. 27.995 Fuel valves.
(a) There must be a positive, quick-acting valve to shut off fuel to each
engine individually.
(b) The control for this valve must be within easy reach of appropriate
crewmembers.
(c) Where there is more than one source of fuel supply there must be means
for independent feeding from each source.
(d) No shutoff valve may be on the engine side of any firewall.
Sec. 27.997 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 itself,
unless adequate strength 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 fuel system 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 must--
(1) Discharge clear of all parts of the rotorcraft;
(2) Have manual or automatic means to assure 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.
(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 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.
(c) The oil cooling provisions for each engine must be able to maintain the
oil inlet temperature to that engine at or below the maximum established
value. This must be shown by flight tests.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-23, 53 FR
34213, Sept. 2, 1988]
Sec. 27.1013 Oil tanks.
Each oil tank must be designed and installed so that--
(a) It can withstand, without failure, each vibration, inertia, fluid, and
structural load expected in operation;
(b) [Reserved]
(c) Where used with a reciprocating engine, it has an expansion space of
not less than the greater of 10 percent of the tank capacity or 0.5 gallon,
and where used with a turbine engine, it has an expansion space of not less
than 10 percent of the tank capacity.
(d) It is impossible to fill the tank expansion space inadvertently with
the rotorcraft in the normal ground attitude;
(e) Adequate venting is provided; and
(f) There are means in the filler opening to prevent oil overflow from
entering the oil tank compartment.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-9, 39 FR
35461, Oct. 1, 1974]
Sec. 27.1015 Oil tank tests.
Each oil tank must be designed and installed so that it can withstand,
without leakage, an internal pressure of 5 p.s.i., except that each
pressurized oil tank used with a turbine engine must be designed and
installed so that it can withstand, without leakage, an internal pressure of
5 p.s.i., plus the maximum operating pressure of the tank.
[Amdt. 27-9, 39 FR 35462, Oct. 1, 1974]
Sec. 27.1017 Oil lines and fittings.
(a) Each oil line must be supported to prevent excessive vibration.
(b) Each oil line connected to components of the rotorcraft between which
relative motion could exist must have provisions for flexibility.
(c) Flexible hose must be approved.
(d) Each oil line must have an inside diameter of not less than the inside
diameter of the engine inlet or outlet. No line may have splices between
connections.
Sec. 27.1019 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. 27.1305(r).
(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. 27-20, 49 FR 6849, Feb. 23, 1984]
Sec. 27.1027 Transmissions and gearboxes: General.
(a) Pressure lubrication systems for transmissions and gearboxes must
comply with the engine oil system requirements of Secs. 27.1013 (except
paragraph (c)), 27.1015, 27.1017, 27.1021, and 27.1337(d).
(b) Each pressure lubrication system must have an oil strainer or filter
through which all of the lubricant flows and must--
(1) 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;
(2) Be equipped with a means to indicate collection of contaminants on the
filter or strainer at or before opening of the bypass required by paragraph
(b)(3) of this section; and
(3) Be equipped with a bypass constructed and installed so that--
(i) The lubricant will flow at the normal rate through the rest of the
system with the strainer or filter completely blocked; and
(ii) The release of collected contaminants is minimized by appropriate
location of the bypass to ensure that collected contaminants are not in the
bypass flowpath.
(c) For each lubricant tank or sump outlet supplying lubrication to rotor
drive systems and rotor drive system components, a screen must be provided 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) of this section. The requirements of paragraph (b) do not apply
to screens installed at lubricant tank or sump outlets.
(d) Splash-type lubrication systems for rotor drive system gearboxes must
comply with Secs. 27.1021 and 27.1337(d).
[Amdt. 27-23, 53 FR 34213, Sept. 2, 1988]
Cooling
Sec. 27.1041 General.
(a) Each powerplant cooling system must be able to maintain the
temperatures of powerplant components within the limits established for these
components under critical surface (ground or water) and flight operating
conditions for which certification is required and after normal shutdown.
Powerplant components to be considered include but may not be limited to
engines, rotor drive system components, auxiliary power units, and the
cooling or lubricating fluids used with these components.
(b) Compliance with paragraph (a) of this section must be shown in tests
conducted under the conditions prescribed in that paragraph.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-23, 53 FR
34213, Sept. 2, 1988]
Sec. 27.1043 Cooling tests.
(a) General. For the tests prescribed in Sec. 27.1041(b), 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) For reciprocating engines, the fuel used during the cooling tests must
be of the minimum grade approved for the engines, and the mixture settings
must be those normally used in the flight stages for which the cooling tests
are conducted.
(4) The test procedures must be as prescribed in Sec. 27.1045.
(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 power-plant 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) General. For each stage of flight, the cooling tests must be conducted
with the rotorcraft--
(1) In the configuration most critical for cooling; and
(2) Under the conditions most critical for cooling.
(b) Temperature stabilization. For the purpose of the cooling tests, a
temperature is "stabilized" when its rate of change is less than two degrees
F. per minute. The following component and engine fluid temperature
stabilization rules apply:
(1) For each rotorcraft, and for each stage of flight--
(i) The temperatures must be stabilized under the conditions from which
entry is made into the stage of flight being investigated; or
(ii) If the entry condition normally does not allow temperatures to
stabilize, operation through the full entry condition must be conducted
before entry into the stage of flight being investigated in order to allow
the temperatures to attain their natural levels at the time of entry.
(2) For each helicopter during the takeoff stage of flight, the climb at
takeoff power must be preceded by a period of hover during which the
temperatures are stabilized.
(c) Duration of test. For each stage of flight the tests must be continued
until--
(1) The temperatures stabilize or 5 minutes after the occurrence of the
highest temperature recorded, as appropriate to the test condition;
(2) That stage of flight is completed; or
(3) An operating limitation is reached.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-23, 53 FR
34214, Sept. 2, 1988]
Induction System
Sec. 27.1091 Air induction.
