16. LANDING CLIMB: ALL-ENGINES-OPERATING - § 25.119 .
a. Explanation. Section 25.119(a) states that the engines are to be set at the power or thrust that is available 8 seconds after initiating movement of the power or thrust controls from the minimum flight idle position to the goaround power or thrust setting. The procedures given below are for the determination of this maximum thrust for showing compliance with the climb requirements of § 25.119.
b. Procedures .
(1) The engines should be trimmed to the low side of the idle trim band, if applicable, as defined in the airplane maintenance manual. The effect of any variation in the idle fuel flow schedule for engines with electronic fuel controllers is typically negligible (but any such claim should be adequately substantiated).
(2) At the most adverse test altitude , not to exceed the maximum field elevation for which certification is sought plus 1,500 ft., and with the most adverse bleed configuration expected in normal operations, stabilize the airplane in level flight with symmetrical power on all engines, landing gear down, flaps in the landing position, at a speed of 1.3 VS0. Retard the throttle(s) of the test engine(s) to flight idle and determine the time needed to reach a stabilized r.p.m., as defined below, for the test engine(s) while maintaining level flight or the minimum rate of descent obtainable with the thrust of the remaining engine(s) not greater than maximum continuous thrust (MCT).
Engine flight idle r.p.m. is considered to be stabilized when the initial rapid deceleration of all rotors is completed.
This has usually been 8-20 seconds. This can be determined in the cockpit as the point where rapid movement of the tachometer ceases. For some airplanes it may be desirable to determine the deceleration time from plots of
r.p.m. versus time.
(3) For the critical air bleed configuration , stabilize the airplane in level flight with symmetric power on all engines, landing gear down, flaps in the landing position, at a speed of 1.3 V S0, simulating the estimated minimum climb-limited landing weights (goal is low airspeed for low idle r.p.m.) at an altitude sufficiently above the selected test altitude so that time to descend to the test altitude with the throttles closed equals the appropriate engine r.p.m. stabilization time determined in paragraph (2) above. Retard the throttles to the flight idle position and descend at 1.3 VS0 to approximately the test altitude. When the appropriate time has elapsed, rapidly advance the power or thrust controls to the go-around power or thrust setting. The power or thrust controls may first be advanced to the forward stop and then retarded to the go-around power or thrust setting. At the applicant’s option, additional less critical bleed configurations may be tested.
(4) The thrust that is available 8 seconds after the initiation of movement of the power or thrust controls from the minimum flight idle position, in accordance with paragraph (3) above, will be the maximum permitted for showing compliance with the landing climb requirements of § 25.119(a), and Section 4T. 119(a) of SR-422B (see Appendix 1) for each of the bleed combinations tested in accordance with paragraph (3) above. If AFM performance is presented such that there is no accountability for various bleed conditions, the thrust obtained with the most critical airbleed shall be used for landing climb performance for all operations. The effects of anti-ice bleed must be accounted for.
17. CLIMB: ONE-ENGINE-INOPERATIVE - § 25.121 .
a. Explanation. None.
b. Procedures .
(1) Two methods for establishing one-engine-inoperative climb performance follow:
(i) Reciprocal heading climbs are conducted at several thrust-to-weight conditions from which the performance for the AFM is extracted. These climbs are flown with the wings nominally level. Reciprocal climbs may not be necessary if inertial corrections (or another equivalent means) are applied to account for wind gradients.
(ii) Drag polars and one-engine-inoperative yaw drag data are obtained for expansion into AFM climb performance. These data are obtained with the wings nominally level. Reciprocal heading check climbs are conducted to verify the predicted climb performance. These check climbs may be flown with the wings maintained
in a near level attitude. Reciprocal climbs may not be necessary if inertial corrections (or another equivalent means) are applied to account for wind gradients.
