§ 4b.152-1 Procedure for demonstrating stability during landing (FAA policies which apply to § 4b.152). (a) Configuration. This test should be conducted in the configuration that follows: Weight-Maximum landing. C. G. position-Most forward and/or most forward at maximum landing weight and most aft. Wing flaps-Maximum landing position. Engines Power off, propellers windmilling. (b) Test procedure and required data. The airplane should be trimmed at a speed of 1.4 V., at any optional altitude (see § 4b.100-3 (c)). Static longitudinal stability should be demonstrated at speeds from just above the stall to 1.8 Vs, (or placard speed). The elevator control force necessary to maintain each speed should be recorded at approximate even increments of velocity within the above speed range. The speed range representing the maximum friction band from which the airplane will not return to the trim speed should be recorded. For aircraft having an increase in stable elevator control force gradient with forward C. G. travel, only the maximum force need be checked at forward C. G. Lateral and directional trim should also be checked (see § 4b.141-1). The following data should be recorded: Weight. C. G. position. Wing flap position Landing gear position. Engines, rpm and manifold pressure. Ambient air temperature. Elevator control force. [Supp. 24, 19 F. R. 4458, July 20, 1954] § 4b.153 Stability during approach. The stick force curve shall have a stable slope at all speeds between 1.1 V11 and 1.8 Vs, with: (a) Wing flaps in sea level approach position, (b) Landing gear retracted, (c) Maximum landing weight, (d) The airplane trimmed at 1.4 V11 and with power sufficient to maintain level flight at this speed. § 4b.153-1 Procedure for demonstrating stability during approach (FAA policies which apply to § 4b.153). (a) Configuration. This test should be conducted in the configuration that follows: Weight-Maximum landing. C. G. position-Most aft. Wing flaps-Approach position. Engines Power required for level flight at 1.4 V1 Cowl flaps-Approach position. (b) Test procedure and required data. The same procedures and data as outlined in § 4b.152-1 (b) should be followed in demonstrating stability for the approach configuration. Supp. 24, 19 F. R. 4458, July 20, 1954] § 4b.154 Stability during climb. The stick force curve shall have a stable slope at all speeds between 85 and 115 percent of the speed at which the airplane is trimmed with: (a) Wing flaps retracted, (d) 75 percent of maximum continuous power for reciprocating-enginepowered airplanes; and maximum power thrust selected by the applicant as an operating limitation for use during climb (see § 4b.718) for turbine-enginepowered airplanes. (e) The airplane trimmed at the best rate-of-climb speed, except that the speed need not be less than 1.4 Vs,. [15 F. R. 3543, June 8, 1950, as amended by Amdt. 4b-2, 20 F. R. 5305, July 26, 1955] § 4b.154-1 Procedure for demonstrating stability during climb (FAA policies which apply to § 4b.154). This test should be conducted in the configuration that follows: (a) Configuration. Weight-Maximum take-off. (b) Test procedure and required data. The airplane should be trimmed at the best rate of climb speed except that this speed need not be less than 1.4 V11. Static longitudinal stability should be demonstrated at speeds from just above the stall to the speed at which the control force becomes excessive but not to exceed VNE. The curve of elevator control force vs. speed should have a stable slope between 85 percent and 115 percent of the trim speed and no reversal of elevator control force should occur throughout the speed range tested. Further test procedures and data to be recorded should be the same as are specified in § 4b.152-1 (b). [Supp. 24, 19 F.R. 4458, July 20, 1954, as amended by Supp. 34, 22 F.R. 6963, Aug. 29, 1957] § 4b.155 Stability during cruising. (a) Landing gear retracted. Between 1.3 V., and VNE the stick force curve shall have a stable slope at all speeds obtainable with a stick force not in excess of 50 pounds with: (1) Wing flaps retracted, (3) 75 percent of maximum continuous power, or the maximum cruising power selected by the applicant as an operating limitation (see § 4b.718), whichever is the greater, except that the power need not exceed that required at VNO, (b) Landing The (4) The airplane trimmed for level flight with the power specified in subparagraph (3) of this paragraph. extended. gear stick force curve shall have a stable slope at all speeds between 1.3 V1, and the speed at which the airplane is trimmed, except that the range of speeds need not exceed that obtainable with a stick force of 50 pounds with: (1) Wing flaps retracted, (2) Maximum take-off weight, (3) 75 percent of maximum continuous power, or the maximum cruising power selected by the applicant as an operating limitation, whichever