(A) Unless the rotorcraft is capable of hovering with one engine inoperative at the desired Category A weight, the LDP becomes largely a function of the runway length required for landing. If landing conditions to be scheduled include considerable runway length (on the order of 1,000 feet), the LDP may be defined at a relatively high speed allowing transition to a takeoff safety speed near VY which will allow the maximum amount of weight for compliance with the balked landing climb requirements of § 29.85(b)/§ 29.67(a)(1). In this case, the requirements of § 29.67(a)(2) usually become limiting. If the runway length is small, LDP will typically be at a lower speed and may be at a higher altitude to allow balked landing transition within the available distance. Landing weight may need to be reduced to allow landing from the lower speed or higher altitude decision point for shorter landing distances. In this case the requirements of § 29.67(a)(1) may be limiting. The climb performance and climb speeds required by § 29.67(a)(1) and (2) should be established prior to Category A landing tests.
(B) The one-engine-inoperative landing is similar in many respects to the height-velocity tests described in paragraph AC 29.79. Most of the comments, cautio ns, and techniques for HV also apply here even though typical flight conditions at LDP are less critical than limiting HV points due to a lower power level and an established rate of descent. The approach is made at a predetermined speed and one engine is made inoperative prior to LDP. After the LDP, speed is reduced and the rotorcraft is flared to a conventional one engine inoperative landing. Depending on the landing characteristics and landing profile, the flare may be initiated either prior or subsequent to the 50 foot elevation utilized in determining landing distance. Testing should include an engine failure such that recognition is at the LDP with a 1-second pilot delay to ensure safe landing capability for this critical case. A sufficient number of acceptable runs should be accomplished to provide confidence in the results. Typically ten acceptable runs are adequate.
(C) The balked landing portion of the landing profile is addressed under § 29.85, Balked Landing: Category A. For an explanation of that requirement and a discussion of those test procedures, refer to paragraph AC 29.77.
(iii) Power. Power used for demonstrating performance should be limited to minimum specification values on the operating engine(s). This may be accomplished by adjustment of the engine topping to minimum specification values including consideration of temperature effects on engine power. If the management of engine power at topping is beyond normal pilot capability, the validity of the Cat A procedure must also be established with representative in-service characteristics. The method used for simulating engine failure must be representative of the power decay characteristics that will occur during a real, sudden engine failure and acceleration of the remaining engine(s). In order to cushion the OEI landing, it is acceptable for the engine and transmission transient range in order to droop the rotor provided performance credit is not taken for this additional power above the maximum permitted rating and it can be shown that the engine(s) will remain within these limits in all Page B - 83
AC 29-2C, Chg 2 4/25/06 conditions requested by the applicant. Any excursion beyond established transient limits in this flight phase should be substantiated to the extent that it does not constitute an immediate hazard to the rotorcraft.
(iv) Aircraft Loading. Aft center of gravity is usually most critical for landing distance determination because visibility constraints limit the degree to which the pilot can flare the rotorcraft for landing. If a weight effect is shown, a minimum of two weights should be flown at each test altitude. One weight should be the maximum weight for prevailing conditions and the other should provide a sufficient spread to validate weight accountability.
(v) Extrapolation. Landing data may be extrapolated along an established W/σ line to the maximum gross weight of the rotorcraft. However, extrapolation will not be considered valid if landing gear loads are marginally acceptable at actual landing weights below the W/σ limit. If no marginal areas are apparent and an acceptable analytical method is used, performance data may be extrapolated up to 4,000 feet density altitude from test conditions. (See paragraph AC 29.45.)
(vi) Ambient Conditions. Appropriate test limits for ambient conditions such as wind and temperature are contained in paragraph AC 29.45. Test data should be corrected for existing wind conditions during landing distance tests. Credit for headwind conditions may be given during flight manual data expansion.
Paragraph AC 29.1587 details allowable wind credit.
(vii) All Engine Out Landing.
