(ii) For pitch captures, the airplane is trimmed for a specified flight condition. The pilot aggressively captures 5 degrees pitch attitude (or 10 degrees if the airplane is already trimmed above 5 degrees).
The pilot then makes a series of aggressive pitch captures of 5 degree increments in both directions, and then continues this procedure with ten degree increments in both directions. An airplane with more capability can continue the procedure with larger pitch excursions. If possible, the initial conditions for each maneuver should be such that the airplane will remain within ± 1,000 feet and ± 10 knots of the specified flight condition during the maneuver; however, large angle captures at high-speed conditions will inevitably produce larger speed and altitude changes. If the airplane should get too far from the specified condition during a task, it should be retrimmed for the specified condition before starting the next maneuver.
(iii) The other kinds of captures are usually done in a similar manner, with some minor differences.
G captures can be done from a constant-g turn or pull ups and pushovers using ± 0.2 g and ± 0.5 g. Heading captures can be used to evaluate the yaw controller alone (usually small heading changes of 5 degrees or less).
(iv) Bank angle captures are also commonly done using bank-to-bank rolls. Starting from a 15 degree bank angle, the pilot aggressively rolls and captures the opposite 15 degree bank angle (total bank angle change of 30 degrees). The pilot then rolls back and captures 15 degrees bank in the original direction. This procedure should continue for a few cycles. The procedure is then repeated using 30 degree bank angles, and then repeated again using 45 degree bank angles. A variation of this is to capture wings-level from the initial bank condition.
(v) Where suitable, combined conditions could be used as described in the task shown in paragraph (vii), below, in which a target g and bank angle are tightly tracked until the target pitch attitude and heading are captured.
(vi) The following upset and/or collision avoidance maneuvers have been found to be effective in evaluating A-PC susceptibility when the airplane is flying at high altitude under conditions of low g to buffet onset, typically 0.3g. This emphasis on cruise susceptibility stems from operational experiences, but should not be interpreted as placing less emphasis on other flight phases.:
(A) Trim for level flight at long-range cruise Mach number. Initiate a slight climb and slow the aircraft while leaving power set. Push the nose over and set up a descending turn with 30 to 40 degrees of bank and approximately 10 degrees nose below the horizon, or as appropriate, to accelerate to the initial trim speed. At the initial trim airspeed initiate a 1.5 g to 1.67 g (not to exceed deterrent buffet) pull up and establish a turn in the opposite direction to a heading which will intercept the initial course on which the airplane was trimmed. Establish a pitch attitude which will provide a stabilized climb back to the initial trim altitude. The pilot may use the throttles as desired during this maneuver and should pick a target g, bank angle, heading, and pitch attitude to be used prior to starting the maneuver. The target g and bank angle should be set and tightly tracked until the target pitch attitude and heading are obtained respectively. The stabilized steady heading climb should be tightly tracked for an adequate amount of time to allow the pilot to assess handling qualities, even through the initial trim altitude and course if required. The pilot should qualitatively evaluate the airplane during both the gross acquisition and fine tracking portions of this task while looking for any tendency towards A-PC in accordance with the criteria in paragraph 20d(6).
(B) This maneuver should be repeated in the nose-down direction by accelerating to M MO from the trim condition 10 degrees nose down and then recover as above.
(C) Trim for level flight as above. Initiate a 1.5 g to 1.67 g ( not to exceed deterrent buffet) pull-up and approximately a 30 degree bank turn. Once the target g is set, transition the aircraft to approximately a
0.5 g pushover and reverse the turn to establish an intercept heading to the initial course. Using power as required, set up a stabilized steady heading descent to intercept the initial course and altitude used for the trimmed condition.
The pilot may continue the heading and descent through the initial conditions to allow more tracking time if needed.
Attempt to precisely set and track bank angle, g, heading, and pitch attitude as appropriate. The pilot should qualitatively evaluate the airplane during both the gross acquisition and fine tracking portions of this task while looking for an A-PC tendency in accordance with paragraph 20d(6).
(5) Fine Tracking Tasks .
