(11) Section 29.143(c) requires a minimum 17-knot control capability for hover and takeoff in winds from any azimuth. Control capability in wind from 0 to at least 17 knots must also be shown for any other appropriate maneuver near the ground such as rolling takeoffs for wheeled rotorcraft. These requirements must be met at all altitudes approved for takeoff and landing. On helicopters incorporating a tail rotor, efficiency of the tail rotor decreases with altitude so that a given sideward flight condition requires more pedal deflection, a higher tail rotor blade angle, and more horsepower. Hence, directional capability in sideward flight (or at critical wind azimuth) is most critical during testing at a high altitude site. Hover controllability, height-velocity, and hover performance are the three regulatory requirements that ordinarily determine the shape of the limiting weight-altitude-temperature (WAT) curve for takeoff and landing at high density altitudes. For Category A rotorcraft, of course, one engine inoperative climb performance requirements may also influence the WAT limit curve.
(111) Section 29.143(d) requires adequate controllability when an engine fails. This requirement specifies conditions under which engine failure testing must be conducted and includes minimum required delay times. * (A) For rotorcraft which meet the engine isolation requirements of Category A, demonstration of sudden complete single-engine failure is required at critical conditions throughout the flight envelope including hover, takeoff, engine lure trant to ling xium continue fer for takeoff power where appropriate) including reasonable engine torque splits. For multiengined Category A installations (3 or more engines) subsequent engine failures should be conducted utilizing the same criteria as that used for first engine failure. The applicant may limit his flight envelope for subsequent failures. Initial or sequential engine failure tests are ordinarily much less severe than the "last" engine failure test required by § 29.75(b) (5). The conditions for last engine failure are maximum continuous power, or 30-minute power if that rating is approved, level flight, and sudden engine failure with the same pilot delay of l second or normal pilot reaction time, whichever is greater.
(B) For Category B, demonstration of sudden complete power failure is required at critical conditions throughout the flight envelope. This includes speeds from zero to VNE (power-on) and conditions of hover, takeoff, climb at Vy, and high speed flight up to VNE. Maximum continuous power is specified prior to the failure, and for cruise conditions, no pilot reaction may begin prior to 1 second or normal pilot reaction time (whichever is greater).
Normal pilot control motions which are part of the pilot task of flight path control are not subject to the 1-second restriction on pilot reaction. The term "oruise" also includes cruise climb and cruise descent conditions. Normal pilot reaction times are used elsewhere. Although this requirement specifies maximum continuous (MC) power, it does not limit engine failure testing to MC power.
takeoff power rating is authorized for hover or takeoff, engine failure testing must also be accomplished for those conditions in order to comply with § 29.63(c). Following power failure, rotor speed, flapping, and aircraft dynamic characteristios must stay within structurally approved limits.
Chap 2 Par 81 153
AC 29-2, CHG 1 11/19/84 * Section 29.143(e) addresses the special case in which a VNE (power-off) 1s established at an airspeed value less than VNE (power-on). For this case, engine failure tests are still required at speeds up to and including VNE (power-on) and the rotorcraft must be capable of being slowed to VNE (power-off) in a controlled manner with normal pilot reactions and skill. There 18, however, no controllability requirement for stabilized power-off flight at speeds above 1.1 VNE (power-off) when VNE (power-off) 1s established per § 29.1505(c).
Application of the controllability requirement for pitch, roll, and yaw at speeds of 1.1 VNE (power-off) and below is similar to that described above for power-on testing at VNE. Sufficient directional control must exist to allow straight flight in autorotation during all approved maneuvers including 30 degrees banked turns up to VNE (power-off) with some small additional allowance for gust control, Adequate controllability margins must exist in all axes throughout the approved autorotative flight envelope. Testing to VNE at MC power per § 29.143(b), 1.1 VNE at power for .9 VH per $ 29.175(b) or 9 29.1505, and to
1.1 VNE (power-off) in autorotation per § 29.143(e) should be sufficient to assure adequate control margin during a descent condition at high speed and low power. The high speed, power-on descent condition should be checked for adequate control margin as a "maneuver appropriate to the type." There has been one instance where insufficient directional pedal was available to maintain a reasonable trimmed sideslip angle with low power at very high speeds, and a case where there was insufficient forward and lateral cyclic available to reach the power on VNE. The insufficient directional pedal margin was due to the offset vertical stabilizers. The lack of cyclic stick margin was because the cyclic stick migrated to the right as power was reduced, and the control limits were ciroular. This provided less total available forward cyclic stick travel when the cyclic was moved right and forward about 45° from the center position. Each of the above helicopters was certificated with a rate of descent limitation to preclude operation in the control limited area.
