Title 14—Aeronautics and Space--Volume 1
PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES
§ 25.1   Applicability.
(a) This part prescribes airworthiness standards for the issue of type certificates, and changes to those certificates, for transport category airplanes.
(b) Each person who applies under Part 21 for such a certificate or change must show compliance with the applicable requirements in this part.
§ 25.2   Special retroactive requirements.
The following special retroactive requirements are applicable to an airplane for which the regulations referenced in the type certificate predate the sections specified below—
(a) Irrespective of the date of application, each applicant for a supplemental type certificate (or an amendment to a type certificate) involving an increase in passenger seating capacity to a total greater than that for which the airplane has been type certificated must show that the airplane concerned meets the requirements of:
(1) Sections 25.721(d), 25.783(g), 25.785(c), 25.803(c)(2) through (9), 25.803 (d) and (e), 25.807 (a), (c), and (d), 25.809 (f) and (h), 25.811, 25.812, 25.813 (a), (b), and (c), 25.815, 25.817, 25.853 (a) and (b), 25.855(a), 25.993(f), and 25.1359(c) in effect on October 24, 1967, and
(2) Sections 25.803(b) and 25.803(c)(1) in effect on April 23, 1969.
(b) Irrespective of the date of application, each applicant for a supplemental type certificate (or an amendment to a type certificate) for an airplane manufactured after October 16, 1987, must show that the airplane meets the requirements of § 25.807(c)(7) in effect on July 24, 1989.
(c) Compliance with subsequent revisions to the sections specified in paragraph (a) or (b) of this section may be elected or may be required in accordance with § 21.101(a) of this chapter.
§ 25.3   Special provisions for ETOPS type design approvals.
(a) Applicability. This section applies to an applicant for ETOPS type design approval of an airplane:
(1) That has an existing type certificate on February 15, 2007; or
(2) For which an application for an original type certificate was submitted before February 15, 2007.
(b) Airplanes with two engines. (1) For ETOPS type design approval of an airplane up to and including 180 minutes, an applicant must comply with § 25.1535, except that it need not comply with the following provisions of Appendix K, K25.1.4, of this part:
(i) K25.1.4(a), fuel system pressure and flow requirements;
(ii) K25.1.4(a)(3), low fuel alerting; and
(iii) K25.1.4(c), engine oil tank design.
(2) For ETOPS type design approval of an airplane beyond 180 minutes an applicant must comply with § 25.1535.
(c) Airplanes with more than two engines. An applicant for ETOPS type design approval must comply with § 25.1535 for an airplane manufactured on or after February 17, 2015, except that, for an airplane configured for a three person flight crew, the applicant need not comply with Appendix K, K25.1.4(a)(3), of this part, low fuel alerting.
§ 25.4   Definitions.
(a) For the purposes of this part, the following general definitions apply:
(1) Certification maintenance requirement means a required scheduled maintenance task established during the design certification of the airplane systems as an airworthiness limitation of the type certificate or supplemental type certificate.
(2) Significant latent failure is a latent failure that, in combination with one or more specific failures or events, would result in a hazardous or catastrophic failure condition.
(b) For purposes of this part, the following failure conditions, in order of increasing severity, apply:
(1) Major failure condition means a failure condition that would reduce the capability of the airplane or the ability of the flightcrew to cope with adverse operating conditions, to the extent that there would be—
(i) A significant reduction in safety margins or functional capabilities,
(ii) A physical discomfort or a significant increase in flightcrew workload or in conditions impairing the efficiency of the flightcrew,
(iii) Physical distress to passengers or cabin crew, possibly including injuries, or
(iv) An effect of similar severity.
(2) Hazardous failure condition means a failure condition that would reduce the capability of the airplane or the ability of the flightcrew to cope with adverse operating conditions, to the extent that there would be—
(i) A large reduction in safety margins or functional capabilities,
(ii) Physical distress or excessive workload such that the flightcrew cannot be relied upon to perform their tasks accurately or completely, or
(iii) Serious or fatal injuries to a relatively small number of persons other than the flightcrew.
(3) Catastrophic failure condition means a failure condition that would result in multiple fatalities, usually with the loss of the airplane.
(c) For purposes of this part, the following failure conditions in order of decreasing probability apply:
(1) Probable failure condition means a failure condition that is anticipated to occur one or more times during the entire operational life of each airplane of a given type.
(2) Remote failure condition means a failure condition that is not anticipated to occur to each airplane of a given type during its entire operational life, but which may occur several times during the total operational life of a number of airplanes of a given type.
(3) Extremely remote failure condition means a failure condition that is not anticipated to occur to each airplane of a given type during its entire operational life, but which may occur a few times during the total operational life of all airplanes of a given type.
(4) Extremely improbable failure condition means a failure condition that is not anticipated to occur during the total operational life of all airplanes of a given type.
§ 25.5   Incorporations by reference.
(a) The materials listed in this section are incorporated by reference in the corresponding sections noted. These incorporations by reference were approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR part 51. These materials are incorporated as they exist on the date of the approval, and notice of any change in these materials will be published in the Federal Register. The materials are available for purchase at the corresponding addresses noted below, and all are available for inspection at the National Archives and Records Administration (NARA). For information on the availability of this material at NARA, call 202-741-6030, or go to: http://www.archives.gov/federal-register/cfr/ibr-locations.html.
(b) The following materials are available for purchase from the following address: The National Technical Information Services (NTIS), Springfield, Virginia 22166.
(1) Fuel Tank Flammability Assessment Method User's Manual, dated May 2008, document number DOT/FAA/AR-05/8, IBR approved for § 25.981 and Appendix N. It can also be obtained at the following Web site: http://www.fire.tc.faa.gov/systems/fueltank/FTFAM.stm.
(2) [Reserved]
§ 25.21   Proof of compliance.
(a) Each requirement of this subpart must be met at each appropriate combination of weight and center of gravity within the range of loading conditions for which certification is requested. This must be shown—
(1) By tests upon an airplane of the type for which certification is requested, or by calculations based on, and equal in accuracy to, the results of testing; and
(2) By systematic investigation of each probable combination of weight and center of gravity, if compliance cannot be reasonably inferred from combinations investigated.
(b) [Reserved]
(c) The controllability, stability, trim, and stalling characteristics of the airplane must be shown for each altitude up to the maximum expected in operation.
(d) Parameters critical for the test being conducted, such as weight, loading (center of gravity and inertia), airspeed, power, and wind, must be maintained within acceptable tolerances of the critical values during flight testing.
(e) If compliance with the flight characteristics requirements is dependent upon a stability augmentation system or upon any other automatic or power-operated system, compliance must be shown with §§ 25.671 and 25.672.
(f) In meeting the requirements of §§ 25.105(d), 25.125, 25.233, and 25.237, the wind velocity must be measured at a height of 10 meters above the surface, or corrected for the difference between the height at which the wind velocity is measured and the 10-meter height.
(g) The requirements of this subpart associated with icing conditions apply only if the applicant is seeking certification for flight in icing conditions.
(1) Paragraphs (g)(3) and (4) of this section apply only to airplanes with one or both of the following attributes:
(i) Maximum takeoff gross weight is less than 60,000 lbs; or
(ii) The airplane is equipped with reversible flight controls.
(2) Each requirement of this subpart, except §§ 25.121(a), 25.123(c), 25.143(b)(1) and (2), 25.149, 25.201(c)(2), 25.239, and 25.251(b) through (e), must be met in the icing conditions specified in Appendix C of this part. Section 25.207(c) and (d) must be met in the landing configuration in the icing conditions specified in Appendix C, but need not be met for other configurations. Compliance must be shown using the ice accretions defined in part II of Appendix C of this part, assuming normal operation of the airplane and its ice protection system in accordance with the operating limitations and operating procedures established by the applicant and provided in the airplane flight manual.
(3) If the applicant does not seek certification for flight in all icing conditions defined in Appendix O of this part, each requirement of this subpart, except §§ 25.105, 25.107, 25.109, 25.111, 25.113, 25.115, 25.121, 25.123, 25.143(b)(1), (b)(2), and (c)(1), 25.149, 25.201(c)(2), 25.207(c), (d), and (e)(1), 25.239, and 25.251(b) through (e), must be met in the Appendix O icing conditions for which certification is not sought in order to allow a safe exit from those conditions. Compliance must be shown using the ice accretions defined in part II, paragraphs (b) and (d) of Appendix O, assuming normal operation of the airplane and its ice protection system in accordance with the operating limitations and operating procedures established by the applicant and provided in the airplane flight manual.
(4) If the applicant seeks certification for flight in any portion of the icing conditions of Appendix O of this part, each requirement of this subpart, except §§ 25.121(a), 25.123(c), 25.143(b)(1) and (2), 25.149, 25.201(c)(2), 25.239, and 25.251(b) through (e), must be met in the Appendix O icing conditions for which certification is sought. Section 25.207(c) and (d) must be met in the landing configuration in the Appendix O icing conditions for which certification is sought, but need not be met for other configurations. Compliance must be shown using the ice accretions defined in part II, paragraphs (c) and (d) of Appendix O, assuming normal operation of the airplane and its ice protection system in accordance with the operating limitations and operating procedures established by the applicant and provided in the airplane flight manual.
§ 25.23   Load distribution limits.
(a) Ranges of weights and centers of gravity within which the airplane may be safely operated must be established. If a weight and center of gravity combination is allowable only within certain load distribution limits (such as spanwise) that could be inadvertently exceeded, these limits and the corresponding weight and center of gravity combinations must be established.
(b) The load distribution limits may not exceed—
(1) The selected limits;
(2) The limits at which the structure is proven; or
(3) The limits at which compliance with each applicable flight requirement of this subpart is shown.
§ 25.25   Weight limits.
(a) Maximum weights. Maximum weights corresponding to the airplane operating conditions (such as ramp, ground or water taxi, takeoff, en route, and landing), environmental conditions (such as altitude and temperature), and loading conditions (such as zero fuel weight, center of gravity position and weight distribution) must be established so that they are not more than—
(1) The highest weight selected by the applicant for the particular conditions; or
(2) The highest weight at which compliance with each applicable structural loading and flight requirement is shown, except that for airplanes equipped with standby power rocket engines the maximum weight must not be more than the highest weight established in accordance with appendix E of this part; or
(3) The highest weight at which compliance is shown with the certification requirements of Part 36 of this chapter.
(b) Minimum weight. The minimum weight (the lowest weight at which compliance with each applicable requirement of this part is shown) must be established so that it is not less than—
(1) The lowest weight selected by the applicant;
(2) The design minimum weight (the lowest weight at which compliance with each structural loading condition of this part is shown); or
(3) The lowest weight at which compliance with each applicable flight requirement is shown.
§ 25.27   Center of gravity limits.
The extreme forward and the extreme aft center of gravity limitations must be established for each practicably separable operating condition. No such limit may lie beyond—
(a) The extremes selected by the applicant;
(b) The extremes within which the structure is proven; or
(c) The extremes within which compliance with each applicable flight requirement is shown.
§ 25.29   Empty weight and corresponding center of gravity.
(a) The empty weight and corresponding center of gravity must be determined by weighing the airplane with—
(1) Fixed ballast;
(2) Unusable fuel determined under § 25.959; and
(3) Full operating fluids, including—
(i) Oil;
(ii) Hydraulic fluid; and
(iii) Other fluids required for normal operation of airplane systems, except potable water, lavatory precharge water, and fluids intended for injection in the engine.
(b) The condition of the airplane at the time of determining empty weight must be one that is well defined and can be easily repeated.
§ 25.31   Removable ballast.
Removable ballast may be used on showing compliance with the flight requirements of this subpart.
§ 25.33   Propeller speed and pitch limits.
(a) The propeller speed and pitch must be limited to values that will ensure—
(1) Safe operation under normal operating conditions; and
(2) Compliance with the performance requirements of §§ 25.101 through 25.125.
(b) There must be a propeller speed limiting means at the governor. It must limit the maximum possible governed engine speed to a value not exceeding the maximum allowable r.p.m.
(c) The means used to limit the low pitch position of the propeller blades must be set so that the engine does not exceed 103 percent of the maximum allowable engine rpm or 99 percent of an approved maximum overspeed, whichever is greater, with—
(1) The propeller blades at the low pitch limit and governor inoperative;
(2) The airplane stationary under standard atmospheric conditions with no wind; and
(3) The engines operating at the takeoff manifold pressure limit for reciprocating engine powered airplanes or the maximum takeoff torque limit for turbopropeller engine-powered airplanes.
§ 25.101   General.
(a) Unless otherwise prescribed, airplanes must meet the applicable performance requirements of this subpart for ambient atmospheric conditions and still air.
(b) The performance, as affected by engine power or thrust, must be based on the following relative humidities;
(1) For turbine engine powered airplanes, a relative humidity of—
(i) 80 percent, at and below standard temperatures; and
(ii) 34 percent, at and above standard temperatures plus 50 °F.
(2) For reciprocating engine powered airplanes, a relative humidity of 80 percent in a standard atmosphere. Engine power corrections for vapor pressure must be made in accordance with the following table:
(c) The performance must correspond to the propulsive thrust available under the particular ambient atmospheric conditions, the particular flight condition, and the relative humidity specified in paragraph (b) of this section. The available propulsive thrust must correspond to engine power or thrust, not exceeding the approved power or thrust less—
(1) Installation losses; and
(2) The power or equivalent thrust absorbed by the accessories and services appropriate to the particular ambient atmospheric conditions and the particular flight condition.
(d) Unless otherwise prescribed, the applicant must select the takeoff, en route, approach, and landing configurations for the airplane.
(e) The airplane configurations may vary with weight, altitude, and temperature, to the extent they are compatible with the operating procedures required by paragraph (f) of this section.
(f) Unless otherwise prescribed, in determining the accelerate-stop distances, takeoff flight paths, takeoff distances, and landing distances, changes in the airplane's configuration, speed, power, and thrust, must be made in accordance with procedures established by the applicant for operation in service.
(g) Procedures for the execution of balked landings and missed approaches associated with the conditions prescribed in §§ 25.119 and 25.121(d) must be established.
(h) The procedures established under paragraphs (f) and (g) of this section must—
(1) Be able to be consistently executed in service by crews of average skill;
(2) Use methods or devices that are safe and reliable; and
(3) Include allowance for any time delays, in the execution of the procedures, that may reasonably be expected in service.
(i) The accelerate-stop and landing distances prescribed in §§ 25.109 and 25.125, respectively, must be determined with all the airplane wheel brake assemblies at the fully worn limit of their allowable wear range.
§ 25.103   Stall speed.
(a) The reference stall speed, V SR , is a calibrated airspeed defined by the applicant. V SR may not be less than a 1-g stall speed. V SR is expressed as:
is first a maximum during the maneuver prescribed in paragraph (c) of this section. In addition, when the maneuver is limited by a device that abruptly pushes the nose down at a selected angle of attack (e.g., a stick pusher), V CL MAX may not be less than the speed existing at the instant the device operates;
(b) V CL MAX is determined with:
(1) Engines idling, or, if that resultant thrust causes an appreciable decrease in stall speed, not more than zero thrust at the stall speed;
(2) Propeller pitch controls (if applicable) in the takeoff position;
(3) The airplane in other respects (such as flaps, landing gear, and ice accretions) in the condition existing in the test or performance standard in which V SR is being used;
(4) The weight used when V SR is being used as a factor to determine compliance with a required performance standard;
(5) The center of gravity position that results in the highest value of reference stall speed; and
(6) The airplane trimmed for straight flight at a speed selected by the applicant, but not less than 1.13V SR and not greater than 1.3V SR .
(c) Starting from the stabilized trim condition, apply the longitudinal control to decelerate the airplane so that the speed reduction does not exceed one knot per second.
(d) In addition to the requirements of paragraph (a) of this section, when a device that abruptly pushes the nose down at a selected angle of attack (e.g., a stick pusher) is installed, the reference stall speed, V SR , may not be less than 2 knots or 2 percent, whichever is greater, above the speed at which the device operates.
§ 25.105   Takeoff.
(a) The takeoff speeds prescribed by § 25.107, the accelerate-stop distance prescribed by § 25.109, the takeoff path prescribed by § 25.111, the takeoff distance and takeoff run prescribed by § 25.113, and the net takeoff flight path prescribed by § 25.115, must be determined in the selected configuration for takeoff at each weight, altitude, and ambient temperature within the operational limits selected by the applicant—
(1) In non-icing conditions; and
(2) In icing conditions, if in the configuration used to show compliance with § 25.121(b), and with the most critical of the takeoff ice accretion(s) defined in appendices C and O of this part, as applicable, in accordance with § 25.21(g):
(i) The stall speed at maximum takeoff weight exceeds that in non-icing conditions by more than the greater of 3 knots CAS or 3 percent of V SR ; or
(ii) The degradation of the gradient of climb determined in accordance with § 25.121(b) is greater than one-half of the applicable actual-to-net takeoff flight path gradient reduction defined in § 25.115(b).
(b) No takeoff made to determine the data required by this section may require exceptional piloting skill or alertness.
(c) The takeoff data must be based on—
(1) In the case of land planes and amphibians:
(i) Smooth, dry and wet, hard-surfaced runways; and
(ii) At the option of the applicant, grooved or porous friction course wet, hard-surfaced runways.
(2) Smooth water, in the case of seaplanes and amphibians; and
(3) Smooth, dry snow, in the case of skiplanes.
(d) The takeoff data must include, within the established operational limits of the airplane, the following operational correction factors:
(1) Not more than 50 percent of nominal wind components along the takeoff path opposite to the direction of takeoff, and not less than 150 percent of nominal wind components along the takeoff path in the direction of takeoff.
(2) Effective runway gradients.
§ 25.107   Takeoff speeds.
(a) V 1 must be established in relation to V EF as follows:
(1) V EF is the calibrated airspeed at which the critical engine is assumed to fail. V EF must be selected by the applicant, but may not be less than V MCG determined under § 25.149(e).
(2) V 1 , in terms of calibrated airspeed, is selected by the applicant; however, V 1 may not be less than V EF plus the speed gained with critical engine inoperative during the time interval between the instant at which the critical engine is failed, and the instant at which the pilot recognizes and reacts to the engine failure, as indicated by the pilot's initiation of the first action (e.g., applying brakes, reducing thrust, deploying speed brakes) to stop the airplane during accelerate-stop tests.
(b) V 2MIN, in terms of calibrated airspeed, may not be less than—
(1) 1.13 V SR for—
(i) Two-engine and three-engine turbopropeller and reciprocating engine powered airplanes; and
(ii) Turbojet powered airplanes without provisions for obtaining a significant reduction in the one-engine-inoperative power-on stall speed;
(2) 1.08 V SR for—
(i) Turbopropeller and reciprocating engine powered airplanes with more than three engines; and
(ii) Turbojet powered airplanes with provisions for obtaining a significant reduction in the one-engine-inoperative power-on stall speed; and
(3) 1.10 times V MC established under § 25.149.
(c) V 2 , in terms of calibrated airspeed, must be selected by the applicant to provide at least the gradient of climb required by § 25.121(b) but may not be less than—
(1) V 2MIN ;
(2) V R plus the speed increment attained (in accordance with § 25.111(c)(2)) before reaching a height of 35 feet above the takeoff surface; and
(3) A speed that provides the maneuvering capability specified in § 25.143(h).
(d) V MU is the calibrated airspeed at and above which the airplane can safely lift off the ground, and con- tinue the takeoff. V MU speeds must be selected by the applicant throughout the range of thrust-to-weight ratios to be certificated. These speeds may be established from free air data if these data are verified by ground takeoff tests.
