Chap 2 19 generalized either as speed increments or speed ratios. The changes in airspeed due to wind profile are superimposed on these speed changes.
Example: Takeoff Test Data Given:
o True airspeed at liftoff, VLOF 144 kts.
o True airspeed at 35 ft. above takeoff surface 147 kts.
o Test wind at liftoff 5.0 kts.
o Test wind with airplane 35 ft. above takeoff surface 6.4 kts.
Results:
o Speed change due to airplane dynamic performance, zero wind speed change (V35-VLOF) = (147-6.4)-(144-5.0) 1.6 kts.
Example: Takeoff--Airplane Flight Manual Data Given:
o Factored wind at liftoff 16.3 kts.
o Factored wind with airplane 35 ft. above takeoff surface 20.8 kts.
o Zero wind speed change, (V35-VLOF) 1.6 kts.
o Zero wind speed change, (VLOF-VR) 0.5 kts.
o True airspeed required at 35 ft. 150 kts.
Results:
o Ground speed required at 35 ft. 150-20.8 = 129.2 kts.
o Ground speed at liftoff 129.2-1.6 = 127.6 kts.
o True airspeed at liftoff 127.6+16.3 = 143.9 kts.
o Ground speed at rotation 127.6-0.5 = 127.1 kts.
o True airspeed at rotation (for distance calculations) 127.1+16.3 = 143.4 kts.
o Airplane flight manual rotation speed, true airspeed 150-0.5-1.6 = 147.9 kts.
NOTE: The indicated airspeed at rotation (VR), given in the AFM, should be predicated on the required speed at 35 ft. minus the speed increment (V35-VR) assuming no airspeed change due to wind profile (i.e., use the "zero wind speed change" increments shown above). The proper rotation speed, assuming the airplane will gain 4.5 knots due to wind profile, is 143.4 knots. However, if this airspeed increase does not materialize, the airplane will be slow by 4.5 knots at 35 feet. It is therefore more desirable to allow the airplane to attain this 4.5 knots on the ground. Any reduction in field length margin of safety by this increase in VR is adequately compensated for by the 50 percent wind factor.
Chap 2 20 This also eliminates a requirement to adjust VR for wind. However, in calculating the AFM field lengths, the airspeed attained due to wind profile may be included.
Example: Landing--Airplane Flight Manual 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 (V50-VTD) 4.0 kts.
o True airspeed required at 50 ft. 130 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 Data for a Range of Airport Elevations.
(i) These guidelines are applicable to expanding takeoff data above the altitude at which the basic or verifying tests were obtained.
(ii) In general, takeoff data may be extrapolated above and below the altitude at which the basic test data was obtained without additional conservatism within certain constraints. The extrapolation limits and constraints for expansion of the measured takeoff performance parameters (e.g., data used to develop V/VS=f(T/W)) to operating altitudes is addressed in sub-paragraph 3a(9)(iii), below. Extrapolation of takeoff performance data may be further limited by engine data extrapolation limits, typically to 3,000 feet above the highest altitude at which engine parameters were evaluated for the pertinent power/thrust setting. (See paragraph 9 of this AC for more information on engine testing and data extrapolation.) 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 the AC.
(iii) When the basic takeoff test data are obtained between sea level and approximately 3,000 ft., the maximum allowable extrapolation limits are 6,000 ft. above, and 3,000 ft. below, the test field elevation. If it is desired to extrapolate beyond these limits, when proven data reduction and expansion methods are used, one of the two following procedures may be employed:
(A) Extrapolation of Performance Data for a Range of Altitudes When Verifying Tests are Not Conducted. The approval of performance data for airport elevations beyond the maximum elevation permitted by basic tests may be allowed without conducting verifying tests, if the calculated data include a conservative factor.
This conservatism should result in an increase of the calculated takeoff distance at the
Chap 2 21 desired airport elevations by an amount equal to zero percent for the highest airport elevation approved on the results of the basic tests, and an additional cumulative 2 percent incremental factor for each 1,000 ft. of elevation above the highest airport elevation approved for zero percent conservatism. The 2 percent incremental factor should have a straight line variation with altitude. When performance data are calculated for the effects of altitude under this procedure, the following provisions are applicable:
(1) Previously established calculation procedures should be used, taking into account all known variables.
(2) The calibrated installed engine power for the pertinent speed and altitude ranges should be used.
(3) The brake kinetic energy limits established by airplane ground tests should not be exceeded.
(B) Extrapolation of Performance Data When Verifying Tests Are Conducted.
