(A) A stick force of at least 50 pounds to reach limit strength in steady maneuver or wind-up turns is considered acceptable to demonstrate adequate minimum force at limit strength in the absence of deterrent buffeting. If heavy buffeting occurs before the limit strength condition is reached, a somewhat lower stick force at limit strength may be acceptable. The acceptability of a stick force of less than 50 pounds at the limit strength condition will depend upon the intensity of the buffet, the adequacy of the warning margin (i.e., the load factor increment between the heavy buffet and the limit strength condition), and the stick force characteristics. In determining the limit strength condition for each critical component, the contribution of buffet loads to the overall maneuvering loads should be taken into account.
(B) This minimum stick force applies in the enroute configuration with the air plane trimmed for straight flight, at all speeds above the minimum speed at which the limit strength condition can be achieved without stalling. No minimum stick force is specified for other configurations, but the requirements of § 25.143(f) are applicable in these conditions.
(iii) Stick Force Characteristics.
(A) At all points within the buffet onset boundary determined in accordance with § 25.251(e), but not including speeds above V FC/MFC, the stick force should increase progressively with increasing load factor.
Any reduction in stick force gradient with change of load factor should not be so large or abrupt as to impair significantly the ability of the pilot to maintain control over the load factor and pitch attitude of the airplane.
(B) Beyond the buffet onset boundary, hazardous stick force characteristics should not be encountered within the permitted maneuvering envelope as limited by paragraph 20e(2)(iii)(C). It should be possible, by use of the primary longitudinal control alone, to rapidly pitch the airplane nose down so as to regain the initial trimmed conditions. The stick force characteristics demonstrated should comply with the following:
(1) For normal acceleration increments of up to 0.3g beyond buffet onset, where these can be achieved, local reversal of the stick force gradient may be acceptable, provided that any tendency to pitch up is mild and easily controllable.
(2) For normal acceleration increments of more than 0.3g beyond buffet onset, where these can be achieved, more marked reversals of the stick force gradient may be acceptable. It should be possible to contain any pitch-up tendency of the airplane within the allowable maneuvering limits, without applying push forces to the control column and without making a large and rapid forward movement of the control column.
(C) In flight tests to satisfy paragraphs 20f(2)(iii)(A) and (B), the load factor should be increased until either:
(1) The level of buffet becomes sufficient to provide a strong and effective deterrent to any further increase of the load factor; or
(2) Further increase of the load factor requires a stick force in excess of 150 pounds (or in excess of 100 pounds when beyond the buffet onset boundary) or is impossible because of the limitations of the control system; or
(3) The positive limit maneuvering load factor established in compliance with §
25.337(b) is achieved.
(iv) Negative Load Factors. It is not intended that a detailed flight test assessment of the maneuvering characteristics under negative load factors should necessarily be made throughout the specified range of conditions. An assessment of the characteristics in the normal flight envelope involving normal accelerations from 1g to zero g will normally be sufficient. Stick forces should also be assessed during other required flight testing involving negative load factors. Where these assessments reveal stick force gradients that are unusually low, or that are subject to significant variation, a more detailed assessment, in the most critical of the specified conditions, will be required. This may be based on calculations, provided they are supported by adequate flight test or wind tunnel data.
21. LONGITUDINAL CONTROL - § 25.145 .
a. Explanation.
(1) Section 25.145(a) requires that there be adequate longitudinal control to promptly pitch the airplane nose down from at, or near the stall to return to the original trim speed. The intent is to insure sufficient pitch control if inadvertently slowed to the point of stall. Though § 25.145(a)(4) requires testing at power settings between idle and maximum continuous, it is not intended that airplanes should be taken to the point of stall with high power settings; this situation is addressed by basing the minimum test speed on stall warning plus a pilot recognition time, as described in paragraph 21(b)(1)(ii).
* (2) Section 25.145(b) requires changes to be made in flap position, power, and speed without undue effort when retrimming is impractical. The purpose is to ensure that any of these changes are possible assuming that the pilot finds it necessary to devote at least one hand to the initiation of the desired operation without being overpowered by the primary airplane controls. The objective is to show that an excessive change in trim does not result from the application or removal of power or the extension or retraction of wing flaps. The presence of gated positions on the flap control does not affect the requirement to demonstrate full flap extensions and retractions without changing the trim control. Compliance with § 25.145(b) also requires that the relation of control force to speed be such that reasonable changes in speed may be made without encountering very high control forces.
