Prior to Amendment 25-42 to § 25.149, the regulation required that rudder control forces must not exceed 180 lbs.
With the adoption of Amendment 25-42, rudder control forces became limited to 150 lbs. The relationships between VEF, V1, and VMCG are discussed in paragraph 10, Takeoff and Takeoff Speeds, and paragraph 11, Accelerate-Stop Distance.
b. Procedures .
(1) General.
(i) Prior to beginning the minimum control speed tests, an evaluation should be conducted to determine which engine's failure will result in the largest asymmetric yawing moment (i.e., the "critical" engine).
This is typically done by setting one outboard engine to maximum thrust, setting the corresponding opposite engine at idle, and decelerating with wings level until full rudder is required. By alternating power on/power off from left to right, the critical engine can be defined as the idle engine that requires the highest minimum speed to maintain a constant heading with full rudder deflection.
(ii) For propeller-driven airplanes, V MCA, VMCG, and VMCL (and VMCL-2, as applicable) should be determined by rendering the critical engine(s) inoperative and allowing the propeller to attain the position it automatically assumes. However, for some engine/propeller installations, a more critical drag condition could be produced as the result of a failure mode that results in a partial power condition that does not actuate the automatic propeller drag reduction system (e.g., autofeather system). One example is a turbopropeller installation that can have a fuel control failure, which causes the engine to go to flight idle, resulting in a higher asymmetric yawing moment than would result from an inoperative engine. In such cases, the minimum control speed tests must be conducted using the most critical failure mode. For propeller-driven airplanes where V MCA is based on operation of a propeller drag reduction system, V MCA should also be defined with the critical engine at idle to address the training situation where engine failure is simulated by retarding the critical engine to idle. If V MCA at idle is more than one knot greater than for the engine failure with an operating drag reduction system, the idle engine V MCA should be included in the Normal Procedures section of the Airplane Flight Manual (AFM) as advisory information to maintain the level of safety in the aforementioned training situation.
(iii) Airplane Flight Manual values of V MCA, VMCG, and VMCL (and VMCL-2, as applicable) should be based on the maximum net thrust reasonably expected for a production engine. These speeds should not be based on specification thrust, since this value represents the minimum thrust guaranteed by the engine manufacturer, and the resulting minimum control speeds will not be representative of what could be achieved in operation. The maximum thrust used for scheduled AFM minimum control speeds should represent the high side of the tolerance band, but may be determined by analysis instead of tests.
(iv) When determining VMCA, VMCL, and VMCL,-2, consideration should be given to the adverse effect of maximum approved lateral fuel imbalance on lateral control availability. This is especially of concern if tests or analysis show that the lateral control available is the determining factor of a particular V MC.
(2) Minimum Control Speeds - Air (VMCA).
(i) To comply with the V MCA requirements, the following two conditions must be satisfied:
(Separate tests are usually conducted to show compliance with these two requirements.)
(A) The stabilized (static) cond ition where constant heading is maintained without exceeding a 5-degree bank angle, and (B) The dynamic condition in which control is maintained without exceeding a heading change of 20 degrees.
(ii) Static Test Procedure and Required Data.
(A) To determine VMCA, use the configuration specified in § 25.149, except that V MCA is normally determined at minimum weight in order to minimize the stall speed and because static V MCA decreases
with increased weight if a 5-degree bank angle is used. The requirement of § 25.149(c) that V MCA not exceed 1.2VS is based on VS at maximum sea level takeoff weight. With the critical engine inoperative, the corresponding opposite engine should be adjusted to maximum takeoff power/thrust, and the airspeed decreased until heading can just be maintained with full rudder and no more than a 5degree bank into the operating engine. For airplanes with more than two engines, the inboard engine(s) may be set to any thrust necessary to assist in developing the desired level of asymmetric thrust, or to achieve the desired flightpath angle (normally level flight).
(B) If the maximum asymmetric thrust that is permitted by the AFM operating limitations was maintained at the test day V MCA, and the rudder pedal force did not exceed the limit specified in § 25.149(d), the resulting speed may be used as the single value of V MCA for the airplane. If, at the option of the applicant, the AFM value of VMCA is to vary with pressure altitude and temperature, the test day minimum control speed and the corresponding thrust should be used to calculate an equivalent yawing moment coefficient (C N). This CN value may then be used to calculate V MCA as a function of takeoff thrust, thus permitting V MCA to be scheduled as a function of pressure altitude and temperature for takeoff data expansion and presentation in the AFM (See Appendix 3 for further discussion of V MCA correction).
