AC 29.73. § 29.73 (Amendment 29-3) PERFORMANCE AT MINIMUM OPERATING SPEED. HOVER PERFORMANCE FOR ROTORCRAFT .
(For performance at minimum operating speed and for hover performance after Amendment 38, see § 29.49 and paragraph AC 29.49).
a. Explanation.
(1) For the purpose of this manual, the word “hover” applies to a rotorcraft that is airborne at a given altitude over a fixed geographical point regardless of wind. Pure hover is accomplished only in still air.
(2) The regulatory requirement for hover performance, § 29.73, refers to hover in ground effect (IGE). For some applications, such as external load operations, hover performance out-of-ground effect (OGE) is necessary; however, it is not required by this section. Hover OGE is that condition, where an increase in height above the ground will not require additional power to hover. Hover OGE is the absence of measurable ground effect. It can be less than one rotor diameter at low gross weight increasing significantly at high gross weights. The lowest OGE hover height at gross weight may be approximated by placing the lowest part of the vehicle 1 ½ rotor diameters above the surface.
(3) The objective of hover performance tests is to determine the power required to hover at different gross weights, ambient temperatures, and pressure altitudes.
Using nondimensional power coefficients (CP) and thrust coefficients (CT) for normalizing and presenting test results, a minimum amount of data are required to cover the rotorcraft’s operating envelope.
(4) Hover performance tests must be conducted over a sufficient range of pressure altitudes and weights to cover the approved ranges of those variables for takeoff and landings. Additional data should be acquired during cold ambient temperatures, especially at high altitudes, to account for possible Mach effects.
(5) The minimum hover height for which data should be obtained and subsequently presented in the flight manual should be the same height consistent with the minimum hover height demonstrated during the takeoff tests. Refer to paragraph AC 29.51 for the procedure to determine the minimum allowable hover height.
b. Procedures.
(1) Two methods of acquiring hover performance data are the tethered and free flight techniques. The tethered technique is accomplished by tethering the rotorcraft to the ground using a cable and load cell. The load cell and cable are attached to the ground tie-down and to the rotorcraft cargo hook. The load cell is used to measure the rotorcraft’s pull on the cable. Hover heights are based on skid or wheel height above Page B - 69
the ground. During tethered hover tests, the rotorcraft should be at light gross weight.
The rotorcraft will be stabilized at a fixed power setting and rotor speed at the appropriate skid or wheel height. Once the required data are obtained, power should be varied from the minimum to the maximum allowed at various rotor RPM. This technique will produce a large CT/CP spread. The load cell reading is recorded for each stabilized point. The total thrust the rotor produces is the rotorcraft’s gross weight, weight of the cables and load cell plus cable tension. Care must be taken that the cable tension does not exceed the cargo hook limit or load capacity of the tie-down. For s ome rotorcraft, it may be necessary to ballast the rotorcraft to a heavy weight in order to record high power hover data.
(2) The pilot maintains the rotorcraft in position so that the cable and load cell are perpendicular to the ground. To insure the cable is vertical, two outside observers, one forward of the rotorcraft and one to one side, can be used. Either hand signals or radio can be used to direct the pilot. The observers should be provided with protective equipment. This can also be accomplished by attaching two accelerometers to the load cell which sense movement along the longitudinal and lateral axes. Any displacement of the load cell will be reflected on instrumentation in the cockpit and by reference to this instrumentation, the rotorcraft can be maintained in the correct position. Increased caution should be utilized as tethered hover heights are decreased because the rotorcraft may become more difficult to control precisely. The tethered hover technique is especially useful for OGE hover performance data because the rotorcraft’s internal weight is low and the cable and load cell can be jettisoned in the event of an engine failure or other emergency.
(3) To obtain consistent data, the wind velocity should be less than 3 knots or less as there are no accurate methods of correcting hover data for wind effects. Large rotorcraft with high downwash velocities may tolerate higher wind velocities. The parameters usually recorded at each stabilized condition are:
(i) Engine torques.
(ii) Rotor speed.
(iii) Ambient temperatures.
(iv) Pressure altitude.
