13 is adequately compensated for by the 50 percent wind factor. This also eliminates a requirement to adjust V R for wind. However, in calculating the AFM field lengths, the airspeed attained due to wind profile may be included.
Example: Landing--AFM Data
Given:
o Factored wind with airplane 50 ft. above landing surface 21.7 kts.
o Factored wind with airplane on landing surface 16.3 kts.
o Zero wind speed change for 50 ft. to touchdown (V 50-VTD) 4.0 kts.
o True airspeed required at 50 ft. 130.0 kts.
Results:
o Ground speed at 50 ft. 130-21.7 = 108.3 kts.
o Ground speed at touchdown 108.3-4.0 = 104.3 kts.
o True airspeed at touchdown 104.3+16.3 = 120.6 kts.
(8) Expansion of Takeoff and Landing Data for a Range of Airport Elevations.
(a) These guidelines apply to expanding AFM takeoff and landing data above and below the altitude at which the airplane takeoff and landing performance tests are conducted.
(b) Historically, limits were placed on the extrapolation of takeoff data. In the past, takeoff data could generally be extrapolated 6,000 feet above and 3,000 feet below the test field elevation when proven testing and data reduction methods were used. For extrapolations beyond these limits, a 2 percent takeoff distance penalty was to be applied for every additional 1,000 feet extrapolation. Such limitations were generally not applied to extrapolation of landing data, provided the effect of the higher true airspeed on landing distance was taken into account.
(c) Since then, considerably more experience has been gained both in terms of modeling airplane and propulsion system (i.e., turbine engines and propellers, where appropriate) performance and in verifying the accuracy of these models for determining high (and low) altitude takeoff and landing performance. This experience has shown that the soundness of the extrapolation is primarily a function of the accuracy of the propulsion system performance model and its integration with the airplane drag model. The basic aerodynamic characteristics of the airplane do not change significantly with altitude or ambient temperature, and any such effects are readily taken into account by standard airplane performance modeling practices.
(d) As a result, with installed propulsion system performance characteristics that have been adequately defined and verified, airplane takeoff and landing performance data obtained at one field elevation may be extrapolated to higher and lower altitudes within the limits of the operating envelope without applying additional performance conservatisms. It should be noted, however, that extrapolation of the propulsion system data used in the
14 determination and validation of propulsion system performance characteristics is typically limited to 3,000 feet above the highest altitude at which propulsion system parameters were evaluated for the pertinent power/ thrust setting. (See paragraph 9 of this AC for more information on an acceptable means of establishing and verifying installed propulsion system performance characteristics.)
(e) Note that certification testing for operation at airports that are above 8,000 feet should also include functional tests of the cabin pressurization system in accordance with paragraph 87b(3) of this AC. Consideration should be given to any sensitivity to, or dependency upon airport altitude, such as: engine and APU starting, passenger oxygen, autopilot, autoland, autothrottle system thrust set/operation.
(9) Tailwind Takeoff and Landing .
(a) Wind Velocities of 10 Knots or Less - Approval may be given for performance, controllability, and engine operating characteristics for operations in reported tailwind velocities up to 10 knots, measured at a 10 meter height, without specific flight tests.
(b) Wind Velocities Greater than 10 Knots.
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 velocity factor, required by §§ 25.105(d)(1) and 25.125(f), affords a satisfactory method for determination of airplane takeoff and landing performance information and limitations up to a limiting tailwind velocity of 15 knots when using a flight test data base obtained under zero wind conditions.
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.
2 Control Characteristics . Airplane control characteristics should be evaluated under the following conditions with the center of gravity at the aft limit and the test tailwind velocity equal to the proposed limit tailwind factored by 150 percent:
(aa) Takeoff. At light weight with maximum approved takeoff flap deflection, both all-engines operating and one-engine inoperative takeoffs should be evaluated.
(bb) Landing. Approach and landing at light weight with maximum approved landing flap deflection.
