This AC paragraph is final and will be included in the upcoming Change 2 update.
AC 29.45. § 29.45 (Amendment 29-24) PERFORMANCE - GENERAL.
a. Explanation.
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b. Procedures.
(1) Winds For Testing.
(i) Allowable wind conditions will vary with the type of test and will also be different for different types and gross weight rotorcraft. For example, higher winds may be tolerated for takeoff and landing distance tests but not for hover performance.
Likewise higher winds may be tolerated during hover performance testing on large, heavy rotorcraft with high lower rotor downwash velocities than for smaller rotorcraft with rotor downwash velocities. Generally, unless the effects of wind on hover performance tests can be determined and accounted for, hover performance testing should be conducted in winds of 3 knots or less.
(ii) Past experience has shown that a steady wind of 0 to 10 knots will result in acceptable takeoff and landing performance if distances are corrected for the winds measured during these tests. This is not the case for vertical takeoffs and landings. To obtain consistent and repeatable vertical performance data, the same general wind criteria used to obtain hover performance; i.e., up to 3 knots, should be used for vertical performance determination. In actuality, a rotorcraft may exhibit reduced IGE hover performance in winds from 3 to 15 knots due to partial immersion of the main rotor in its own vortex. Since the height-speed envelope determination is affected by wind just as vertical takeoff and landing performance are, the same allowable winds for testing should be adhered to for HV testing; i.e., 0 to 3 knots. For Category A testing, the effects of crosswind and/or tailwind should also be considered up to the maximum for which Category A certification is requested.
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This AC paragraph is final and will be included in the upcoming Change 2 update.
AC 29.59A (AC's 29.60, 29.61, & 29.62) §§ 29.59 (29.60, 29.61 and 29.62) (Amendment 29-39) TAKEOFF PATH, DISTANCE AND REJECTED TAKEOFF; GROUND LEVEL AND ELEVATED HELIPORT: CATEGORY A
(For § 29.59 prior to Amendment 39, see paragraph AC 29.59.)
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b. Procedures.
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(7) Power. Power used for demonstrating performance should be limited to minimum specification values on the operating engine(s). This may be accomplished by adjustment of the engine topping (maximum power available) to minimum specification values including consideration of temperature effects on engine power. If topping results in unrepresentative engine power management, the validity of the Cat A procedure must also be established with representative in-service characteristics. The method used for simulating engine failure must be representative of the power decay characteristics that will occur during a real, sudden engine failure and acceleration of the remaining engine(s). In order to cushion a rejected take-off, it is acceptable for the engine and transmission transient range to be entered in order to droop the rotor provided performance credit is not taken for this additional power above the maximum permitted rating and it can be shown that the engine(s) will remain within these limits in all conditions requested by the applicant. Any excursion beyond established transient limits in this flight phase should be substantiated to the extent that it does not constitute an immediate hazard to the rotorcraft.
(8) Turbine engine power does not vary directly with density altitude (HD). At a given HD, turbine engine power available varies with ambient temperature. Turbine engines typically produce less horsepower as ambient temperature is increased (pressure altitude decreases) at a given density altitude, although some engines produce less horsepower at extremely cold temperatures. In either event, if one test sequence is to be utilized for a given HD, it would be appropriate to restrict test power to the lowest value attainable from a minimum specification engine through the approved ambient temperature range at the density altitude of the test. To attain maximum weights for varying ambient conditions, the applicant may utilize a parametric mapping of power available, pressure altitude, and temperature effects. For this case, engine topping may be adjusted throughout a range appropriate to the test HD.
(9) Aircraft Loading. Both forward and aft CG extremes should be briefly checked to determine the critical loading for takeoff distances. Forward center of gravity is usually critical for continued takeoff distance tests while aft CG may be critical for the rejected takeoff due to forward and downward field of view. A minimum of two weights should be flown at each altitude if the manufacturer elects to schedule field length variation as a function of gross weight. One weight should be the maximum weight for Page 2
This AC paragraph is final and will be included in the upcoming Change 2 update.
prevailing conditions and the other weight(s) should be low enough to attain a sufficient spread to verify weight effect.
(10) Extrapolation. Takeoff and landing data may be extrapolated up to 4000 feet along an established W/σ line, to the maximum gross weight of the rotorcraft.
However, extrapolation will not be considered valid if unacceptable or marginally acceptable landing gear loads are experienced during testing at weights below the W/σ limit. See paragraph AC 29.77b(5) for further discussion of landing gear loads.
(11) Ambient Conditions. Appropriate test limits for ambient conditions such as wind and temperature are contained in paragraph AC 29.45. Test data should be corrected for existing wind conditions during takeoff distance testing. Credit for headwind conditions may be given during flight manual data expansion. Refer to paragraph AC 29.1587(a)(3)(iii) under “Wind Accountability” for allowable wind credit.
Care should be applied in considering headwind credit for vertical operations as previous experience has resulted in difficulty collecting meaningful, repeatable data.
(12) Vertical Takeoffs.