(1) If data approval is desired for a greater range of airport elevations, the performance may be calculated from the basic test data up to the maximum airport elevation, provided verifying tests are conducted at appropriate elevations to substantiate the validity of the calculations. The actual airplane performance data from the verifying tests should correspond closely to the calculated performance values.
(2) For the verifying tests, it has been found that normally three takeoffs at maximum weights for the elevations tested will provide adequate verification.
(3) If verifying tests substantiate the expanded takeoff data, the data may be further expanded up to 6,000 ft. above the altitude at which the verifying tests were conducted. At altitudes higher than 6,000 ft. above the verifying test altitude, the 2 percent per 1,000 ft. cumulative factor discussed in paragraph 3a(9)(iii)(A), above, should be applied, starting at zero percent at the verifying test altitude plus 6,000 feet.
(9) Expansion of Landing Data for a Range of Airport Elevations.
(i) For turbine powered airplanes without propellers, the landing distance data may be extrapolated to more than 6,000 ft. above the test altitude without the penalty described in paragraph 3a(9), provided the true airspeed effect on the distance is accounted for.
(ii) For turbopropeller and reciprocating engine-powered airplanes, the extrapolated performance should include the conservative factor described in paragraph 3a(9). This may not be necessary if it can be shown that the effects of altitude on the
Chap 2 22 landing performance, as affected by propeller drag, are known from data available for the airplane configuration.
(10) Tailwind Takeoff and Landing.
(i) 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.
(ii) Wind Velocities Greater than 10 Knots.
(A) 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(e), 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.
(B) 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:
(1) Takeoff. At light weight with maximum approved takeoff flap deflection, both all-engines operating and one-engine inoperative takeoffs should be evaluated.
(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, shimmy, flight director and/or autopilot Instrument Landing System (ILS) approaches, Ground Proximity Warning Systems (GPWS) sink rate modes, etc.
(4) If engine idle thrust is increased to account for the increased tailwind velocity, ensure that deviations above the glideslope are recoverable.
(C) Weight Limits. In accordance with the requirements of §§
25.105(d)(1) and 25.125(e), maximum takeoff weight and maximum quick turnaround
Chap 2 23 weight should be determined using brake energies and tire speeds, as appropriate, calculated with the limit tailwind velocity factored by 150 percent.
(D) 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 setting procedure used for AFM performance (typically completed by approximately 80 knots), both manually and automatically (autothrottle).
(3) Reverse thrust operations.
(E) 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 15. Climb: General
11. Accelerate-Stop Distance 16. Landing Climb
12. Takeoff Path 17. Climb: One Engine Inoperative
13. Takeoff Distance and Takeoff Run 18. Enroute Flight Path
14. Takeoff 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 is available, the data may be obtained from interpolation of flight data obtained at no less than four flap settings which are within a constant configuration of other lift devices. If the span of flap settings is small and previously obtained data provides 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)).
(i) 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.
Chap 2 24
(ii) 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 stall warning (see paragraph 29d(2)(ix) of this AC) and the maximum operating airspeed for the configuration.
4.-7. [RESERVED]
Chap 2 25
8. PROPELLER SPEED AND PITCH LIMITS - § 25.33.
a. Explanation. None.
b. Procedures. The tachometers and the airspeed indicating system of the test airplane must have been calibrated in 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 r.p.m. position, determine that the maximum takeoff power settings do not exceed the rated takeoff r.p.m. of each engine during takeoff and climb at the best rate-of-climb speed.
(4) With the propeller governors made inoperative by mechanical means, maximum power, no-wind, static r.p.m.'s must be determined. On reciprocating engines 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. On turbopropeller engines, the engine speeds must 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, the data must be corrected to these conditions by an acceptable means. A no-wind 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, the low-pitch stop setting and also the high-pitch stop setting must be measured. 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. These blade angles must be included in the Type Certificate Data Sheet.
Chap 2 26 Section 2. PERFORMANCE
9. GENERAL - § 25.101
a. Explanation - Engine 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….” Therefore, engine parameter evaluations, engine torque, thrust lapse, propeller thrust, and throttle sensitivity, especially with the Electronic Control Unit off (if installed), should be accounted for at all altitudes up to and including the maximum approved takeoff altitude.
b. Procedures . This aspect of the rule may be accomplished by actual flight test, with approved takeoff ratings, at the maximum desired takeoff altitude, or with acceptable parameter variation analysis associated with tests conducted at less than the maximum objective takeoff altitude.
