111 3 If static VMCA is near stall speed at the minimum practicable test weight, or if the thrust-to-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 VMCA using a sudden engine cut because of the dynamics of the rapid pitch down 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 VR to be not less than 1.05 VMCA, and there is some additional speed increase prior to lift off, a transport airplane is typically never airborne below approximately 1.08 VMCA.
Therefore, instead of using the dynamic method to define VMCA for these aircraft with high T/W or stall speed coincident with VMCA, 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 1.08 VSR or 1.1 VMCA (static), whichever is greater. During the entry to, and recovery from this maneuver, all the requirements of § 25.149(d) must be met.
4 For airplanes with rudder travel-limited VMCA’s that have increased power or 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 power or thrust has not increased more than 10 percent above the level at which dynamic VMCA had previously been demonstrated. (See Appendix 6).
(3) Minimum Control Speed - Ground (VMCG) - § 25.149(e).
(a) It must be demonstrated that, when the critical engine is suddenly made inoperative at VMCG 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 pre-engine-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.
(b) Work up tests may be conducted by abruptly retarding the critical engine to idle to determine the airplane asymmetric control characteristics and provide data from which an estimate of VMCG can be made. Due to the engine spindown characteristics with the critical engine retarded to idle, the speed will not, in general, be representative of the VMCG speed that would be obtained with a fuel cut. Therefore, the certification tests for VMCG should be conducted using fuel cuts. Starting from a speed comfortably above the estimated VMCG and with the maximum takeoff power or thrust level to be certified, several fuel cuts should be made at decreasing calibrated airspeeds to establish the minimum airspeed at which the lateral deviation is less than or equal to 30 ft. VMCG is determined for zero crosswind conditions.
However, in light crosswind test conditions the VMCG value determined should be that which is appropriate to the adverse crosswind or, at the applicant’s option, may be corrected to a zero crosswind value using runs made on reciprocal headings.
(c) During determination of VMCG, engine failure recognition should be provided by:
1 The pilot feeling a distinct change in the directional tracking characteristics of the airplane; or
2 The pilot seeing a directional divergence of the airplane with respect to the view outside the airplane.
(d) Directional 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. Pilot input to controls to supplement the rudder effectiveness should not be used. Care should also be taken not to inadvertently apply brake pressure during large rudder deflections, as this will invalidate the test data.
(e) VMCG testing should be conducted at the most critical weight in the range where VMCG may impact AFM V1 speeds.
(f) 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.
(g) VMCG testing should not be conducted on runways with excessive crowning
(i.e., cross-runway slope) unless the effects of such crowning are determined to be conservative.
(h) For airplanes with certification bases prior to Amendment 25-42, VMCG 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, smooth (i.e., not grooved or PFC) runway. The test(s) should include engine failure at or near a minimum VEF 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 (VMCL) - § 25.149(f).
(a) This section is intended to ensure that 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 usually consists of an engine failing after the power or thrust has been increased to perform a go-around from an all-engines-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 critical 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 VMCL to roll the airplane through an angle of 20 degrees, in the direction necessary to initiate a turn away from the
113 inoperative engine, in not more than five seconds when starting from a steady straight flight condition.
(b) 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)(b) and (c) for VMCA may be used to determine VMCL, 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.
(c) In accordance with § 25.149(f)(5) 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.
(d) At the option of the applicant, a one-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 VMCL(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 (VMCL-2) - § 25.149(g).
(a) For airplanes with three or more engines, VMCL-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.
(b) In accordance with § 25.149(g)(5) for propeller driven airplanes, the propeller 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.
(c) Conduct this test using the 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 one-engine-inoperative approach and landing configurations.
The following demonstrations should be conducted to determine VMCL-2:
1 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 VMCL-2 speed is established by the procedures presented in paragraphs 23b(2)(b) and
(c) for VMCA, except that flap and trim settings should be appropriate to the approach and landing
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.