* 16. § 21.35 FLIGHT TESTS.
a • Explanation•
(l) This section outlines the requirements of the applicant for aircraft type certification and should be used in conjunction with Order 8110.4, Section
5. section zi.35 requires, in part, that the applicant conduct sufficient flight tests to show compliance with the flight requirements throughout the proposed flight envelope. The results of the applicant's flight test should be submitted to the FAA in report form for evaluation to determine what verification flight tests the FAA may elect to conduct. The report should conclude that in the applicant's opinion the test aircraft complies with the applicable certification requirements. The FAA verification flight test should include, but not be limited to, the critical or marginal results contained in the applicant's flight test report. The FAA's role in the certification effort is not envisioned to be one of conducting day-to-day routine flight tests with the applicant, but only to verify his results through limited sampling. In certain tests, such as high altitude testing at a remote mountain site, there is an advantage in conducting flight tests concurrently with the applicant. Additionally, the FAA can provide technical flight test assistance to the applicant in certain cases. This can be done after a cursory review and a letter of authorization is issued to the flight test crew.
(2) Preflight Test Planning. After the applicant's flight test report is reviewed, it should be determined what FAA engineering flight tests are necessary. These tests are normally specified in the Type Inspection Authorization (TIA). At the same time the FAA must know and agree to the applicant's proposed means of data aquisition, reduction, and expansion of the flight test data, The adequacy of the test instrumentation should be evaluated prior to official type certification tests (reference paragraph 24).
(3) Order of Testing. The Federal Aviation Regulations are so worded that the results of some flight tests have a definite bearing on the conduct of other tests. For this reason, and to minimize retesting, careful attention should be given to the order of testing. The exact order of testing will be determined only by considering the particular rotorcraft and test program involved. lests which are particularly important in the early stages of the program are:
(1) Airspeed calibration: All tests involving airspeed depend upon the calibration.
(11) Engine power available determination.
(111) Engine cooling.
(4) Test Groupings.
Weight and c.g.: In addition to the regulatory relationship of one test to another, efficient testing requires that consideration be given to the acoomplishment of as many tests on a single flight as can be accommodated successfully.
Chap 1 Par 16 21
AC 29-2, CHG 1 11/19/84 * (11) Special instrumentation. Similarly, consideration should be given to grouping of tests that involve special instrumentation. Examples of these are takeoff and landing tests which usually require group equipment to record horizontal distance, height, and time. Ground calibration of the airspeed indicating system can be accomplished at the same time. It is the applicant's responsibility to provide the necessary instrumentation.
b. Procedures.
(1) Type Certification Flight Tests.
(1) Prior to initiating official FAA flight tests, a conformity inspection of the test aircraft must be accomplished. This is needed to assure that the test aircraft is in the proper configuration or "conforms" to the engineering drawings and documents that have been submitted to FAA, evaluated, and approved. It is absolutely essential to know the configuration being tested in any engineering flight evaluation. Conformity inspection prior to TIA flight tests assures that testing will not be wasted because of configuration uncertainties.
Certification Handbook 8110.4, paragraph 67, contains a requirement that the applicant must keep the FAA advised of any configuration changes to the aircraft. The manufacturing inspector should keep the FAA flight test pilot apprised of any change which may affect safety of the test aircraft or may influence test results.
(111) Results of the conformity inspection and the engineering flight test program must be documented. This is normally done in the Type Inspection Report (TIR). Results may be documented in any acceptable engineering format.
The report should be in sufficient detail to clearly show how compliance with each appropriate section of the rules was determined.
(iv) The flight test pilot must assure that the FAA manufacturing inspector and certification engineer are aware of all configuration changes found necessary as a result of FAA tests. The manufacturing inspector is responsible for assuring that all changes are incorporated into production drawings after the design data reflecting the change have been approved by the certification engineer.
Additional flight test responsibilities, procedures, and requirements during the certification flight test process are contained in Certification Handbook 8110.4, Section 5, Flight.
(2) Function and Reliability Tests.
(1) A comprehensive and systematic check of all aircraft components must be made to assure that they perform their intended function and are reliable.
(11) Function and reliability (F&R) testing must be accomplished on an aircraft which is in conformity with the approved production configuration.
F&R testing should follow the type certification testing described in paragraph
(1) above to assure that significant changes resulting from type certification tests can be incorporated on the aircraft prior to F&R tests.
