3 Chapter 1 - General
1. Applicability - § 25.1 [Reserved]
2. Special Retroactive Requirements - § 25.2 [Reserved]
Chapter 2 - Flight
Section 1. General
3. Proof of Compliance - § 25.21.
a. Explanation . In an effort to provide the necessary guidelines for the flight test evaluation of transport category airplanes, without producing a cumbersome document, this AC assumes a conventional transport airplane configuration. In general, a conventional airplane configuration is one with distinct wing and fuselage elements that are joined together, aftmounted horizontal and vertical stabilizers that are attached to the fuselage, and propulsion provided either by turbojet/turbofan engines that do not provide any significant increase in lift due to their operation or engine-driven propellers. The effects of non-conventional airplane configurations (e.g., blown flaps) on the compliance methods should be evaluated and determined based on the intent of the guidelines presented for conventional airplane configurations.
(1) Section 25.21(a) - Proof of Compliance .
( a) The burden of showing compliance with the flight requirements for an airworthiness certificate or a type certificate rests with the applicant. The applicant should, at his own expense and risk, conduct such official flight tests as required by the FAA to demonstrate compliance with the applicable requirements. During the certification process, the applicant should make available the airplane, as well as all of the personnel and equipment necessary to obtain and process the required data.
( b) If the airplane flight characteristics or the required flight data are affected by weight and/or center of gravity (c.g.), the compliance data must be presented for the most critical weight and c.g. position per § 25.21(a). Unless the applicant shows that the allowable c.g. travel in one or more axes (e.g., lateral fuel imbalance) has a negligible effect on compliance with the airworthiness requirements, the applicant must substantiate compliance at the critical c.g.
( c) The gross weight and c.g. tolerances specified in paragraphs 3a(3)(b)1 and 3 are test tolerances and are not intended to allow compliance to be shown at less than critical conditions.
(d) Section 21.35(a)(3) requires that the test airplane be in conformity with its type design specifications.
1 This means that the test airplane must be in conformity with its type design specification as relates to the particular test being conducted.
2 Any deviation from conformity must be clearly shown to be of no consequence to the particular test being conducted. For example, if the slip resistant escape surface required by § 25.810(c) is not installed, the applicant must show that its presence would have no effect on measured airplane performance and flight characteristics.
(e) Section 21.35(b)(2) requires the applicant to conduct sufficient flight testing the FAA finds necessary to determine whether there is reasonable assurance that the airplane, its components, and its equipment are reliable and function properly. Appendix 2 to this AC provides guidance for showing compliance with this requirement.
(2) Section 25.21(c) - Proof of Compliance (Altitude Effect on Flight Characteristics ).
(a) Any of the flying qualities, including controllability, stability, trim, and stall characteristics, affected by altitude must be investigated for the most adverse altitude conditions approved for operations.
(b) Consideration should be given in the test program to aerodynamic control system changes with altitude (e.g., control throws or auto slats which are sometimes inhibited by Mach number at altitude).
(3) Section 25.21(d) - Proof of Compliance (Flight Test Tolerances).
(a) To allow for variations from precise test values, acceptable tolerances during flight testing must be maintained. The purpose of these tolerances is to allow for variations in flight test values from which data are acceptable for reduction to the value desired. They are not intended for tests to be routinely scheduled at the lower weights, or to allow for compliance to be shown at less than the critical condition; nor are they to be considered as allowable inaccuracy of measurement. As an example, when demonstrating stability with a specified trim speed of
1.3 VSR1, the trim speed may be 1.3 VSR1 +3 knots or 3 percent; however, no positive tolerance is permitted when demonstrating the minimum prescribed trim speed of 1.3 VSR1 (Ref. § 25.161).
(b) Where variation in the parameter on which a tolerance is allowed will have an effect on the results of the test, the results should be corrected to the most critical value of that parameter within the operating envelope being approved. If such a correction is impossible or impractical (e.g., performance at forward c.g.), the average test conditions should assure that the measured characteristics represent the actual critical value.
1 Weight limits. Figure 3-1, below, presents weight tolerances that have been found acceptable for the specified flight tests. Many flight tests need to be conducted at or very near the maximum operating weight for the airplane configuration, particularly those tests used to establish airplane flight manual (AFM) performance information. As noted in paragraph
(a) above, the purpose of the test tolerances is to allow for variations in flight test values, not to routinely schedule tests at less than critical weight conditions or to allow for compliance to be shown at less than the critical weight condition.
