(i) This segment represents the flight test measured average time from touchdown to pilot activation of the first deceleration device. For AFM data expansion, use the longer of 1 second or the test time.
(ii) This segment represents the flight test measured average test time from pilot activation of the first deceleration device to pilot activation of the second deceleration device. For AFM data expansion, use the longer of 1 second or the test time.
(iii) Step is repeated until pilot activation of all deceleration devices has been completed and the airplane is in the full braking configuration.
(4) For approved automatic deceleration devices (e.g., autobrakes or auto-spoilers, etc.) for which performance credit is sought for AFM data expansion, established times determined during certification testing may be used without the application of the 1-second minimum time delay required in the appropriate segment above.
(5) It has been considered acceptable to expand the airborne portion of the landing distance in terms of a fixed airborne time, independent of airplane weight or approach speed.
(6) Assumptions to be made in assessing the effect of wind on landing distance are discussed in paragraph 3 of this AC.
h. Parametric Analysis Data Reduction . The following is an acceptable method of converting the test data to a mathematical model for the parametric analysis method of air distance described in paragraph 19b(3).
Test Data for Each Test Point:
R/S50 = Rate of sink at 50 ft. above landing surface, Ft/Sec R/STD = Rate of sink at touchdown, Ft/Sec V50 = True airspeed at 50 ft. above landing surface, Ft/Sec VTD = True airspeed at touchdown, Ft/Sec t = Air time 50 ft. to touchdown, Sec The multiple linear regression analysis as outlined below is used to solve for the constant of the two independent variable equations:
50/t = a + b(R/S 50) + (c)(R/STD) To maintain the same units for all variables, the dependent variable is chosen as 50/t.
The test values of all the test points, 1 through n, are processed as follows, where n equals the number of test points and R1 through R13 are the regression coefficients:
R1 = 1n∑ R/S50 R2 = 1n∑ (R/S50)2 R3 = 1n∑ R/STD R4 = 1n∑ (R/STD)2 R5 = 1n∑ (R/S50)(R/STD) R6 = 1n∑ (50/t) R7 = 1n∑ (R/S50)(50/t) R8 = 1n∑ (R/STD)(50/t) R9 = (n)(R2)-(R1)2 R10 = (n)(R8)-(R3)(R6) R11 = (n)(R5)-(R1)(R3) R12 = (n)(R7)-(R1)(R6) R13 = (n)(R4)-(R3)2 c = ((R9)(R10)-(R11)(R12))/((R9)(R13) -(R11)2) b = ((R12)-(c)(R11))/R9 a = ((R6)-(b)(R1)-(c)(R3))/n In the same manner, determine the values of the constants, a, b, and c, in an equation for speed reduction between 50 ft. and touchdown by replacing 50/t with (V 50/VTD) for each test run.
After determining the values of the constants, the two equations are used to calculate the time from 50 ft. to touchdown and V50/VTD for the desired conditions of -3.5 degrees flight path and R/S TD = 8 ft/Sec. The R/S50 is calculated from the approach path and V 50.
After VTD is determined, the air distance may be determined for the average flare speed and t.
Example:
Test Data :
Run R/S50 R/STD V50 VTD t 1 13.4 6.1 219 214 5.6 2 10.9 1.8 223 218 8.5 3 7.9 5.8 209 201 7.4 4 8.3 2.3 213 206 9.6 5 9.8 4.1 218 212 7.5 Results :
50/t = 1.0432 + .3647(R/S 50) + .4917(R/STD) V50/VTD = 1.05508 - .003198(R/S50) + .001684(R/S TD) For conditions of V 50 = 220, flight path = -3.5 degrees, R/S TD = 8.0, the resultants are:
R/S50 = 13.43 V50/VTD = 1.0256 t = 5.063 sec. Air Distance = 1100 ft.
Section 3. CONTROLLABILITY AND MANEUVERABILITY
20. GENERAL - § 25.143 .
a. Explanation. The purpose of § 25.143 is to verify that any operational maneuvers conducted within the operational envelope can be accomplished smoothly with average piloting skill and without exceeding any airplane structural limits. Control forces should not be so high that the pilot cannot safely maneuver the airplane. Also, the forces should not be so light that it would take exceptional skill to maneuver the airplane without over-stressing it or losing control. The airplane response to any control input should be predictable to the pilot.
