started at the desired W/σ with maximum continuos power, or at VNE, in level flight. A series of points should be taken, reducing airspeed 10 to 15 knots between points, with the lowest speed point at approximately 20 to 30 knots. Weight should be computed for each point and the test altitude adjusted to maintain a constant W/σ. After the data are reduced to standard day conditions, the minimum power required airspeed will result in the airspeed for maximum rate of climb. However, aircraft stability may suggest that a higher climb speed may be used for VY.
(3) Prior to the flight test, the rotorcraft should be ballasted to the desired gross weight and the critical center of gravity. The airspeed should be stabilized prior to data acquisition. Data to be recorded includes time, altitude, airspeed, ambient temperature, engine parameters, torque(s), rotor RPM, fuel reading, aircraft heading, external configuration, etc. Power setting, weight, and climb airspeed should be planned prior to flight. For some turboshaft engines, temperature and/or engine speed limits may be reached prior to a limiting torque. The test team should verify that the resulting power utilized in these tests closely approximates the power producing capabilities of installed minimum specification engine.
c. Procedure to Determine all Engine Operating Climb Performance.
(1) Background. Continuous climbs are conducted at the appropriate climb airspeeds as outlined above in order to obtain the rotorcraft’s climb performance for the flight manual. By -products are a qualitative evaluation of the rotorcraft handling characteristics in a climb and engine data to assist in the determination of installed power available.
(2) Techniques. The techniques used to determine this performance may be the same as those used in the VY determination. The climbs are conducted on reciprocal headings at the established airspeed(s) through the target altitude range.
The same parameters are recorded. The rotorcraft will usually climb very rapidly during the first few thousand feet; therefore, the data acquisition method should be timely if accurate results are expected. This procedure is usually repeated at weight extremes.
The resulting data should then be corrected for power and weight. Power and weight corrections are satisfactory, provided the test powers and weights closely approximate the target values to make the weight and power corrections accurate. Once this data is finalized and corrected for all the flight test variables, interpolation for intermediate weights can be made with a high degree of reliability. If the rotorcraft has any stability augmentation system, vent systems, etc., which may influence the climb performance, then it should be accounted for. Caution should be taken that anti -ice, air-conditioning, etc., are not on unless the performance is being established specifically for those conditions.
AC 29.67. § 29.67 (Amendment 29-34) CLIMB: ONE ENGINE INOPERATIVE.
a. Explanation.
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(1) Section 29.67 requires that Category A rotorcraft must be capable of a steady rate-of-climb without ground effect, of at least 100 feet per minute for all combinations of weight, altitude, temperature, and center of gravity for which takeoffs are to be scheduled. The rate- of-climb is determined with the critical engine inoperative and the remaining engine(s) operating within approved operating limits. The landing gear is extended and the airspeed is the takeoff safety speed (VTOSS) selected by the applicant.
(2) In addition, the steady rate-of-climb must be at least 150 feet per minute at 1,000 feet above the takeoff surface for which takeoffs are to be scheduled. The rate-of-climb will be determined with the critical engine inoperative and the remaining engine(s) at maximum continuous or the 30-minute minimum specification installed power available values. The landing gear is retracted and the airspeed is that selected by the applicant.
b. Procedures.
(1) One of the acceptable procedures used to obtain the required climb performance is similar to the all engine climb performance determination (paragraph AC 29.65) except that the VTOSS and the Category A climb speed may be selected by the applicant for different weights and ambient conditions. The Category A climb speed could be a single speed, vary as VY does, or actually be VY. Making a Category A climbout speed equal to VY should be encouraged to simplify cockpit procedures. The required results are the allowable weight, altitude, and temperature combinations wherein the rotorcraft is capable of demonstrating 100 feet per minute rate-of-climb at VTOSS and 150 feet per minute rate-of-climb at 1,000 feet above the takeoff surface. Either of these two climb requirements may establish the maximum allowable takeoff weight.