(v) Fuel used (or remaining).
(vi) Load cell reading.
(vii) Generator(s) load.
As a technique, it is recommended the rotorcraft be loaded to a center of gravity near the hook to minimize fuselage angle changes with varying powers. All tethered hover Page B - 70
data should be verified by a limited spotcheck using the free flight technique. The free flight technique as contained in paragraph b(4) below will determine if any problems, such as load cell malfunctions have occurred. The free flight hover data must fall within the allowable scatter of the tethered data.
(4) If there are no provisions or equipment to conduct tethered hover tests, the free flight technique is also a valid method. The disadvantage of this technique as the primary source of data acquisition is that it is very time consuming. In addition a certain element of safety is lost OGE in the event of emergency. The rotorcraft must be reballasted to different weights to allow the maximum CT/CP spread. When using the free flight technique, either as a primary data source or to substantiate the tethered technique, the same considerations for wind, recorded parameters, etc., as used in the tethered technique apply. Free flight hover tests should be conducted at CG extremes to verify any CG effects. If the rotorcraft has any stability augmentation system which may influence hover performance, it must be accounted for.
(5) It is extremely difficult to determine when a rotorcraft is hovering OGE at high altitudes above ground level since there is no ground reference. In a true hover, the rotorcraft will drift with the wind. Numerous techniques have been tried to allow OGE hover data acquisition at high altitudes, all of which have resulted in much data scatter. Until a method is proposed and found acceptable to the FAA/AUTHORITY, OGE hover data must be obtained at the various altitude sites where IGE hover data is obtained. Hover performance can usually be extrapolated up to a maximum of 4,000 feet.
AC 29.75. § 29.75 (Amendment 29-17) LANDING.
a. Explanation.
(1) This rule incorporates all of the landing performance requirements for transport category rotorcraft. It consolidates requirements for landing data, Category A landing, Category A flight data, and Category B landing. Parallel takeoff requirements are located in four separate sections of the rule, §§ 29.51 through 29.63. As such, to assure necessary subjects are treated separately, the following discussion will be separated into three parts: (a) a general discussion of basic landing distance requirements, (b) Category A requirements (including vertical landing), and
(c) Category B requirements.
(2) All landing performance data are corrected to a smooth, dry, hard, level landing surface condition. As with other flight maneuvers, landings must be accomplished with acceptable flight and ground characteristics using normal pilot skills.
The rule states that Category A and B landing data must be determined at each approved WAT (Weight, Altitude, Temperature) condition. Reasonable sampling and extrapolation methods are, of course, allowed. General guidance on those subjects is given in paragraph AC 29.45. As in other performance areas, engines must be Page B - 71
operated within approved limits. Power considerations are the same as those described under paragraph b(2)(ii)(C).
(3) Unlike fixed wing aircraft, rotorcraft typically require significantly more landing surface area with an engine inoperative than with all engines operating.
Because of this characteristic, the landing distance requirements are met with at least one engine inoperative to assure the most conservative landing distance measurement is achieved.
b. Procedures.
(1) Category A Requirements.
(i) Explanation. The Category A certification concept limits landing weight to a value that will allow the rotorcraft, following an engine failure at the landing decision point (LDP), to land within the available runway or to execute a balked landing, descending no lower than 35 feet above the landing surface. See figure AC 29.75-1.
(A) LDP. The Category A landing profile begins with an assumed engine failure at or prior to the LDP. The LDP is typically defined in terms of airspeed, rate of descent, and altitude above the landing surface. The approach path angle can be defined by LDP airspeed and rate of descent values. Definition of the LDP should include an approach angle because both the landing distance and the missed approach path are significantly influenced by landing approach angle. At any point in the single engine approach path down to and including the LDP, the pilot may elect to land or to execute a balked landing and he is assured both an adequate surface area for OEI landing and adequate climb capability for an OEI balked landing. Said another way, if an engine fails at any point down to and including the LDP, the pilot may safely elect to land or to “go around” by executing a balked landing. The LDP must be defined to permit acceleration to VTOSS at an altitude no lower than 35 feet above the landing surface. The LDP represents a “commit” point for landing. Prior to the LDP in the one engine inoperative approach, the pilot has a choice, he may either land or fly away.
