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. Prior to Amendment 29-24, hover controllability, height-velocity, and hover performance were the three regulatory requirements that ordinarily determined the shape of the limiting weight-altitudetemperature (WAT) curve for takeoff and landing. For category A performance rotorcraft operations, of course, the one engine inoperative (OEI) climb performance requirements may also influence the WAT limit curve. Amendment 29-24 allows, under certain conditions, the deletion of any hover controllability condition determined under § 29.143(c) from becoming an operating limitation. Section 29.1587 of Amendment 29- 24 provides a means wherein category B certificated rotorcraft (in accordance with the requirements of § 29.1, effective with Amendment 29-21) may not be limited by the hover controllability requirements of § 29.143(c). Section 29.1583(g) requirements for category A certificated rotorcraft are unchanged from past regulatory requirements in that if the hover controllability requirements of § 29.143(c) result in the most restrictive envelope it will be published as an operating limitation. Section 29.1587(b) provides a means wherein category B certificated rotorcraft, as defined in § 29.1, may not be restricted in its utilization. It allows such rotorcraft to publish the maximum takeoff and landing capabilities of the rotorcraft, provided something other than the 17 knot hover controllability requirement is not limiting. This may be zero wind IGE hover performance or any other performance the applicant elects to use if the maximum safe wind for operations near the ground is provided. Rotorcraft certificated prior to Amendment 29- 24 can update their certification basis to take advantage of this provision. If an applicant with a previously type certificated rotorcraft elects to update to this later amendment, caution should be taken to verify that the HV information is done in accordance with Amendment 29-21; that all engine out landing capabilities are satisfactorily accounted for at the new proposed gross weight, altitude, temperature combinations; that takeoff and landing information is provided; and that sufficient information is provided to properly advise the crew of the rotorcraft’s capabilities when utilizing this increased performance capabilities.
(iii) 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, climb at VY, and high speed flight up to VNE. Entry c onditions for the first engine failure are engine or transmission limiting maximum continuous power (MCP) (or take-off 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 the 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).
Page B -104.1
AC 29-2C, Chg 4 5/1/2014 The conditions for last-engine failure are MCP, 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 category B powerplant installation rotorcraft, demonstration of sudden complete power failure is required at critical conditions throughout the flight envelope. This includes speeds from zero to V NE (power-on) and conditions of hover, takeoff and climb at VY. MCP 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 part 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’s feet being 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 “cruise” also includes cruise climb and cruise descent conditions. Normal pilot reaction times are used elsewhere. Although this requirement specifies MCP, it does not limit engine failure testing to MCP. If a 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 characteristics must stay within structurally approved limits.
(iv) Section 29.143(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 VNE (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 § 29.1505(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 V NE at MCP per § 29.143(b), 1.1 VNE at power for 0.9 VH per § 29.175(b) or § 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 Page B – 104.2
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 (RFM).
(3) The required controllability and maneuvering capabilities must also be considered following the failure of automatic equipment used in the control system (§ 29.672). 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 surfaces, 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 or softover malfunction should be included in the RFM.
b. Procedures.
Page B – 104.3
AC 29-2C, Chg 4 5/1/2014
(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 and CG variation as fuel is burned.
(2) The critical condition for VNE controllability testing is ordinarily aft CG, MCP, 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 MCP 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 V NE controllability test is normally accomplished shortly after the
1.1 VNE (or 1.1 VH) point obtained during stability tests required by § 29.175(b).
Controllability must be satisfactory for both conditions. If V NE varies with altitude or temperature, VNE for existing ambient conditions is utilized for the test. Extremes of the altitude or 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. Testing will be carried out at the power required to achieve stabilized flight conditions. With rotorcraft that are operating in conditions such that the gross weight is limited by the power available, there should always be adequate tail rotor pedal authority to maintain yaw control when using the maximum approved all engines operating (AEO) power, which is takeoff power (TOP) for most designs.
(i) Where the rotorcraft is capable of operating at maximum gross weight with less than maximum approved power, it is appropriate to examine the rotorcraft characteristics with small amounts of additional power applied above the trim power required to allow for typical power variations experienced during normal use of the Page B – 104.4
rotorcraft. For example, maneuvering, turbulence, or rotor governing characteristics may cause the pilot to use power in excess of power required for trim. The maximum power excursions should not exceed maximum approved AEO power, excluding any transient range.
