3-5 needed to conduct safe testing cannot be installed on the production airplane configuration, or the weight of such instrumentation still prevents the minimum weight from being obtained, consider the use of simulation to extend the results obtained at the minimum practical test weight. (See paragraph 3.1.2.6 of this AC.)
3.1.4.3.3 For follow-on airplane certification programs involving an increase in the maximum allowable gross weight, the test weight limits of table 3-1 have been applied as extrapolation limits on the original test data in order to minimize additional testing. For the test weight tolerance limits to be applied in this manner, the original test data must be from an existing certificated database for an aerodynamically similar model of the same airplane type. The tolerance limit should be applied to the maximum weight at which the original testing was conducted, not to the maximum certified weight.
3.1.4.3.4 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 table 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 3-1 below for an example of how this can be done.) If the resulting equivalent gross weight does not exceed the certificated maximum landing weight by more than the five percent weight extrapolation limit specified in table 3-1, the reduced flap certification may be eligible for a reduced flight test program (e.g., limited to stall speed verification, handling characteristics, and a 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.
3-6 Figure 3-1. Equivalent Weight Extrapolation ∆ G.W. Not greater than 5% Gross Weight - Pounds ∆ KE Max. Landing Wt.
“Equivalent” Gross Wt.
Flap Brake Energy Ori ginal Reduced Flap
3.1.4.4 Wind Limits.
A wind velocity limit of 10 knots (from any direction) or 0.11 VSR1 (whichever is lower) is considered the maximum acceptable for obtaining valid takeoff and landing flight test data. Takeoff and landing performance data obtained under runway wind conditions greater than 5 knots may be inconsistent and unreliable because winds of that magnitude are likely to be unsteady. However, performance data obtained with winds between 5 and 10 knots should not necessarily be discarded. Their validity should be checked against data obtained in conditions with lesser winds. Wind velocity should be measured at the height of the wing mean aerodynamic chord (MAC), as determined with the airplane in a static ground attitude.
When measuring test wind velocity at the wing MAC height, a height of six feet above the ground should be considered as a minimum measurement height to avoid possible measurement inaccuracies due to surface interference.
3.1.4.5 CG Limits.
A test tolerance of ±7 percent of the total CG range is intended to allow for inflight CG movement. This tolerance is only acceptable when the test data scatter is on both sides of the limiting CG or when adjusting the data from the test CG to the limit CG is acceptable. If compliance with a requirement is marginal at a test condition that is inside of the CG limits, the test should be repeated at the CG limits.
3.1.4.6 Airspeed Limits.
Normally, tests conducted within 3 percent or 3 knots (whichever is the higher) of the desired test speed are considered acceptable.
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3.1.4.7 Thrust/Power Limits.
Thrust critical tests, such as minimum control speeds, should be conducted at the highest thrust (or 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 (or power) corrections should be limited to 5 percent of test day thrust (or power), unless a detailed analysis is performed.
3.1.4.8 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 should always be conducted under controlled conditions and with the flight test crew’s full knowledge of the situation.
Examples of such flights are:
3.1.4.8.1 Takeoff at greater than maximum takeoff weight to reach a test area at the maximum takeoff weight.
3.1.4.8.2 Landing at greater than maximum landing weights during the course of conducting takeoff tests.
3.1.4.8.3 Flights to obtain data for future approvals beyond that substantiated for the initial type design.
3.1.4.8.4 Table 3-2 below indicates the cases for which corrections are normally allowed. Any corrections to flight test data should be made by methods that are agreed to by the FAA.
3-8 Table 3-2. Test Parameters that Normally can be Corrected
3.1.4.9 All instrumentation used in the flight test program should be appropriately calibrated and acceptable to the FAA test team.
