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4.11.3.5.6 The time sequence for the actions necessary to obtain the recommended level of propeller reverse thrust should be demonstrated by flight test. The time sequence used to determine the landing distances should reflect the most critical case relative to the time needed to obtain selected propeller reverse thrust.
4.11.3.5.7 The response times of the affected airplane systems to pilot inputs should be taken into account, for example, delays in system operation, such as interlocks and power lever detents that prevent the pilot from immediately selecting propeller reverse thrust. The effects of transient response characteristics, such as propeller reverse thrust engine spin-up, should also be included.
4.11.3.5.8 To enable a pilot of average skill to consistently obtain the recommended level of propeller reverse thrust under typical in-service conditions, a lever position that incorporates tactile feedback (e.g., a detent or stop) should be provided. If tactile feedback is not provided, a conservative level of propeller reverse thrust should be assumed.
4.11.3.5.9 The applicant should demonstrate that exceptional skill is not required to maintain directional control on a wet runway. The propeller reverse thrust procedures may specify a speed at which the propeller reverse thrust is cancelled in order to maintain directional controllability.
4.11.3.5.10 Compliance with the requirements of §§ 25.901(b)(2), 25.901(c),
25.1309(b), and 25.1309(c) will be accepted as providing compliance with the “safe and reliable” requirements of §§ 25.101(h)(2) and 25.125(c)(3).
4.11.4 Instrumentation and Data.
Instrumentation should include a means to record the airplane’s glide path relative to the ground, and the ground roll against time, in a manner that permits determining the horizontal and vertical distance time-histories. The appropriate data to permit analysis of these time-histories should also be recorded.
4.11.5 Landing on Unpaved Runways.
Guidance material for evaluation of landing on unpaved runways is contained in chapter 42 of this AC.
4.11.6 Automatic Braking Systems.
Guidance material relative to evaluation of auto-brake systems is provided in paragraph 15.4.9 of this AC.
4.11.7 AFM Landing Distances.
4.11.7.1 In accordance with § 25.101(i), AFM landing distances must be determined with all the airplane wheel brake assemblies at the fully worn limit of their allowable wear range. The brakes may be in any wear state
4-72 during the flight tests used to determine the landing distances, as long as a suitable combination of airplane and dynamometer tests is used to determine the landing distances corresponding to fully worn brakes.
Alternatively, the relationship between brake wear and stopping performance established during accelerate-stop testing may be used if it encompasses the brake wear conditions and energies achieved during the airplane flight tests used to establish the landing distances.
4.11.7.2 In deriving the scheduled distances, the time delays shown in figure 4-19 below should be assumed.
Figure 4-19. Landing Time Delays
4.11.7.2.1 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.
4.11.7.2.2 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.
4.11.7.2.3 Segment is repeated until pilot activation of all deceleration devices has been completed and the airplane is in the full braking configuration.
4.11.7.3 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.
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4.11.7.4 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.
4.11.7.5 Assumptions to be made in assessing the effect of wind on landing distance are discussed in paragraph 3.1 of this AC.
4.11.8 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 4.11.2.3.