(ii) The rotorcraft should be flown both in ground effect (IGE) and out of ground effect (OGE), with the most adverse wind speed and direction for directional control within the flight envelope proposed, using power variations above trim that might be expected during normal use of the rotorcraft. Consideration should be given to the amount of excess power available, the ease with which power can be controlled via the collective, the effect of tail rotor control inputs on power required, and the characteristics of the rotorcraft if the limits of directional control are approached. There should be no tendency to deviate rapidly or suddenly in yaw. This assessment is normally conducted in conjunction with the critical azimuth testing.
(4) Prior to engine failure testing, it is mandatory that the pilot be fully aware of the 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 § 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.
(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 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 Page B – 105
AC 29-2C, Chg 4 9/17/2009 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 § 29.75 (see section 29.75 of this AC).
AC 29.143A. § 29.143 (Amendment 29- 51) Controllability and Maneuverability.
a. Explanation. Amendment 29- 51 made a minor clarification to assure that inground-effect (IGE) controllability is demonstrated at all wind speeds up to 17 knots, for all azimuths. In many rotorcraft, the entry into the regime of translational lift requires the most power, thus potentially causing control difficulties, and frequently occurs at speeds less than 17 knots. The amendment also requires that out -of-ground-effect (OGE) controllability be determined up to a speed of at least 17 knots at a weight selected by the applicant. The amendment clarifies the intent of Amendment 29- 21 and Amendment 29-24 with respect to removing hover controllability as a limit.
Section 29.25 is amended to assure that appropriate weight limitations are incorporated into the Rotorcraft Flight Manual (RFM) when the relieving provisions of the previous amendments are adopted by an applicant. The previous amendment and associated AC material indicated that certain Category B rotorcraft were relieved from providing, as a limitation, the conditions of § 29.143(c). In practice, the 17- knot controllability requirement was still treated as a limitation, but, as indicated in the amended § 29.25, additional limits could be included, when demonstrated, that allowed for something other than 17-knot all azimuth controllability. The established weight, altitude, and temperature charts, including any associated wind constraints, could be contained in the performance section of the flight manual when the appropriate reference to those charts were included in the limitations section of the RFM. In addition, the relief of Amendments 29-21 and 29-24 were only intended for those category B rotorcraft with nine or less passenger seats. All the policy material pertaining to this section remains in effect with the following changes:
(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 controllability and for maneuverability are summarized in § 29.143(a) which is 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 rotorcraft, the hover condition clearly applies under "any maneuver appropriate to the type."
(2) Paragraphs (b) through (e) in § 29.143 include more specific flight conditions and highlight the typical areas of concern during a flight test program.
(i) Section 29.143(b) specifies flight at VNE with critical weight, center of gravity (CG), rotor RPM, and power. Adequate cyclic authority must remain at V NE for Page B – 106
nose down pitching of the rotorcraft and for adequate roll control. Nose down pitching 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 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 nose down 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.
(ii) Section 29.143(c) requires a minimum control capability for hover and takeoff in winds from zero to at least 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 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. Prior to Amendment 29-24, hover controllability, height-velocity, and hover performance were the three regulatory requirements that ordinarily determined the shape of the limiting weight-altitude-temperature (WAT) curve for takeoff and landing. For Category A performance rotorcraft operations, of course, the one-engine-inoperative (OEI) climb performance requirements may also influence the WAT limit curve. Amendment 29- 24 allows, under certain conditions, the deletion of any hover controllability condition determined under § 29.143(c) from becoming an operating limitation. Section 29.1587 of Amendment 29-24 provides a means wherein Category B certificated rotorcraft (in accordance with the requirements of § 29.1, effective with Amendment 29-21) may not be limited by the hover controllability requirements of § 29.143(c). Section 29.1583(g) requirements for Category A certificated rotorcraft are unchanged from past regulatory requirements in that if the hover controllability requirements of § 29.143(c) result in the most restrictive envelope it will be published as an operating limitation. Section
29.1587(b) provides a means wherein Category B certificated rotorcraft, as defined in § 29.1, may not be restricted in its utilization. Section 29.1587(b) allows some Category B RFMs to include maximum takeoff and landing performance information, provided that something other than the 17-knot hover controllability requirement is not limiting. This may be zero wind IGE hover performance or any other performance the applicant elects to use, if the maximum safe wind for operations near the ground is provided. Rotorcraft certificated prior to Amendment 29-24 can update their certification basis to take advantage of this provision. If an applicant with a previously type certificated rotorcraft elects to update to this later amendment, caution should be taken to verify that the Page B – 106.1
AC 29-2C, Chg 4 9/17/2009 height-velocity information is done in accordance with Amendment 29-21; that all engine out landing capabilities are satisfactorily accounted for at the new proposed gross weight, altitude, temperature combinations; that takeoff/landing information is provided;
and that sufficient information is provided to properly advise the crew of the rotorcraft's capabilities when utilizing this increased performance capabilities.
