must be positively damped with controls free and must be controllable with normal use of the primary controls without requiring exceptional piloting skill.
7.4.2 Procedures.
7.4.2.1 Dynamic Longitudinal Stability.
7.4.2.1.1 The test for longitudinal dynamic stability is accomplished by a rapid movement or pulse of the longitudinal control in a nose up and nose down direction at a rate and degree necessary to obtain a short period pitch response from the airplane.
7.4.2.1.2 Dynamic longitudinal stability should be checked at a sufficient number of points in each configuration to assure compliance at all operational speeds.
7.4.2.2 Dynamic Lateral-Directional Stability.
7.4.2.2.1 A typical test for lateral-directional dynamic stability is accomplished by a rudder doublet input at a rate and amplitude that will excite the lateral-directional response (i.e., Dutch roll). The control input should be in phase with the airplane’s oscillatory response.
7.4.2.2.2 Dynamic lateral-directional stability should be checked under all conditions and configurations. If critical, special emphasis should be placed on adverse wing fuel loading conditions.
7.4.2.3 Airplanes Equipped with Stability Augmentation Systems (SAS).
7.4.2.3.1 In the event a SAS is required for the airplane to show compliance with § 25.181(a) or (b), it must meet the requirements of §§ 25.671 and 25.672.
Additionally:
7.4.2.3.2 If the airplane is equipped with only one SAS (i.e., a single strand system), in accordance with § 25.672, compliance with the dynamic stability requirements of § 25.181(a) or (b), as applicable, must be shown throughout the normal operating flight envelope to be certificated with the SAS operating, and in a reduced, practical operating flight envelope that will permit continued safe flight and landing with the SAS inoperative.
7.4.2.3.3 If the airplane is equipped with more than one SAS, the resulting effects of SAS failure should be considered when determining whether or not the primary and any redundant SAS should be operating simultaneously for showing compliance with the dynamic stability requirements of § 25.181(a) or (b). If the primary and redundant SAS are dissimilar, the functional capability (i.e., control authority) of the redundant SAS should be considered with regard to restricting the operating envelope after failure of the primary SAS. At the applicant’s option, however, compliance with 7-11
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AC 25-7D § 25.181(a) or (b) may still be demonstrated to a reduced flight envelope with no SAS operating as described in paragraph 7.4.2.3.2 above.
7.4.2.3.4 Regardless of the SAS redundancy, the airplane should be safely controllable at the point of system failure or malfunction anywhere in the approved operating flight envelope of the airplane. Accordingly, it should be demonstrated that the airplane remains controllable during transition from the operating SAS to any redundant SAS, and during transition from anywhere in the normal operating envelope to the reduced practical operating envelope of § 25.672(c), if applicable. Airplane controllability should be demonstrated to meet the following levels as defined by the FAA HQRM. (The FAA HQRM is described in appendix E of this AC.)
1. In the normal operating flight envelope with the SAS operating, the handling qualities should be “satisfactory” (SAT) as defined by the FAA HQRM.
2. At the point of SAS failure in the normal operating envelope, the airplane should be “controllable” (CON), as defined by the FAA HQRM, during the short term transitory period required to attain a speed and configuration that will permit compliance with paragraph 3 below.
3. During transition from the primary SAS to a redundant SAS, or from the normal operating envelope to a reduced, practical operating envelope (where applicable), the handling qualities should be “adequate” (ADQ) as defined by the HQRM.
4. In the reduced, practical operating flight envelope that will permit continued safe flight and landing, the handling qualities should be “satisfactory” (SAT) as defined by the HQRM.
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AC 25-7D CHAPTER 8. FLIGHT: STALLS
8.1 Stall Testing.
8.1.1 Applicable 14 CFR Regulations.
Section 25.21(c), Proof of compliance.
Section 25.103, Stall speed.
Section 25.143, Controllability and maneuverability—General.
Section 25.201, Stall demonstration.
Section 25.203, Stall characteristics.
Section 25.207, Stall warning.
8.1.2 Explanation.
8.1.2.1 The purpose of stall testing is threefold:
8.1.2.1.1 To define the reference stall speeds and how they vary with weight, altitude, and airplane configuration.
8.1.2.1.2 To demonstrate that handling qualities are adequate to allow a safe recovery from the highest angle-of-attack attainable in normal flight (stall characteristics).
8.1.2.1.3 To determine that there is adequate pre-stall warning (either aerodynamic or artificial) to allow the pilot time to recover from any probable high angle-of-attack condition without inadvertently stalling the airplane.
