109 be considered, including fully closing the power lever of the failed engine in conjunction with maintaining the go-around power setting on the operating engine(s).
Section 4. Trim
24. Trim - § 25.161 [Reserved]
Section 5. Stability
25. General - § 25.171. [Reserved]
26. Static Longitudinal Stability and Demonstration of Static Longitudinal Stability - §§ 25.173 and 25.175.
a. Explanation .
(1) Section 25.173 - Static Longitudinal Stability .
(a) Compliance with the general requirements of § 25.173 is determined from a demonstration of static longitudinal stability under the conditions specified in § 25.175.
(b) The requirement is to have a pull force to obtain and maintain speeds lower than trim speed, and a push force to obtain and maintain speeds higher than trim speed. There may be no force reversal at any speed that can be obtained, except lower than the minimum for steady, unstalled flight or, higher than the landing gear or wing flap operating limit speed or VFC/MFC, whichever is appropriate for the test configuration. The required trim speeds are specified in § 25.175.
(c) When the control force is slowly released from any speed within the required test speed range, the airspeed must return to within 10 percent of the original trim speed in the climb, approach, and landing conditions, and return to within 7.5 percent of the trim speed in the cruising condition specified in § 25.175 (free return).
(d) The average gradient of the stick force versus speed curves for each test configuration may not be less than one pound for each 6 knots for the appropriate speed ranges specified in § 25.175. Therefore, after each curve is drawn, draw a straight line from the intersection of the curve and the required maximum speed to the trim point. Then draw a straight line from the intersection of the curve and the required minimum speed to the trim point.
The slope of these lines must be at least one pound for each 6 knots. The local slope of the curve must remain stable for this range.
NOTE: Due to different design features of individual airplanes, there may be cases where the local slope gradient deviates somewhat from that specified by
110 § 25.173. When this occurs, an investigation should be performed to determine if a finding of equivalent safety can be made based on pilot evaluation.
(2) Section 25.175, Demonstration of Static Longitudinal Stability , specifically defines the flight conditions, airplane configurations, trim speed, test speed ranges, and thrust settings to be used in demonstrating compliance with the longitudinal stability requirements.
b. Procedures .
(1) For the demonstration of static longitudinal stability, the airplane should be trimmed in smooth air at the conditions required by the regulation. Aft c.g. loadings are generally most critical. After obtaining trim speed, apply a light pull force and stabilize at a slower speed. Continue this process in acceptable increments, depending on the speed spread being investigated, until reaching the minimum speed for steady, unstalled flight or the minimum required as appropriate for the configuration. A continuous pull force from the trim speed is required on each series of test points to eliminate hysteresis effects. At the end of the required speed range, the force should be gradually relaxed to allow the airplane to return slowly toward the trim speed and zero stick force. Depending on the amount of friction in the control system, the eventual speed at which the airplane stabilizes will normally be less than the original trim speed. The new speed, called the free return speed, must meet the requirements of § 25.173.
(2) Starting again at the trim speed , push forces should be gradually applied and gradually relaxed in the same manner as described in paragraph (1), above.
(3) The above techniques result in several problems in practice. One effect of changing airspeed is a change of altitude, with a corresponding change in Mach number and thrust or power output. Consequently, a reasonably small altitude band, limited to +3,000 ft., should be used for the complete maneuver. If the altitude band is exceeded, regain the original trim altitude by changing the power setting and flap and gear position, but without changing the trim setting. Then continue the push or pull maneuver in the original configuration. Testing somewhat beyond the required speed limits in each direction assures that the resulting data at least extends to the required speed ranges. It will also be noted in testing that while holding force constant at each data point, the airspeed and instantaneous vertical speed vary in a cyclic manner. This is due to the long period (phugoid) oscillation. Care should be exercised in eventing the data point, since it may be biased by this phugoid oscillation. Averaging these oscillating speeds at each
Change 1 111 data point is an acceptable method of eliminating this effect. Extremely smooth air improves the quality of the test data, with critical areas requiring the best of smooth air. In-bay and cross-bay wing fuel shift is another problem experienced in some airplanes. In-bay fuel shift occurs rapidly with pitch angle; therefore, consideration should be given to testing with fuel loadings that provide the maximum shift since it is generally destabilizing. Slower, cross-bay fuel shift, or burn from an aft tank, can influence the measured stability but usually only because of the time required to obtain the data points. This testing induced instability should be removed from the data before evaluation of the required slope.