(A) Several procedures can be utilized to demonstrate compliance with the all-engine-out landing requirement. As discussed in the explanation portion o f this paragraph, §§ 29.79 and 29.83 each require that a landing from autorotation be possible. The maneuver is entered by smoothly reducing power at an optimum autorotation airspeed at a safe height above the landing surface. All -engine-out landing tests should be initiated at light weight with a gradual buildup to the limiting weight conditions. If a complete company test program has documented all -engine-out landings to the GW/σ limit, the buildup conditions during verification test may be decreased. If not, buildup testing should be initiated at light weight. This test is ordinarily conducted at mid center of gravity. Typically, all altitudes may be approved with two weight limit landings - one at sea level and one near maximum takeoff and landing altitude.
(B) Demonstrated compliance with this requirement is intended to show that an autorotative descent rate can be arrested, and forward speed at touchdown can be controlled to a reasonable value (less than 40 KTAS is recommended) 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 collective may be needed to prevent rotor overspeed conditions when the Page B - 84
rotorcraft is flared for landing. Also, when testing the final maximum weight points, the pilot should anticipate a need for considerable collective pitch to control rotor overspeed during autorotative descent, particularly at high altitude WAT limiting conditions. Some designs incorporate features which 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 cut off.
(C) 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. Other methods may be considered if they lead to reasonable assurance that descent can be arrested and forward speed 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. Considerations for weight and altitude extrapolation are the same as those for HV testing (see paragraph AC 29.79). The all -engine-inoperative landing test is ordinarily done in conjunction with height velocity tests because ground and onboard instrumentation requirements are the same for both tests.
(D) 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 to verify that the engine power being delivered was not materially influencing landing capability or landing distances. The flare is maintained as long as is reasonable to dissipate speed and build RPM. Rotor RPM should stay with allowable limits. Aft center of gravity is ordinarily critical due to visibility and flare- ability.
Following the flare, the rotorcraft is allowed to touch down in a landing attitude. Rotor RPM at touchdown should be recorded, and it should be within allowable structural limits.
(viii) Vertical Landings. The reader should be familiar with the preceding discussion of conventional Category A, landing profiles because duplicate information is not repeated here. A typical vertical landing profile is shown in figure AC 29.75A-2.
This profile is equally applicable to both ground level and elevated heliport sites. The profile begins at a stabilized single engine approach condition. It should be possible to make a safe OEI landing or go-around at any point prior to the LDP unless alternate Page B - 85
AEO approach procedures are presented in the Flight Manual according to paragraph AC 29.75b(1)(i)(A). It is possible to have two landing techniques: an “offset” one, which schedules drop down for elevated heliports (but still ensure 15 feet radial deck edge clearance), and a “straight in” approach which utilizes the ground level heliport criteria. These techniques should be stipulated as such in the Flight Manual. At the LDP the aircraft becomes committed to landing. A safe landing should be possible in case of an engine failure at any point before or after the LDP. Testing should include a simulated failure at LDP with a 1-second delay or normal pilot response time, whichever is longer, and subsequent landing within the allowable area. The landing distance is the distance from the point at which the lowest portion of the rotorcraft reaches 50 feet above the landing surface to the forward-most point after coming to a stop (including main rotor tip path). The LDP becomes very important for landing on small, elevated heliports. The LDP should be clearly defined and Flight Manual instructions should carefully explain any pilot procedures. An illustration similar to figure AC 29.75A-2 with somewhat more detailed information is most useful. Night OEI landings should be conducted to verify suitable visibility for both internal and external vertical landing cues. The minimum elevated heliport size demonstrated for the OEI approach procedure and for alternate AEO approach procedures (when provided) should also be provided in the Flight Manual.
c. Category B Requirements.
(1) Explanation. Section 29.83 contains the Category B landing requirements.
Landing distance is measured from the 50-foot point to the point at which the rotorcraft is completely stopped (approximately 3 knots for water landings). The approach speed is selected by the applicant. Appropriate ambient conditions and allowable extrapolation are discussed under paragraph AC 29.45.
(2) Procedures.
