3-15 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 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.
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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.
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.
4-1 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 system’s (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 in order 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
4-2 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
4-3 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
4-4 all weight, altitude, and temperature conditions within the operational limits selected by the applicant.
4.2.2 Procedures: General.
4.2.2.1 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.
4.2.2.2 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 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.
4.2.2.3 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.
4.2.3 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).
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4.2.4 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.
4.2.5 Procedures: Section 25.107(b)—Minimum Takeoff Safety Speed (V2MIN).
4.2.5.1 V2MIN, in terms of calibrated airspeed, cannot be less than:
4.2.5.1.1 1.1 times the VMC defined in § 25.149.
4.2.5.1.2 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 control, blown flaps, etc.). The value of VSR to be used in determining V2MIN is the stall speed in the applicable takeoff configuration, landing gear retracted, except for those airplanes with a fixed landing gear or for gear-down dispatch.
4.2.5.2 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 stall speed reduction through the use of such things as boundary layer control, blown flaps, etc.
4.2.5.3 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.
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4.2.6 Procedures: Section 25.107(c)—Takeoff Safety Speed (V2).
4.2.6.1 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 and a speed that provides the maneuvering capability specified in § 25.143(h). In addition, § 25.111(c)(2) stipulates that the airplane must reach V2 before it is 35 feet above the takeoff surface and continue at a speed not less than V2 until it is 400 feet above the takeoff surface. These requirements were first expressed in Special Civil Air Regulation No. SR-422, Turbine-Powered Transport Category Airplanes of Current Design (SR-422A), paragraphs 4T.114(b)(4) and (c)(3) and 4T.116(e). The concern that the regulation change was addressing was the overshoot of V2 after liftoff under the previous requirement that the airplane attain V2 on, or near, the ground. The intent of the current requirement is to allow an acceleration to V2 after liftoff but not to allow a decrease in the field length required to attain a height of 35 feet above the takeoff surface by attaining a speed greater than V2, under low drag ground conditions, and using the excess kinetic energy to attain the 35-foot height.
4.2.6.2 In the case of turbojet powered airplanes, when most of the one-engine-inoperative data have been collected using throttle chops, V2, and its relationship to VR, should be substantiated by at least a limited number of fuel cuts at VEF. For derivative programs not involving a modification that would affect thrust decay characteristics, demonstrations of fuel cuts may be unnecessary.
4.2.6.3 For propeller-driven airplanes, the use of fuel cuts can be more important in order to ensure that the takeoff speeds and distances are obtained with the critical engine’s propeller attaining the position it would during a sudden engine failure. The number of tests that should be conducted using fuel cuts depends on the correlation obtained with the throttle chop data and substantiation that the data analysis methodology adequately models the effects of a sudden engine failure.
4.2.7 Procedures: Section 25.107(d)—Minimum Unstick Speed (VMU).
4.2.7.1 Section 25.107(d) states, “VMU speeds must be selected by the applicant.”
An applicant can either determine the lowest possible VMU speeds or select
4-7 a higher speed that supports the takeoff performance targets of the airplane. Regardless of how the applicant selects the VMU speeds, compliance must be shown with § 25.107(d), (e)(1)(iv), (e)(3), and (e)(4) to show that the selected VMU speeds allow the airplane to safely lift off the ground and continue the takeoff.
4.2.7.2 An applicant should comply with § 25.107(d) by conducting VMU tests with all engines operating and also with one engine inoperative. During these tests, the takeoff should be continued until the airplane is out of ground effect. The airplane pitch attitude should not be decreased after liftoff.
4.2.7.3 VMU testing to demonstrate the lowest VMU speed is a maximum performance flight test maneuver, and liftoff may occur very near the angle-of-attack for maximum lift coefficient. Also, even though pitch attitude may be held fairly constant during the maneuver, environmental conditions and transiting through ground effect may result in changes in angle-of-attack. It is permissible to lift off at a speed that is below the normal stall warning speed, provided no more than light buffet is encountered.
4.2.7.3.1 It is important for the flight test team to understand the control laws and any transitions between control laws during takeoff (e.g., based on weight on wheels) for an electronic flight control system that may present unique hazards that should be taken into account.