(C) If maximum allowable takeoff thrust could not be developed at the flight test conditions, but maximum rudder deflection was achieved, then the V MCA value corresponding to sea level standard day maximum asymmetric thrust may be calculated from the C N attained at the test value of V MCA. Extrapolation using this constant C N method is limited to 5 percent of the test day asymmetric thrust, and is only permitted if the rudder pedal force at the test day V MCA was not more than 95 percent of the limit value specified in § 25.149(d). For extrapolation beyond 5 percent thrust, a more rigorous analysis is required, which includes all the applicable stability and control terms (See Appendix 3 for further discussion of V MCA correction).
(D) If VMCA could not be achieved due to stall buffet, or excessive rudder pedal force, a parametric investigation should be undertaken to determine whether V MCA is limited by stall speed, maximum rudder deflection, or maximum allowable rudder pedal force (See Appendix 4).
(iii) Dynamic Test Procedures and Required Data.
(A) After the static VMCA tests have been completed, dynamic engine cuts should be evaluated at a series of decreasing airspeeds to show that sudden engine failure at any speed down to the static V MCA value meets the requirements of § 25.149. The dynamic V MCA test is conducted by applying the maximum approved power/thrust to all outboard engines, stabilizing at the test airspeed, and then cutting fuel to the critical engine. It must be possible to recover to a constant heading, without deviating more than 20 degrees from the original heading, while maintaining the test airspeed, without reducing power/thrust on the operating engine(s), and without exceeding the rudder pedal force limit of § 25.149(d). If the dynamic tests result in a V MCA greater than the static value, the increment between the static and dynamic V MCA at the same altitude should be added to the sea level extrapolated value. If the dynamic value is less than the static value, the static V MCA must be used for the AFM data expansion.
(B) If static VMCA is near stall speed at the minimum practicable test weight, or if the thrustto-weight ratio (T/W) results in a trimmed pitch attitude of more than 20 degrees, it is not feasible to attempt to accurately define a quantitative value of V MCA using a sudden engine cut because of the dynamics of the rapid pitchdown maneuver required, and the hazard associated with a potential spin entry. Additionally, an extreme nose up attitude followed by an engine cut is not representative of an operational takeoff engine failure. Since §
25.107(e)(1)(ii) requires V R to be not less than 1.05V MCA, and there is some additional speed increase prior to lift off, a transport airplane is typically never airborne below approximately 1.08V MCA. Therefore, instead of using the dynamic method to define V MCA for these aircraft with high T/W or stall speed coincident with V MCA, it is more appropriate for a dynamic engine cut to be evaluated only for acceptable controllability, and at a more representative speed. For these airplanes, a dynamic engine cut should be evaluated at an airspeed of either 1.1V S or 1.1VMCA (static), whichever is greater. During the entry to, and recovery from this maneuver, all the requirements of § 25.149(d) must be met.
(C) For airplanes with rudder travel-limited VMCA’s that have increased thrust engines installed, with no changes to the airframe’s geometric layout or dimensions, it may not be necessary to conduct dynamic VMCA flight testing if the thrust has not increased more than 10 percent above the level at which dynamic VMCA had previously been demonstrated (see Appendix 3 of this AC).
(3) Minimum Control Speed - Ground (VMCG) - § 25.149(e).
(i) It must be demonstrated that, when the critical engine is suddenly made inoperative at V MCG during the takeoff ground roll, the airplane is safely controllable if the takeoff is continued. During the demonstration, the airplane must not deviate more than 30 ft. (25 ft. prior to Amendment 25-42) from the preengine-cut projected ground track. The critical engine)for ground minimum control speed testing should be determined during the takeoff ground run using techniques similar to these described in paragraph 23b(1). If there is a significant difference in left and right rudder deflection, the loss of asymmetric propeller disc loading, due to near zero angle of attack during the takeoff roll, could result in the critical engine being on the opposite side of the airplane relative to the airborne minimum control speed tests.
(ii) Tests may be conducted by abruptly retarding the c ritical engine to idle to establish the minimum value of VMCG. At least one fuel cut should be made at each maximum asymmetric thrust level desired to be certificated to investigate the more rapid thrust decay associated with this type of engine failure. At the applicant's option, to account for crosswind test conditions, the runs may be made on reciprocal headings, or an analytical correction may be applied to determine the zero crosswind value of V MCG.
(iii) During determination of V MCG, engine failure recognition should be provided by:
(A) The pilot feeling a distinct change in the directional tracking characteristics of the airplane; or (B) The pilot seeing a directional divergence of the airplane with respect to the view outside the airplane.
(iv) Control of the airplane should be accomplished by use of the rudder only. All other controls, such as ailerons and spoilers, should only be used to correct any alterations in the airplane attitude and to maintain a wings level condition. Use of those controls to supplement the rudder effectiveness should not be allowed. Care should also be taken not to inadvertently apply brake pressure during large rudder deflections, as this will invalidate the test data.