(a) The air induction system for each engine must supply the air required
by that engine under the operating conditions and maneuvers for which
certification is requested.
(b) Each cold air induction system opening must be outside the cowling if
backfire flames can emerge.
(c) If fuel can accumulate in any air induction system, that system must
have drains that discharge fuel--
(1) Clear of the rotorcraft; and
(2) Out of the path of exhaust flames.
(d) For turbine engine powered rotorcraft--
(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 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.
(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 degrees F., and with the engines at 75 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 degrees F.;
(2) Each rotorcraft with sea level engines using carburetors tending to
prevent icing has a sheltered alternate source of air, and that the preheat
supplied to the alternate air intake is not less than that provided by the
engine cooling air downstream of the cylinders;
(3) Each rotorcraft with altitude engines using conventional venturi
carburetors has a preheater capable of providing a heat rise of 120 degrees
F.; and
(4) Each rotorcraft with altitude engines using carburetors tending to
prevent icing has a preheater that can provide a heat rise of--
(i) 100 degrees F.; or
(ii) If a fluid deicing system is used, at least 40 degrees F.
(b) Turbine engine. (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 Part 29 of this
chapter; 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 gram 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
superchargers 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 operating 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))
For each exhaust system--
(a) There must be means for thermal expansion of manifolds and pipes;
(b) There must be means to prevent local hot spots;
(c) Exhaust gases must discharge clear of the engine air intake, fuel
system components, and drains;
(d) 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;
(e) Exhaust gases may not impair pilot vision at night due to glare;
(f) 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;
(g) Each exhaust heat exchanger must incorporate means to prevent blockage
of the exhaust port after any internal heat exchanger failure.
(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. 5074, 29 FR 15695, Nov. 24, 1964 as amended by Amdt. 27-12, 42 FR
15045, Mar. 17, 1977]
Sec. 27.1123 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 operations.
(c) Exhaust piping connected to components between which relative motion
could exist must have provisions for flexibility.
[Amdt. 27-11, 41 FR 55470, Dec. 20, 1976]
Powerplant Controls and Accessories
Sec. 27.1141 Powerplant controls: general.
(a) Powerplant controls must be located and arranged under Sec. 27.777 and
marked under Sec. 27.1555.
(b) Each flexible powerplant control must be approved.
(c) 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.
(d) For turbine engine powered rotorcraft, no single failure or
malfunction, or probable combination thereof, in any powerplant control
system may cause the failure of any powerplant function necessary for safety.
(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. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-12, 42 FR
15045, Mar. 17, 1977; Amdt. 27-23, 53 FR 34214, Sept. 2, 1988]
Sec. 27.1143 Engine controls.
(a) There must be a separate power control for each engine.
(b) Power controls must be grouped and arranged to allow--
(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) 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.
(e) 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.
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) There must be means to quickly shut off all ignition by the grouping of
switches or by a master ignition control.
(b) 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. 752, 775, 49 U.S.C. 1354(a), 1421, and
1423; sec. 6(c), 49 U.S.C. 1655(c))
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-12, 42 FR
15045, Mar. 17, 1977]
Sec. 27.1147 Mixture controls.
If there are mixture controls, each engine must have a separate control and
the controls must be arranged to allow--
(a) Separate control of each engine; and
(b) Simultaneous control of all engines.
Sec. 27.1163 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) 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) 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 must be fire resistant, except that flammable fluid
tanks and supports which are part of and attached to the engine must be
fireproof or be enclosed by a fireproof shield unless damage by fire to any
non-fireproof part will not cause leakage or spillage of flammable fluid.
Components must be shielded or located so as to safeguard against the
ignition of leaking flammable fluid. 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) 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.
(c) Each flammable fluid drain and vent must discharge clear of the
induction system air inlet.
(a) Each fuel tank must be isolated from the engines by a firewall or
shroud.
(b) Each tank or reservoir, other than a fuel tank, that is part of a
system containing flammable fluids or gases must be isolated from the engine
by a firewall or shroud, unless 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 isolated from the engines.
(c) There must be at least one-half inch of clear airspace between each
tank and each firewall or shroud isolating that tank, unless equivalent means
are used to prevent heat transfer from each engine compartment to the
flammable fluid.
Each compartment containing any part of the powerplant installation must
have provision for ventilation.
Sec. 27.1189 Shutoff means.
(a) There must be means to shut off each line carrying flammable fluids
into the engine compartment, except--
(1) Lines, fittings, and components forming an intergral part of an engine;
(2) For oil systems for which all components of the system, including oil
tanks, are fireproof or located in areas not subject to engine fire
conditions; and
(3) For reciprocating engine installations only, engine oil system lines in
installation using engines of less than 500 cu. in. displacement.
(b) There must be means to guard against inadvertent operation of each
shutoff, and to make it possible for the crew to reopen it in flight after it
has been closed.
(c) Each shutoff valve and its control must be designed, located, and
protected to function properly under any condition likely to result from an
engine fire.
(a) Each engine, including the combustor, turbine, and tailpipe sections of
turbine engines 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 a controlled landing: and
(2) Not protected under Sec. 27.861.
(b) Each auxiliary power unit and combustion heater, and any other
combustion equipment to be used in flight, must be isolated from the rest of
the rotorcraft by firewalls, shrouds, or equivalent means.
(c) In meeting paragraphs (a) and (b) of this section, account must be
taken of the probable path of a fire as affected by the airflow in normal
flight and in autorotation.
(d) Each firewall and shroud must be constructed so that no hazardous
quantity of air, fluids, or flame can pass from any engine compartment to
other parts of the rotorcraft.