An evaluation of the emergency descent capability of the helicopter should be made, either analytically or through flight test. Areas of consideration are the rate of descent available, the maximum approved altitude, and the time before a catastrophic failure following the loss of transmission oil pressure, a fuselage fire, or other similar failure. Each helicopter should have the capability to descend to sea level and land from the maximum certificated altitude within the time period established as safe following a critical failure.
b. Procedures.
Flight test instrumentation should include ambient parameters, all flight control positions, rotor rpm, main and tail rotor flapping (if appropriate), engine power instruments, and throttle position. Flight controls that are projected to be near their limits of authority should be rigged to the most adverse production tolerance. A very accurate weight and balance computation is needed along with a precise knowledge of the aircraft's weight/c.g. variation as fuel is burned.
Chap 2 154 Par 81
The critical condition for VNE controllability testing is ordinarily aft c.g., MC power, and minimum power-on rotor rpm, although power and rpm variations should be specifically evaluated to verify their effects.
turbine engine is sensitive to ambient temperatures which affect the engine's ability to produce rated maximun continuous torque. Flight tests conducted at ambient temperatures that cause the turbine temperature to limit maximum continuous power would not produce the same results obtained at the same density altitude at colder ambient temperatures where maximum continuous torque would be limiting. Forward c.g. should be spot checked for any "tuck under" tendency at high speed. The VNE controllability test is normally accomplished shortly after the 1.1 VNE (or 1.1 Vn) point obtained during stability tests required by § 29.175(b). Controllability must be satisfactory for both conditions. If VNE varies with altitude or temperature, VNE for existing ambient conditions is utilized for the test. Extremes of the altitude/temperature envelope should be analyzed and investigated by flight test.
(3) The critical condition for controllability testing in a hover is ordinarily forward c.g. at maximum weight with minimum power-on rotor rpm. For rearward flight testing of configurations where the forward c.g. limit varies with weight, low or high gross weight may be critical. Lateral c.g. limits should also be investigated. A calibrated pace vehicle is needed to assure stabilized flight conditions. Surface winds should be less than 3 knots throughout the test sequence. Testing can be done in higher stabilized wind conditions (gusting less than 3 knots); however, these conditions are very difficult to find and the method is very time consuming due to the necessity of waiting for stabilized winds.
Testing in calm winds is preferred. Hover controllability testing should be accomplished with the lowest portion of the helicopter at the published hover height above ground level; however, the test altitude above the ground may be increased to provide reasonable ground clearance. Although the necessary yaw response will vary somewhat from model to model, sufficient control power should be available to permit a clearly recognizable yaw response after full directional control displacement when the helicopter is held in the most critical position relative to wind.
(4) Prior to engine failure testing, it is mandatory that the pilot be fully aware of his engine, drive system, and rotor limits. These limits were established during previous ground and flight tests and they should be specified in the TIA. Particular attention should be given to minimum stablized and minimum transient rotor rpm limits. These values must be included in the TIA and should be approached gradually with a build-up in time delay unless the company testing has completely validated all pertinent aspects of engine failure testing. On Category A installations the maximum power output of each engine must be limited so that when an engine fails and the remaining engine(s) assume the additional load, the remaining engine(s) are not damaged by excessive power extraction and over-temping. This is needed for compliance with § 29.903(b). The propulsion engineer should have assured that this feature was properly addressed in the engine and drive system substantiation; however, it must be assumed that for some period of time the pilot may extract maximum available power from the remaining engine(s) when an engine fails during critical flight maneuvers. Substantiation of this feature should be accomplished primarily by engine and drive system ground tests.
