(7) For FAA/AUTHORITY tests, the minimum required crew and minimum instrument panel display 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 should be c onsistent with the takeoff and landing weight, altitude, temperature (WAT) limit curve which will be placed in the rotorcraft flight manual (RFM). 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/ (see Figure AC 29.79-2, 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 29.79-2, Part B). Another me thod of presentation is to show varying weights at a constant density altitude (see Figure AC 29.79-2, 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/.
(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.
(11) Rotorcraft certificated prior to Amendment 29-21 were required to have the resulting HV diagram as an operating limitation. This limitation restricted opportunities when operating large rotorcraft in various utility applications. Subsequently, Amendment 29-21 allows, under certain conditions, the HV diagram to be placed in the Flight Manual Performance Information Section instead of the Limitations Section.
Specifically, the rotorcraft must be: (1) certificated for a maximum gross weight of 20,000 pounds or less; (2) configured with nine passenger seats or less; and
(3) certificated in Category B. Testing must be completed with the aircraft at the maximum gross weight at sea level. For altitudes above sea level, the test aircraft must be at a weight no less than the highest weight the rotorcraft can hover out of ground effect (OGE). Rotorcraft certificat ed prior to Amendment 29-21 can update their certification basis to take advantage of this provision.
b. Procedures.
(1) Instrumentation.
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AC 29-2C, Chg 4 5/1/2014
(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, 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 photo panels 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 di agram 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 manufacturer 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 limit. For cruise or “fly-in” points above the knee, the appropriate condition is power required for level flight. Rotor speed at execution of the engine failure should be the minimum speed appropriate to the flight condition.
(ii) The applicable power failure conditions are listed in § 29.79(b). Power should be completely cut for category B 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 adj ustment as ambient conditions change. The power can be takeo ff power (TOP), 2 1/2-minute power, or Page B - 94
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 failure tests. Care must be taken to assure 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 extrapolatio n is not permitted for HV. Therefore, the test weight must be closely controlled. Ba llast 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 required 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 del ay, 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 hov er 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 mu st 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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AC 29-2C, Chg 4 5/1/2014
(ii) Low Hover. From the low hover position there is no flare capability and little time for collective reaction. No ti me 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 very high, unconservative low hover altitudes that are 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 pilo t 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. C onsiderable 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 poin ts 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 high speed touchdown. Test points with excessive gear loads, above average 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 s hould 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 - 96
response, etc. The HV curve should be pr esented 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. Engine failu res 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. See related discussion under AC 29-2C, section 29.63.
(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 (H) ✓-KNEE ~-TAKEOFF
CORRIDOR
HIGH SPEED
SPEED (V) AC 29-2C, Chg 4 5/1/2014 FIGURE AC 29.79-1 HEIGHT- VELOCITY (HV) DIAGRAM Page B - 98
Approx Constant GWe Height Velocity Weight — A
CONSTANT HV DIAGRAM, VARIABLE WEIGHT Increasing Density Altitude Height Velocity
CONSTANT WEIGHT Increasing Gross Weight Height Velocity
CONSTANT DENSITY ALTITUDE FIGURE AC 29.79-2 ALTITUDENEIGHT CCOUNTABILITY Page B - 99
AC 29.85. § 29.85 (Amendment 29-39) BALKED LANDING: CATEGORY A (For Balked Landing prior to Amendment 39, see § 29.77 and paragraph AC 29.77.)
a. Explanation. Amendment 29-39 revised and relocated the original § 29.77 as a new § 29.85. The guidance material of paragraph AC 29.77 does not apply to rotorcraft certified with Amendment 29-39 or later. This rule has two distinct portions.
(1) Section 29.85(a) states that the rotorcraft must be capable of transitioning smoothly from each approved Category A approach condition to a missed approach with one engine inoperative (OEI). Although not specifically stated in the rule, this requirement must be met for any point prior to the landing decision point (LDP).