is the greater, except that the power need not exceed that required for level flight at LE V (4) The airplane trimmed for level flight with the power specified in subparagraph (3) of this paragraph. [15 F. R. 3543, June 8, 1950, as amended by Amdt. 4b-2, 20 F. R. 5305, July 26, 1955] § 4b.155-1 Procedure for demonstrat ing stability during cruising (FAA policies which apply to § 4b.155). (a) Cruising, landing gear retracted, 4b.155(a)-(1) Configuration. This test should be conducted in the configuration that follows: (2) Test procedure and required data. The airplane should be trimmed at the speed for level flight with the power specified in § 4b.155 (b) (3). Static longitudinal stability should be demonstrated at speeds from just above the stall to the speed at which the control forces become excessive (50 lbs.), but not to exceed VNE. Further test procedure and data to be recorded should be the same as are specified in § 4b.152-1 (b). (b) Cruising, landing gear extended, 84b.155 (b)-(1) Configuration. This test should be conducted in the configuration that follows: Weight-Maximum take-off. Engines-The power specified in § 4b.155(b) (3). Cowl flaps-Appropriate for flight condition. (2) Test procedure and required data. The airplane should be trimmed at the speed for level flight with the power specified in § 4b.155 (b) (3). Static longitudinal stability should be demonstrated at speeds from just above the stall to the speed at which the control forces become excessive (50 lbs.), but not to exceed VLE. Further test procedures and data to be recorded should be the same as are specified in § 4b.152-1 (b). [Supp. 24, 19 F.R. 4458, July 20, 1954, as amended by Supp. 34, 22 F.R. 6963, Aug. 29, 1957] § 4b.156 Dynamic longitudinal stability. Any short period oscillation occurring between stalling speed and maximum permissible speed appropriate to the configuration of the airplane shall be heavily damped with the primary controls free and in a fixed position. § 4b.156-1 Procedure for demonstrating dynamic longitudinal stability (FAA policies which apply to § 4b.156). Damping of accelerations and movement of the control should be noted when: (a) The control column is quickly offset and immediately released and (b) The control column is quickly offset and immediately returned to the trim position and held in this position. [Supp. 24, 19 F. R. 4458, July 20, 1954] § 4b.157 Static directional and lateral stability. (a) The static directional stability, as shown by the tendency to recover from a skid with rudder free, shall be positive with all landing gear and flap positions and symmetrical power conditions, at all speeds from 1.2 V., up to the operating limit speed. (b) The static lateral stability, as shown by the tendency to raise the low wing in a sideslip with the aileron controls free and with all landing gear and flap positions and symmetrical power conditions, shall: (1) Be positive at the operating limit speed, (2) Not be negative at a speed equal to 1.2 V11. (c) In straight steady sideslips (unaccelerated forward slips) the aileron and rudder control movements and forces shall be substantially proportional to the angle of sideslip, and the factor of proportionality shall lie between limits found necessary for safe operation throughout the range of sideslip angles appropriate to the operation of the airplane. At greater angles up to that at which the full rudder control is employed or a rudder pedal force of 180 pounds is obtained, the rudder pedal forces shall not reverse, and increased rudder deflection shall produce increased angles of sideslip. Sufficient bank shall accompany sideslipping to indicate clearly any departure from steady unyawed flight, unless a yaw indicator is provided. [15 F. R. 3543, June 8, 1950, as amended by Amdt. 4b-6, 17 F. R. 1089, Feb. 5, 1952] § 4b.157-1 Procedure for demonstrating static directional and lateral stability (FAA policies which apply to § 4b.157). (a) Motion involving roll. No real motion of the airplane involving roll is possible without yaw also being involved, and vice versa. In showing compliance with § 4b.157 the rolling and yawing stability should be investigated separately. (b) Directional stability. Directional stability should be investigated by starting from steady flight in the required configuration and deflecting the rudder at a fairly rapid rate by the amount required to maintain a steady skid with the airplane yawed approximately 20° (as read on the directional gyro) while the wings are maintained level by use of the ailerons, and the speed held constant by means of the elevator control. When the steady condition has been established, the rudder should be released and, if the airplane is directionally stable, it should cease to skid; i. e., the yaw should decrease to approximately zero and, if also laterally stable the aileron deflection