(e) Each opening in the firewall or shroud must be sealed with close-
fitting, fireproof grommets, bushings, or firewall fittings.
(f) Each firewall and shroud must be fireproof and protected against
corrosion.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-2, 22 FR
964, Jan. 26, 1968]
Sec. 27.1193 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) There must be means for rapid and complete drainage of each part of the
cowling or engine compartment in the normal ground and flight attitudes.
(c) No drain may discharge where it might cause a fire hazard.
(d) Each cowling and engine compartment covering must be at least fire
resistant.
(e) 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.
(f) A means of retaining 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
structural or mechanical failure of the normal retention means, unless such
failure is extremely improbable.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-23, 53 FR
34214, Sept. 2, 1988]
Sec. 27.1194 Other surfaces.
All surfaces aft of, and near, powerplant compartments, other than tail
surfaces not subject to heat, flames, or sparks emanating from a powerplant
compartment, must be at least fire resistant.
[Amdt. 27-2, 33 FR 964, Jan. 26, 1968]
Sec. 27.1195 Fire detector systems.
Each turbine engine powered rotorcraft must have approved quick-acting fire
detectors in numbers and locations insuring prompt detection of fire in the
engine compartment which cannot be readily observed in flight by the pilot in
the cockpit.
[Amdt. 27-5, 36 FR 5493, Mar. 24, 1971]
Subpart F--Equipment
General
Sec. 27.1301 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. 27.1303 Flight and navigation instruments.
The following are the required flight and navigation instruments:
(a) An airspeed indicator.
(b) An altimeter.
(c) A magnetic direction indicator.
Sec. 27.1305 Powerplant instruments.
The following are the required powerplant instruments:
(a) A carburetor air temperature indicator, for each engine having a
preheater that can provide a heat rise in excess of 60 deg. F.
(b) A cylinder head temperature indicator, for each--
(1) Air cooled engine;
(2) Rotorcraft with cooling shutters; and
(3) Rotorcraft for which compliance with Sec. 27.1043 is shown in any
condition other than the most critical flight condition with respect to
cooling.
(c) A fuel pressure indicator, for each pump-fed engine.
(d) A fuel quantity indicator, for each fuel tank.
(e) A manifold pressure indicator, for each altitude engine.
(f) An oil temperature warning device to indicate when the temperature
exceeds a safe value in each main rotor drive gearbox (including any
gearboxes essential to rotor phasing) having an oil system independent of the
engine oil system.
(g) An oil pressure warning device to indicate when the pressure falls
below a safe value in each pressure-lubricated main rotor drive gearbox
(including any gearboxes essential to rotor phasing) having an oil system
independent of the engine oil system.
(h) An oil pressure indicator for each engine.
(i) An oil quantity indicator for each oil tank.
(j) An oil temperature indicator for each engine.
(k) At least one tachometer to indicate the r.p.m. of each engine and, as
applicable--
(1) The r.p.m. of the single main rotor;
(2) The common r.p.m. of any main rotors whose speeds cannot vary
appreciably with respect to each other; or
(3) The r.p.m. of each main rotor whose speed can vary appreciably with
respect to that of another main rotor.
(l) A low fuel warning device for each fuel tank which feeds an engine.
This device must--
(1) Provide a warning to the flightcrew when approximately 10 minutes of
usable fuel remains in the tank; and
(2) Be independent of the normal fuel quantity indicating system.
(m) Means to indicate to the flightcrew the failure of any fuel pump
installed to show compliance with Sec. 27.955.
(n) A gas temperature indicator for each turbine engine.
(o) Means to enable the pilot to determine the torque of each turboshaft
engine, if a torque limitation is established for that engine under Sec.
27.1521(e).
(p) For each turbine engine, an indicator to indicate the functioning of
the powerplant ice protection system.
(q) An indicator for the fuel filter required by Sec. 27.997 to indicate
the occurrence of contamination of the filter at the degree established by
the applicant in compliance with Sec. 27.955.
(r) For each turbine engine, a warning means for the oil strainer or filter
required by Sec. 27.1019, 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. 27.1019(a)(2).
(s) An indicator to indicate the functioning of any selectable or
controllable heater used to prevent ice clogging of fuel system components.
(t) For rotorcraft for which a 30-second/2-minute OEI power rating is
requested, a means must be provided to alert the pilot when the engine is at
the 30-second and the 2-minute OEI power levels, when the event begins, and
when the time interval expires.
(u) For each turbine engine utilizing 30-second/2-minute OEI power, a
device or system must be provided for use by ground personnel which--
(1) Automatically records each usage and duration of power at the 30-second
and 2-minute OEI levels;
(2) Permits retrieval of the recorded data;
(3) Can be reset only by ground maintenance personnel; and
(4) Has a means to verify proper operation of the system or device.
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).
The following is the required miscellaneous equipment:
(a) An approved seat for each occupant.
(b) An approved safety belt for each occupant.
(c) A master switch arrangement.
(d) An adequate source of electrical energy, where electrical energy is
necessary for operation of the rotorcraft.
(e) Electrical protective devices.
Sec. 27.1309 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 equipment, systems, and installations of a multiengine rotorcraft
must be designed to prevent hazards to the rotorcraft in the event of a
probable malfunction or failure.
(c) The equipment, systems, and installations of single-engine rotorcraft
must be designed to minimize hazards to the rotorcraft in the event of a
probable malfunction or failure.
(d) In showing compliance with paragraph (a), (b), or (c) of this section,
the effects of lightning strikes on the rotorcraft must be considered in
accordance with Sec. 27.610.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-21, 49 FR
44435, Nov. 6, 1984]
Instruments: Installation
Sec. 27.1321 Arrangement and visibility.
(a) Each flight, navigation, and powerplant instrument for use by any pilot
must be easily visible to him.
