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.
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(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.
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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.
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АС 27-1A 7/30/97
84. § 27.171 STABILITY: GENERAL.
a. Explanation. This section is intended to require a manageable pilot workload for the minimum crew under foreseeable operating conditions.
b. Procedures.
(1) Compliance with the requirements of this section can often be obtained for the VFR condition without any specific or designated flight testing. If the rotorcraft is marginal in regard to pilot strain and fatigue, the FAA/AUTHORITY pilot should be assured, through special tests if necessary, that the aircraft can be satisfactorily flown throughout the maximum endurance capabilities of the rotorcraft including night and turbulence conditions if those are critical. This test should be conducted with minimum required systems in the aircraft and with minimum flightcrew.
(2) Reasonable failure conditions which add to pilot workload, strain, and fatigue should be evaluated (electrical, hydraulic, and mechanical failures, etc.). The necessary times associated with flight with a failed system must be appropriate to the flight manual procedures for each failure. A failure condition requiring immediate landing would obviously require shorter evaluation time than a condition allowing continued flight to destination.
(3) IFR approvals necessitate a careful evaluation of Paragraphs b (1) and (2) above. In FR operations, weather conditions frequently necessitate continued flight to destination or diversion to alternate airports with critical failures. Immediate landing may not be feasible. The evaluating pilot must ensure pilot strain and fatigue are acceptable during typical flight profiles for each type of operation to be approved.
85. § 27.173 (Amendment 27-21) STATIC LONGITUDINAL STABILITY.
a. Explanation.
(1) This rule contains control requirements for both stability and control.
Paragraph (a) contains the basic control philosophy necessary for all civil aircraft.
Forward motion of the cyclic control must produce increasing speeds, and aft motion must result in decreasing speeds. For rotorcraft, this is accomplished with throttle and collective held constant. Rotorcraft with either highly stable or highly unstable static longitudinal stability characteristics can typically comply with the basic requirement for control sense of motion. However, the intent and interpretation of this paragraph is to provide a stable stick position versus airspeed gradient. Therefore, a stabilized airspeed less than the trim speed requires a cyclic stick position aft of the trim stick position, and a stabilized airspeed greater than the trim speed requires a cyclic stick position forward of the trim speed stick position.
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(2) The remainder of § 27.173, through reference to § 27.175, contains the basic control position requirements necessary to establish a minimum level of static longitudinal stability. Positive stability is found for conditions of climb, cruise, and autorotation in § 27.175 by requiring a stable stick position gradient through a specified speed range. A defined level of instability is permitted for the hovering condition.
b. Procedures.
(1) The control requirement of this section is so essential to basic flight mechanics that compliance may be found during conventional flight testing for compliance with other portions of the regulations. No special or designated testing should be required.
(2) The procedures necessary to ensure compliance with the stability requirements of this section are contained under § 27.175. Refer to Paragraph 86 of this advisory circular for an explanation of detailed flight test procedures.
86. § 27.175 (Amendment 27-21) DEMONSTRATION OF STATIC LONGITUDINAL STABILITY.
a. Explanation.
(1) This rule incorporates the specific flight requirements for demonstration of static longitudinal stability. Specific loadings, configurations, power levels, and speed ranges are stated for conditions of climb, cruise, autorotation, and hover.
(2) Some rotorcraft in forward flight experience significant changes in engine power with changes in airspeed even though collective and throttle controls are held fixed and altitude remains relatively constant. For these cases, the guidance in § 27.173 which states that throttle and collective pitch must be held constant is appropriate for administration of this rule, and the specified power in § 27.175(a), (b), and (c) should be considered as power established at initial trim conditions. This will result in slightly higher or lower torque readings at "off trim" conditions. Collective and throttle controls are held constant when obtaining data during climb, cruise, and autorotation tests.
(3) The effects of rotor RPM on autorotative static stability should be determined and positive stability demonstrated for the most critical RPM. Values for RPM can be expected to change as airspeed is varied from the "trimmed" condition.
The manufacturer's recommended autorotation airspeed is ordinarily used for trim.
