(c) The acceptable HQ ratings for A-PC tendencies is shown in figure 12 of Appendix 6. As described in that appendix, the minimum HQ rating, and consequently the pass/fail criteria, varies with the flight envelope, atmospheric disturbance considered, and failure state. For example, figure 20-3 below shows a handling qualities matrix for a tracking task with the airplane at aft c.g. trimmed in flight conditions giving 1.3 g to buffet onset.
Figure 20-3 Example of Acceptable HQ Rating For A-PC Tendencies
Airplane at aft c.g. trimmed in conditions giving 1.3 g to buffet onset
AIRSPEED M LRC M LRC M LRC M LRC LOAD FACTOR RANGE
0.8 TO 1.3
-1.0 TO 2.5
0.8 TO 1.3
-1.0 TO 2.5 BUFFET LEVEL
ONSET DETERRENT ONSET DETERRENT
TURBULENCE LIGHT LIGHT LIGHT LIGHT
FAILURE
NONE
NONE
IMPROBABLE
FAILURE OF SAS
IMPROBABLE
FAILURE OF SAS FLIGHT
ENVELOPE NFE LFE NFE LFE
MINIMUM
PERMITTED
HQ RATING
SAT
ADQ
ADQ
CON
SAT = Satisfactory ADQ = Adequate CON = Controllable NFE = Normal flight envelope LFE = Limit flight envelope SAS = Stability augmentation system M LRC = Long range cruise mach number
e. Maneuvering Characteristics - § 25.143(g) .
(1) General . An acceptable means of compliance with the requirement that stick forces may not be excessive when maneuvering the airplane is to demonstrate that, in a turn for 0.5g incremental normal acceleration (0.3g above 20,000 feet) at speeds up to VFC/MFC, the average stick force gradient does not exceed 120 pounds per g.
(2) Interpretive Material .
(a) The objective of § 25.143(g) is to ensure that the limit strength of any critical component on the airplane would not be exceeded in maneuvering flight. In much of the structure, the load sustained in maneuvering flight can be assumed to be directly proportional to the load factor applied. However, this may not be the case for some parts of the structure (e.g.,
92 the tail and rear fuselage). Nevertheless, it is accepted that the airplane load factor will be a sufficient guide to the possibility of exceeding limit strength on any critical component if a structural investigation is undertaken whenever the design positive limit maneuvering load factor is closely approached. If flight testing indicates that the positive design limit maneuvering load factor could be exceeded in steady maneuvering flight with a 50 pound stick force, the airplane structure should be evaluated for the anticipated load at a 50 pound stick force. The airplane will be considered to have been overstressed if limit strength has been exceeded in any critical component. For the purposes of this evaluation, limit strength is defined as the lesser of either the limit design loads envelope increased by the available margins of safety, or the ultimate static test strength divided by 1.5.
(b) Minimum Stick Force to Reach Limit Strength.
1 A stick force of at least 50 pounds to reach limit strength in steady maneuvers 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.
2 This minimum stick force applies in the en route configuration with the airplane 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(g) are applicable in these conditions.
(c) Stick Force Characteristics.
1 At all points within the buffet onset boundary determined in accordance with § 25.251(e), but not including speeds above VFC/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.
2 Beyond the buffet onset boundary, hazardous stick force characteristics should not be encountered within the permitted maneuvering envelope as limited by paragraph 20e(2)(c)3. 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:
(aa) 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.
93 (bb) 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.
3 In flight tests to satisfy paragraphs 20e(2)(c)(1 ) and (2), the load factor should be increased until either:
(aa) The level of buffet becomes sufficient to provide a strong and effective deterrent to any further increase of the load factor; or
(bb) 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
(cc) The positive limit maneuvering load factor established in compliance with § 25.337(b) is achieved.
(d) 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.
f. Thrust or Power Setting for Maneuver Capability Demonstrations. The effect of thrust or power on maneuver capability is normally a function of only the thrust-to-weight ratio.
Therefore, for those configurations in which the weight, altitude, temperature (WAT)-limited thrust or power setting is prescribed, it is usually acceptable to use the thrust or power setting that is consistent with a WAT-limited climb gradient at the test conditions of weight, altitude, and temperature. However, if the maneuver margin to stall warning (or other characteristic that might interfere with normal maneuvering) is reduced with increasing thrust or power, the critical conditions of both thrust or power and thrust-to-weight ratio should be taken into account when demonstrating the required maneuvering capabilities.
