AC 27-1B, Chg 4 9/17/2009 AC 27.73. § 27.73 PERFORMANCE AT MINIMUM OPERATING SPEED (thru amendment 27-43; see § 27..49 effective with amendment 27-44).
All of the policy material pertaining to AC 27.73 remains in effect through Amendment
27-43. See § 27.49 and AC 27.49 effective with Amendment 27-44.
a. Explanation.
(1) 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. For the purpose of this manual, the word "hover" will mean pure hover.
(2) The regulatory requirement for hover performance, § 27.73, 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 weight. The lowest OGE hover height at gross weight may be approximated by placing the lowest part of the vehicle one and one-half rotor diameters above the surtace.
(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 minimizes the amount of data 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 landing. Additional data should be acquired during cold ambient temperatures, especially at high altitudes, to account for possible Mach effects.
(5) The hover ceiling 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 27.51 for the procedure to determine this hover height.
b. Procedures.
(1) Two methods of acquiring hover performance data are the tethered and the 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 the ground. During tethered hover tests, the rotorcraft should be at light Page B - 32
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 C/Cp spread. The load cell reading is recorded for each stabilized point. The total thrust the rotor produces is equal to the rotorcraft's gross weight plus the 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 some 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 cables and load cell are perpendicular to the ground. To ensure 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. Positioning can also be accomplished by attaching two accelerometers to the load cell which sense angle or 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 as there are no accurate methods of correcting hover data for wind effects. Rotorcraft with high downwash velocities may tolerate higher wind velocities. The parameters usually recorded at each stabilized condition are:
Engine torque.
(ii) Rotor speed.
(iii) Ambient temperatures.
(iv) Pressure altitude.
(v) Fuel used (or remaining).
(vi) Load cell reading.
(vii) Generator(s) load.
(viii) Wind speed and direction.
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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 data should be verified by a limited spotcheck using the free flight technique. The free flight technique as contained in paragraph AC 27.73b(4) 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 an emergency. The rotorcraft must be reballasted to different weights to allow the maximum C 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 are obtained. Hover performance can usually be extrapolated up to a maximum of 4,000 feet.
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AC 27.75. § 27.75 (Amendment 27-14) LANDING.
a. Explanation.
(1) This rule incorporates all of the landing requirements for Part 27 rotorcraft.
(2) As with other flight maneuvers, landings must be accomplished with acceptable flight and ground characteristics using normal pilot skills. Reasonable sampling and extrapolation methods are, of course, allowed. General guidance on those subjects is given in paragraph AC 27.45. As in other performance areas, engines must be operated within approved limits.
(3) Landing. Approach and landing path requirements are stated in general terms in paragraphs (a)(1) and (a)(2) of § 27.75. The approach path must allow smooth transition for a one-engine-inoperative landing and adequate clearance from potentially hazardous HV combinations.
(4) All-engine-out landing. Section 27.75(b) contains the certification requirement for "last" engine failure and all-engines-inoperative landing. The rule states that it must be possible to make a safe landing after complete power failure during normal cruise. It is not intended that all engines be failed simultaneously, although complete power failure has occurred in twin-engine rotorcraft with Category A engine isolation. This requirement assures 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 VNE engine failure testing where instrumentation and critical loading have been established for those test conditions. The all-engine-out landing is ordinarily conducted in conjunction with an HV or landing distance phase where ground instrumentation and safety equipment are available.
b. Procedures.
(1) Instrumentation/Equipment. Aircraft instrumentation may include engine and flight parameters, control positions, power lever position, and landing gear loads. 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 the support of a fire engine is highly desirable.
(2) The one-engine-inoperative landing is similar in many respects to the HV tests described in paragraph AC 27.79. Most of the comments, cautions, and techniques for HV also apply here even though the typical flight conditions 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 with one engine inoperative.
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AC 27-1B, Chg 4 3/31/2014 The speed is reduced and the rotorcraft is flared to a conventional one-engine-inoperative landing.
(3) 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 system it becomes vitally important to check topping power prior to each test sequence.
(4) Aircraft Loading. Aft center of gravity is usually most critical because visibility constraints limit the degree to which the pilot can see the landing surface during the flare. 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.
(5) All-engine-out landing.
Several procedures can be utilized to demonstrate compliance with the all-engine-out landing requirement. As discussed in the explanation portion of this paragraph, § 27.75(b) contains two separate requirements. One is the ability to transition safely into autorotation after failure of the last operative engine. The second aspect of this rule requires that a landing from autorotation be possible. The second requirement is discussed below. The maneuver is entered by smoothly reducing power at an optimum autorotation airspeed at a safe height above the landing surface. If a complete company test program has documented an all-engine-out landing to the GW/o (gross weight/density ratio) limit, 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.
