AC 27-1B, Chg 4 3/31/2014 failure tests. Care must be taken to ensure that this power value is no more than that which would be delivered by a minimum specification engine under the ambient conditions to be approved.
(4) Test Loadings. Weight extrapolation is not permitted for HV. Therefore, the test weight must be closely controlled. Ballast or fuel should be added frequently to maintain the weight within -1 to +5 percent when testing final points. Ordinarily, tests are conducted at a mid-center of gravity unless a particular loading is expected to be particularly critical.
(5) Landing Gear Loads.
(i) Instrumented landing gear can be a great help in evaluating test results. This information can be telemetered to a ground station or otherwise recorded and displayed for direct reference following each landing.
(ii) Any landing which results in permanent deformation of aircraft structure or landing gear beyond allowable maintenance limits is considered an unsatisfactory test point.
(6) Piloting Considerations. In verifying the HV diagram, the minimum certificated instrument panel display and minimum crew should be used in order not to mislead the operational pilot who has no test equipment available and may have no copilot to assist.
Three distinctly different flight profiles are utilized in developing the diagram.
(i) High Hover. A stabilized OGE hover condition prior to power failure is essential. A minimum 1-second time delay between power failure and initial control actuation is utilized. Following the time delay, the primary concern is to quickly lower collective and to gain sufficient airspeed to allow an effective flare approaching touchdown. While the immediate development of airspeed is necessary, the dive angle must be reasonable and must be representative of that expected in service. While initial aircraft attitude will vary between models and with changing conditions, 10°-20° has been previously applied as a maximum allowable nose down pitch attitude. Use of greater attitudes could result in a diagram which is difficult to achieve and unrealistic for operations in service. Initial testing should start relatively high with gradual lowering of height to the final high hover altitude. A stabilized OGE hover condition prior to power failure is essential. If a stabilized high hover condition cannot be achieved prior to the engine cut, then this point should be tested from a minimum level flight speed. This will result in an open-ended HV diagram. A smoke source or balloon on a long cord is highly desirable since the wind can vary significantly from surface observations to typical high hover altitudes. Vertical speed must be very near zero at the throttle chop.
Any climb or sink rate can have a significant influence on the success of the test point.
Use of a radar altimeter with a cross check to barometric altitude is essential.
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(ii) Low Hover. From the low hover position there is no flare capability and little time for collective reaction. No time delay is applied other than normal pilot reaction.
For typical designs the collective may not be lowered after power failure. Lowering of the collective is not permitted because it is not a pilot action which could be expected if an engine failed without notice during a hovering condition in service. Initial lowering of collective immediately after power failure can result in a very high, unconservative low hover height that is unrealistic for operational conditions. If, however, a design is such that a 1-second pilot delay after power failure could be achieved without any appreciable descent, a slight lowering of collective could be allowed.
(ili) Takeoff Corridor. Normal pilot reaction is applied when the engine is made inoperative. At low speeds, collective may be lowered quickly to retain RPM and minimize the time between power failure and ground contact. If airspeed is sufficient for an effective flare, the aircraft is flared to reduce airspeed, retain rotor RPM, and control vertical speed prior to touchdown. Considerable surface area may be needed for a sliding or rolling stop.
(iv) Additional Considerations. The "in-between" points utilize similar techniques. The cruise or "fly-in" points are similar to the high hover point although the steep initial pitch attitudes are not needed as altitude is decreased and airspeed is increased along the curve. The low speed points along the takeoff corridor are similar to the low hover point except that the collective may be quickly lowered and some flare capability may be used as the "knee" is approached. The pilot should be proficient in all normal autorotation landings before conducting HV tests in a single-engine rotorcraft.
(7) Ground Support. Motion picture or theodolite coverage and ground safety equipment are necessary. Communication capability among these elements should be provided. Use of a phototheodolite to compare height and speed with cockpit observations is very desirable.
(8) Verifying the HV Diagram.
(I) A sufficient number of test points must be flown to verify the diagram. The key areas are the knee, high altitude hover, low altitude hover, and low altitude high speed flight. Test points with excessive gear loads, exceptional skill requirements, winds above permissible levels, rotor droop below approved minimum transient RPM, damage to the rotorcraft, excessive power, incorrect time delay, etc., cannot be accepted.
(ii) After the HV diagram is defined, it should be ascertained that the corridor permits takeoffs within +5 knots of the recommended takeoff profile.
(9) Flight Manual. The flight manual should list any procedures which may apply to specific points (e.g., high speed points) and test conditions, such as runway surface, wave height for amphibious tests, marginal areas of controllability or landing gear Page B - 41
AC 27-1B, Chg 4 3/31/2014 response, etc. The HV curve should be presented in the RFM using actual altitude above ground level and indicated airspeed.
(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 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.