(2) Section 29.85(b) requires that the LDP be defined so that it will permit transition to a safe climb condition in the event a balked landing is necessary. (See figure AC 29.75A-1.) The safe climb conditi ons are defined in § 29.67(a)(1) and (2). A clearly defined balked landing profile similar to the Category A takeoff profile should be established. The balked landing profile must insure compliance with the climb performance requirements of §§ 29.67(a)(1) and 29.67(a)(2).
b. Procedures.
(1) Instrumentation. Instrumentatio n requirements are similar to those for Category A takeoff. A ground station with positioning capability is needed along with on-board instrumentation of engine and flight parameters.
(2) Balked Landing Profiles. One engine inoperative balked landing profiles during approach must be conducted at conditions down to and including the LDP. The LDP should be designated so that the balked landing profile may be completed with the rotorcraft clearing the landing surface by a minimum of 15 feet. Fifteen feet should be considered the absolute minimum clearance allowed with greater clearances required for some rotorcraft dependent on rotorcraft geometry and performance characteristics.
For elevated or ground level heliports, with significantly lower LDP heights than 100 feet, the minimum clearance is 15 feet vertically and radially. These minimum heights would need to be demonstrated with variations in piloting techniques and with pilot recognition and reaction times for engine failures occurring before/after LDP. The distance from the LDP to the point in the balked landing profile at which a minimum of 35 feet above the landing surface is attained at VTOSS in a climbing posture should be recorded. This distance should be com pared against the landing distance determined under § 29.81 to assure the balked landing maneuver can be completed within the designated landing area. This is especially important for future steep angle, low speed Category A approaches to heliports.
(3) Handling Qualities. Handling qu alities features in the balked landing transition should be carefully evaluated. Characteristics such as excessive nose down pitching with power application or excessive engine lag should not be approved.
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(4) Climb Performance. In accordance with this rule, the climb requirements of § 29.67(a)(1) and (2) must also be met in the event a balked landing is made. See paragraphs AC 29.65 and AC 29.67.
AC 29.87. § 29.87 (Amendment 29-39) LIMITING HEIGHT-SPEED ENVELOPE.
(For Limiting Height-Speed Envelope prior to Amendment 39, see § 29.79 and paragraph AC 29.79.)
a. Explanation. Amendment 39 redesignated § 29.79 as § 29.87.
b. Procedures. The guidance material presented in paragraph AC 29.79 continues to apply.
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AC 29-2C, Chg 4 5/1/2014
SUBPART B - FLIGHT
FLIGHT CHARACTERISTICS AC 29.141. § 29.141 (Amendment 29-24) FLIGHT CHARACTERISTICS GENERAL.
a. Explanation.
(1) This regulation prescribes the general flight characteristics required for certification of a transport 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 § 29.141(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 V Y, 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 §§ 29.141 through 29.161 are met at takeoff power levels, no flight manual entry is needed. Obviously, if transmission limits for maximum continuous (MC) and takeoff power are coincident, 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 § 29.143(c) would be in question. The same consideration must be given to all rotorcraft controls and moveable aerodynamic surfaces where questionable compliance with the regulations may exist. If the rotor -induced vibration characteristics of the rotorcraft are significantly affected and require time-consuming rigging for such things as acceptable ride comfort, then the rotor(s) should be rigged to the allowable extreme tolerance limits to determine compliance, for example, with § 29.251.
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(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 29.695 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 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, it should be considered if this failure mode should also be investigated during high altitude testing at Max GW/σ. 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 29.141(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, § 29.141(c) requires additional characteristics for night and IFR flight. The appropriate flight test procedures are included in other portions of this guidance.
b. Procedures. None.
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AC 29-2C, Chg 4 5/1/2014 AC 29.143. § 29.143 (Amendment 29-24) CONTROLLABILITY AND MANEUVERABILITY.
a. Explanation.
(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 control and for maneuverability are summarized in § 29.143(a) which is largely self-explanatory. During the assessment carried out under § 29.143(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 § 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.” The rotorcraft must still meet the stability requirements of Subpart B and, if applicable, Appendix B.
(2) Paragraphs (b) through (e), § 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 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 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.
(ii) Section 29.143(c) requires a minimum 17-knot control capability for hover and takeoff in winds 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 rotorcraft incorporating a tail rotor, Page B - 104