(e) V R, in terms of calibrated airspeed, must be selected in accordance with the conditions of paragraphs (e)(1) through (4) of this section:
(1) V R may not be less than—
(i) V 1 ;
(ii) 105 percent of V MC ;
(iii) The speed (determined in accordance with § 25.111(c)(2)) that allows reaching V 2 before reaching a height of 35 feet above the takeoff surface; or
(iv) A speed that, if the airplane is rotated at its maximum practicable rate, will result in a V LOF of not less than —
(A) 110 percent of V MU in the all-engines-operating condition, and 105 percent of V MU determined at the thrust-to-weight ratio corresponding to the one-engine-inoperative condition; or
(B) If the V MU attitude is limited by the geometry of the airplane ( i.e., tail contact with the runway), 108 percent of V MU in the all-engines-operating condition, and 104 percent of V MU determined at the thrust-to-weight ratio corresponding to the one-engine-inoperative condition.
(2) For any given set of conditions (such as weight, configuration, and temperature), a single value of V R, obtained in accordance with this paragraph, must be used to show compliance with both the one-engine-inoperative and the all-engines-operating takeoff provisions.
(3) It must be shown that the one-engine-inoperative takeoff distance, using a rotation speed of 5 knots less than V R established in accordance with paragraphs (e)(1) and (2) of this section, does not exceed the corresponding one-engine-inoperative takeoff distance using the established V R . The takeoff distances must be determined in accordance with § 25.113(a)(1).
(4) Reasonably expected variations in service from the established takeoff procedures for the operation of the airplane (such as over-rotation of the airplane and out-of-trim conditions) may not result in unsafe flight characteristics or in marked increases in the scheduled takeoff distances established in accordance with § 25.113(a).
(f) V LOF is the calibrated airspeed at which the airplane first becomes airborne.
(g) V FTO , in terms of calibrated airspeed, must be selected by the applicant to provide at least the gradient of climb required by § 25.121(c), but may not be less than—
(1) 1.18 V SR ; and
(2) A speed that provides the maneuvering capability specified in § 25.143(h).
(h) In determining the takeoff speeds V 1 , V R , and V 2 for flight in icing conditions, the values of V MCG , V MC , and V MU determined for non-icing conditions may be used.
§ 25.109   Accelerate-stop distance.
(a) The accelerate-stop distance on a dry runway is the greater of the following distances:
(1) The sum of the distances necessary to—
(i) Accelerate the airplane from a standing start with all engines operating to V EF for takeoff from a dry runway;
(ii) Allow the airplane to accelerate from V EF to the highest speed reached during the rejected takeoff, assuming the critical engine fails at V EF and the pilot takes the first action to reject the takeoff at the V 1 for takeoff from a dry runway; and
(iii) Come to a full stop on a dry runway from the speed reached as prescribed in paragraph (a)(1)(ii) of this section; plus
(iv) A distance equivalent to 2 seconds at the V 1 for takeoff from a dry runway.
(2) The sum of the distances necessary to—
(i) Accelerate the airplane from a standing start with all engines operating to the highest speed reached during the rejected takeoff, assuming the pilot takes the first action to reject the takeoff at the V 1 for takeoff from a dry runway; and
(ii) With all engines still operating, come to a full stop on dry runway from the speed reached as prescribed in paragraph (a)(2)(i) of this section; plus
(iii) A distance equivalent to 2 seconds at the V 1 for takeoff from a dry runway.
(b) The accelerate-stop distance on a wet runway is the greater of the following distances:
(1) The accelerate-stop distance on a dry runway determined in accordance with paragraph (a) of this section; or
(2) The accelerate-stop distance determined in accordance with paragraph (a) of this section, except that the runway is wet and the corresponding wet runway values of V EF and V 1 are used. In determining the wet runway accelerate-stop distance, the stopping force from the wheel brakes may never exceed:
(i) The wheel brakes stopping force determined in meeting the requirements of § 25.101(i) and paragraph (a) of this section; and
(ii) The force resulting from the wet runway braking coefficient of friction determined in accordance with paragraphs (c) or (d) of this section, as applicable, taking into account the distribution of the normal load between braked and unbraked wheels at the most adverse center-of-gravity position approved for takeoff.
(c) The wet runway braking coefficient of friction for a smooth wet runway is defined as a curve of friction coefficient versus ground speed and must be computed as follows:
(1) The maximum tire-to-ground wet runway braking coefficient of friction is defined as:
(2) The maximum tire-to-ground wet runway braking coefficient of friction must be adjusted to take into account the efficiency of the anti-skid system on a wet runway. Anti-skid system operation must be demonstrated by flight testing on a smooth wet runway, and its efficiency must be determined. Unless a specific anti-skid system efficiency is determined from a quantitative analysis of the flight testing on a smooth wet runway, the maximum tire-to-ground wet runway braking coefficient of friction determined in paragraph (c)(1) of this section must be multiplied by the efficiency value associated with the type of anti-skid system installed on the airplane:
(d) At the option of the applicant, a higher wet runway braking coefficient of friction may be used for runway surfaces that have been grooved or treated with a porous friction course material. For grooved and porous friction course runways, the wet runway braking coefficent of friction is defined as either:
(1) 70 percent of the dry runway braking coefficient of friction used to determine the dry runway accelerate-stop distance; or
(2) The wet runway braking coefficient defined in paragraph (c) of this section, except that a specific anti-skid system efficiency, if determined, is appropriate for a grooved or porous friction course wet runway, and the maximum tire-to-ground wet runway braking coefficient of friction is defined as:
(e) Except as provided in paragraph (f)(1) of this section, means other than wheel brakes may be used to determine the accelerate-stop distance if that means—
(1) Is safe and reliable;
(2) Is used so that consistent results can be expected under normal operating conditions; and
(3) Is such that exceptional skill is not required to control the airplane.
(f) The effects of available reverse thrust—
(1) Shall not be included as an additional means of deceleration when determining the accelerate-stop distance on a dry runway; and
(2) May be included as an additional means of deceleration using recommended reverse thrust procedures when determining the accelerate-stop distance on a wet runway, provided the requirements of paragraph (e) of this section are met.
(g) The landing gear must remain extended throughout the accelerate-stop distance.
(h) If the accelerate-stop distance includes a stopway with surface characteristics substantially different from those of the runway, the takeoff data must include operational correction factors for the accelerate-stop distance. The correction factors must account for the particular surface characteristics of the stopway and the variations in these characteristics with seasonal weather conditions (such as temperature, rain, snow, and ice) within the established operational limits.
(i) A flight test demonstration of the maximum brake kinetic energy accelerate-stop distance must be conducted with not more than 10 percent of the allowable brake wear range remaining on each of the airplane wheel brakes.
§ 25.111   Takeoff path.
(a) The takeoff path extends from a standing start to a point in the takeoff at which the airplane is 1,500 feet above the takeoff surface, or at which the transition from the takeoff to the en route configuration is completed and V FTO is reached, whichever point is higher. In addition—
(1) The takeoff path must be based on the procedures prescribed in § 25.101(f);
(2) The airplane must be accelerated on the ground to V EF, at which point the critical engine must be made inoperative and remain inoperative for the rest of the takeoff; and
(3) After reaching V EF, the airplane must be accelerated to V 2 .
(b) During the acceleration to speed V 2 , the nose gear may be raised off the ground at a speed not less than V R . However, landing gear retraction may not be begun until the airplane is airborne.
(c) During the takeoff path determination in accordance with paragraphs (a) and (b) of this section—
(1) The slope of the airborne part of the takeoff path must be positive at each point;
(2) The airplane must reach V 2 before it is 35 feet above the takeoff surface and must continue at a speed as close as practical to, but not less than V 2 , until it is 400 feet above the takeoff surface;
(3) At each point along the takeoff path, starting at the point at which the airplane reaches 400 feet above the takeoff surface, the available gradient of climb may not be less than—
(i) 1.2 percent for two-engine airplanes;
(ii) 1.5 percent for three-engine airplanes; and
(iii) 1.7 percent for four-engine airplanes.
(4) The airplane configuration may not be changed, except for gear retraction and automatic propeller feathering, and no change in power or thrust that requires action by the pilot may be made until the airplane is 400 feet above the takeoff surface; and
(5) If § 25.105(a)(2) requires the takeoff path to be determined for flight in icing conditions, the airborne part of the takeoff must be based on the airplane drag:
(i) With the most critical of the takeoff ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), from a height of 35 feet above the takeoff surface up to the point where the airplane is 400 feet above the takeoff surface; and
(ii) With the most critical of the final takeoff ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), from the point where the airplane is 400 feet above the takeoff surface to the end of the takeoff path.
(d) The takeoff path must be determined by a continuous demonstrated takeoff or by synthesis from segments. If the takeoff path is determined by the segmental method—
(1) The segments must be clearly defined and must be related to the distinct changes in the configuration, power or thrust, and speed;
(2) The weight of the airplane, the configuration, and the power or thrust must be constant throughout each segment and must correspond to the most critical condition prevailing in the segment;
(3) The flight path must be based on the airplane's performance without ground effect; and
(4) The takeoff path data must be checked by continuous demonstrated takeoffs up to the point at which the airplane is out of ground effect and its speed is stabilized, to ensure that the path is conservative relative to the continous path.
(e) For airplanes equipped with standby power rocket engines, the takeoff path may be determined in accordance with section II of appendix E.
§ 25.113   Takeoff distance and takeoff run.
(a) Takeoff distance on a dry runway is the greater of—
(1) The horizontal distance along the takeoff path from the start of the takeoff to the point at which the airplane is 35 feet above the takeoff surface, determined under § 25.111 for a dry runway; or
(2) 115 percent of the horizontal distance along the takeoff path, with all engines operating, from the start of the takeoff to the point at which the airplane is 35 feet above the takeoff surface, as determined by a procedure consistent with § 25.111.
(b) Takeoff distance on a wet runway is the greater of—
(1) The takeoff distance on a dry runway determined in accordance with paragraph (a) of this section; or
(2) The horizontal distance along the takeoff path from the start of the takeoff to the point at which the airplane is 15 feet above the takeoff surface, achieved in a manner consistent with the achievement of V 2 before reaching 35 feet above the takeoff surface, determined under § 25.111 for a wet runway.
(c) If the takeoff distance does not include a clearway, the takeoff run is equal to the takeoff distance. If the takeoff distance includes a clearway—
(1) The takeoff run on a dry runway is the greater of—
(i) The horizontal distance along the takeoff path from the start of the takeoff to a point equidistant between the point at which V LOF is reached and the point at which the airplane is 35 feet above the takeoff surface, as determined under § 25.111 for a dry runway; or
(ii) 115 percent of the horizontal distance along the takeoff path, with all engines operating, from the start of the takeoff to a point equidistant between the point at which V LOF is reached and the point at which the airplane is 35 feet above the takeoff surface, determined by a procedure consistent with § 25.111.
(2) The takeoff run on a wet runway is the greater of—
(i) The horizontal distance along the takeoff path from the start of the takeoff to the point at which the airplane is 15 feet above the takeoff surface, achieved in a manner consistent with the achievement of V 2 before reaching 35 feet above the takeoff surface, as determined under § 25.111 for a wet runway; or
(ii) 115 percent of the horizontal distance along the takeoff path, with all engines operating, from the start of the takeoff to a point equidistant between the point at which V LOF is reached and the point at which the airplane is 35 feet above the takeoff surface, determined by a procedure consistent with § 25.111.
§ 25.115   Takeoff flight path.
(a) The takeoff flight path shall be considered to begin 35 feet above the takeoff surface at the end of the takeoff distance determined in accordance with § 25.113(a) or (b), as appropriate for the runway surface condition.
(b) The net takeoff flight path data must be determined so that they represent the actual takeoff flight paths (determined in accordance with § 25.111 and with paragraph (a) of this section) reduced at each point by a gradient of climb equal to—
(1) 0.8 percent for two-engine airplanes;
(2) 0.9 percent for three-engine airplanes; and
(3) 1.0 percent for four-engine airplanes.
(c) The prescribed reduction in climb gradient may be applied as an equivalent reduction in acceleration along that part of the takeoff flight path at which the airplane is accelerated in level flight.
§ 25.117   Climb: general.
Compliance with the requirements of §§ 25.119 and 25.121 must be shown at each weight, altitude, and ambient temperature within the operational limits established for the airplane and with the most unfavorable center of gravity for each configuration.
§ 25.119   Landing climb: All-engines-operating.
In the landing configuration, the steady gradient of climb may not be less than 3.2 percent, with the engines at the power or thrust that is available 8 seconds after initiation of movement of the power or thrust controls from the minimum flight idle to the go-around power or thrust setting—
(a) In non-icing conditions, with a climb speed of V REF determined in accordance with § 25.125(b)(2)(i); and
(b) In icing conditions with the most critical of the landing ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), and with a climb speed of V REF determined in accordance with § 25.125(b)(2)(ii).
§ 25.121   Climb: One-engine-inoperative.
(a) Takeoff; landing gear extended. In the critical takeoff configuration existing along the flight path (between the points at which the airplane reaches V LOF and at which the landing gear is fully retracted) and in the configuration used in § 25.111 but without ground effect, the steady gradient of climb must be positive for two-engine airplanes, and not less than 0.3 percent for three-engine airplanes or 0.5 percent for four-engine airplanes, at V LOF and with—
(1) The critical engine inoperative and the remaining engines at the power or thrust available when retraction of the landing gear is begun in accordance with § 25.111 unless there is a more critical power operating condition existing later along the flight path but before the point at which the landing gear is fully retracted; and
(2) The weight equal to the weight existing when retraction of the landing gear is begun, determined under § 25.111.
(b) Takeoff; landing gear retracted. In the takeoff configuration existing at the point of the flight path at which the landing gear is fully retracted, and in the configuration used in § 25.111 but without ground effect:
(1) The steady gradient of climb may not be less than 2.4 percent for two-engine airplanes, 2.7 percent for three-engine airplanes, and 3.0 percent for four-engine airplanes, at V 2 with:
(i) The critical engine inoperative, the remaining engines at the takeoff power or thrust available at the time the landing gear is fully retracted, determined under § 25.111, unless there is a more critical power operating condition existing later along the flight path but before the point where the airplane reaches a height of 400 feet above the takeoff surface; and
(ii) The weight equal to the weight existing when the airplane's landing gear is fully retracted, determined under § 25.111.
(2) The requirements of paragraph (b)(1) of this section must be met:
(i) In non-icing conditions; and
(ii) In icing conditions with the most critical of the takeoff ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), if in the configuration used to show compliance with § 25.121(b) with this takeoff ice accretion:
(A) The stall speed at maximum takeoff weight exceeds that in non-icing conditions by more than the greater of 3 knots CAS or 3 percent of V SR ; or
(B) The degradation of the gradient of climb determined in accordance with § 25.121(b) is greater than one-half of the applicable actual-to-net takeoff flight path gradient reduction defined in § 25.115(b).
(c) Final takeoff. In the en route configuration at the end of the takeoff path determined in accordance with § 25.111:
(1) The steady gradient of climb may not be less than 1.2 percent for two-engine airplanes, 1.5 percent for three-engine airplanes, and 1.7 percent for four-engine airplanes, at V FTO with—
(i) The critical engine inoperative and the remaining engines at the available maximum continuous power or thrust; and
(ii) The weight equal to the weight existing at the end of the takeoff path, determined under § 25.111.
(2) The requirements of paragraph (c)(1) of this section must be met:
(i) In non-icing conditions; and
(ii) In icing conditions with the most critical of the final takeoff ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), if in the configuration used to show compliance with § 25.121(b) with the takeoff ice accretion used to show compliance with § 25.111(c)(5)(i):
(A) The stall speed at maximum takeoff weight exceeds that in non-icing conditions by more than the greater of 3 knots CAS or 3 percent of V SR ; or
(B) The degradation of the gradient of climb determined in accordance with § 25.121(b) is greater than one-half of the applicable actual-to-net takeoff flight path gradient reduction defined in § 25.115(b).
(d) Approach. In a configuration corresponding to the normal all-engines-operating procedure in which V SR for this configuration does not exceed 110 percent of the V SR for the related all-engines-operating landing configuration:
(1) The steady gradient of climb may not be less than 2.1 percent for two-engine airplanes, 2.4 percent for three-engine airplanes, and 2.7 percent for four-engine airplanes, with—
(i) The critical engine inoperative, the remaining engines at the go-around power or thrust setting;
(ii) The maximum landing weight;
(iii) A climb speed established in connection with normal landing procedures, but not exceeding 1.4 V SR ; and
(iv) Landing gear retracted.
(2) The requirements of paragraph (d)(1) of this section must be met:
(i) In non-icing conditions; and
(ii) In icing conditions with the most critical of the approach ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g). The climb speed selected for non-icing conditions may be used if the climb speed for icing conditions, computed in accordance with paragraph (d)(1)(iii) of this section, does not exceed that for non-icing conditions by more than the greater of 3 knots CAS or 3 percent.
§ 25.123   En route flight paths.
(a) For the en route configuration, the flight paths prescribed in paragraph (b) and (c) of this section must be determined at each weight, altitude, and ambient temperature, within the operating limits established for the airplane. The variation of weight along the flight path, accounting for the progressive consumption of fuel and oil by the operating engines, may be included in the computation. The flight paths must be determined at a speed not less than V FTO , with—
(1) The most unfavorable center of gravity;
(2) The critical engines inoperative;
(3) The remaining engines at the available maximum continuous power or thrust; and
(4) The means for controlling the engine-cooling air supply in the position that provides adequate cooling in the hot-day condition.
(b) The one-engine-inoperative net flight path data must represent the actual climb performance diminished by a gradient of climb of 1.1 percent for two-engine airplanes, 1.4 percent for three-engine airplanes, and 1.6 percent for four-engine airplanes—
(1) In non-icing conditions; and
(2) In icing conditions with the most critical of the en route ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), if:
(i) A speed of 1.18 “V SR0 with the en route ice accretion exceeds the en route speed selected for non-icing conditions by more than the greater of 3 knots CAS or 3 percent of V SR ; or
(ii) The degradation of the gradient of climb is greater than one-half of the applicable actual-to-net flight path reduction defined in paragraph (b) of this section.
(c) For three- or four-engine airplanes, the two-engine-inoperative net flight path data must represent the actual climb performance diminished by a gradient of climb of 0.3 percent for three-engine airplanes and 0.5 percent for four-engine airplanes.
§ 25.125   Landing.
(a) The horizontal distance necessary to land and to come to a complete stop (or to a speed of approximately 3 knots for water landings) from a point 50 feet above the landing surface must be determined (for standard temperatures, at each weight, altitude, and wind within the operational limits established by the applicant for the airplane):
(1) In non-icing conditions; and
(2) In icing conditions with the most critical of the landing ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), if V REF for icing conditions exceeds V REF for non-icing conditions by more than 5 knots CAS at the maximum landing weight.
(b) In determining the distance in paragraph (a) of this section:
(1) The airplane must be in the landing configuration.
(2) A stabilized approach, with a calibrated airspeed of not less than V REF , must be maintained down to the 50-foot height.
(i) In non-icing conditions, V REF may not be less than:
(A) 1.23 V SR 0;
(B) V MCL established under § 25.149(f); and
(C) A speed that provides the maneuvering capability specified in § 25.143(h).