Chap 2 Par 81 155
AC 29-2, CHG 1 11/19/84 * (5) Longitudinal cyclic authority at VNE with any power setting must permit suitable nose down pitching of the rotorcraft. If the remaining control travel is considered marginal, tests should include applications up to full control deflection to assess the remaining authority. Some knowledge of the aircraft's response to turbulence is useful in assessing the remaining margin. As a minimum, the rotorcraft must have adequate margin available to overcome a moderate turbulent gust and must not have any divergent characteristic which requires full deflection of the primary recovery control to arrest aircraft motion. If other controls must be utilized to overcome adverse aircraft motion, the results are unacceptable; e.g., if a pitch up tendency resulting from an actual or simulated moderate turbulent gust cannot be satisfactorily overcome by remaining forward cyclic, the use of throttle or collective controls to assist the recovery is not an acceptable procedure; however, the use of lateral oyolic to correct roll in conjuction with forward cyclic to correct pitchup is satisfactory. Obviously during the conduct of these tests, all available However, compliance with this section requires that recovery must be shown by use of only the primary control for each axis of aircraft motion. techniques should be utilized when the pilot finds himself "out of control."
(6) Cyclic control authority in autorotation must be sufficient to allow adequate flare capability and landing under the all engine inoperative requirements of § 29.75(b) (5) and (c). See paragraph 67 of this advisory circular.
82. § 29.161 (through Amendment 29-19) TRIM CONTROL.
a. Explanation.
The pilot has many tasks to perform with each hand during sustained flight conditions. The trim requirement is intended to provide the pilot with a reference cyclic control position for the given flight condition, reduce the physical demands to maintain a given flight condition, and allow the pilot to release the cyclic control for brief periods of time to perform other cockpit duties. A primary flight control which can move when released imposes an additional pilot workload by requiring a continuous hands-on condition. It is not intended to require that control forces be reduced to zero by the trim control during dynamic maneuvers such as takeoff acceleration.
characteristics. One popular method of meeting this requirement is through the use of control balance springs in conjunction with a small amount of built-in control system friction. Other methods include use of friction, magnetic brakes,
(3) This regulation is not intended to require zero friction or zero breakout force in the control system, nor is it intended to require automatic control'recentering. The regulation, in fact, specifically prohibits excessive high friction or high breakout forces which would produce undesirable discontinuities in the primary control force gradient.
Chap 2 156 Par 81
* 140. § 29.351 (through Amendment 29-19) YAWING CONDITIONS.
a. Explanation. The rule requires proof of a rotorcraft "structural" yaw 500331 or sideslip design envelope. This sideslip envelope must cover minimum forward speed or hover to VH or VNE, whichever 18 less.
structurally safe for the thrust capability of the directional control system.
(1) The rotorcraft structure must be designed to withstand the loads for the specified yaw conditions. The standard does not require a structural flight demonstration. It is a structural design standard.
(2) Maximum displacement of the directional control, except as limited by pilot effort (130 pounds; § 29.397(a)), is required for the conditions cited in the rule. A control system rate limiter or a yaw damper may be used. The total displacement is therefore a function of time as well as the maximum effort applied (130 pounds).
(1) At low airspeeds, 90° yaw (sideward flight) should be the design limit.
(11) At high airspeeds, stabilized yaw angle (stabilized sideslip) must be substantiated as stated in the rule.
(111) At high airspeeds, the maximum tail rotor thrust will be combined with the vertical (directional) stabilizer surface load, if a stabilizer is used, as specified by § 29.351 (b) (1).
(iv) At high airspeeds, while the rotorcraft is in the sideslip condition, the directional control is then returned to the neutral position, attendent with the flight condition. The tail rotor thrust will be added to the restoring force of the vertical stablizer.
(v) Both right and left yaw conditions should be proven.
(3) The tail rotor attachment structure must comply with § 29.403.
(4) The vertical stabilizing surface must also comply with § 29.413.
b. Procedures•
(1) Many of the current single main rotor rotorcraft designs have vertical (directional) stabilizing surfaces. These surfaces may be solely vertical stabilizing fins as on the Bell Model 206, or a swept vertical extension of the tail boom as on the Hiller Model FH1100. The Hiller FHl100 tail surface houses the tail rotor drive shaft and the tail rotor output gearbox.
(1) For vertical stabilizers, the airloads may be assumed independent of the tail rotor thrust.
Chap 2 Par 140 235
AC 29-2, CHG 1 11/19/84 * (11) For vertical stablizers that house the tail rotor output gearbox, such as the Hiller Model FH1100, the tail surface air loads will add to or subtract from the tail rotor thrust according to the flight condition under consideration NOTE: For one example: At stablized yaw to the right (left pedal depressed to limit) (§ 29.351(b) (2)), the tail rotor thrust moment should equal the restoring moment of the tail boom, vertical stabilizer and main rotor torque. As stated by § 29.351(b)(3), the tail rotor thrust moment then i9 added to the vertical stabilizer restoring moment. The addition of tail rotor thrust (§ 29.351(b) (3)) and vertical stabilizer load is generally one of the critical design conditions for the fuselage/tail boom.