Chap 1 22 Par 16
11/19/84 ILLUSTRATIONS AC 29-2, CHG 1 The first block of digits in the figure number denotes the associated paragraph number. The second block of digits denotes the figure number within the paragraph. The figure is located at the end of the paragraph.
Figure No. Title Page 58-1 Shaft Horsepower vs. Turbine Outlet Temperature - Sea Level Standard Day 99 58-2 Uninstalled Takeoff Power Available 100 58-3 Installed Takeoff Power Available 101 Power Assurance Check Chart 102 Takeoff Performance Category A 105 Category A Vertical Takeoff Profile, Ground Level Heliport 112 Category A Vertical Takeoff Profile, Pinnacle 113 Conventional Takeoff Profile, Category B 116 Autorotational Characteristics - Typical 121 67-1 Category A Conventional Landing - Clear Heliport 132 67-2 Category A Vertioal Landing 132 69-1 Height-Velocity (HV) Diagram 139 69-2 Altitude/Weight Accountability 140 85-1 Static Longitudinal Stability 162 '110~7 Vibration Demonstration Points 195 137-1 Load Factor - Gross Weight Curve 232 249-1 Example of Casting Load Sheet, Retract Actuator Support - 407 Landing Gear 386-1 Continuous Icing - Temperature vs. Altitude Limits 649 386-2 Intermittent Icing - Temperature vs. Altitude Limits 650 386-3 Continuous Icing - Liquid Water Content vs. Drop Diameter 651 386-4 Intermittent Icing - Liquid Water Content vs. Drop Diameter 652 425-1 Cantilevered Shaft, First Critical Speed 687 425-2 Shaft Between Support Bearings, First Critical Speed 687 * 516-1 Additional Cooling Test Point 815 518-1 All-Engine-Critical Altitude 821 * 775-1 Helicopter Dynamic Stability Requirements for IFR 1252 XIII
* (111) All components of the rotorcraft should be periodically operated in sequences and combinations likely to ocour in service. Ground inspections should be made at appropriate intervals to identify potential failure conditions;
however, no special maintenance beyond that described in the aircraft maintenance manual should be allowed.
(iv) A complete record of defects and failures should be maintained along with required servicing of aircraft fluid levels. Results of this record should be consistent with inspection and servicing information provided in the airoraft maintenance manual.
(v) A certain portion of the F&R test program may emphasize systems, operating conditions, or environments found particularly marginal during type certification tests.
(v1) See Handbook 8110.4, paragraph 166(c), for additional information and procedures.
17.-23. RESERVED.
24. § 21.39 FLIGHT TEST INSTRUMENT CALIBRATION AND CORRECTION REPORT •
a. Explanation. It is the applicant's responsibility to provide instrumentation for all parameters needed to show compliance with the airworthiness regulations.
(1) For those data which are necessary to show compliance with the regulations, a permanent record should be established. A permanent record is acceptable in either graphical or photographic form, and in some instance a manual recording may be satisfactory.
(2) Regardless of the record form, the accuracy of the record must be established by reference to a laboratory standard traceable to the National Bureau of Standards.
(3) If multiplexing is used, the time base must be synchronized to a reference point from which the magnitude of each parameter can unquestionably be determined. Also, the sampling rate should be sufficiently frequent to assure that the maxima, minima, and trends of magnitude of the parameter are recorded with respect to time.
b. Procedure. Prior to conducting flight tests, the FAA flight test team should review the applicant's flight test instrumentation calibration and correction report.
Chap 1 Par 16 23
AC 29-2, CHG 1 11/19/84 * (1) Normally the frequency of instrument calibration should not exceed 90 days. However, the frequency of recalibration varies with the consistency of the instrumentation under consideration. For example, cyclic and collective position is sometimes calibrated immediately before and after a flight where these parameters are used to provide oritical flight data. Six months is a typical interval for recording/signal conditioning and nonstrain gage sensors, while one year 18 typical for strain gaged components. Also, environmental effects such as vibration, humidity, temperature, etc., should be considered when determining whether recalibration is necessary.
(2) The highest and lowest magnitude of the parameter being recorded should be considered when establishing the scale for instrumentation. Ideally, the highest magnitude throughout the flight would fall on the maxinum indicating point of the recording. •
25.-30. RESERVED.
