(a) With landing gear extended and wing flaps retracted, (b) With landing gear extended and wing flaps extended under the forward center of gravity position approved with the maximum weight, (c) With landing gear extended and wing flaps extended under the most forward center of gravity position approved, regardless of weight. (iii) During level flight at any speed from 0.9 Vn to Vr or 1.4 Vs, with landing gear and wing flaps retracted. (b) In addition to the above, multiengine airplanes shall maintain longitudinal and directional trim at a speed between V, and 1.4 V., during climbing flight with the critical of two or more engines inoperative, with: (1) The other engine(s) operating at maximum continuous power, (2) The landing gear retracted, (4) Bank not in excess of 5 degrees. [21 F.R. 3339, May 22, 1956, as amended by Amdt. 3-5, 24 F.R. 7066, Sept. 1, 1959] STABILITY §3.113 General. The airplane shall be longitudinally, directionally, and laterally stable in accordance with the following sections. Suitable stability and control "feel" (static stability) shall be required in other conditions normally encountered in service, if flight tests show such stability to be necessary for safe operation. §3.114 Static longitudinal stability. In the configurations outlined in §3.115 and with the airplane trimmed as indicated, the characteristics of the elevator control forces and the friction within the control system shall be such that: (a) A pull shall be required to obtain and maintain speeds below the specified trim speed and a push to obtain and maintain speeds above the specified trim speed. This shall be so at any speed which can be obtained without excessive control force, except that such speeds need not be greater than the appropriate maximum permissible speed or less than the minimum speed in steady unstalled flight. (b) The air speed shall return to within 10 percent of the original trim speed when the control force is slowly released from any speed within the limits defined in paragraph (a) of this section. §3.115 Specific conditions. In conditions set forth in this section, within the speeds specified, the stable slope of stick force versus speed curve shall be such that any substantial change in speed is clearly perceptible to the pilot through a resulting change in stick force. (a) Approach. The stick force curve shall have a stable slope and the stick force shall not exceed 40 lbs. at any speed between 1.1 V., and 1.8 Vs, with: (1) Wing flaps in the landing position, (4) Airplane trimmed at 1.5 V11 and power on as required to maintain a 3 degree angle of descent. (b) Climb. The stick force curve shall have a stable slope at all speeds between 1.2 Vs, and 1.6 Vs, with: 1 (1) Wing flaps retracted, (2) Landing gear retracted, (4) 75 percent of maximum continuous power, (5) The airplane trimmed at 1.4 V.,. (c) Cruising. (1) Between 1.3 V1, and the maximum permissible speed, the stick force curve shall have a stable slope at all speeds obtainable with a stick force not in excess of 40 pounds with: (i) Landing gear retracted, (ii) Wing flaps retracted. (iii) Maximum weight, (iv) 75 percent of maximum continuous power, (v) The airplane trimmed for level flight with 75 percent of the maximum ($) continuous power. (2) Same as subparagraph (1) of this paragraph, except that the landing gear shall be extended and the level flight trim speed need not be exceeded. [21 F.R. 3339, May 22, 1956, as amended by Amdt. 3-5, 24 F.R. 7066, Sept. 1, 1959] §3.116 Instrumented stick force meas urements. Instrumented stick force measurements need not be made when changes in speed are clearly reflected by changes in stick forces and the maximum forces obtained in the above conditions are not = excessive. §3.117 Dynamic longitudinal stability. Any short period oscillation occurring between stalling speed and maximum permissible speed shall be heavily damped with the primary controls (1) free, and (2) in a fixed position. §3.118 Directional and lateral stability. (a) Three-control airplanes. (1) The static directional stability, as shown by the tendency to recover from a skid with rudder free, shall be positive for all landing gear and flap positions appropriate to the takeoff, climb, cruise, and approach #configurations, with symmetrical power up to maximum continuous power, and at all speeds from 1.2 V11 up to the maximum permissible speed for the configuration being investigated. The angle of Eskid for these tests shall be appropriate to as the type of airplane. At greater angles of skid up to that at which full rudder is employed or a control force limit specified in § 3.106 is obtained, whichever occurs first, and at speeds from 1.2 V11 to Vp, the rudder pedal force shall not reverse. (2) The static lateral stability, shown by the tendency to raise the low wing in a sideslip, shall be positive for all landing gear and flap positions with symmetrical power up to 75 percent maximum continuous power at all speeds above 1.2 V11 up to the maximum permissible speed for the configuration investigated but shall not be negative at a speed of 1.2 V11. The angle of sideslip for these tests shall be appropriate to the type of airplane but in no case shall the sideslip be less than that obtained with 10 degrees of bank. (3) In straight steady sideslips at a speed of 1.2 V11 for all gear and flap positions and for all symmetrical power conditions up to 50 percent maximum continuous power, the aileron and rudder control movements and forces shall increase steadily, but not necessarily in constant proportion, as the angle of sideslip is increased up to the maximum appropriate to the type of airplane. At greater angles up to that at which the full rudder or aileron control is employed or a control force limit specified by § 3.106 is obtained, the rudder pedal force shall > not reverse. Sufficient bank shall accompany sideslipping to prevent departure from a constant heading. Rapid entry into or recovery from a maximum sideslip shall not result in uncontrollable flight characteristics. (4) Any short-period oscillation occurring between stalling speed and maximum permissible speed shall be heavily damped with the primary controls (i) free and (ii) in a fixed position. (b) Two-control (or simplified) airplanes. (1) The directional stability shall be shown to be adequate by demonstrating that the airplane in all configurations can be rapidly rolled from a 45-degree bank to a 45-degree bank in the opposite direction without exhibiting dangerous skidding characteristics. (2) Lateral stability shall be shown to be adequate by demonstrating that the airplane will not assume a dangerous attitude or speed when all the controls are abandoned for a period of 2 minutes. This demonstration shall be made in moderately smooth air with the airplane trimmed for straight level flight at 0.9 Vn (or at Vc, if lower), flaps and gear retracted, and with rearward center of gravity loading. (3) Any short period oscillation occurring between the stalling speed and the maximum permissible speed shall be heavily damped with the primary controls (i) free and (ii) in a fixed position. [21 F.R. 3339, May 22, 1956, as amended by Amdt. 3-5, 24 F.R. 7066, Sept. 1, 1959] §3.118-2 Large displacements of flight controls in directional and lateral stability tests (FAA policies which apply to § 3.118). (a) In performing flight tests to determine compliance with §3.118, it should be borne in mind that the airplane structural requirements do not provide for large displacements of the flight controls at high speeds. Full application of rudder and aileron controls should be confined to speeds below the design maneuvering speed Vp. The following rules (approximations) will serve as a guide for the maximum permissible control surface deflections at speeds above Vp. (This does not imply that these maximum deflections must be used in the tests at high speeds). (1) The permissible rudder angle decreases approximately according to the ratio (VD/V), where V is the speed of the test. (2) The permissible aileron deflection decreases approximately at the ratio (Vp/V), up to the design cruising speed, 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 tal level shall be entered in the Airplane Flight Manual. (f) A clear and distinctive stall warning shall precede the stalling of the air-d plane, with the flaps and landing gear in any position, both in straight and turning flight. The stall warning shall wer 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 1.5Vs, 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] ཐོ §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 Indications of stall warnings (FAA policies which apply to §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: (i) 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. (iii) Inability to hold wing level. [Supp. 10, 16 F. R. 3284, Apr. 14, 1951] §3.121 Climbing stalls. 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 5 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 appro- i priate 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 |