AC 25-7D CHG 1
3.1.10.2 Tailwind Velocities Greater than 10 Knots, up to 15 knots.
3.1.10.2.1 Performance.
It is considered that takeoff, rejected takeoff, and landing distances, measured in tailwind conditions greater than 10 knots, are unreliable for use in determining airplane performance. Wind conditions of such magnitude are generally not sufficiently consistent over the length of the runway or over the time period required to perform the test maneuver. The 150 percent operational tailwind factor, required by §§ 25.105(d)(1) and
25.125(f), provides a satisfactory level of safety for operation in tailwinds up to 15 knots when using AFM data based on flight tests in nominally calm wind conditions. Tailwind approvals for greater than 15 knots should include dedicated takeoff and landing testing in elevated tailwind conditions to define AFM performance models.
Note: The design requirements of § 25.479, Level landing conditions, also require the effects of increased contact speeds to be investigated if approval for landings with tailwinds greater than 10 knots is desired.
3.1.10.2.2 Control Characteristics.
The test tailwind velocity for demonstrating handling qualities should be equal to the proposed limit tailwind increased by 5 knots. The intent of the 5-knot increase is to provide adequate margin for wind variability in operations, including currency of the wind data, averaging of the data by the measuring and reporting method, and the highly variable nature of higher wind conditions. Therefore, the test wind condition of 5 knots higher than the proposed tailwind limit should be an averaged or smoothed wind speed, not a peak wind speed. (Refer to Section 9.5.2.3 – Guidance to § 25.237 for explanation on averaged crosswind speed, to be adapted to tailwind.) Airplane control characteristics should be evaluated under the following conditions with the CG at the aft limit and the test mean tailwind velocity equal to the proposed limit tailwind increased by 5 knots:
1. Takeoff. Both all-engines-operating and one-engine-inoperative
(i.e., with a simulated failure of the critical engine at the engine failure speed) takeoffs should be evaluated at a light weight with maximum approved takeoff flap deflection.
2. Landing. Approach and landing at light weight with maximum approved landing flap deflection.
3. Determination of the increased ground speed effect on gear vibration or shimmy, and flight director, or autopilot instrument landing system (ILS) approaches, terrain awareness warning system (TAWS) sink rate modes, etc.
4. Deviations above the glideslope should be shown as recoverable.
In particular, it should be shown that the approach speed can be maintained, in tailwind conditions, on a glide path that is 1° steeper 3-13
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than a typical 3° glideslope. This can be shown by either flight test or analysis. Whatever method is used for glideslope recovery compliance, the actual tailwind (i.e., without correction to 10 meters height) need not be higher than the proposed tailwind limit increased by 5 knots.
3.1.10.2.3 Weight Limits.
Consistent with the requirements of §§ 25.105(d)(1) and 25.125(f), the maximum takeoff and maximum quick turnaround weights should be determined using brake energies and tire speeds, as appropriate, calculated with the limit tailwind velocity factored by 150 percent.
3.1.10.2.4 Engine Operating Characteristics.
Satisfactory engine operation should be demonstrated at the limit tailwind velocity factored by 150 percent. The demonstrations should include:
1. Zero groundspeed operation.
2. Takeoff power or thrust setting procedure used for AFM performance (typically completed by approximately 80 knots), both manually and automatically (autothrottle).
3. Reverse thrust operations.
3.1.10.2.5 Airplane Flight Manual.
The AFM should contain a statement that the limitation for tailwinds greater than 10 knots reflects the capability of the airplane as evaluated in terms of airworthiness but does not constitute approval for operation in tailwinds exceeding 10 knots.
3.1.11 Procedures.
3.1.11.1 The performance-related flight test procedures are discussed in the following paragraphs of this AC:
Paragraph 4.2, Takeoff and Takeoff Speeds—§§ 25.105 and 25.107.
Paragraph 4.3, Accelerate-Stop Distance—§ 25.109.
Paragraph 4.4, Takeoff Path—§ 25.111.
Paragraph 4.5, Takeoff Distance and Takeoff Run—§ 25.113.
