(10) Night Evaluation. If a rotorcraft is to be certified for night operation, a night evaluation is required. Simulated engine failures should be conducted along the recommended takeoff path. Landings should also be qualitatively evaluated with an engine failed. Engine failures at critical HV conditions are not required. The intent is to show adequate visibility using aircraft and/or runway lights without requiring a duplication of the daytime HV test program.
(11) Water Landings. For amphibious float-equipped rotorcraft, day and night water landings should be conducted under critical loading conditions with an engine failed. Engine failures should be conducted along the recommended takeoff path.
Engine failures at critical HV conditions are not required. The intent is to show similarity to test results over land without requiring a duplication of the HV test program.
Par 69 100
7/30/97/ AC 27-1A HEIGHT LEVEL FLIGHT (H)
- KNEE
TAKEOFF
TAKEOFF CORRIDOR
HIGH SPEED
SPEED (V) FIGURE 69-1. HEIGHT - VELOCITY DIAGRAM Par 69 101
• Approx.
Constant GWo Height Velocity LA- Weight
CONSTANT HV DIAGRAM, VARIABLE WEIGHT Increasing Density Altitude Height Velocity
- B
CONSTANT WEIGHT Increasing Gross Weight Height Velocity
- C-
CONSTANT DENSITY ALTITUDE FIGURE 69-2. ALTITUDEMEIGHT ACCOUNTABILITY Par 69 102
69A. § 27.79 (Amendment 27-21) LIMITING HEIGHT-SPEED ENVELOPE.
a. Explanation. Amendment 27-21 to the regulation redefines the required weight for establishing the HV envelope at altitudes above sea level.
b. Procedures. All of the policy material pertaining to this section remains in effect. In addition, the following applies:
(1) The rotorcraft height-velocity envelope should be established for the maximum gross weight at sea level. At altitudes above sea level, the envelope should be established at not less than the maximum operating weight or OGE hover weight, whichever is lower. If a weight below the OGE hover weight is selected, by definition, that selected weight becomes the maximum operating weight for the rotorcraft at that altitude.
(2) If the HV envelope is established for a maximum altitude less than 7,000 feet, by definition, the maximum takeoff and landing altitude for the rotorcraft may be no higher than that maximum HV altitude. Hover performance information should not be presented for altitudes above the maximum altitude for which the HV envelope is established.
70.-79. RESERVED.
Par 69A 103(thru 113)
SECTION 4. FLIGHT CHARACTERISTICS
80. § 27.141 (Amendment 27-21) GENERAL.
a. Explanation.
(1) This section prescribes the general flight characteristics required for certification of a normal category rotorcraft. Specifically, it states that the rotorcraft shall comply with the flight characteristics requirements at all approved operating altitudes, gross weights, center of gravity locations, airspeeds, power, and rotor speed conditions for which certification is requested. The reference to altitude in § 27.141(a)(1) refers to 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 flight characteristics required for instrument flight are contained in Paragraph 775 of this advisory circular.
(2) Generally, the aircraft structural (load level) survey accounts for takeoff power values at speeds up to and including Vy. At speeds above Vy, maximum continuous power is assumed. Stress to rotating components usually increases 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 Rotorcraft Flight Manual must limit takeoff power to speeds of Vy and below. If takeoff power is structurally substantiated throughout the flight envelope, and appropriate portions of the controllability, maneuverability, and trim requirements of §§ 27.141 through 27.161 are met at takeoff power levels, no flight manual entry is needed. Obviously if transmission limits for maximum continuous (MC) and takeoff power coincide, no special action is needed.
(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 §27.143(c) would be in question. The same consideration must be given to all rotorcraft controls and movable aerodynamic surfaces where questionable compliance with the regulations may exist. If the rotor-induced vibration characteristics of the rotorcraft 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 § 27.251.
114 Par 80
7/30/97 АС 27-1A
(4) During the FAA/AUTHORITY 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 document the results of these evaluations.
Section 27.141(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 § 27.141(b). These requirements must also be met during appropriate engine failure conditions for each category of rotorcraft.
(5) For night or IFR approval, § 27.141 (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. § 27.143 (Amendment 27-21) CONTROLLABILITY AND MANEUVERABILITY.
a. Explanation.
(1) This regulation contains the basic controllability requirements for normal category rotorcraft. It also specifies a minimum maneuvering capability for required conditions of flight. The general requirements for control and for maneuverability are summarized in § 27.143(a) which is largely self-explanatory. The hover condition is not specifically addressed in § 27.143(a)(2) so that the general requirement may remain applicable to all rotorcraft types, including those without hover capability. For rotorcraft, the hover condition clearly applies under "any maneuver appropriate to the type."