(її) 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 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 Page 34.2
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.
(iii) The intent of this rule is to demonstrate controlled touchdown conditions and freedom from loss of control or apparent hazard to occupants when landing with all engines failed. In these cases compliance can be demonstrated by leaving throttles in the idle position and ensuring no power is delivered to the drive train.
Also, computer analysis may be used in conjunction with simulated in-flight checks to give reasonable assurance that an actual safe touchdown can be accomplished.
Another method may be to make a power recovery after flare effectiveness of the rotorcraft has been determined. 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. 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.
(6) Prior to conducting these tests, the crew should be familiar with the engine inoperative landing characteristics of the rotorcraft. The flight profile may be entered in the same manner as a straight-in practice autorotation. It is recommended that for safety reasons idle power be used if a "needle split" (no engine power to the rotor) can be achieved. In some cases, a low engine idle adjustment has been set to assure needle split is attained. In other cases a temporary detent between idle and cutoff was used on the throttle. In a third case the engine was actually shut down on sample runs 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 must stay within allowable limits. Aft center of gravity is ordinarily critical due to visibility and flarability. Following the flare, the rotorcraft is allowed to touch down in a landing attitude. Rotor RPM at touchdown should be recorded and it must be within allowable structural limits.
AC 27.75A. § 27.75 (Amendment 27-44) Landing.
a. Explanation. Amendment 27-44 rewords § 27.75(a) by replacing the word "glide" with "autorotation" and further clarifies that the one-engine-inoperative (OEI) approach is to be performed from an established OEI approach. The OEI approach should be made utilizing a normal helicopter approach angle of approximately 6° or some other angle determined to be acceptable to the FAA/Authority and properly referenced in the Rotorcraft Flight Manual (RFM).
b. Procedures. The policy material pertaining to this section remains in effect with the following changes and additions:
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AC 27-1B, Chg 4 9/17/2009
(1) Instrumentation/Equipment. Aircraft instrumentation may include engine and flight parameters, control positions, power lever position, and landing gear loads. 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 the support of a fire engine is highly desirable.
(2) The OEI landing is similar in many respects to the Height-Velocity (HV) tests described in AC 27.79 of this advisory circular. Most of the comments, cautions, and techniques for HV also apply here, even though the typical flight conditions 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 with OEI. The speed is reduced and the rotorcraft is flared to a conventional OEI landing. To show compliance, full OEl landing should be demonstrated from sea level to the maximum altitude capability of the rotorcraft or 7,000 feet, whichever is less, at the maximum landing weight, without damage. For altitudes above 7,000 feet to the maximum takeoff and landing altitude, compliance with this requirement is shown by demonstrating that the OEI descent rate and forward speed can be reasonably controlled.
(3) Power. Power should be limited to minimum specification values on the operating engines). This may be accomplished by adjustment of engine topping to minimum specification values for the range of the 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 system, it becomes vitally important to check topping power prior to each test sequence.
(4) Aircraft Loading. Aft CG is usually the most critical because visibility constraints limit the degree to which the pilot can see the landing surface during the flare. 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.
(5) All-engine-out landing.
(1) Several procedures can be utilized to demonstrate compliance with the all-engine-out landing requirement of § 27.75(a)(1) and § 27.75(b). These paragraphs contain two separate requirements. One is the ability to transition safely into autorotation after failure of the last operative engine. The second aspect of this rule requires that a landing from autorotation be possible. The second requirement is discussed below. The maneuver is entered by smoothly reducing power at an optimum autorotation airspeed at a safe height above the landing surface. Typically a full autorotation landing to touchdown is demonstrated at sea level standard day for the maximum landing gross weight for that altitude. For altitudes above standard sea level, Page 36
a demonstration of flare effectiveness with power recovery at the maximum landing gross weight corresponding to the altitude satisfies the requirement for touchdown. The flare must reduce the autorotative descent rate and forward speed to a reasonable value (l.e., autorotative descent rate and forward speed consistent with that demonstrated at sea level). If a complete company test program has documented the all-engine-out landing capability, verification tests may be initiated at those limiting weight conditions. If not, buildup testing should be initiated at light weight.
(ii) Demonstrated compliance with this requirement is intended to show that the autorotative descent rate and forward speed at touchdown (less than 40 knots true airspeed (KTAS) is recommended) can be controlled to a reasonable value to ensure a reasonable chance of survivability for the all engine failure condition. On multiengine rotorcraft, rotor inertia is typically lower than for single-engine rotorcraft. RPM decays rapidly when the last engine is made inoperative. Due to this relatively low inertia level, considerable application of collective pitch may be needed to prevent rotor overspeed conditions when the rotorcraft is flared for landing. Also, when testing the final maximum weight points, the pilot should anticipate a need for considerable application of collective pitch to control rotor overspeed during autorotative descent, particularly at high altitude. Some designs incorporate features that may lead to rotorcraft damage in testing this requirement (e.g., droop stop breakage or loss of directional control with skids) if landings are conducted to a full stop with the engines off.