(i) These tasks may be used to assess the airplane's A-PC susceptibility when flying in turbulent atmospheric conditions. In this task, a tracking target is displayed which commands pitch and roll changes for the evaluation pilot to follow. Whatever visual cue is used (e.g., HUD, flight director, etc.), it should present the tracking task without filtering, smoothing, or bias. The pitch and roll commands should be combinations of steps and ramps. The sequence of pitch and roll commands should be designed so as to keep the airplane within ± 1,000 feet of the test altitude and within ±10 knots of the test airspeed. The sequence should be long enough and complex enough that the pilot cannot learn to anticipate the commands. The unfamiliarity is intended to help keep the test pilot's gain high and to preclude inadvertent pilot compensation while accomplishing the task. Such compensation, along with reduced gains, could mask any A-PC tendencies.
(ii) Even though these fine tracking tasks will provide insight into A-PC susceptibility of a conventional airplane when flying in turbulence, other considerations apply to augmented airplane types. For example, structural load alleviation systems that use the same flight control surface as the pilot will limit the pilot's control authority in turbulent atmospheric conditions. Under these circumstances of rate or position limiting, A-PC tendencies will be more critical as previously discussed. Therefore, specific evaluations for turbulent atmospheric conditions with these systems operating are necessary for these airplane types.
(iii) For single axis tasks, it has been found that aural commands given in a timed sequence provide an adequate cue in the event it is not possible to modify the flight director to display the pitch commands.
(iv) Based on A-PC events seen in service, high altitude tracking tasks (with up to approximately ± 4° pitch excursions from trim occurring at varying intervals of approximately 2 to 5 seconds) have been effective in evaluating A-PC susceptibility. These tasks have been used where the airplane is flying under conditions of low g margin to buffet onset. The following time history is a pictorial representation of a sample task in MIL-STD-1797A that has the desired attributes for high altitude A-PC evaluations:
FIGURE 20-1. SAMPLE PITCH TRACKING TASK t (sec) delta pitch -4 -3 -2 -1 0 1 2 3 4 0 5 10 15 20 25 30 35 40 45 50 55 60 65 70 75 80 85 90
(6) A-PC Assessment Criteria .
* (i) The evaluation of an airplane for A-PC susceptibility will be conducted using the FAA Handling Qualities Rating Method (HQRM). Tasks should be designed to focus on any A -PC tendencies that may exist. Figure 20-2 contains the descriptive material associated with A-PC characteristics and its relationship to the PIO Rating Scale called out in the U.S. Military Standard.
(ii) Figure 20-2 provides the FAA Handling Qualities Rating descriptions of airplane motions that may be seen during the conduct of specific A-PC tasks or during tests throughout the entire certification flight test program. The italicized phrases highlight major differences between rating categories in the table.
* FIGURE 20-2. A-PC RATING CRITERIA AND COMPARISON TO MIL STANDARD FAA HQ
RATING A-PC CHARACTERISTICS DESCRIPTION MIL 1797A STD.
PIO RATING SCALE NO TENDENCY FOR PILOT TO INDUCE UNDESIRABLE MOTION. 1
SAT UNDESIRABLE MOTIONS (OVERSHOOTS) TEND TO OCCUR
WHEN PILOT INITIATES ABRUPT MANEUVERS OR ATTEMPTS TIGHT CONTROL. THESE MOTIONS CAN BE PREVENTED OR ELIMINATED BY PILOT TECHNIQUE. (NO MORE THAN MINIMAL
PILOT COMPENSATION REQUIRED) 2 ADQ
UNDESIRABLE MOTIONS (UNPREDICTABILITY OR OVER
CONTROL) EASILY INDUCED WHEN PILOT INITIATES ABRUPT MANEUVERS OR ATTEMPTS TIGHT CONTROL.
THESE MOTIONS CAN BE PREVENTED OR ELIMINATED BUT
ONLY AT SACRIFICE TO TASK PERFORMANCE OR THROUGH CONSIDERABLE PILOT ATTENTION AND EFFORT. (NO MORE
THAN EXTENSIVE PILOT COMPENSATION REQUIRED) 3 CON
OSCILLATIONS TEND TO DEVELOP WHEN PILOT INITIATES ABRUPT MANEUVERS OR ATTEMPTS TIGHT CONTROL.
ADEQUATE PERFORMANCE IS NOT ATTAINABLE AND PILOT MUST REDUCE GAIN TO RECOVER. (PILOT CAN RECOVER BY
MERELY REDUCING GAIN) 4 UNSAT
DIVERGENT OSCILLATIONS TEND TO DEVELOP WHEN PILOT
INITIATES ABRUPT MANEUVERS OR ATTEMPTS TIGHT CONTROL. PILOT MUST OPEN LOOP BY RELEASING OR FREEZING THE CONTROLLER.