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HEIGHT LEVEL
XHJ FLIGHT -KNEE
- TAKEOFF
TAKEOFF CORRIDOR
HIGH SPEED
SPEED (V) FIGURE AC 27.79-1 HEIGHT-VELOCITY (HV) DIAGRAM Page B - 43
AC 27-1B, Chg 4 3/31/2014
• Approx.
Height Constant GW Velocity Weight
CONSTANT HY DIAGRAM, VARIABLE WEIGHT Increasing Density Altitude Height Velocity — B
CONSTANT WEIGHT Increasing Gross Weight Height Velocity
CONSTANT DENSITY ALTITUDE FIGURE AC 27.79-2 ALTITUDEWEIGHT ACCOUNTABILITYT Page B - 44
AC 27.79A.§ 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.
AC 27.87. § 27.87 (Former § 27.79) (Amendment 27-44) Height-Speed Envelope.
(For § 27.79 prior to Amendment 27-44, see AC 27.79) AC 27.87. § 27.87 (Amendment 27-44) HEIGHT-SPEED ENVELOPE.
a. Explanation. Amendment 27-44, in addition to some minor text changes, clarifies the one-engine-inoperative (OEI) engine power to be used (for multi-engine rotorcraft) when demonstrating the requirement. The engine power of the remaining engine is the minimum uninstalled specification power after it is corrected for installation losses. The methods of determining this power are established in the general performance paragraph of this AC.
b. Procedures. The policy material pertaining to the procedures outlined in this section remain in effect.
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AC 27.87A. § 27.87 (AMENDMENT 27-51) HEIGHT-VELOCITY ENVELOPE.
a. Explanation. Amendment 27-51 changed the title of the section from "Heightspeed envelope" to "Height-velocity envelope" to be consistent with the title nomenclature of § 29.87, which is the equivalent requirement for transport category rotorcraft. The title of § 29.87 was changed from "Limiting height-speed envelope" to "Height-velocity envelope" at amendment 29-39.
b. Procedures. The policy pertaining to the procedures in this section remains unchanged. For purposes of this AC, the terms "height-velocity" and "height-speed" are equivalent.
Page B - 45.1
SUBPART B - FLIGHT
FLIGHT CHARACTERISTICS AC 27.141. § 27.141 (Amendment 27-21) FLIGHT CHARACTERISTICS-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 must 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 Appendix B of this AC.
(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 (RFM) 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 Page B - 46
(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 assessment and to carefully document the results of these evaluations.
Section 27.695 requires an alternate system allowing continued safe flight and landing following any single failure of a control system hydraulic boost system. This assessment of 'safe' should take into account not only residual post-failure control loads but also workload and pilot fatigue considerations. The following is suggested as an appropriate test sequence, conducted by a range of pilots, for a VFR approval:
(i) Simulated hydraulic failure at critical flight conditions and Max GW.
(ii) Establish level flight at a cruise speed >Vy.
(iii) Fly for approximately 30 minutes (to assess workload and account for pilot strength variations), demonstrating ability to climb or descend and small bank angle turns. This also allows for the possibility that the hydraulic failure occurs over water or other undesirable landing area.
(iv) Land at a suitable area using a recommended landing technique, appropriate to the emergency.
(5) Because control loads typically increase at higher altitudes, failure of the control hydraulic boost system should also be investigated during high altitude testing, including landing at 7000' HD, at Max GW/o. Where approval for any other type of operation is requested (e.g., IFR, category A), an appropriate test sequence must be proposed to the FAA/Authority. Section 27.141(b) provides the regulatory requirements for these strength and skill requirements, as well as a smooth transition capability between appropriate flight conditions. These requirements must also be met during appropriate engine failure conditions for each category of rotorcraft. Flight characteristics and pilot workload should be evaluated in all expected flight conditions, including actual turbulence.
(6) For night or IFR approval, § 27.141(c) requires additional characteristics for night and IF flight. The appropriate flight test procedures are included in other portions of this guidance.
b. Procedures. None.
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AC 27-1B, Chg 4 3/31/2014 AC 27.143. § 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. During the assessment carried out under § 27.143(a) (2)(v) for rotorcraft in glide (i.e., autorotation), in addition to controllability and maneuverability, it should be shown by flight testing that the directional stability characteristics are sufficient to allow the pilot to control the rotorcraft without undue attention to heading at the speeds for minimum rate of descent and best angle of glide. This should be evaluated at normal trim conditions and in turns up to 30° angle of bank. The ability to generate a sideslip must be evaluated throughout the autorotation speed envelope. 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." The rotorcraft must still meet the stability requirements of Subpart B and, if applicable, Appendix B.
(2) Paragraphs (b) through (e), § 27.143, include more specific 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. 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 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 whether 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.
(i) 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 Page B - 48
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.
(iii) 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 that 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 (MCP) (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 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 § 27.75(b).
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 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. 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 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 tor 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. Following power failure, Page B - 49
AC 27-1B, Chg 4 3/31/2014 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 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 § 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 MCP 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 (RFM).
(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), Page B - 50