(ii) In icing conditions, V REF may not be less than:
(A) The speed determined in paragraph (b)(2)(i) of this section;
(B) 1.23 V SR0 with the most critical of the landing ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), if that speed exceeds V REF selected for non-icing conditions by more than 5 knots CAS; and
(C) A speed that provides the maneuvering capability specified in § 25.143(h) with the most critical of the landing ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g).
(3) Changes in configuration, power or thrust, and speed, must be made in accordance with the established procedures for service operation.
(4) The landing must be made without excessive vertical acceleration, tendency to bounce, nose over, ground loop, porpoise, or water loop.
(5) The landings may not require exceptional piloting skill or alertness.
(c) For landplanes and amphibians, the landing distance on land must be determined on a level, smooth, dry, hard-surfaced runway. In addition—
(1) The pressures on the wheel braking systems may not exceed those specified by the brake manufacturer;
(2) The brakes may not be used so as to cause excessive wear of brakes or tires; and
(3) Means other than wheel brakes may be used if that means—
(i) Is safe and reliable;
(ii) Is used so that consistent results can be expected in service; and
(iii) Is such that exceptional skill is not required to control the airplane.
(d) For seaplanes and amphibians, the landing distance on water must be determined on smooth water.
(e) For skiplanes, the landing distance on snow must be determined on smooth, dry, snow.
(f) The landing distance data must include correction factors for not more than 50 percent of the nominal wind components along the landing path opposite to the direction of landing, and not less than 150 percent of the nominal wind components along the landing path in the direction of landing.
(g) If any device is used that depends on the operation of any engine, and if the landing distance would be noticeably increased when a landing is made with that engine inoperative, the landing distance must be determined with that engine inoperative unless the use of compensating means will result in a landing distance not more than that with each engine operating.
§ 25.143   General.
(a) The airplane must be safely controllable and maneuverable during—
(1) Takeoff;
(2) Climb;
(3) Level flight;
(4) Descent; and
(5) Landing.
(b) It must be possible to make a smooth transition from one flight condition to any other flight condition without exceptional piloting skill, alertness, or strength, and without danger of exceeding the airplane limit-load factor under any probable operating conditions, including—
(1) The sudden failure of the critical engine;
(2) For airplanes with three or more engines, the sudden failure of the second critical engine when the airplane is in the en route, approach, or landing configuration and is trimmed with the critical engine inoperative; and
(3) Configuration changes, including deployment or retraction of deceleration devices.
(c) The airplane must be shown to be safely controllable and maneuverable with the most critical of the ice accretion(s) appropriate to the phase of flight as defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), and with the critical engine inoperative and its propeller (if applicable) in the minimum drag position:
(1) At the minimum V 2 for takeoff;
(2) During an approach and go-around; and
(3) During an approach and landing.
(d) The following table prescribes, for conventional wheel type controls, the maximum control forces permitted during the testing required by paragraph (a) through (c) of this section:
(e) Approved operating procedures or conventional operating practices must be followed when demonstrating compliance with the control force limitations for short term application that are prescribed in paragraph (d) of this section. The airplane must be in trim, or as near to being in trim as practical, in the preceding steady flight condition. For the takeoff condition, the airplane must be trimmed according to the approved operating procedures.
(f) When demonstrating compliance with the control force limitations for long term application that are prescribed in paragraph (d) of this section, the airplane must be in trim, or as near to being in trim as practical.
(g) When maneuvering at a constant airspeed or Mach number (up to V FC /M FC ), the stick forces and the gradient of the stick force versus maneuvering load factor must lie within satisfactory limits. The stick forces must not be so great as to make excessive demands on the pilot's strength when maneuvering the airplane, and must not be so low that the airplane can easily be overstressed inadvertently. Changes of gradient that occur with changes of load factor must not cause undue difficulty in maintaining control of the airplane, and local gradients must not be so low as to result in a danger of overcontrolling.
(h) The maneuvering capabilities in a constant speed coordinated turn at forward center of gravity, as specified in the following table, must be free of stall warning or other characteristics that might interfere with normal maneuvering:
1 A combination of weight, altitude, and temperature (WAT) such that the thrust or power setting produces the minimum climb gradient specified in § 25.121 for the flight condition.
2 Airspeed approved for all-engines-operating initial climb.
3 That thrust or power setting which, in the event of failure of the critical engine and without any crew action to adjust the thrust or power of the remaining engines, would result in the thrust or power specified for the takeoff condition at V 2 , or any lesser thrust or power setting that is used for all-engines-operating initial climb procedures.
(i) When demonstrating compliance with § 25.143 in icing conditions—
(1) Controllability must be demonstrated with the most critical of the ice accretion(s) for the particular flight phase as defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g);
(2) It must be shown that a push force is required throughout a pushover maneuver down to a zero g load factor, or the lowest load factor obtainable if limited by elevator power or other design characteristic of the flight control system. It must be possible to promptly recover from the maneuver without exceeding a pull control force of 50 pounds; and
(3) Any changes in force that the pilot must apply to the pitch control to maintain speed with increasing sideslip angle must be steadily increasing with no force reversals, unless the change in control force is gradual and easily controllable by the pilot without using exceptional piloting skill, alertness, or strength.
(j) For flight in icing conditions before the ice protection system has been activated and is performing its intended function, it must be demonstrated in flight with the most critical of the ice accretion(s) defined in Appendix C, part II, paragraph (e) of this part and Appendix O, part II, paragraph (d) of this part, as applicable, in accordance with § 25.21(g), that:
(1) The airplane is controllable in a pull-up maneuver up to 1.5 g load factor; and
(2) There is no pitch control force reversal during a pushover maneuver down to 0.5 g load factor.
§ 25.145   Longitudinal control.
(a) It must be possible, at any point between the trim speed prescribed in § 25.103(b)(6) and stall identification (as defined in § 25.201(d)), to pitch the nose downward so that the acceleration to this selected trim speed is prompt with
(1) The airplane trimmed at the trim speed prescribed in § 25.103(b)(6);
(2) The landing gear extended;
(3) The wing flaps (i) retracted and (ii) extended; and
(4) Power (i) off and (ii) at maximum continuous power on the engines.
(b) With the landing gear extended, no change in trim control, or exertion of more than 50 pounds control force (representative of the maximum short term force that can be applied readily by one hand) may be required for the following maneuvers:
(1) With power off, flaps retracted, and the airplane trimmed at 1.3 V SR1 , extend the flaps as rapidly as possible while maintaining the airspeed at approximately 30 percent above the reference stall speed existing at each instant throughout the maneuver.
(2) Repeat paragraph (b)(1) except initially extend the flaps and then retract them as rapidly as possible.
(3) Repeat paragraph (b)(2), except at the go-around power or thrust setting.
(4) With power off, flaps retracted, and the airplane trimmed at 1.3 V SR1 , rapidly set go-around power or thrust while maintaining the same airspeed.
(5) Repeat paragraph (b)(4) except with flaps extended.
(6) With power off, flaps extended, and the airplane trimmed at 1.3 V SR1 , obtain and maintain airspeeds between V SW and either 1.6 V SR1 or V FE , whichever is lower.
(c) It must be possible, without exceptional piloting skill, to prevent loss of altitude when complete retraction of the high lift devices from any position is begun during steady, straight, level flight at 1.08 V SR1 for propeller powered airplanes, or 1.13 V SR1 for turbojet powered airplanes, with—
(1) Simultaneous movement of the power or thrust controls to the go-around power or thrust setting;
(2) The landing gear extended; and
(3) The critical combinations of landing weights and altitudes.
(d) If gated high-lift device control positions are provided, paragraph (c) of this section applies to retractions of the high-lift devices from any position from the maximum landing position to the first gated position, between gated positions, and from the last gated position to the fully retracted position. The requirements of paragraph (c) of this section also apply to retractions from each approved landing position to the control position(s) associated with the high-lift device configuration(s) used to establish the go-around procedure(s) from that landing position. In addition, the first gated control position from the maximum landing position must correspond with a configuration of the high-lift devices used to establish a go-around procedure from a landing configuration. Each gated control position must require a separate and distinct motion of the control to pass through the gated position and must have features to prevent inadvertent movement of the control through the gated position. It must only be possible to make this separate and distinct motion once the control has reached the gated position.
§ 25.147   Directional and lateral control.
(a) Directional control; general. It must be possible, with the wings level, to yaw into the operative engine and to safely make a reasonably sudden change in heading of up to 15 degrees in the direction of the critical inoperative engine. This must be shown at 1.3 V S R1 for heading changes up to 15 degrees (except that the heading change at which the rudder pedal force is 150 pounds need not be exceeded), and with—
(1) The critical engine inoperative and its propeller in the minimum drag position;
(2) The power required for level flight at 1.3 V S R1, but not more than maximum continuous power;
(3) The most unfavorable center of gravity;
(4) Landing gear retracted;
(5) Flaps in the approach position; and
(6) Maximum landing weight.
(b) Directional control; airplanes with four or more engines. Airplanes with four or more engines must meet the requirements of paragraph (a) of this section except that—
(1) The two critical engines must be inoperative with their propellers (if applicable) in the minimum drag position;
(2) [Reserved]
(3) The flaps must be in the most favorable climb position.
(c) Lateral control; general. It must be possible to make 20° banked turns, with and against the inoperative engine, from steady flight at a speed equal to 1.3 V S R1, with—
(1) The critical engine inoperative and its propeller (if applicable) in the minimum drag position;
(2) The remaining engines at maximum continuous power;
(3) The most unfavorable center of gravity;
(4) Landing gear (i) retracted and (ii) extended;
(5) Flaps in the most favorable climb position; and
(6) Maximum takeoff weight.
(d) Lateral control; roll capability. With the critical engine inoperative, roll response must allow normal maneuvers. Lateral control must be sufficient, at the speeds likely to be used with one engine inoperative, to provide a roll rate necessary for safety without excessive control forces or travel.
(e) Lateral control; airplanes with four or more engines. Airplanes with four or more engines must be able to make 20° banked turns, with and against the inoperative engines, from steady flight at a speed equal to 1.3 V S R1, with maximum continuous power, and with the airplane in the configuration prescribed by paragraph (b) of this section.
(f) Lateral control; all engines operating. With the engines operating, roll response must allow normal maneuvers (such as recovery from upsets produced by gusts and the initiation of evasive maneuvers). There must be enough excess lateral control in sideslips (up to sideslip angles that might be required in normal operation), to allow a limited amount of maneuvering and to correct for gusts. Lateral control must be enough at any speed up to V FC / M FC to provide a peak roll rate necessary for safety, without excessive control forces or travel.
§ 25.149   Minimum control speed.
(a) In establishing the minimum control speeds required by this section, the method used to simulate critical engine failure must represent the most critical mode of powerplant failure with respect to controllability expected in service.
(b) V MC is the calibrated airspeed at which, when the critical engine is suddenly made inoperative, it is possible to maintain control of the airplane with that engine still inoperative and maintain straight flight with an angle of bank of not more than 5 degrees.
(c) V MC may not exceed 1.13 V SR with—
(1) Maximum available takeoff power or thrust on the engines;
(2) The most unfavorable center of gravity;
(3) The airplane trimmed for takeoff;
(4) The maximum sea level takeoff weight (or any lesser weight necessary to show V MC );
(5) The airplane in the most critical takeoff configuration existing along the flight path after the airplane becomes airborne, except with the landing gear retracted;
(6) The airplane airborne and the ground effect negligible; and
(7) If applicable, the propeller of the inoperative engine—
(i) Windmilling;
(ii) In the most probable position for the specific design of the propeller control; or
(iii) Feathered, if the airplane has an automatic feathering device acceptable for showing compliance with the climb requirements of § 25.121.
(d) The rudder forces required to maintain control at V MC may not exceed 150 pounds nor may it be necessary to reduce power or thrust of the operative engines. During recovery, the airplane may not assume any dangerous attitude or require exceptional piloting skill, alertness, or strength to prevent a heading change of more than 20 degrees.
(e) V MCG , the minimum control speed on the ground, is the calibrated airspeed during the takeoff run at which, when the critical engine is suddenly made inoperative, it is possible to maintain control of the airplane using the rudder control alone (without the use of nosewheel steering), as limited by 150 pounds of force, and the lateral control to the extent of keeping the wings level to enable the takeoff to be safely continued using normal piloting skill. In the determination of V MCG , assuming that the path of the airplane accelerating with all engines operating is along the centerline of the runway, its path from the point at which the critical engine is made inoperative to the point at which recovery to a direction parallel to the centerline is completed may not deviate more than 30 feet laterally from the centerline at any point. V MCG must be established with—
(1) The airplane in each takeoff configuration or, at the option of the applicant, in the most critical takeoff configuration;
(2) Maximum available takeoff power or thrust on the operating engines;
(3) The most unfavorable center of gravity;
(4) The airplane trimmed for takeoff; and
(5) The most unfavorable weight in the range of takeoff weights.
(f) V MCL , the minimum control speed during approach and landing with all engines operating, is the calibrated airspeed at which, when the critical engine is suddenly made inoperative, it is possible to maintain control of the airplane with that engine still inoperative, and maintain straight flight with an angle of bank of not more than 5 degrees. V MCL must be established with—
(1) The airplane in the most critical configuration (or, at the option of the applicant, each configuration) for approach and landing with all engines operating;
(2) The most unfavorable center of gravity;
(3) The airplane trimmed for approach with all engines operating;
(4) The most favorable weight, or, at the option of the applicant, as a function of weight;
(5) For propeller airplanes, the propeller of the inoperative engine in the position it achieves without pilot action, assuming the engine fails while at the power or thrust necessary to maintain a three degree approach path angle; and
(6) Go-around power or thrust setting on the operating engine(s).
(g) For airplanes with three or more engines, V MCL-2 , the minimum control speed during approach and landing with one critical engine inoperative, is the calibrated airspeed at which, when a second critical engine is suddenly made inoperative, it is possible to maintain control of the airplane with both engines still inoperative, and maintain straight flight with an angle of bank of not more than 5 degrees. V MCL-2 must be established with—
(1) The airplane in the most critical configuration (or, at the option of the applicant, each configuration) for approach and landing with one critical engine inoperative;
(2) The most unfavorable center of gravity;
(3) The airplane trimmed for approach with one critical engine inoperative;
(4) The most unfavorable weight, or, at the option of the applicant, as a function of weight;
(5) For propeller airplanes, the propeller of the more critical inoperative engine in the position it achieves without pilot action, assuming the engine fails while at the power or thrust necessary to maintain a three degree approach path angle, and the propeller of the other inoperative engine feathered;
(6) The power or thrust on the operating engine(s) necessary to maintain an approach path angle of three degrees when one critical engine is inoperative; and
(7) The power or thrust on the operating engine(s) rapidly changed, immediately after the second critical engine is made inoperative, from the power or thrust prescribed in paragraph (g)(6) of this section to—
(i) Minimum power or thrust; and
(ii) Go-around power or thrust setting.
(h) In demonstrations of V MCL and V MCL-2 —
(1) The rudder force may not exceed 150 pounds;
(2) The airplane may not exhibit hazardous flight characteristics or require exceptional piloting skill, alertness, or strength;
(3) Lateral control must be sufficient to roll the airplane, from an initial condition of steady flight, through an angle of 20 degrees in the direction necessary to initiate a turn away from the inoperative engine(s), in not more than 5 seconds; and
(4) For propeller airplanes, hazardous flight characteristics must not be exhibited due to any propeller position achieved when the engine fails or during any likely subsequent movements of the engine or propeller controls.
§ 25.161   Trim.
(a) General. Each airplane must meet the trim requirements of this section after being trimmed, and without further pressure upon, or movement of, either the primary controls or their corresponding trim controls by the pilot or the automatic pilot.
(b) Lateral and directional trim. The airplane must maintain lateral and directional trim with the most adverse lateral displacement of the center of gravity within the relevant operating limitations, during normally expected conditions of operation (including operation at any speed from 1.3 V SR 1 to V MO /M MO ).
(c) Longitudinal trim. The airplane must maintain longitudinal trim during—
(1) A climb with maximum continuous power at a speed not more than 1.3 V SR 1 , with the landing gear retracted, and the flaps (i) retracted and (ii) in the takeoff position;
(2) Either a glide with power off at a speed not more than 1.3 V SR1 , or an approach within the normal range of approach speeds appropriate to the weight and configuration with power settings corresponding to a 3 degree glidepath, whichever is the most severe, with the landing gear extended, the wing flaps (i) retracted and (ii) extended, and with the most unfavorable combination of center of gravity position and weight approved for landing; and
(3) Level flight at any speed from 1.3 V SR 1 , to V MO /M MO, with the landing gear and flaps retracted, and from 1.3 V SR 1 to V LE with the landing gear extended.
(d) Longitudinal, directional, and lateral trim. The airplane must maintain longitudinal, directional, and lateral trim (and for the lateral trim, the angle of bank may not exceed five degrees) at 1.3 V SR 1 during climbing flight with—
(1) The critical engine inoperative;
(2) The remaining engines at maximum continuous power; and
(3) The landing gear and flaps retracted.
(e) Airplanes with four or more engines. Each airplane with four or more engines must also maintain trim in rectilinear flight with the most unfavorable center of gravity and at the climb speed, configuration, and power required by § 25.123(a) for the purpose of establishing the en route flight paths with two engines inoperative.
§ 25.171   General.
The airplane must be longitudinally, directionally, and laterally stable in accordance with the provisions of §§ 25.173 through 25.177. In addition, suitable stability and control feel (static stability) is required in any condition normally encountered in service, if flight tests show it is necessary for safe operation.
§ 25.173   Static longitudinal stability.
Under the conditions specified in § 25.175, the characteristics of the elevator control forces (including friction) must be as follows:
(a) A pull must be required to obtain and maintain speeds below the specified trim speed, and a push must be required to obtain and maintain speeds above the specified trim speed. This must be shown at any speed that can be obtained except speeds higher than the landing gear or wing flap operating limit speeds or V FC /M FC, whichever is appropriate, or lower than the minimum speed for steady unstalled flight.
(b) The airspeed must return to within 10 percent of the original trim speed for the climb, approach, and landing conditions specified in § 25.175 (a), (c), and (d), and must return to within 7.5 percent of the original trim speed for the cruising condition specified in § 25.175(b), when the control force is slowly released from any speed within the range specified in paragraph (a) of this section.
(c) The average gradient of the stable slope of the stick force versus speed curve may not be less than 1 pound for each 6 knots.
(d) Within the free return speed range specified in paragraph (b) of this section, it is permissible for the airplane, without control forces, to stabilize on speeds above or below the desired trim speeds if exceptional attention on the part of the pilot is not required to return to and maintain the desired trim speed and altitude.
§ 25.175   Demonstration of static longitudinal stability.
Static longitudinal stability must be shown as follows:
(a) Climb. The stick force curve must have a stable slope at speeds between 85 and 115 percent of the speed at which the airplane—
(1) Is trimmed, with—
(i) Wing flaps retracted;
(ii) Landing gear retracted;
(iii) Maximum takeoff weight; and
(iv) 75 percent of maximum continuous power for reciprocating engines or the maximum power or thrust selected by the applicant as an operating limitation for use during climb for turbine engines; and
(2) Is trimmed at the speed for best rate-of-climb except that the speed need not be less than 1.3 V SR 1 .