DISTURBANCE OR NORMAL PILOT CONTROL MAY CAUSE DIVERGENT OSCILLATION. PILOT MUST OPEN CONTROL LOOP BY RELEASING OR FREEZING THE CONTROLLER.
6 * SAT = Satisfactory ADQ = Adequate CON = Controllable UNSAT = Unsatisfactory or Failed
* (iii) The required H.Q. rating for A-PC tendencies is shown in Figure 12 of Appendix 7. As described in that appendix, the minimum H.Q. 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 MINIMUM PERMITTED HQ RATING FOR A-PC TENDENCIES Airplane at aft c.g. trimmed in conditions giving 1.3 g to buffet onset
AIRSPEED MLRC MLRC MLRC MLRC
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 IMPROBABLE 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 MLRC = Long range cruise mach number
e. Tailplane stall investigation with ice accretion .
(1) Configuration:
(i) All normal combinations of wing flaps and landing gear, except the cruise configuration.
(ii) Critical weight and c.g. for the test wing flap and landing gear position.
(iii) Speeds from 1.2V S or VREF-5 kts, as appropriate to the wing flap position, up to the maximum speed to be encountered operationally in a given flap/gear configuration that will not result in exceeding V FE or VLE, as applicable, during the recovery.
(iv) Power or thrust: Flight idle to maximum takeoff.
(v) Icing condition: The applicant should specify the critical ice case(s) to be investigated in terms of location, shape, thickness, and texture. FAA concurrence on the critical ice shape(s) should be obtained prior to testing. The critical ice case(s) should include an allowance for any time delays in activation of the ice protection system associated with ice detection or observation systems, or that may be reasonably expected in service. If ice accretion on the wings can result in increased airflow angle of attack on the horizontal tail, ice shapes will also be required on the wing surface, and should be representative of the ice accumulation that would normally be expected in the icing condition defined to be critical for the tailplane. It should be noted that ice accreted with the flaps retracted may result in a more critical condition than ice accreted with the flaps extended. Ice accretion thickness need not be greater than that resulting from flight in the maximum continuous atmospheric icing conditions defined in Appendix C to Part 25.
(2) Zero g Pushover Maneuver: The test procedure described below is essentially a nose down pitching maneuver. Execution of this maneuver should be preceded by experimentation to determine initial entry speeds and pitch attitudes that will result in the airplane achieving the target load factor and airspeed as it pitches through approximately level flight.
(i) Trim the airplane in the test configurations and at the test s peeds prescribed in paragraph 20e(1), above.
(ii) Initial testing should be conducted by moving the control column forward at a slow rate while evaluating for force lightening or force reversal. Continue the test by incrementally increasing the rate of control movement until a zero g flight condition is obtained or, if limited by elevator power, to the lowest load factor attainable.
WARNING: The test maneuver described above may result in a sudden and violent loss of pitch stability or control due to aerodynamic stall of the horizontal stabilizer and/or elevators. High stick forces may be required to effect recovery; consequently, the non-flying pilot should be prepared to immediately retract the flaps and lend assistance in overcoming any high nose down stick forces.
(iii) A push longitudinal control force must be required throughout the test maneuver.
(iv) The airplane should demonstrate suitable controllability and maneuverability throughout the maneuver with no force reversal and no tendency to diverge in pitch.
(3) Steady State Sideslip Maneuver: The following test procedure is intended to determine if the airplane is susceptible to tailplane stall resulting from detached flow on the horizontal tail as induced by a lateral flow component.
(i) For the test conditions described in paragraph 20e(1), establish the airplane in a straight, steady sideslip, up to the sideslip angle appropriate to normal operation of the airplane as used to demonstrate compliance with § 25.177(c).
(ii) Changes in longitudinal control force to maintain speed with increasing sideslip should be progressive with no reversals or sudden discontinuities.
f. Maneuvering Characteristics - § 25.143(f) .
(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 V FC/MFC, the average stick force gradient does not exceed 120 pounds per
g.
(2) Interpretive Material .
(i) The objective of § 25.143 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., 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.
(ii) Minimum Stick Force to Reach Limit Strength.
(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.