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 ensure sufficient pitch control for a prompt recovery if inadvertently slowed to the point of stall. Although this requirement must be met with power off and at maximum continuous thrust or power, there is no intention to require stall demonstrations with thrust or power above that specified in § 25.101(a)(2). Instead of performing a full stall at maximum continuous power or thrust, compliance may be assessed by demonstrating sufficient static longitudinal stability and nose down control margin when the deceleration is ended at least one second past stall warning during a one knot per second deceleration. The static longitudinal stability during the maneuver and the nose down control power remaining at the end of the maneuver must be sufficient to assure compliance with the requirement.
(2) Section 25.145(b) requires changes to be made in flap position, power, and speed without undue effort when re-trimming is not practical. 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 goaround 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.
(a) 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:
1 The expected relative frequency of use of the available go-around configurations.
2 The effects of selecting the incorrect high-lift device control position.
3 The potential for the pilot to select the incorrect control position, considering the likely situations for use of the different go-around positions.
4 The extent to which the gate(s) aid the pilot in quickly and accurately selecting the correct position of the high-lift devices.
(b) 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).
(a) Configuration:
1 Maximum weight or a lighter weight if considered more critical.
2 Aft c.g. position.
3 Landing gear extended.
4 Wing flaps retracted and extended to the maximum landing position.
5 Engine power at idle and maximum continuous.
(b) Test procedure: The airplane must be trimmed at the speed for each configuration as prescribed in § 25.103(b)(6). The airplane should then be decelerated at 1 knot per second with wings level. For tests at idle power, the applicant must demonstrate 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).
(a) Configuration:
1 Maximum landing weight or a lighter weight if considered more critical.
2 Critical c.g. position.
3 Wing flaps retracted.
4 Landing gear extended.
5 Engine power at flight idle.
(b) Test procedure: The airplane must be trimmed at a speed of 1.3 V SR. The flaps must be extended to the maximum landing position as rapidly as possible while maintaining approximately 1.3 VSR for the flap position existing at each instant throughout the maneuver. The control forces must 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).
(a) Configuration:
1 Maximum landing weight or a lighter weight if considered more critical.
2 Critical c.g. position.
3 Wing flaps extended to maximum landing position.
4 Landing gear extended.
5 Engine power at flight idle and the go-around power or thrust setting.
(b) Test procedure: With the airplane trimmed at 1.3 V SR, the flaps must be retracted to the full up position while maintaining approximately 1.3 VSR for the flap position existing at each instant throughout the maneuver. The longitudinal control force must not exceed 50 lbs. throughout the maneuver without changing the trim control.
(4) Longitudinal control , power application, § 25.145(b)(4) & (5).
(a) Configuration:
1 Maximum landing weight or a lighter weight if considered more critical.
2 Critical c.g. position.
3 Wing flaps retracted and extended to the maximum landing position.
4 Landing gear extended.
5 Engine power at flight idle.
(b) Test procedure: The airplane must be trimmed at a speed of 1.3 V SR. Quickly set go-around power or thrust while maintaining the speed of 1.3 VSR. The longitudinal control force must not exceed 50 pounds throughout the maneuver without changing the trim control.
(5) Longitudinal control , airspeed variation, § 25.145(b)(6).
(a) Configuration:
1 Maximum landing weight or a lighter weight if considered more critical.
2 Most forward c.g. position.
3 Wing flaps extended to the maximum landing position.
4 Landing gear extended.
5 Engine power at flight idle.
(b) Test Procedure: The airplane must be trimmed at a speed of 1.3 V SR. The speed should then be reduced to VSW and then increased to 1.6 VSR, or the maximum flap extended speed, VFE, whichever is lower. The longitudinal control force must not be greater than
98 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).
(a) Configuration:
1 Critical combinations of maximum landing weights and altitudes.
2 Critical c.g. position.
3 Wing flaps extended to the maximum landing position and gated position, if applicable.
4 Landing gear extended.
5 Engine power for level flight at a speed of 1.08 V SR for propeller driven airplanes, or 1.13VSR for turbojet powered airplanes.
(b) Test procedure. With the airplane stable in level flight at a speed of 1.08 V SR for propeller driven airplanes, or 1.13 VSR 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).
(a) Configuration:
1 Critical combinations of takeoff weight and forward and aft c.g. limits.
2 Wing flaps in all takeoff positions.
3 All engines operating at maximum takeoff power or thrust.
(b) The airplane should 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 (e) require an airplane to be easily controllable with the critical engine(s) inoperative. Section 25.147(d) further requires that lateral control be sufficient to provide a roll rate necessary for safety, without excessive control forces or travel, at the speeds likely to be used with one engine inoperative. Compliance can normally be demonstrated in the takeoff configuration at V2 speed, because this condition is usually the most critical.