Paragraph 4.6, Takeoff Flight Path—§ 25.115.
Paragraph 4.7, Climb: General—§ 25.117.
Paragraph 4.8, Landing Climb: All-Engines-Operating—§ 25.119.
Paragraph 4.9, Climb: One-Engine-Inoperative—§ 25.121.
Paragraph 4.10, En Route Flight Paths—§ 25.123.
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Paragraph 4.11, Landing—§ 25.125.
3.1.11.2 Performance Data for Multiple Flap or Additional Flap Positions.
If approval of performance data is requested for flap settings at which no test data are available, the data may be obtained from interpolation of flight data obtained at no less than four flap settings that are within a constant configuration of other lift devices. If the span of flap settings is small and previously obtained data provide sufficient confidence (i.e., the shapes of the curves are known and lend themselves to accurate interpolation), data from three flap settings may be acceptable.
3.1.11.3 Flight Characteristics for Abnormal Configurations.
See § 25.671(c).
3.1.11.3.1 For purposes of this AC, an abnormal configuration is an operational configuration that results from any single failure, or any combination of failures not shown to be improbable.
3.1.11.3.2 Flight characteristics for abnormal configurations may be determined by test or analysis to assure that the airplane is capable of continued safe flight and landing. Flight tests, if required, should be conducted at the critical conditions of altitude, weight, CG, and engine power or thrust associated with the configuration, and at the most critical airspeed between the speed reached one second after stall warning occurs (see paragraph 8.1.5.2.8 of this AC) and the maximum operating airspeed for the configuration.
3.2 Load Distribution Limits—§ 25.23. [Reserved]
3.3 Weight Limits and Center of Gravity Limits—§§ 25.25 and 25.27. [Reserved]
3.4 Empty Weight and Corresponding Center of Gravity—§ 25.29. [Reserved]
3.5 Removable Ballast—§ 25.31.
3.5.1 Explanation.
None.
3.5.2 Procedures.
Ballast may be carried during the flight tests whenever it is necessary to achieve a specific weight and CG location. Consideration should be given to the vertical as well as horizontal location of the ballast in cases where it may have an appreciable effect on 3-15
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the flying qualities of the airplane. The strength of the supporting structures should be considered in order to make sure they do not fail as a result of the anticipated loads that may be imposed during the particular tests. As required by § 21.35(a), applicants must show that these structures comply with the applicable structural requirements of part 25 before conducting flight tests with these structures in place.
3.6 Propeller Speed and Pitch Limits—§ 25.33.
3.6.1 Explanation.
None.
3.6.2 Procedures.
The tachometers and the airspeed indicating system of the test airplane should have been calibrated within the last six months. With that prerequisite satisfied, the following should be accomplished:
3.6.2.1 Determine that the propeller speeds and pitch settings are safe and satisfactory during all tests that are conducted in the flight test program within the certification limits of the airplane, engine, and propeller. This includes establishing acceptable low pitch (flight idle) blade angles on turbopropeller airplanes and verifying that propeller configurations are satisfactory at VMO/MMO to prevent propeller overspeed.
3.6.2.2 Determine that the propeller speeds and pitch settings are safe and satisfactory during all tests that are conducted to satisfy the performance requirements.
3.6.2.3 With the propeller governors operative and the propeller controls in full high revolutions per minute (RPM) position, determine that the maximum takeoff power settings do not exceed the rated takeoff RPM of each engine during takeoff and climb at the best rate-of-climb speed.
3.6.2.4 With the propeller governors made inoperative by mechanical means, determine the maximum power, no-wind, static RPMs. With the propeller governors operating on the low pitch stop, the engine speeds must not exceed 103 percent of the maximum allowable takeoff RPM or 99 percent of an approved maximum overspeed, as required by § 25.33(c). On turbopropeller engines, the engine speeds should not exceed the maximum engine speeds allowed by engine and propeller type designs. Note which systems were disabled and how the disablement was done. If maximum takeoff power torque or sea level standard conditions cannot be obtained on the test day, correct the data to these conditions by an acceptable means. A no-wind condition is considered to be a wind of 5 knots or less.