(b) Cruise. Static longitudinal stability must be shown in the cruise condition as follows:
(1) With the landing gear retracted at high speed, the stick force curve must have a stable slope at all speeds within a range which is the greater of 15 percent of the trim speed plus the resulting free return speed range, or 50 knots plus the resulting free return speed range, above and below the trim speed (except that the speed range need not include speeds less than 1.3 V SR 1 , nor speeds greater than V FC /M FC, nor speeds that require a stick force of more than 50 pounds), with—
(i) The wing flaps retracted;
(ii) The center of gravity in the most adverse position (see § 25.27);
(iii) The most critical weight between the maximum takeoff and maximum landing weights;
(iv) 75 percent of maximum continuous power for reciprocating engines or for turbine engines, the maximum cruising power selected by the applicant as an operating limitation (see § 25.1521), except that the power need not exceed that required at V MO / M MO ; and
(v) The airplane trimmed for level flight with the power required in paragraph (b)(1)(iv) of this section.
(2) With the landing gear retracted at low speed, the stick force curve must have a stable slope at all speeds within a range which is the greater of 15 percent of the trim speed plus the resulting free return speed range, or 50 knots plus the resulting free return speed range, above and below the trim speed (except that the speed range need not include speeds less than 1.3 V SR 1 , nor speeds greater than the minimum speed of the applicable speed range prescribed in paragraph (b)(1), nor speeds that require a stick force of more than 50 pounds), with—
(i) Wing flaps, center of gravity position, and weight as specified in paragraph (b)(1) of this section;
(ii) Power required for level flight at a speed equal to ( V MO + 1.3 V SR 1 )/2; and
(iii) The airplane trimmed for level flight with the power required in paragraph (b)(2)(ii) of this section.
(3) With the landing gear extended, the stick force curve must have a stable slope at all speeds within a range which is the greater of 15 percent of the trim speed plus the resulting free return speed range, or 50 knots plus the resulting free return speed range, above and below the trim speed (except that the speed range need not include speeds less than 1.3 V SR 1 , nor speeds greater than V LE, nor speeds that require a stick force of more than 50 pounds), with—
(i) Wing flap, center of gravity position, and weight as specified in paragraph (b)(1) of this section;
(ii) 75 percent of maximum continuous power for reciprocating engines or, for turbine engines, the maximum cruising power selected by the applicant as an operating limitation, except that the power need not exceed that required for level flight at V LE ; and
(iii) The aircraft trimmed for level flight with the power required in paragraph (b)(3)(ii) of this section.
(c) Approach. The stick force curve must have a stable slope at speeds between V SW and 1.7 V SR 1 , with—
(1) Wing flaps in the approach position;
(2) Landing gear retracted;
(3) Maximum landing weight; and
(4) The airplane trimmed at 1.3 V SR 1 with enough power to maintain level flight at this speed.
(d) Landing. The stick force curve must have a stable slope, and the stick force may not exceed 80 pounds, at speeds between V SW and 1.7 V SR 0 with—
(1) Wing flaps in the landing position;
(2) Landing gear extended;
(3) Maximum landing weight;
(4) The airplane trimmed at 1.3 V SR0 with—
(i) Power or thrust off, and
(ii) Power or thrust for level flight.
(5) The airplane trimmed at 1.3 V SR 0 with power or thrust off.
§ 25.177   Static lateral-directional stability.
(a) The static directional stability (as shown by the tendency to recover from a skid with the rudder free) must be positive for any landing gear and flap position and symmetric power condition, at speeds from 1.13 V SR1 , up to V FE , V LE , or V FC /M FC (as appropriate for the airplane configuration).
(b) The static lateral stability (as shown by the tendency to raise the low wing in a sideslip with the aileron controls free) for any landing gear and flap position and symmetric power condition, may not be negative at any airspeed (except that speeds higher than V FE need not be considered for flaps extended configurations nor speeds higher than V LE for landing gear extended configurations) in the following airspeed ranges:
(1) From 1.13 V SR1 to V MO /M MO .
(2) From V MO /M MO to V FC /M FC , unless the divergence is—
(i) Gradual;
(ii) Easily recognizable by the pilot; and
(iii) Easily controllable by the pilot.
(c) The following requirement must be met for the configurations and speed specified in paragraph (a) of this section. In straight, steady sideslips over the range of sideslip angles appropriate to the operation of the airplane, the aileron and rudder control movements and forces must be substantially proportional to the angle of sideslip in a stable sense. This factor of proportionality must lie between limits found necessary for safe operation. The range of sideslip angles evaluated must include those sideslip angles resulting from the lesser of:
(1) One-half of the available rudder control input; and
(2) A rudder control force of 180 pounds.
(d) For sideslip angles greater than those prescribed by paragraph (c) of this section, up to the angle at which full rudder control is used or a rudder control force of 180 pounds is obtained, the rudder control forces may not reverse, and increased rudder deflection must be needed for increased angles of sideslip. Compliance with this requirement must be shown using straight, steady sideslips, unless full lateral control input is achieved before reaching either full rudder control input or a rudder control force of 180 pounds; a straight, steady sideslip need not be maintained after achieving full lateral control input. This requirement must be met at all approved landing gear and flap positions for the range of operating speeds and power conditions appropriate to each landing gear and flap position with all engines operating.
§ 25.181   Dynamic stability.
(a) Any short period oscillation, not including combined lateral-directional oscillations, occurring between 1.13 V SR and maximum allowable speed appropriate to the configuration of the airplane must be heavily damped with the primary controls—
(1) Free; and
(2) In a fixed position.
(b) Any combined lateral-directional oscillations (“Dutch roll”) occurring between 1.13 V SR and maximum allowable speed appropriate to the configuration of the airplane must be positively damped with controls free, and must be controllable with normal use of the primary controls without requiring exceptional pilot skill.
§ 25.201   Stall demonstration.
(a) Stalls must be shown in straight flight and in 30 degree banked turns with—
(1) Power off; and
(2) The power necessary to maintain level flight at 1.5 V SR1 (where V SR1 corresponds to the reference stall speed at maximum landing weight with flaps in the approach position and the landing gear retracted).
(b) In each condition required by paragraph (a) of this section, it must be possible to meet the applicable requirements of § 25.203 with—
(1) Flaps, landing gear, and deceleration devices in any likely combination of positions approved for operation;
(2) Representative weights within the range for which certification is requested;
(3) The most adverse center of gravity for recovery; and
(4) The airplane trimmed for straight flight at the speed prescribed in § 25.103(b)(6).
(c) The following procedures must be used to show compliance with § 25.203;
(1) Starting at a speed sufficiently above the stalling speed to ensure that a steady rate of speed reduction can be established, apply the longitudinal control so that the speed reduction does not exceed one knot per second until the airplane is stalled.
(2) In addition, for turning flight stalls, apply the longitudinal control to achieve airspeed deceleration rates up to 3 knots per second.
(3) As soon as the airplane is stalled, recover by normal recovery techniques.
(d) The airplane is considered stalled when the behavior of the airplane gives the pilot a clear and distinctive indication of an acceptable nature that the airplane is stalled. Acceptable indications of a stall, occurring either individually or in combination, are—
(1) A nose-down pitch that cannot be readily arrested;
(2) Buffeting, of a magnitude and severity that is a strong and effective deterrent to further speed reduction; or
(3) The pitch control reaches the aft stop and no further increase in pitch attitude occurs when the control is held full aft for a short time before recovery is initiated.
§ 25.203   Stall characteristics.
(a) It must be possible to produce and to correct roll and yaw by unreversed use of the aileron and rudder controls, up to the time the airplane is stalled. No abnormal nose-up pitching may occur. The longitudinal control force must be positive up to and throughout the stall. In addition, it must be possible to promptly prevent stalling and to recover from a stall by normal use of the controls.
(b) For level wing stalls, the roll occurring between the stall and the completion of the recovery may not exceed approximately 20 degrees.
(c) For turning flight stalls, the action of the airplane after the stall may not be so violent or extreme as to make it difficult, with normal piloting skill, to effect a prompt recovery and to regain control of the airplane. The maximum bank angle that occurs during the recovery may not exceed—
(1) Approximately 60 degrees in the original direction of the turn, or 30 degrees in the opposite direction, for deceleration rates up to 1 knot per second; and
(2) Approximately 90 degrees in the original direction of the turn, or 60 degrees in the opposite direction, for deceleration rates in excess of 1 knot per second.
§ 25.207   Stall warning.
(a) Stall warning with sufficient margin to prevent inadvertent stalling with the flaps and landing gear in any normal position must be clear and distinctive to the pilot in straight and turning flight.
(b) The warning must be furnished either through the inherent aerodynamic qualities of the airplane or by a device that will give clearly distinguishable indications under expected conditions of flight. However, a visual stall warning device that requires the attention of the crew within the cockpit is not acceptable by itself. If a warning device is used, it must provide a warning in each of the airplane configurations prescribed in paragraph (a) of this section at the speed prescribed in paragraphs (c) and (d) of this section. Except for the stall warning prescribed in paragraph (h)(3)(ii) of this section, the stall warning for flight in icing conditions must be provided by the same means as the stall warning for flight in non-icing conditions.
(c) When the speed is reduced at rates not exceeding one knot per second, stall warning must begin, in each normal configuration, at a speed, V SW , exceeding the speed at which the stall is identified in accordance with § 25.201(d) by not less than five knots or five percent CAS, whichever is greater. Once initiated, stall warning must continue until the angle of attack is reduced to approximately that at which stall warning began.
(d) In addition to the requirement of paragraph (c) of this section, when the speed is reduced at rates not exceeding one knot per second, in straight flight with engines idling and at the center-of-gravity position specified in § 25.103(b)(5), V SW , in each normal configuration, must exceed V SR by not less than three knots or three percent CAS, whichever is greater.
(e) In icing conditions, the stall warning margin in straight and turning flight must be sufficient to allow the pilot to prevent stalling (as defined in § 25.201(d)) when the pilot starts a recovery maneuver not less than three seconds after the onset of stall warning. When demonstrating compliance with this paragraph, the pilot must perform the recovery maneuver in the same way as for the airplane in non-icing conditions. Compliance with this requirement must be demonstrated in flight with the speed reduced at rates not exceeding one knot per second, with—
(1) The most critical of the takeoff ice and final takeoff ice accretions defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), for each configuration used in the takeoff phase of flight;
(2) The most critical of the en route ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), for the en route configuration;
(3) The most critical of the holding ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), for the holding configuration(s);
(4) The most critical of the approach ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), for the approach configuration(s); and
(5) The most critical of the landing ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), for the landing and go-around configuration(s).
(f) The stall warning margin must be sufficient in both non-icing and icing conditions to allow the pilot to prevent stalling when the pilot starts a recovery maneuver not less than one second after the onset of stall warning in slow-down turns with at least 1.5 g load factor normal to the flight path and airspeed deceleration rates of at least 2 knots per second. When demonstrating compliance with this paragraph for icing conditions, the pilot must perform the recovery maneuver in the same way as for the airplane in non-icing conditions. Compliance with this requirement must be demonstrated in flight with—
(1) The flaps and landing gear in any normal position;
(2) The airplane trimmed for straight flight at a speed of 1.3 V SR ; and
(3) The power or thrust necessary to maintain level flight at 1.3 V SR .
(g) Stall warning must also be provided in each abnormal configuration of the high lift devices that is likely to be used in flight following system failures (including all configurations covered by Airplane Flight Manual procedures).
(h) The following stall warning margin is required for flight in icing conditions before the ice protection system has been activated and is performing its intended function. Compliance must be shown using the most critical of the ice accretion(s) defined in Appendix C, part II, paragraph (e) of this part and Appendix O, part II, paragraph (d) of this part, as applicable, in accordance with § 25.21(g). The stall warning margin in straight and turning flight must be sufficient to allow the pilot to prevent stalling without encountering any adverse flight characteristics when:
(1) The speed is reduced at rates not exceeding one knot per second;
(2) The pilot performs the recovery maneuver in the same way as for flight in non-icing conditions; and
(3) The recovery maneuver is started no earlier than:
(i) One second after the onset of stall warning if stall warning is provided by the same means as for flight in non-icing conditions; or
(ii) Three seconds after the onset of stall warning if stall warning is provided by a different means than for flight in non-icing conditions.
(i) In showing compliance with paragraph (h) of this section, if stall warning is provided by a different means in icing conditions than for non-icing conditions, compliance with § 25.203 must be shown using the accretion defined in appendix C, part II(e) of this part. Compliance with this requirement must be shown using the demonstration prescribed by § 25.201, except that the deceleration rates of § 25.201(c)(2) need not be demonstrated.
§ 25.231   Longitudinal stability and control.
(a) Landplanes may have no uncontrollable tendency to nose over in any reasonably expected operating condition or when rebound occurs during landing or takeoff. In addition—
(1) Wheel brakes must operate smoothly and may not cause any undue tendency to nose over; and
(2) If a tail-wheel landing gear is used, it must be possible, during the takeoff ground run on concrete, to maintain any attitude up to thrust line level, at 75 percent of V SR 1 .
(b) For seaplanes and amphibians, the most adverse water conditions safe for takeoff, taxiing, and landing, must be established.
§ 25.233   Directional stability and control.
(a) There may be no uncontrollable ground-looping tendency in 90° cross winds, up to a wind velocity of 20 knots or 0.2 V SR 0 , whichever is greater, except that the wind velocity need not exceed 25 knots at any speed at which the airplane may be expected to be operated on the ground. This may be shown while establishing the 90° cross component of wind velocity required by § 25.237.
(b) Landplanes must be satisfactorily controllable, without exceptional piloting skill or alertness, in power-off landings at normal landing speed, without using brakes or engine power to maintain a straight path. This may be shown during power-off landings made in conjunction with other tests.
(c) The airplane must have adequate directional control during taxiing. This may be shown during taxiing prior to takeoffs made in conjunction with other tests.
§ 25.235   Taxiing condition.
The shock absorbing mechanism may not damage the structure of the airplane when the airplane is taxied on the roughest ground that may reasonably be expected in normal operation.
§ 25.237   Wind velocities.
(a) For land planes and amphibians, the following applies:
(1) A 90-degree cross component of wind velocity, demonstrated to be safe for takeoff and landing, must be established for dry runways and must be at least 20 knots or 0.2 V SR0 , whichever is greater, except that it need not exceed 25 knots.
(2) The crosswind component for takeoff established without ice accretions is valid in icing conditions.
(3) The landing crosswind component must be established for:
(i) Non-icing conditions, and
(ii) Icing conditions with the most critical of the landing ice accretion(s) defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g).
(b) For seaplanes and amphibians, the following applies:
(1) A 90-degree cross component of wind velocity, up to which takeoff and landing is safe under all water conditions that may reasonably be expected in normal operation, must be established and must be at least 20 knots or 0.2 V SR 0 , whichever is greater, except that it need not exceed 25 knots.
(2) A wind velocity, for which taxiing is safe in any direction under all water conditions that may reasonably be expected in normal operation, must be established and must be at least 20 knots or 0.2 V SR0 , whichever is greater, except that it need not exceed 25 knots.
§ 25.239   Spray characteristics, control, and stability on water.
(a) For seaplanes and amphibians, during takeoff, taxiing, and landing, and in the conditions set forth in paragraph (b) of this section, there may be no—
(1) Spray characteristics that would impair the pilot's view, cause damage, or result in the taking in of an undue quantity of water;
(2) Dangerously uncontrollable porpoising, bounding, or swinging tendency; or
(3) Immersion of auxiliary floats or sponsons, wing tips, propeller blades, or other parts not designed to withstand the resulting water loads.
(b) Compliance with the requirements of paragraph (a) of this section must be shown—
(1) In water conditions, from smooth to the most adverse condition established in accordance with § 25.231;
(2) In wind and cross-wind velocities, water currents, and associated waves and swells that may reasonably be expected in operation on water;
(3) At speeds that may reasonably be expected in operation on water;
(4) With sudden failure of the critical engine at any time while on water; and
(5) At each weight and center of gravity position, relevant to each operating condition, within the range of loading conditions for which certification is requested.
(c) In the water conditions of paragraph (b) of this section, and in the corresponding wind conditions, the seaplane or amphibian must be able to drift for five minutes with engines inoperative, aided, if necessary, by a sea anchor.
§ 25.251   Vibration and buffeting.
(a) The airplane must be demonstrated in flight to be free from any vibration and buffeting that would prevent continued safe flight in any likely operating condition.
(b) Each part of the airplane must be demonstrated in flight to be free from excessive vibration under any appropriate speed and power conditions up to V DF /M DF . The maximum speeds shown must be used in establishing the operating limitations of the airplane in accordance with § 25.1505.
(c) Except as provided in paragraph (d) of this section, there may be no buffeting condition, in normal flight, including configuration changes during cruise, severe enough to interfere with the control of the airplane, to cause excessive fatigue to the crew, or to cause structural damage. Stall warning buffeting within these limits is allowable.
(d) There may be no perceptible buffeting condition in the cruise configuration in straight flight at any speed up to V MO / M MO, except that stall warning buffeting is allowable.
(e) For an airplane with M D greater than .6 or with a maximum operating altitude greater than 25,000 feet, the positive maneuvering load factors at which the onset of perceptible buffeting occurs must be determined with the airplane in the cruise configuration for the ranges of airspeed or Mach number, weight, and altitude for which the airplane is to be certificated. The envelopes of load factor, speed, altitude, and weight must provide a sufficient range of speeds and load factors for normal operations. Probable inadvertent excursions beyond the boundaries of the buffet onset envelopes may not result in unsafe conditions.
§ 25.253   High-speed characteristics.
(a) Speed increase and recovery characteristics. The following speed increase and recovery characteristics must be met:
(1) Operating conditions and characteristics likely to cause inadvertent speed increases (including upsets in pitch and roll) must be simulated with the airplane trimmed at any likely cruise speed up to V MO / M MO . These conditions and characteristics include gust upsets, inadvertent control movements, low stick force gradient in relation to control friction, passenger movement, leveling off from climb, and descent from Mach to airspeed limit altitudes.
(2) Allowing for pilot reaction time after effective inherent or artificial speed warning occurs, it must be shown that the airplane can be recovered to a normal attitude and its speed reduced to V MO / M MO, without—
(i) Exceptional piloting strength or skill;
(ii) Exceeding V D / M D, V DF / M DF, or the structural limitations; and
(iii) Buffeting that would impair the pilot's ability to read the instruments or control the airplane for recovery.
(3) With the airplane trimmed at any speed up to V MO /M MO , there must be no reversal of the response to control input about any axis at any speed up to V DF /M DF . Any tendency to pitch, roll, or yaw must be mild and readily controllable, using normal piloting techniques. When the airplane is trimmed at V MO /M MO , the slope of the elevator control force versus speed curve need not be stable at speeds greater than V FC /M FC , but there must be a push force at all speeds up to V DF /M DF and there must be no sudden or excessive reduction of elevator control force as V DF /M DF is reached.
(4) Adequate roll capability to assure a prompt recovery from a lateral upset condition must be available at any speed up to V DF /M DF .
(5) With the airplane trimmed at V MO /M MO , extension of the speedbrakes over the available range of movements of the pilot's control, at all speeds above V MO /M MO , but not so high that V DF /M DF would be exceeded during the maneuver, must not result in:
(i) An excessive positive load factor when the pilot does not take action to counteract the effects of extension;
(ii) Buffeting that would impair the pilot's ability to read the instruments or control the airplane for recovery; or
(iii) A nose down pitching moment, unless it is small.
