CHAPTER 8. AIRWORTHINESS-MISCELLANEOUS ITEMS
228. Design and Function of Artificial Stall Warning and
Identification Systems701
229. Reduced and Derated Thrust Takeoff Operations704
230. Runway Gradients Greater than +2 Percent705
231. Performance Requirements for Flight in Icing Conditions707
232. Flying Qualities in Natural Icing Conditions708
233. Criteria for Approval of Steep Approach to Landing709
234. Takeoff and Landing on Unpaved Runways713
235. Accountability of Performance De gradation Relative Both Minor Design Changes (External Exposure) and
Configuration Deviation List (CDL) Items716
236. Configuration Deviation List (CDL)717
237. Spare Engine Pod 719
238. Authorization for Ferry Flight with One Engine Inoperative719
239. Instrument Landing System Weather Minima720
240. Takeoff Performance Credit for Alternate Forward Center of Gravity Limits721
241. Performance Credit for Automatic Propeller Drag Reduction Devices724
242. Airplane Backing Using Reverse Thru st725
APPENDICES Appendix 1 History of Jet Transport Performance Standards (CAR 4b Special Regulations 422, 422A and 422B) Appendix 2 History of Accelerate-Stop Transition Time Development Appendix 3 Correction of Air Minimum Control Speed (V MCA) to Standard Conditions Appendix 4 Rudder Pedal Force-Limited Air Minimum Control Speed (V MCA) Appendix 5 One-G Stall Compliance Basis Appendix 6 Sample Flight Test Program [RESERVED Appendix 7 FAA Handling Qualities Rating System Appendix 8 Functional and Reliability Tests (§ 21.35(b)(2))
CHAPTER 1 – GENERAL
1.-2. [RESERVED]
Chap 2 12 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 advisory circular assumes a conventional transport airplane configuration. In general, a “conventional” airplane configuration is one with: distinct wing and fuselage elements that are joined together; aft-mounted 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.
(i) 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.
(ii) If the airplane flight characteristics or the required flight data are affected by weight and/or center of gravity (c.g.), the compliance data should be presented for the most critical weight and c.g. position. Unless it can be shown that the allowable c.g. travel in one or more axes (e.g., lateral fuel imbalance) has a negligible effect on the airworthiness requirements, compliance at the critical c.g. must be substantiated.
(iii) The gross weight and c.g. tolerances specified in paragraphs 3a(3)(ii)(A) and (D) are test tolerances and are not intended to allow compliance to be shown at less than critical conditions.
(iv) Section 25.21(a)(1) requires, in part, that the test airplane be in conformity with its type design specifications.
(A) This is interpreted as meaning the test airplane must be in conformity with its type design specification as relates to the particular test being conducted.
(B) 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
Chap 2 13 escape surface required by §25.810(c) is not installed, it must be shown that its presence would have no effect on measured airplane performance and flight characteristics.
(2) Section 25.21(c) - Proof of Compliance (Altitude Effect on Flight Characteristics).
(i) Any of the flying qualities, including controllability, stability, trim, and stall characteristics, affected by altitude should be investigated for the most adverse altitude conditions approved for operations and limited, as appropriate, in accordance with § 25.1527.
(ii) 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).
(i) To allow for variations from precise test values, certain tolerances during flight testing should 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 1.4VS1, the trim speed may be
1.4VS1 +3 knots or 3 percent; however, no positive tolerance is permitted when demonstrating the minimum prescribed trim speed of 1.4VS1 (Ref. § 25.161).
(ii) 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.
(A) 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 performance information. As noted in paragraph (i), 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 that the critical weight condition.
Chap 2 14 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
(1) 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.
(2) 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 includes 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.
(3) 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 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
Chap 2 15 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) C26c ("Aircraft Wheels and Wheel-Brake Assemblies, with Addendum I," dated May 18, 1984), wheel fusible plug energy limits, quick turnaround considerations, the maximum demonstrated kinetic energy level, wheel spin-up loads, etc.
FIGURE 3-2. EQUIVALENT WEIGHT EXTRAPOLATION (B) Wind Limits. For takeoff and landing tests, a wind velocity limit of 10 knots (from any direction) or 0.12VS1 (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.
(C) 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.
(D) Airspeed Limits. Normally, tests conducted within 3 percent or 3 knots (whichever is the higher) of the desired test speed are considered acceptable.
Chap 2 16 (E) 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 are limited to 5 percent of test day thrust, unless a detailed analysis is performed (see Appendix 3 of this AC).
(iii) 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:
(A) Takeoff at greater than maximum takeoff weight for the purpose of reaching a test area at the maximum takeoff weight.
(B) Landing at greater than maximum landing weights during the course of conducting takeoff tests.
(C) Flights to obtain data for future approvals beyond that substantiated for the initial type design.
(iv) 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 --- ---
(v) All instrumentation used in the flight test program should be appropriately calibrated and acceptable to the FAA test team.
Chap 2 17
(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 VW2 = Wind velocity at H2 VW1 = 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.
0 10 20 30 40 50 60 70 0 10 20 30 40 50 60
WIND VELOCITY - KNOTS 10 meters
Chap 2 18 Example: Test Data Given:
o Height of mean-aerodynamic-chord with airplane on surface 8 ft.
o Height of wind measurement 6 ft.
o Measured wind velocity 4.8 kts.
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 Airplane Flight Manual (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 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
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