and force required to hold the wings level should also approach zero. The test should be conducted by executing skids both to the right and left, recording in each case the time required from the release of the rudder controls and the number of oscillations, if any, involved to recover to steady level flight. (c) Lateral stability. Lateral stability should be investigated by starting from steady flight in the required configuration and banking the airplane approximately 20° (as read on the gyro horizon) by means of the ailerons, while maintaining the original heading by means of the rudder, and the original speed by means of the longitudinal trimming device. When the steady slipping condition has been established, the aileron control should be released. If the airplane is laterally stable, it should cease to slip; i. e., the wing should return to an approximately level attitude, and the rudder deflection and pedal force required to maintain the heading should approach zero. The test should be conducted by executing slips from both to right and left, and in each case the time required from the release of the aileron control and the number of oscillations, if any, involved to recover to steady level flight should be recorded. (d) Additional test. In addition to the directional and lateral stability tests, § 4b.157 (c) contains provisions which should be used to test the airplane for rudder over balance. (e) Static directional stability test, § 4b.157 (a) and (c). CAUTION: Prior to conducting this test and that in paragraph (f) of this section, complete agreement should be reached between the applicant and the FAA Flight Test Agent to insure that the severity of control application will not result in loads exceeding the design limitations. (1) Configuration. This test should be conducted in the configurations that follow: Maximum take-off weight with wing flaps retracted. Maximum landing weight with wing flaps extended. C. G. position-Most aft. Wing flaps-Retracted and maximum landing position. Landing gear-Retracted and extended. (2) Test procedure and required data. The following tests should be conducted at the altitude deemed most critical for the combination of power and aerodynamic damping effect: (i) The airplane should be yawed slowly to the left and right using ailerons to hold wings level, and, when controls are released slowly, the tendency of airplane to recover from the skid should be noted. (ii) The qualitative proportionality of rudder and aileron deflection and force during steady straight sideslips should be noted. (iii) Damping of yawing and movement of control should be noted when the rudder is quickly offset and immediately released and when the rudder is quickly offset and immediately returned and held in the trim position. (3) The tests in subparagraph (2) of this paragraph should be conducted in the following configurations: (i) Flaps in landing position and gear extended, at 1.2 V11, power off, and 75 percent maximum continuous power. (ii) The flaps and gear retracted at 1.2 V11, and Ve with 75 percent maximum continuous power. (iii) Flaps and gear retracted at 1.2 V11, with power off. (4) The following data should be recorded for the tests specified in subparagraphs (2) and (3) of this paragraph: Weight. C. G. position. Wing flap position. Landing gear position. Engines, rpm and manifold pressure. Ambient air temperature. Air speed at VFE, 1.2 Vs, and VC. (f) Static lateral stability test, § 4b.157 (b)—(1) Configuration. This test should be conducted in the configurations that follow: Maximum take-off weight with flaps retracted. Maximum landing weight with flaps extended. C. G. position-Most aft. Wing flaps-Retracted and maximum landing position. Landing gear-Retracted and extended. (2) Test procedure and required data. The following tests should be conducted in the configurations specified in paragraph (e) (3) of this section and at the altitude deemed most critical for the combination of power and aerodynamic damping effect: (i) Starting from steady straight flight the airplane should be banked 20° while a constant heading is held; the aileron control should then be released. The stability as evidenced by the tendency to raise the low wing should be positive at high speed and should not be negative at 1.2 V11. (ii) Damping of rolling motion and movement of controls should be noted when the aileron is quickly offset and immediately released and also when the aileron is quickly offset and immediately returned and held in the trim position. (iii) The same data as specified in paragraph (e) (4) of this section should be recorded at air speeds of 1.2 V., and Vc. [Supp. 24, 19 F.R. 4458, July 20, 1954, as amended by Supp. 34, 22 F.R. 6963, Aug. 29, 1957] § 4b.158 Dynamic directional and lateral stability. Any short period oscillation occurring between stalling speed and