(1) One acceptable method of demonstrating high altitude engine operability and performance is to simulate the maximum operating altitude conditions by overboosting (i.e., using a higher than normal power setting) at a lower altitude.
(2) It is the applicant’s responsibility, in conjunction with the engine manufacturer, to define the test plan for the engine parameter evaluation up to the maximum approved takeoff altitude. Supportive data from the engine manufacturer on the same “type or series” of engine may also be included to reduce the amount of testing conducted.
(3) Extrapolation of engine performance data to 3,000 feet above the highest altitude tested (up to the maximum takeoff altitude to be approved) has been accepted, provided the supportive data, including the flight testing accomplished, establishes a consistent baseline.
(4) Detailed guidance for the testing necessary to show compliance with § 25.101(c) may be found in Advisory Circular 25-XX, “Turbojet Operations at High Altitude Airports,” dated (to be released).
10. TAKEOFF AND TAKEOFF SPEEDS - §§ 25.105 AND 25.107 (PART 25 AS AMENDED THROUGH AMENDMENT 25-42) .
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 234) of this Advisory Circular (AC).
b. Procedures . Although the following speed definitions are given in terms of calibrated airspeed, the Airplane Flight Manual (AFM) presentations shall 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 V 1 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 177 of this AC.
(1) Section 25.107(a)(1) - Engine Failure Speed ( VEF). The engine failure speed (V EF) is defined as the calibrated airspeed at which the critical engine is assumed to fail and must be selected by the
Chap 2 27 applicant. VEF cannot be less than the ground minimum control speed (V MCG) as described in §§ 25.149(e) and 25.107(a)(1).
* (2) Section 25.107(a)(2) - V 1. V1 may not be less than V EF plus the speed gained with the critical engine inoperative during the time interval between V EF and the instant at which the pilot takes action after recognizing the engine failure (with the proviso that the time interval may not be less than one second). This action is indicated by pilot initiation of the first action to stop the airplane, such as brakes, throttles, spoilers, etc., during accelerate-stop tests, or by the first control input during V MCG testing. The applicant may choose the sequence of pilot actions. Refer to paragraph 11 of this AC, addressing § 25.109, for a more complete description of rejected takeoff transition procedures and associated time delays. If it becomes evident in expanding the takeoff data for presentation in the AFM that excessive variation in V 1 exists due to the many performance variables involved (variations of +1.5 knots or +100 ft. have been found acceptable), then measures must be taken to reduce the variability displayed in the AFM presentation. Examples of such measures are field length factors, or increments, and multiple web charts (accelerate-go/stop, V 1/VR) for a particular configuration.
(3) Section 25.107(b) - Minimum Takeoff Safety Speed (V 2MIN).
(i) V2MIN, in terms of calibrated airspeed, cannot be less than:
(A) 1.1 times the VMCA defined in § 25.149.
(B) 1.2 times VS 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 V S to be used in determining V 2-MIN 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.
(ii) V2MIN may be reduced to 1.15 times V S for turbopropeller and reciprocating enginepowered 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.
(iii) For propeller-driven airplanes, the diff erence 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) Section 25.107(c) - Takeoff Safety Speed (V2 ). 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 V EF using an established rotation speed (V R). During the takeoff speeds demonstration, V 2 should be continued to an altitude sufficient to assure stable conditions beyond the 35-ft height. V 2 cannot be less than V 2MIN.
In addition, V2 cannot be less than the liftoff speed, V LOF, as defined in § 25.107(f). In accordance with §
25.107(c), V2 in terms of calibrated airspeed “...may not be less than V R plus the speed increment attained before reaching a height of 35 feet above the takeoff surface.” 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 V 2 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
Chap 2 28 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 V 2 speed.” The concern that the regulation change was addressing was the overshoot of V 2 after liftoff when it was required that the airplane attain V 2 on, or near, the ground.
It was therefore the intent of the regulation to allow an acceleration to V 2 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 V 2, under low drag ground conditions, and using the excess kinetic energy to attain the 35 foot height.
(i) In the case of turbojet powered airplanes, when the bulk of the one-engineinoperative data have been determined with idle cuts, V 2, and its relationship to V R, should be substantiated by a limited number of fuel cuts at V EF. 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.
(ii) For propeller-drive n 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).
(i) 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
(vi) below for further discussion regarding the setting of stick pusher and stall warning systems during V MU 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.