4.2.7.3.2 An artificial stall warning system (e.g., a stick shaker) may be disabled during VMU testing, although doing so will require extreme caution and depend upon a thorough knowledge of the airplane’s stall characteristics, both in and out of ground effect.
4.2.7.3.3 If the airplane is equipped with a stick pusher, angle-of-attack limiter, or other system that may affect the conduct of the test, the angle of attack setting for activation of the system may be selected by the applicant and differ from the nominal setting. The system may alternatively be disabled or its activation delayed for test purposes until a safe altitude is reached.
However, for airplanes equipped with a stick pusher that is not designed to be inhibited during takeoff, the VMU test demonstrations will need to be assessed and will only remain valid if the stick pusher would not have activated with the angle-of-attack indication means set at the lowest angle within production tolerances.
4.2.7.4 In lieu of conducting one-engine-inoperative VMU tests, the applicant may conduct all-engines-operating VMU tests if all pertinent factors that would be associated with an actual one-engine-inoperative VMU test are simulated or otherwise taken into account. To take into account all pertinent factors,
4-8 it may be necessary to adjust the resulting VMU test values analytically.
The factors to be accounted for should include at least the following:
Thrust/weight ratio for the one-engine-inoperative range.
Controllability (may be related to one-engine-inoperative free air tests, such as minimum control speed in the air (VMCA)).
Increased drag due to use of lateral/directional control systems.
Reduced lift due to use of devices such as wing spoilers for lateral control.
Adverse effects of use of any other systems or devices on control, drag, or lift.
4.2.7.5 The number of VMU tests needed may be minimized by testing only the critical all-engines-operating and one-engine-inoperative thrust/weight ratios, provided the VMU speeds determined at these critical conditions are used for the range of thrust/weights appropriate to the all-engines-operating and one-engine-inoperative configurations. The critical thrust/weight is established by correcting, to the VMU speed, the thrust that results in the airplane achieving its limiting one-engine-inoperative climb gradient at the normally scheduled speed and in the appropriate configuration.
4.2.7.6 Amendment 25-42, effective March 1, 1978, revised §§ 25.107(d) and
25.107(e)(1)(iv) in order to permit the one-engine-inoperative VMU to be determined by all-engines-operating tests at the thrust/weight ratio corresponding to the one-engine-inoperative condition. As revised, § 25.107(d) specifies that VMU must be selected for the range of thrust/weight ratios to be certificated, rather than for the all-enginesoperating and one-engine-inoperative conditions as was previously required. In determining the all-engines-operating thrust/weight ratio that corresponds to the one-engine-inoperative condition, consider trim and control drag differences between the two configurations in addition to the effect of the number of engines operating. The minimum thrust/weight ratio to be certificated is established by correcting, to the VMU speed, the thrust that results in the airplane achieving its limiting engine-out climb gradient in the appropriate configuration and at the normally scheduled speed.
4.2.7.7 To conduct the VMU tests, rotate the airplane as necessary to achieve the VMU attitude. It is acceptable to use some additional nose-up trim over the normal trim setting during VMU demonstrations. If additional nose-up trim is required, the additional considerations of paragraph 4.2.7.8 below apply.
VMU is the speed at which the weight of the airplane is completely supported by aerodynamic lift and thrust forces. Some judgment may be necessary on airplanes that have tilting main landing gear bogies.
4-9 Determining the liftoff point from gear loads and wheel speeds has been found acceptable in past programs. After liftoff, the airplane should be flown out of ground effect. During liftoff and the subsequent climbout, the airplane should be fully controllable.
4.2.7.8 VMU Testing for Airplanes having Limited Pitch Control Authority.
4.2.7.8.1 For some airplanes with limited pitch control authority, it may not be possible, at forward CG and normal trim, to rotate the airplane to a liftoff attitude where the airplane could otherwise perform a clean flyaway at a minimum speed had the required attitude been achieved. This may occur only over a portion of the takeoff weight range in some configurations.