(iii) Section 29.143(e) 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 that 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, 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 (MCP) (or takeoff power where appropriate) including reasonable engine torque splits. For multiengine Category A installations with three or more engines, the 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 MCP 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 Category B powerplant installation rotorcraft, demonstration of sudden complete power failure is required at critical conditions throughout the flight envelope. This includes speeds from zero to V NE (power-on) and conditions of hover, takeoff, and climb at VY. MCP 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 part 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 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 "cruise" also includes cruise climb and cruise descent conditions. Normal pilot reaction times are used elsewhere. Although this requirement specifies MCP, it does not limit engine failure testing to MCP. If a 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, the rotor speed, flapping, and aircraft dynamic characteristics must stay within structurally approved limits.
Page B – 106.2
(iv) Section 29.143(f) 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 VNE (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 § 29.1505(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 V NE at MC power per § 29.143(b) and § 29.175(c), and to 1.1 VNE (power-off) in autorotation per § 29.143(f) 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 RFM.
(3) The required controllability and maneuvering capabilities must also be considered following the failure of automatic equipment used in the control system (§ 29.672). 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 Page B – 106.3
AC 29-2C, Chg 4 9/17/2009 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. A n 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 surfaces, 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 or softover malfunction should be included in the RFM.
b. Procedures. The policy material pertaining to this section remains in effect with the following changes and additions:
(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 MCP 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 V NE controllability test is normally accomplished shortly after the 1.1 VNE (or 1.1 VH ) point obtained during stability tests required by § 29.175(b).
Controllability must be satisfactory for both conditions. If V NE 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) Controllability
(i) 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 Page B – 106.4
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. IGE 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. OGE testing should be done with the rotor at a predetermined height above the ground at which it has been determined that there is no ground effect. 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.
(A) Testing will normally be carried out at the power required to achieve stabilized flight conditions. However, it is also important to show that yaw control remains adequate to allow normal power changes that might be required in normal operational maneuvers typical for the type and use of the rotorcraft. With rotorcraft that are operating in conditions in which the gross weight is limited by the power available, there should always be adequate tail rotor pedal control available to maintain yaw control when using up to Take-off Power. However, this will not be the case if the rotorcraft weight in the low speed flight envelope is limited by yaw control system capability. There may be other conditions where adequate yaw control is not available at high power, for example a rotorcraft which is limited by the CAT A weight (for rotorcraft certificated to § 29.1 (c)).
(B) To cover the case where excess power is available, it is appropriate to examine the rotorcraft characteristics with some small amounts of additional power applied. This will account for typical power variations that will be experienced during normal use of the rotorcraft. For example, maneuvering or turbulence will cause the pilot to use some of the excess power available. The rotorcraft should be flown, both IGE and OGE, with the most adverse wind speed and direction for directional control within the flight envelope proposed. Use power variations above trim that might be expected during normal use of the rotorcraft giving consideration to the amount of excess power available, the ease with which power can be controlled by collective, and the characteristics of the rotorcraft if the limits of directional control are approached.
There should be no tendency to deviate rapidly or suddenly in yaw. This assessment is normally conducted in conjunction with the critical azimuth testing.
(C) It may be appropriate to provide flight manual information on the directional control characteristics, including any relevant maximum power above which it could be expected that directional control might not be maintained.
Page B – 106.5
AC 29-2C, Chg 4 9/17/2009
(ii) Comprehensive controllability tests are typically conducted at low, intermediate (~7000 feet Hd), and high tests sites, with prepared landing surfaces, in conjunction with takeoff, landing, and performance testing.
(iii) Alternatively, a predicted controllability model developed for high altitude may be used if verified by limited flight testing with steady ambient winds. The extrapolation guidelines in AC 29.45 b(2) are still applicable. These high altitude controllability tests could typically be conducted in conjunction with takeoff, landing and performance tests.
(iv) Controllability can usually be extrapolated up to a maximum of 2,000 feet above the highest test site altitude.
Note: Engine operating characteristics must be considered during the limited altitude tests.
(4) Prior to engine failure testing, the pilot should 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 should 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 should 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 exceeding a temperature limitation. 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.
(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 cyclic to correct roll in conjunction with forward cyclic to correct pitch-up is satisfactory. Obviously Page B – 106.6