Chap 1 24 (thru 60) Par 24
SECTION 4. FLIGHT CHARACTERISTICS
80. § 29.141 (through Amendment 29-19) FLIGHT CHARACTERISTICS - GENERAL.
a. Explanation,
(1) This regulation prescribes the general flight characteristics required for certification of a transport category helicopter. Specifically, it states that the helicopter shall comply with the flight oharacteristics the reference to "altitude" in § 29.141(a) (1) is correctly interpreted as "density altitude ‚' Density altitude is, of course, a function of pressure altitude and ambient temperature, hence the need to account for ambient temperature effects.
Additional fiight characteristics required for instrument flight are contained in paragraph 775 of this advisory circular.
(2) Generally the aircraft structural (load level) survey accounts for inoreases with airspeed and power. If the takeoff power rating exceeds the maximum continuous power rating, and the structural survey has been conducted under the assumption that takeoff power is not used at speeds above Vy, the otorcraft Flight Manual must limit takeoff power to speeds of Vy and below. I akeoff power is structurally substantiated throughout the flight envelope, and appropriate portions of the controllability, maneuverability, and trim requirements of §$ 29.141 through 29.161 are met at takeoff power levels, no flight manual entry is needed. Obviously if transmission limits for MC and
(3) During the flight characteristics testing, the controls must be rigged in accordance with the approved rigging instructions and tolerances. The control system rigging must be known prior to testing. In addition to the normal rigging procedures, any programmed control surfaces which may be operated by dynamic pressure, electronics, etc., must also be calibrated. During the flight test program, it is frequently necessary to rig a control, such as the swashplate or tail rotor blade angle, to the allowable critical extreme of the tolerance band. For example, it would be necessary to rig the tail rotor to the minimum allowable blade angle if meeting the requirements of § 29.143(c) would be in question. The same consideration must be given to all helicopter controls and moveable aerodynamic surfaces where questionable compliance with the regulations may exist. If the rotor-induced vibration characteristics of the helicopter are significantly affected and require time-consuming rigging for such things as acceptable ride comfort, then the rotor(s) should be rigged to the allowable extreme tolerance limits to determine compliance, for example, with § 29.251.
AC 29-2, CHG 1 11/19/84
(4) During the FAA flight test program, the crew should be especially alert for conditions requiring great attentiveness, high skill levels, or exceptional strength. If any of these features appear marginal, it is advisable to obtain another pilot's opinion and to carefully dooument the results of these evaluations. Section 29.141(b) provides the regulatory basis for these strength and skill requirements. The general requirements for a smooth transition capability between appropriate flight conditions are also included in § 29.141(b). These requirements must also be met during appropriate engine failure conditions for each category of rotorcraft.
(5) For night or IFR approval, § 29.141(c) contains the general regulatory reference which requires additional characteristics for night and IFR flight. The appropriate flight test procedures are included in other portions of this order.
81. § 29.143 (through Amendment 29-19) CONTROLLABILITY AND MANEUVERABILITY.
a. Explanation.
(1) This regulation contains the basic controllability requirements for transport rotorcraft. It also specifies a minimum maneuvering capability for required conditions of flight. The general requirements for control and for maneuverability are summarized in § 29.143(a) which is largely self-explanatory.
The hover condition is not specifically addressed in § 29.143(a) (2) so that the general requirement may remain applicable to all rotorcraft types, including those without hover capability. For helicopters, the hover condition clearly applies under "any maneuver appropriate to the type."
*
(2) Paragrapha (b) through (e), § 29.143, include more specific flight conditions and highlight the typical areas of concern during a flight test program.
Section 29.143(b) specifies flight at VNE with oritical weight, center of gravity (c.g.), rotor rpm, and power. Adequate cyclic authority must remain at VNE for nose down pitohing of the rotorcraft and for adequate roll control. Nose down pitohing capability is needed for control of gust response and to allow necessary flight path changes in a nose down direction.
Roll control is needed for gust response and for normal maneuvering of the aircraft. In the past, 10 percent control margin has been applied as an appropriate minimum control standard. The required amount of control power, however, has very little to do with any fixed percentage of remaining control There are foreseeable designs for which 5 percent remaining is adequate and others for which 20 percent may not be enough. The key is, can the remaining longitudinal control travel at VNE generate a clearly positive nose down pitching moment, and will the remaining lateral travel allow at least 30-degree banked turns at reasonable roll rates? Moderate lateral control reversals should be included in this evaluation and since available roll control can diminish with sideslip, reasonable out of trim conditions (directionally) should be investigated. This "control remaining" philosophy must also be applied for other flight conditions specified in this section.
Chap 2 152 Par 80
(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