(b) Maximum speed for stability characteristics, V FC /M FC . V FC /M FC is the maximum speed at which the requirements of §§ 25.143(g), 25.147(f), 25.175(b)(1), 25.177(a) through (c), and 25.181 must be met with flaps and landing gear retracted. Except as noted in § 25.253(c), V FC /M FC may not be less than a speed midway between V MO /M MO and V DF /M DF , except that, for altitudes where Mach number is the limiting factor, M FC need not exceed the Mach number at which effective speed warning occurs.
(c) Maximum speed for stability characteristics in icing conditions. The maximum speed for stability characteristics with the most critical of the ice accretions defined in Appendices C and O of this part, as applicable, in accordance with § 25.21(g), at which the requirements of §§ 25.143(g), 25.147(f), 25.175(b)(1), 25.177(a) through (c), and 25.181 must be met, is the lower of:
(1) 300 knots CAS;
(2) V FC ; or
(3) A speed at which it is demonstrated that the airframe will be free of ice accretion due to the effects of increased dynamic pressure.
§ 25.255   Out-of-trim characteristics.
(a) From an initial condition with the airplane trimmed at cruise speeds up to V MO /M MO, the airplane must have satisfactory maneuvering stability and controllability with the degree of out-of-trim in both the airplane nose-up and nose-down directions, which results from the greater of—
(1) A three-second movement of the longitudinal trim system at its normal rate for the particular flight condition with no aerodynamic load (or an equivalent degree of trim for airplanes that do not have a power-operated trim system), except as limited by stops in the trim system, including those required by § 25.655(b) for adjustable stabilizers; or
(2) The maximum mistrim that can be sustained by the autopilot while maintaining level flight in the high speed cruising condition.
(b) In the out-of-trim condition specified in paragraph (a) of this section, when the normal acceleration is varied from + 1 g to the positive and negative values specified in paragraph (c) of this section—
(1) The stick force vs. g curve must have a positive slope at any speed up to and including V FC /M FC ; and
(2) At speeds between V FC /M FC and V DF /M DF the direction of the primary longitudinal control force may not reverse.
(c) Except as provided in paragraphs (d) and (e) of this section, compliance with the provisions of paragraph (a) of this section must be demonstrated in flight over the acceleration range—
(1) −1 g to + 2.5 g; or
(2) 0 g to 2.0 g, and extrapolating by an acceptable method to −1 g and + 2.5 g.
(d) If the procedure set forth in paragraph (c)(2) of this section is used to demonstrate compliance and marginal conditions exist during flight test with regard to reversal of primary longitudinal control force, flight tests must be accomplished from the normal acceleration at which a marginal condition is found to exist to the applicable limit specified in paragraph (b)(1) of this section.
(e) During flight tests required by paragraph (a) of this section, the limit maneuvering load factors prescribed in §§ 25.333(b) and 25.337, and the maneuvering load factors associated with probable inadvertent excursions beyond the boundaries of the buffet onset envelopes determined under § 25.251(e), need not be exceeded. In addition, the entry speeds for flight test demonstrations at normal acceleration values less than 1 g must be limited to the extent necessary to accomplish a recovery without exceeding V DF /M DF .
(f) In the out-of-trim condition specified in paragraph (a) of this section, it must be possible from an overspeed condition at V DF /M DF to produce at least 1.5 g for recovery by applying not more than 125 pounds of longitudinal control force using either the primary longitudinal control alone or the primary longitudinal control and the longitudinal trim system. If the longitudinal trim is used to assist in producing the required load factor, it must be shown at V DF /M DF that the longitudinal trim can be actuated in the airplane nose-up direction with the primary surface loaded to correspond to the least of the following airplane nose-up control forces:
(1) The maximum control forces expected in service as specified in §§ 25.301 and 25.397.
(2) The control force required to produce 1.5 g.
(3) The control force corresponding to buffeting or other phenomena of such intensity that it is a strong deterrent to further application of primary longitudinal control force.
§ 25.301   Loads.
(a) Strength requirements are specified in terms of limit loads (the maximum loads to be expected in service) and ultimate loads (limit loads multiplied by prescribed factors of safety). Unless otherwise provided, prescribed loads are limit loads.
(b) Unless otherwise provided, the specified air, ground, and water loads must be placed in equilibrium with inertia forces, considering each item of mass in the airplane. These loads must be distributed to conservatively approximate or closely represent actual conditions. Methods used to determine load intensities and distribution must be validated by flight load measurement unless the methods used for determining those loading conditions are shown to be reliable.
(c) If deflections under load would significantly change the distribution of external or internal loads, this redistribution must be taken into account.
§ 25.302   Interaction of systems and structures.
For airplanes equipped with systems that affect structural performance, either directly or as a result of a failure or malfunction, the influence of these systems and their failure conditions must be taken into account when showing compliance with the requirements of subparts C and D of this part. These criteria are only applicable to structure whose failure could prevent continued safe flight and landing.
(a) General. The applicant must use the following criteria in determining the influence of a system and its failure conditions on the airplane structure.
(b) System fully operative. With the system fully operative, the following criteria apply:
(1) The applicant must derive limit loads for the limit conditions specified in subpart C of this part, taking into account the behavior of the system up to the limit loads. System nonlinearities must be taken into account.
(2) The applicant must show that the airplane meets the strength requirements of subparts C and D of this part, using the appropriate factor of safety to derive ultimate loads from the limit loads defined in paragraph (b)(1) of this section. The effect of nonlinearities must be investigated sufficiently beyond limit conditions to ensure the behavior of the system presents no detrimental effects compared to the behavior below limit conditions. However, conditions beyond limit conditions need not be considered when it can be shown that the airplane has design features that will not allow it to exceed those limit conditions.
(3) [Reserved]
(c) System in the failure condition. For any system failure condition not shown to be extremely improbable or that results from a single failure, the following criteria apply:
(1) At the time of occurrence. The applicant must establish a realistic scenario, starting from 1g level flight conditions, and including pilot corrective actions, to determine the loads occurring at the time of failure and immediately after failure.
(i) For static strength substantiation, the airplane must be able to withstand the ultimate loads determined by multiplying the loads in paragraph (c)(1) of this section by a factor of safety that is related to the probability of occurrence of the failure. The factor of safety (F.S.) is defined in Figure 1.
(ii) For residual strength substantiation, the airplane must be able to withstand two thirds of the ultimate loads defined in paragraph (c)(1)(i) of this section. For pressurized cabins, these loads must be combined with the normal operating differential pressure.
(iii) [Reserved]
(iv) Failures of the system that result in forced structural vibrations (oscillatory failures) must not produce loads that could result in detrimental deformation of primary structure.
(2) For the continuation of the flight. For the airplane, in the system failed state and considering any appropriate reconfiguration and flight limitations, the following apply:
(i) The loads derived from the following conditions at speeds up to V C /M C , or the speed limitation prescribed for the remainder of the flight must be determined:
(A) the limit symmetrical maneuvering conditions specified in §§ 25.331 and 25.345,
(B) the limit gust and turbulence conditions specified in §§ 25.341 and 25.345,
(C) the limit rolling conditions specified in § 25.349 and the limit unsymmetrical conditions specified in §§ 25.367 and 25.427(b) and (c),
(D) the limit yaw maneuvering conditions specified in § 25.351,
(E) the limit ground loading conditions specified in §§ 25.473 and 25.491, and
(F) any other subpart C of this part load condition for which a system is specifically installed or tailored to reduce the loads of that condition.
(ii) For static strength substantiation, each part of the structure must be able to withstand the loads in paragraph (c)(2)(i) of this section multiplied by a factor of safety that depends on the probability of being in this failure condition. The factor of safety is defined in Figure 2.
(iii) For residual strength substantiation, the airplane must be able to withstand two thirds of the ultimate loads defined in paragraph (c)(2)(ii) of this section. For pressurized cabins, these loads must be combined with the normal operating differential pressure.
(iv) If the loads induced by the failure condition have a significant effect on fatigue or damage tolerance then their effects must be taken into account.
(v)-(vi) [Reserved]
(3) [Reserved]
(d) Failure indications. For system failure detection and indication, the following apply:
(1) The system must be checked for failure conditions evaluated under paragraph (c) of this section that degrade the structural capability below the level required by subparts C (excluding § 25.302) and D of this part or that reduce the reliability of the remaining system. As far as practicable, these failures must be indicated to the flightcrew before flight.
(2) The existence of any failure condition evaluated under paragraph (c) of this section that results in a factor of safety between the airplane strength and the loads of subpart C of this part below 1.25 must be indicated to the flightcrew.
(e) Dispatch with known failure conditions. If the airplane is to be dispatched in a known system failure condition that affects structural performance or affects the reliability of the remaining system to maintain structural performance, then the Master Minimum Equipment List must ensure the provisions of § 25.302 are met for the dispatched condition and for any subsequent failures. Flight limitations and operational limitations may be taken into account in establishing Qj as the combined probability of being in the dispatched failure condition and the subsequent failure condition for the safety margins in Figure 2. No reduction in these safety margins is allowed if the subsequent system failure rate is greater than 10 −3 per flight hour.
§ 25.303   Factor of safety.
Unless otherwise specified, a factor of safety of 1.5 must be applied to the prescribed limit load which are considered external loads on the structure. When a loading condition is prescribed in terms of ultimate loads, a factor of safety need not be applied unless otherwise specified.
§ 25.305   Strength and deformation.
(a) The structure must be able to support limit loads without detrimental permanent deformation. At any load up to limit loads, the deformation may not interfere with safe operation.
(b) The structure must be able to support ultimate loads without failure for at least 3 seconds. However, when proof of strength is shown by dynamic tests simulating actual load conditions, the 3-second limit does not apply. Static tests conducted to ultimate load must include the ultimate deflections and ultimate deformation induced by the loading. When analytical methods are used to show compliance with the ultimate load strength requirements, it must be shown that—
(1) The effects of deformation are not significant;
(2) The deformations involved are fully accounted for in the analysis; or
(3) The methods and assumptions used are sufficient to cover the effects of these deformations.
(c) Where structural flexibility is such that any rate of load application likely to occur in the operating conditions might produce transient stresses appreciably higher than those corresponding to static loads, the effects of this rate of application must be considered.
(d) [Reserved]
(e) The airplane must be designed to withstand any vibration and buffeting that might occur in any likely operating condition up to V D /M D , including stall and probable inadvertent excursions beyond the boundaries of the buffet onset envelope. This must be shown by analysis, flight tests, or other tests found necessary by the Administrator.
(f) Unless shown to be extremely improbable, the airplane must be designed to withstand any forced structural vibration resulting from any failure, malfunction or adverse condition in the flight control system. These must be considered limit loads and must be investigated at airspeeds up to V C /M C .
§ 25.307   Proof of structure.
(a) Compliance with the strength and deformation requirements of this subpart must be shown for each critical loading condition. Structural analysis may be used only if the structure conforms to that for which experience has shown this method to be reliable. In other cases, substantiating tests must be made to load levels that are sufficient to verify structural behavior up to loads specified in § 25.305.
(b)-(c) [Reserved]
(d) When static or dynamic tests are used to show compliance with the requirements of § 25.305(b) for flight structures, appropriate material correction factors must be applied to the test results, unless the structure, or part thereof, being tested has features such that a number of elements contribute to the total strength of the structure and the failure of one element results in the redistribution of the load through alternate load paths.
§ 25.321   General.
(a) Flight load factors represent the ratio of the aerodynamic force component (acting normal to the assumed longitudinal axis of the airplane) to the weight of the airplane. A positive load factor is one in which the aerodynamic force acts upward with respect to the airplane.
(b) Considering compressibility effects at each speed, compliance with the flight load requirements of this subpart must be shown—
(1) At each critical altitude within the range of altitudes selected by the applicant;
(2) At each weight from the design minimum weight to the design maximum weight appropriate to each particular flight load condition; and
(3) For each required altitude and weight, for any practicable distribution of disposable load within the operating limitations recorded in the Airplane Flight Manual.
(c) Enough points on and within the boundaries of the design envelope must be investigated to ensure that the maximum load for each part of the airplane structure is obtained.
(d) The significant forces acting on the airplane must be placed in equilibrium in a rational or conservative manner. The linear inertia forces must be considered in equilibrium with the thrust and all aerodynamic loads, while the angular (pitching) inertia forces must be considered in equilibrium with thrust and all aerodynamic moments, including moments due to loads on components such as tail surfaces and nacelles. Critical thrust values in the range from zero to maximum continuous thrust must be considered.
§ 25.331   Symmetric maneuvering conditions.
(a) Procedure. For the analysis of the maneuvering flight conditions specified in paragraphs (b) and (c) of this section, the following provisions apply:
(1) Where sudden displacement of a control is specified, the assumed rate of control surface displacement may not be less than the rate that could be applied by the pilot through the control system.
(2) In determining elevator angles and chordwise load distribution in the maneuvering conditions of paragraphs (b) and (c) of this section, the effect of corresponding pitching velocities must be taken into account. The in-trim and out-of-trim flight conditions specified in § 25.255 must be considered.
(b) Maneuvering balanced conditions. Assuming the airplane to be in equilibrium with zero pitching acceleration, the maneuvering conditions A through I on the maneuvering envelope in § 25.333(b) must be investigated.
(c) Maneuvering pitching conditions. The following conditions must be investigated:
(1) Maximum pitch control displacement at V A . The airplane is assumed to be flying in steady level flight (point A 1 , § 25.333(b)) and the cockpit pitch control is suddenly moved to obtain extreme nose up pitching acceleration. In defining the tail load, the response of the airplane must be taken into account. Airplane loads that occur subsequent to the time when normal acceleration at the c.g. exceeds the positive limit maneuvering load factor (at point A 2 in § 25.333(b)), or the resulting tailplane normal load reaches its maximum, whichever occurs first, need not be considered.
(2) Checked maneuver between V A and V D . Nose-up checked pitching maneuvers must be analyzed in which the positive limit load factor prescribed in § 25.337 is achieved. As a separate condition, nose-down checked pitching maneuvers must be analyzed in which a limit load factor of 0g is achieved. In defining the airplane loads, the flight deck pitch control motions described in paragraphs (c)(2)(i) through (iv) of this section must be used:
(i) The airplane is assumed to be flying in steady level flight at any speed between V A and V D and the flight deck pitch control is moved in accordance with the following formula:
(ii) For nose-up pitching maneuvers, the complete flight deck pitch control displacement history may be scaled down in amplitude to the extent necessary to ensure that the positive limit load factor prescribed in § 25.337 is not exceeded. For nose-down pitching maneuvers, the complete flight deck control displacement history may be scaled down in amplitude to the extent necessary to ensure that the normal acceleration at the center of gravity does not go below 0g.
(iii) In addition, for cases where the airplane response to the specified flight deck pitch control motion does not achieve the prescribed limit load factors, then the following flight deck pitch control motion must be used:
(iv) In cases where the flight deck pitch control motion may be affected by inputs from systems (for example, by a stick pusher that can operate at high load factor as well as at 1g), then the effects of those systems shall be taken into account.
(v) Airplane loads that occur beyond the following times need not be considered:
(A) For the nose-up pitching maneuver, the time at which the normal acceleration at the center of gravity goes below 0g;
(B) For the nose-down pitching maneuver, the time at which the normal acceleration at the center of gravity goes above the positive limit load factor prescribed in § 25.337;
(C) t max. .
§ 25.333   Flight maneuvering envelope.
(a) General. The strength requirements must be met at each combination of airspeed and load factor on and within the boundaries of the representative maneuvering envelope ( V-n diagram) of paragraph (b) of this section. This envelope must also be used in determining the airplane structural operating limitations as specified in § 25.1501.
(b) Maneuvering envelope.
§ 25.335   Design airspeeds.
The selected design airspeeds are equivalent airspeeds (EAS). Estimated values of V S 0 and V S 1 must be conservative.
(a) Design cruising speed, V C . For V C, the following apply:
(1) The minimum value of V C must be sufficiently greater than V B to provide for inadvertent speed increases likely to occur as a result of severe atmospheric turbulence.
(2) Except as provided in § 25.335(d)(2), V C may not be less than V B + 1.32 U REF (with U REF as specified in § 25.341(a)(5)(i)). However V C need not exceed the maximum speed in level flight at maximum continuous power for the corresponding altitude.
(3) At altitudes where V D is limited by Mach number, V C may be limited to a selected Mach number.
(b) Design dive speed, V D . V D must be selected so that V C / M C is not greater than 0.8 V D / M D, or so that the minimum speed margin between V C / M C and V D / M D is the greater of the following values:
(1) From an initial condition of stabilized flight at V C / M C, the airplane is upset, flown for 20 seconds along a flight path 7.5° below the initial path, and then pulled up at a load factor of 1.5 g (0.5 g acceleration increment). The speed increase occurring in this maneuver may be calculated if reliable or conservative aerodynamic data is used. Power as specified in § 25.175(b)(1)(iv) is assumed until the pullup is initiated, at which time power reduction and the use of pilot controlled drag devices may be assumed;
(2) The minimum speed margin must be enough to provide for atmospheric variations (such as horizontal gusts, and penetration of jet streams and cold fronts) and for instrument errors and airframe production variations. These factors may be considered on a probability basis. The margin at altitude where M C is limited by compressibility effects must not less than 0.07M unless a lower margin is determined using a rational analysis that includes the effects of any automatic systems. In any case, the margin may not be reduced to less than 0.05M.
(c) Design maneuvering speed V A . For V A , the following apply:
(1) V A may not be less than V S 1 √n where—
(i) n is the limit positive maneuvering load factor at V C ; and
(ii) V S 1 is the stalling speed with flaps retracted.
(2) V A and V S must be evaluated at the design weight and altitude under consideration.
(3) V A need not be more than V C or the speed at which the positive C N max curve intersects the positive maneuver load factor line, whichever is less.
(d) Design speed for maximum gust intensity, V B .
(1) V B may not be less than
(2) At altitudes where V C is limited by Mach number—
(i) V B may be chosen to provide an optimum margin between low and high speed buffet boundaries; and,
(ii) V B need not be greater than V C .
(e) Design flap speeds, V F . For V F , the following apply:
(1) The design flap speed for each flap position (established in accordance with § 25.697(a)) must be sufficiently greater than the operating speed recommended for the corresponding stage of flight (including balked landings) to allow for probable variations in control of airspeed and for transition from one flap position to another.
(2) If an automatic flap positioning or load limiting device is used, the speeds and corresponding flap positions programmed or allowed by the device may be used.
(3) V F may not be less than—
(i) 1.6 V S 1 with the flaps in takeoff position at maximum takeoff weight;
(ii) 1.8 V S 1 with the flaps in approach position at maximum landing weight, and
(iii) 1.8 V S 0 with the flaps in landing position at maximum landing weight.
(f) Design drag device speeds, V DD . The selected design speed for each drag device must be sufficiently greater than the speed recommended for the operation of the device to allow for probable variations in speed control. For drag devices intended for use in high speed descents, V DD may not be less than V D . When an automatic drag device positioning or load limiting means is used, the speeds and corresponding drag device positions programmed or allowed by the automatic means must be used for design.
§ 25.337   Limit maneuvering load factors.
(a) Except where limited by maximum (static) lift coefficients, the airplane is assumed to be subjected to symmetrical maneuvers resulting in the limit maneuvering load factors prescribed in this section. Pitching velocities appropriate to the corresponding pull-up and steady turn maneuvers must be taken into account.
(b) The positive limit maneuvering load factor n for any speed up to Vn may not be less than 2.1 + 24,000/ ( W + 10,000) except that n may not be less than 2.5 and need not be greater than 3.8—where W is the design maximum takeoff weight.