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(C) When used, the back-up technique usually requires the pilot to keep sufficient portions of the helipad in view and involves a rearward movement from the takeoff point to the TDP. In such cases the rearward horizontal distance required should be established as the distance from the rearmost point of the rotorcraft at the initiation of takeoff to the rearmost part of the rotorcraft at TDP. As stated in AC 29.45, crosswinds and tailwinds should be considered if requested by the applicant. Typically, this will require flight-testing to evaluate performance, pilot workload, field-of-view, and visual cueing.
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(13) Night Operations.
(i) A minimum of three normal takeoffs (and landings) should be conducted to ensure that aircraft lighting (internal and external) is adequate to allow normal Category A operations at night.
(ii) Engine failures should be simulated from points along the requested takeoff and landing profiles. Night OEI rejected takeoffs and continued takeoffs from the TDP and OEI landings from the LDP should be conducted at the requested WAT limiting conditions to ensure adequate night field of view, suitability of aircraft external lighting, and meets the Category A profiles.
This AC paragraph is final and will be included in the upcoming Change 2 update.
(iii) If special airfield marking or lighting is used as a reference or to define the TDP, the aircraft external lighting should be evaluated to assure that the airfield marking or lighting is adequate for night operations.
This AC paragraph is final and will be included in the upcoming Change 2 update.
AC 29.75A. (AC's 29.77, 29.79, 29.81, & 29.83) §§ 29.75, 29.77, 29.79, 29.81, and
29.83 (Amendment 29-39) LANDING.
(For § 29.77 and § 29.79 prior to Amendment 39, see paragraphs AC 29.77 and AC 29.79 respectively.)
a. Explanation.
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b. Procedures - Category A Requirements.
(1) Explanation. The Category A certification concept limits landing weight to a value that will allow the rotorcraft, following an engine failure at the landing decision point (LDP), to land within the available area or to execute a balked landing descending no lower than 15 feet (or higher depending on rotorcraft geometry and performance characteristics) above the landing surface. For elevated heliports the rotorcraft may descend below the landing surface, but all parts of the rotorcraft must clear the heliport and other obstacles by not less than 15 feet. These minimum heights should be demonstrated with variations in piloting techniques and with pilot recognition and reaction times for engine failures occurring before and after the LDP. See figure AC 29.75A-1. For additional information addressing the OEI landing case at night, refer to AC 29.59A.b(12).
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(2) Procedures.
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(iii) Power. Power used for demonstrating performance should be limited to minimum specification values on the operating engine(s). This may be accomplished by adjustment of the engine topping to minimum specification values including consideration of temperature effects on engine power. If the management of engine power at topping is beyond normal pilot capability, the validity of the Cat A procedure must also be established with representative in-service characteristics. The method used for simulating engine failure must be representative of the power decay characteristics that will occur during a real, sudden engine failure and acceleration of the remaining engine(s). In order to cushion the OEI landing, it is acceptable for the engine and transmission transient range in order to droop the rotor provided performance credit is not taken for this additional power above the maximum permitted rating and it can be shown that the engine(s) will remain within these limits in all conditions requested by the applicant. Any excursion beyond established transient Page 5
This AC paragraph is final and will be included in the upcoming Change 2 update.
limits in this flight phase should be substantiated to the extent that it does not constitute an immediate hazard to the rotorcraft.
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This AC paragraph is final and will be included in the upcoming Change 2 update.
AC 29.337. § 29.337 (Amendment 29-30) LIMIT MANEUVERING LOAD FACTOR.
a. Explanation. The rotorcraft must be designed and substantiated to load factors as specified to provide a minimum level of structural integrity of the rotorcraft airframe and rotors.
(1) A range of design positive load factors from +3.5 to +2.0 may be used.
(2) A range of design negative load factors from -1.0 to -0.5 may be used.
(3) Load factors inside the range of +3.5 to -1.0 may be used provided the probability of exceeding the design load factors is shown by analysis and flight tests to be extremely remote, and the selected load factors are appropriate to each weight condition between design maximum and minimum weights.
(4) Load factors exceeding these “minimums” may be used.
b. Procedures.
(1) The applicant may elect to substantiate the rotorcraft for a design maneuvering load factor less than +3.5 and more than -1.0. Whenever this option is used, an analytical study and flight demonstration are required. Maximum available rotor lift with both power on and power off must be considered when substantiating maneuver load factors less than the specified values.
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This AC paragraph is final and will be included in the upcoming Change 2 update.
AC 29.631. § 29.631 (Amendment 29-40) BIRD STRIKE.
a. Explanation. Amendment 40 adds requirements for continued safe flight and landing after a bird strike. Compliance with § 29.631 must be shown for a 2.2 lb (1.0 kg) bird at a relative velocity equal to the lesser of VNE or VH at altitudes up to 8,000 feet.
For Category A certification, the rotorcraft must be capable of continued safe flight and landing after the described bird strike. For Category B certification, the rotorcraft must be capable of a safe landing after the bird strike.
b. Procedures. For compliance with FAR 29.631, it should be demonstrated by test or analysis supported by test evidence that,
(1) The windshields will withstand the bird strike, without penetration, and,
(2) The rotorcraft is capable of continued safe flight and landing following impact with a 2.2-lb (1.0 kg) bird at VNE or VH (whichever is the lesser) at altitudes up to 8,000 feet. Areas of impact that are of particular interest include flight control surfaces (which includes main and tail rotors) and exposed flight control system components.