(i1i) For vertical stablizers or fins that have an offset incidence angie with respect to the helicopter axis, the vertical fin moment is added, or subtracted as applicable, to the tail rotor thrust moment. The condition stated in § 29.351(b)(1) may result in adding the fin load to the tail rotor thrust.
(Iv) Low airspeed maneuvers, such as sideward, rearward, and hover turns over a spot, typically impose insignificant aerodynamic loads on the fugelage and/or tail boom. The aerodynamic loads at VH or VNE, whichever is required, are generally the significant aerodynamic design loads.
A rational assessment of the various yaw conditions may be used to reduce the load deviation and analysis to the critical rotorcraft design conditions.
(vi) The helicopter structure shall be analyzed or tested for loads derived from the critical design conditions.
(vi1) A simple structural design envelope may be derived from these design data. If the right or left yaw limits are not very different, common, conservative design limits may be used. A sample yaw/forward speed diagram, as derived from design analysis of the characteristics of a hypothetical helicopter, is presented in Figure 140-l. A table of values would also suffice. This figure reflects characteristics which include a 90° yaw when the directional control inputs are applied at low airspeeds (up to 30 knots presumably the maximum are applied at VH, with a straight line variation from 30 knot forward apeed to • sideward flight speed of which this aircraft is capable) and 10° yaw when they Chap 2 236 Par 140
90° -90° Yaw Degrees 10° Max.
30 knot sideward Flight Airspeed FIGURE 140-1 * (vii1) During flight test evaluations, yaw angles have been measured using a yaw angle probe (swiveling vane type) on a nose boom. Both a visual readout for the pilot and a record, such as an oscillograph trace, have been used. This test may be conducted in the flight test program or in the flight load survey program. This record should confirm the yaw angle used in design as conservative with respeot to operational and actual flight characteristics. This test is not a requirement however.
Chap 2 Par 140 237
AC 29-2, CHG 1 11/19/84
141. § 29.361 (through Amendment 29-19) ENGINE TORQUE.
a. Explanation.
(1) The rotorcraft should be designed for limit engine to que values, as prescribed by the rule, to account for maximum engine torque, including transients and torsional oscillations. The rule recognized that reciprocating (piston) engines generate higher torque oscillations than turbine engines.
(2) A factor of 1.25 applies to maximum continuous power for turbine engines.
(3) Torque factors are also specified for reciprocating engines having two or more cylinders § 29.361a.2. The appropriate torque factor applies to takeoff power torque as well as maximum continuous power and other power conditions.
(4) Paragraph 206 of this document concerns $ 29.549(e) that contains standards associated with 2 1/2-minute power rating.
D• Procedures•
(1) The engine torque associated with the maximum continuous power condition must be multiplied by the appropriate torque factor to obtain the engine torque value used for structural substantiation purposes of the rotorcraft.
(2) The torque values associated with the minimum power-on rpm limit should be used. Maximum power-on speed limit will result in a lower torque value when calculating torque from design horsepower values. However, due to piston engine power output characteristics, an engine may produce a higher torque at higher engine speeds contrary to the previous statement. The torque factor should account for this characteristic.
142.-151. RESERVED.
Chap 2 238 (thru 248) Par 141
SECTION 7. STRENGTH REQUIREMENTS - GENERAL * 121. § 29.301 (through Amendment 29-19) LOADS.
a. Explanation,
(1) The rule is a general statement concerning limit and ultimate loads, and the application of these loads to the rotorcraft.
(2) Ultimate loads are limit loads multiplied by the prescribed factors of safety.
(3) The specified loads must be distributed appropriately or conservatively and signficant changes in distribution of the loads, as a result of deflection, must be taken into account.
b. Procedures. The design criteria report and/or design loads report must contain data that comply with the rule.
122. § 29.303 (through Amendment 29-19) FACTOR OF SAFETY.
a• Explanation.