(2) Paragraphs (b) through (e), § 27.143, include more specitic flight conditions and highlight the typical areas of concern during a flight test program.
(i) Section 27.143(b) specifies flight at VNE With critical weight, center of gravity (CG), rotor RPM, and power. Adequate cyclic authority must remain at VNE for nosedown pitching of the rotorcraft and for adequate roll control. Nosedown pitching capability is needed for control of gust response and to allow necessary flight path changes in a nosedown direction. Roll control is needed for gust response and for normal maneuvering of the aircraft. In the past, 10 percent control travel 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 travel. 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 nosedown pitching moment, and will the remaining lateral travel allow at least 30° 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.
Par 80 115
(it) Section 27.143(c) requires a minimum control capability for hover and takeoff in winds of 17 knots from any azimuth. Control capability in wind from zero 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 from standard sea level conditions to the maximum altitude capability of the rotorcraft or 7,000 feet, whichever is less. On rotorcraft 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.
Section 27.143(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 transport Category A, demonstration of sudden complete single-engine failure is required at critical conditions throughout the flight envelope including hover, takeoff, climb at Vy, and high speed flight up to VNE. Entry conditions for the first engine failure are engine or transmission limiting maximum continuous power (or takeoff power where appropriate) including reasonable engine torque splits. For multiengine Category A installations (three 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 § 27.75(b).
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 1 second or normal pilot reaction time, whichever is greater.
(B) For rotorcraft without transport Category A engine isolation, 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, and climb at Vy. Maximum continuous power is specified prior to the failure for the cruise condition. Power levels appropriate to the maneuver should be used for other conditions. The corrective action time delay for the cruise failure should be 1-second or normal pilot reaction time (whichever is greater). Cyclic and directional control motions which are apart of the pilot task of flight path control are normally not subject to the 1-second restriction; however, the delay is always applied to the collective control for the cruise failure. If the aircraft flying qualities and cyclic trim configuration would encourage routine release of the cyclic control to complete other cockpit tasks during cruise flight, consideration should be given to also holding cyclic fixed for the 1-second delay. Although the same philosophy could be extended to the directional controls, the likelihood of the pilot having his feet away from the pedals is much lower, unless the aircraft has a heading hold feature. Rotor speed at execution of the cruise condition power failure should be the minimum power-on value. The term 116 Par 81
7130/97 AC 27-1A "cruise" 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. If a takeoff power rating is authorized for hover or takeoff, engine failure testing must also be accomplished for those conditions. Following power failure, rotor speed, flapping, and aircraft dynamic characteristics must stay within structurally approved limits.
(iv) Section 27.143(e) addresses the special case in which a VNE (power-off)is 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 NE (power-off) in a controlled manner with normal pilot reactions and skill. There is, however, no controllability requirement for stabilized power-off flight at speeds above 1.1 VNE (power-off) when VNE (power-off) is established per § 27.1505(c).
(v) 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° 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 § 27.143(b), 1.1 VNE at power for 0.9 VH per § 27.175(b) or § 27.1505, and to 1.1 VNE (power-off) in autorotation per § 27.143(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 circular. 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 rotorcraft was certificated with a rate of descent limitation to preclude operation in the control-limited area.
(vi) An evaluation of the emergency descent capability of the rotorcraft 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 or other similar failure. Each rotorcraft 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. If the time period does not permit a sea level landing, the maximum height above the terrain must be specified in the limitation section of the Rotorcraft Flight Manual.
Par 81 117
(3) The required controllability and maneuvering capabilities must also be considered following the failure of automatic equipment used in the control system (§ 27.672). Examples include stability augmentation systems (SAS), stability and control augmentation systems (SCAS), automatic flight control systems (AFCS), devices to provide or improve longitudinal static stability such as a pitch bias actuator (PBA), yaw dampers, and fly-by-wire elevator or stabilator surfaces. These systems all use actuators of some type, and they are subject to actuator softover and hardover malfunctions. The flight control system should be evaluated to determine whether an actuator jammed in an extreme position would result in reduced control margins.
Generally, if the flight control system stops are between the actuator and the cockpit control, the control margin will be affected. If the control stops are between the actuator and the rotor head, the control margins may not be affected, but the location of the cockpit control may be shifted. This could produce interference with other items in the cockpit. An example of this would be a lateral actuator jammed hardover causing a leftward shift in the cyclic stick position. Interference between the cyclic stick, the pilot's leg, and the collective pitch control could reduce the left lateral control available and reduce left sideward flight capability. In the case of fly-by-wire surtaces, both the high speed forward flight controllability and the rearward flight capabilities could be affected.