The static RPM should be the average obtained with a direct crosswind from the left and a direct crosswind from the right.
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3.6.2.5 If the above determinations are satisfactory, then measure the low-pitch stop setting and the high-pitch stop setting. These data may have been obtained from the propeller manufacturer and may be used, provided the pitch stops have not been changed since the manufacturer delivered the propeller. If measured, the blade station should be recorded. Include these blade angles in the type certificate data sheet.
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CHAPTER
4. FLIGHT: PERFORMANCE
4.1 General—§ 25.101.
4.1.1 Explanation of Propulsion System Behavior.
Section 25.101(c) requires that airplane performance “correspond to the propulsive thrust available under the particular ambient atmospheric conditions, the particular flight conditions….” The propulsion systems (i.e., turbine engines and propellers, where appropriate) installed performance characteristics are primarily a function of engine power or thrust setting, airspeed, propeller efficiency (where applicable), altitude, and ambient temperature. Determine the effects of each of these variables to establish the thrust available for airplane performance calculations.
4.1.2 Procedures.
4.1.2.1 The intent is to develop a model of propulsion system performance that covers the approved flight envelope. Further, it should be shown that the combination of the propulsion system performance model and the airplane performance model is validated by the takeoff performance test data, climb performance tests, and tests used to determine airplane drag.
Installed propulsion system performance characteristics may be established via the following tests and analyses:
Steady-state engine power (or thrust) setting versus power (or thrust) testing. See paragraph 4.1.2.2.
Lapse rate takeoff testing to characterize the behavior of power or thrust setting, rotor speeds, propeller effects (i.e., torque, RPM, and blade angle), or gas temperature as a function of time, thermal state, or airspeed, as appropriate. See paragraph 4.1.2.3.
Power/thrust calculation substantiation. See paragraph 4.1.2.4.
Effects of ambient temperature. See paragraph 4.1.2.5.
4.1.2.2 Steady-State Engine Power (or Thrust) Setting versus Power (or Thrust) Testing.
Engines should be equipped with adequate instrumentation to allow the determination of thrust (or power). Data should be acquired to validate the model, including propeller-installed thrust, if applicable, over the range of power or thrust settings, altitudes, temperatures, and airspeeds for which approval is sought. Although it is not possible to definitively list or foresee all of the types of instrumentation that might be considered adequate for determining thrust (or power) output, two examples used in past certification programs are (1) engine pressure rakes, with engines calibrated in a ground test cell, and (2) fan speed, with engines calibrated in a ground test cell and the calibration data validated by the use of a flying test bed. In any case, the applicant should substantiate the adequacy 4-1
of the instrumentation to be used for determining the thrust (or power) output.
4.1.2.3 Lapse Rate Takeoff Testing to Characterize the Behavior of Power or Thrust Setting, Rotor Speeds, Propeller Effects, or Gas Temperature as a Function of Time, Thermal State, or Airspeed.
These tests should include the operation of an automatic takeoff thrust control system (ATTCS), if applicable, and should cover the range of power or thrust settings for which approval is sought.
4.1.2.3.1 Data for higher altitude power or thrust settings may be acquired via overboost (i.e., operating at a higher-than-normal power or thrust setting for the conditions) with the consent of the engine and propeller manufacturer(s), when applicable. When considering the use of overboost on turbopropeller propulsion system installations to stimulate higher altitude and ambient temperature range conditions, the capability to achieve an appropriate simulation should be evaluated based on the engine and propeller control system(s) and aircraft performance and structural considerations. Engine (gearbox) torque, rotor speed, or gas temperature limits, including protection devices to prohibit or limit exceedances, may prevent the required amount of overboost needed for performance at the maximum airport altitude sought for approval. Overboost may be considered as increased torque, reduced propeller speed, or a combination of both, in order to achieve the appropriate blade angle for the higher altitude and ambient temperature range simulation. Consideration for extrapolations will depend on the applicant’s substantiation of the proper turbopropeller propulsion system simulated test conditions.
4.1.2.3.2 Lapse rate characteristics should be validated by takeoff demonstrations at the maximum airport altitude for which takeoff approval is being sought.