Though generally associated with the inability of the pitch control surfaces to provide adequate pitching moment to rotate the airplane to the desired pitch attitude at low thrust/weight ratio conditions, the same phenomenon may occur at high thrust/weight ratio conditions for airplanes with high thrust lines (e.g., aft engines mounted high on the fuselage). When limited pitch control authority is clearly shown to be the case, VMU test conditions may be modified to allow testing aft of the forward CG limit and/or with use of more airplane nose-up trim than normal. The VMU data determined with this procedure should be corrected to those values representative of the appropriate forward limit; the variation of VMU with CG may be assumed to be like the variation of free air stalling speed with CG.
Although the development of scheduled takeoff speeds may proceed from these corrected VMU data, additional tests are required (see paragraph 4.2.7.8.2 below) to check that the relaxed VMU criteria have not neglected problems that might arise from operational variations in rotating airplanes with limited pitch control authority.
4.2.7.8.2 In the following assurance test, the airplane should demonstrate safe flyaway characteristics: Minimum speed liftoff should be demonstrated at the critical forward CG limit with normal trim. For airplanes with a cutback forward CG at heavy weight, two weight/CG conditions should be considered. The heavy weight tests should be conducted at maximum structural or maximum sea level climb-limited weight with the associated forward CG. The full forward CG tests should be conducted at the highest associated weight. Alternatively, testing may be conducted at a single weight if an analysis is provided that identifies the critical weight/CG combination with regard to limited pitch attitude capability for liftoff.
These assurance tests should be conducted at the thrust/weight ratio that is most critical for attaining a pitch attitude that will provide a minimum liftoff speed.
4.2.7.8.3 For airplanes that are limited by low thrust/weight conditions, tests should be conducted at the minimum thrust/weight ratio for both the simulated one-engine-inoperative test (i.e., symmetrical reduced thrust) case and the all-engines-operating case.
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4.2.7.8.4 For airplanes that are limited by high thrust/weight conditions, tests should be conducted at the highest thrust/weight ratio within the airplane’s operating envelope for both the simulated one-engine-inoperative case
(i.e., symmetrical reduced thrust) and the all-engines-operating case.
4.2.7.8.5 One acceptable test technique is to hold full nose-up control column as the airplane accelerates. As pitch attitude is achieved to establish the minimum liftoff speed, pitch control may be adjusted to prevent over-rotation, but the liftoff attitude should be maintained as the airplane flies off the ground and out of ground effect.
4.2.7.8.6 The resulting liftoff speeds are acceptable if the test proves successful and the liftoff speed is at least 5 knots below the normally scheduled liftoff speed.
4.2.7.8.7 This minimum 5 knot margin from the scheduled liftoff speed provides some leeway for operational variations such as mistrim, CG errors, etc., that could further limit the elevator authority. The reduced VMU margins arising from this test, relative to those specified in § 25.107(e)(1)(iv), are considered acceptable because of the reduced probability of a pitch control authority-limited airplane getting into a high drag condition due to over-rotation.
4.2.7.9 VMU Testing for Geometry Limited Airplanes.
4.2.7.9.1 For airplanes that are geometry limited (i.e., the minimum possible VMU speeds are limited by tail contact with the runway), § 25.107(e)(1)(iv)(B) allows the VMU to VLOF speed margins to be reduced to 108 percent and 104 percent for the all-engines-operating and one-engine-inoperative conditions, respectively. The VMU demonstrated should be sound and repeatable.
4.2.7.9.2 An airplane that is deemed to be geometry limited at the conditions tested is expected to be geometry limited over its entire takeoff operating envelope. If this is not the case, the airplane is not considered geometry limited and the reduced VMU to VLOF speed margins do not apply.
4.2.7.9.3 One acceptable means for demonstrating compliance with §§ 25.107(d) and 25.107(e)(1)(iv) with respect to the capability for a safe liftoff and fly-away from the geometry limited condition is to show that at the lowest thrust-to-weight ratio for the all-engines-operating condition:
1. In the speed range from 96 to 100 percent of the actual liftoff speed), the aft under-surface of the airplane should be in contact with the runway. Because of the dynamic nature of the test, it is recognized that contact will probably not be maintained during this entire speed range, so some judgment is necessary. It has been found acceptable for