(ili) The intent of this rule is to demonstrate controlled touchdown conditions and freedom from loss of control or apparent hazard to occupants when landing with all engines failed. In these cases, compliance can be demonstrated by leaving throttles in the idle position and ensuring no power is delivered to the drive train. Also, computer analysis may be used in conjunction with simulated in-flight checks to give reasonable assurance that an actual safe touchdown can be accomplished. Another method may be to make a power recovery after flare effectiveness of the rotorcraft has been determined, showing that the rate of descent and forward speed can be controlled to allow for safe landing. Other methods may be considered if they lead to reasonable assurance that rate of descent and forward speed can be controlled to allow safe landing with no injury to occupants when landing on a prepared surface with all engines failed. Regardless of the method(s) used to comply with this requirement, careful planning and analyses are very important due to the potentially hazardous aspects of power off simulation and landing of a multiengine rotorcraft totally without power. The OE landing test is ordinarily done in conjunction with height velocity tests because ground and onboard instrumentation requirements are the same for both tests.
(6) Prior to conducting these tests, the crew should be familiar with the engine inoperative landing characteristics of the rotorcraft. The flight profile may be entered in the same manner as a straight-in practice autorotation. It is recommended that for safety reasons, idle power be used if a "needle split" (no engine power to the rotor) can be achieved. In some cases, a low engine idle adjustment has been set to assure needle split is attained. In other cases, a temporary detent between idle and cutoff was used on the throttle. In a third case, the engine was actually shut down on sample runs Page B - 36.1
AC 27-1B, Chg 4 9/17/2009 to verify that the engine power being delivered was not materially influencing landing capability or landing distances. The flare is maintained as long as it is reasonable to dissipate speed and build RPM. Rotor RPM must stay within allowable limits. Aft CG is ordinarily critical due to visibility and pitch attitude effectiveness in a flare. Following the flare, the rotorcraft is allowed to touch down in a landing attitude. Rotor RPM at touchdown should be recorded, and it must be within allowable structural limits.
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AC 27.79. § 27.79 LIMITING HEIGHT-SPEED ENVELOPE.
All of the policy material pertaining to this AC section remains in effect through Amendment 27-43. See § 27.87 and this AC's section 27.87 effective with Amendment
27-44.
a. Explanation.
(1) The height-speed envelope is normally referred to as the height-velocity (HV) diagram. It defines an envelope of airspeed and height above the ground from which a safe power-off or one engine inoperative (OEI) landing cannot be made. The diagram normally consists of three portions: (a) the level flight (cruise) portion, (b) the takeoff portion, and (c) the high speed portion (see Figure AC 27.79-1). The high speed portion is omitted on occasions when it can be shown that the rotorcraft can suffer an engine failure at low altitude and high speed (up to VH) and make a successful landing or climb out on the remaining engine(s).
(2) Power failure, engine failure, throttle chop, or other similar terms used in this discussion mean a simulated engine failure. The actual shutdown of an engine to simulate an engine failure should not be necessary if the simulated procedure ensures that the engine power is suddenly removed from driving the rotor and remains so. The normal fuel control deceleration schedule is usually satisfactory for the power removal for turbine engines but the flight or ground idle speed may have to be set lower than normal for HV testing.
(3) The avoid areas of the HV diagram are separated by the takeoff corridor. This corridor should be wide enough to consistently permit a takeoff flight path clear of the HV diagram using normal pilot skill. The takeoff corridor should always permit a minimum of +5 knots clearance from critical portions of the diagram.
(4) The knee of the curve separates the takeoff portion from the cruise portion and is defined as the highest speed point on the low speed portion of the HV envelope.
Altitudes above this point are considered cruise, or "fly-in," points, and these test points require a minimum time delay of 1 second between throttle chop and control actuation (reference § 27.143(d)). Altitudes below the knee represent takeoff profile points. For test points in the takeoff portion, use takeoff power (or a lower power selected by the applicant as an operating procedure) and normal pilot reaction time for corrective control actuation.
(5) Since the HV diagram may represent the limiting capabilities of the rotorcraft, each test point should be approached with caution. The manufacturer's buildup program should be reviewed to determine the amount of conservatism in the HV diagram (if any). It should be remembered that the operational pilot will be operating at or near the HV diagram without the benefit of a buildup program. Buildup testing is necessary, and it is most important to vary only one parameter at a time to prevent surprises. Light weight testing is ordinarily conducted first. High and low hover points Page B - 37
AC 27-1B, Chg 4 3/31/2014 are approached from above and below respectively. Portions near the knee are initially evaluated at high speed with subsequent backing down of the speed. In most rotorcraft the effective flare airspeed is critical. At airspeeds slightly below this value, the ability to arrest and control descent rates through use of an aft cyclic flare may be greatly diminished. Extreme care should be exercised when "backing down" to lower speeds.