Alternatively, if overboost (see paragraph above) is used to simulate the power or thrust setting parameters of the maximum airport altitude for which takeoff approval is sought, the takeoff demonstrations of lapse rate characteristics can be performed at an airport altitude up to 3,000 feet lower than the maximum airport altitude.
4.1.2.4 Power/Thrust Calculation Substantiation.
Installed power or thrust should be calculated via a mathematical model of the propulsion system, or other appropriate means, adjusted as necessary to match the measured inflight performance characteristics of the installed propulsion system. The propulsion system mathematical model should define the relationship of power or thrust to the power or thrust setting parameter over the range of power or thrust settings, airspeeds, altitudes, and temperatures for which approval is sought. For turbojet airplanes, the propulsion system mathematical model should be substantiated by ground tests in which thrust is directly measured via a calibrated load cell or equivalent means. For turbopropeller airplanes, the engine power 4-2
measurement should be substantiated by a calibrated dynamometer or equivalent means, the engine jet thrust should be established by an acceptable engine model, and the propeller thrust and power characteristics should be substantiated by wind tunnel testing or equivalent means.
4.1.2.5 Effects of Ambient Temperature.
The flight tests of paragraph 4.1.2.2 of this AC will typically provide data over a broad range of ambient temperatures. Additional data may be obtained from other flight or ground tests of the same type or series of engine. The objective is to confirm that the propulsion system model accurately reflects the effect of temperature over the range of ambient temperatures for which approval is being sought (operating envelope).
Because thrust (or power) data can usually be normalized versus temperature using either dimensionless variables (e.g., theta exponents or a thermodynamic cycle model), it is usually unnecessary to obtain data over the entire ambient temperature range. There is no needed to conduct additional testing if:
The data show that the behavior of power or thrust and limiting parameters versus ambient temperature can be predicted accurately, and Analysis based upon the test data shows that the propulsion system will operate at rated power or thrust without exceeding propulsion system limits.
4.1.2.6 Extrapolation of propulsion system performance data to 3,000 feet above the highest airport altitude tested (but no higher than the maximum takeoff airport altitude to be approved) is acceptable, provided the supporting data, including flight test and propulsion system operations data (e.g., engine and propeller control, limits exceedance, and surge protection devices scheduling), substantiates the proposed extrapolation procedures.
Considerations for extrapolation depend upon an applicant’s determination, understanding, and substantiation of the critical operating modes of the propulsion system. This understanding includes a determination and quantification of the effects that propulsion system installation and variations in ambient conditions have on these modes.
4.2 Takeoff and Takeoff Speeds—§§ 25.105 and 25.107.
4.2.1 Explanation.
Section 25.105 specifies the conditions that must be considered in determining the takeoff speeds, accelerate-stop distances, takeoff path, takeoff distance, and takeoff run in accordance with part 25 requirements. The primary objective of the takeoff tests required by § 25.107 is to determine the takeoff speeds for all takeoff configurations at 4-3
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all weight, altitude, and temperature conditions within the operational limits selected by the applicant.
4.2.2 Background Information.
This information is not related to means of compliance for § 25.105 and § 25.107 but is included to increase awareness of the risks associated with in-ground-effect stall during execution of the required flight testing. The information is based upon previous industry experience and recommended best practices, much of which resulted from investigations following a flight test accident during takeoff performance testing. It is recommended that the following considerations be included in the applicant’s flight test hazard assessment in preparation for the takeoff performance flight testing.
Ground Effect Considerations. It is important to understand the aerodynamic characteristics of the airplane in ground effect to avoid an inadvertent stall while operating at high angles of attack in close proximity to the ground. Ground effect is the phenomenon that modifies a body’s aerodynamic characteristics when it is generating lift close to the ground.
The proximity of the ground suppresses downwash and wing tip vortices and affects the flow between the airplane and the ground, causing a slowing of the airstream under the wing, producing a positive pressure or buoyancy field.