(v) VMCG testing should be conducted at the heaviest weight where V MCG may impact the AFM V 1 schedule.
(vi) VMCG testing should be conducted at aft c.g. and with the nose wheel free to caster, to minimize the stabilizing effect of the nose gear. If the nose wheel does not caster freely, the test may be conducted with enough nose up elevator applied to lift the nose wheel off the runway.
(vii) For airplanes with certification bases prior to Amendment 25-42, V MCG values may be demonstrated with nose wheel rudder pedal steering operative for dispatch on wet runways. The test should be conducted on an actual wet runway. The test(s) should include engine failure at or near a minimum V EF associated with minimum VR to demonstrate adequate controllability during rotation, liftoff, and the initial climbout. The VMCG values obtained by this method are applicable for wet or dry runways only, not for icy runways.
(4) Minimum Control Speed During Approach and Landing V MCL - § 25.149(f).
(i) This section is intended to ensure t hat the airplane is safely controllable following an engine failure during an all-engines-operating approach and landing. From a controllability standpoint, the most critical case consists of an engine failing after the power or thrust has been increased to perform a go-around from an allengines-operating approach. Section 25.149(f) requires the minimum control speed to be determined that allows a pilot of average skill and strength to retain control of the airplane after the engine becomes inoperative and to maintain straight flight with less than 5 degrees of bank angle. Section 25.149(h) requires that sufficient lateral control be available at V MCL to roll the airplane through an angle of 20 degrees, in the direction necessary to initiate a turn away from the inoperative engine, in not more than five seconds when starting from a steady straight flight condition.
(ii) Conduct this test using the most critical of the all-engines-operating approach and landing configurations or, at the option of the applicant, each of the all-engines-operating approach and landing configurations. The procedures given in paragraphs 23b(2)(ii) and (iii) for V MCA may be used to determine V MCL, except that flap and trim settings should be appropriate to the approach and landing 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 requirements of §§ 25.149(f) and (h) must be demonstrated.
(iii) For propeller driven airplanes , the propeller must be in the position it achieves without pilot action following engine failure, assuming the engine fails while at the power or thrust necessary to maintain a three degree approach path angle.
(iv) At the option of the applicant, a o ne-engine-inoperative landing minimum control speed, VMCL(1 out) , may be determined in the conditions appropriate to an approach and landing with one engine having failed before the start of the approach. In this case, only those configurations recommended for use during an approach and landing with one engine inoperative need be considered. The propeller of the inoperative engine, if applicable, may be feathered throughout. The resulting value of V MCL(1 out) may be used in determining the recommended procedures and speeds for a one-engine-inoperative approach and landing.
(5) Minimum Control Speed with Two Inoperative Engines During Approach and Landing (V MCL-2) - §
25.149(g).
(i) For airplanes with three or more engines, V MCL-2 is the minimum speed for maintaining safe control during the power or thrust changes that are likely to be made following the failure of a second critical engine during an approach initiated with one engine inoperative.
(ii) For propeller driven airplanes, the propelle r of the engine that is inoperative at the beginning of the approach may be in the feathered position. The propeller of the more critical engine must be in the position it automatically assumes following engine failure.
(iii) Conduct this test using t he most critical approved one-engine-inoperative approach or landing configuration (usually the minimum flap deflection), or at the option of the applicant, each of the approved oneengine-inoperative approach and landing configurations. The following demonstrations are required to determine VMCL-2:
(A) With the power or thrust on the operating engines set to maintain a minus 3 degree glideslope with one critical engine inoperative, the second critical engine is made inoperative and the remaining operating engine(s) are advanced to the go-around power or thrust setting. The V MCL-2 speed is established by the procedures presented in paragraphs 23b(2)(ii) and (iii) for VMCA, except that flap and trim settings should be appropriate to the approach and landing 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 V MCL-2 requirements of §§ 25.149(g) and (h) must be demonstrated.
(B) With power on the operating eng ines set to maintain a minus 3 degree glideslope, with one critical engine inoperative:
(1) Set the airspeed at the value determined above in step (A) 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
(2) Make the second critical engine inoperative and retard the remaining operating engine(s) to minimum available power without changing the directional trim. The V MCL-2 determined in paragraph (A) is acceptable if constant heading can be maintained without exceeding a 5 degree bank angle and the limiting conditions of § 25.149(h).
(C) Starting from a stea dy 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.
(iv) 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 V MCL-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 engines.
Section 4. TRIM
24. [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.
