114 configurations, the power or thrust on the operating engine(s) should be set to the go-around power or thrust setting, and compliance with all VMCL-2 requirements of § 25.149(g) and (h) must be demonstrated.
2 With power or thrust on the operating engines set to maintain a minus 3 degree glideslope, with one critical engine inoperative:
(aa) Set the airspeed at the value determined above in step (1) and, with zero bank angle, maintain a constant heading using trim to reduce the control force to zero. If full trim is insufficient to reduce the control force to zero, full trim should be used plus control deflection as required; and
(bb) Make the second critical engine inoperative and retard the remaining operating engine(s) to minimum available power or thrust without changing the directional trim.
The VMCL-2 determined in paragraph (1) is acceptable if constant heading can be maintained without exceeding a 5 degree bank angle and the limiting conditions of § 25.149(h).
3 Starting from a steady straight flight condition, demonstrate that sufficient lateral control is available at VMCL-2 to roll the airplane through an angle of 20 degrees in the direction necessary to initiate a turn away from the inoperative engines in not more than five seconds. This maneuver may be flown in a bank-to-bank roll through a wings level attitude.
(d) At the option of the applicant, a two-engines-inoperative landing minimum control speed, VMCL-2(2 out), may be determined in the conditions appropriate to an approach and landing with two engines having failed before the start of the approach. In this case, only those configurations recommended for use during an approach and landing with two engines inoperative need be considered. The propellers of the inoperative engines, if applicable, may be feathered throughout. The values of VMCL-2 or VMCL-2(2 out) should be used as guidance in determining the recommended procedures and speeds for a two-engines-inoperative approach and landing.
(6) Autofeather Effects. Where an autofeather or other drag limiting system is installed, and will be operative at approach power settings, its operation may be assumed in determining the propeller position achieved when the engine fails. Where automatic feathering is not available, the effects of subsequent movements of the engine and propeller controls should 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.
a. Explanation. Adequate trim capability should be provided for any flight condition that it is reasonable to assume will be maintained steadily for any appreciable time.
b. Procedures.
(1) The trim requirements specify the ranges of speed and the airplane configurations at which the airplane must be able to maintain trim.
(2) All weights, from the minimum in-flight weight to the maximum takeoff weight, should be considered. For airplanes with unpowered controls, the lowest weight may be more critical since this results in the lowest airspeed.
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.
(2) Section 25.175, Demonstration of Static Longitudinal Stability, specifically defines the flight conditions, airplane configurations, trim speed, test speed ranges, and power or thrust settings to be used in demonstrating compliance with the longitudinal stability requirements.
b. Procedures.
(1) Stabilized Method.
(a) 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 stabilizing at the trim speed, apply a light pull force and stabilize at a slower speed. Continue this process in increments, the size of the speed increment being dependent on the speed spread being investigated, until reaching the minimum speed for steady, unstalled flight or the minimum speed appropriate for the configuration. A continuous pull force should be used from the trim speed 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.
(b) Starting again at the trim speed, and with the airplane in trim, push forces should be gradually applied and gradually relaxed in the same manner as described in paragraph
(a), above.
(c) 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 power or thrust output. Consequently, a reasonably small altitude band, limited to +3,000 ft., should be used for the complete maneuver. If this altitude band is exceeded, regain the original trim altitude by changing the power or thrust setting and flap and gear position as necessary, 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 covers at least 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 defining and evaluating the data point, since it may be biased by this phugoid oscillation. Averaging these oscillating speeds at each data point is an acceptable method of eliminating this effect. Extremely smooth air improves the quality of the test data. In-bay and cross-bay wing fuel shift is another issue 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 evaluating the slope of the stick force versus speed.
(2) Acceleration-Deceleration Method.
(a) Trim at the desired airspeed and note the power or thrust setting. Without changing pitch trim, increase power or thrust to accelerate the airplane to the extreme speed of the desired data band. Using elevator control as needed, maintain approximately a constant altitude. Then, without changing pitch trim, quickly reset the power or thrust to the original power setting and allow the airplane to decelerate at a constant altitude back to the original trim speed. Obtain longitudinal static stability data during the deceleration to trim speed with the power and the pitch trim position the same as the original trim data point.
(b) Obtain data below the trim speed in a similar manner, by reducing power or thrust to decelerate the airplane to the lowest speed in the data band. Using elevator control as needed without changing pitch trim, maintain approximately a constant altitude. Then, without changing pitch trim, quickly reset the power to the original power setting, and record the data during the level flight acceleration back to trim speed. If, because of thrust/drag relationships, the airplane has difficulty returning to the trim conditions, small altitude changes within ± 2,000 feet can also be used to coax an airplane back to trim speed. Level flight is preferred, if possible.
Obtain speed and elevator stick force data approximately every 10 knots of speed change.
(3) 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).
118 Figure 26-1. Longitudinal Static Stability
(4) Examples of “local reversals” are given in Figure 26-2. Curves A and C depict a local gradient reversal within the required speed range. Even though it might be argued that the “average gradient” meets the one pound in six knots criterion, the gradient reversals would render these characteristics unacceptable. Curve B depicts a situation in which the gradient reverses, but only outside the required speed range. In addition, Curve B demonstrates a situation in which the local gradient does not always meet the required one pound in six knots, even though the average gradient does.
119 Figure 26-2. Local Reversal
Pull Push Speed A Unsat.
B Sat.* Required Average Gradient Free Return Speed Long Control Force B Sat.* Required Speed Range C Unsat.
A Unsat.
C Unsat.
Trim Speed Free Return Speed
* zero slope at end of speed range
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. Prior to Amendment 25-72, a separate demonstration of positive static directional stability was required by § 25.177(a) for any landing gear and flap position and symmetrical power or thrust condition at speeds from 1.13 VSR1 up to VFE, VLE, or VFC/MFC, as appropriate for the airplane configuration.
(2) Static Lateral Stability. Positive static lateral stability is defined as the tendency to raise the low wing in a sideslip with hands off the roll controls. Prior to Amendment 25-72, a separate demonstration was required by § 25.177(b) to show that static lateral stability was not negative in any landing gear and flap position and symmetrical power or thrust 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 was 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. Also, the factor of proportionality must lie between limits found necessary for safe operation. Section 25.177(c) also states that 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. This means that if using one-half of the available rudder control input takes less than 180 pounds of force, then compliance must be based on using one-half of the available rudder control input. If application of 180 pounds of rudder control force results in using less than onehalf of the available rudder control input, then compliance must be based on applying 180 pounds of rudder control force. 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 or thrust 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 = arc sin (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.
(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 V2+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 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).
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 the roll controls. 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 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 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 long-term 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.
(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.
(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 (i.e., 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 5 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.