The impact of ground effect is largely dependent on the height of the wing above the ground and the magnitude of the resultant pressure field that exists between the underside of the wing and the ground plane. When the distance between the wing and ground is small (typically 5 to 10 percent of wingspan), the pressure field and extent of upwash/downwash suppression is large. When a significant blockage effect is experienced, the excess flow that cannot pass under the wing traverses the wing upper surface causing increased suction on the wing leading edge and increased lift at low angles of attack; this effect grows with larger and more powerful trailing edge flaps and their proximity to ground at liftoff attitude. The increased leading-edge suction tends to steepen the adverse pressure gradient behind the suction peak that can promote early aerodynamic flow separation, especially if the wing is already characterized by a leading-edge stall. It is especially important to note that when this effect is large, it will not only reduce the stall angle-of-attack, but it can also consequently decrease the maximum lift of the airplane in ground effect below the value observed in free air.
It is also important to understand the effect that the Mach number has on the stall angle-of-attack in the range expected for takeoff. Experience has shown that reductions in the stall angle-of-attack with an increasing Mach number, observed in free air conditions, can also be considered applicable in ground effect.
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Due to remaining uncertainties in the in-ground-effect stall angle-ofattack, the test pilots involved in the takeoff testing should be familiar with the free air stall characteristics of the airplane, any predicted changes in the stall characteristics in ground-effect, and have established stall recovery techniques (e.g., reduction in angle-of-attack, advancing the throttles to increase thrust) in the event that a stall is encountered, including allowance for stall angle-of-attack hysteresis to re-attach the airflow in the recovery.
4.2.2.1 Ground Effect Estimation Methods: Industry experience has shown that a reduction in stall angle-of-attack in ground effect relative to free air of 4 to 5 degrees for takeoff flap deflections is a reasonable estimate in the absence of more configuration-specific data, though due to configuration differences there is no assurance that this estimate will be conservative in all cases. Wind tunnel results obtained at lower Reynolds numbers than full-scale flight should be used with extreme caution when predicting ground effect. While the “linear range” effects of such sub-scale testing can be accurate, the impact of ground effect on stall may not be captured at low Reynolds numbers. Computational fluid dynamics (CFD) results have been shown to produce useful indications of in-ground-effect stall, but the undertaking of predicting stall angles with CFD is, in general, not trivial. Before using wind tunnel or CFD predicted in-ground-effect stall angles, comparisons should be made with results from free air stall flight testing to understand and/or improve the accuracy of the predictions.
Caution should be exercised to ensure that conservative margins are identified for flight test use.
4.2.3 Procedures: General.
4.2.3.1 Section 25.105(c)(1) requires the takeoff performance data to be determined for smooth, dry and wet, hard-surfaced runways.
Paragraph 4.3 of this AC describes methods for determining the accelerate-stop distances required by § 25.109. Paragraph 4.5 describes methods for determining the takeoff distance and takeoff run required by § 25.113.
4.2.3.2 In accordance with § 25.101(f), testing for determining the accelerate-stop distances, takeoff flight paths, and takeoff distances should be accomplished using procedures established by the applicant for operation in service. In accordance with §25.101(h), these procedures must be able to be consistently executed in service by crews of average skill, use methods or devices that are safe and reliable, and include allowances for any time delays in the execution of the procedures that may reasonably be expected in service. These requirements prohibit the use of exceptional piloting techniques, such as higher control force inputs or higher pitch 4-5
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rates than would occur in operational service, from being used to generate unrealistic takeoff distances. The intent of these requirements is to establish takeoff performance representative of that which can reasonably be expected to be achieved in operational service.
4.2.3.3 Attention should be paid to all potential sources of airspeed error, but special consideration should be given to airplanes with electronic instruments in the cockpit that apply electronic filtering to the airspeed data. This filtering, which causes a time delay in the airspeed indication, can be a source of significant systematic error in the presentation of airspeed to the flightcrew. With normal takeoff acceleration, the airplane will be at a higher speed than is indicated by the cockpit instrument, which can result in longer distances than are presented in the AFM, particularly in the event of a rejected takeoff near the indicated V1 speed. The effects of any time delays caused by electronic filtering, pneumatic system lag, or other sources should be adequately addressed in the AFM speed and distance presentations. Further explanation of airspeed lag, particularly pertaining to airplanes with electronic instruments in the cockpit, and procedures for calibrating the airspeed indicating system (§ 25.1323(b)) are presented in paragraph 33.3 of this AC.