(i) Landing Distance. Aft center of gravity is ordinarily critical due to field-of-view and flare ability. For wheeled rotorcraft, the brakes are applied to an incipient skid for most efficient stopping. For rotorcraft on skids, the collective should be lowered as soon as characteristics allow in order to place a greater weight on the landing skids. These procedures would be appropriate flight manual entries to show how landing distances can be realized. For flight manual purposes, the landing distance should include the horizontal distance from the point at which the lowest part of the rotorcraft first reaches 50 feet above the landing surface to the point at the foremost part of the rotorcraft (including rotor tip path) after coming to a stop. Multiengine rotorcraft incorporating Category A engine isolation features may elect to show compliance with § 29.79 and § 29.81. A sufficient number of acceptable runs should be accomplished to provide confidence in the results. Typically ten acceptable runs are adequate. If a weight effect on landing distance is to be shown, a minimum of two weight extremes are normally tested.
(ii) All-Engine-Out Landing.
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(A) Several procedures can be utilized to demonstrate compliance with the all-engine-out landing requirement. Section 29.83(c) requires that a landing from autorotation be possible. The maneuver is entered by smoothly reducing power at an optimum autorotation airspeed at a safe height above the landing surface.
All-engine-out landing tests should be initiated at light weight with a gradual buildup to the limiting weight conditions. If a complete company test program has documented all-engine-out landings to the GW/σ limit, the buildup conditions during verification test may be decreased. This test is ordinarily conducted at mid center of gravity. Typically, all altitudes may be approved with two weight limit landings - one at sea level and one near maximum takeoff and landing altitude.
(B) Demonstrated compliance with this requirement is intended to show that an autorotative descent rate can be arrested, and forward speed at touchdown can be controlled to a reasonable value (less than 40 KTAS is recommended) 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 low rotor inertia, considerable collective 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 collective pitch to control rotor overspeed during autorotative descent, particularly at high altitude WAT limiting conditions.
(C) 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. Other methods may be considered if they lead to reasonable assurance that descent can be arrested and forward speed 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. Considerations for weight and altitude extrapolation are the same as those for HV testing (see paragraph AC 29.79). The all -engine-inoperative landing test is ordinarily done in conjunction with height velocity tests because ground and onboard instrumentation requirements are the same for both tests.
(D) 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 Page B - 87
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 to verify that the engine power being delivered as not materially influencing landing capability or landing distances. The flare is maintained as long as is reasonable to dissipate speed and build RPM. Rotor RPM should stay with allowable li mits. Aft center of gravity is ordinarily critical due to visibility and flareability. Following the flare, the rotorcraft is allowed to touch down in a landing attitude. Rotor RPM at touchdown should be recorded, and it should be within allowable structural limits.
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9/30/99 АС 29-2C VroSs LDP 50 tt.
LANDING DISTANCE FIGURE AC 29.75A-1 CATEGORY A CONVENTIONAL LANDING - CLEAR HELIPORT Page B - 89
АС 29-2C 9/30/99 VrOSS Y TOSS 15 ft FIGURE AC 29.75A-2 CATEGORY A VERTICAL LANDING Page B - 90
AC 29.77. § 29.77 (Amendment 29-24) BALKED LANDING: CATEGORY A (For § 29.77 after Amendment 38, see paragraph AC 29.75A)
a. Explanation. This ru le has two distinct portions.
(1) Section 29.77(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 29.77(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 29.75-1.) The safe climb conditions are defined in § 29.67(a)(1) and (2). This suggests establishing a clearly defined balked landing profile similar to the Category A takeoff profile established under § 29.59. 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 up to and including the LDP. The LDP should be designated so that the balked landing profile may be completed with the rotorcraft descending no lower than 35 feet above the landing surface. 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 compared 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.
(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.
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AC 29-2C, Chg 4 5/1/2014 AC 29.79. § 29.79 (Amendment 29-21) LIMITING HEIGHT-SPEED ENVELOPE.
(For § 29.79 after Amendment 29-38, see section 29.75A of this AC.)
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 29.79-1). T he 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) Engine power considerations are similar to those in previous takeoff and landing requirements (see sections 29.53, 29.63, and 29.75 of this AC).
(3) The prohibited sections 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 § 29.143(d)). Altit udes 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 limitation) and normal pilot reaction time.
(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 t hat 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 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 airs peeds 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 exerci sed 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.
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(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