(2) For multiengine Category B rotorcraft with engine isolation, the steady rate of climb or descent must be determined at VY, using maximum continuous power and 30-minute power if that rating is approved. Appropriate performance data must be included in the Rotorcraft Flight Manual to cover variations in gross weight, altitude, and temperature.
(3) Since climb performance testing is normally conducted separately from Category A and B takeoff performance testing, it is imperative the engine power(s), rotor RPM, and aircraft configuration be the same as those used during the takeoff testing to ensure the climb performance demonstrated will be that attainable immediately after an engine failure during takeoff. The allowable pilot/crew actions during the Category A takeoff and climbout maneuver must be thoroughly evaluated. The pilot’s full attention is required to control the rotorcraft during this phase of flight. Permitting the pilot to readjust (beep) the rotor RPM during this phase of flight should be considered only if such adjustment can be accomplished without a significant increase in pilot workload.
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(4) A typical sequence for selecting the various speeds to comply with this requirement is as follows:
(i) Conduct sawtooth climbs at the various airspeeds (VY and below) up to the proposed takeoff and landing altitudes. From this a determination can be made regarding the maximum allowable weight that will result in a rate of climb of 150 feet per minute at the selected VY for the proposed ambient conditions.
(ii) At the same time determine the minimum value of VTOSS that will result in 100 feet per minute rate of climb at the maximum weight determined in (b)(4)(i).
AC 29.67A. § 29.67 (Amendment 29-39) CLIMB: ONE ENGINE INOPERATIVE.
a. Explanation.
(1) Amendment 29-39 expanded the OEI rate of climb requirements. The guidance material presented in paragraph AC 29.67 does not apply to rotorcraft certified with Amendment 29-39 or later. Section 29.67 requires that Category A rotorcraft should be capable of a steady rate-of-climb without ground effect 200 feet above the takeoff surface, of at least 100 feet per minute for all combinations of weight, altitude, temperature, and center of gravity for which takeoffs are to be scheduled. The rate-of-climb is determined with the critical engine inoperative and the remaining engine(s) operating within approved operating limits. The landing gear is extended and the airspeed is the takeoff safety speed (VTOSS) selected by the applicant.
(2) The steady rate-of-climb should be at least 150 feet per minute at 1,000 feet above the takeoff surface for which takeoffs are to be scheduled. The rate- of-climb will be determined with the critical engine inoperative and the remaining engine(s) at maximum continuous or the 30-minute minimum specification installed power available values. The landing gear is retracted and the airspeed is that selected by the applicant.
(3) Additionally, the steady state rate of climb or descent must be determined with the critical engine inoperative and the remaining engines at OEI maximum continuous power and at 30-minute OEI power if applicable. This performance must be scheduled throughout the ranges of weight, altitude and temperatures for which certification is requested with the landing gear retracted, at an airspeed selected by the applicant.
b. Procedures.
(1) One of the acceptable procedures used to obtain the required climb performance is similar to the all engine climb performance determination (paragraph AC 29.65) except that the VTOSS and the Category A climb speed may be selected by the applicant for different weights and ambient conditions. The Category A climb speed could be a single speed, vary as VY does, or actually be VY. Making a Category A climbout speed equal to VY should be encouraged to simplify cockpit Page B - 65
procedures. The required results are the allowable weight, altitude, and temperature combinations wherein the rotorcraft is capable of demonstrating 100 feet per minute rate-of-climb at VTOSS at a height of 200 feet above the takeoff surface and 150 feet per minute rate-of-climb at 1,000 feet above the takeoff surface. Either of these two climb requirements may establish the maximum allowable takeoff weight.
(2) For multiengine Category B rotorcraft with engine isolation, the steady rate of climb or descent should be determined at VY, using maximum continuous power, maximum continuous OEI power, and 30-minute power if that rating is approved.
Appropriate performance data should be included in the Rotorcraft Flight Manual to cover variations in gross weight, altitude, and temperature.
