under two conditions: (1) With power off, and (2) With a power setting of not less than that required to show compliance with the provisions of §3.85 (a) for airplanes of more than 6,000 pounds maximum weight, or with 90 percent of maximum continuous power for airplanes of 6,000 pounds or less maximum weight. (b) In either condition required by paragraph (a) of this section it shall be possible, with flaps and landing gear in any position, with center of gravity in the position least favorable for recovery, and with appropriate airplane weights, to show compliance with the applicable requirements of paragraphs (c) through (f) of this section. (c) For airplanes having independently controlled rolling and directional controls, it shall be possible to produce and to correct roll by unreversed use of the rolling control and to produce and correct yaw by unreversed use of the directional control up until the time the airplane pitches in the maneuver prescribed in paragraph (g) of this section. (d) For two-control airplanes having either interconnected lateral and directional controls or for airplanes having only one of these controls, it shall be possible to produce and to correct roll by unreversed use of the rolling control without producing excessive yaw up until the time the airplane pitches in the maneuver prescribed in paragraph (g) of this section. (e) During the recovery portion of the maneuver, it shall be possible to prevent more than 15 degrees roll or yaw by the normal use of controls, and any loss of altitude in excess of 100 feet or any pitch in excess of 30 degrees below level shall be entered in the Airplane Flight Manual. (f) A clear and distinctive stall warning shall precede the stalling of the airplane, with the flaps and landing gear in any position, both in straight and turning flight. The stall warning shall begin at a speed exceeding that of stalling by not less than 5 but not more than 10 miles per hour and shall continue until the stall occurs. (g) In demonstrating the qualities required by paragraphs (c) through (f) of this section, the procedure set forth in subparagraphs (1) and (2) of this paragraph shall be followed. (1) With trim controls adjusted for straight flight at 1.5V, or at the minimum trim speed, whichever is higher, the speed shall be reduced by means of the elevator control until the speed is slightly above the stalling speed; then (2) The elevator control shall be pulled back at a rate such that the airplane speed reduction does not exceed 1 mile per hour per second until a stall is produced as evidenced by an uncontrollable downward pitching motion of the airplane, or until the control reaches the stop. Normal use of the elevator control for recovery shall be allowed after such pitching motion has unmistakably developed. [21 F.R. 3339, May 22, 1956, as amended by Amdt. 3-5, 24 F.R. 7066, Sept. 1, 1959] § 3.120-1 Measuring loss of altitude during stall (FAA policies which apply to § 3.120). To meet the requirements of §3.120, pertaining to the maximum loss of altitude permitted during the stall, it is necessary that a suitable method be used for the purpose of measuring such loss during the investigation of stalls. Unless special features of an individual type being investigated render the following instructions inapplicable, the procedure described shall be used for this purpose: (a) The standard procedure for approaching a stall shall be used as specified in §3.120. (b) The loss of altitude encountered in the stall (power on or power off) shall be the distance as observed on the sensitive altimeter testing installation from the moment the airplane pitches to the observed altitude reading at which horizontal flight has been regained. (c) Power used during the recovery portions of a stall maneuver may be that which, at the discretion of the inspector, would be likely used by a pilot under normal operating conditions when executing this particular maneuver. However, the power used to regain level flight shall not be applied until the airplane has regained flying control at a speed of approximately 1.2 Vs1. This means that in the investigation of stalls with the critical engine inoperative, the power may be reduced on the operating engine(s) before reapplying power on the operating engine or engines for the purpose of regaining level flight. [Supp. 1, 12 F. R. 3435, May 28, 1947, as amended by Amdt. 1, 14 F. R. 36, Jan. 5, 1949] § 3.120-2 to Indications of stall warnings (FAA policies which apply § 3.120). (a) No precise and complete description of the various warnings that would comply with § 3.120 can be given at this time, but the following lists of items may be used as a guide: (1) Satisfactory items include: (1) Buffeting, which may be defined as general shaking or vibration of the airplane, elevator nibble, aileron nibble, rudder nibble, audible indications such as oil canning of structural members or covering