(b) For each multiengine rotorcraft, identical powerplant instruments must
be located so as to prevent confusion as to which engine each instrument
relates.
(c) Instrument panel vibration may not damage, or impair the readability or
accuracy of, any instrument.
(d) 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) of the Dept. of
Transportation Act (49 U.S.C. 1655(c)))
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964; 29 FR 17885, Dec. 17, 1964, as
amended by Amdt. 27-13, 42 FR 36971, July 18, 1977]
Sec. 27.1322 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. 27-11, 41 FR 55470, Dec. 20, 1976]
Sec. 27.1323 Airspeed indicating system.
(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) The airspeed indicating system must be calibrated in flight at forward
speeds of 20 knots and over.
(c) At each forward speed above 80 percent of the climbout speed, the
airspeed indicator must indicate true airspeed, at sea level with a standard
atmosphere, to within an allowable installation error of not more than the
greater of--
(1) +/-3 percent of the calibrated airspeed; or
(2) Five knots.
(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. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-13, 42 FR
36972, July 18, 1977]
Sec. 27.1325 Static pressure systems.
(a) Each instrument with static air case connections must be vented so that
the influence of rotorcraft speed, the opening and closing of windows,
airflow variation, and moisture or other foreign matter does not seriously
affect its accuracy.
(b) 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 the altimeter when on the primary static system
by more than 50 feet, a correction card must be provided for the alternate
static system.
(c) Except as provided in paragraph (d) of this section, if the static
pressure system incorporates both a primary and an alternate static pressure
source, the means for selecting one or the other source must be designed so
that--
(1) When either source is selected, the other is blocked off; and
(2) Both sources cannot be blocked off simultaneously.
(d) For unpressurized rotorcraft, paragraph (c)(1) of this section does not
apply if it can be demonstrated that the static pressure system calibration,
when either static pressure source is selected is not changed by the other
static pressure source being open or blocked.
(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. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-13, 42 FR
36972, July 18, 1977]
Sec. 27.1327 Magnetic direction indicator.
(a) Except as provided in paragraph (b) of this section--
(1) Each magnetic direction indicator must be installed so that its
accuracy is not excessively affected by the rotorcraft's vibration or
magnetic fields; and
(2) The compensated installation may not have a deviation, in level flight,
greater than 10 degrees on any heading.
(b) A magnetic nonstabilized direction indicator may deviate more than 10
degrees due to the operation of electrically powered systems such as
electrically heated windshields if either a magnetic stabilized direction
indicator, which does not have a deviation in level flight greater than 10
degrees on any heading, or a gyroscopic direction indicator, is installed.
Deviations of a magnetic nonstabilized direction indicator of more than 10
degrees must be placarded in accordance with Sec. 27.1547(e).
(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)))
[Amdt. 27-13, 42 FR 36972, July 18, 1977]
Sec. 27.1329 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 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.
[Amdt. 27-21, 49 FR 44435, Nov. 6, 1984]
Sec. 27.1335 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) of the Dept. of
Transportation Act (49 U.S.C. 1655(c)))
[Amdt. 27-13, 42 FR 36972, July 18, 1977]
Sec. 27.1337 Powerplant instruments.
(a) Instruments and instrument lines.
(1) Each powerplant instrument line must meet the requirements of Secs.
27.- 961 and 27.993.
(2) Each line carrying flammable fluids under pressure must--
(i) Have restricting orifices or other safety devices at the source of
pressure to prevent the escape of excessive fluid if the line fails; and
(ii) Be installed and located so that the escape of fluids would not create
a hazard.
(3) Each powerplant 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. Each fuel quantity indicator must be installed
to clearly indicate to the flight crew the quantity of fuel in each tank in
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. 27.959;
(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; and
(3) 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 means to indicate the quantity of
oil in each tank--
(1) On the ground (including during the filling of each tank); and
(2) In flight, if there is an oil transfer system or reserve oil supply
system.
(e) Rotor drive system transmissions and gearboxes utilizing ferromagnetic
materials must be equipped with chip detectors designed to indicate or reveal
the presence of ferromagnetic particles resulting from damage or excessive
wear. Chip detectors must--
(1) Incorporate means to indicate the accumulation of ferromagnetic
particles on the magnetic poles; or
(2) Be readily removable for inspection of the magnetic poles for metallic
chips. Means must be provided to prevent loss of lubricant in the event of
failure of the retention device for removable chip detector components.
(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. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-12, 42 FR
15046, Mar. 17, 1977; Amdt. 27-23, 53 FR 34214, Sept. 2, 1988]
Electrical Systems and Equipment
Sec. 27.1351 General.
(a) Electrical system capacity. Electrical equipment must be adequate for
its intended use. In addition--
(1) Electric power sources, their transmission cables, and their associated
control and protective devices must be able to furnish the required power at
the proper voltage to each load circuit essential for safe operation; and
(2) Compliance with paragraph (a)(1) of this section must be shown by an
electrical load analysis, or by electrical measurements that take into
account the electrical loads applied to the electrical system, in probable
combinations and for probable durations.
(b) Function. For each electrical system, the following apply:
(1) Each system, when installed, must be--
(i) Free from hazards in itself, in its method of operation, and in its
effects on other parts of the rotorcraft; and
(ii) Protected from fuel, oil, water, other detrimental substances, and
mechanical damage.
(2) Electric power sources must function properly when connected in
combination or independently.
(3) No failure or malfunction of any source may impair the ability of any
remaining source to supply load circuits essential for safe operation.
(4) Each electric power source control must allow the independent operation
of each source.