After passing the LDP he no longer has sufficient energy to assure transition to a balked landing condition without contacting the landing surface. If an engine fails after LDP in a normal (all engine) landing the pilot is committed to land. The LDP and landing approach path must be defined such that the critical areas of the height-velocity diagram are avoided. A typical LDP for conventional Category A profiles is 100 feet above the landing surface. LDP should be specified in terms of both actual altitude above the landing surface and indicated barometric altitude. Speed at the LDP should be specified in terms of indicated airspeed.
(B) Landing distance. Approach and landing path requirements are stated in general terms in paragraphs (b)(2) and (4) of § 29.75. The approach path must allow smooth transition for one engine inoperative landing and for balked landing maneuvers and must allow adequate clearance from potentially hazardous HV combinations. Paragraph (b)(4)(ii) implies that a less restrictive HV envelope may exist Page B - 72
for the Category A approach condition in comparison to that determined under high power conditions in § 29.79. The manufacturer may elect to use this added capability.
The added capability arises from the fact that lower power levels, a lower collective setting, and an established rate of descent accompany typical approach conditions as opposed to the more critical high power conditions of § 29.79. Landing distance is measured from a point 50 feet (25 feet for VTOL) above the landing surface to a stop.
For flight manual purposes, the distance is from the point at which the lowest part of the rotorcraft first reaches 50 feet (25 for VTOL) to the foremost point of the rotorcraft (including rotor tip path) after coming to a stop.
(C) All engine out landing. Section 29.75(b)(5) contains the Category A certification requirement for “last” engine failure and all engine inoperative landing. The rule states that it must be possible to make a safe landing on a prepared surface after complete power failure during normal cruise. It is not intended that all engines be failed simultaneously. See paragraph AC 29.143a(2)(iii)(A) for the Category A sequential engine failure criteria. The conditions for last engine failure are maximum continuous power or 30-minute power if that rating is approved, “wings” level flight, and sudden engine failure with a pilot delay of 1 second or normal pilot recognition time, whichever is greater. Complete power failure has occurred in twin engine Category A rotorcraft.
This requirement ensures that in the event of cockpit mismanagement, fuel exhaustion, improper maintenance, fuel contamination, or unforeseen mechanical failures, a safe autorotation entry can be made and a safe power-off landing can be affected. Two separate aspects of this rule are normally evaluated at different times during the test program. The last engine failure is normally evaluated during cruise or V NE engine failure testing where instrumentation and critical loading have been established for those test conditions. See discussion under paragraph AC 29.143. The all engine out landing is ordinarily conducted in conjunction with an HV or Category A landing distance phase where ground instrumentation and safety equipment are available. The rotorcraft must be capable of conducting the all engine out landing at the takeoff and landing WAT limiting conditions up to the maximum altitude approved for takeoff and landing.
(ii) Procedures.
(A) Instrumentation/Equipment. Instrumentation requirements are basically the same as those for Category A takeoff. A photo theodolite, grid camera, or other position measuring equipment is needed, along with a ground station to measure wind, OAT, and humidity (if applicable). A two- way communication system between the aircraft and the position measuring equipment is essential. Aircraft instrumentation should include engine and flight parameters, control positions, power lever position, landing gear loads, and a method for synchronizing power cuts between the external light normally used for photo theodolite or camera, and onboard instrumentation. A record of rotor RPM at touchdown is necessary to assure it does not exceed transient limits. Rotor RPM at touchdown may be lower than the minimum transient limit for flight, provided stress limits are not exceeded. A crash recovery team with support of a fire engine is highly desirable.
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(B) Establishing the LDP.
(1) 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 § 29.77(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.
(2) 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, cautions, 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 at the LDP with a 1-second pilot delay to assure safe landing capability for this critical case. A minimum of five acceptable runs for distance should be flown by the FAA/AUTHORITY pilot. These may be averaged with an equal number of acceptable runs by the company pilot.