(cc) Determination of the increased ground speed effect on gear vibration, shimmy, flight director and/or autopilot instrument landing system (ILS) approaches, ground proximity warning systems (GPWS) sink rate modes, etc.
(dd) If engine idle thrust is increased to account for the increased tailwind velocity, ensure that deviations above the glideslope are recoverable.
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.
4 Engine Operating Characteristics . Satisfactory engine operation should be demonstrated at the limit tailwind velocity factored by 150 percent. The demonstrations should include:
(aa) Zero groundspeed operation.
(bb) Takeoff power setting procedure used for AFM performance (typically completed by approximately 80 knots), both manually and automatically (autothrottle).
(cc) Reverse thrust operations.
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.
b. Procedures.
(1) The performance-related flight test procedures are discussed in each of the following paragraphs of this AC:
10. Takeoff and Takeoff Speeds
11. Accelerate-Stop Distance
12. Takeoff Path
13. Takeoff Distance and Takeoff Run
14. Takeoff Flight Path
15. Climb: General
16. Landing Climb
17. Climb: One Engine Inoperative
18. En Route Flight Path
19. Landing
(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
16 previously obtained data provide sufficient confidence (i.e., the shape of the curves are known and lend themselves to accurate interpolation), data from three flap settings may be acceptable.
(3) Flight Characteristics for Abnormal Configurations (Ref. § 25.671(c)).
(a) 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.
(b) 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, c.g., and engine thrust associated with the configuration, and at the most critical airspeed between the speed reached one second after stall warning occurs (see paragraph 29e(2)(h) of this AC) and the maximum operating airspeed for the configuration.
4. Load Distribution Limits - § 25.23 [Reserved].
5. Weight Limits and Center of Gravity Limits - §§ 25.25 and 25.27 [Reserved].
6. Empty Weight and Corresponding Center of Gravity - § 25.29 [Reserved].
7. Removable Ballast - § 25.31 [Reserved].
8. Propeller Speed and Pitch Limits - § 25.33.
a. Explanation. None.
b. Procedures. The tachometers and the airspeed indicating system of the test airplane should have been calibrated within the last six months. With those conditions satisfied, the following should be accomplished:
(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.
(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) With the propeller governors operative and the propeller controls in full high revolutions per minute (r.p.m.) position, determine that the maximum takeoff power settings do
17 not exceed the rated takeoff r.p.m. of each engine during takeoff and climb at the best rate-ofclimb speed.
(4) With the propeller governors made inoperative by mechanical means, determine the maximum power, no-wind, static r.p.m.’s. With the propeller governors operating on the low pitch stop, the engine speeds must not exceed 103 percent of the maximum allowable takeoff
r.p.m. 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 nowind condition is considered to be a wind of 5 knots or less. The static r.p.m. should be the average obtained with a direct crosswind from the left and a direct crosswind from the right.
(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.
Section 2. Performance
9. General - § 25.101.
a. Explanation . - Propulsion System Behavior. Section 25.101(c) requires that airplane “performance must 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 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.
b. Procedures.
(1) The intent is to develop a model of propulsion system performance that covers the approved flight envelope. Furthermore, 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:
(a) Steady-state engine power setting vs. thrust (or power) 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
18 thrust, if applicable, over the range of power 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 of the instrumentation to be used for determining the thrust (or power) output.
(b) Lapse rate takeoff testing to characterize the behavior of power 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. These tests should include the operation of an automatic takeoff thrust control system (ATTCS), if applicable, and should cover the range of power settings for which approval is sought.
1 Data for higher altitude power settings may be acquired via overboost
(i.e., operating at a higher than normal power 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 simulate 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.
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 1 above) is used to simulate the 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.
(c) Thrust calculation substantiation. Installed 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 thrust to the power setting parameter over the range of power setting, airspeed, altitude, and temperature 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 measurement should be substantiated by a calibrated dynamometer or equivalent means, the engine jet thrust should be
19 established by an acceptable engine model, and the propeller thrust and power characteristics should be substantiated by wind tunnel testing or equivalent means.
