a demonstration of flare effectiveness with power recovery at the maximum landing gross weight corresponding to the altitude satisfies the requirement for touchdown. The flare must reduce the autorotative descent rate and forward speed to a reasonable value (i.e., autorotative descent rate and forward speed consistent with that demonstrated at sea level). If a complete company test program has documented the all-engine-out landing capability, verification tests may be initiated at those limiting weight conditions. If not, buildup testing should be initiated at light weight.
(ii) Demonstrated compliance with this requirement is intended to show that the autorotative descent rate and forward speed at touchdown (less than 40 knots true airspeed (KTAS) is recommended) can be controlled to a reasonable value to ensure a reasonable chance of survivability for the all engine failure condition. On multiengine rotorcraft, rotor inertia is typically lower than for single-engine rotorcraft. RPM decays rapidly when the last engine is made inoperative. Due to this relatively low inertia level, considerable application of collective pitch may be needed to prevent rotor overspeed conditions when the rotorcraft is flared for landing. Also, when testing the final maximum weight points, the pilot should anticipate a need for considerable application of collective pitch to control rotor overspeed during autorotative descent, particularly at high altitude. Some designs incorporate features that may lead to rotorcraft damage in testing this requirement (e.g., droop stop breakage or loss of directional control with skids) if landings are conducted to a full stop with the engines off.
(iii) The intent of this rule is to demonstrate controlled touchdown conditions and freedom from loss of control or apparent hazard to occupants when landing with all engines failed. In these cases, compliance can be demonstrated by leaving throttles in the idle position and ensuring no power is delivered to the drive train. Also, computer analysis may be used in conjunction with simulated in-flight checks to give reasonable assurance that an actual safe touchdown can be accomplished. Another method may be to make a power recovery after flare effectiveness of the rotorcraft has been determined, showing that the rate of descent and forward speed can be controlled to allow for safe landing. Other methods may be considered if they lead to reasonable assurance that rate of descent and forward speed can be controlled to allow safe landing with no injury to occupants when landing on a prepared surface with all engines failed. Regardless of the method(s) used to comply with this requirement, careful planning and analyses are very important due to the potentially hazardous aspects of power off simulation and landing of a multiengine rotorcraft totally without power. The OEI landing test is ordinarily done in conjunction with height velocity tests because ground and onboard instrumentation requirements are the same for both tests.
(6) Prior to conducting these tests, the crew should be familiar with the engine inoperative landing characteristics of the rotorcraft. The flight profile may be entered in the same manner as a straight-in practice autorotation. It is recommended that for safety reasons, idle power be used if a "needle split" (no engine power to the rotor) can be achieved. In some cases, a low engine idle adjustment has been set to assure needle split is attained. In other cases, a temporary detent between idle and cutoff was used on the throttle. In a third case, the engine was actually shut down on sample runs Page B - 36.1
AC 27-1B, Chg 4 9/17/2009 to verify that the engine power being delivered was not materially influencing landing capability or landing distances. The flare is maintained as long as it is reasonable to dissipate speed and build RPM. Rotor RPM must stay within allowable limits. Aft CG is ordinarily critical due to visibility and pitch attitude effectiveness in a flare. Following the flare, the rotorcraft is allowed to touch down in a landing attitude. Rotor RPM at touchdown should be recorded, and it must be within allowable structural limits.
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AC 27.79. § 27.79 LIMITING HEIGHT-SPEED ENVELOPE.
All of the policy material pertaining to this AC section remains in effect through Amendment 27-43. See § 27.87 and this AC’s section 27.87 effective with Amendment
27-44.
a. Explanation.
(1) The height-speed envelope is normally referred to as the height-velocity (HV) diagram. It defines an envelope of airspeed and height above the ground from which a safe power-off or one engine inoperative (OEI) landing cannot be made. The diagram normally consists of three portions: (a) the level flight (cruise) portion, (b) the takeoff portion, and (c) the high speed portion (see Figure AC 27.79-1). The high speed portion is omitted on occasions when it can be shown that the rotorcraft can suffer an engine failure at low altitude and high speed (up to VH) and make a successful landing or climb out on the remaining engine(s).
(2) Power failure, engine failure, throttle chop, or other similar terms used in this discussion mean a simulated engine failure. The actual shutdown of an engine to simulate an engine failure should not be necessary if the simulated procedure ensures that the engine power is suddenly removed from driving the rotor and remains so. The normal fuel control deceleration schedule is usually satisfactory for the power removal for turbine engines but the flight or ground idle speed may have to be set lower than normal for HV testing.
(3) The avoid areas of the HV diagram are separated by the takeoff corridor. This corridor should be wide enough to consistently permit a takeoff flight path clear of the HV diagram using normal pilot skill. The takeoff corridor should always permit a minimum of ±5 knots clearance from critical portions of the diagram.
(4) The knee of the curve separates the takeoff portion from the cruise portion and is defined as the highest speed point on the low speed portion of the HV envelope.
Altitudes above this point are considered cruise, or “fly-in,” points, and these test points require a minimum time delay of 1 second between throttle chop and control actuation (reference § 27.143(d)). Altitudes below the knee represent takeoff profile points. For test points in the takeoff portion, use takeoff power (or a lower power selected by the applicant as an operating procedure) and normal pilot reaction time for corrective control actuation.