(4) The resulting pilot longitudinal force test points should be plotted versus airspeed to show the positive stable gradient of static longitudinal stability and that there are no “local” reversals in the stick force vs. airspeed relationship over the range of airspeeds tested. This plot should also show the initial trim point and the two return-to-trim points to evaluate the return-totrim characteristics (see figure 26-1).
Figure 26-1. Longitudinal Static Stability
27. Static Directional and Lateral Stability - § 25.177.
a. Explanation .
(l) Static Directional Stability . Positive static directional stability is defined as the tendency to recover from a skid with the rudder free. Positive static directional stability is required by § 25.177(a) for any landing gear and flap position and symmetrical power condition at speeds from 1.13 VSR1 up to VFE, VLE, or VFC/MFC, as appropriate for the airplane configuration.
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(2) Static Lateral Stability . Positive static lateral stability is defined as the tendency to raise the low wing in a sideslip with the aileron controls free. Section 25.177(b) requires that static lateral stability not be negative in any landing gear and flap position and symmetrical power condition at speeds from 1.13 VSR1 to VFE, VLE, or VMO/MMO, as appropriate for the airplane configuration. At speeds from VMO/MMO to VFC/MFC, negative static lateral stability is permitted by § 25.177(b), if the divergence is:
(a) Gradual;
(b) Easily recognizable by the pilot; and
(c) Easily controllable by the pilot.
(3) Steady Straight Sideslips .
(a) 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 and that this factor of proportionality must lie between limits necessary for safe operation. Section 25.177(c) also states that the range of sideslip angles evaluated must 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 pounds. By cross-reference to § 25.177(a), § 25.177(c) requires that these steady, straight sideslip criteria must be met for all landing gear and flap positions and symmetrical power conditions at speeds from 1.13 VSR1 to VFE, VLE, or VFC/MFC, as appropriate for the configuration.
(b) Experience has shown that an acceptable method for determining the appropriate sideslip angle for the operation of a transport category airplane is provided by the following equation:
ß = arcsin (30/V)
where ß = Sideslip angle, and V = Airspeed (KTAS)
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 degrees 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 degrees may not be appropriate for airplanes with low approach speeds or high crosswind capability.
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(c) A lower sideslip angle than that provided in paragraph 27a(3)(b) may be used if it is substantiated that the lower value conservatively 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.
(d) 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 pounds, 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 one-half 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.
(4) Full Rudder Sideslips.
(a) 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 pounds 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.
(b) 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 pounds 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.
(c) Section 25.177(d) states that the criteria listed in paragraph 27a(4)(a) 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:
1 For takeoff configurations, speeds from V 2+xx (airspeed approved for allengines-operating initial climb) to VFE or VLE, as appropriate, and takeoff power/thrust;
2 For flaps up configurations, speeds from 1.23 V SR to VLE or VMO/MMO, as appropriate, and power from idle to maximum continuous power/thrust;
Change 1 114 3 For approach configurations, speeds from 1.23 V SR to VFE or VLE, as appropriate, and power from idle to go-around power/thrust; and
4 For landing configurations, speeds from V REF-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).
b. Procedures . The test conditions should include each flap and landing gear configuration as described in paragraphs 27a(1) through 27a(4) at an altitude appropriate to each configuration.
(1) Basic Tests for Static Directional and Lateral Stability .
(a) 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 ailerons. When the rudder is released, the airplane should tend to return to straight flight.
(b) 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 aileron. When the ailerons 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 degrees 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 longterm effects, such as those due to spanwise fuel movement, need not be taken into account.
(2) 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 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 27a(3)(b)) or the sideslip angle associated with one-half of the available rudder control input (as limited by a rudder control force of 180 pounds), whichever is greater.
(a) 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.
(b) In lieu of conducting each of the separate qualitative tests described in paragraph 27b(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.
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(3) Full Rudder Sideslips .
(a) 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 pounds) to investigate the potential for rudder lock and lack of directional stability.
(b) 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 pounds. If full lateral control is reached before reaching the rudder control limit or 180 pounds of rudder control force, continue the test to the rudder limiting condition in a non-steady heading sideslip maneuver.
(4) Control Limits. The control limits approved for the airplane should not be exceeded when conducting the flight tests required by § 25.177.
