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.
(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