Figure 3-1. Weight Tolerance Limits
Flight Test Conditions Weight Tolerance Limit +5% +10% Stall Speeds X Stall Characteristics X Climb Performance X Takeoff Flight Paths X Landing Braking Distance X Landing Air Distance X Takeoff Distance & Speed X Accelerate-Stop Distance X Maximum Energy RTOs X Minimum Unstick Speed X
(aa) For follow-on airplane certification programs (e.g., minor gross weight increase), where it is desired to use an existing certificated data base from an aerodynamically similar model of the same airplane type, the test weight tolerance limits of figure 3-1 have also been applied as extrapolation limits to minimize additional test requirements.
(bb) For follow-on airplane certification programs, test data may be extrapolated to higher gross weights than specified in figure 3-1, if the available test data include an adequate range of weights and an appropriate number of points at the highest weight tested.
If the test data analysis verifies the predicted effect of weight, then other extrapolation limits may apply.
(cc) Equivalent weight extrapolation limits. For follow-on airplane certification programs where it is desired to increase a maximum operating weight based on existing certified performance parameters that have weight as one of their independent terms, those parameters should be examined for equivalent compliance with the weight tolerance limits of figure 3-1. An example would be the reduction of an airplane’s landing flap position, to one approved on a similar model of the same airplane type, which would incur an increase in landing speeds and brake energy, relative to the original certificated landing flap, at any given weight.
The brake energy, at the maximum certificated landing weight, should be calculated for the reduced landing flap. This brake energy should account for the increased landing speeds and reduced aerodynamic drag associated with the reduced flap setting. It should then be determined what equivalent gross weight would have rendered that brake energy with the original landing flap (see figure 3-2). The resulting equivalent gross weight should not exceed the certificated
7 maximum landing weight by more than the five percent weight extrapolation limit specified in figure 3-1. If the equivalent gross weight (at the existing landing flap setting) is not more than five percent greater than the certificated maximum landing weight, the reduced flap certification may be eligible for a limited flight test program (e.g., stall speed verification, handling characteristics, and qualitative landing demonstration). Further limitations may be imposed by the criteria of technical standard order (TSO) C135a (“Transport Airplane Wheels and Wheel and Brake Assemblies,” dated July 1, 2009).
Figure 3-2. Equivalent Weight Extrapolation
2 Wind Limits. For takeoff and landing tests, a wind velocity limit of 10 knots (from any direction) or 0.11 VSR1 (whichever is lower) at the height of the mean aerodynamic chord (MAC), as determined with the airplane in a static ground attitude, has been considered the maximum acceptable. When test wind velocity is to be measured at the wing MAC height, a measurement height of six feet above the ground should be considered as a limit to avoid possible measurement inaccuracies due to surface interference. Because of likely unsteady wind conditions, it is generally considered that takeoff and landing performance data obtained under runway wind conditions greater than 5 knots are likely to be inconsistent and unreliable. If performance data are obtained with winds greater than 5 knots, these data should not necessarily be ignored.
3 C.G. Limits. A test tolerance of + 7 percent of the total c.g. range is intended to allow some practical relief for inflight c.g. movement. This relief is only acceptable when the test data general scatter is on both sides of the limiting c.g. or when c.g. correction from test c.g. to limit c.g. is acceptable.
4 Airspeed Limits. Normally, tests conducted within 3 percent or 3 knots (whichever is the higher) of the desired test speed are considered acceptable.
5 Thrust Limits. Thrust critical tests, such as minimum control speeds, should be conducted at the highest thrust (power) allowable on the engine given the constraints of temperature and altitude. It is then permitted to calculate further corrections to allow extrapolation of data to cover the entire operating envelope. These thrust corrections should be limited to 5 percent of test day thrust, unless a detailed analysis is performed. (See Appendix 7 of this AC).
(c) Because the tolerance values normally are not considered in the airplane design substantiation, it is not the purpose of these tolerances to allow flights at values in excess of those authorized in the type design. If such flights are to be conducted, adequate structural substantiation for the flight conditions should be available. These flights, however, are always conducted under controlled conditions and with the flight test crew’s full cognizance of the situation. Examples of such flights are:
1 Takeoff at greater than maximum takeoff weight for the purpose of reaching a test area at the maximum takeoff weight.
2 Landing at greater than maximum landing weights during the course of conducting takeoff tests.
3 Flights to obtain data for future approvals beyond that substantiated for the initial type design.
(d) The table in figure 3-3 indicates the cases in which corrections are normally allowed. Any corrections to flight test data should be made by methods that are agreed to by the FAA.