(2) If full rudder with wings level cannot maintain constant heading, small bank angles into the operating engine(s), with full rudder, should be used to maintain constant heading. Unless the landing lights automatically retract with engine failure, testing should be conducted with the lights extended for § 25.121(a) Takeoff; landing gear extended, § 25.121(b) Takeoff; landing gear retracted, and § 25.121(d) Approach.
(3) The climb performance tests with landing gear extended, in accordance with § 25.121(a), may be conducted with the landing gear and gear doors in the position they finally achieve after “gear down” selection. The critical configuration for the landing gear extended climb is considered to be that which presents the largest frontal area to the local airflow. This would normally be with no weight on the landing gear (full strut extension and trucks tilted) and all gear doors open. However, since the takeoff path will be determined by measurement of continuous takeoffs, or checked by continuous takeoffs if constructed by the segmental method (Ref. § 25.111(d)), any nonconservatism arising from the gear doors “closed” climb data will be evident. Also, some measure of conservatism is added to the landing gear extended climb performance by the requirement of § 25.111(d)(3) for the takeoff path data to be based on the airplane’s performance without ground effect.
(4) If means, such as variable intake doors , are provided to control powerplant cooling air supply during takeoff, climb, and en route flight, they should be set in a position that will maintain the temperature of major powerplant components, engine fluids, etc., within the established limits. The effect of these procedures should be included in the climb performance of the airplane. These provisions apply for all ambient temperatures up to the highest operational temperature limit for which approval is desired. (Reference: § 25.1043)
(5) The latter parts of §§ 25.121(a)(1) and (b)(1) , which state “...unless there is a more critical power operating condition existing later along the flight path...” are intended to cover those cases similar to where a wet engine depletes its water and reverts to dry engine operation. This is not intended to cover normal altitude thrust lapse rates above the point where the landing gear is fully retracted. (Reference: Preamble to SR-422A)
(6) Section 25.121(d) requires that the stalling speed for the approach configuration, landing gear retracted, not exceed 110 percent of the stalling speed for the related landing configuration, landing gear extended.
This stall speed ratio requirement is to ensure that an adequate margin above the stall speed in the selected approach configuration is maintained during flap retraction. To achieve this stall speed spread requirement, it is permissible to arbitrarily increase the landing flap stalling speed, V S0, to show compliance. The AFM must, however, base the landing speed on the increased stalling speed, and the landing distance demonstrations and the AFM landing field length requirements must also be predicated on the increased speed. The stall warning requirements of § 25.207 must be established for the adjusted stalling speed. However, the § 25.203 stall characteristics requirements must still be met at the normal stall speed. (An alternative to raising the landing flap stall speed, V S0, is to simply increase the landing speed stall ratio to a value greater than 1.3.)
18. EN ROUTE FLIGHT PATHS - § 25.123 .
a. Explanation. This guidance is intended for showing compliance with the requirements of § 25.123 and application to the operating requirements of §§ 121.191 and 121.193, which specify the clearances over terrain and obstructions required of the net en route flight paths subsequent to the failure of one or two engines.
b. Procedures .
(1) Sufficient en route climb performance data should be presented in the AFM to permit the determination of the net climb gradient and the net flight path in accordance with §§ 25.123(b) and (c) for all gross weights, altitudes, and ambient temperatures within the operating limits of the airplane. This en route climb performance data should be presented for altitudes up to the all-engines-operating ceiling to permit the calculation of drift-down data in the event of an en route engine failure.
(2) Fuel Consumption Accountability . The effect of the variation of the airplane’s weight along the flight path due to the progressive consumption of fuel may be taken into account using fuel flow rates obtained from airplane manufacturers’ test data. If measured fuel flow data is unavailable, a conservative fuel flow rate not greater than 80 percent of the engine specification flow rate at maximum continuous thrust (MCT) may be used.
(3) The procedures and flight conditions upon which the en route flight path data are based should be provided to the flightcrew. Credit for fuel dumping, if available and included in the flightcrew procedures, may be used to achieve the performance capability presented in the AFM. A conservative analysis should be used in taking into account the ambient conditions of temperature and wind existing along the flight path. All performance should be based on the net flight path and with MCT on the operating engine(s).