(1) The maximum forces given in the table in § 25.143(c) for pitch and roll control for short term application are applicable to maneuvers in which the control force is only needed for a short period. Where the maneuver is such that the pilot will need to use one hand to operate other controls (such as the landing flare or a goaround, or during changes of configuration or power resulting in a change of control force that must be trimmed out) the single-handed maximum control forces will be applicable. In other cases (such as takeoff rotation, or maneuvering during en route flight), the two-handed maximum forces will apply.
(2) Short-term and long-term for ces should be interpreted as follows:
(i) Short-term forces are the initial stabilized control forces that result from maintaining the intended flight path following configuration changes and normal transitions from one flight condition to another, or from regaining control following a failure. It is assumed that the pilot will take immediate action to reduce or eliminate such forces by re-trimming or changing configuration or flight conditions, and consequently short-term forces are not considered to exist for any significant duration. They do not include transient force peaks that may occur during the configuration change, change of flight conditions, or recovery of control following a failure.
(ii) Long-term forces are those control forces that r esult from normal or failure conditions and that cannot readily be trimmed out or eliminated.
(3) Under the general requirements of § 25.143(a) for the airplane to "be safely controllable and maneuverable...," the potential for tailplane stall should be investigated. Elevator hinge moment becomes more sensitive to angle of attack as the tailplane stall is approached. This is particularly noticeable with elevators having significant aerodynamic balancing (leading edge and/or horn balance). In flight, these effects become apparent by lightening stick pull forces in stall maneuvers and lightening stick push forces in nose down pitching maneuvers.
Tailplane leading edge surface condition and deposits of frost and ice can also have a significant effect on the available angle of attack range resulting in decreased tailplane stall margins. The loss of control, transient or sustained, due to flow separation at the tailplane, can thus occur with both manually operated and irreversible powered controls. Experience has shown that airplanes with non-powered elevators are the most at risk to encounter the phenomenon of large stick-force changes, but it is considered that all airplanes should be assessed for controllability problems in this respect. The possibility of encountering a tailplane stall condition is increased considerably during operations in icing conditions. Experience has shown that even a thin layer of rough sandpaper ice on the tailplane surfaces can cause longitudinal control problems. Consequently, a pushover maneuver, similar to that described in paragraph 20e, should be evaluated with regard to tailplane stall, with expected in-service frost and ice accretions present on the tailplane surfaces. The applicant must determine the critical icing accretion with regard to location, shape, thickness, and texture with FAA concurrence. For airplanes unlikely to be critical in certain icing conditions, it may be possible to show adequate longitudinal controllability by means other than flight test (see paragraph 20e(1)(v).
(4) In conducting the controllability and maneuverability tests of §§ 25.143, 25.145 and 25.147, consideration should be given to the requirements of § 25.671(c). Section 25.671(c) requires the airplane to "be shown by analysis, tests, or both, to be capable of continued safe flight and landing after any....failures or jamming in the flight control systems and surfaces (including trim, lift, drag, and feel systems), within the normal flight envelope...." This would include any single failure, or any combination of failures not shown to be extremely improbable. The specific concerns for possible flight test evaluation are as follows:
(i) Unless safe flight characteristics are demonstrated with functionally related high lift or drag devices retracted on one side of the airplane and extended on the other, the motion of the devices on opposite sides
of the plane of symmetry must be synchronized by a mechanical interconnection or approved equivalent means as required by § 25.701 (Flap and slat interconnection).
(ii) The "safe flight characteristics" referred to in paragraph ( i), above, are interpreted as being the retention of adequate stall margins and controllability, following the described failure conditions, without requiring exceptional piloting skill or strength.
(iii) Additionally, no hazardous change in altitude, or attitude, should occur during transition to the asymmetric condition considering likely transition rates.
(iv) Cases of primary flight control jams (i .e.; pitch, roll, and yaw) to be considered should include the following:
(A) For pitch controls:
(1) A jam at the initiation of the takeoff rotation, and
(2) A jam at the most critical peak elevator angle reached during a normal rotation maneuver.