(3) Since climb performance testing is normally conducted separately from Category A and B takeoff performance testing, it is imperative the engine power(s), rotor RPM, and aircraft configuration be the same as those used during the takeoff testing to ensure the climb performance demonstrated will be that attainable immediately after an engine failure during takeoff. The allowable pilot/crew actions during the Category A takeoff and climbout maneuver should be thoroughly evaluated. The pilot’s full attention is required to control the rotorcraft during this phase of flight. Permitting the pilot to readjust (beep) the rotor RPM during this phase of flight should be considered only if such adjustment can be accomplished without a significant increase in pilot workload.
(4) A typical sequence for selecting the various speeds to comply with this requirement is as follows:
(i) Conduct sawtooth climbs at the various airspeeds (VY and below) up to the proposed takeoff and landing altitudes. From this, a determination can be made regarding the maximum allowable weight that will result in a rate of climb of 150 feet per minute at the selected VY for the proposed ambient conditions.
(ii) At the same time, determine the minimum value of VTOSS that will result in 100 feet per minute rate of climb at the maximum weight determined in b(4)i.
AC 29.71. § 29.71 (Amendment 29-12) ROTORCRAFT ANGLE OF GLIDE:
CATEGORY B.
a. Explanation.
(1) Performance capabilities during stabilized autorotative descent are useful pilot tools to assist in the management of a Category B rotorcraft when all engines fail.
This information is also useful in determining the suitability of available landing areas along a given route segment.
(2) Two speeds are of particular importance, the speed for minimum rate of descent and the speed for best angle of glide. These speeds are required as flight manual entries per § 29.1587. The speed for minimum rate of descent is useful for Page B - 66
engine failure conditions at higher altitudes and the pilot is required to perform some time-related task, engine restart, float inflation, radio calls, etc. The speed for best angle of glide is a somewhat higher speed that is of particular use when it is necessary to reach a distant landing area. This speed, with appropriate rotor RPM, provides the maximum horizontal distance available from a particular altitude assuming zero wind conditions.
(3) A third speed, recommended autorotation speed, may be provided in addition to minimum rate of descent speed and maximum glide angle speed. The recommended speed for autorotation is usually optimized to assure an effective flare capability and yet be slow enough to allow a controlled, relatively slow touchdown condition. Recommended autorotation speed is ordinarily between the minimum rate of descent and maximum glide angle speeds. The recommended autorotation speed may be provided in the Rotorcraft Flight Manual. The relationship between minimum rate of descent, best glide angle, and recommended autorotation speed is shown in figure AC 29.71-1.
(4) Forward center of gravity is usually critical, however, center of gravity effects should be spot-checked to confirm this for a given design.
b. Procedures.
(1) Tests are conducted at speeds which bracket the anticipated speeds for minimum rate of descent and best glide angle. On a power required plot, the speed for minimum power required approximates the speed for minimum rate of descent. The speed for maximum range glide may be estimated by drawing a tangent from the origin to the power required curve.
(2) Autorotative performance tests may be conducted in conjunction with the climb performance tests. The required data are similar for both tests and it is sometimes convenient and efficient to run alternating climbs and descents through a desired altitude band. Descents should be conducted on reciprocal headings and results averaged in the same manner as climb performance tests.
(3) A reduction in rotor RPM from the normal power-on value may enhance autorotative performance. If the applicant wishes to develop autorotative performance at RPM values significantly below the governing or power-on range, the practicality of reducing and controlling RPM at the lower value and of then increasing RPM as a landing is approached, must be considered. At low weights and low density altitudes, full down collective may automatically produce lower RPM values and this condition is, of course, acceptable provided the approved power-off RPM range is not exceeded.
(4) Care must be taken to make certain that no engine power is delivered to the rotor drive system since a very small amount of power can have a large effect on descent performance.
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RATE OF
DESCENT RECOMMENDED
AUTOROTATIONAL SPEED
BEST GLIDE ANGLE
MINIMUM
RATE OF DESCENT
AIRSPEED FIGURE AC 29.71-1 AUTOROTATIONAL CHARACTERISTICS - TYPICAL Page B - 68
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