19. LANDING - § 25.125 .
a. Explanation.
(1) The landing distance is the horizontal distance from the point at which the main gear of the airplane is 50 ft. above the landing surface (treated as a horizontal plane through the touchdown point) to the point at which the airplane is brought to a stop. (For water landings, a speed of approximately 3 knots is considered “stopped.”) In this AC, the distance is treated in two parts: the airborne distance from 50 ft. to touchdown, and the ground distance from touchdown to stop. The latter may be further subdivided into a transition phase and a full braking phase if the applicant prefers this method of analysis.
(2) The term VREF used in this AC means the landing threshold speed (i.e., speed at 50 ft. height) scheduled in the AFM for normal operations. The minimum value of V REF is specified in § 25.125(a)(2) as 1.3V S, which provides an adequate margin above the stall speed to allow for likely speed variations during an approach in low turbulence. If the landing demonstrations are unable to show the acceptability of the minimum approach speed, and the tests are predicated on the use of a V REF greater than the minimum 1.3 V S, the landing distance data presented in the AFM must be based upon the higher approach speed.
(3) The engines should be set to the high side of the flight idle trim band, if applicable, for the landing flight tests. The effect of any variation in the idle fuel flow schedule for engines with electronic fuel controllers is typically negligible (but any such claim should be adequately substantiated).
b. Procedures for Determination of the Airborne Distance . Three acceptable means of compliance are described in paragraphs (1), (2), and (3) below. These differ from the “traditional” method in which steep approaches and high touchdown sink rates were permitted. Such a demonstration of maximum performance is no longer considered acceptable. However, the distances obtained using that method have resulted in a satisfactory operational safety record. The methods given here allow credit for the amount of testing an applicant is prepared to conduct, such that if the method described in paragraph 19b(3) (the most complex) is chosen, distances typical of those from the “traditional” method should be obtained, but without incurring the associated risks during testing.
NOTE. If it is determined that the constraints on approach angle and touchdown rate-of-sink described in paragraphs (2) and (3), below, are not appropriate due to novel or unusual features of future transport category airplane design, new criteria may be established. Such a change would be acceptable only if it is determined that an equivalent level of safety to existing performance standards and operational procedures is maintained.
(1) Experience shows an upper bound to the Part 25 zero-wind airborne distances achieved in past certifications and, similarly, a minimum speed loss. These are approximated by the following:
Air Distance (feet) = 1.55 (V REF-80)1.35 +800 where VREF is in knots TAS Touchdown Speed = V REF-3 knots An applicant may choose to use these relationships to establish landing distance in lieu of measuring airborne distance and speed loss. If an applicant chooses to use these relationships, it must be shown by test or analysis that they do not result in air distances or touchdown speeds that are nonconservative.
(2) If an applicant chooses to measure airborne distance or time, at least six tests covering the landing weight range are required for each airplane configuration for which certification is desired. These tests should meet the following criteria:
(i) A stabilized approach, targeting a glideslop e of -3 degrees and an indicated airspeed of V REF, should be maintained for a sufficient time prior to reaching a height of 50 feet above the landing surface to simulate a continuous approach at this speed. During this time, there should be no appreciable change in the power setting, pitch attitude, or rate of descent. The average glideslope of all landings used to show compliance should not be steeper than -3 degrees.
(ii) Below 50 feet, there should be no nose depression by use of the longitudinal c ontrol and no change in configuration, except for reduction in power.
(iii) The target rate of descent at touchdown should not exceed 6 feet per second. Target values cannot be achieved precisely; however, the average touchdown rate of descent should not exceed 6 feet per second.
(3) If the applicant conducts enough tests to allow a parametric analysis (or equivalent method) that establishes, with sufficient confidence, the relationship between airborne distance (or time) as a function of the rates of descent at 50 feet and touchdown, the 14 CFR part 25 airborne distances may be based on an approach angle of -
3.5 degrees, and a touchdown sink rate of 8 feet per second (see paragraph 19g for a sample of this analysis method).