4.2.4 Procedures: Section 25.107(a)(1)—Engine Failure Speed (VEF).
The engine failure speed (VEF) is defined as the calibrated airspeed at which the critical engine is assumed to fail and must be selected by the applicant. VEF cannot be less than the ground minimum control speed (VMCG).
4.2.5 Procedures: Section 25.107(a)(2)—V1.
V1 may not be less than VEF plus the speed gained with the critical engine inoperative during the time interval between VEF and the instant at which the pilot takes action after recognizing the engine failure. This is indicated by pilot application of the first deceleration means such as brakes, throttles, spoilers, etc. during accelerate-stop tests.
The applicant may choose the sequence of events. Refer to paragraph 4.3 of this AC, addressing § 25.109, for a more complete description of rejected takeoff (RTO) transition procedures and associated time delays.
4.2.6 Procedures: Section 25.107(b)—Minimum Takeoff Safety Speed (V2MIN).
4.2.6.1 V2MIN, in terms of calibrated airspeed, cannot be less than:
4.2.6.1.1 1.1 times the VMC defined in § 25.149.
4.2.6.1.2 1.13 times VSR for two-engine and three-engine turbopropeller and reciprocating engine-powered airplanes and for all turbojet airplanes that do not have provisions for obtaining a significant reduction in the one-engine-inoperative power-on stalling speed (i.e., boundary layer 4-6
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control, blown flaps, etc.). The value of VSR to be used in determining V2MIN is the free air reference stall speed in the applicable takeoff configuration, landing gear retracted, except for those airplanes with a fixed landing gear or for gear-down dispatch.
4.2.6.2 V2MIN may be reduced to 1.08 times VSR for turbopropeller and reciprocating engine-powered airplanes with more than three engines, and turbojet powered airplanes with adequate provisions for obtaining significant power-on reference stall speed reduction through the use of such things as boundary layer control, blown flaps, etc.
4.2.6.3 For propeller-driven airplanes, the difference between the two margins, based upon the number of engines installed on the airplane, is because the application of power ordinarily reduces the stalling speed appreciably. In the case of the two-engine propeller-driven airplane, at least half of this reduction is eliminated by the failure of an engine. The difference in the required factors therefore provides approximately the same margin over the actual stalling speed under the power-on conditions that are obtained after the loss of an engine, no matter what the number of engines (in excess of one) may be. Unlike the propeller-driven airplane, the turbojet/turbofan powered airplane does not show any appreciable difference between the power-on and power-off stalling speed. This is due to the absence of the propeller, which ordinarily induces a slipstream with the application of power causing the wing to retain its lift to a speed lower than the power-off stalling speed. The applicant’s selection of the two speeds specified will influence the nature of the testing required in establishing the takeoff flight path.
4.2.7 Procedures: Section 25.107(c)—Takeoff Safety Speed (V2).
4.2.7.1 V2 is the calibrated airspeed that is attained at or before the airplane reaches a height of 35 feet above the takeoff surface after an engine failure at VEF using an established rotation speed (VR). From the liftoff point, the takeoff surface extends to the end of the takeoff distance continuing at the same slope as the runway. During the takeoff speeds demonstration, V2 should be continued to an altitude sufficient to assure stable conditions beyond the 35-foot height. V2 cannot be less than V2MIN. In addition, V2 cannot be less than the liftoff speed, VLOF, which is defined in § 25.107(f).
In accordance with § 25.107(c), V2, in terms of calibrated airspeed, may not be less than VR plus the speed increment attained before reaching a height of 35 feet above the takeoff surface using a takeoff maneuver that can be executed consistently by crews of average skill per the requirement of § 25.101(h)(1), and a speed that provides the maneuvering capability specified in § 25.143(h). Due to the constraints on VR and VLOF specified in § 25.107(e), and accounting for other constraints required for safe 4-7