(c) Generating system. There must be at least one generator if the system
supplies power to load circuits essential for safe operation. In addition--
(1) Each generator must be able to deliver its continuous rated power;
(2) Generator voltage control equipment must be able to dependably regulate
each generator output within rated limits;
(3) Each generator must have a reverse current cutout designed to
disconnect the generator from the battery and from the other generators when
enough reverse current exists to damage that generator; and
(4) Each generator must have an overvoltage control designed and installed
to prevent damage to the electrical system, or to equipment supplied by the
electrical system, that could result if that generator were to develop an
overvoltage condition.
(d) Instruments. There must be means to indicate to appropriate crewmembers
the electric power system quantities essential for safe operation of the
system. In addition--
(1) For direct current systems, an ammeter that can be switched into each
generator feeder may be used; and
(2) If there is only one generator, the ammeter may be in the battery
feeder.
(e) 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.
(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. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 41 FR
55470, Dec. 20, 1976; Amdt. 27-13, 42 FR 36972, July 18, 1977]
Sec. 27.1353 Storage battery design and installation.
(a) Each storage battery must be designed and installed as prescribed in
this section.
(b) Safe cell temperatures and pressures must be maintained during any
probable charging and discharging condition. No uncontrolled increase in cell
temperature may result when the battery is recharged (after previous complete
discharge)--
(1) At maximum regulated voltage or power;
(2) During a flight of maximum duration; and
(3) Under the most adverse cooling condition likely to occur in service.
(c) Compliance with paragraph (b) of this section must be shown by test
unless experience with similar batteries and installations has shown that
maintaining safe cell temperatures and pressures presents no problem.
(d) No explosive or toxic gases emitted by any battery in normal operation,
or as the result of any probable malfunction in the charging system or
battery installation, may accumulate in hazardous quantities within the
rotorcraft.
(e) No corrosive fluids or gases that may escape from the battery may
damage surrounding structures or adjacent essential equipment.
(f) Each nickel cadmium battery installation capable of being used to start
an engine or auxiliary power unit must have provisions to prevent any
hazardous effect on structure or essential systems that may be caused by the
maximum amount of heat the battery can generate during a short circuit of the
battery or of its individual cells.
(g) Nickel cadmium battery installations capable of being used to start an
engine or auxiliary power unit must have--
(1) A system to control the charging rate of the battery automatically so
as to prevent battery overheating;
(2) A battery temperature sensing and over-temperature warning system with
a means for disconnecting the battery from its charging source in the event
of an over-temperature condition; or
(3) A battery failure sensing and warning system with a means for
disconnecting the battery from its charging source in the event of battery
failure.
(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. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-13, 42 FR
36972, July 18, 1977; Amdt. 27-14, 43 FR 2325, Jan. 16, 1978]
Sec. 27.1357 Circuit protective devices.
(a) Protective devices, such as fuses or circuit breakers, must be
installed in each electrical circuit other than--
(1) The main circuits of starter motors; and
(2) Circuits in which no hazard is presented by their omission.
(b) A protective device for a circuit essential to flight safety may not be
used to protect any other circuit.
(c) Each resettable circuit protective device ("trip free" device in which
the tripping mechanism cannot be overridden by the operating control) must be
designed so that--
(1) A manual operation is required to restore service after trippling; and
(2) If an overload or circuit fault exists, the device will open the
circuit regardless of the position of the operating control.
(d) If the ability to reset a circuit breaker or replace a fuse is
essential to safety in flight, that circuit breaker or fuse must be located
and identified so that it can be readily reset or replaced in flight.
(e) If fuses are used, there must be one spare of each rating, or 50
percent spare fuses of each rating, whichever is greater.
(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. 5074, 29 FR 15695, Nov. 24, 1964; 29 FR 17885, Dec. 17, 1964, as
amended by Amdt. 27-13, 42 FR 36972, July 18, 1977]
Sec. 27.1361 Master switch.
(a) There must be a master switch arrangement to allow ready disconnection
of each electric power source from the main bus. The point of disconnection
must be adjacent to the sources controlled by the switch.
(b) Load circuits may be connected so that they remain energized after the
switch is opened, if they are protected by circuit protective devices, rated
at five amperes or less, adjacent to the electric power source.
(c) The master switch or its controls must be installed so that the switch
is easily discernible and accessible to a crewmember in flight.
Sec. 27.1365 Electric cables.
(a) Each electric connecting cable must be of adequate capacity.
(b) Each cable that would overheat in the event of circuit overload or
fault must be at least flame resistant and may not emit dangerous quantities
of toxic fumes.
Sec. 27.1367 Switches.
Each switch must be--
(a) Able to carry its rated current;
(b) Accessible to the crew; and
(c) Labeled as to operation and the circuit controlled.
Lights
Sec. 27.1381 Instrument lights.
The instrument lights must--
(a) Make each instrument, switch, and other devices 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. 27.1383 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. 27.1385 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. 27.1387 through 27.1397.
(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. 27.1387 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. 25.1385(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. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-7, 36 FR
21278, Nov. 5, 1971]
Sec. 27.1389 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. 27.1391.
(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. 27.1393, where I is the minimum intensity
prescribed in Sec. 27.1391 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. 27.1395,
except that higher intensities in overlaps may be used with main beam
intensities substantially greater than the minima specified in Secs. 27.1391
and 27.1393, if the overlap intensities in relation to the main beam
intensities do not adversely affect signal clarity. When the peak intensity
of the forward position lights is greater than 100 candles, the maximum
overlap intensities between them may exceed the values in Sec. 27.1395 if the
overlap intensity in Area A is not more than 10 percent of peak position
light intensity and the overlap intensity in Area B is not more than 2.5
percent of peak position light intensity.