(3) The balked landing portion of the landing profile is addressed under § 29.77, Balked Landing: Category A. For an explanation of that requirement and a discussion of those test procedures refer to paragraph AC 29.71.
(C) Power. Power should be limited to minimum specification values on the operating engine(s). This may be accomplished by adjustment of engine topping to minimum specification values for the range of atmospheric variables to be approved.
This is frequently done by installing an adjustable device in the throttle linkage with a control in the cockpit so that engine topping can be accurately adjusted for varying ambient conditions. With such a device in the control sy stem it becomes vitally important to check topping power prior to each test sequence.
(D) 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 Page B - 74
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.
(E) Extrapolation. Weight cannot be extrapolated above test weight. See discussion under Height-Velocity Testing in paragraph AC 29.79. If no marginal areas are apparent and an acceptable analytical method is used, performance data may be extrapolated ±4,000 feet density altitude from test conditions. (See paragraph AC 29.45.)
(F) Ambient Conditions. Appropriate test limits for ambient conditions such as wind and temperature are contained in paragraph AC 29.45. Test data must be corrected for existing wind conditions during landing distance tests. Credit for headwind conditions may be given during flight manual data expansion. Paragraph AC 29.45 details allowable wind credit.
(G) All engine out landing.
(1) Several procedures can be utilized to demonstrate compliance with the all engine out landing requirement. As discussed in the explanation portion of this paragraph, § 29.75(b) contains two separate requirements. One is the ability to transition safely into autorotation after failure of the last operative engine. This requirement is discussed in paragraph AC 29.143. The second aspect of this rule requires that a landing from autorotation be possible on a prepared surface. The second requirement is discussed below. The maneuver is entered by smoothly reducing power at an optimum autorotation airspeed at a safe height above a prepared landing surface. If a complete company test program has documented an all engine out landing to the GW/σ (gross weight/density ratio) limit for takeoff and landing at each altitude, verification tests may be initiated at those limiting weight conditions. 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.
(2) 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 assure a reasonable chance of survivability for the all engine failure condition. The touchdown speed (less than 40 KIAS is recommended) should be consistent with the type design limits including landing gear capability, aircraft visibility, and any other factors affecting repeatability of the maneuver. On Category A rotorcraft, rotor inertia is typically much lower than for single engine rotorcraft. RPM decays rapidly when the last engine is made inoperative. Also, due to this relatively low inertia level, considerable collective may be needed to prevent rotor overspeed conditions when the rotorcraft is flared for landing. Also, when testing 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 Page B - 75
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.
(3) 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 assuring 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 Category A rotorcraft totally without power. Considerations for weight and altitude extrapolation are the same as those for HV testing (reference 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.
(H) 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 29.75-2. This profile is equally applicable to both ground level and pinnacle sites. The profile begins at a stabilized single engine approach condition. It must be possible to make a safe OEI landing or go-around at any point prior to the LDP. At the LDP the aircraft becomes committed to landing. A safe landing must 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 LDP is typically well above the 25-foot point from which landing distance is measured. The landing distance is the distance from the point at which the lowest portion of the rotorcraft reaches 25 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 must be clearly defined and flight manual instructions should carefully explain any pilot procedures. An illustration similar to figure AC 29.75-2 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.
c. Category B Requirements.
(1) Explanation. Section 29.75(c) contains the Category B landing requirements. For rotorcraft that do not meet the Category A powerplant installation requirements of this part, landing tests are conducted with all engines inoperative in an autorotative descent condition. Landing distance is measured from the 50- foot point to the point at which the rotorcraft is completely stopped (approximately 3 knots for water Page B - 76
landings). The autorotative approach speed is selected by the applicant. The landing maneuver is similar to that referred to during normal training flights as a practice autorotation. As in HV tests, care must be taken to assure no power is delivered to the rotor drive system during these tests. A small amount of power can have a significant effect on landing test results. Multiengine rotorcraft incorporating Category A engine isolation features may conduct landing distance tests with only one engine inoperative using the procedures prescribed above for Category A. For these rotorcraft the one engine inoperative condition typically results in much shorter distances due both to a much lower speed at the 50-foot point and the added power available for flaring and cushioning the landing. Instrumentation requirements are the same as those described under Category A above. Appropriate ambient conditions and allowable extrapolation are discussed under paragraph AC 29.45.