8.1.2.2 During this testing, the angle-of-attack should be increased at least to the point where the behavior of the airplane gives the pilot a clear and distinctive indication through the inherent flight characteristics or the characteristics resulting from the operation of a stall identification device
(e.g., a stick pusher) that the airplane is stalled.
8.1.3 Stall Demonstration—§ 25.201.
8.1.3.1 The airplane is considered to be fully stalled when any one or a combination of the characteristics listed below occurs to give the pilot a clear and distinctive indication to cease any further increase in angle-of-attack, at which time recovery should be initiated using normal techniques.
8.1.3.1.1 The pitch control reaches the aft stop and is held full aft for two seconds, or until the pitch attitude stops increasing, whichever occurs later. In the case of turning flight stalls, recovery may be initiated once the pitch 8-1
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AC 25-7D
control reaches the aft stop when accompanied by a rolling motion that is not immediately controllable (provided the rolling motion complies with § 25.203(c)).
8.1.3.1.2 An uncommanded, distinctive, and easily recognizable nose down pitch that cannot be readily arrested. This nose down pitch may be accompanied by a rolling motion that is not immediately controllable, provided that the rolling motion complies with § 25.203(b) or (c), as appropriate.
8.1.3.1.3 The airplane demonstrates an unmistakable, inherent aerodynamic warning of a magnitude and severity that is a strong and effective deterrent to further speed reduction. This deterrent level of aerodynamic warning
(i.e., buffet) should be of a much greater magnitude than the initial buffet ordinarily associated with stall warning. An example is a large transport airplane that exhibits “deterrent buffet” with flaps up and is characterized by an intensity that inhibits reading cockpit instruments and would require a strong determined effort by the pilot to increase the angle-of-attack any further.
8.1.3.1.4 The activation point of a stall identification device that provides one of the characteristics listed above. See paragraph 42.1 of this AC for additional guidance material on demonstrating compliance with the regulatory requirements of part 25 for stall identification systems.
8.1.3.2 It should be recognized that the point at which the airplane is considered stalled may vary, depending on the airplane configuration (e.g., flaps, gear, CG, and gross weight). In any case, the angle-of-attack should be increased until one or more of these characteristics is reached for all likely combinations of variables.
8.1.4 Stall Speeds.
8.1.4.1 Background.
Since many of the regulations pertaining to performance and handling qualities specify trim speeds and other variables that are functions of stall speeds, it is desirable to accomplish the stall speed testing early in the program, so the data are available for subsequent testing. Because of this interrelationship between the stall speeds and other critical performance parameters, it is essential that accurate measurement methods be used.
Most standard airplane pitot-static systems are unacceptable for stall speed determination. These tests require the use of properly calibrated instruments and usually require a separate test airspeed system.
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8.1.4.2 Configuration.
8.1.4.2.1 Stall speeds should be determined for all aerodynamic configurations to be certificated for use in the takeoff, en route, approach, and landing configurations.
8.1.4.2.2 The CG positions to be used should be those that result in the highest stall speeds for each weight (forward CG in most cases).
8.1.4.2.3 Sufficient testing should be conducted to determine the effects of weight on stall speed. Altitude effects (compressibility, Reynolds Number) may also be considered if credit for variations in these parameters is sought by the applicant. If stall speeds are not to be defined as a function of altitude, then all stall speed testing should be conducted at a nominal altitude no lower than 1,500 feet above the maximum approved takeoff and landing altitude. (See paragraph 8.1.4.5.7 of this AC.)
8.1.4.3 Procedures.
8.1.4.3.1 The airplane should be trimmed for hands-off flight at a speed 13 percent to 30 percent above the anticipated VSR, with the engines at idle and the airplane in the configuration for which the stall speed is being determined.
Then, using only the primary longitudinal control for speed reduction, maintain a constant deceleration (entry rate) until the airplane is stalled, as defined in § 25.201(d) and paragraph 8.1.3.1 of this AC. Following the stall, engine power or thrust may be used as desired to expedite recovery.
8.1.4.3.2 A sufficient number of stalls (normally four to eight) should be accomplished at each critical combination of weight, altitude, CG, and external configuration. The intent is to obtain enough data to determine the stall speed at an entry rate not exceeding 1.0 knot/second. During the maneuver for determining stall speeds, the flight controls should be operated smoothly in order to achieve good data quality rather than trying to maintain a constant entry rate because experience has shown that adjusting the flight controls to maintain a constant entry rate leads to fluctuations in load factor and significant data scatter.
8.1.4.3.3 During the stall speed testing, the stall characteristics of the airplane must also satisfy the requirements of § 25.203(a) and (b).