(c) The negative limit maneuvering load factor—
(1) May not be less than −1.0 at speeds up to V C ; and
(2) Must vary linearly with speed from the value at V C to zero at V D .
(d) Maneuvering load factors lower than those specified in this section may be used if the airplane has design features that make it impossible to exceed these values in flight.
§ 25.341   Gust and turbulence loads.
(a) Discrete Gust Design Criteria. The airplane is assumed to be subjected to symmetrical vertical and lateral gusts in level flight. Limit gust loads must be determined in accordance with the provisions:
(1) Loads on each part of the structure must be determined by dynamic analysis. The analysis must take into account unsteady aerodynamic characteristics and all significant structural degrees of freedom including rigid body motions.
(2) The shape of the gust must be:
(3) A sufficient number of gust gradient distances in the range 30 feet to 350 feet must be investigated to find the critical response for each load quantity.
(4) The design gust velocity must be:
(5) The following reference gust velocities apply:
(i) At airplane speeds between V B and V C : Positive and negative gusts with reference gust velocities of 56.0 ft/sec EAS must be considered at sea level. The reference gust velocity may be reduced linearly from 56.0 ft/sec EAS at sea level to 44.0 ft/sec EAS at 15,000 feet. The reference gust velocity may be further reduced linearly from 44.0 ft/sec EAS at 15,000 feet to 20.86 ft/sec EAS at 60,000 feet.
(ii) At the airplane design speed V D : The reference gust velocity must be 0.5 times the value obtained under § 25.341(a)(5)(i).
(6) The flight profile alleviation factor, F g , must be increased linearly from the sea level value to a value of 1.0 at the maximum operating altitude defined in § 25.1527. At sea level, the flight profile alleviation factor is determined by the following equation:
(7) When a stability augmentation system is included in the analysis, the effect of any significant system nonlinearities should be accounted for when deriving limit loads from limit gust conditions.
(b) Continuous turbulence design criteria. The dynamic response of the airplane to vertical and lateral continuous turbulence must be taken into account. The dynamic analysis must take into account unsteady aerodynamic characteristics and all significant structural degrees of freedom including rigid body motions. The limit loads must be determined for all critical altitudes, weights, and weight distributions as specified in § 25.321(b), and all critical speeds within the ranges indicated in § 25.341(b)(3).
(1) Except as provided in paragraphs (b)(4) and (5) of this section, the following equation must be used:
(2) Values of A must be determined according to the following formula:
(3) The limit turbulence intensities, U σ , in feet per second true airspeed required for compliance with this paragraph are—
(i) At airplane speeds between V B and V C :
(ii) At speed V D : U σ is equal to 1/2 the values obtained under paragraph (b)(3)(i) of this section.
(iii) At speeds between V C and V D : U σ is equal to a value obtained by linear interpolation.
(iv) At all speeds, both positive and negative incremental loads due to continuous turbulence must be considered.
(4) When an automatic system affecting the dynamic response of the airplane is included in the analysis, the effects of system non-linearities on loads at the limit load level must be taken into account in a realistic or conservative manner.
(5) If necessary for the assessment of loads on airplanes with significant non-linearities, it must be assumed that the turbulence field has a root-mean-square velocity equal to 40 percent of the U σ values specified in paragraph (b)(3) of this section. The value of limit load is that load with the same probability of exceedance in the turbulence field as A U σ of the same load quantity in a linear approximated model.
(c) Supplementary gust conditions for wing-mounted engines. For airplanes equipped with wing-mounted engines, the engine mounts, pylons, and wing supporting structure must be designed for the maximum response at the nacelle center of gravity derived from the following dynamic gust conditions applied to the airplane:
(1) A discrete gust determined in accordance with § 25.341(a) at each angle normal to the flight path, and separately,
(2) A pair of discrete gusts, one vertical and one lateral. The length of each of these gusts must be independently tuned to the maximum response in accordance with § 25.341(a). The penetration of the airplane in the combined gust field and the phasing of the vertical and lateral component gusts must be established to develop the maximum response to the gust pair. In the absence of a more rational analysis, the following formula must be used for each of the maximum engine loads in all six degrees of freedom:
§ 25.343   Design fuel and oil loads.
(a) The disposable load combinations must include each fuel and oil load in the range from zero fuel and oil to the selected maximum fuel and oil load. A structural reserve fuel condition, not exceeding 45 minutes of fuel under the operating conditions in § 25.1001(e) and (f), as applicable, may be selected.
(b) If a structural reserve fuel condition is selected, it must be used as the minimum fuel weight condition for showing compliance with the flight load requirements as prescribed in this subpart. In addition—
(1) The structure must be designed for a condition of zero fuel and oil in the wing at limit loads corresponding to—
(i) A maneuvering load factor of + 2.25; and
(ii) The gust and turbulence conditions of § 25.341(a) and (b), but assuming 85% of the gust velocities prescribed in § 25.341(a)(4) and 85% of the turbulence intensities prescribed in § 25.341(b)(3).
(2) Fatigue evaluation of the structure must account for any increase in operating stresses resulting from the design condition of paragraph (b)(1) of this section; and
(3) The flutter, deformation, and vibration requirements must also be met with zero fuel.
§ 25.345   High lift devices.
(a) If wing flaps are to be used during takeoff, approach, or landing, at the design flap speeds established for these stages of flight under § 25.335(e) and with the wing flaps in the corresponding positions, the airplane is assumed to be subjected to symmetrical maneuvers and gusts. The resulting limit loads must correspond to the conditions determined as follows:
(1) Maneuvering to a positive limit load factor of 2.0; and
(2) Positive and negative gusts of 25 ft/sec EAS acting normal to the flight path in level flight. Gust loads resulting on each part of the structure must be determined by rational analysis. The analysis must take into account the unsteady aerodynamic characteristics and rigid body motions of the aircraft. The shape of the gust must be as described in § 25.341(a)(2) except that—
(b) The airplane must be designed for the conditions prescribed in paragraph (a) of this section, except that the airplane load factor need not exceed 1.0, taking into account, as separate conditions, the effects of—
(1) Propeller slipstream corresponding to maximum continuous power at the design flap speeds V F, and with takeoff power at not less than 1.4 times the stalling speed for the particular flap position and associated maximum weight; and
(2) A head-on gust of 25 feet per second velocity (EAS).
(c) If flaps or other high lift devices are to be used in en route conditions, and with flaps in the appropriate position at speeds up to the flap design speed chosen for these conditions, the airplane is assumed to be subjected to symmetrical maneuvers and gusts within the range determined by—
(1) Maneuvering to a positive limit load factor as prescribed in § 25.337(b); and
(2) The vertical gust and turbulence conditions prescribed in § 25.341(a) and (b).
(d) The airplane must be designed for a maneuvering load factor of 1.5 g at the maximum take-off weight with the wing-flaps and similar high lift devices in the landing configurations.
§ 25.349   Rolling conditions.
The airplane must be designed for loads resulting from the rolling conditions specified in paragraphs (a) and (b) of this section. Unbalanced aerodynamic moments about the center of gravity must be reacted in a rational or conservative manner, considering the principal masses furnishing the reacting inertia forces.
(a) Maneuvering. The following conditions, speeds, and aileron deflections (except as the deflections may be limited by pilot effort) must be considered in combination with an airplane load factor of zero and of two-thirds of the positive maneuvering factor used in design. In determining the required aileron deflections, the torsional flexibility of the wing must be considered in accordance with § 25.301(b):
(1) Conditions corresponding to steady rolling velocities must be investigated. In addition, conditions corresponding to maximum angular acceleration must be investigated for airplanes with engines or other weight concentrations outboard of the fuselage. For the angular acceleration conditions, zero rolling velocity may be assumed in the absence of a rational time history investigation of the maneuver.
(2) At V A, a sudden deflection of the aileron to the stop is assumed.
(3) At V C, the aileron deflection must be that required to produce a rate of roll not less than that obtained in paragraph (a)(2) of this section.
(4) At V D, the aileron deflection must be that required to produce a rate of roll not less than one-third of that in paragraph (a)(2) of this section.
(b) Unsymmetrical gusts. The airplane is assumed to be subjected to unsymmetrical vertical gusts in level flight. The resulting limit loads must be determined from either the wing maximum airload derived directly from § 25.341(a), or the wing maximum airload derived indirectly from the vertical load factor calculated from § 25.341(a). It must be assumed that 100 percent of the wing air load acts on one side of the airplane and 80 percent of the wing air load acts on the other side.
§ 25.351   Yaw maneuver conditions.
The airplane must be designed for loads resulting from the yaw maneuver conditions specified in paragraphs (a) through (d) of this section at speeds from V MC to V D . Unbalanced aerodynamic moments about the center of gravity must be reacted in a rational or conservative manner considering the airplane inertia forces. In computing the tail loads the yawing velocity may be assumed to be zero.
(a) With the airplane in unaccelerated flight at zero yaw, it is assumed that the cockpit rudder control is suddenly displaced to achieve the resulting rudder deflection, as limited by:
(1) The control system on control surface stops; or
(2) A limit pilot force of 300 pounds from V MC to V A and 200 pounds from V C /M C to V D /M D , with a linear variation between V A and V C /M C .
(b) With the cockpit rudder control deflected so as always to maintain the maximum rudder deflection available within the limitations specified in paragraph (a) of this section, it is assumed that the airplane yaws to the overswing sideslip angle.
(c) With the airplane yawed to the static equilibrium sideslip angle, it is assumed that the cockpit rudder control is held so as to achieve the maximum rudder deflection available within the limitations specified in paragraph (a) of this section.
(d) With the airplane yawed to the static equilibrium sideslip angle of paragraph (c) of this section, it is assumed that the cockpit rudder control is suddenly returned to neutral.
§ 25.353   Rudder control reversal conditions.
Airplanes with a powered rudder control surface or surfaces must be designed for loads, considered to be ultimate, resulting from the yaw maneuver conditions specified in paragraphs (a) through (e) of this section at speeds from V MC to V C /M C . Any permanent deformation resulting from these ultimate load conditions must not prevent continued safe flight and landing. The applicant must evaluate these conditions with the landing gear retracted and speed brakes (and spoilers when used as speed brakes) retracted. The applicant must evaluate the effects of flaps, flaperons, or any other aerodynamic devices when used as flaps, and slats-extended configurations, if they are used in en route conditions. Unbalanced aerodynamic moments about the center of gravity must be reacted in a rational or conservative manner considering the airplane inertia forces. In computing the loads on the airplane, the yawing velocity may be assumed to be zero. The applicant must assume a pilot force of 200 pounds when evaluating each of the following conditions:
(a) With the airplane in unaccelerated flight at zero yaw, the flightdeck rudder control is suddenly and fully displaced to achieve the resulting rudder deflection, as limited by the control system or the control surface stops.
(b) With the airplane yawed to the overswing sideslip angle, the flightdeck rudder control is suddenly and fully displaced in the opposite direction, as limited by the control system or control surface stops.
(c) With the airplane yawed to the opposite overswing sideslip angle, the flightdeck rudder control is suddenly and fully displaced in the opposite direction, as limited by the control system or control surface stops.
(d) With the airplane yawed to the subsequent overswing sideslip angle, the flightdeck rudder control is suddenly and fully displaced in the opposite direction, as limited by the control system or control surface stops.
(e) With the airplane yawed to the opposite overswing sideslip angle, the flightdeck rudder control is suddenly returned to neutral.
§ 25.361   Engine and auxiliary power unit torque.
(a) For engine installations—
(1) Each engine mount, pylon, and adjacent supporting airframe structures must be designed for the effects of—
(i) A limit engine torque corresponding to takeoff power/thrust and, if applicable, corresponding propeller speed, acting simultaneously with 75% of the limit loads from flight condition A of § 25.333(b);
(ii) A limit engine torque corresponding to the maximum continuous power/thrust and, if applicable, corresponding propeller speed, acting simultaneously with the limit loads from flight condition A of § 25.333(b); and
(iii) For turbopropeller installations only, in addition to the conditions specified in paragraphs (a)(1)(i) and (ii) of this section, a limit engine torque corresponding to takeoff power and propeller speed, multiplied by a factor accounting for propeller control system malfunction, including quick feathering, acting simultaneously with 1g level flight loads. In the absence of a rational analysis, a factor of 1.6 must be used.
(2) The limit engine torque to be considered under paragraph (a)(1) of this section must be obtained by—
(i) For turbopropeller installations, multiplying mean engine torque for the specified power/thrust and speed by a factor of 1.25;
(ii) For other turbine engines, the limit engine torque must be equal to the maximum accelerating torque for the case considered.
(3) The engine mounts, pylons, and adjacent supporting airframe structure must be designed to withstand 1g level flight loads acting simultaneously with the limit engine torque loads imposed by each of the following conditions to be considered separately:
(i) Sudden maximum engine deceleration due to malfunction or abnormal condition; and
(ii) The maximum acceleration of engine.
(b) For auxiliary power unit installations, the power unit mounts and adjacent supporting airframe structure must be designed to withstand 1g level flight loads acting simultaneously with the limit torque loads imposed by each of the following conditions to be considered separately:
(1) Sudden maximum auxiliary power unit deceleration due to malfunction, abnormal condition, or structural failure; and
(2) The maximum acceleration of the auxiliary power unit.
§ 25.362   Engine failure loads.
(a) For engine mounts, pylons, and adjacent supporting airframe structure, an ultimate loading condition must be considered that combines 1g flight loads with the most critical transient dynamic loads and vibrations, as determined by dynamic analysis, resulting from failure of a blade, shaft, bearing or bearing support, or bird strike event. Any permanent deformation from these ultimate load conditions must not prevent continued safe flight and landing.
(b) The ultimate loads developed from the conditions specified in paragraph (a) of this section are to be—
(1) Multiplied by a factor of 1.0 when applied to engine mounts and pylons; and
(2) Multiplied by a factor of 1.25 when applied to adjacent supporting airframe structure.
§ 25.363   Side load on engine and auxiliary power unit mounts.
(a) Each engine and auxiliary power unit mount and its supporting structure must be designed for a limit load factor in lateral direction, for the side load on the engine and auxiliary power unit mount, at least equal to the maximum load factor obtained in the yawing conditions but not less than—
(1) 1.33; or
(2) One-third of the limit load factor for flight condition A as prescribed in § 25.333(b).
(b) The side load prescribed in paragraph (a) of this section may be assumed to be independent of other flight conditions.
§ 25.365   Pressurized compartment loads.
For airplanes with one or more pressurized compartments the following apply:
(a) The airplane structure must be strong enough to withstand the flight loads combined with pressure differential loads from zero up to the maximum relief valve setting.
(b) The external pressure distribution in flight, and stress concentrations and fatigue effects must be accounted for.
(c) If landings may be made with the compartment pressurized, landing loads must be combined with pressure differential loads from zero up to the maximum allowed during landing.
(d) The airplane structure must be designed to be able to withstand the pressure differential loads corresponding to the maximum relief valve setting multiplied by a factor of 1.33 for airplanes to be approved for operation to 45,000 feet or by a factor of 1.67 for airplanes to be approved for operation above 45,000 feet, omitting other loads.
(e) Any structure, component or part, inside or outside a pressurized compartment, the failure of which could interfere with continued safe flight and landing, must be designed to withstand the effects of a sudden release of pressure through an opening in any compartment at any operating altitude resulting from each of the following conditions:
(1) The penetration of the compartment by a portion of an engine following an engine disintegration;
(2) Any opening in any pressurized compartment up to the size H o in square feet; however, small compartments may be combined with an adjacent pressurized compartment and both considered as a single compartment for openings that cannot reasonably be expected to be confined to the small compartment. The size H o must be computed by the following formula:
(3) The maximum opening caused by airplane or equipment failures not shown to be extremely improbable.
(f) In complying with paragraph (e) of this section, the fail-safe features of the design may be considered in determining the probability of failure or penetration and probable size of openings, provided that possible improper operation of closure devices and inadvertent door openings are also considered. Furthermore, the resulting differential pressure loads must be combined in a rational and conservative manner with 1-g level flight loads and any loads arising from emergency depressurization conditions. These loads may be considered as ultimate conditions; however, any deformations associated with these conditions must not interfere with continued safe flight and landing. The pressure relief provided by intercompartment venting may also be considered.
(g)(1) Except as provided in paragraph (g)(2) of this section, bulkheads, floors, and partitions in pressurized compartments for occupants must be designed to withstand the conditions specified in paragraph (e) of this section. In addition, reasonable design precautions must be taken to minimize the probability of parts becoming detached and injuring occupants while in their seats.
(2) Partitions adjacent to the opening specified in paragraph (e)(2) of this section need not be designed to withstand that condition provided—
(i) Failure of the partition would not interfere with continued safe flight and landing; and
(ii) Designing the partition to withstand the condition specified in paragraph (e)(2) of this section would be impractical.
§ 25.367   Unsymmetrical loads due to engine failure.
(a) The airplane must be designed for the unsymmetrical loads resulting from the failure of the critical engine. Turbopropeller airplanes must be designed for the following conditions in combination with a single malfunction of the propeller drag limiting system, considering the probable pilot corrective action on the flight controls:
(1) At speeds between V MC and V D, the loads resulting from power failure because of fuel flow interruption are considered to be limit loads.
(2) At speeds between V MC and V C, the loads resulting from the disconnection of the engine compressor from the turbine or from loss of the turbine blades are considered to be ultimate loads.
(3) The time history of the thrust decay and drag build-up occurring as a result of the prescribed engine failures must be substantiated by test or other data applicable to the particular engine-propeller combination.
(4) The timing and magnitude of the probable pilot corrective action must be conservatively estimated, considering the characteristics of the particular engine-propeller-airplane combination.
(b) Pilot corrective action may be assumed to be initiated at the time maximum yawing velocity is reached, but not earlier than two seconds after the engine failure. The magnitude of the corrective action may be based on the control forces specified in § 25.397(b) except that lower forces may be assumed where it is shown by anaylsis or test that these forces can control the yaw and roll resulting from the prescribed engine failure conditions.
§ 25.371   Gyroscopic loads.
The structure supporting any engine or auxiliary power unit must be designed for the loads, including gyroscopic loads, arising from the conditions specified in §§ 25.331, 25.341, 25.349, 25.351, 25.473, 25.479, and 25.481, with the engine or auxiliary power unit at the maximum rotating speed appropriate to the condition. For the purposes of compliance with this paragraph, the pitch maneuver in § 25.331(c)(1) must be carried out until the positive limit maneuvering load factor (point A 2 in § 25.333(b)) is reached.
§ 25.373   Speed control devices.
If speed control devices (such as spoilers and drag flaps) are installed for use in en route conditions—
(a) The airplane must be designed for the symmetrical maneuvers prescribed in §§ 25.333 and 25.337, the yawing maneuvers in § 25.351, and the vertical and lateral gust and turbulence conditions prescribed in § 25.341(a) and (b) at each setting and the maximum speed associated with that setting; and
(b) If the device has automatic operating or load limiting features, the airplane must be designed for the maneuver and gust conditions prescribed in paragraph (a) of this section, at the speeds and corresponding device positions that the mechanism allows.
§ 25.391   Control surface loads: General.
The control surfaces must be designed for the limit loads resulting from the flight conditions in §§ 25.331, 25.341(a) and (b), 25.349, and 25.351, considering the requirements for—
(a) Loads parallel to hinge line, in § 25.393;
(b) Pilot effort effects, in § 25.397;
(c) Trim tab effects, in § 25.407;
(d) Unsymmetrical loads, in § 25.427; and
(e) Auxiliary aerodynamic surfaces, in § 25.445.