Normal operation of a yaw stability augmentation system (SAS) should be considered in accordance with the normal operating procedures. Roll response with all engines operating, § 25.147(f), should be satisfactory for takeoff, approach, landing, and high speed configurations.
Any permissible configuration that could affect roll response should be evaluated.
b. Procedures . The following test procedures outline an acceptable means for demonstrating compliance with § 25.147.
(1) Directional Control - General, § 25.147(a) .
(a) Configuration:
1 Maximum landing weight.
2 Most aft c.g. position.
3 Wing flaps extended to the approach position.
100 4 Landing gear retracted.
5 Yaw SAS on, and off if applicable.
6 Operating engine(s) at the power for level flight at 1.3 V SR, but not more than maximum continuous power.
7 Inoperative engine that would be most critical for controllability, with propeller feathered, if applicable.
(b) Test Procedure. The airplane must be trimmed in level flight at the most critical altitude in accordance with § 25.21(c). Make reasonably sudden changes in heading to the left and right of up to 15 degrees (but not using more than 150 lbs. rudder force), using ailerons to maintain approximately wings level flight. 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) .
(a) Configuration:
1 Maximum landing weight.
2 Most forward c.g. position.
3 Wing flaps in the most favorable climb position (normally retracted).
4 Landing gear retracted.
5 Yaw SAS on, and off if applicable.
6 Operating engines at the power required for level flight at 1.3 V SR1, but not more than maximum continuous power.
7 Two inoperative engines that would be most critical for controllability with (if applicable) propellers feathered.
(b) Test Procedure. The procedure outlined in subparagraph 22b(1)(b), above, is applicable to this test.
(3) Lateral Control - General, § 25.147(c) .
(a) Configuration:
1 Maximum takeoff weight.
2 Most aft c.g. position.
3 Wing flaps in the most favorable climb position.
4 Landing gear retracted and extended.
5 Yaw SAS on, and off if applicable.
6 Operating engine(s) at maximum continuous power.
7 The inoperative engine that would be most critical for controllability, with the propeller feathered, if applicable.
(b) Test Procedure. With the airplane trimmed at 1.3 V SR1, turns with a bank angle of 20 degrees must be demonstrated with and against the inoperative engine from a steady climb at 1.3 VSR1. It should not take exceptional piloting skill to make smooth, predictable turns.
(4) Lateral Control – Roll Capability, § 25.147(d).
(a) Configuration:
1 Maximum takeoff weight.
2 Most aft c.g. position.
3 Wing flaps in the most critical takeoff position.
4 Landing gear retracted.
5 Yaw SAS on, and off, if applicable.
6 Operating engine(s) at maximum takeoff power.
7 The inoperative engine that would be most critical for controllability, with propellers feathered, if applicable.
(b) Test Procedure. With the airplane in trim, or as nearly as possible in trim, for straight flight at V2, establish a steady 30 degree banked turn. Demonstrate that the airplane can be rolled to a 30 degree bank angle in the other direction in not more than 11 seconds. The rudder may be used to the extent necessary to minimize sideslip. Demonstrate this maneuver in the most adverse direction. The maneuver may be unchecked, that is, the pilot need not apply a control input to stop the roll until after the 30 degree bank angle is achieved. Care should be taken to prevent excessive sideslip and bank angle during the recovery.
(5) Lateral Control - Four or More Engines, § 25.147(e) .
(a) Configuration:
1 Maximum takeoff weight.
2 Most aft c.g. position.
3 Wing flaps in the most favorable climb position.
4 Landing gear retracted and extended.
5 Yaw SAS on, and off if applicable.
6 Operating engines at maximum continuous power.
7 Two inoperative engines most critical for controllability, with propellers feathered, if applicable.
(b) Test Procedure: The procedure outlined in paragraph 22b(3)(b) is applicable to this test.
(6) Lateral Control - All Engines Operating, § 25.147(f) .
(a) Configuration: All configurati ons within the flight envelope for normal operation.
(b) 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 (VMCG), and during approach and landing (VMCL and VMCL-2). The VMC (commonly referred to as VMCA) requirements are specified in § 25.149(a), (b), (c) and (d); the VMCG requirements are described in § 25.149(e); and the VMCL 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