(3) Section 25.145(c) contains requirements associated primarily with attempting a go-around maneuver from the landing configuration. Retraction of the high-lift devices from the landing configuration should not result in a loss of altitude if the power or thrust controls are moved to the go-around setting at the same time that flap/slat retraction is begun. The design features involved with this requirement are the rate of flap/slat retraction, the presence of any flap gates, and the go-around power or thrust setting. The go-around power or thrust setting should be the same as is used to comply with the approach and landing climb performance requirements of §§ 25.121(d) and 25.119, and the controllability requirements of §§ 25.145(b)(3), 25.145(b)(4), 25.145( b)(5), 25.149(f), and
25.149(g). The controllability requirements may limit the go-around power or thrust setting.
* (4) Section 25.145(d) provides requirements for demonstrating compliance with § 25.145(c) when gates are installed on the flap selector. Section 25.145(d) also specifies gate design requirements. Flap gates, which prevent the pilot from moving the flap selector through the gated position without a separate and distinct movement of the selector, allow compliance with these requirements to be demonstrated in segments. High lift device retraction must be demonstrated beginning from the maximum landing position to the first gated position, between gated positions, and from the last gated position to the fully retracted position.
(i) If gates are provided, § 25.145(d) requires the first gate from the maximum landing position to be to be located at a position corresponding to a go-around configuration. If there are multiple go-around configurations, the following criteria should be considered when selecting the location of the gate:
(A) The expected relative frequency of use of the available go-around configurations.
(B) The effects of selecting the incorrect high-lift device control position.
(C) The potential for the pilot to select the incorrect control position, considering the likely situations for use of the different go-around positions.
(D) The extent to which the gate(s) aid the pilot in quickly and accurately selecting the correct position of the high-lift devices.
(ii) Regardless of the location of any gates, initiating a go-around from any of the approved landing positions should not result in a loss of altitude. Therefore, § 25.145(d) requires that compliance with § 25.145(c) be demonstrated for retraction of the high-lift devices from each approved landing position to the control position(s) associated with the high-lift device configuration(s) used to establish the go-around procedure(s) from that landing position. A separate demonstration of compliance with this requirement should only be necessary if there is a gate between an approved landing position and its associated go-around position(s). If there is more than one associated
go-around position, conducting this test using the go-around configuration with the most retracted high-lift device position should suffice, unless there is a more critical case. If there are no gates between any of the landing flap positions and their associated go-around positions, the demonstrations discussed in paragraph 21a(4) above should be sufficient to show compliance with this provision of § 25.145(d).
b. Procedures . The following test procedures outline an acceptable means for demonstrating compliance with § 25.145. These tests may be conducted at an optional altitude in accordance with § 25.21(c). Where applicable, the conditions should be maintained on the engines throughout the maneuver.
(l) Longitudinal control recovery, § 25.145(a):
(i) Configuration:
(A) Maximum weight or a lighter weight if considered more critical.
(B) Aft c.g. position.
(C) Landing gear extended.
(D) Wing flaps retracted and extended to the maximum landing position.
(E) Engine power at idle and maximum continuous.
(ii) Test procedure: The airplane should be trimmed at the speed for each configuration as prescribed in § 25.103(b)(1). The airplane should then be decelerated at 1 knot per second with wings level. For tests at idle power, it should be demonstrated that the nose can be pitched down from any speed between the trim speed and the stall. Typically, the most critical point is at the stall when in stall buffet. The rate of speed increase during the recovery should be adequate to promptly return to the trim point. Data from the stall characteristics testing can be used to evaluate this capability at the stall. For tests at maximum continuous power, the maneuver need not be continued for more than one second beyond the onset of stall warning. However, the static longitudinal stability characteristics during the maneuver, and the nose down control power remaining at the end of the maneuver, must be sufficient to assure that a prompt recovery to the trim speed could be attained if the airplane is slowed to the point of stall.
(2) Longitudinal control , flap extension, § 25.145(b)(1).
(i) Configuration:
(A) Maximum landing weight or a lighter weight if considered more critical.
(B) Critical c.g. position.
(C) Wing flaps retracted.
(D) Landing gear extend ed.
(E) Engine power at flight idle.
(ii) Test procedure: The airplane should be trimmed at a speed of 1.4V S. The flaps should be extended to the maximum landing position as rapidly as possible while maintaining approximately 1.4V S for the flap position existing at each instant throughout the maneuver. The control forces should not exceed 50 lbs. (the maximum force for short term application that can be applied readily by one hand) throughout the maneuver without changing the trim control.
(3) Longitudinal control , flap retraction, §§ 25.145(b)(2) & (3).
(i) Configuration:
(A) Maximum landing weight or a lighter weight if considered more critical.
(B) Critical c.g. position.