(6) In addition to the on-board and ground instrumentation, a motion picture camera or other position measuring equipment should cover each run.
(7) For FAA/AUTHORITY tests, the minimum required crew and the minimum instrument panel display presented for certification should be used. Ground safety equipment should be provided.
(8) This test is the least predictable of all the performance items. Therefore, the expansion and extrapolation of test data are questionable. Weight may not be extrapolated to higher values. In order to extrapolate HV data to higher altitudes, any analytical method must have FAA/AUTHORITY approval. In lieu of pure analytical methods, simulations have been used successfully, especially for multiengine rotorcraft.
In either case, the maximum allowable extrapolation should be limited to 2,000 feet density altitude (HD). HV test weights for normal category rotorcraft are the maximum weight at sea level and some lesser weight at high density altitudes. The high density altitude HV curve needs to be defined only to 7,000 feet and may be a lower altitude if the rotorcraft does not have the performance capabilities to attain 7,000 feet. A weight less than the maximum weight may be used to define the high density altitude HV curve, but this weight should not be less than the maximum weight that will allow hovering out of ground effect (OGE). For a given diagram, typical weight reductions that are necessary as altitude is increased can be conservatively estimated by maintaining a constant gross weight divided by density ratio, GW/o (see Figure AC 27.79-2, Part A). If weight is not varied, an enlarged HV diagram is required for safe power-off landing as density altitude is increased (see Figure AC 27.79-2, Part B).
Another method of presentation is to show varying weights at a constant density altitude (see Figure AC 27.79-2, Part C.)
(9) The FAA accepts, as a method of extrapolation, a weight penalty of 3% for each 1000 feet above the permitted 2000 feet extrapolation. This weight penalty has been applied to extrapolations along a constant gross weight divided by density ratio, GWIG.
(10) Vertical takeoff and landing (VTOL) testing normally does not require separate HV testing. The takeoff and landing tests take on the combined characteristics of takeoff, landing, and HV tests.
b. Procedures.
(1) Instrumentation.
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(i) Ground Station. The ground station must have equipment and instrumentation to determine wind direction and velocity, outside air temperature, and if the test rotorcraft has reciprocating engines, humidity. Since the tests must be conducted in winds of 2 knots or less, a smoke generator is highly recommended to show both flightcrew and ground crew personnel the wind direction and velocity at any given time. Additionally, the location of the ground station should be such that it is free of rotor downwash at all times. Motion picture or phototheodolite and radio equipment will be necessary to properly conduct the test program. The use of telemetry equipment is desirable if the location of the test site and the magnitude of the test program make it practical.
(ii) Airborne Equipment (Test Rotorcraft). Necessary installed test equipment may include photopanels or recorders for recording engine parameters, control positions, landing gear loads, landing gear deflections, airspeed, altitude, and other variables. An external light attached to the rotorcraft (or any other means of identifying the engine failure point to the ground camera or phototheodolite) is needed to identify the exact time of engine failure and may also be used to synchronize the ground recorder with the airborne recorded data.
(2) Analytical Prediction. The HV diagram can be estimated by analytical means and this is recommended prior to test. HV, however, is the least predictable of all rotorcraft performance and because of this, the expansion and extrapolation of test data must be done with great care. Test weight may not be extrapolated. All test points should be approached conservatively with some speed or altitude margin. If the applicant has conducted a comprehensive HV flight test program to validate his analytical predictions, much preliminary testing can be eliminated. In any case, the maximum allowable extrapolation from flight test conditions is 2,000 feet density altitude, and an approved analytical or simulation method must be utilized for extrapolation.
(3) Power.
(i) The appropriate power level before engine failure for the low and high hover points is simply the power required to hover at the prevailing hover conditions. The appropriate power condition prior to failure of the engine for points below the knee is takeoff power or a lower value if approved as an operating procedure. For cruise or "flyin" points above the knee, the appropriate condition is power required for level flight.
(il) The applicable power failure conditions are listed in § 29.79(b). Power should be completely cut for normal category rotorcraft. For multiengine rotorcraft that comply with the category A engine isolation design requirements, the HV envelope may be determined with OE and the desired topping power (for the remaining engine(s)) should be set prior to the test. This power value will need adjustment as ambient conditions change. The power can be takeoff power (TOP), 2 ½-minute power, or some calculated lower power for simulating hot day or higher density altitude conditions.
Power is verified and recorded by the pilot by "topping" the engine(s) prior to engine Page B - 39