(1) Unless otherwise provided by FAR Part 29, a factor of safety of 1.5 is required and is applied as stated in the rule. This safety margin will assure that the design strength of the rotorcraft is greater than the design loads contained in FAR Part 29.
(2) Other rules, $$ 29.561 (b)(3) and 29.787(c), specify use of defined ultimate inertia forces for protection of occupants.
b. Procedures.
(1) The design criteria report and/or design loads report must contain data that include the appropriate factor of safety.
(2) The factor of safety multiplies the limit external and inertia loads. The rule does allow the application of this factor to the resulting 'limit internal" stresses if it is more conservative.
123. § 29.305 (through Amendment 29-19) STRENGTH AND DEFORMATION.
a. Explanation.
(1) This general rule defines, in relative terms, allowable deformation for limit and ultimate loads.
(2) The structure must support, in a static test, ultimate loads for 3 seconds without failure or dynamic tests simulating actual load application may * be used.
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AC 29-2, CHG 1 11/19/84
(3) Section 29.307 concerns proof of the structure and requires certain specified tests. This rule also allows substantiation by struotural analysis.
See paragraph 124 of this advisory circular.
b. Procedures. Any test results, static or dynamic, must satisfy the limitations or acceptance criteria contained in the rule.
(1) Any test proposals submitted for approval that are used to demonstrate compliance with sections of FAR Part 29 must contain the criteria stated in the rule.
(11) Any test results reports shall contain data and information showing the test results comply with the standard.
Chap 2 198 Par 123
124. § 29.307 (through Amendment 29-19) PROOF OF STRUCTURE.
a• Explanation.
(1) The rule requires compliance for each critical loading condition.
Certain tests must be conducted as specified. Additional tests for new or unusual design features may be required as noted in subparagraph (b) (6).
(2) Structural analysis may be used only if the structure conforms to those for which experience has shown this method to be reliable.
(3) Fatigue substantiation requirements are explained further in paragraph 230, § 29.571 of this advisory ciroular.
b. Procedures.
(1) The design criteria and/or design loads report should contain typical or representative loading conditions from which the oritical loading conditions will be selected for analytical substantiation in struotural (static and fatigue) reports and dynamics (vibration and stability) reports and fatigue, statio, dynamio, or operational test reports.
(2) Whenever tests are used or required, a test proposal or plan must be approved prior to the tests. The test article must have received conformity inspections and must have been acoepted by the FAA for the test. Test fixtures and instrumentation must also be acceptable to the FAA (using DER's as appropriate) prior to start of the test. The quality control office of the applicant or other qualified pesonnel may be authorized to conduct inspections of the test fixtures and instrumentation rather than the FAA or DER performing this task. The test proposal may be used to define and to authorize the means to accomplish inspection of the test fixtures and instrumentation. Unnecessary drawings, such as test fixture details, or layering of approvals is not intended or envisaged by this policy. Drawings, sketches or photographs have been used by the FAA to control and to assure correct location, direction, and magnitude of loads and other oritical test parameters.
(3) Structural analysis has been accepted for rotorcraft in place of static tests. Generally the helicopter airframe should have frequency placements remote to predominate rotor excitation sources, including rotor harmonics, to avoid undesirable and possibly excessive vibration and potentially high operating stress levels due to this vibration. During the flight load measurement program conducted under § 29.571, critical loaded areas or critical joints may be instrumented with strain gages or other stress strain measuring devices. This actual flight data may be compared to the analytical data to verify accuracy.
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AC 29-2, CHG 1 11/19/84
(4) Subparagraph (b) of the rule specifies certain tests. Test proposals must be approved prior to conducting official FAA tests. Other paragraphs in this advisory circular pertain to those tests.
125. § 29.309 (through Amendment 29-19) DESIGN LIMITATIONS.
a. Explanation.
(1) The rule requires an orderly selection and presentation of the basic structural design limitations of the rotorcraft. The applicant must establish these structural limitations to facilitate design of the rotorcraft.
(2) Refer to the rule for the specifio requirements.
b. Procedures.
(1) The design criteria and/or design load report should contain the design limits specified.
(2) These items are structural design limits. Other requirements may result in narrowing the ranges of type design limits or in reducing limits. It is not necessary to revise structural design criteria limits to agree with more conservative operational limits established during the certification program. The operational limits may be subsequently expanded by additional flight tests to agree with design limits.
126.-135. RESERVED.
Chap 2 200 (thru 230) Par 124