§ 25.393   Loads parallel to hinge line.
(a) Control surfaces and supporting hinge brackets must be designed for inertia loads acting parallel to the hinge line.
(b) In the absence of more rational data, the inertia loads may be assumed to be equal to KW, where—
(1) K = 24 for vertical surfaces;
(2) K = 12 for horizontal surfaces; and
(3) W = weight of the movable surfaces.
§ 25.395   Control system.
(a) Longitudinal, lateral, directional, and drag control system and their supporting structures must be designed for loads corresponding to 125 percent of the computed hinge moments of the movable control surface in the conditions prescribed in § 25.391.
(b) The system limit loads of paragraph (a) of this section need not exceed the loads that can be produced by the pilot (or pilots) and by automatic or power devices operating the controls.
(c) The loads must not be less than those resulting from application of the minimum forces prescribed in § 25.397(c).
§ 25.397   Control system loads.
(a) General. The maximum and minimum pilot forces, specified in paragraph (c) of this section, are assumed to act at the appropriate control grips or pads (in a manner simulating flight conditions) and to be reacted at the attachment of the control system to the control surface horn.
(b) Pilot effort effects. In the control surface flight loading condition, the air loads on movable surfaces and the corresponding deflections need not exceed those that would result in flight from the application of any pilot force within the ranges specified in paragraph (c) of this section. Two-thirds of the maximum values specified for the aileron and elevator may be used if control surface hinge moments are based on reliable data. In applying this criterion, the effects of servo mechanisms, tabs, and automatic pilot systems, must be considered.
(c) Limit pilot forces and torques. The limit pilot forces and torques are as follows:
1 The critical parts of the aileron control system must be designed for a single tangential force with a limit value equal to 1.25 times the couple force determined from these criteria.
2 D = wheel diameter (inches).
3 The unsymmetrical forces must be applied at one of the normal handgrip points on the periphery of the control wheel.
§ 25.399   Dual control system.
(a) Each dual control system must be designed for the pilots operating in opposition, using individual pilot forces not less than—
(1) 0.75 times those obtained under § 25.395; or
(2) The minimum forces specified in § 25.397(c).
(b) The control system must be designed for pilot forces applied in the same direction, using individual pilot forces not less than 0.75 times those obtained under § 25.395.
§ 25.405   Secondary control system.
Secondary controls, such as wheel brake, spoiler, and tab controls, must be designed for the maximum forces that a pilot is likely to apply to those controls. The following values may be used:
Pilot Control Force Limits (Secondary Controls)
*Limited to flap, tab, stabilizer, spoiler, and landing gear operation controls.
§ 25.407   Trim tab effects.
The effects of trim tabs on the control surface design conditions must be accounted for only where the surface loads are limited by maximum pilot effort. In these cases, the tabs are considered to be deflected in the direction that would assist the pilot, and the deflections are—
(a) For elevator trim tabs, those required to trim the airplane at any point within the positive portion of the pertinent flight envelope in § 25.333(b), except as limited by the stops; and
(b) For aileron and rudder trim tabs, those required to trim the airplane in the critical unsymmetrical power and loading conditions, with appropriate allowance for rigging tolerances.
§ 25.409   Tabs.
(a) Trim tabs. Trim tabs must be designed to withstand loads arising from all likely combinations of tab setting, primary control position, and airplane speed (obtainable without exceeding the flight load conditions prescribed for the airplane as a whole), when the effect of the tab is opposed by pilot effort forces up to those specified in § 25.397(b).
(b) Balancing tabs. Balancing tabs must be designed for deflections consistent with the primary control surface loading conditions.
(c) Servo tabs. Servo tabs must be designed for deflections consistent with the primary control surface loading conditions obtainable within the pilot maneuvering effort, considering possible opposition from the trim tabs.
§ 25.415   Ground gust conditions.
(a) The flight control systems and surfaces must be designed for the limit loads generated when the airplane is subjected to a horizontal 65-knot ground gust from any direction while taxiing and while parked. For airplanes equipped with control system gust locks, the taxiing condition must be evaluated with the controls locked and unlocked, and the parked condition must be evaluated with the controls locked.
(b) The control system and surface loads due to ground gust may be assumed to be static loads, and the hinge moments H must be computed from the formula:
(c) The hinge moment factor K for ground gusts must be taken from the following table:
* A positive value of K indicates a moment tending to depress the surface, while a negative value of K indicates a moment tending to raise the surface.
(d) The computed hinge moment of paragraph (b) of this section must be used to determine the limit loads due to ground gust conditions for the control surface. A 1.25 factor on the computed hinge moments must be used in calculating limit control system loads.
(e) Where control system flexibility is such that the rate of load application in the ground gust conditions might produce transient stresses appreciably higher than those corresponding to static loads, in the absence of a rational analysis substantiating a different dynamic factor, an additional factor of 1.6 must be applied to the control system loads of paragraph (d) of this section to obtain limit loads. If a rational analysis is used, the additional factor must not be less than 1.2.
(f) For the condition of the control locks engaged, the control surfaces, the control system locks, and the parts of any control systems between the surfaces and the locks must be designed to the resultant limit loads. Where control locks are not provided, then the control surfaces, the control system stops nearest the surfaces, and the parts of any control systems between the surfaces and the stops must be designed to the resultant limit loads. If the control system design is such as to allow any part of the control system to impact with the stops due to flexibility, then the resultant impact loads must be taken into account in deriving the limit loads due to ground gust.
(g) For the condition of taxiing with the control locks disengaged, or where control locks are not provided, the following apply:
(1) The control surfaces, the control system stops nearest the surfaces, and the parts of any control systems between the surfaces and the stops must be designed to the resultant limit loads.
(2) The parts of the control systems between the stops nearest the surfaces and the flight deck controls must be designed to the resultant limit loads, except that the parts of the control system where loads are eventually reacted by the pilot need not exceed:
(i) The loads corresponding to the maximum pilot loads in § 25.397(c) for each pilot alone; or
(ii) 0.75 times these maximum loads for each pilot when the pilot forces are applied in the same direction.
§ 25.427   Unsymmetrical loads.
(a) In designing the airplane for lateral gust, yaw maneuver and roll maneuver conditions, account must be taken of unsymmetrical loads on the empennage arising from effects such as slipstream and aerodynamic interference with the wing, vertical fin and other aerodynamic surfaces.
(b) The horizontal tail must be assumed to be subjected to unsymmetrical loading conditions determined as follows:
(1) 100 percent of the maximum loading from the symmetrical maneuver conditions of § 25.331 and the vertical gust conditions of § 25.341(a) acting separately on the surface on one side of the plane of symmetry; and
(2) 80 percent of these loadings acting on the other side.
(c) For empennage arrangements where the horizontal tail surfaces have dihedral angles greater than plus or minus 10 degrees, or are supported by the vertical tail surfaces, the surfaces and the supporting structure must be designed for gust velocities specified in § 25.341(a) acting in any orientation at right angles to the flight path.
(d) Unsymmetrical loading on the empennage arising from buffet conditions of § 25.305(e) must be taken into account.
§ 25.445   Auxiliary aerodynamic surfaces.
(a) When significant, the aerodynamic influence between auxiliary aerodynamic surfaces, such as outboard fins and winglets, and their supporting aerodynamic surfaces, must be taken into account for all loading conditions including pitch, roll, and yaw maneuvers, and gusts as specified in § 25.341(a) acting at any orientation at right angles to the flight path.
(b) To provide for unsymmetrical loading when outboard fins extend above and below the horizontal surface, the critical vertical surface loading (load per unit area) determined under § 25.391 must also be applied as follows:
(1) 100 percent to the area of the vertical surfaces above (or below) the horizontal surface.
(2) 80 percent to the area below (or above) the horizontal surface.
§ 25.457   Wing flaps.
Wing flaps, their operating mechanisms, and their supporting structures must be designed for critical loads occurring in the conditions prescribed in § 25.345, accounting for the loads occurring during transition from one flap position and airspeed to another.
§ 25.459   Special devices.
The loading for special devices using aerodynamic surfaces (such as slots, slats and spoilers) must be determined from test data.
§ 25.471   General.
(a) Loads and equilibrium. For limit ground loads—
(1) Limit ground loads obtained under this subpart are considered to be external forces applied to the airplane structure; and
(2) In each specified ground load condition, the external loads must be placed in equilibrium with the linear and angular inertia loads in a rational or conservative manner.
(b) Critical centers of gravity. The critical centers of gravity within the range for which certification is requested must be selected so that the maximum design loads are obtained in each landing gear element. Fore and aft, vertical, and lateral airplane centers of gravity must be considered. Lateral displacements of the c.g. from the airplane centerline which would result in main gear loads not greater than 103 percent of the critical design load for symmetrical loading conditions may be selected without considering the effects of these lateral c.g. displacements on the loading of the main gear elements, or on the airplane structure provided—
(1) The lateral displacement of the c.g. results from random passenger or cargo disposition within the fuselage or from random unsymmetrical fuel loading or fuel usage; and
(2) Appropriate loading instructions for random disposable loads are included under the provisions of § 25.1583(c)(2) to ensure that the lateral displacement of the center of gravity is maintained within these limits.
(c) Landing gear dimension data. Figure 1 of appendix A contains the basic landing gear dimension data.
§ 25.473   Landing load conditions and assumptions.
(a) For the landing conditions specified in § 25.479 to § 25.485 the airplane is assumed to contact the ground—
(1) In the attitudes defined in § 25.479 and § 25.481;
(2) With a limit descent velocity of 10 fps at the design landing weight (the maximum weight for landing conditions at maximum descent velocity); and
(3) With a limit descent velocity of 6 fps at the design take-off weight (the maximum weight for landing conditions at a reduced descent velocity).
(4) The prescribed descent velocities may be modified if it is shown that the airplane has design features that make it impossible to develop these velocities.
(b) Airplane lift, not exceeding airplane weight, may be assumed unless the presence of systems or procedures significantly affects the lift.
(c) The method of analysis of airplane and landing gear loads must take into account at least the following elements:
(1) Landing gear dynamic characteristics.
(2) Spin-up and springback.
(3) Rigid body response.
(4) Structural dynamic response of the airframe, if significant.
(d) The landing gear dynamic characteristics must be validated by tests as defined in § 25.723(a).
(e) The coefficient of friction between the tires and the ground may be established by considering the effects of skidding velocity and tire pressure. However, this coefficient of friction need not be more than 0.8.
§ 25.477   Landing gear arrangement.
Sections 25.479 through 25.485 apply to airplanes with conventional arrangements of main and nose gears, or main and tail gears, when normal operating techniques are used.
§ 25.479   Level landing conditions.
(a) In the level attitude, the airplane is assumed to contact the ground at forward velocity components, ranging from V L1 to 1.25 V L2 parallel to the ground under the conditions prescribed in § 25.473 with—
(1) V L1 equal to V S0 (TAS) at the appropriate landing weight and in standard sea level conditions; and
(2) V L2 equal to V S0 (TAS) at the appropriate landing weight and altitudes in a hot day temperature of 41 degrees F. above standard.
(3) The effects of increased contact speed must be investigated if approval of downwind landings exceeding 10 knots is requested.
(b) For the level landing attitude for airplanes with tail wheels, the conditions specified in this section must be investigated with the airplane horizontal reference line horizontal in accordance with Figure 2 of Appendix A of this part.
(c) For the level landing attitude for airplanes with nose wheels, shown in Figure 2 of Appendix A of this part, the conditions specified in this section must be investigated assuming the following attitudes:
(1) An attitude in which the main wheels are assumed to contact the ground with the nose wheel just clear of the ground; and
(2) If reasonably attainable at the specified descent and forward velocities, an attitude in which the nose and main wheels are assumed to contact the ground simultaneously.
(d) In addition to the loading conditions prescribed in paragraph (a) of this section, but with maximum vertical ground reactions calculated from paragraph (a), the following apply:
(1) The landing gear and directly affected attaching structure must be designed for the maximum vertical ground reaction combined with an aft acting drag component of not less than 25% of this maximum vertical ground reaction.
(2) The most severe combination of loads that are likely to arise during a lateral drift landing must be taken into account. In absence of a more rational analysis of this condition, the following must be investigated:
(i) A vertical load equal to 75% of the maximum ground reaction of § 25.473 must be considered in combination with a drag and side load of 40% and 25% respectively of that vertical load.
(ii) The shock absorber and tire deflections must be assumed to be 75% of the deflection corresponding to the maximum ground reaction of § 25.473(a)(2). This load case need not be considered in combination with flat tires.
(3) The combination of vertical and drag components is considered to be acting at the wheel axle centerline.
§ 25.481   Tail-down landing conditions.
(a) In the tail-down attitude, the airplane is assumed to contact the ground at forward velocity components, ranging from V L1 to V L2 parallel to the ground under the conditions prescribed in § 25.473 with—
(1) V L 1 equal to V S 0 (TAS) at the appropriate landing weight and in standard sea level conditions; and
(2) V L 2 equal to V S 0 (TAS) at the appropriate landing weight and altitudes in a hot day temperature of 41 degrees F. above standard.
(3) The combination of vertical and drag components considered to be acting at the main wheel axle centerline.
(b) For the tail-down landing condition for airplanes with tail wheels, the main and tail wheels are assumed to contact the ground simultaneously, in accordance with figure 3 of appendix A. Ground reaction conditions on the tail wheel are assumed to act—
(1) Vertically; and
(2) Up and aft through the axle at 45 degrees to the ground line.
(c) For the tail-down landing condition for airplanes with nose wheels, the airplane is assumed to be at an attitude corresponding to either the stalling angle or the maximum angle allowing clearance with the ground by each part of the airplane other than the main wheels, in accordance with figure 3 of appendix A, whichever is less.
§ 25.483   One-gear landing conditions.
For the one-gear landing conditions, the airplane is assumed to be in the level attitude and to contact the ground on one main landing gear, in accordance with Figure 4 of Appendix A of this part. In this attitude—
(a) The ground reactions must be the same as those obtained on that side under § 25.479(d)(1), and
(b) Each unbalanced external load must be reacted by airplane inertia in a rational or conservative manner.
§ 25.485   Side load conditions.
In addition to § 25.479(d)(2) the following conditions must be considered:
(a) For the side load condition, the airplane is assumed to be in the level attitude with only the main wheels contacting the ground, in accordance with figure 5 of appendix A.
(b) Side loads of 0.8 of the vertical reaction (on one side) acting inward and 0.6 of the vertical reaction (on the other side) acting outward must be combined with one-half of the maximum vertical ground reactions obtained in the level landing conditions. These loads are assumed to be applied at the ground contact point and to be resisted by the inertia of the airplane. The drag loads may be assumed to be zero.
§ 25.487   Rebound landing condition.
(a) The landing gear and its supporting structure must be investigated for the loads occurring during rebound of the airplane from the landing surface.
(b) With the landing gear fully extended and not in contact with the ground, a load factor of 20.0 must act on the unsprung weights of the landing gear. This load factor must act in the direction of motion of the unsprung weights as they reach their limiting positions in extending with relation to the sprung parts of the landing gear.
§ 25.489   Ground handling conditions.
Unless otherwise prescribed, the landing gear and airplane structure must be investigated for the conditions in §§ 25.491 through 25.509 with the airplane at the design ramp weight (the maximum weight for ground handling conditions). No wing lift may be considered. The shock absorbers and tires may be assumed to be in their static position.
§ 25.491   Taxi, takeoff and landing roll.
Within the range of appropriate ground speeds and approved weights, the airplane structure and landing gear are assumed to be subjected to loads not less than those obtained when the aircraft is operating over the roughest ground that may reasonably be expected in normal operation.
§ 25.493   Braked roll conditions.
(a) An airplane with a tail wheel is assumed to be in the level attitude with the load on the main wheels, in accordance with figure 6 of appendix A. The limit vertical load factor is 1.2 at the design landing weight and 1.0 at the design ramp weight. A drag reaction equal to the vertical reaction multiplied by a coefficient of friction of 0.8, must be combined with the vertical ground reaction and applied at the ground contact point.
(b) For an airplane with a nose wheel the limit vertical load factor is 1.2 at the design landing weight, and 1.0 at the design ramp weight. A drag reaction equal to the vertical reaction, multiplied by a coefficient of friction of 0.8, must be combined with the vertical reaction and applied at the ground contact point of each wheel with brakes. The following two attitudes, in accordance with figure 6 of appendix A, must be considered:
(1) The level attitude with the wheels contacting the ground and the loads distributed between the main and nose gear. Zero pitching acceleration is assumed.
(2) The level attitude with only the main gear contacting the ground and with the pitching moment resisted by angular acceleration.
(c) A drag reaction lower than that prescribed in this section may be used if it is substantiated that an effective drag force of 0.8 times the vertical reaction cannot be attained under any likely loading condition.
(d) An airplane equipped with a nose gear must be designed to withstand the loads arising from the dynamic pitching motion of the airplane due to sudden application of maximum braking force. The airplane is considered to be at design takeoff weight with the nose and main gears in contact with the ground, and with a steady-state vertical load factor of 1.0. The steady-state nose gear reaction must be combined with the maximum incremental nose gear vertical reaction caused by the sudden application of maximum braking force as described in paragraphs (b) and (c) of this section.
(e) In the absence of a more rational analysis, the nose gear vertical reaction prescribed in paragraph (d) of this section must be calculated according to the following formula:
§ 25.495   Turning.
In the static position, in accordance with figure 7 of appendix A, the airplane is assumed to execute a steady turn by nose gear steering, or by application of sufficient differential power, so that the limit load factors applied at the center of gravity are 1.0 vertically and 0.5 laterally. The side ground reaction of each wheel must be 0.5 of the vertical reaction.
§ 25.497   Tail-wheel yawing.
(a) A vertical ground reaction equal to the static load on the tail wheel, in combination with a side component of equal magnitude, is assumed.
(b) If there is a swivel, the tail wheel is assumed to be swiveled 90° to the airplane longitudinal axis with the resultant load passing through the axle.
(c) If there is a lock, steering device, or shimmy damper the tail wheel is also assumed to be in the trailing position with the side load acting at the ground contact point.
§ 25.499   Nose-wheel yaw and steering.
(a) A vertical load factor of 1.0 at the airplane center of gravity, and a side component at the nose wheel ground contact equal to 0.8 of the vertical ground reaction at that point are assumed.
(b) With the airplane assumed to be in static equilibrium with the loads resulting from the use of brakes on one side of the main landing gear, the nose gear, its attaching structure, and the fuselage structure forward of the center of gravity must be designed for the following loads:
(1) A vertical load factor at the center of gravity of 1.0.
(2) A forward acting load at the airplane center of gravity of 0.8 times the vertical load on one main gear.
(3) Side and vertical loads at the ground contact point on the nose gear that are required for static equilibrium.
(4) A side load factor at the airplane center of gravity of zero.
(c) If the loads prescribed in paragraph (b) of this section result in a nose gear side load higher than 0.8 times the vertical nose gear load, the design nose gear side load may be limited to 0.8 times the vertical load, with unbalanced yawing moments assumed to be resisted by airplane inertia forces.
(d) For other than the nose gear, its attaching structure, and the forward fuselage structure, the loading conditions are those prescribed in paragraph (b) of this section, except that—
(1) A lower drag reaction may be used if an effective drag force of 0.8 times the vertical reaction cannot be reached under any likely loading condition; and
(2) The forward acting load at the center of gravity need not exceed the maximum drag reaction on one main gear, determined in accordance with § 25.493(b).