roughness in riding qualities of the airplane due to aerodynamic disturbances, etc. (ii) Stall warning instrument, either visual or aural. A visual instrument could be either a light or a dial. (iii) Stick force, defined as heavy. (2) Unsatisfactory items include: (ii) Inability to hold heading. When stalled from an excessive climb attitude it shall be possible to recover from this maneuver without exceeding the limiting air speed or the allowable acceleration limit. § 3.121-1 Climbing stall flight tests for limited control airplanes (FAA inter pretations which apply to § 3.121). (a) This requirement is intended to draw particular attention to any stall recovery characteristics that might be encountered when a limited control airplane is completely stalled from an extremely nose high attitude, either intentionally or inadvertently. In practice it is possible that the elevator control travel could be limited to such an extent that stalls could not be obtained at the normal rate of deceleration used in testing. However, if the airplane was pulled up into a very steep climbing attitude from reasonably high speed flight either power on or power off, and held in this attitude, excessive pitching may occur. At the same time, the limited elevator travel may retard recovery from the pitched attitude until excessively high speeds are obtained. These characteristics would normally be considered under § 3.106; however, it appears wise to call particular attention to the control characteristics that might result from these flight configurations on limited control airplanes. (b) Although Form ACA-283-03, item A, (3), (a), indicates that take-off power should be used for these tests, this is not a mandatory requirement. In this regard it is to be noted that although §3.121 is entitled "Climbing Stalls", it specifically states: ". . . when stalled from an excessive climb attitude”, thus a specified application of power is not required. For example, flight tests recently conducted on several aircraft have indicated that the power-off configuration was critical since the stall resulted in greater pitch and less elevator control. The technique used for inducing such stalls consisted of stalling the airplane (power off) in as steep a climbing attitude as possible without falling into a whip stall, or other flight maneuver that might overstress the structure. (Form ACA-283-03 will be revised at the next printing, so that the power found to be critical can be recorded in a space that will be provided for this purpose.) [Supp. 10, 16 F. R. 3284, Apr. 14, 1951] § 3.122 Turning flight stalls. When stalled during a coordinated 30degree banked turn with 75 percent maximum continuous power on all engines, flaps and landing gear retracted, it shall be possible to recover to normal level flight without encountering excessive loss of altitude, uncontrollable rolling characteristics, or uncontrollable spinning tendencies. These qualities shall be demonstrated by performing the following maneuver: After a steady curvilinear level coordinated flight condition in a 30-degree bank is established and while maintaining the 30-degree bank, the airplane shall be stalled by steadily and progressively tightening the turn with the elevator control until the airplane is stalled or until the elevator has reached its stop. When the stall has fully developed, recovery to level flight shall be made with normal use of the controls. §3.123 One-engine-inoperative stalls. Multiengine airplanes shall not display any undue spinning tendency and shall be safely recoverable without applying power to the inoperative engine when stalled with: (a) The critical engine inoperative, (b) Flaps and landing gear retracted, (c) The remaining engines operating at up to 75 percent of maximum continuous power, except that the power need not be greater than that at which the use of maximum control travel just holds the wings laterally level in approaching the stall. The operating engines may be throttled back during the recovery from the stall. SPINNING §3.124 Spinning. (a) Category N. All airplanes of 4,000 lbs. or less maximum weight shall recover from a one-turn spin with the controls applied normally for recovery in not more than one additional turn and without exceeding either the limiting air speed or the limit positive maneuvering load factor for the airplane. In addition, there shall be no excessive back pressure either during the spin or in the recovery. It shall not be possible to obtain uncontrollable spins by means of any possible use of the controls. Compliance with these requirements shall be demonstrated at any permissible combination of weight and center of gravity positions obtainable with all or any part of the designed useful load. All airplanes in category N, regardless of weight, shall be placarded against spins or demonstrated to be "characteristically incapable of spinning" in which case they