Easily controllable by the pilot.
7.3.1.3 Steady Straight Sideslips.
7.3.1.3.1 Section 25.177(c) requires, in steady, straight sideslips throughout the range of sideslip angles appropriate to the operation of the airplane, that the aileron and rudder control movements and forces be proportional to the angle of sideslip. Also, the factor of proportionality must lie between limits found necessary for safe operation. Section 25.177(c) also states th at that the range of sideslip angles evaluated mu st include those sideslip angles resulting from the lesser of: (1) one-half of the available rudder control input; and (2) a rudder control force of 180 lbs. This means that if using one-half of the available rudder control input takes less than 180 lbs of force, then compliance mu st be based on using one-half of the available rudder control input. If application of 180 lbs of rudder control force results in using less than one-half of the available rudder control input, then compliance must be based on applying 180 lbs of rudder control force. By cross-reference to § 25.177(a), § 25.177(c) requires that these steady, straight sideslip criteria be met for all landing gear and flap positions and symme tric al power or thrust conditions at speeds from
1.13 VSR1 to VFE, VLE, or VFC/MFC, as appropriate for the configuration.
7.3.1.3.2 Experience has shown that an acceptable metho d for determining th e appropriate sideslip angle for the operation of a transport category airplane is provided by the following equation:
Where:
7.3.1.3.3 Recognizing that smaller sideslip angles are appropriate as speed is increased, this equation provides sideslip angle as a function of airspeed.
The equation is based on the theoretical sideslip value for a 30-knot crosswind, but has been shown to conservatively represent (i.e., exceed) the sideslip angles achieved in maximum crosswind takeoffs and landings and minimum static and dynamic control speed testing for a variety of transport category airplanes. Experience has also shown that a maximum sideslip angle of 15° is generally appropriate for most transport category airplanes even though the equation above may provide a higher sideslip angle. However, limiting the maximum sideslip angle to 15° may not be appropriate for airplanes with low approach speeds or high crosswind capability.
7.3.1.3.4 A lower sideslip angle than that provided in paragraph 7.3.1.3.2 above may be used if it is substantiated that the lower value conservatively 7-6
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AC 25-7D covers all crosswind conditions, engine failure scenarios, and other conditions where sideslip may be experienced within the approved operating envelope. Conversely, a higher value should be used for airplanes where test evidence indicates that a higher value would be appropriate to the operation of the airplane.
7.3.1.3.5 For the purpose of showing compliance with the requirement out to sideslip angles associated with the lesser of (1) one half of the available rudder control input; and (2) a rudder control force of 180 lbs; there is no need to consider a rudder control input beyond that corresponding to full available rudder surface travel. Some rudder control system designs may limit the available rudder surface deflection such that full deflection for the particular flight condition, or the maximum commanded sideslip angle for the flight condition, is reached before the rudder control reaches onehalf of its available travel. In such cases, further rudder control input is unnecessary as it would not result in a higher sideslip angle, and therefore would not affect compliance with the rule.
7.3.1.4 Full Rudder Sideslips.
7.3.1.4.1 At sideslip angles greater than those appropriate for normal operation of the airplane, up to the sideslip angle at which full rudder control input is used or a rudder control force of 180 lbs is obtained, § 25.177(d) requires that the rudder pedal control may not reverse, and increased rudder deflection must be needed for increased angles of sideslip. The goals of this higher-than-normal sideslip angle test are to show that at full rudder control input, or at maximum expected pilot effort (1) the rudder control force does not reverse, and (2) increased rudder deflection must be needed for increased angles of sideslip, thus demonstrating freedom from rudder lock or fin stall, and adequate directional stability for maneuvers involving large rudder inputs.
7.3.1.4.2 Compliance with this requirement should be shown using straight, steady sideslips. However, if full lateral control input is reached before full rudder control travel or a rudder control force of 180 lbs is reached, the maneuver may be continued in a non-steady heading (i.e., rolling and yawing) maneuver. Care should be taken to prevent excessive bank angles that may occur during this maneuver.
7.3.1.4.3 Section 25.177(d) states that the criteria listed in paragraph 7.3.1.4.1 above must be met at all approved landing gear and flap positions for the range of operating speeds and power conditions appropriate to each landing gear and flap position with all engines operating. The range of operating speeds and power conditions appropriate to each landing gear and flap position with all engines operating should be consistent with the following:
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AC 25-7D
1. For takeoff configurations, speeds from V2+xx (airspeed approved for all-engines-operating initial climb) to VFE or VLE, as appropriate, and takeoff power/thrust;
2. For flaps up configurations, speeds from 1.23 VSR to VLE or VMO/MMO, as appropriate, and power from idle to maximum continuous power/thrust;
3. For approach configurations, speeds from 1.23 VSR to VFE or VLE, as appropriate, and power from idle to go-around power/thrust; and
4. For landing configurations, speeds from VREF-5 knots to VFE or VLE, as appropriate, with power from idle to go-around power/thrust at speeds from VREF to VFE/VLE, and idle power at VREF-5 knots (to cover the landing flare).