Flight control systems that incorporate automatic devices should be thoroughly evaluated for critical areas. Every failure condition that is questionable should be flight tested with the appropriate actuator fixed in the critical failure position. These failures may require limitations of the flight envelope. Any procedure or limitation that must be observed to compensate for an actuator hardover and/or softover malfunction should be included in the Rotorcraft Flight Manual.
b. Procedures.
(1) 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/CG variation as fuel is burned.
(2) The critical condition for VNE controllability testing is ordinarily aft CG, MC power, and minimum power-on rotor RPM, although power and RPM variations should be specifically evaluated to verify their effects. The turbine engine is sensitive to ambient temperatures which affect the engine's ability to produce rated maximum 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 CG 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 1.1VH) point obtained during stability tests 118 Par 81
required by § 27.175(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 CG at maximum weight with minimum power-on rotor RPM. For rearward flight testing of configurations where the forward CG limit varies with weight, low or high gross weight may be critical. Lateral CG 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 rotorcraft 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 rotorcraft 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 stabilized 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 § 27.903(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.
(5) Longitudinal cyclic authority at VNE With any power setting must permit suitable nosedown 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 Par 81 119
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 cyclic to correct roll in conjunction with forward cyclic to correct pitchup is satisfactory.
Obviously during the conduct of these tests, all available techniques should be utilized when the pilot finds himself "out of control." However, compliance with this section requires that recovery must be shown by use of only the primary control for each axis of aircraft motion.
(6) Cyclic control authority in autorotation must be sufficient to allow adequate flare capability and landing under the all-engine-inoperative requirements of § 27.75(b).
See Paragraph 67 of this AC.
82. § 27.151 (Amendment 27-21) FLIGHT CONTROLS.
a. Explanation. Excessive breakout or preload in the flight controls produces control system force discontinuities which result in increased workload and even controllability problems for the pilot. Similarly, excessive freeplay results in lost motion which increases pilot workload and, in an extreme case, could lead to a hazardous pilot-induced oscillation. In some designs friction can provide a positive contribution to the function of the flight controls (e.g., masking aerodynamic teedback in reversible systems). At some point, friction will have a detrimental effect on the pilot's ability to properly control the machine. In the case of an irreversible design equipped with an artificial force feel system in pitch and roll, excessive friction can mask a shallow force gradient making positive stick centering and control force static stability difficult if not impossible to demonstrate. In such an instance, the initial choice of fixes might include implementation of a steeper force gradient or addition of a force preload. Unfortunately, these solutions often lead to the kinds of problems discussed earlier. Care must therefore be exercised during the initial design phase to ensure that the components and characteristics of the flight control system are well matched.
b. Procedures. Regardless of the flight control system sophistication, it is important that the test pilot understand the system configuration prior to flight evaluation. Appropriate mechanical characteristics should be documented. For VFR aircraft, the mechanical characteristics are typically assessed in flight on a qualitative basis. If a controllability or workload problem is identified, a more detailed investigation would be necessary. Since IF certification rules include specific trim and force requirements, a more quantitative investigation of mechanical characteristics is normally conducted. The constantly varying feedback forces of reversible flight control systems generally make such designs unsuitable for IFR application. Irreversible system mechanical characteristics can often be partially documented on the ground with external hydraulic and electrical power supplies connected to the aircraft.
120 Par 81
Knowledge of the breakout, friction, and force gradient characteristics prior to flight can be useful to the pilot during flight evaluation of the system.
83. § 27.161 (Amendment 27-21) TRIM CONTROL.
a. Explanation.
(1) The pilot has many tasks to perform with each hand during sustained flight conditions. The trim requirement is intended to reduce the physical demands to maintain a given flight 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.
(2) A number of devices may be used to produce the necessary trim 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, bungees, and irreversible mechanical schemes.
(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.
b. Procedures.
(1) If comprehensive company flight test data are available, compliance with this requirement can quickly be found by spot checking extreme center of gravity loadings. Trim tests can ordinarily be done during the course of other flight test activities. To conduct the test, briefly release the control at the required flight conditions and determine that the control does not move. The words "any appropriate speed" ordinarily include any speed from hover to VH. If the control system trim device might be subject to temperature or humidity effects, these should be investigated at a minimum of two altitude extremes and during several test phases.
(2) If a pilot controllable variable friction device is incorporated, compliance with this requirement must be shown at the minimum adjustable value. The maximum value of adjustable friction should not completely lock the flight controls.
(3) Continued compliance with this requirement should be ensured through a production procedure. If minimum friction or centering springs are used, it is desirable for the manufacturer to include some adjustment capability for production differences.
The explanation and procedures discussed here are applicable for VFR approval under § 27.161. For additional FR trim requirements, refer to Paragraph 775 of this advisory circular.
Par 82 121