(d) Effects of ambient temperature. The flight tests of paragraph 9b(l)(a) above 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 need to conduct additional testing if:
1 The data show that the behavior of thrust and limiting parameters versus ambient temperature can be predicted accurately, and
2 Analysis based upon the test data shows that the propulsion system will operate at rated thrust without exceeding propulsion system limits.
(2) 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.
10. Takeoff and Takeoff Speeds - §§ 25.105 and 25.107.
a. Explanation. The primary objective of the takeoff tests required by § 25.107 is to determine the takeoff speed schedule for all takeoff configurations at all weight, altitude, and temperature conditions within the operational limits selected by the applicant. The provisions of § 25.105 are self evident and are not repeated or amplified in this discussion. Guidance material for unpaved runway evaluation is contained in Chapter 8 (paragraph 232) of this AC.
b. Procedures. Although the following speed definitions are given in terms of calibrated airspeed, the AFM presentations should be given in terms of indicated airspeed. 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. During a 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
20 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 177 of this AC.
(1) 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).
(2) Section 25.107(a)(2) - Takeoff Decision Speed (V1). The takeoff decision speed (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 device such as brakes, throttles, spoilers, etc. during accelerate-stop tests, or by the first control input during VMCG testing. The applicant may choose the sequence of events. Refer to paragraph 11 of this AC, addressing § 25.109, for a more complete description of rejected takeoff (RTO) transition procedures and associated time delays. If it becomes evident in expansion of takeoff data for the AFM that excessive variation in V1 exists, resulting from the many performance variables involved (variations of +1.5 knots or +100 ft. have been found acceptable), then measures should be taken to ensure that scheduled performance variations are not excessive. Examples of such measures are small field length factors, or increments, and multiple web charts (acceleratego/stop, V1/VR) for a particular configuration.
(3) Section 25.107(b) - Minimum Takeoff Safety Speed (V 2MIN).
(a) V 2MIN, in terms of calibrated airspeed, cannot be less than:
1 1.1 times the V MCA defined in § 25.149.
2 1.13 times V SR for two-engine and three-engine turbopropeller and reciprocating engine-powered airplanes and for all turbojet airplanes that do not have provisions for obtaining significant reduction in the one-engine inoperative power-on stalling speed (i.e., boundary layer control, blown flaps, etc.). The value of VS 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.
(b) V 2MIN 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.
(c) 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 poweron 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) Section 25.107(c) - Takeoff Safety Speed (V 2). V2 is the calibrated airspeed that is attained at or before the airplane reaches a height of 35 ft. above the takeoff surface after an engine failure at VEF using an established rotation speed (VR). During the takeoff speeds demonstration, V2 should be continued to an altitude sufficient to assure stable conditions beyond the 35 ft 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). Section 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 SR-422A, paragraphs 4T.114(b)(4) and (c)(3) and 4T.116(e). The intent of these requirements is discussed in the preamble to SR-422A, which states, in part, “For these reasons, this regulation permits the airplane to lift off the ground at a speed lower than the V2 speed.” The concern that the regulation change was addressing was the overshoot of V2 after liftoff when it was required that the airplane attain V2 on, or near, the ground. It was therefore the intent of the regulation 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.
(a) In the case of turbojet powered airplanes, when the bulk of the one-engineinoperative data have been determined with idle cuts, V2, and its relationship to VR, should be substantiated by a limited number of fuel cuts at VEF. For derivative programs not involving a new or modified engine type (i.e., a modification that would affect thrust decay characteristics), fuel cuts are unnecessary if thrust decay characteristics have been adequately substantiated.
(b) 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, if any, depends on the correlation obtained with the idle cut data and substantiation that the data analysis methodology adequately models the effects of a sudden engine failure.
(5) Section 25.107(d) - Minimum Unstick Speed (V MU).