(4) Hovering is considered a flight maneuver for which the pilot repeatedly adjusts collective to maintain an approximately constant altitude above the ground. For hover stability tests, collective and throttle adjustments are made as necessary to Par 85 123
АС 27-1А 7/30/97 maintain an approximately constant height above the ground. Also, a limited amount of negative longitudinal control travel is allowed with changes in speed.
b. Procedures
(1) Instrumentation.
Sensitive control position instrumentation is mandatory. Engine power parameters should be recorded at trim. For testing of minor modifications or when using a "before and after" method, a tape measure or a stick plotting board may be utilized. A stick plotting board consists of a level surface with a clean sheet of paper on it attached to the cockpit or seat structure. The installation must not interfere when the flight controls are fully displaced. A recording pencil is attached to the cyclic control by an offsetting arm in such a manner that it can be pushed down on the board to record relative cyclic position at key times during test maneuvers. The Figure 86-1 plot is a typical presentation of longitudinal static stability.
(її) Other necessary parameters include pressure altitude, ambient temperature, and indicated airspeed (pace vehicle or theodolite speed for hover tests).
For hover tests, hover height (radar altitude if available) and surface winds should be documented. Two-way communication with a pace vehicle is highly desirable. Ground safety equipment is desirable.
(2) Ambient Conditions. Smooth air is necessary for stability testing. Allowable wind conditions for hover stability testing are the same as those for hover controllability tests. Extrapolation is covered in Paragraph 58 of this advisory circular.
(3) Loading. Aft center of gravity (CG) is ordinarily critical for longitudinal stability testing, although high speed flight and hover should be checked at full forward CG and maximum weight. At aft CG, light or heavy weight conditions can be critical.
The manufacturer's flight data should be reviewed to determine critical loading conditions.
(4) Conducting The Test.
The rotorcraft should be established in the desired configuration and flight condition (climb, cruise, autorotation) with the required power and rotor speed at the trim airspeed. The collective stick should be fixed in that position, usually by applying sufficient friction to ensure that it is not inadvertently moved. For autorotative tests, a rotor speed should be selected so that the variations in rotor speed as airspeed and altitude change do not exceed the allowable limits. This point is recorded as the trim point. Airspeed is then increased or decreased in about 10-knot increments, stabilizing on each speed and recording the data. At least two points on each side of the trim speed should be taken.
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(i¡) The cruise test should be accomplished by first determining V. (level flight speed at maximum continuous power) at the test altitude. Then reduce power to establish a level flight trimmed condition at 0.9 VH (or 0.9 VNE if lower). This point is then recorded as the trim point. The collective pitch and throttle must remain fixed at the trim setting for the remainder of the test. The airspeed is then varied above and below the trim speed using the cyclic control to climb or dive slightly.
(iii) For climb and autorotation tests, conduct fixed collective tests through an altitude band (usually ‡2,000 feet), first increasing airspeed as data points are collected, then decreasing speed through the same altitude band. It will probably not be possible to obtain the required data on one pass through the altitude band. If repeated passes are required, a trim point should be taken at the beginning of each pass unless very sensitive collective pitch position information is available in the cockpit. Generally, it will be possible to acquire all the high speed points on one pass and the low speed points on the second.
(iv) If extremely precise results are required, an alternate method of testing can be used to acquire the data at a constant altitude. For cruise, data can be obtained by alternating airspeeds above and below the trim speed to arrive in the vicinity of the test altitude as the point is recorded. This method results in very precise data because collective and throttle are not moved as airspeed is changed at a constant altitude. A typical sequence of speeds that could produce these results would be: 150 (V), 135 (0.9Vu) trim speed, 125, 145, 115, 155, 105, and 165.
(v) For rotorcraft with high rates of climb, a series of climbs, each at a different speed, may be required through a given altitude, utilizing sensitive instrumentation to ensure collective position is the same for each data point. In autorotation, a similar case arises and a series of descents, each at a different speed, may be required through a given altitude band, using sensitive instrumentation to ensure a repeatable collective position.