(5) Since the HV diagram may represent the limiting capabilities of the rotorcraft, each test point should be approached with caution. The manufacturer’s buildup program should be reviewed to determine the amount of conservatism in the HV diagram (if any). It should be remembered that the operational pilot will be operating at or near the HV diagram without the benefit of a buildup program. Buildup testing is necessary, and it is most important to vary only one parameter at a time to prevent surprises. Light weight testing is ordinarily conducted first. High and low hover points Page B - 37
AC 27-1B, Chg 4 3/31/2014 are approached from above and below respectively. Portions near the knee are initially evaluated at high speed with subsequent backing down of the speed. In most rotorcraft the effective flare airspeed is critical. At airspeeds slightly below this value, the ability to arrest and control descent rates through use of an aft cyclic flare may be greatly diminished. Extreme care should be exercised when “backing down” to lower speeds.
(6) In addition to the on-board and ground instrumentation, a motion picture camera or other position measuring equipment should cover each run.
(7) For FAA/AUTHORITY tests, the minimum required crew and the minimum instrument panel display presented for certification should be used. Ground safety equipment should be provided.
(8) This test is the least predictable of all the performance items. Therefore, the expansion and extrapolation of test data are questionable. Weight may not be extrapolated to higher values. In order to extrapolate HV data to higher altitudes, any analytical method must have FAA/AUTHORITY approval. In lieu of pure analytical methods, simulations have been used successfully, especially for multiengine rotorcraft.
In either case, the maximum allowable extrapolation should be limited to 2,000 feet density altitude (HD). HV test weights for normal category rotorcraft are the maximum weight at sea level and some lesser weight at high density altitudes. The high density altitude HV curve needs to be defined only to 7,000 feet and may be a lower altitude if the rotorcraft does not have the performance capabilities to attain 7,000 feet. A weight less than the maximum weight may be used to define the high density altitude HV curve, but this weight should not be less than the maximum weight that will allow hovering out of ground effect (OGE). For a given diagram, typical weight reductions that are necessary as altitude is increased can be conservatively estimated by maintaining a constant gross weight divided by density ratio, GW/σ (see Figure AC 27.79-2, Part A). If weight is not varied, an enlarged HV diagram is required for safe power-off landing as density altitude is increased (see Figure AC 27.79-2, Part B).
Another method of presentation is to show varying weights at a constant density altitude (see Figure AC 27.79-2, Part C.)
(9) The FAA accepts, as a method of extrapolation, a weight penalty of 3% for each 1000 feet above the permitted 2000 feet extrapolation. This weight penalty has been applied to extrapolations along a constant gross weight divided by density ratio, GW/σ.
(10) Vertical takeoff and landing (VTOL) testing normally does not require separate HV testing. The takeoff and landing tests take on the combined characteristics of takeoff, landing, and HV tests.
b. Procedures.
(1) Instrumentation.
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(i) Ground Station. The ground station must have equipment and instrumentation to determine wind direction and velocity, outside air temperature, and if the test rotorcraft has reciprocating engines, humidity. Since the tests must be conducted in winds of 2 knots or less, a smoke generator is highly recommended to show both flightcrew and ground crew personnel the wind direction and velocity at any given time. Additionally, the location of the ground station should be such that it is free of rotor downwash at all times. Motion picture or phototheodolite and radio equipment will be necessary to properly conduct the test program. The use of telemetry equipment is desirable if the location of the test site and the magnitude of the test program make it practical.
(ii) Airborne Equipment (Test Rotorcraft). Necessary installed test equipment may include photopanels or recorders for recording engine parameters, control positions, landing gear loads, landing gear deflections, airspeed, altitude, and other variables. An external light attached to the rotorcraft (or any other means of identifying the engine failure point to the ground camera or phototheodolite) is needed to identify the exact time of engine failure and may also be used to synchronize the ground recorder with the airborne recorded data.
(2) Analytical Prediction. The HV diagram can be estimated by analytical means and this is recommended prior to test. HV, however, is the least predictable of all rotorcraft performance and because of this, the expansion and extrapolation of test data must be done with great care. Test weight may not be extrapolated. All test points should be approached conservatively with some speed or altitude margin. If the applicant has conducted a comprehensive HV flight test program to validate his analytical predictions, much preliminary testing can be eliminated. In any case, the maximum allowable extrapolation from flight test conditions is 2,000 feet density altitude, and an approved analytical or simulation method must be utilized for extrapolation.
(3) Power.
(i) The appropriate power level before engine failure for the low and high hover points is simply the power required to hover at the prevailing hover conditions. The appropriate power condition prior to failure of the engine for points below the knee is takeoff power or a lower value if approved as an operating procedure. For cruise or “flyin” points above the knee, the appropriate condition is power required for level flight.
(ii) The applicable power failure conditions are listed in § 29.79(b). Power should be completely cut for normal category rotorcraft. For multiengine rotorcraft that comply with the category A engine isolation design requirements, the HV envelope may be determined with OEI and the desired topping power (for the remaining engine(s)) should be set prior to the test. This power value will need adjustment as ambient conditions change. The power can be takeoff power (TOP), 2 ½-minute power, or some calculated lower power for simulating hot day or higher density altitude conditions.
Power is verified and recorded by the pilot by “topping” the engine(s) prior to engine Page B - 39
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.
(iii) 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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FIGURE AC 27.79-1 HEIGHT-VELOCITY (HV) DIAGRAM
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AC 27-1B, Chg 4 3/31/2014
• Approx.
Height Ho Constant GW/% Velocity LA Weight
CONSTANT HV 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 27-1B, Chg 9 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.
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