(a) An applicant should comply with § 25.107(d) by conducting minimum unstick speed (VMU) determination tests with all engines operating and with one engine inoperative. If a stick pusher is installed, it should normally be active and set to the minimum angle-of-attack side of its rigging tolerance band. If desired, artificial stall warning systems may be disabled for this demonstration. (See paragraph (f) below for further discussion regarding the setting of stick pusher and stall warning systems during VMU testing.) During this demonstration, the takeoff should be continued until the airplane is out of ground effect. The airplane pitch attitude should not be decreased after liftoff.
(b) In lieu of conducting actual one-engine-inoperative VMU tests, the applicant may conduct all-engines-operating VMU tests that simulate and account for all pertinent factors that would be associated with an actual one-engine-inoperative VMU test. To account fully for pertinent factors, it may be necessary to adjust the resulting VMU test values analytically. The factors to be accounted for should include at least the following:
1 Thrust/weight ratio for the one-engine-inoperative range.
2 Controllability (may be related to one-engine-inoperative free air tests, such as VS, VMCA, etc.).
3 Increased drag due to lateral/directional control systems.
4 Reduced lift due to devices such as wing spoilers used for lateral control.
5 Adverse effects of any other systems or devices on control, drag, or lift.
(c) The number of VMU tests required may be minimized by testing only the critical all-engines-operating and one-engine-inoperative thrust/weight ratios, provided that 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.
(d) 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 allengines-operating tests at the thrust/weight ratio corresponding to the one-engine-inoperative
23 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-engines-operating and one-engineinoperative conditions as was previously required. In determining the all-engines-operating thrust/weight ratio that corresponds to the one-engine-inoperative condition, consideration should be given to 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.
(e) 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 (g), 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. 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.
(f) V MU testing 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 consequential 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. 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. If the airplane is equipped with a stick pusher, for flight test safety reasons it should normally be active and set to the minimum angle-of-attack side of its rigging tolerance band. However, depending on the airplane’s stall characteristics and the stick pusher design, disabling the pusher or delaying activation of the system until a safe altitude is reached may be the safer course. Again, this decision should be made only with a thorough knowledge of the airplane’s stall characteristics combined with a complete understanding of the stick pusher design.
(g) V MU Testing for Airplanes Having Limited Pitch Control Authority.
1 For some airplanes with limited pitch control authority, it may not be possible, at forward c.g. 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. 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 c.g. limit and/or with use of more airplane nose-up trim than normal. The VMU data determined with this procedure should be
Change 1 24 corrected to those values representative of the appropriate forward limit; the variation of VMU with c.g. may be assumed to be like the variation of free air stalling speed with c.g. Although the development of scheduled takeoff speeds may proceed from these corrected VMU data, additional tests are required (see paragraph 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.
2 In the following assurance test, the airplane should demonstrate safe flyaway characteristics.
(aa) Minimum speed liftoff should be demonstrated at the critical forward
c.g. limit with normal trim. For airplanes with a cutback forward c.g. at heavy weight, two weight/c.g. 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 c.g.
The full forward c.g. tests should be conducted at the highest associated weight. These tests should be conducted at minimum thrust/weight for both the simulated one-engine-inoperative test (symmetrical reduced thrust) and the all-engines-operating case.
(bb) 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 overrotation, but the liftoff attitude should be maintained as the airplane flies off the ground and out of ground effect.
(cc) Resulting liftoff speeds do not affect AFM speed schedules if the test proves successful and the resulting liftoff speed is at least 5 knots below the normally scheduled liftoff speed. Adjustments should be made to the scheduled VR, forward c.g. limit, etc., if necessary, to achieve this result.
(dd) This minimum 5 knot reduction below the scheduled liftoff speed provides some leeway for operational variations such as mis-trim, c.g. errors, etc., that could further limit the elevator authority. The reduced VMU margins of the minimum liftoff speeds demonstrated in this test, relative to those specified in § 25.107(e)(1)(iv), result from the reduced probability of a pitch control authority-limited airplane getting into a high drag condition due to overrotation.
(h) VMU Testing for Geometry Limited Airplanes.
1 For airplanes that are geometry limited (i.e., the minimum possible V MU 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.
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