Figure 3-3. Test Parameters That Normally Can Be Corrected
Flight Test Condition Correctable Parameters Wt. C.G. Airspeed Altitude Power/ Thrust Wind Airspeed calibration X --- --- --- --- --- Stall speeds X X --- --- X --- Climb performance X X X X X --- Landing performance X --- X X --- X Takeoff performance X X --- X X X Accelerate-stop perf. X X --- X X X Minimum control speed --- --- --- --- X --- Minimum unstick speed X X X --- X --- Buffet boundary X X --- X --- ---
(e) All instrumentation used in the flight test program should be appropriately calibrated and acceptable to the FAA test team.
(4) Section 25.21(f) - Proof of Compliance (Wind Measurement and Corrections).
The relationship between the wind measured at one height and the corresponding wind at another height may be obtained by the following equation:
VW2 = VW1(H2/H1)1/7
Where: H = Height above the runway surface V W2 = Wind velocity at H2 V W1 = Wind velocity at H1
This equation is presented graphically below. Values of H less than 5 feet should not be used in this relationship.
Figure 3-4. Wind Profile Variation
(5) Wind Profile Variation for Test Data. The performance data of airplanes should be obtained in such a manner that the effect of wind on the test data may be determined. The test wind velocity should be corrected from the recorded height above the test surface to the height of the airplane wing mean-aerodynamic-chord. If the wind profile variation is not measured, the variation may be calculated using the equation in paragraph (4) above. The following examples are methods of handling wind profile variation data. Other methods have also been found acceptable.
Example: Test Data
Given:
o Height of mean-aerodynamic-chord with airplane on surface 8.0 ft.
o Height of wind measurement 6.0 ft.
o Measured wind velocity 4.8 kts.
0 10 20 30 40 50 60 70 0 1 02 03 04 05 06 0
WIND VELOCITY - KNOTS
HEIGHT ABOVE FIELD ELEVATION - FEET 10 meters
Results:
o Test wind velocity with airplane 50 ft. above landing surface
4.8((50 + 8)/6) 1/7 = 6.6 kts.
o Test wind velocity with airplane 35 ft. above takeoff surface
4.8((35 + 8)/6) 1/7 = 6.4 kts.
o Test wind velocity with airplane on surface
4.8(8/6) 1/7 = 5.0 kts.
(6) Wind Profile Variation for AFM Data. When expanding the data to the AFM conditions, the result should include the effective velocity, at the airplane’s wing meanaerodynamic-chord, which corresponds to the wind condition as measured at 10 meters (32.81 ft.) above the takeoff surface, and corrected for wind factors of § 25.105(d)(1).
Example: Airplane Flight Manual Data
Given:
o Height of mean-aerodynamic-chord with airplane on surface 8.0 ft.
o Reported head wind at 10 meters 40.0 kts.
o Section 25.105(d)(1) wind factor 0.5
Results:
o Factored wind velocity with airplane 50 ft. above landing surface
(0.5)(40)((50 + 8)/32.81) 1/7 = 21.7 kts.
o Factored wind velocity with airplane 35 ft. above takeoff surface
(0.5)(40)((35 + 8)/32.81) 1/7 = 20.8 kts.
o Factored wind velocity with airplane on surface
(0.5)(40)(8/32.81) 1/7 = 16.3 kts.
(7) Airplane Airspeed Variation Due to Wind Profile Variation With Simultaneous Speed Changes Due to Airplane Dynamic Performance. In the reduction of test data and in the expansion of such data to AFM conditions, the increase or decrease of speed due to the dynamic effect of the forces on the airplane are manifested by only the change in ground speed. These changes in speed (zero wind speed changes) may be generalized either as speed increments or
12 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, V LOF 144.0 kts.
o True airspeed at 35 ft. above takeoff surface 147.0 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 ( V 35-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, (V 35-VLOF) 1.6 kts.
o Zero wind speed change, (V LOF-VR) 0.5 kts.
o True airspeed required at 35 ft. 150.0 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 (V R), given in the AFM, should be predicated on the required speed at 35 ft. minus the speed increment (V 35-VR) assuming no airspeed change due to wi nd 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 pr ofile, 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 marg in of safety by this increase in V R
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.