3.1.5 Section 25.21(f)—Proof of Compliance: Wind Measurement and Corrections.
3.1.5.1 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 𝑉𝑊2 = 𝑊𝑖𝑛𝑑 𝑣𝑒𝑙𝑜𝑐𝑖𝑡𝑦 𝐻2 𝑉𝑊1 = 𝑊𝑖𝑛𝑑 𝑣𝑒𝑙𝑜𝑐𝑖𝑡𝑦 𝐻1
3.1.5.2 This equation is presented graphically below. Values of H less than 5 feet should not be used in this relationship.
Flight Test Condition Correctable Parameters Weight CG 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 performance X X --- X X X Minimum control speed --- --- --- --- X --- Minimum unstick speed X X X --- X --- Buffet boundary X X --- X --- ---
3-9 Figure 3-2. Wind Profile Variation
3.1.6 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 3.1.5 above. The following examples are methods of handling wind profile variation data. Other methods have also been found acceptable.
Example: Test Data Given:
Height of mean aerodynamic chord with airplane on surface 8.0 feet Height of wind measurement 6.0 feet Measured wind velocity 4.8 knots Results:
Test wind velocity with airplane 50 feet above landing surface
4.8((50 + 8)/6)1/7 = 6.6 knots Test wind velocity with airplane 35 feet above takeoff surface
4.8((35 + 8)/6)1/7 = 6.4 knots Test wind velocity with airplane on surface 4.8(8/6)1/7 = 5.0 knots 0 10 20 30 40 50 60 70 0 10 20 30 40 50 60
HEIGHT ABOVE FIELD ELEVATION - FEET
WIND VELOCITY - KNOTS 10 meters
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3.1.7 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 mean aerodynamic chord, which corresponds to the wind condition as measured at 10 meters (32.81 feet) above the takeoff surface, and corrected for the wind factors of § 25.105(d)(1).
Example: AFM Data Given:
Height of mean aerodynamic chord with airplane on surface
8.0 feet Reported headwind at 10 meters 40.0 knots Section 25.105(d)(1) wind factor 0.5 Results:
Factored wind velocity with airplane 50 feet above landing surface
(0.5)(40)((50 + 8)/32.81)1/7 = 21.7 knots Factored wind velocity with airplane 35 feet above takeoff surface
(0.5)(40)((35 + 8)/32.81)1/7 = 20.8 knots Factored wind velocity with airplane on surface (0.5)(40)(8/32.81)1/7 = 16.3 knots
3.1.8 Airplane Airspeed Variation Due to Wind Profile Variation Combined with 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 shown only by the change in ground speed. These changes in ground speed may be generalized either as speed increments or speed ratios. The changes in airspeed due to wind profile variation are superimposed on these speed changes.
Example: Determination of True Airspeed from Ground Speed—Takeoff Test Data Given:
Ground speed at liftoff, VLOF 139.0 knots Ground speed at 35 feet above takeoff surface 140.6 knots Speed change due to airplane dynamic performance 1.6 knots Test headwind at liftoff 5.0 knots Test headwind with airplane 35 feet above takeoff surface 6.4 knots Results:
True airspeed at liftoff, VLOF 139.0 + 5.0 = 144.0 knots True airspeed at 35 feet above takeoff surface 140.6 + 6.4 147.0 knots
3-11 Example: Determination of Rotation Speed from True Airspeed at 35-Foot Height—AFM Data Given:
Factored headwind at liftoff 16.3 knots Factored headwind with airplane 35 feet above takeoff surface 20.8 knots Ground speed change, (V35 – VLOF) 1.6 knots Ground speed change, (VLOF – VR) 0.5 knots True airspeed required at 35 feet 150.0 knots Results:
Ground speed required at 35 feet 150 - 20.8 = 129.2 knots Ground speed at liftoff 129.2 - 1.6 = 127.6 knots True airspeed at liftoff 127.6 + 16.3 = 143.9 knots Ground speed at rotation 127.6 - 0.5 = 127.1 knots True airspeed at rotation (for AFM speed and distances) 127.1+16.3 = 143.4 knots
Example: Landing—AFM Data Given:
Factored headwind with airplane 50 feet above landing surface 21.7 knots Factored headwind with airplane on landing surface 16.3 knots Ground speed change for 50 feet to touchdown (V50 – VTD) 4.0 knots True airspeed required at 50 feet 130 knots Results:
Ground speed at 50 feet 130 - 21.7 = 108.3 knots Ground speed at touchdown 108.3 - 4.0 = 104.3 knots True airspeed at touchdown 104.3 + 16.3 = 120.6 knots