(i) The air distance or air time established by this method may not be less than 90 percent of the lowest demonstrated value obtained using the target values for approach angle and touchdown sink rate specified in paragraph (ii), below. Test data with approach angles steeper than -3.5 degrees, or touchdown sink rates greater than 8 feet per second, may not be used to satisfy this requirement.
(ii) In order to determine the parametric relationships, it is recommended that test targets should span approach angles from -2.5 degrees to –3.5 degrees, and sink rates at touchdown from 2-6 ft. per second. Target speed for all tests should be V REF.
(iii) Below 50 feet, there should be no nose depression by use of the longitudinal control and no change in configuration that requires action by the pilot, except for reduction in power.
(iv) If an acceptable method of analysis is developed by an applicant to statistically establish a satisfactory confidence level for the resulting parametric relationships, then 12 tests, in each aerodynamic configuration for which certification is desired, will be sufficient. More tests will be necessary if the distribution of the data does not give sufficient confidence in the parametric correlation. Past experience has shown that a total of 40 landings would establish a satisfactory confidence level without further analysis. Autolands may be included in the analysis but should not comprise more than half of the data points. If it is apparent that configuration is not a significant variable, all data may be included in a single parametric analysis.
(v) If an applicant proposes any other method as being equivalent to a parametric analysis, that method should be based on a developed mathematical model that employs performance-related variables such as thrust, attitude, angle of attack, and load factor to adequately reproduce the flight test trajectory and airspeed variation from the 50 foot point to touchdown. Such a mathematical model should be validated by not less than 12 tests in each aerodynamic configuration for which certification is desired, and be justified by a comparison of tested and calculated landing airborne distances.
(vi) For the same aerodynamic configuration as previously certificated--if new tests are necessary to substantiate performance to a weight higher than that permitted by the extrapolation limits of § 25.21(d), two landings per configuration will be required for each 5 percent increase in landing weight, with a maximum total requirement of six landings. These may be merged with previous certification tests for parametric analysis, whether the previous certification was conducted by this method or not. If a new aerodynamic configuration is proposed, the 12 tests per configuration described in paragraph (iv), above, must be conducted.
(vii) In calculating the AFM landing distances, the speed loss from 50 feet to touchdown, as a percentage of VREF, may be determined using the conditions described in paragraph 19b(3).
(4) Whichever method is chosen to establish airborne distances, satisfactory flight characteristics must be demonstrated in the flare maneuver when a final approach speed of V REF-5 knots is maintained down to 50 feet.
(i) Below 50 feet, the application of longitudinal control to initiate flar e should occur at the same altitude as for a normal “on-speed” landing; no nose depression should be made and power should not be increased to facilitate the flare.
(ii) All power levers should be in their minimum flight idle position prior to touchdow n.
(iii) The normal flare technique should be used such that the touchdown speed should be at least 5 knots less than the touchdown speed used to establish the landing distance and the rate of descent at touchdown should not be greater than 6 feet per second.
(iv) This demonstration must be performed at both maximum landing weight and near minimum landing weight.
(v) These VREF-5 knots landing demonstration must not require the use of high control forces or full control deflections.
c. Procedures for Determination of the Transition and Stopping Distances .
(1) The transition distance extends from the initial touchdown point to the point where all approved deceleration devices are operative. The stopping distance extends from the end of transition to the point where the airplane is stopped. The two phases may be combined if the applicant prefers this method of analysis.
(2) If sufficient data are not available , there should be a minimum of six landings in the primary landing configuration. Experience has shown that if sufficient data are available for the airplane model to account for variation of braking performance with weight, lift, drag, ground speed, torque limit, etc., at least two test runs are necessary for each configuration when correlation for multiple configurations is being shown.