For airplanes that have stall identification devices for which the angle-of-attack for activation is biased by angle-of-attack rate, some additional considerations are necessary. The stall speeds are normalized against an average airspeed deceleration rate, as described in paragraph 8.1.4.5.5. However, stall identification systems generally activate at a specific angle-of-attack, biased by an instantaneous angle-ofattack rate. Therefore, longitudinal control manipulation by the pilot during the stall maneuver, close to the stall identification system activation 8-3
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AC 25-7D point, can advance or delay its activation without appreciably affecting the average stall entry airspeed rate. To minimize scatter in the stall speed versus entry rate data, the pilot should attempt to maintain a stable angleof-attack rate or pitch rate (not necessarily a fixed airspeed deceleration rate), until the stall identification system activates. The resulting timehistory of angle-of-attack data should be smooth and without discontinuities. A cross plot of airspeed deceleration rate, as defined in paragraph 8.1.4.5.5, versus angle-of-attack rate for all related test points, will show the general trend of this relationship for each flap setting. Any points that do not follow this general trend should not be used in establishing the stall speed.
8.1.4.4 Thrust Effects on Stall Speed.
8.1.4.4.1 Stall speeds are typically determined with the thrust levers at idle;
however, it is necessary to verify by test or analysis that engine idle thrust does not result in appreciably lower stall speeds than would be obtained at zero thrust. Prior to amendment 25-108, a negative idle thrust at the stall, which slightly increases stall speeds, was considered acceptable, but applicants were not required to base stall speeds on idle thrust. With the adoption of amendment 25-108, it became a requirement to base stall speeds on idle thrust, except where that thrust level results in a significant decrease in stall speeds. If idle thrust results in a significant decrease in stall speeds, then stall speeds cannot be based on more than zero thrust.
8.1.4.4.2 To determine whether thrust effects on stall speed are significant, at least three stalls should be conducted at one flap setting, with thrust set to approximately the value required to maintain level flight at 1.5 VSR in the selected configuration.
8.1.4.4.3 These data may then be extrapolated to a zero thrust condition to determine the effects of idle thrust on stall speeds. (See figure 8-1 below.)
If the difference between idle thrust and zero thrust stall speed is 0.5 knots or less, the effect may be considered insignificant.
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AC 25-7D Figure 8-1. Thrust Effect on Stall Speed
THRUST EFFECT ON
STALL SPEED
STALL SPEED ZERO THRUST THRUST FOR 1.5 VSR
IN LEVEL FLIGHT IDLE THRUST
ENGINE THRUST
8.1.4.4.4 The effects of engine power on stall speeds for a turbopropeller airplane can be evaluated in a similar manner. Stall speed flight tests should be accomplished with engines idling and the propellers in the takeoff position. Engine torque, engine RPM, and estimated propeller efficiency can be used to predict the thrust associated with this configuration.
8.1.4.5 Data Reduction and Presentation.
The following is an example of how the data obtained during the stall speed testing may be reduced to standard conditions. Other methods may be found acceptable.
8.1.4.5.1 Record the indicated airspeed from the flight test airspeed system throughout the stall and correct these values to equivalent airspeed. Also record load factor normal to the flight path. Typically, the load factor data would be obtained from a sufficient number of accelerometers capable of resolving the flight path load factor. It may be possible to obtain acceptable data using one accelerometer aligned along the expected 1-g stall pitch angle. More likely, it will take at least two accelerometers, one aligned along the fuselage longitudinal axis and one aligned at 90° to that axis, as well as a means to determine the angle between the flight path and the fuselage longitudinal axis.
8.1.4.5.2 Calculate the airplane lift coefficient (CL) from the equation given below and plot it as a time history throughout the stall maneuver:
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AC 25-7D Where:
8.1.4.5.3 The maximum lift coefficient CLMAX is defined as the maximum value of CL achieved during the stall test. Where the time history plot of CL exhibits multiple peak values, CLMAX normally corresponds to the first maximum. However, the peak corresponding to the highest CL achieved may be used for CLMAX, provided it represents usable lift, meaning that it does not occur after deterrent buffet or other stall identification cue (ref.
§ 25.201(d)). There should also typically be a noticeable break in a plot of the load factor normal to the flight path near the point at which CLMAX is reached. The analysis to determine CLMAXshould disregard any transient or dynamic increases in recorded load factor, such as might be generated by abrupt control inputs that do not reflect the lift capability of the airplane.
The load factor normal to the flight path should be maintained at nominally 1.0 until CLMAX i s reached. (See figure 8-2 below.)