4.11.8.1 Test data for each test point:
𝑅 𝑆50⁄ = 𝑅𝑎𝑡𝑒 𝑜𝑓 𝑠𝑖𝑛𝑘 𝑎𝑡 50 𝑓𝑒𝑒𝑡 𝑎𝑏𝑜𝑣𝑒 𝑙𝑎𝑛𝑑𝑖𝑛𝑔 𝑠𝑢𝑟𝑓𝑎𝑐𝑒 (𝑓𝑡/𝑠𝑒𝑐) 𝑅 𝑆𝑇𝐷⁄ = 𝑅𝑎𝑡𝑒 𝑜𝑓 𝑠𝑖𝑛𝑘 𝑎𝑡 𝑡𝑜𝑢𝑐ℎ𝑑𝑜𝑤𝑛 (𝑓𝑡/𝑠𝑒𝑐) 𝑉50 = 𝑇𝑟𝑢𝑒 𝑎𝑖𝑟𝑠𝑝𝑒𝑒𝑑 𝑎𝑡 50 𝑓𝑒𝑒𝑡 𝑎𝑏𝑜𝑣𝑒 𝑙𝑎𝑛𝑑𝑖𝑛𝑔 𝑠𝑢𝑟𝑓𝑎𝑐𝑒 (𝑓𝑡/𝑠𝑒𝑐) 𝑉𝑇𝐷 = 𝑇𝑟𝑢𝑒 𝑎𝑖𝑟𝑠𝑝𝑒𝑒𝑑 𝑎𝑡 𝑡𝑜𝑢𝑐ℎ𝑑𝑜𝑤𝑛 (𝑓𝑡/𝑠𝑒𝑐) 𝑡 = 𝐴𝑖𝑟 𝑡𝑖𝑚𝑒 50 𝑓𝑒𝑒𝑡 𝑡𝑜 𝑡𝑜𝑢𝑐ℎ𝑑𝑜𝑤𝑛 (𝑠𝑒𝑐)
4.11.8.2 The multiple linear regression analysis as outlined below is used to solve for the constants in the following equation:
50 𝑡⁄ = 𝑎 + 𝑏(𝑅 𝑆50⁄ ) + 𝑐(𝑅 𝑆𝑇𝐷⁄ )
4.11.8.3 The form of the dependent variable being solved in the above equation is 50/t, rather than just t, in order to maintain the same units for all variables.
4.11.8.4 The test values of all the test points, 1 through n, are used to determine the constants a, b, and c in the above equation as follows, where n equals the number of test points and R1 through R13 are the regression coefficients:
𝑅1 = ∑ 𝑅 𝑆50⁄ 𝑛 1
𝑅2 = ∑(𝑅 𝑆50⁄ )2 𝑛 1
𝑅3 = ∑ 𝑅 𝑆𝑇𝐷⁄ 𝑛 1
𝑅4 = ∑(𝑅 𝑆𝑇𝐷⁄ )2 𝑛 1
4-74 𝑅5 = ∑(𝑅 𝑆50⁄ )(𝑅 𝑆𝑇𝐷⁄ ) 𝑛 1
𝑅6 = ∑(50 𝑡⁄ ) 𝑛 1
𝑅7 = ∑(𝑅 𝑆50⁄ )(50 𝑡⁄ ) 𝑛 1
𝑅8 = ∑(𝑅 𝑆𝑆𝑇𝐷⁄ )(50 𝑡⁄ ) 𝑛 1
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 𝑎 = [(𝑅9)(𝑅10) − (𝑅11)(𝑅12)]/[(𝑅9)(𝑅13) − (𝑅11)2] 𝑏 = [(𝑅12) − (𝑐)(𝑅11)]/𝑅9 𝑐 = [(𝑅6) − (𝑏)(𝑅1) − (𝑐)(𝑅3)]/𝑛
4.11.8.5 Using the same regression coefficient relationships, determine the values of the constants, a, b, and c, for the speed reduction between 50 feet and touchdown (V50/VTD) by using the value of (V50/VTD) for (50/t) for each test point.
4.11.8.6 After determining the values of the constants, use the above equation for (50/t) to calculate the time from 50 feet to touchdown for the target conditions of a -3.5° flight path angle and R/STD = 8 ft/sec. Use a value of (R/S50) calculated from the approach path and V50. Then, using the same equation, but substituting (V50/VTD) for (50/t) and using the constants determined for (V50/VTD), calculate (V50/VTD).
4.11.8.7 After VTD is determined (from V50/VTD and V50), the air distance may be determined for the average flare speed and air time.