(5) Flight Test Safety Concerns. In planning for and conducting the full rudder sideslips, items relevant to flight test safety should be considered, including:
(a) Inadvertent stalls,
(b) Effects of sideslip on stall protection systems,
(c) Actuation of stick pusher, including the effects of sideslip on angle-of-attack sensor vanes,
(d) Heavy buffet,
(e) Exceeding flap loads or other structural limits,
(f) Extreme bank angles,
(g) Propulsion system behavior (e.g., propeller stress, fuel and oil supply, and inlet stability),
(h) Minimum altitude for recovery,
(i) Resulting roll rates when the aileron limit is exceeded, and
(j) Position errors and effects on electronic or augmented flight control systems, especially when using the airplane’s production airspeed system.
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(k) Rudder loads, particularly those that may occur with dynamic rudder inputs.
28. Dynamic Stability - § 25.181.
a. Explanation .
(1) 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.
(2) Dynamic Longitudinal Stability .
(a) 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.
(b) Heavily damped means that the oscillation has decreased to 1/10 the initial amplitude within approximately two cycles after completion of the control input.
(c) Short period oscillations must be heavily damped, both with controls free and controls fixed.
(3) Dynamic Lateral-Directional Stability . The evaluation of the dynamic lateraldirectional stability should include any combined lateral-directional oscillation (“Dutch roll”) occurring over the speed range appropriate to the airplane configuration. This oscillation must be positively damped with controls free and must be controllable with normal use of the primary controls without requiring exceptional piloting skill.
b. Procedures .
(1) Dynamic Longitudinal Stability .
(a) 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.
(b) Dynamic longitudinal stability should be checked at a sufficient number of points in each configuration to assure compliance at all operational speeds.
(2) Dynamic Lateral-Directional Stability .
(a) 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 (“Dutch roll”). The control input should be in phase with the airplane’s oscillatory response.
(b) 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.
(3) Airplanes Equipped with Stability Augmentation Systems (SAS) . 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:
(a) 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.
(b) 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 § 25.181(a) or (b) may still be demonstrated to a reduced flight envelope with no SAS operating as described in paragraph 28b(3)(a), above.
(c) 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 6 to 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.
118 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.
Section 6. Stalls
29. Stall Testing.
a. The applicable Code of Federal Regulations (CFR) are as follows:
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
b. Explanation .
(1) The purpose of stall testing is threefold:
(a) To define the reference stall speeds and how they vary with weight, altitude, and airplane configuration.
(b) To demonstrate that handling qualities are adequate to allow a safe recovery from the highest angle-of-attack attainable in normal flight (stall characteristics).
(c) 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.
(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.
c. Stall Demonstration - § 25.201.
(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.
(a) 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 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)).
(b) 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.
(c) 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.
(d) The activation point of a stall identification device that provides one of the characteristics listed above. Paragraph 228 of this AC presents guidance material for demonstrating compliance with the regulatory requirements of part 25 for stall identification systems.
(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, center of gravity, 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.
d. Stall Speeds .
(1) Background .
(a) 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.
(b) Prior to amendment 25-108, the stall speed defined in § 25.103 was the minimum speed attained in the stalling maneuver. For many high speed swept wing transport category airplanes, the resulting stall speed often occurs at a load factor normal to the flight path considerably less than one, which leads to inconsistent and unrepeatable reference stall speeds.
Pilot technique can also significantly influence the rate and magnitude of any nose down pitch occurring at the stall, thereby contributing to inconsistencies in the determination of the minimum speed obtained in the stalling maneuver. Since part 25 defines operating speeds as multiples of the stall speed, the resulting operating speed margins to stall may not be representative of the actual lift margin available (i.e., the margin to the speed at which wing lift alone can support the weight of the airplane in 1-g flight). To ensure that operating speed and maneuvering margins are directly related to wing lift margin, amendment 25-108 redefined the reference stall speed as the 1-g stall speed, which is the speed at which the wing is generating maximum usable lift in a 1-g flight condition.
(c) Since the 1-g stall speed is generally higher than the minimum speed obtained in the stalling maneuver, retaining the existing multiplying factors for determining the minimum operating speeds would have resulted in higher minimum operating speeds. To retain approximately equivalent operating speeds that have been found acceptable, the multiplying factors applied to the reference stall speed were reduced by approximately six percent.
(d) Although the reduced multiplying factors were intended to result in roughly equivalent operating speeds, there is one class of airplanes for which a significantly lower operating speed would be obtained. Airplanes equipped with a device that abruptly pushes the