(B) For roll and yaw controls, considered separately, a jam in the control position necessary to initiate all normal turns.
b. General Test Requirements . Compliance with § 25.143 is primarily a qualitative determination by the pilot during the course of the flight test program. The control forces required and airplane response should be evaluated during changes from one flight condition to another and during maneuvering flight. The forces required should be appropriate to the flight condition being evaluated. For example, during an approach for landing, the forces should be light and the airplane responsive in order that adjustments in the flight path can be accomplished with a minimum of workload. In cruise flight, forces and airplane response should be such that inadvertent control input does not result in exceeding limits or in undesirable maneuvers. Longitudinal control forces should be evaluated during accelerated flight to ensure a positive stick force with increasing normal acceleration. Forces should be heavy enough at the limit load factor to prevent inadvertent excursions beyond the design limit. Sudden engine failures should be investigated during any flight condition or in any configuration considered critical, if not covered by another section of Part 25. Control forces considered excessive should be measured to verify compliance with the maximum control force limits specified in § 25.143(c). Allowance should be made for delays in the initiation of recovery action appropriate to the situation.
c. Controllability Following Engine Failure . Section 25.143(b)(1) requires the airplane to be controllable following the sudden failure of the critical engine. To show compliance with this requirement, the demonstrations described in paragraphs (1) and (2), below, should be made with engine failure (simulated by fuel cuts) occurring during straight, wings level flight. To allow for likely in-service delays in initiating recovery action, no action should be taken to recover control for two seconds following pilot recognition of engine failure. The recovery action should not necessitate movement of the engine, propeller, or trim controls, and should not result in excessive control forces. Additionally, the airplane will be considered to have reached an unacceptable attitude if the bank angle exceeds 45 degrees during the recovery. These tests may be conducted using throttle slams to idle, with actual fuel cuts repeated only for those tests found to be critical.
(1) At each tak eoff flap setting at the initial all-engine climb speed (i.e., V 2 + 10 knots) with:
(i) All engines operating at maximum takeoff power or thrust prior to failure of the critical engine;
(ii) All propeller controls (if applicable) in the takeoff pos ition;
(iii) The landing gear retracted; and
(iv) The airplane trimmed at the prescribed initial flight condition.
(2) With the wing flaps retracted at a speed of 1.3 V S with:
(i) All engines operating at maximum continuous power or thrust prior to failure of the critical engine;
(ii) All propeller controls in the enroute position;
(iii) The landing gear retracted; and
(iv) The airplane trimmed at the prescribed initial flight condition.
d. Airplane-Pilot Coupling (A-PC) .
(1) Explanation.
(i) Sections 25.143(a) and (b) require that the airplane be safely controllable and maneuverable without exceptional piloting skill and without danger of exceeding the airplane limiting load factor under any probable operating conditions. Service history events have indicated that modern transport category airplanes can be susceptible to airplane-pilot coupling under certain operating conditions and would not meet the intent of this requirement.
(ii) The classic airplane-pilot coupli ng (A-PC) situation, commonly referred to as "pilot-induced oscillations (PIO)," is considered to occur when some airplane's response metric is approximately 180 degrees out of phase with the pilot's control input. However, A-PC events with 180 degrees phase relationships are not the only conditions in which the airplane may exhibit closed-loop (pilot-in-the-loop) characteristics that are unacceptable for operation within the normal, operational, or limit flight envelopes. Others include unpredictability of the airplane's response to the pilot's control input. This may be due to nonlinearities in the control system, actuator rate or position limiting not sensed by the pilot through the flight controls, or changing pitch response at high altitude as the airplane maneuvers into and out of Mach buffet. Artificial trim and feel systems which produce controllers with too small a displacement and light force gradients may also lead to severe over control. This is especially true in a dynamic environment of high altitude turbulence or upsets in which the autopilot disconnects. This places the airplane in the hands of the unsuspecting pilot in conditions of only a small g or airspeed margin to buffet onset and with very low aerodynamic damping. These characteristics, while not 180 ° out of phase PIOs per se, may be hazardous and should be considered under the more general description of airplane-pilot coupling tendencies
(iii) Some of the A-PC tendency characteristics described in parag raph (ii) above are attributes of transport airplanes (e.g., low frequency short period, large response lags) that are recognized by part 25. Limits are placed on some of these individual attributes by part 25 (e.g., stick force per g, heavily damped short period) to assure satisfactory open-loop characteristics. However, service reports from recent years have indicated that certain operating envelope conditions, combined with triggering events, can result in airplane-pilot coupling incidents.
Some of the conditions that have led to these A-PCs include fuel management systems that permit extended operations with a center of gravity at or near the aft limit, operating at weight/speed/altitude conditions that result in reduced margins to buffet onset combined with tracking tasks such as not exceeding speed limitations and severe buffet due to load factor following an upset, and control surface rate or position limiting.