shall be so designated. (See paragraph (d) of this section.) (b) Category U. Airplanes in this category shall comply with either the entire requirements of paragraph (a) of this section or the entire requirements of paragraph (c) of this section. (c) Category A. All airplanes in this category shall be capable of spinning and shall comply with the following: (1) At any permissible combination of weight and center of gravity position obtainable with all or part of the design useful load, the airplane shall recover from a six-turn spin, or from any point in a six-turn spin, in not more than 11⁄2 additional turns after the application of the controls in the manner normally used for recovery. (2) It shall be possible to recover from the maneuver prescribed in subparagraph (1) of this paragraph without exceeding either the limiting air speed or the limit positive maneuvering load factor of the airplane. (3) It shall not be possible to obtain uncontrollable spins by means of any possible use of the controls. (4) A placard shall be placed in the cockpit of the airplane setting forth the use of the controls required for recovery from spinning maneuvers. (d) Category NU. When it is desired to designate an airplane as a type "characteristically incapable of spinning,” the flight tests to demonstrate this characteristic shall also be conducted with: (1) A maximum weight percent in excess of the weight for which approval is desired, (2) A center of gravity at least 3 percent aft of the rearmost position for which approval is desired, (3) An available up-elevator travel 4 degrees in excess of that to which the elevator travel is to be limited by appropriate stops. (4) An available rudder travel 7 degrees, in both directions, in excess of that to which the rudder travel is to be limited by appropriate stops. § 3.124-1 Spin tests for category N airplanes (FAA interpretations which apply to § 3.124(a)). If during recovery from a one-turn flaps-down spin the airplane exceeds the placard flap speed or limit load factor, it is permissible to retract the flaps during recovery to avoid exceeding these limits. [Supp. 10, 16 F. R. 3284, Apr. 14, 1951] §3.124-2 Spin tests for category A airplanes (FAA interpretations which apply to § 3.124(c)). If during recovery from a one-turn flaps-down spin the airplane exceeds the placard flap speed or limit load factor, it is permissible to retract the flaps during recovery to avoid exceeding these limits. In addition the airplane is to be placarded "International spins with flaps down prohibited." [Supp. 10, 16 F. R. 3284, Apr. 14, 1951] GROUND AND WATER CHARACTERISTICS §3.143 Requirements. All airplanes shall comply with the requirements of §§ 3.144 to 3.147. § 3.144 Longitudinal stability and control. There shall be no uncontrollable tendency for landplanes to nose over in any operating condition reasonably expected for the type, or when rebound occurs during landing or take-off. Wheel brakes shall operate smoothly and shall exhibit no undue tendency to induce nosing over. Seaplanes shall exhibit no dangerous or uncontrollable porpoising at any speed at which the airplane is normally operated on the water. §3.145 Directional stability and control. (a) There shall be no uncontrollable looping tendency in 90-degree cross winds up to a velocity equal to 0.2 Vo at any speed at which the aircraft may be expected to be operated upon the ground or water. (b) All landplanes shall be demonstrated to be satisfactorily controllable with no exceptional degree of skill or alertness on the part of the pilot in power-off landings at normal landing speed and during which brakes or engine power are not used to maintain a straight path. (c) Means shall be provided for adequate directional control during taxying. §3.146 Shock absorption. The shock-absorbing mechanism shall not produce damage to the structure when the airplane is taxied on the roughest ground which it is reasonable to expect the airplane to encounter in normal operation. §3.147 Spray characteristics. For seaplanes, spray during taxiing, takeoff, and landing shall at no time dangerously obscure the vision of the pilots nor produce damage to the propeller or other parts of the airplane. FLUTTER AND VIBRATION §3.159 Flutter and vibration. All parts of the airplane shall be demonstrated to be free from flutter and excessive vibration under all speed and power conditions appropriate to the operation of the airplane up to at least the minimum value permitted for Va in §3.184. There shall also be no buffeting condition in any normal flight condition severe enough to interfere with the satisfactory control of the airplane or to cause excessive fatigue to the crew or result in structural damage. However, buffeting as stall warning is considered desirable and discouragement of this type of buffeting is not intended. Subpart C-Strength Requirements GENERAL (a) Strength requirements are specified