4-75 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 + 0.3647(R/S50) + 0.4917(R/STD) V50/VTD = 1.05508 - 0.003198(R/S50) + 0.001684(R/STD) For conditions of V50 = 220 ft/sec; flight path = -3.5°; R/STD = 8.0 ft/sec; the results are:
R/S50 = 13.43 ft/sec V50/VTD = 1.0256 t = 5.063 sec Air distance = 1100 ft
5-1 CHAPTER 5. FLIGHT: CONTROLLABILITY AND MANEUVERABILITY
5.1 General—§ 25.143.
5.1.1 Explanation.
5.1.1.1 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 encountering a stall warning or other characteristics that might interfere with normal maneuvering, or 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.
5.1.1.2 The maximum forces given in the table in § 25.143(d) 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 during the landing flare or a go-around, or during changes of configuration or power/thrust resulting in a change of control force that needs to 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.
5.1.1.3 Short-term and long-term forces should be interpreted as follows:
5.1.1.3.1 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.
5.1.1.3.2 Long-term forces are those control forces that result from normal or failure conditions that cannot readily be trimmed out or eliminated.
5.1.1.4 In conducting the controllability and maneuverability tests to show compliance with § 25.143 at speeds between VMO/MMO and VFC/MFC, the airplane should be trimmed at VMO/MMO.
5.1.1.5 Modern wing designs can exhibit a significant reduction in maximum lift capability with increasing Mach number. The magnitude of this Mach
5-2 number effect depends on the design characteristics of the particular wing.
For wing designs with a large Mach number effect, the maximum bank angle that can be achieved while retaining an acceptable stall margin can be significantly reduced. Because the effect of Mach number can be significant, and because it can also vary greatly for different wing designs, the multiplying factors applied to VSR may be insufficient to ensure that adequate maneuvering capability exists at the minimum operating speeds.
To address this issue, § 25.143(h) was added by amendment 25-108 to require a minimum bank angle capability in a coordinated turn without encountering stall warning or any other characteristic (including the envelope protection features of fly-by-wire flight control systems or automatic power or thrust increases) that might interfere with normal maneuvering. The maneuvering requirements consist of the minimum bank angle capability the FAA deems adequate for the specified regimes of flight combined with additional bank angle capability to provide a safety margin for various operational factors. These operational factors include both potential environmental conditions (e.g., turbulence, wind gusts) and an allowance for piloting imprecision (e.g., inadvertent overshoots). The FAA considers the automatic application of power or thrust by an envelope protection feature to be a feature that might interfere with normal maneuvering because it will result in a speed increase and flight path deviation, as well as potentially increasing crew workload due to the unexpected power or thrust increase.
5.1.2 General Test Requirements.
5.1.2.1 Compliance with § 25.143 (a) through (g) 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(d). Allowance should be made for delays in the initiation of recovery action appropriate to the situation.
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5.1.2.2 Since § 25.143(h) involves a target speed, bank angle, and maximum value of thrust/power setting, not all flight test conditions to demonstrate compliance will necessarily result in a constant-altitude, thrust-limited turn. In cases with positive excess power or thrust, a climbing condition at the target bank and speed is acceptable. Alternately, if desired, the power or thrust may be reduced to less than the maximum allowed, so that compliance is shown with a completely stabilized, constant-altitude turn.
With the airplane stabilized in a coordinated turn, holding power or thrust and speed, increase bank angle at constant airspeed until compliance is shown. For cases with negative excess power or thrust (e.g., the landing configuration case), a constant-altitude slow-down maneuver at the target bank angle has been shown to be a suitable technique. With the airplane descending at VREF in wings-level flight on a 3° glide path, trim and throttle position are noted. The airplane is then accelerated to VREF + 10 to 20 knots in level flight. The original trim and throttle conditions are reset as the airplane is rolled into a constant-altitude slow-down turn at the target bank angle. Throttles can be manipulated between idle and the marked position to vary slow-down rate as desired. Compliance is shown when the airplane decelerates through VREF in the turn without encountering a stall warning or other characteristic that might interfere with normal maneuvering.