(2) Procedures. Prior to conducting these tests the crew should be familiar with the engine inoperative landing characteristics of the rotorcraft. For Category B rotorcraft without engine isolation, 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 landing flare may be initiated prior to the 50- foot point. The flare is maintained as long as is reasonable to dissipate speed and build RPM. Rotor RPM must stay within allowable limits. Aft center of gravity is ordinarily critical due to visibility and flare-ability. Following the flare, the rotorcraft is allowed to touchdown in a landing attitude. Rotor RPM at touchdown should be recorded and it must be within allowable structural limits. 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. For Category B rotorcraft with engine isolation, the landing procedures are as described for Category A landing.
When conducting Category B landings utilizing Category A “procedures,” § 29.75(b)(2) can be misleading. No transition capability to balked landing is intended for Category B rotorcraft. Section 29.77, Balked Landing, Category A, applies only to Category A rotorcraft and not to Category B rotorcraft which incorporates Category A “design” features. Five acceptable landing runs should be flown by the FAA/AUTHORITY pilot at each test weight. Results may be averaged with an equal number of company runs. If a weight effect on landing distance is to be shown, a minimum of two weight extremes are normally tested.
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VrOSs Vy LDP VTOSS 50 FT.
35.F
LANDING DISTANCE FIGURE AC 29.75-1 CATEGORY A CONVENTIONAL LANDING - CLEAR HELIPORT Page B - 78
VTOSS J LDP VTOSS 35 ft 25 ft. Landing Distance FIGURE AC 29.75-2 CATEGORY A VERTICAL LANDING Page B - 79
AC 29-2C, Chg 4 5/1/2014 AC 29.75A. (AC's 29.77, 29.79, 29.81, & 29.83) §§ 29.75, 29.77,
29.79, 29.81, and 29.83 (Amendment 29-39) LANDING.
(For § 29.77 and § 29.79 prior to Amendment 29-39, see paragraphs AC 29.77 and AC 29.79 respectively.)
a. Explanation.
(1) Amendment 29-39 revised and relocated many of the landing requirements of Part 29. Changes were made to the general landing requirements of § 29.75. New requirements were added for designating a landing decision point (LDP) in § 29.77.
The original § 29.79 was redesignated as a new § 29.87. Category A landing requirements were established in a new § 29.79. Requirements were added to determine landing distances in a new § 29.81. Revised Category B landing requirements were relocated from § 29.75(c) into a new § 29.83. The guidance material from paragraph AC 29.75 does not apply to rotorcraft certified with Amendment 29-39 or later.
(2) These rules incorporate all of the landing performance requirements for transport category rotorcraft. They contain the requirements for landing data, Category A landing, and Category B landing. Parallel takeoff requirements are located in eight separate sections of the rule, §§ 29.51 through 29.63. As such, to ensure that necessary subjects are treated separately, the following discussion will be separated into three parts: (a) a general discussion of basic landing distance requirements,
(b) Category A requirements (including vertical landing), and (c) Category B requirements.
(3) All landing performance data are corrected to a smooth, dry, hard, level landing surface condition. As with other flight maneuvers, landings should be accomplished with acceptable flight and ground characteristics using normal pilot skills.
The rule states that Category A and B landing data should be determined at each approved WAT (Weight, Altitude, Temperature) condition. Reasonable sampling and extrapolation methods are, of course, allowed. General guidance on those subjects is given in paragraph AC 29.45. As in other performance areas, engines should be operated within approved limits. Power considerations are the same as those described under paragraph b(1)(ii)(C).
(4) Unlike fixed-wing aircraft, rotorcraft typically require significantly more landing surface area with an engine inoperative than with all engines operating.
Because of this characteristic, the Category A landing distance requirements are met with at least one engine inoperative to ensure the most conservative landing distance measurement is achieved.
b. Procedures - Category A Requirements.
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