(C) Wing flaps extended to maximum landing p osition.
(D) Landing gear extended.
(E) Engine power at flight idle and the go-around power or thrust setting.
(ii) Test procedure: With the airplane trimmed at 1.4V S, the flaps should be retracted to the full up position while maintaining approximately 1.4V S for the flap position existing at each instant throughout the maneuver. The longitudinal control force should not exceed 50 lbs. throughout the maneuver without changing the trim control.
(4) Longitudinal control , power application, §§ 25.145(b)(4) & (5)
(i) Configuration:
(A) Maximum landing weight or a lighter weight if considered more critical.
(B) Critical c.g. position.
(C) Wing flaps retracted and extended to the maximum landing position.
(D) Landing gear extended.
(E) Engine power at flight idle.
(ii) Test procedure: The airplane should be trimmed at a speed of 1.4V S. Quickly set go-around power or thrust while maintaining the speed of 1.4V S. The longitudinal control force should not exceed 50 pounds throughout the maneuver without changing the trim control.
(5) Longitudinal control , airspeed variation, § 25.145(b)(6).
(i) Configuration:
(A) Maximum landing weight or a lighter weight if considered more critical.
(B) Most forward c.g. position.
(C) Wing flaps extended to the maximum landing position.
(D) Landing gear extended.
(E) Engine power at flight idle.
(ii) Test Procedure: The airplane should be trimmed at a speed of 1.4V S. The speed should then be reduced to 1.1V S and then increased to 1.7V S, or the flap placard speed, V FE, whichever is lower. The longitudinal control force should not be greater than 50 lbs. Data from the static longitudinal stability tests in the landing configuration at forward c.g., § 25.175(d), may be used to show compliance with this requirement.
(6) Longitudinal control , flap retraction and power application, § 25.145(c).
(i) Configuration:
(A) Critical combinations of maximum landing weights and al titudes.
(B) Critical c.g. position.
(C) Wing flaps extended to the maximum landing position and gated position, if applicable.
(D) Landing gear extended.
(E) Engine power for level flight at a speed of 1.1V S for propeller driven airplanes, or 1.2V S for turbojet powered airplanes.
* (ii) Test procedure. With the airplane stable in level flight at a speed of 1.1V S for propeller driven airplanes, or 1.2V S for turbojet powered airplanes, retract the flaps to the full up position, or the next gated position, while simultaneously setting go-around power or thrust. Use the same power or thrust as is used to comply with the performance requirement of § 25.121(d), as limited by the applicable controllability requirements. It must be possible, without requiring exceptional piloting skill, to prevent losing altitude during the maneuver. Trimming is permissible at any time during the maneuver. If gates are provided, conduct this test beginning from the maximum landing flap position to the first gate, from gate to gate, and from the last gate to the fully retracted position. If there is a gate between any landing position and its associated go-around position(s), this test should also be conducted from that landing position through the gate to the associated go-around position. If there is more than one associated go-around position, this additional test should be conducted using the go-around position corresponding to the most retracted flap position, unless another position is more critical. Keep the landing gear extended throughout the test.
*
(7) Longitudinal control , extreme out-of-trim takeoff conditions, §§ 25.107(e)(4) and 25.143(a)(1).
(i) Configuration:
(A) Critical combinations of takeoff weight and forward and aft c.g. limits.
(B) Wing flaps in all takeoff positions.
(C) All engines operating at maximum takeoff power or thrust.
(ii) The airplane sh ould be loaded to weight and c.g. combinations representing critical corners of the takeoff envelope for both forward and aft c.g. limits. The longitudinal trim should be set for the extreme opposite c.g. (e.g., load to forward c.g. limit at a given weight and set the longitudinal trim for the aft c.g. limit at that weight) as presented in the takeoff trim "green-band" including the takeoff warning system rigging tolerance.
Accomplish a takeoff at normal operating speeds and evaluate the control forces and airplane responses to control inputs. In accordance with § 25.107(e)(4), this out-of-trim takeoff configuration must not result in any unsafe flight characteristics.
22. DIRECTIONAL AND LATERAL CONTROL - § 25.147 .
a. Explanation.
(1) Sections 25.147(a) and (b) provide criteria for investigation of the airplane to determine if it may have dangerous characteristics such as rudder lock or loss of directional control if it is maneuvered with the rudder only, maintaining wings level, when one or two critical engines are inoperative. Some yaw should be possible into the operating engine(s). It should also be possible to make reasonably sudden heading changes of up to 15 degrees, as limited by rudder force or deflection, toward the inoperative engine(s). The intention of the requirement is that the airplane can be yawed as prescribed without the need for application of bank angle. Small variations of bank angle that are inevitable in a realistic flight test demonstration are acceptable.