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Figure 8-2. CIMAX and Load Factor
2.2
1.8 CLHAX
1.4-293388/2988808|8288|833|33888 FACTOR АТ 1.0 g BREAK § 0.6
8.1.4.5.4 Correct the CIMAX obtained for each stall, if necessary, from the test CG position to the targeted CG position, and for any thrust effects, using the equation:
CIMAX = CIMAX (Pest CG position) [1 + (MAC/ L) (CGstd - CGeest)] - ACL Where:
MAC = Wing mean aerodynamic chord length (inches) It = Effective tail length, measured between the wing 25 percent MAC and the stabilizer 25 percent MAC (inches) CGstd = CG position resulting in the highest value of reference stall speed, normally the forward CG limit at the pertinent weight (percent MAC/100) CGtest = Actual test CG position (percent MAC/100) ACp = Change in C1 due to engine thrust (if effect of idle thrust is greater than 0.5 knots in stall speed 8-7
8.1.4.5.5 Determine the stall entry rate, which is defined as the slope of a straight line connecting the stall speed and an airspeed 10 percent above the stall speed, for each stall test. Because CLMAX is relatively insensitive to stall entry rate, a rigorous investigation of entry rate effects should not be necessary.
8.1.4.5.6 For each approved configuration, construct a plot of CLMAX versus weight.
(See figure 8-3 below.)
Figure 8-3. CLMAX versus Weight and Flap Setting CL MAX
8.1.4.5.7 Flight test safety concerns usually dictate the lowest test altitude for determining stall speeds. The test data should then be expanded to lower altitudes, and hence lower Mach numbers, to cover the operational envelope of the airplane. Since CLMAX usually increases as the Mach number is reduced, simple expansion of the flight test data could result in extrapolating to a higher CL LMAX than tested. The expansion of C MAX versus Mach number data is only permitted up to the highest CLMAX demonstrated within the range of W/’s tested, unless the continuation of the trend of higher CLMAX with decreasing Mach number is substantiated with other test data. For example, data obtained at a more aft CG position or with power on can be used for this purpose if CG and thrust effects can be accounted for. Data from another airplane in the same family with the 8-8
same wing and showing the same general trend of CLMAX Versus Mach
(e.g., a lighter weight variant) may also be used if shown to be applicable.
8.1.4.5.8 The reference stall speed, VsR, is a calibrated airspeed defined by the applicant. Vsk may not be less than the 1-g stall speed and is expressed as:
VSR 2 VGLMAX VRzw Where:
CIMAX = 295.37 (nzw)(W) /CIMAYS + AVc.
if the stalling maneuver is limited by a device that commands an abrupt nose down pitch (e.g., a stick pusher), CLux may not be less than the speed existing at the instant the instant the device operates.
AVc = Compressibility correction (i.e., the difference between equivalent airspeed and calibrated airspeed) W = Airplane weight (Ibs) Nzw = Airplane load factor normal to the flight path CLMAX = Value of CLMAX Corresponding to the chosen weight (see Figure 8-4 below) 5 = Reference wing area (ft?)
8.1.4.5.9 Construct a plot of reference stall speed versus weight for each flap/gear configuration. (See figure 8-4 below).
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Figure 8-4. Stall Speed versus Weight and Flap Setting
A B C D
INCREASING
FLAP SETTING WEIGHT
STALL SPEED (KEAS)
8.1.4.5.10 For airplanes equipped with a device that abruptly pushes the nose down at a selected angle-of-attack (e.g., a stick pusher), VSR must not be less than the greater of 2 knots or 2 percent above the speed at which the device activates (§ 25.103(d)).
8.1.4.5.11 In showing compliance with § 25.103(d) for airplanes equipped a device that abruptly pushes the nose down at a selected angle-of-attack (e.g., a stick pusher), the speed at which the device operates need not be corrected to 1 g. Requiring a load factor correction of the device activation speed to the 1-g condition would unnecessarily increase the stringency of § 25.103(d). For example, it would be possible for the device activation speed to be assessed as higher than VSR (or at least closer to VSR than would be obtained without correcting to the 1 g condition). Test procedures should be in accordance with paragraph 8.1.4.3.1 to ensure that no abnormal or unusual pilot control input is used to obtain an artificially low speed at which the device first activates.
8.1.5 Stall Characteristics—§ 25.203.
8.1.5.1 Background.
To assure a safe and expeditious recovery from an unintentional stall, it should not require any unusual piloting technique to successfully demonstrate compliance with § 25.203, nor should it require exceptional skill or repeated practice by the test pilot. The behavior of the airplane 8-10