Sec. 27.1391 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 10 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. 27.1393 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 l
0 deg. 1.00
0 deg. to 5 deg. 0.90
5 deg. to 10 deg. 0.80
10 deg. to 15 deg. 0.70
15 deg. to 20 deg. 0.50
20 deg. to 30 deg. 0.30
30 deg. to 40 deg. 0.10
40 deg. to 90 deg. 0.05
Sec. 27.1395 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. 27.1389(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. 27.1397 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--0.400x";
Where "yc" is the "y" coordinate of the Planckian radiator for the value of
"x" considered.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-6, 36 FR
12972, July 10, 1971]
Sec. 27.1399 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 nautical 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.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-2, 33 FR
964, Jan. 26, 1968]
Sec. 27.1401 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 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. 27.1397.
(e) Light intensity. The minimum light intensities in any vertical plane,
measured with the red filter (if used) and expressed in terms of "effective"
intensities, must meet the requirements of paragraph (f) of this section. The
following relation must be assumed:
t2
INTEGRAL I(t)dt
t1
Ie =
--------------
0.2+(t2-t1)
where:
Ie=effective intensity (candles).
I(t)=instantaneous intensity as a function of time.
t2-t1=flash time interval (seconds).
Normally, the maximum value of effective intensity is obtained when t2 and t1
are chosen so that the effective intensity is equal to the instantaneous
intensity at t2 and t1.
(f) Minimum effective intensities for anticollision 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. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-6, 36 FR
12972, July 10, 1971; Amdt. 27-10, 41 FR 5290, Feb. 5, 1976]
Safety Equipment
Sec. 27.1411 General.
(a) Required safety equipment to be used by the crew in an emergency, such
as flares and automatic liferaft releases, must be readily accessible.
(b) Stowage provisions for required safety equipment must be furnished and
must--
(1) Be arranged so that the equipment is directly accessible and its
location is obvious; and
(2) Protect the safety equipment from damage caused by being subjected to
the inertia loads specified in Sec. 27.561.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 41 FR
55470, Dec. 20, 1976]
Sec. 27.1413 Safety belts.
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)))
(a) Emergency flotation and signaling equipment required by any operating
rule in this chapter must meet the requirements of this section.
(b) Each raft and each life preserver must be approved and must be
installed so that it is readily available to the crew and passengers. The
storage provisions for life preservers must accommodate one life preserver
for each occupant for which certification for ditching is requested.
(c) Each raft released automatically or by the pilot must be attached to
the rotorcraft by a line to keep it alongside the rotorcraft. This line must
be weak enough to break before submerging the empty raft to which it is
attached.
(d) Each signaling device must be free from hazard in its operation and
must be installed in an accessible location.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 41 FR
55470, Dec. 20, 1976]
Sec. 27.1419 Ice protection.
(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 Part 29 of this chapter 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 indentified 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. 27-19, 48 FR 4389, Jan. 31, 1983]
Sec. 27.1435 Hydraulic systems.
(a) Design. Each hydraulic system and its elements must withstand, without
yielding, any structural loads expected in addition to hydraulic loads.
(b) Tests. Each system must be substantiated by proof pressure tests. When
proof tested, no part of any system may fail, malfunction, or experience a
permanent set. The proof load of each system must be at least 1.5 times the
maximum operating pressure of that system.
(c) Accumulators. No hydraulic accumulator or pressurized reservoir may be
installed on the engine side of any firewall unless it is an integral part of
an engine.
Sec. 27.1457 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 adjusted or
supplemented so 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 microphone 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. 27-22, 53 FR 26144, July 11, 1988]
Sec. 27.1459 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. 27.1323, 27.1325,
and 27.1327 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 preflight checking of the
recorder for proper recording of data in the storage medium;
(5) Except for recorders powered solely by the engine-driven electrical
generator system, there is an automatic means to simultaneously stop a
recorder that has a data erasure feature and prevent each erasure feature
from functioning, within 10 minutes after any crash impact; and
(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 they are not likely to be separated during crash impact.
[Amdt. 27-22, 53 FR 26144, July 11, 1988]
Sec. 27.1461 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. 27-2, 33 FR 964, Jan. 26, 1968]
Subpart G--Operating Limitations and Information
Sec. 27.1501 General.
(a) Each operating limitation specified in Secs. 27.1503 through 27.1525
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.
27.1541 through 27.1589.
(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) of the Dept. of
Transportation Act (49 U.S.C. 1655(c)))
[Amdt. 27-14, 43 FR 2325, Jan. 16, 1978]
Operating Limitations
Sec. 27.1503 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. 27.1505 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. 27.309;
(ii) 0.9 times the maximum speed shown under Secs. 27.251 and 27.629; or
(iii) 0.9 times the maximum speed substantiated for advancing blade tip
mach number effects.
(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. 27.65(b) for single engine helicopters; and
(ii) Sec. 27.67 for multiengine helicopters.
(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) of the 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. 27.309(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. 27.33(a)(1) and (b)(1).
Sec. 27.1519 Weight and center of gravity.
The weight and center of gravity limitations determined under Secs. 27.25
and 27.27, respectively, must be established as operating limitations.
[Amdt. 27-2, 33 FR 965, Jan. 26, 1968, as amended by Amdt. 27-21, 49 FR
44435, Nov. 6, 1984]
Sec. 27.1521 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 determined by the rotor design; or
(ii) The maximum value shown during the type tests;
(2) The maximum allowable manifold pressure (for reciprocating engines);
(3) The time limit for the use of the power corresponding to the
limitations established in paragraphs (b)(1) and (2) of this section;
(4) If the time limit in paragraph (b)(3) of this section exceeds two
minutes, the maximum allowable cylinder head, coolant outlet, or oil
temperatures;
(5) The gas temperature limits for turbine engines over the range of
operating and atmospheric conditions for which certification is requested.
(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. 27.1509(c); and
(3) The gas temperature limits for turbine engines over the range of
operating and atmospheric conditions for which certification is requested.