(ii) In lieu of conducting actual one-engine-in operative VMU tests, the applicant may conduct all-engines-operating V MU tests that simulate and account for all pertinent factors that would be associated with an actual one-engine-inoperative V MU test. To account fully for pertinent factors, it may be necessary to adjust the resulting V MU test values analytically. The factors to be accounted for must at least include the following:
(A) Thrust/weight ratio for the one-engine-inoperative range.
(B) Controllability (may be related to one-engine-i noperative free air tests, such as VS, VMCA, etc.).
(C) Increased drag due to lateral/directional control systems.
(D) Reduced lift due to devices such as wing spoilers used for lateral control.
(E) Adverse effects of any other systems or de vices on control, drag, or lift.
(iii) The number of VMU tests required may be minimized by testing only the critical allengines-operating and one-engine-inoperative thrust/weight ratios, provided that the V MU speeds determined at these critical conditions are used for the range of thrust/weights appropriate to the allengines-operating and one-engine-inoperative configurations. The critical thrust/weight is established by correcting, to the V MU speed, the thrust that results in the airplane achieving its limiting one-engineinoperative climb gradient at the normally scheduled speed and in the appropriate configuration.
(iv) 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 V MU to be determined by all-engines-
Chap 2 29 operating tests at the thrust/weight ratio corresponding to the one-engine-inoperative condition. As revised, § 25.107(d) specifies that V MU must be selected for the range of thrust/weight ratios to be certificated, rather than for the all-engines-operating 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, 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 V MU 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.
(v) To conduct the VMU tests, rotate the airplane as necessary to achieve the V MU attitude.
It is acceptable to use some additional nose-up trim over the normal trim setting during V MU demonstrations. If additional nose-up trim is required, the additional considerations of paragraph (vii), 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.
(vi) VMU 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 V MU 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.
(vii) VMU Testing for Airplanes Having Limited Pitch Control Authority.
(A) For some airplanes with lim ited 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, V MU 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 V MU data determined with this procedure should be 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 V MU data, additional tests are required (see paragraph (B) below) to check that the relaxed V MU criteria have not neglected problems that might arise from operational variations in rotating airplanes with limited pitch control authority.
(B) In the following assurance test, the airplane must demonstrate safe flyaway characteristics.
(1) Minimum speed liftoff should be demonstrated at the c ritical 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
Chap 2 30 for both the simulated one-engine-inoperative test (symmetrical reduced thrust) and the all-enginesoperating case.
(2) 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.
(3) 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 V R, forward c.g. limit, etc., if necessary, to achieve this result.
(4) 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 V MU 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.
(viii) VMU Testing for Geometry Limited Airplanes.
(A) For airplanes that are geometry limited, the 11 0 percent of VMU required by §
25.107(e)(1)(iv) may be reduced to an operationally acceptable value of 108 percent on the basis that equivalent airworthiness is provided for the geometry-limited airplane. For acceptance of the 108 percent of VMU liftoff speed, safeguards protecting the geometry-limited airplane against both overrotation on the ground and in the air must be provided. Also, the airplane should be geometry limited to the extent that a maximum gross weight takeoff with the tail dragging will result in a clean liftoff and fly-away in the allengines-operating condition. During such a takeoff for the all-engines-operating condition, the resulting distance to the 35 ft. height must not be greater than 105 percent of the normal takeoff distance un der similar weight, altitude, and temperature conditions before the 15 percent margin is added. Lastly, the V MU demonstrated must be sound and repeatable. Compliance with these criteria should be formally addressed in a finding of equivalent safety.
(B) The criteria for demonstrating the capability for a clean liftoff and fly-away are as follows:
(1) The airplane’s pitch attitude from a speed of 96 percent of the actual liftoff speed must be within 5 percent (in degrees) of the tail-dragging attitude to the point of liftoff.
(2) During the above speed range (96 to 100 percent of the actual liftoff speed), the aft under-surface of the airplane must have achieved actual runway contact. It has been found acceptable in tests for contact to exist approximately 50 percent of the time that the airplane is in this speed range.
(3) Beyond the point of liftoff to a height of 35 ft., the airplane’s pitch attitude must not decrease below that at the point of liftoff, or the speed must not increase more than 10 percent.
(4) The airplane shall be at the critical thrust/weight condition, as defined in §
25.107(d), but with all engines operating.
(ix) VMU for a Stretched Version of a Tested Airplane.
(A) It has been considered that the VMU speed represents the aerodynamic potential minimum liftoff speed related to a safe liftoff and fly-away, even though the body contact attitude may