(2) Sections 25.147(c) and (d) require an airplane to be easily controllable with the critical engine(s) inoperative. Roll response, § 25.147(e), should be satisfactory for takeoff, approach, landing, and high speed configurations. Any permissible configuration that could affect roll response should be evaluated.
b. Procedures .
(1) Directional Control - General, § 25.147(a) .
(i) Configuration:
(A) Maximum landing weight.
(B) Most aft c.g. position.
(C) Wing flaps extended to the approa ch position.
(D) Landing gear retracted.
(E) Yaw SAS on, and off if applicable.
(F) Operating engine(s) at the power for level flight at 1.4V S, but not more than maximum continuous power.
(G) Inoperative engine that would be most cr itical for controllability, with propeller feathered, if applicable.
(ii) Test Procedure. The airplane should be trimmed in level flight at the most critical altitude in accordance with § 25.21(c). Reasonably sudden changes in heading to the left and right, using ailerons to maintain approximately wings level flight, should be made demonstrating a change of up to 15 degrees or that at which 150 lbs. rudder force is required. The airplane should be controllable and free from any hazardous characteristics during this maneuver. For the airplane equipped with a rudder boost system, the evaluation should be done without rudder boost if the boost system can be inoperative.
(2) Directional Control - Four or More Engines, § 25.147(b) .
(i) Configurat ion:
(A) Maximum landing weight.
(B) Most forward c.g. position.
(C) Wing flaps in the most favorable climb position (normally retracted).
(D) Landing gear retracted.
(E) Yaw SAS on, and off if applicable.
(F) Operating engines at the power required for level flight at 1.4V S1, but not more than maximum continuous power.
(G) Two inoperative engines that would be most critical for controllability with (if applicable) propellers feathered.
(ii) Test Procedure. Th e procedure outlined in subparagraph 22b(1)(ii), above, is applicable to this test.
(3) Lateral Control - General, § 25.147(c) .
(i) Configuration:
(A) Maximum takeoff weight.
(B) Most aft c.g. position.
(C) Wing flaps in the most favorable climb position.
(D) Landing gear retracted and extended.
(E) Yaw SAS on, and off if applicable.
(F) Operating engine(s) at maximum continuous power.
(G) The inoperative engine that would be most critical for controllabili ty, with the propeller feathered, if applicable.
(ii) Test Procedure. With the airplane trimmed at 1.4V S, turns with a bank angle of 20 degrees should be demonstrated with and against the inoperative engine from a steady climb at 1.4V S1. It should not take exceptional piloting skill to make smooth, predictable turns.
(4) Lateral Control - Four or More Engines, § 25.147(d) .
(i) Configuration:
(A) Maximum takeoff weight.
(B) Most aft c.g. position.
(C) Wing flaps in the most favorable climb position.
(D) Landing gear retracted and extended.
(E) Yaw SAS on, and off if applicable.
(F) Operating engines at maximum continuous power.
(G) Two inoperative engines most critical for controllability, with propell ers feathered, if applicable.
(ii) Test Procedure: The procedure outlined in paragraph 22b(3)(ii) is applicable to this test.
(5) Lateral Control - All Engines Operating, § 25.147(e) .
(i) Configuration: All configurations within the flight e nvelope for normal operation.
(ii) Test Procedure: This is primarily a qualitative evaluation that should be conducted throughout the test program. Roll performance should be investigated throughout the flight envelope, including speeds to VFC/MFC, to ensure adequate peak roll rates for safety, considering the flight condition, without excessive control force or travel. Roll response during sideslips expected in service should provide maneuvering capabilities adequate to recover from such conditions. Approach and landing configurations should be carefully evaluated to ensure adequate control to compensate for gusts and wake turbulence while in close proximity to the ground.
23. MINIMUM CONTROL SPEED - § 25.149 .
a. Explanation. Section 25.149 defines requirements for minimum control speeds during takeoff climb (VMC), during takeoff ground roll (V MCG), and during approach and landing (V MCL and VMCL-2). The VMC (commonly referred to as V MCA) requirements are specified in §§ 25.149(a), (b), (c) an d (d); the VMCG requirements are described in § 25.149(e); and the V MCL and VMCL-2 requirements are covered in §§ 25.149(f), (g) and (h). Section 25.149(a) states that "...the method used to simulate critical engine failure must represent the most critical mode of powerplant failure with respect to controllability expected in service." That is, the thrust loss from the inoperative engine must be at the rate that would occur if an engine suddenly became inoperative in service.
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.