(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) Turboshaft engine torque. For rotorcraft with main rotors driven by
turboshaft engines, and that do not have a torque limiting device in the
transmission system, the following apply:
(1) A limit engine torque must be established if the maximum torque that
the engine can exert is greater than--
(i) The torque that the rotor drive system is designed to transmit; or
(ii) The torque that the main rotor assembly is designed to withstand in
showing compliance with Sec. 27.547(e).
(2) The limit engine torque established under paragraph (e)(1) of this
section may not exceed either torque specified in paragraph (e)(1) (i) or
(ii) of this section.
(f) Ambient temperature. For turbine engines, ambient temperature
limitations (including limitations for winterization installations, if
applicable) must be established as the maximum ambient atmospheric
temperature at which compliance with the cooling provisions of Secs. 27.1041
through 27.1045 is shown.
(g) 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 after failure of an engine. 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 demonstrated during the type tests;
(2) The maximum allowable gas temperature; and
(3) The maximum allowable torque.
(h) 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 demonstrated during the type tests;
(2) The maximum allowable gas temperature; and
(3) The maximum allowable torque.
(i) 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 demonstrated during the type tests;
(2) The maximum allowable gas temperature; and
(3) The maximum allowable torque.
(j) 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
A27.4 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.
(k) 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
A27.4 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) of the 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).
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. 27.1525.
Sec. 27.1525 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. 27-21, 49 FR 44435, Nov. 6, 1984]
Sec. 27.1527 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) of the Dept. of
Transportation Act (49 U.S.C. 1655(c)))
[Amdt. 27-14, 43 FR 2325, Jan. 16, 1978]
Sec. 27.1529 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. 27-18, 45 FR 60177, Sept. 11, 1980]
Markings and Placards
Sec. 27.1541 General.
(a) The rotorcraft must contain--
(1) The markings and placards specified in Secs. 27.1545 through 27.1565,
and
(2) Any additional information, instrument markings, and placards required
for the safe operation of rotorcraft with unusual design, operating or
handling characteristics.
(b) Each marking and placard prescribed in paragraph (a) of this section--
(1) Must be displayed in a conspicuous place; and
(2) May not be easily erased, disfigured, or obscured.
Sec. 27.1543 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. 27.1545 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 rotocraft other than helicopters, at VNE; and
(ii) For helicopters at 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) of the 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 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.
(e) If a magnetic nonstabilized direction indicator can have a deviation of
more than 10 degrees caused by the operation of electrical equipment, the
placard must state which electrical loads, or combination of loads, would
cause a deviation of more than 10 degrees when turned on.
(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. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-13, 42 FR
36972, July 18, 1977]
Sec. 27.1549 Powerplant instruments.
For each required powerplant instrument, as appropriate to the type of
instrument--
(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. 27.1553 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. 27.1555 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 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; 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 be red. A placard must be near each emergency exit
control and must clearly indicate the location of that exit and its method of
operation.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as amended by Amdt. 27-11, 41 FR
55471, Dec. 20, 1976]
Sec. 27.1559 Limitations placard.
There must be a placard in clear view of the pilot that specifies the kinds
of operations (such as VFR, IFR, day, night, or icing) for which the
rotorcraft is approved.
[Amdt. 27-21, 49 FR 44435, Nov. 6, 1984]
Sec. 27.1561 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.
Sec. 27.1565 Tail rotor.
Each tail rotor must be marked so that its disc is conspicuous under normal
daylight ground conditions.
[Amdt. 27-2, 33 FR 965, Jan. 26, 1968]
Rotorcraft Flight Manual and Approved Manual Material
Sec. 27.1581 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. 27.1583 through 27.1589.
(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. 27.1583
through 27.1589, that is appropriate to the rotorcraft, must be furnished,
verified, and approved, and must be segregated, identified, 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) of the Dept. of
Transportation Act (49 U.S.C. 1655(c)))
[Amdt. 27-14, 43 FR 2325, Jan. 16, 1978]
Sec. 27.1583 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. 27.1521.
(2) Explanation of the limitations, when appropriate.
(3) Information necessary for marking the instruments required by Secs.
27.1549 through 27.1553.
(c) Weight and loading distribution. The weight and center of gravity
limits required by Secs. 27.25 and 27.27, 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. 27.1523
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) [Reserved]
(g) Altitude. The altitude established under Sec. 27.1527 and an
explanation of the limiting factors 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) of the Dept. of
Transportation Act (49 U.S.C. 1655(c)))
(a) Parts of the manual containing operating procedures must have
information concerning any normal and emergency procedures and other
information necessary for safe operation, including takeoff and landing
procedures and associated airspeeds. The manual must contain any pertinent
information including--
(1) The kind of takeoff surface used in the tests and each appropriate
climbout speed; and
(2) The kind of landing surface used in the tests and appropriate approach
and glide airspeeds.
(b) For multiengine rotorcraft, information identifying each operating
condition in which the fuel system independence prescribed in Sec. 27.953 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.
27.1505(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. 27.1353 (g)(2) or
(g)(3), the operating procedures for disconnecting the battery from its
charging source must be furnished.
(e) If the unusable fuel supply in any tank exceeds five percent of the
tank capacity, or one 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) The airspeeds and rotor speeds for minimum rate of descent and best
glide angle as prescribed in Sec. 27.71 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) of the Dept. of
Transportation Act (49 U.S.C. 1655(c)))
(a) The rotorcraft must be furnished with the following information,
determined in accordance with Secs. 27.51 through 27.79 and 27.143(c):
(1) Enough information to determine the limiting height-speed envelope.