7.3.2 Procedures.
The test conditions should include each flap and landing gear configuration as described in paragraphs 7.3.1.1 through 7.3.1.4 of this AC at an altitude appropriate to each configuration.
7.3.2.1 Basic Tests for Static Directional and Lateral Stability.
7.3.2.1.1 Static Directional Stability.
To check static directional stability with the airplane in the desired configuration and stabilized at the trim speed, the airplane is slowly yawed in both directions while maintaining the wings level with the roll controls.
When the rudder is released, the airplane should tend to return to straight flight.
7.3.2.1.2 Static Lateral Stability.
To check lateral stability with a particular configuration and trim speed, conduct steady, straight sideslips at the trim speed by maintaining the airplane heading with rudder and banking with the roll controls. When the roll controls are released, with the rudder held fixed, the low wing should tend to return to level. Initial bank angle should be appropriate to type;
however, it is recommended that it should not be less than 10° or that necessary to maintain the steady, straight sideslip with one-half rudder deflection, whichever occurs first. Roll control centering by the pilot should not be permitted during this evaluation. The intent of this testing is to evaluate the short-term response of the airplane; therefore, long-term effects, such as those due to spanwise fuel movement, need not be taken into account.
7.3.2.1.3 Steady, Straight Sideslips.
Steady, straight sideslips should be conducted in each direction to show that the aileron and rudder control movements and forces are substantially proportional to the angle of sideslip in a stable sense, and that the factor of 7-8
proportionality is within the limits found necessary for safe operation.
These tests should be conducted at progressively greater sideslip angles up to the sideslip angle appropriate to the operation of the airplane (see paragraph 7.3.1.3.2) or the sideslip angle associated with one-half of the available rudder control input (as limited by a rudder control force of 180 lbs), whichever is greater.
7.3.2.1.4 When determining the rudder and aileron control forces, the controls should be relaxed at each point to find the minimum force needed to maintain the control surface deflection. If excessive friction is present, the resulting low forces will indicate the airplane does not have acceptable stability characteristics.
7.3.2.1.5 Instead of conducting each of the separate qualitative tests described in paragraph 7.3.2.1, the applicant may use recorded quantitative data showing aileron and rudder control force and position versus sideslip (left and right) to the appropriate limits in the steady heading sideslips conducted to show compliance with § 25.177(c). If the control force and position versus sideslip indicates positive dihedral effect and positive directional stability, compliance with § 25.177(a) and (b) will have been successfully demonstrated.
7.3.2.2 Full Rudder Sideslips.
7.3.2.2.1 Rudder lock is that condition where the rudder over-balances aerodynamically and either deflects fully with no additional pilot input or does not tend to return to neutral when the pilot input is released. It is indicated by a reversal in the rudder control force as sideslip angle is increased. Full rudder sideslips are conducted to determine the rudder control forces and deflections out to sideslip angles associated with full rudder control input (or as limited by a rudder control force of 180 lbs) to investigate the potential for rudder lock and lack of directional stability.
7.3.2.2.2 To check for positive directional stability and for the absence of rudder lock, conduct steady heading sideslips at increasing sideslip angles until obtaining full rudder control input or a rudder control force of 180 lbs. If full lateral control is reached before reaching the rudder control limit or 180 lbs of rudder control force, continue the test to the rudder limiting condition in a non-steady heading sideslip maneuver.
7.3.2.3 Control Limits.
The control limits approved for the airplane should not be exceeded when conducting the flight tests required by § 25.177.
7.3.2.4 Flight Test Safety Concerns.
In planning for and conducting the full rudder sideslips, items relevant to flight test safety should be considered, including:
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AC 25-7D
Inadvertent stalls, Effects of sideslip on stall protection systems, Actuation of stick pusher, including the effects of sideslip on angle-of-attack sensor vanes, Heavy buffet, Exceeding flap loads or other structural limits, Extreme bank angles, Propulsion system behavior (e.g., propeller stress, fuel and oil supply, and inlet stability), Minimum altitude for recovery, Resulting roll rates when the aileron limit is exceeded, Position errors and effects on electronic or augmented flight control systems, especially when using the airplane’s production airspeed system, and Rudder loads, particularly those that may occur with dynamic rudder inputs.
7.4 Dynamic Stability—§ 25.181.
7.4.1 Explanation.
The dynamic stability tests described in this section should be conducted over the speed range of 1.13 VSR to VFE, VLE or VFC/MFC, as appropriate.
7.4.1.1 Dynamic Longitudinal Stability.
7.4.1.1.1 The short period oscillation is the first oscillation the pilot sees after disturbing the airplane from its trim condition with the pitch control (as opposed to the long period (phugoid)). Care should be taken that the control movement used to excite the motion is not too abrupt.
7.4.1.1.2 Heavily damped means that the oscillation has decreased to 1/10 the initial amplitude within approximately two cycles after completion of the control input.
7.4.1.1.3 Short period oscillations must be heavily damped, both with controls free and controls fixed.
7.4.1.2 Dynamic Lateral-Directional Stability.
The evaluation of the dynamic lateral-directional stability should include any combined lateral-directional oscillation (“Dutch roll”) occurring over the speed range appropriate to the airplane configuration. This oscillation 7-10
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.
8-2
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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