5.1.2.3 If stall warning is provided by an artificial stall warning system, the effect of production tolerances on the stall warning system should be considered when evaluating compliance with the maneuvering capability requirements of § 25.143(h). See paragraph 8.1.6.2.6 of this AC for more information.
5.1.3 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 5.1.3.1 and 5.1.3.2 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° 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.
5.1.3.1 At each takeoff flap setting at the initial all-engine climb speed (e.g., V2 + 10 knots) with:
5.1.3.1.1 All engines operating at maximum takeoff power or thrust prior to failure of the critical engine;
5-4
5.1.3.1.2 All propeller controls (if applicable) in the takeoff position;
5.1.3.1.3 The landing gear retracted; and
5.1.3.1.4 The airplane trimmed at the prescribed initial flight condition.
5.1.3.2 With the wing flaps retracted at a speed of 1.23 VSR with:
5.1.3.2.1 All engines operating at maximum continuous power or thrust prior to failure of the critical engine;
5.1.3.2.2 All propeller controls in the en route position;
5.1.3.2.3 The landing gear retracted; and
5.1.3.2.4 The airplane trimmed at the prescribed initial flight condition.
5.1.4 Pilot Induced Oscillations (PIO).
5.1.4.1 Explanation.
5.1.4.1.1 Section 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.
5.1.4.1.2 The classic PIO is considered to occur when an airplane’s response is approximately 180° out of phase with the pilot’s control input. However, PIO events with 180° 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 a classic 180° out of phase PIO per se, may be hazardous and should be considered under the more general description of airplane-pilot coupling tendencies.
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5.1.4.1.3 Some of the PIO tendency characteristics described in paragraph 5.1.4.1.2 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 PIOs include fuel management systems that permit extended operations with a CG 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.
5.1.4.1.4 This service experience has shown that compliance with only the quantitative, open-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 should be evaluated by test pilots conducting high-gain (wide-bandwidth), closed-loop tasks to determine that the potential of encountering adverse PIO tendencies is minimal.
5.1.4.1.5 For the most part, these tasks should 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.
5.1.4.2 Special Considerations.
5.1.4.2.1 The certification team should understand the flight control system and airplane design.
5.1.4.2.2 The applicant should explain why the design is not conducive to a PIO problem and how this is to be shown in both developmental and certification flight tests.
5.1.4.2.3 The applicant should explain what has been done during the development flight test experience and any design changes that were required for PIO problems.
5.1.4.2.4 The certification flight test program should be tailored to the specific airplane design and to evaluate the airplane in conditions that were found
5-6 to be critical during its development program and PIO analytical assessment.
5.1.4.2.5 The FAA flight test pilots should also continuously evaluate the airplane for PIO 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.
5.1.4.2.6 Since the evaluation of flying qualities under § 25.143(a) and (b) is basically qualitative, especially evaluations of PIO 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.
5.1.4.3 Procedures: Flight Test.
5.1.4.3.1 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 PIO characteristics is needed as described in paragraph 5.1.4.6. The tasks are used only to increase the pilot’s gain, which is a prerequisite for exposing PIO 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.
5.1.4.3.2 Tasks for a specific certification project should be based on operational situations, flight testing maneuvers, or service difficulties that have produced PIO 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/CG combinations.
5.1.4.3.3 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
5-7 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 PIO 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.
5.1.4.4 Capture Tasks.
5.1.4.4.1 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.
5.1.4.4.2 For pitch captures, the airplane is trimmed for a specified flight condition.
The pilot aggressively captures 5° pitch attitude (or 10° if the airplane is already trimmed above 5°). The pilot then makes a series of aggressive pitch captures of 5° increments in both directions, and then continues this procedure with 10° 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 re-trimmed for the specified condition before starting the next maneuver.
5.1.4.4.3 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° or less).
5.1.4.4.4 Bank angle captures are also commonly done using bank-to-bank rolls.
Starting from a 15° bank angle, the pilot aggressively rolls and captures the opposite 15° bank angle (total bank angle change of 30°). The pilot