(2) Information relative to--
(i) The hovering ceilings and the steady rates of climb and descent, as
affected by any pertinent factors such as airspeed, temperature, and
altitude;
(ii) The maximum safe wind for operation near the ground. If there are
combinations of weight, altitude, and temperature for which performance
information is provided and at which the rotorcraft cannot land and takeoff
safely with the maximum wind value, those portions of the operating envelope
and the appropriate safe wind conditions shall be identified in the flight
manual;
(iii) For reciprocating engine-powered rotorcraft, the maximum atmospheric
temperature at which compliance with the cooling provisions of Secs. 27.1041
through 27.1045 is shown; and
(iv) Glide distance as a function of altitude when autorotating at the
speeds and conditions for minimum rate of descent and best glide as
determined in Sec. 27.71.
(b) The Rotorcraft Flight Manual must contain--
(1) In its performance information section any pertinent information
concerning the takeoff weights and altitudes used in compliance with Sec.
27.51; and
(i) Any pertinent information concerning the takeoff procedure, including
the kind of takeoff surface used in the tests and each appropriate climb-out
speed; and
(ii) Any pertinent landing procedures, including the kind of landing
surface used in the tests and appropriate approach and glide airspeeds; and
(2) The horizontal takeoff distance determined in accordance with Sec.
27.65(a)(2)(i).
(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)))
There must be loading instructions for each possible loading condition
between the maximum and minimum weights determined under Sec. 27.25 that can
result in a center of gravity beyond any extreme prescribed in Sec. 27.27,
assuming any probable occupant weights. Appendix A Part
27--Instructions for Continued Airworthiness
A27.1 General.
(a) This appendix specifies requirements for the preparation of
Instructions for Continued Airworthiness as required by Sec. 27.1529.
(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 appliance 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.
A27.2 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.
A27.3 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 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 problem 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,
discarded recommendations, and torque values.
(g) A list of special tools needed.
A27.4 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. 27.571. 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 inspections
and other maintenance required under Secs. 43.16 and 91.403 of the Federal
Aviation Regulations unless an alternative program has been FAA approved."
[Amdt. 27-17, 45 FR 60178, Sept. 11, 1980, as amended by Amdt. 27-24, 54 FR
34329, Aug. 18, 1989]
Effective Date Note: At 54 FR 34329, Aug. 18, 1989, Sec. A27.4 in Appendix
A, Part 27 was amended by changing the cross reference "Sec. 91.163" to "Sec.
91.403", effective August 18, 1989
Appendix B to Part 27--Airworthiness Criteria for Helicopter Instrument
Flight
I. General. A normal 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 paragraph IV
(b) or (c) of this appendix, as appropriate. The stick force must vary with
speed so that any substantial speed change results in a stick force clearly
perceptible to the pilot. For single-pilot approval, 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 paragraph IV(b) of the this
appendix.
(b) For single-pilot approval:
(1) Climb. Stability must be shown in climb throughout the speed range 20
knots either side of trim with--
(i) The helicopter trimmed at VYI;
(ii) Landing gear retracted (if retractable); and
(iii) Power required for limit climb rate (at least 1,000 fpm) at VYI or
maximum continuous power, whichever is less.
(2) 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--
(i) The helicopter trimmed and power adjusted for level flight at 0.9 VH or
0.9 VNEI, whichever is lower; and
(ii) Landing gear retracted (if retractable).
(3) 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--
(i) the helicopter trimmed and power adjusted for level flight at 1.1
VMINI; and
(ii) Landing gear retracted (if retractable).
(4) Descent. Stability must be shown throughout the speed range 20 knots
either side of trim with--
(i) The helicopter trimmed at 0.8 VH or 0.8 VNEI (or 0.8 VLE for the
landing gear extended case), whichever is lower;
(ii) Power required for 1,000 fpm descent at trim speed; and
(iii) Landing gear extended and retracted, if applicable.
(5) 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--
(i) The helicopter trimmed at the recommended approach speed or speeds;
(ii) Landing gear extended and retracted, if applicable; and
(iii) Power required to maintain a 3 deg. glide path and power required to
maintain the steepest approach gradient for which approval is requested.
(c) Helicopters approved for a minimum crew of two pilots must comply with
the provisions of paragraphs IV(b)(2) and IV(b)(5) of this appendix.
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 cyclic movement with sideslip must not be excessive.
VI. Dynamic stability. (a) For single-pilot approval--
(1) Any oscillation having a period of less than 5 seconds must damp to 1/2
amplitude in not more than one cycle.
(2) 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.
(3) Any oscillation having a period of 10 seconds or more but less than 20
seconds must be damped.
(4) Any oscillation having a period of 20 seconds or more may not achieve
double amplitude in less than 20 seconds.
(5) Any aperiodic response may not achieve double amplitude in less than 6
seconds.
(b) For helicopters approved with a minimum crew of two pilots--
(1) Any oscillation having a period of less than 5 seconds must damp to 1/2
amplitude in not more than two cycles.
(2) Any oscillation having a period of 5 seconds or more but less than 10
seconds must be damped.
(3) Any oscillation having a period of 10 seconds or more may not achieve
double amplitude in less than 10 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 27 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.
VIII. Equipment, systems, and installation. The basic equipment and
installation must comply with Secs. 29.1303, 29.1431, and 29.1433 through
Amendment 29-14, with the following exceptions and additions:
(a) Flight and Navigation Instruments. (1) A magnetic gyro-stablized
direction indicator instead of a 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). For two-pilot configurations, one pilot's primary indicator
may be designated for this purpose. If standby batteries are provided, they
may be charged from the aircraft electrical system if adequate isolation is
incorporated.
(b) Miscellaneous requirements. (1) Instrument systems and other systems
essential for IFR flight that could be adversely affected by icing must be
adequately protected when exposed to the continuous and intermittent maximum
icing conditions defined in Appendix C of Part 29 of this chapter, 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 crewmember 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.