maximum permissible speed appropriate to the configuration of the airplane shall be heavily damped with the primary controls free and in a fixed position. § 4b.158-1 Procedure for demonstrating dynamic directional and lateral stability (FAA policies which apply to § 4b.158). Damping of yawing and movement of the control should be noted during the test procedure in § 4b.157-1(e) (2) (iii). [Supp. 24, 19 F.R. 4459, July 20, 1954] STALLING CHARACTERISTICS § 4b.160 Stalling; symmetrical power. (a) Stalls shall be demonstrated with the airplane in straight flight and in banked turns at 30 degrees, both with power off and with power on. In the power-on conditions the power shall be that necessary to maintain level flight at a speed of 1.6 Vs1, where Vs, corresponds with the stalling speed with flaps in the approach position, the landing gear retracted, and maximum landing weight. (b) The stall demonstration shall be in the following configurations: (1) Wing flaps and landing gear in any likely combination of positions, (2) Representative weights within the range for which certification is sought, (3) The center of gravity in the most adverse position for recovery. (c) The stall demonstration shall be conducted as follows: (1) With trim controls adjusted for straight flight at a speed of 1.4 Vs,, the speed shall be reduced by means of the elevator control until it is steady at slightly above stalling speed; after which the elevator control shall be applied at a rate such that the airplane speed reduction does not exceed one mile per hour per second until the airplane is stalled or, if the airplane is not stalled, until the control reaches the stop. (2) The airplane shall be considered stalled when, at an angle of attack measurably greater than that of maximum lift, the inherent flight characteristics give a clear indication to the pilot that the airplane is stalled, except that for airplanes demonstrating unmistakable inherent aerodynamic warning associated with the stall in all required configurations, the speed need not be reduced below a value which provides an adequate stall warning margin as defined in § 4b.162. NOTE: A nose-down pitch or a roll which cannot be readily arrested are typical indications that the airplane is stalled. Other indications, such as marked loss of control effectiveness, abrupt change in control force or motion, characteristic buffeting, or a distinctive vibration of the pilot's controls, may be accepted if found in a particular case to be sufficiently clear. Types of inherent aerodynamic warning considered acceptable include characteristics such as buffeting, small amplitude pitch or roll oscillations, distinctive shaking of the pilots' controls, etc. (3) Recovery from the stall shall be effected by normal recovery techniques, starting as soon as the airplane is stalled. (d) During stall demonstration it shall be possible to produce and to correct roll and yaw by unreversed use of the aileron and rudder controls up to the moment the airplane is stalled; there shall occur no abnormal nose-up pitching; and the longitudinal control force shall be positive up to and including the stall. (e) In straight flight stalls the roll occurring between the stall and the completion of the recovery shall not exceed approximately 20 degrees. (f) In turning flight stalls the action of the airplane following the stall shall 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. (g) In both the straight flight and the turning flight stall demonstrations it shall be possible promptly to prevent the airplane from stalling and to recover from the stall condition by normal use of the controls. [Amdt. 4b-5, 16 F. R. 12220, Dec. 4, 1951, as amended by Amdt. 4b-3, 21 F. R. 990, Feb. 11, 1956] § 4b.160-1 Procedure for demonstrating stall tests, symmetrical power (FAA policies which apply to § 4b.160 (c) (2)). (a) Angle-of-attack. The angle-ofattack during the stall maneuver should be increased at least to the point where the following two conditions are satisfied: (1) Attainment of an angle-of-attack measurably greater than that for maximum lift. (2) Clear indication to the pilot through the inherent flight characteristics that the airplane is stalled. (b) Procedure to be used. The following procedure may be used to demonstrate that these two conditions are fulfilled. (1) A photopanel or equivalent method of obtaining continuous records of the following variables at not greater than 4 second intervals should be provided: indicated angle-of-attack, swivel static and shielded total pressure head or equivalent, pressure altitude, pitch and bank angle, normal acceleration, elevator position and force, aileron and rudder position. (2) If it is evident that longitudinal stick force is always positive, that is, no reversal exists down to the stall, then time history of this item should not be required. |