(3) A series of at least six measured landing tests covering the landing weight range should be conducted on the same set of wheels, tires, and brakes in order to substantiate that excessive wear of wheel brakes and tires is not produced in accordance with the provisions of § 25.125(b). The landing tests should be conducted with the normal operating brake pressures for which the applicant desires approval. The main gear tire pressure should be set to not less than the maximum pressure desired for certification corresponding to the specific test weight. Longitudinal control and brake application procedures must be such that they can be consistently applied in a manner that permits the airplane to be de -rotated at a controlled rate to preclude an excessive nose gear touchdown rate and so that the requirements of §§ 25.125(a)(4) and (5) are met. Nose gear touchdown rates in the certification landing tests should not be greater than eight feet per second. Certification practice has not allowed manually applied brakes before all main gear wheels are firmly on the ground.
(4) Airplane operating procedures appropriate for determination of landing distance must be described in the performance section of the AFM.
(5) Propeller pitch position used in determining the normal all-engines-operating landing stopping distance should be established using the criteria of § 25.125(f) for those airplanes that may derive some deceleration benefit from operating engines. Section 25.125(f) states that if the landing distance determined using a “device” that depends on the operation of any engine would be “noticeably increased” when a landing is made with that engine inoperative, the landing distance must be determined with that engine inoperative, unless a “compensating means” will result in one-engine-inoperative landing distances not greater than those with all engines operating. Acceptable interpretations of the terms “device,” “noticeably increased,” and “compensating means” are described below.
(i) If the propeller produces drag at any speed during the stopping phase of the normal allengines-operating landing distance, the maximum drag from this “device” for which performance credit may be taken is that which results from a propeller pitch position that renders zero thrust at zero airspeed. If the normal operational ground idle setting produces negative thrust at zero airspeed, the all-engines-operating stopping distances should be determined using a special flight test power lever stop to limit the propeller blade angle.
(ii) Distances should be measured for landings made with the propeller feathered on one engine, and ground idle selected after touchdown on the operating engines. The airplane configuration for this test, including the ground idle power lever position, should be the same as that used for the all-engines-operating landing
distance determination. The nose wheel should be free to caster, as in V MCG tests, to simulate wet runway surface conditions. Differential braking may be used to maintain directional control. This testing should be conducted at a minimum of three weights that cover the expected range of operational landing weights. If the resulting distances do not exceed the all-engines-operating landing distances by more than two percent (2%), they are not “noticeably increased” and no further testing is required to take performance credit for all-engines-operating disking drag in Airplane Flight Manual (AFM) landing distances.
(iii) If the distances determined in paragraph (ii), above, are more than two percent greater than the all-engines-operating landing distances, it must be shown that a “compensating means” exists in order to take performance credit for the all-engines-operating disking drag. Reverse propeller thrust on the operating engines is considered a “compensating means” if the resulting landing distances, with one propeller feathered, are demonstrated to be not longer than those determined for all-engines-operating with disking drag. The airplane configuration for this test should be the same as that used for the all-engines-operating landing distance determination. The nose wheel should be free to caster, as in V MCG tests, to simulate wet runway surface conditions.
Differential braking may be used to maintain directional control. Reverse thrust should not be selected until one second after nose wheel touchdown. This testing should be conducted at a minimum of three weights that cover the expected range of operational landing weights.
d. 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.
e. Landing on Unpaved Runways . Guidance material for evaluation of landing on unpaved runways is contained in Chapter 8 of this AC.
f. Automatic Braking Systems. Guidance material relative to evaluation of auto-brake systems is provided in paragraph 55c(6) of this AC.
g. Airplane Flight Manual Landing Distances .
(l) As a minimum, the AFM must include data for standard temperature and zero runway gradient showing the variation of landing distance with weight (up to maximum takeoff weight), altitude, and wind. If the airplane is intended for operation under Part 121 of the FAR, the distances presented should include the operational field length factors for both dry and wet runways required by § 121.195.
* (2) 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 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. *
(3) In deriving the scheduled distances, the time delay shown below should be assumed.
(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 .