in terms of limit and ultimate loads. Limit loads are the maximum loads anticipated in service. Ultimate loads are equal to the limit loads multiplied by the factor of safety. Unless otherwise described, loads specified are limit loads. (b) Unless otherwise provided, the specified air, ground, and water loads shall be placed in equilibrium with inertia forces, considering all items of mass in the airplane. All such loads shall be distributed in a manner conor servatively approximating closely representing actual conditions. If deflections under load would change significantly the distribution of external or internal loads, such redistribution shall be taken into account. (c) Simplified structural design criteria shall be acceptable if the Administrator finds that they result in design loads not less than those prescribed in §§ 3.181 through 3.265. § 3.171-1 Design criteria (FAA policies which apply to § 3.171(c)). The Administrator finds that the simplified structural design criteria contained in Appendix A1 to Civil Aeronautics Manual 3, result in design loads not less than those prescribed in §§ 3.181 through 3.265. [Supp. 16, 17 F. R. 11786, Dec. 30, 1952] § 3.171-2 Design loads and load distributions (FAA policies which apply to § 3.171(b)). The simplified method in Appendix D' to Civil Aeronautics Manual 3 may be used to determine the air loads and air load distributions resulting from the use of tip stores for low speed, low altitude (design Mach number less than 0.4; design altitude less than 15,000 ft.) airplanes with small amounts of sweep (i.e., mid-chord angles of sweep less than 15 degrees). [Supp. 30, 22 F. R. 10016, Dec. 13, 1957] §3.172 Factor of safety. The factor of safety shall be 1.5 unless otherwise specified. §3.173 Strength and deformations. The structure shall be capable of supporting limit loads without suffering detrimental permanent deformations. At all loads up to limit loads, the deformation shall be such as not to interfere with safe operation of the airplane. The structure shall be capable of supporting ultimate loads without failure for at least 3 seconds, except that when proof of strength is demonstrated by dynamic tests simulating actual conditions of load application, the 3-second limit does not apply. §3.173-1 Dynamic tests (FAA policies which apply to § 3.173). (a) Section 3.173 permits dynamic testing in lieu of stress analysis or static testing in the proof of compliance of the structure with strength and deformation requirements. In demonstrating, by dynamic tests, proof of strength of landing gears for the stipulated landing conditions contained in §§ 3.245, 3.246, and 3.247, it is necessary to employ a procedure which will not result in the accepting of landing gears weaker than those 1 Not filed for publication in the Office of the Federal Register. qualified for acceptance under present procedures, i.e., stress analysis or static testing. (b) The Administrator will accept, as an adequate procedure for this purpose, the following dynamic tests: The structure shall be dropped a minimum of 10 times from the limit drop height, and at least one time from the ultimate drop height, for each basic design condition for which proof of strength is being made by drop tests. (c) With regard to the extent to which the structure can be proved by dynamic tests, such dynamic tests shall be accepted as proof of strength for only those elements of the structure for which it can be shown that the critical limit and ultimate loads have been reproduced. [Supp. 1, 12 F. R. 3435, May 28, 1947, as amended by Amdt. 1, 14 F. R. 36, Jan. 5, 1949] §3.174 Proof of structure. Proof of compliance of the structure with the strength and deformation requirements of § 3.173 shall be made for all critical loading conditions. Proof of compliance by means of structural analysis will be accepted only when the structure conforms with types for which experience has shown such methods to be reliable. In all other cases substantiating load tests are required. Dynamic tests including structural flight tests shall be acceptable, provided that it is demonstrated that the design load conditions have been simulated. In all cases certain portions of the structure must be subjected to tests as specified in Subpart D of this part. § 3.174-1 Material correction factors (FAA policies which apply to §3.174). (a) In tests conducted for the purpose of establishing allowable strengths of structural elements such as sheet, sheet stringer combinations, riveted joints, etc., test results should be reduced to values which would be met by elements of the structure if constructed of materials having properties equal to design allowable values. Material correction factors in this case may be omitted, however, if sufficient test data are obtained to permit a probability analysis showing that 90 percent or more of the elements will either equal or exceed in strength the |