(iv) This service experience has shown that compliance with only the quantitative, op en-loop (pilot-out-of-the loop) requirements does not guarantee that the required levels of flying qualities are achieved.
Therefore, in order to ensure that the airplane has achieved the flying qualities required by §§ 25.143(a) and (b), the airplane must be evaluated by test pilots conducting high-gain (wide-bandwidth), closed-loop tasks to determine that the potential of encountering adverse A-PC tendencies is minimal.
(v) For the most part, these tasks must be performed in actual flight. However, for conditions that are considered too dangerous to attempt in actual flight (i.e., certain flight conditions outside of the operational flight envelope, flight in severe atmospheric disturbances, flight with certain failure states, etc.), the closed loop evaluation tasks may be performed using a motion base high fidelity simulator if it can be validated for the flight conditions of interest.
(2) Special Considerations .
(i) The certification team should understand the flight control system and airpla ne design.
(ii) The applicant should explain why the design is not conducive to an A-PC problem and how this is to be shown in both developmental and certification flight tests.
(iii) The applicant should explain what has been done during the devel opment flight test experience and any design changes that were required for A-PC problems.
(iv) The certification flight test program should be tailored to the specific airplane design and to evaluate the airplane in conditions that were found to be critical during its development program and A-PC analytical assessment.
(v) The FAA flight test pilots should also continuously evaluate the airplane for A -PC tendencies during the certification program in both the airplane and simulator. This evaluation should include both normal and malfunction states; all certification flight test points; transitions between and recoveries from these flight test points;
and normal, crosswind, and offset landing task evaluations.
(vi) Since the evaluation of flying qualities under §§ 25.143(a) and (b) is basically qualitative, especially evaluations of A-PC susceptibility, the high-gain tasks discussed herein should be accomplished by at least three test pilots. Use of other pilots can provide additional insights into the airplane handling qualities, but for the purpose of demonstrating compliance with this requirement the evaluation pilots should be trained test pilots.
(3) Procedures (Flight Test)
(i) Evaluation of the actual task performance achieved, e.g. , flight technical error, is not recommended as a measure of proof of compliance. Only the pilot's rating of the A-PC characteristics is needed as described in paragraph 20d(6). The tasks are used only to increase the pilot's gain, which is a prerequisite for exposing A-PC tendencies. Although task performance is not used as proof of compliance, task performance should be recorded and analyzed to insure that all pilots seem to be attempting to achieve the same level of performance.
(ii) Tasks for a s pecific certification project should be based on operational situations, flight testing maneuvers, or service difficulties that have produced A-PC events. Task requirements for a specific project will be dictated by the particular airplane and its specific areas of interest as determined by the tailored flight test program mentioned above. Some of these include high altitude upset maneuvers, encounters with turbulence at high altitude in which the autopilot disconnects, crosswind/crossed control landings with and without one engine inoperative, and offset landings to simulate the operational case in which the airplane breaks out of Instrument Meteorological Conditions (IMC) offset from the glideslope and/or localizer beam and the pilot makes a rapid alignment correction.
Tests should be conducted at or near the critical altitude/weight/c.g. combinations.
(iii) Tasks described here may be useful in any given evaluation and have proven to be operationally significant in the past. It is not intended that these are the only tasks that may be used or may be required depending on the scope and focus of the individual evaluation being conducted. Other tasks may be developed and used as appropriate. For example, some manufacturers have used formation tracking tasks successfully in the investigation of these tendencies. For all selected tasks, a build-up approach should be used and all end points should be approached with caution. Capture tasks and fine tracking tasks share many common characteristics but serve to highlight different aspects of any A-PC problem areas that may exist. In some cases, depending on individual airplane characteristics, it may be prudent to look at capture tasks first and then proceed to fine tracking tasks or combined gross acquisition (capture) and fine tracking tasks as appropriate.
(4) Capture Tasks .
(i) Capture tasks are intended to evaluate handling qualities for gross acquisition as opposed to continuous tracking. A wide variety of captures can be done provided the necessary cues are available to the pilot.
Pitch attitude, bank angle, heading, flight path angle, angle of attack, and g captures can be done to evaluate different aspects of the airplane response. These capture tasks can give the pilot a general impression of the handling qualities of the airplane, but because they do not involve closed-loop fine tracking, they do not expose all of the problems that may arise in fine tracking tasks. Capture tasks should not be used as the only evaluation tasks.
(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.