(e) With the airplane at design ramp weight, and the nose gear in any steerable position, the combined application of full normal steering torque and vertical force equal to 1.33 times the maximum static reaction on the nose gear must be considered in designing the nose gear, its attaching structure, and the forward fuselage structure.
§ 25.503   Pivoting.
(a) The airplane is assumed to pivot about one side of the main gear with the brakes on that side locked. The limit vertical load factor must be 1.0 and the coefficient of friction 0.8.
(b) The airplane is assumed to be in static equilibrium, with the loads being applied at the ground contact points, in accordance with figure 8 of appendix A.
§ 25.507   Reversed braking.
(a) The airplane must be in a three point static ground attitude. Horizontal reactions parallel to the ground and directed forward must be applied at the ground contact point of each wheel with brakes. The limit loads must be equal to 0.55 times the vertical load at each wheel or to the load developed by 1.2 times the nominal maximum static brake torque, whichever is less.
(b) For airplanes with nose wheels, the pitching moment must be balanced by rotational inertia.
(c) For airplanes with tail wheels, the resultant of the ground reactions must pass through the center of gravity of the airplane.
§ 25.509   Towing loads.
(a) The towing loads specified in paragraph (d) of this section must be considered separately. These loads must be applied at the towing fittings and must act parallel to the ground. In addition—
(1) A vertical load factor equal to 1.0 must be considered acting at the center of gravity;
(2) The shock struts and tires must be in their static positions; and
(3) With W T as the design ramp weight, the towing load, F TOW, is—
(i) 0.3 W T for W T less than 30,000 pounds;
(ii) ( 6W T + 450,000)/70 for W T between 30,000 and 100,000 pounds; and
(iii) 0.15 W T for W T over 100,000 pounds.
(b) For towing points not on the landing gear but near the plane of symmetry of the airplane, the drag and side tow load components specified for the auxiliary gear apply. For towing points located outboard of the main gear, the drag and side tow load components specified for the main gear apply. Where the specified angle of swivel cannot be reached, the maximum obtainable angle must be used.
(c) The towing loads specified in paragraph (d) of this section must be reacted as follows:
(1) The side component of the towing load at the main gear must be reacted by a side force at the static ground line of the wheel to which the load is applied.
(2) The towing loads at the auxiliary gear and the drag components of the towing loads at the main gear must be reacted as follows:
(i) A reaction with a maximum value equal to the vertical reaction must be applied at the axle of the wheel to which the load is applied. Enough airplane inertia to achieve equilibrium must be applied.
(ii) The loads must be reacted by airplane inertia.
(d) The prescribed towing loads are as follows:
§ 25.511   Ground load: unsymmetrical loads on multiple-wheel units.
(a) General. Multiple-wheel landing gear units are assumed to be subjected to the limit ground loads prescribed in this subpart under paragraphs (b) through (f) of this section. In addition—
(1) A tandem strut gear arrangement is a multiple-wheel unit; and
(2) In determining the total load on a gear unit with respect to the provisions of paragraphs (b) through (f) of this section, the transverse shift in the load centroid, due to unsymmetrical load distribution on the wheels, may be neglected.
(b) Distribution of limit loads to wheels; tires inflated. The distribution of the limit loads among the wheels of the landing gear must be established for each landing, taxiing, and ground handling condition, taking into account the effects of the following factors:
(1) The number of wheels and their physical arrangements. For truck type landing gear units, the effects of any seesaw motion of the truck during the landing impact must be considered in determining the maximum design loads for the fore and aft wheel pairs.
(2) Any differentials in tire diameters resulting from a combination of manufacturing tolerances, tire growth, and tire wear. A maximum tire-diameter differential equal to 2/3 of the most unfavorable combination of diameter variations that is obtained when taking into account manufacturing tolerances, tire growth, and tire wear, may be assumed.
(3) Any unequal tire inflation pressure, assuming the maximum variation to be ±5 percent of the nominal tire inflation pressure.
(4) A runway crown of zero and a runway crown having a convex upward shape that may be approximated by a slope of 1 1/2 percent with the horizontal. Runway crown effects must be considered with the nose gear unit on either slope of the crown.
(5) The airplane attitude.
(6) Any structural deflections.
(c) Deflated tires. The effect of deflated tires on the structure must be considered with respect to the loading conditions specified in paragraphs (d) through (f) of this section, taking into account the physical arrangement of the gear components. In addition—
(1) The deflation of any one tire for each multiple wheel landing gear unit, and the deflation of any two critical tires for each landing gear unit using four or more wheels per unit, must be considered; and
(2) The ground reactions must be applied to the wheels with inflated tires except that, for multiple-wheel gear units with more than one shock strut, a rational distribution of the ground reactions between the deflated and inflated tires, accounting for the differences in shock strut extensions resulting from a deflated tire, may be used.
(d) Landing conditions. For one and for two deflated tires, the applied load to each gear unit is assumed to be 60 percent and 50 percent, respectively, of the limit load applied to each gear for each of the prescribed landing conditions. However, for the drift landing condition of § 25.485, 100 percent of the vertical load must be applied.
(e) Taxiing and ground handling conditions. For one and for two deflated tires—
(1) The applied side or drag load factor, or both factors, at the center of gravity must be the most critical value up to 50 percent and 40 percent, respectively, of the limit side or drag load factors, or both factors, corresponding to the most severe condition resulting from consideration of the prescribed taxiing and ground handling conditions;
(2) For the braked roll conditions of § 25.493 (a) and (b)(2), the drag loads on each inflated tire may not be less than those at each tire for the symmetrical load distribution with no deflated tires;
(3) The vertical load factor at the center of gravity must be 60 percent and 50 percent, respectively, of the factor with no deflated tires, except that it may not be less than 1g; and
(4) Pivoting need not be considered.
(f) Towing conditions. For one and for two deflated tires, the towing load, F TOW, must be 60 percent and 50 percent, respectively, of the load prescribed.
§ 25.519   Jacking and tie-down provisions.
(a) General. The airplane must be designed to withstand the limit load conditions resulting from the static ground load conditions of paragraph (b) of this section and, if applicable, paragraph (c) of this section at the most critical combinations of airplane weight and center of gravity. The maximum allowable load at each jack pad must be specified.
(b) Jacking. The airplane must have provisions for jacking and must withstand the following limit loads when the airplane is supported on jacks—
(1) For jacking by the landing gear at the maximum ramp weight of the airplane, the airplane structure must be designed for a vertical load of 1.33 times the vertical static reaction at each jacking point acting singly and in combination with a horizontal load of 0.33 times the vertical static reaction applied in any direction.
(2) For jacking by other airplane structure at maximum approved jacking weight:
(i) The airplane structure must be designed for a vertical load of 1.33 times the vertical reaction at each jacking point acting singly and in combination with a horizontal load of 0.33 times the vertical static reaction applied in any direction.
(ii) The jacking pads and local structure must be designed for a vertical load of 2.0 times the vertical static reaction at each jacking point, acting singly and in combination with a horizontal load of 0.33 times the vertical static reaction applied in any direction.
(c) Tie-down. If tie-down points are provided, the main tie-down points and local structure must withstand the limit loads resulting from a 65-knot horizontal wind from any direction.
§ 25.521   General.
(a) Seaplanes must be designed for the water loads developed during takeoff and landing, with the seaplane in any attitude likely to occur in normal operation, and at the appropriate forward and sinking velocities under the most severe sea conditions likely to be encountered.
(b) Unless a more rational analysis of the water loads is made, or the standards in ANC-3 are used, §§ 25.523 through 25.537 apply.
(c) The requirements of this section and §§ 25.523 through 25.537 apply also to amphibians.
§ 25.523   Design weights and center of gravity positions.
(a) Design weights. The water load requirements must be met at each operating weight up to the design landing weight except that, for the takeoff condition prescribed in § 25.531, the design water takeoff weight (the maximum weight for water taxi and takeoff run) must be used.
(b) Center of gravity positions. The critical centers of gravity within the limits for which certification is requested must be considered to reach maximum design loads for each part of the seaplane structure.
§ 25.525   Application of loads.
(a) Unless otherwise prescribed, the seaplane as a whole is assumed to be subjected to the loads corresponding to the load factors specified in § 25.527.
(b) In applying the loads resulting from the load factors prescribed in § 25.527, the loads may be distributed over the hull or main float bottom (in order to avoid excessive local shear loads and bending moments at the location of water load application) using pressures not less than those prescribed in § 25.533(c).
(c) For twin float seaplanes, each float must be treated as an equivalent hull on a fictitious seaplane with a weight equal to one-half the weight of the twin float seaplane.
(d) Except in the takeoff condition of § 25.531, the aerodynamic lift on the seaplane during the impact is assumed to be 2/3 of the weight of the seaplane.
§ 25.527   Hull and main float load factors.
(a) Water reaction load factors n W must be computed in the following manner:
(1) For the step landing case
(2) For the bow and stern landing cases
(b) The following values are used:
(1) n W = water reaction load factor (that is, the water reaction divided by seaplane weight).
(2) C 1 = empirical seaplane operations factor equal to 0.012 (except that this factor may not be less than that necessary to obtain the minimum value of step load factor of 2.33).
(3) V S 0 = seaplane stalling speed in knots with flaps extended in the appropriate landing position and with no slipstream effect.
(4) β = angle of dead rise at the longitudinal station at which the load factor is being determined in accordance with figure 1 of appendix B.
(5) W= seaplane design landing weight in pounds.
(6) K 1 = empirical hull station weighing factor, in accordance with figure 2 of appendix B.
(7) r x = ratio of distance, measured parallel to hull reference axis, from the center of gravity of the seaplane to the hull longitudinal station at which the load factor is being computed to the radius of gyration in pitch of the seaplane, the hull reference axis being a straight line, in the plane of symmetry, tangential to the keel at the main step.
(c) For a twin float seaplane, because of the effect of flexibility of the attachment of the floats to the seaplane, the factor K 1 may be reduced at the bow and stern to 0.8 of the value shown in figure 2 of appendix B. This reduction applies only to the design of the carrythrough and seaplane structure.
§ 25.529   Hull and main float landing conditions.
(a) Symmetrical step, bow, and stern landing. For symmetrical step, bow, and stern landings, the limit water reaction load factors are those computed under § 25.527. In addition—
(1) For symmetrical step landings, the resultant water load must be applied at the keel, through the center of gravity, and must be directed perpendicularly to the keel line;
(2) For symmetrical bow landings, the resultant water load must be applied at the keel, one-fifth of the longitudinal distance from the bow to the step, and must be directed perpendicularly to the keel line; and
(3) For symmetrical stern landings, the resultant water load must be applied at the keel, at a point 85 percent of the longitudinal distance from the step to the stern post, and must be directed perpendicularly to the keel line.
(b) Unsymmetrical landing for hull and single float seaplanes. Unsymmetrical step, bow, and stern landing conditions must be investigated. In addition—
(1) The loading for each condition consists of an upward component and a side component equal, respectively, to 0.75 and 0.25 tan β times the resultant load in the corresponding symmetrical landing condition; and
(2) The point of application and direction of the upward component of the load is the same as that in the symmetrical condition, and the point of application of the side component is at the same longitudinal station as the upward component but is directed inward perpendicularly to the plane of symmetry at a point midway between the keel and chine lines.
(c) Unsymmetrical landing; twin float seaplanes. The unsymmetrical loading consists of an upward load at the step of each float of 0.75 and a side load of 0.25 tan β at one float times the step landing load reached under § 25.527. The side load is directed inboard, perpendicularly to the plane of symmetry midway between the keel and chine lines of the float, at the same longitudinal station as the upward load.
§ 25.531   Hull and main float takeoff condition.
For the wing and its attachment to the hull or main float—
(a) The aerodynamic wing lift is assumed to be zero; and
(b) A downward inertia load, corresponding to a load factor computed from the following formula, must be applied:
§ 25.533   Hull and main float bottom pressures.
(a) General. The hull and main float structure, including frames and bulkheads, stringers, and bottom plating, must be designed under this section.
(b) Local pressures. For the design of the bottom plating and stringers and their attachments to the supporting structure, the following pressure distributions must be applied:
(1) For an unflared bottom, the pressure at the chine is 0.75 times the pressure at the keel, and the pressures between the keel and chine vary linearly, in accordance with figure 3 of appendix B. The pressure at the keel (psi) is computed as follows:
(2) For a flared bottom, the pressure at the beginning of the flare is the same as that for an unflared bottom, and the pressure between the chine and the beginning of the flare varies linearly, in accordance with figure 3 of appendix B. The pressure distribution is the same as that prescribed in paragraph (b)(1) of this section for an unflared bottom except that the pressure at the chine is computed as follows:
(c) Distributed pressures. For the design of the frames, keel, and chine structure, the following pressure distributions apply:
(1) Symmetrical pressures are computed as follows:
(2) The unsymmetrical pressure distribution consists of the pressures prescribed in paragraph (c)(1) of this section on one side of the hull or main float centerline and one-half of that pressure on the other side of the hull or main float centerline, in accordance with figure 3 of appendix B.
§ 25.535   Auxiliary float loads.
(a) General. Auxiliary floats and their attachments and supporting structures must be designed for the conditions prescribed in this section. In the cases specified in paragraphs (b) through (e) of this section, the prescribed water loads may be distributed over the float bottom to avoid excessive local loads, using bottom pressures not less than those prescribed in paragraph (g) of this section.
(b) Step loading. The resultant water load must be applied in the plane of symmetry of the float at a point three-fourths of the distance from the bow to the step and must be perpendicular to the keel. The resultant limit load is computed as follows, except that the value of L need not exceed three times the weight of the displaced water when the float is completely submerged:
(c) Bow loading. The resultant limit load must be applied in the plane of symmetry of the float at a point one-fourth of the distance from the bow to the step and must be perpendicular to the tangent to the keel line at that point. The magnitude of the resultant load is that specified in paragraph (b) of this section.
(d) Unsymmetrical step loading. The resultant water load consists of a component equal to 0.75 times the load specified in paragraph (a) of this section and a side component equal to 0.25 tan β times the load specified in paragraph (b) of this section. The side load must be applied perpendicularly to the plane of symmetry of the float at a point midway between the keel and the chine.
(e) Unsymmetrical bow loading. The resultant water load consists of a component equal to 0.75 times the load specified in paragraph (b) of this section and a side component equal to 0.25 tan β times the load specified in paragraph (c) of this section. The side load must be applied perpendicularly to the plane of symmetry at a point midway between the keel and the chine.
(f) Immersed float condition. The resultant load must be applied at the centroid of the cross section of the float at a point one-third of the distance from the bow to the step. The limit load components are as follows:
(g) Float bottom pressures. The float bottom pressures must be established under § 25.533, except that the value of K 2 in the formulae may be taken as 1.0. The angle of dead rise to be used in determining the float bottom pressures is set forth in paragraph (b) of this section.
§ 25.537   Seawing loads.
Seawing design loads must be based on applicable test data.
§ 25.561   General.
(a) The airplane, although it may be damaged in emergency landing conditions on land or water, must be designed as prescribed in this section to protect each occupant under those conditions.
(b) The structure must be designed to give each occupant every reasonable chance of escaping serious injury in a minor crash landing when—
(1) Proper use is made of seats, belts, and all other safety design provisions;
(2) The wheels are retracted (where applicable); and
(3) The occupant experiences the following ultimate inertia forces acting separately relative to the surrounding structure:
(i) Upward, 3.0g
(ii) Forward, 9.0g
(iii) Sideward, 3.0g on the airframe; and 4.0g on the seats and their attachments.
(iv) Downward, 6.0g
(v) Rearward, 1.5g
(c) For equipment, cargo in the passenger compartments and any other large masses, the following apply:
(1) Except as provided in paragraph (c)(2) of this section, these items must be positioned so that if they break loose they will be unlikely to:
(i) Cause direct injury to occupants;
(ii) Penetrate fuel tanks or lines or cause fire or explosion hazard by damage to adjacent systems; or
(iii) Nullify any of the escape facilities provided for use after an emergency landing.
(2) When such positioning is not practical (e.g. fuselage mounted engines or auxiliary power units) each such item of mass shall be restrained under all loads up to those specified in paragraph (b)(3) of this section. The local attachments for these items should be designed to withstand 1.33 times the specified loads if these items are subject to severe wear and tear through frequent removal (e.g. quick change interior items).
(d) Seats and items of mass (and their supporting structure) must not deform under any loads up to those specified in paragraph (b)(3) of this section in any manner that would impede subsequent rapid evacuation of occupants.
§ 25.562   Emergency landing dynamic conditions.
(a) The seat and restraint system in the airplane must be designed as prescribed in this section to protect each occupant during an emergency landing condition when—
(1) Proper use is made of seats, safety belts, and shoulder harnesses provided for in the design; and
(2) The occupant is exposed to loads resulting from the conditions prescribed in this section.
(b) Each seat type design approved for crew or passenger occupancy during takeoff and landing must successfully complete dynamic tests or be demonstrated by rational analysis based on dynamic tests of a similar type seat, in accordance with each of the following emergency landing conditions. The tests must be conducted with an occupant simulated by a 170-pound anthropomorphic test dummy, as defined by 49 CFR Part 572, Subpart B, or its equivalent, sitting in the normal upright position.
(1) A change in downward vertical velocity (Δ v) of not less than 35 feet per second, with the airplane's longitudinal axis canted downward 30 degrees with respect to the horizontal plane and with the wings level. Peak floor deceleration must occur in not more than 0.08 seconds after impact and must reach a minimum of 14g.
(2) A change in forward longitudinal velocity (Δ v) of not less than 44 feet per second, with the airplane's longitudinal axis horizontal and yawed 10 degrees either right or left, whichever would cause the greatest likelihood of the upper torso restraint system (where installed) moving off the occupant's shoulder, and with the wings level. Peak floor deceleration must occur in not more than 0.09 seconds after impact and must reach a minimum of 16g. Where floor rails or floor fittings are used to attach the seating devices to the test fixture, the rails or fittings must be misaligned with respect to the adjacent set of rails or fittings by at least 10 degrees vertically ( i.e. , out of Parallel) with one rolled 10 degrees.
(c) The following performance measures must not be exceeded during the dynamic tests conducted in accordance with paragraph (b) of this section:
(1) Where upper torso straps are used for crewmembers, tension loads in individual straps must not exceed 1,750 pounds. If dual straps are used for restraining the upper torso, the total strap tension loads must not exceed 2,000 pounds.
(2) The maximum compressive load measured between the pelvis and the lumbar column of the anthropomorphic dummy must not exceed 1,500 pounds.
(3) The upper torso restraint straps (where installed) must remain on the occupant's shoulder during the impact.
(4) The lap safety belt must remain on the occupant's pelvis during the impact.
(5) Each occupant must be protected from serious head injury under the conditions prescribed in paragraph (b) of this section. Where head contact with seats or other structure can occur, protection must be provided so that the head impact does not exceed a Head Injury Criterion (HIC) of 1,000 units. The level of HIC is defined by the equation:
(6) Where leg injuries may result from contact with seats or other structure, protection must be provided to prevent axially compressive loads exceeding 2,250 pounds in each femur.
(7) The seat must remain attached at all points of attachment, although the structure may have yielded.
(8) Seats must not yield under the tests specified in paragraphs (b)(1) and (b)(2) of this section to the extent they would impede rapid evacuation of the airplane occupants.
§ 25.563   Structural ditching provisions.
Structural strength considerations of ditching provisions must be in accordance with § 25.801(e).