3.1.9 Expansion of Takeoff and Landing Data for a Range of Airport Elevations.
3.1.9.1 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.
3.1.9.2 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
3-12 extrapolation of landing data, provided the effect of the higher true airspeed on landing distance was taken into account.
3.1.9.3 Since then, considerably more experience has since 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.
3.1.9.4 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 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 4.1 of this AC for more information on an acceptable means of establishing and verifying installed propulsion system performance characteristics.)
3.1.9.5 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 20.1.2.3 of this AC. Consideration should be given to any sensitivity to, or dependency upon airport altitude, such as: engine and auxiliary power unit (APU) starting, passenger oxygen, autopilot, autoland, autothrottle system power/thrust set/operation.
3.1.10 Tailwind Takeoff and Landing.
3.1.10.1 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 without conducting additional flight tests at specific wind speeds.
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3.1.10.2 Wind Velocities Greater than 10 Knots.
3.1.10.2.1 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 factor, required by §§ 25.105(d)(1) and
25.125(f), provides a satisfactory level of safety for operation in tailwinds up to 15 knots when using AFM data based on flight tests in nominally calm 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.
3.1.10.2.2 Control Characteristics.
The test tailwind velocity for demonstrating handling qualities should be equal to the proposed limit tailwind factored by 150 percent. The intent of the 150 percent factor is to provide adequate margin for wind variability in operations, including currency of the wind data, averaging of the data by the measuring and reporting method, and the highly variable nature of higher wind conditions. Therefore, the test wind condition of 150 percent of the proposed tailwind limit should be an averaged or smoothed wind speed, not a peak wind speed. Airplane control characteristics should be evaluated under the following conditions with the CG at the aft limit and the test mean tailwind velocity equal to the proposed limit tailwind factored by 150 percent:
1. Takeoff. Both all-engines-operating and one-engine-inoperative
(i.e., with a simulated failure of the critical engine at the engine failure speed) takeoffs should be evaluated at a light weight with maximum approved takeoff flap deflection.
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 or shimmy, and flight director, or autopilot instrument landing system (ILS) approaches, terrain awareness warning system (TAWS) sink rate modes, etc.
4. If engine idle power or thrust is increased to account for the increased tailwind velocity, ensure that deviations above the glideslope are recoverable.
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3.1.10.2.3 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.
3.1.10.2.4 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 or thrust setting procedure used for AFM performance (typically completed by approximately 80 knots), both manually and automatically (autothrottle).
3. Reverse thrust operations.
3.1.10.2.5 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.
3.1.11 Procedures.
3.1.11.1 The performance-related flight test procedures are discussed in the following paragraphs of this AC:
Paragraph 4.2, Takeoff and Takeoff Speeds—§§ 25.105 and 25.107.
Paragraph 4.3, Accelerate-Stop Distance—§ 25.109.
Paragraph 4.4, Takeoff Path—§ 25.111.
Paragraph 4.5, Takeoff Distance and Takeoff Run—§ 25.113.
Paragraph 4.6, Takeoff Flight Path—§ 25.115.
Paragraph 4.7, Climb: General—§ 25.117.
Paragraph 4.8, Landing Climb: All-Engines-Operating—§ 25.119.
Paragraph 4.9, Climb: One-Engine-Inoperative—§ 25.121.
Paragraph 4.10, En Route Flight Paths—§ 25.123.
Paragraph 4.11, Landing—§ 25.125.