(i) Compliance with the general requirements of § 25.173 is determined from a demonstration of static longitudinal stability under the conditions specified in § 25.175.
(ii) 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.
(iii) 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).
(iv) 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 § 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 slowly return 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 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. 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 condition at speeds from 1.2VS1 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 the aileron controls free. 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 condition at speeds from
1.2VS1 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), providing the divergence was:
(i) Gradual;
(ii) Easily recognizable by the pilot; and
(iii) Easily controllable by the pilot.
(3) Steady Straight Sideslips.
(i) Section 25.177(c) requires, in steady, straight sideslips to 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 must lie between limits necessary for safe operation. Experience has shown that 15 degrees is an appropriate sideslip angle for transport category airplanes. Section 25.177(c) states that compliance with these steady, straight sideslip criteria "must be demonstrated for all landing gear and flap positions and symmetrical power conditions at speeds from 1.2VS1 to VFE, VLE, or VFC/MFC, as appropriate."
(ii) At sideslip angles greater than those appropriate for normal operation of the airplane, up to the sideslip angle at which full rudder travel is used or a rudder pedal force of 180 pounds is obtained, the rudder pedal forces may not reverse and increased rudder deflection must produce increased angles of sideslip. The goals of this higher-than-normal sideslip angle test are to show that at full rudder, or at maximum expected pilot effort: (1) the rudder pedal force does not reverse, and (2) increased rudder deflection produces increased angles of sideslip, thus demonstrating adequate directional stability for maneuvers involving large, dynamic rudder inputs, such as engine failure during takeoff just after V1. Consequently, it is permissible to continue this test beyond full lateral control input.
(iii) Section 25.177(d) (post-Amendment 25-72) stipulates that increased rudder deflection must continue to increase the angle of sideslip at speeds between VMO/MMO and VFC/MFC; but the dihedral effect, as exhibited by aileron deflection opposite the corresponding rudder input, may become negative in this speed range, provided the divergence is:
(A) Gradual;
(B) Easily recognized by the pilot; and (C) Easily controlled by the pilot.
b. Procedures. The test conditions should include each flap and landing gear configuration as described in paragraphs 27a(1), (2) or (3), at both low altitude and the maximum altitude appropriate to each configuration.
(1) Basic Test.
(i) 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.
(ii) 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.
(iii) Absence of Rudder Lock.
(A) Rudder lock is that condition where the rudder over-balances aerodynamically and deflects fully with no additional pilot input.
(B) To check for the absence of rudder lock with a particular configuration and trim speed, conduct steady, straight sideslips (unaccelerated forward slips) while maintaining a desired airplane track.
(C) Aileron and rudder control movements must remain in harmony, and forces must increase in proportion with sideslip angle, up to the limits found necessary for safe operation. At sideslip angles greater than those appropriate for normal operation of the airplane, up to the sideslip angle at which full rudder travel is used or a rudder pedal force of 180 pounds is obtained, the rudder pedal forces may not reverse; and increased rudder pedal deflection must produce increased angles of sideslip. The purpose of the second part of this test, at greater than normal sideslip angles, is to show that no rudder lock tendencies exist up to the rudder deflection limits as may be encountered during a dynamic rudder input to maintain directional control.
(2) Alternative Test. In lieu of conducting each of the qualitative tests described in paragraph 27b(1), the applicant may obtain recorded quantitative data showing aileron and rudder force and position versus sideslip (left and right) to the appropriate limits in constant heading sideslips.
(i) When determining the rudder and aileron forces, the controls should be relaxed at each point to find the minimum force required to hold the control surface deflection. If excessive friction is present, the resulting low forces will indicate the airplane does not have acceptable stability characteristics.
(ii) If the force and position versus sideslip indicates positive dihedral effect, positive directional stability and no rudder lock, compliance with § 25.177 has been successfully demonstrated.
(3) The control limits approved for the airplane should not be exceeded when conducting the flight tests required by § 25.177.
28. DYNAMIC STABILITY - § 25.181.
a. Explanation.
(1) The dynamic stability tests described in this section should be conducted over the speed range of 1.2VS to VFE, VLE or VFC/MFC, as appropriate.
(2) Dynamic Longitudinal Stability.
(i) 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.
(ii) Heavily damped means that the oscillation has decreased to 1/10 the initial amplitude within approximately two cycles after completion of the control input.
(iii) 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.
(i) 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.
(ii) Dynamic longitudinal stability must be checked at a sufficient number of points in each configuration to assure compliance at all operational speeds.
(2) Dynamic Lateral-Directional Stability.
(i) 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.
(ii) Dynamic lateral-directional stability must 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:
(i) If the airplane is equipped with only one SAS (i.e., a single strand system), 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.
(ii) 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,