(vi) Hover tests should be conducted by maintaining an approximately constant altitude above the ground at the hover height established for performance purposes. The test altitude above the ground may be increased to provide reasonable ground clearance during rearward flight. Groundspeed is varied using a pace vehicle, theodolite, or other velocity measuring equipment. A pace vehicle is an aid in maintaining an accurate hover height. The pilot can accurately maintain height by controlling his sight picture of the pace vehicle (level with the roof, antenna, etc.).
Hover stability tests are ordinarily conducted in conjunction with hover controllability tests because instrumentation and facilities are essentially the same.
(vii) Normally, climb, cruise, and autorotation tests should be conducted at low, medium, and high altitudes. See Paragraph 58 for guidance on interpolation and extrapolation. High speed stability has been critical during cold weather testing. In two recent models, VNE at cold temperatures has been limited by the stability requirements Par 86 125
АС 27-1A 7/30/97 of § 27.175(b). Cold weather testing should be accomplished or a conservative approach for advancing blade tip Mach number should be used to limit cold weather VNE to tip Mach number values demonstrated during warm weather testing.
(viii) Hover stability should be verified at low altitude and, if required, at high altitude. Refer to Paragraph 58b(2) for guidance on expansion and extrapolation of altitude.
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O CRUISE
• CLIMB
A AUTOROTATION
• HOVER
SHADED POINTS REPRESENT TRIM 100 90 0000000 000000 LONGITUDINAL CYCLIC POSITION (% FROM FULL AFT) • 20 10 20 40 60 80 100 120 140 160
CALIBRATED AIRSPEED (KNOTS) FIGURE 86-1. STATIC LONGITUDINAL STABILITY Par 86 127
87. § 27.177 (Amendment 27-21) STATIC DIRECTIONAL STABILITY.
a. Explanation. This rule requires that positive static directional stability be demonstrated at the trim airspeeds defined in § 27.175. The trim speed for climb is Vy and for cruise 0.9VH or 0.9VNE (whichever is less).
b. Procedures.
(1) Tests for static directional stability require instrumentation for pedal position and sideslip angle. To obtain accurate sideslip angle and airspeed information, a "yaw boom" is usually installed for the purpose of mounting a sideslip vane and swiveling airspeed pilot head outside the main rotor downwash region of influence. Special care should be taken to ensure that the yaw boom installation has been verified to be structurally adequate and free of dynamic instabilities for all combinations of airspeed and rotor speed likely to be experienced during the static directional evaluation. For some installations, the instrumentation yaw boom may influence the flying qualities of the rotorcraft itself. Thus, it is advisable to correlate yaw string displacement or slip indicator ball widths of skid with yaw boom sideslip angle, and then repeat a few critical points with the yaw boom removed.
(2) For some rotor system designs, the main and tail rotor flapping angle may be a critical instrumentation requirement for static directional testing. Both main and tail rotor flapping may increase dramatically at high airspeeds with increasing sideslip angle. Therefore, for rotor systems exhibiting this characteristic, flapping should be monitored carefully during the sideslip maneuver to avoid exceeding limitations. Static directional stability is normally defined in terms of pedal displacement required to maintain a steady heading sideslip. A single-rotor rotorcraft flying in coordinated flight will exhibit a small inherent sideslip due to tail rotor thrust and fuselage/main rotor sideforces. This condition is normally taken as trim with the inherent sideslip angle noted. Airspeeds should be the trim values described above. The procedures used to establish and maintain the steady heading sideslip can significantly influence test results. A generally accepted technique follows:
(1) Stabilize at the trim point, and note indicated airspeed.
Record trim conditions including inherent sideslip. Maintain fixed collective and throttle for the remainder of the maneuver.
(Ш1) Smoothly apply directional control and coordinate with lateral control to establish the desired sideslip angle. A steady track can best be ensured by maintaining a track over a straight landmark on the ground such as a section line or straight segment of powerline or highway.
(iv) Because drag increases with sideslip, the aircraft will decelerate. The trim airspeed should be maintained by entering a slight dive or decreased rate of climb.
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