3.1.11.2 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
3-15 constant configuration of other lift devices. If the span of flap settings is small and previously obtained data provide sufficient confidence (i.e., the shapes of the curves are known and lend themselves to accurate interpolation), data from three flap settings may be acceptable.
3.1.11.3 Flight Characteristics for Abnormal Configurations.
See § 25.671(c).
3.1.11.3.1 For purposes of this AC, an abnormal configuration is an operational configuration that results from any single failure or any combination of failures not shown to be improbable.
3.1.11.3.2 Flight characteristics for abnormal configurations may be determined by test or analysis to assure that the airplane is capable of continued safe flight and landing. Flight tests, if required, should be conducted at the critical conditions of altitude, weight, CG, and engine power or thrust associated with the configuration, and at the most critical airspeed between the speed reached one second after stall warning occurs (see paragraph 8.1.5.2.8 of this AC) and the maximum operating airspeed for the configuration.
3.2 Load Distribution Limits—§ 25.23. [Reserved]
3.3 Weight Limits and Center of Gravity Limits—§§ 25.25 and 25.27. [Reserved]
3.4 Empty Weight and Corresponding Center of Gravity—§ 25.29. [Reserved]
3.5 Removable Ballast—§ 25.31.
3.5.1 Explanation.
None.
3.5.2 Procedures.
Ballast may be carried during the flight tests whenever it is necessary to achieve a specific weight and CG location. Consideration should be given to the vertical as well as horizontal location of the ballast in cases where it may have an appreciable effect on the flying qualities of the airplane. The strength of the supporting structures should be considered in order to make sure they do not fail as a result of the anticipated loads that may be imposed during the particular tests. As required by § 21.35(a), applicants must show that these structures comply with the applicable structural requirements of part 25 before conducting flight tests with these structures in place.
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3.6 Propeller Speed and Pitch Limits—§ 25.33.
3.6.1 Explanation.
None.
3.6.2 Procedures.
The tachometers and the airspeed indicating system of the test airplane should have been calibrated within the last six months. With that prerequisite satisfied, the following should be accomplished:
3.6.2.1 Determine that the propeller speeds and pitch settings are safe and satisfactory during all tests that are conducted in the flight test program within the certification limits of the airplane, engine, and propeller. This includes establishing acceptable low pitch (flight idle) blade angles on turbopropeller airplanes and verifying that propeller configurations are satisfactory at VMO/MMO to prevent propeller overspeed.
3.6.2.2 Determine that the propeller speeds and pitch settings are safe and satisfactory during all tests that are conducted to satisfy the performance requirements.
3.6.2.3 With the propeller governors operative and the propeller controls in full high revolutions per minute (RPM) position, determine that the maximum takeoff power settings do not exceed the rated takeoff RPM of each engine during takeoff and climb at the best rate-of-climb speed.
3.6.2.4 With the propeller governors made inoperative by mechanical means, determine the maximum power, no-wind, static RPMs. With the propeller governors operating on the low pitch stop, the engine speeds must not exceed 103 percent of the maximum allowable takeoff RPM or 99 percent of an approved maximum overspeed, as required by § 25.33(c). On turbopropeller engines, the engine speeds should not exceed the maximum engine speeds allowed by engine and propeller type designs. Note which systems were disabled and how the disablement was done. If maximum takeoff power torque or sea level standard conditions cannot be obtained on the test day, correct the data to these conditions by an acceptable means. A no-wind condition is considered to be a wind of 5 knots or less.
The static RPM should be the average obtained with a direct crosswind from the left and a direct crosswind from the right.
3.6.2.5 If the above determinations are satisfactory, then measure the low-pitch stop setting and the high-pitch stop setting. These data may have been obtained from the propeller manufacturer and may be used, provided the pitch stops have not been changed since the manufacturer delivered the propeller. If measured, the blade station should be recorded. Include these blade angles in the type certificate data sheet.