This AC paragraph is final and will be included in the upcoming Change 2 update.
AC 29.672. § 29.672 (Amendment 29-24) STABILITY AUGMENTATION, AUTOMATIC, AND POWER-OPERATED SYSTEMS.
a. Explanation.
(1) This rule requires that the pilot be made aware of stability augmentation, automatic, or power-operated system failures which could lead to an unsafe condition.
It should be understood that this requirement applies to stability augmentation and supplementary controls and not the primary flight control system, which is dealt with under § 29.695 and associated advisory material. Examples of clearly distinguishable warnings include, but are not limited to, an obvious aircraft attitude change following the failure or an audio warning tone. A visual indication itself may not be adequate since detection of a visual warning would normally require special pilot attention. The use of devices such as stick pushers or shakers is not acceptable as a warning means since the automatic flight control systems (AFCS) may provide a hands-off capability or normal helicopter vibrations could mask a control shaker. However, this rule is not intended to eliminate the use of such devices for other purposes. Examples of automatic control systems other than a stability augmentation system would be a pitch axis actuator used for the purpose of demonstrating compliance with longitudinal static stability requirements or a fly-by-wire elevator. The design of such systems must not interfere with completion of the control checks described in § 29.671(c). Further, for control systems where a series actuator malfunction could degrade control authority, a means should be provided to the pilot to determine actuator alignment (see § 29.1329(b)).
(2) The corrective flight control input following a system failure should be in the logical direction. For example, a malfunction resulting in a nosedown pitch of the aircraft should require a corrective cyclic control input in the aft direction. The system deactivating means does not have to be located on the primary flight control grips;
however, it should be easily accessible to the pilot. Consideration should be given to the consequences of inadvertent de-selection of the automatic stabilization system, especially if the deactivation control is mounted on a primary control grip. Malfunctions and subsequent recoveries must be shown throughout the operating envelope of the aircraft. In a case where control authority is decreased following a malfunction, a practical flight envelope must be defined wherein compliance with controllability and maneuverability requirements can be demonstrated. This practical flight envelope must be presented in the flight manual. Compliance with trim and stability characteristics is not required following a malfunction; however, a pilot workload assessment should be made to show that these characteristics are not impaired below that needed for continued safe flight and landing.
This AC paragraph is final and will be included in the upcoming Change 2 update.
AC 29.683. § 29.683 OPERATION TESTS.
a. Explanation. The rule requires that the control system be free from jamming, excessive friction, and excessive deflection. An operational test is required in which specified loads are applied at the pilot controls and carried through an operating control system.
b. Procedures.
(1) Compliance with the requirements of this rule is obtained by use of a test setup similar to that used for the limit load tests of § 29.681, except the load reactions at the blades (or surfaces) must allow for movement of the blades (or surfaces) as the system is operated through its operating range.
(2) Fixtures are normally affixed to the surfaces (or replace the surfaces) to allow pulley arrangements which provide for movement under load. These fixtures should be evaluated to assure that system loads up to limit will be applied during the full range of operations of each system.
(3) Each flight control system should be operated through its entire range under a light load and under limit load. As the controls are being operated, the system should be checked for jamming, excessive friction, and excessive deflection. Excessive deflection includes deflection sufficient to contact other systems or structure. Also, if under these limit load conditions the components deflect, the deflection would be considered excessive if there is permanent deformation of any component or supporting structure. Also any deflection that results in an uncorrected condition when the load is released, e.g., if a bellcrank is forced off-center or over-center during load and does not return to the normal position after load release is excessive deflection. Floor panels, wall panels, and other access panels may have to be removed to permit visual checks of the entire control system. However, care should be taken when removing panels so that airframe structure is not weakened enough to deflect from its normal position when test loads are applied to the control system.
This AC paragraph is final and will be included in the upcoming Change 2 update.
AC 29.777. § 29.777 COCKPIT CONTROLS.
a. Explanation. This section defines the general cockpit control requirements.
Cockpit control location and arrangement, with respect to the pilot’s seat, must be designed to accommodate pilots from 5’2” to 6’0” in height. Pilots within this range should be able to reach and operate all required controls and have sufficient clearance with the structure, panels, etc.
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This AC paragraph is final and will be included in the upcoming Change 2 update.
AC 29.1321. § 29.1321 (Amendment 29-21) ARRANGEMENT AND VISIBILITY.
a. Background. This section is the first in a series that concerns the installation of instruments. Specific requirements for individual instruments are addressed in other paragraphs. The instruments should be arranged in a manner so that the pilot may avail himself of the information displayed by the instruments without undue distraction.
Additionally, for instrument flight, the rule requires that the attitude, altitude, airspeed, and compass indicators be grouped in the so-called standard “T.” Instrument location and arrangement, with respect to the pilot’s seat, should be designed to accommodate pilots from 5’2” to 6’0” in height. Pilots within this range should be able to see and, where necessary, reach and operate all of the displays.
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