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Part 23 Appendices
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23.1501 General. 23.1541 General. Airplane Flight Manual and Approved Manual Material 23.1581 General. Appendix A--Simplified
Design Load Criteria Authority: 49 U.S.C. 106(g), 40113, 44701-44702,
44704. |
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(a) Each operating limitation specified in Secs. 23.1505 through 23.1527 and other limitations and information necessary for safe operation must be established. (b) The operating limitations and other information necessary for safe operation must be made available to the crewmembers as prescribed in Secs. 23.1541 through 23.1589. [Amdt. 23-21, 43 FR 2319, Jan. 16, 1978] Sec. 23.1505 Airspeed limitations. (a) The never-exceed speed VNE must be established so that it is-- (1) Not less than 0.9 times the minimum value of VD allowed under Sec. 23.335; and (2) Not more than the lesser of-- (i) 0.9 VD established under Sec. 23.335; or (ii) 0.9 times the maximum speed shown under Sec. 23.251. (b) The maximum structural cruising speed VNO must be established so that it is-- (1) Not less than the minimum value of VC allowed under Sec. 23.335; and (2) Not more than the lesser of-- (i) VC established under Sec. 23.335; or (ii) 0.89 VNE established under paragraph (a) of this section. (c) Paragraphs (a) and (b) of this section do not apply to turbine airplanes or to airplanes for which a design diving speed VD/MD is established under Sec. 23.335(b)(4). For those airplanes, a maximum operating limit speed (VMO/MMO-airspeed or Mach number, whichever is critical at a particular altitude) must be established as a speed that may not be deliberately exceeded in any regime of flight (climb, cruise, or descent) unless a higher speed is authorized for flight test or pilot training operations. VMO/MMO must be established so that it is not greater than the design cruising speed VC/MC and so that it is sufficiently below VD/MD and the maximum speed shown under Sec. 23.251 to make it highly improbable that the latter speeds will be inadvertently exceeded in operations. The speed margin between VMO/MMO and VD/MD or the maximum speed shown under Sec. 23.251 may not be less than the speed margin established between VC/MC and VD/MD under Sec. 23.335(b), or the speed margin found necessary in the flight test conducted under Sec. 23.253. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR 13096, Aug. 13, 1969] Sec. 23.1507 Operating maneuvering speed. The maximum maneuvering speed Vo, must be established as an operating limitation. Vo is a selected speed that is not greater than Vs[square root]n established in Sec. 23.335(c). [Amdt. No. 23-45, 58 FR 42165, Aug. 6, 1993] Sec. 23.1511 Flap extended speed. (a) The flap extended speed VFE must be established so that it is-- (1) Not less than the minimum value of VF allowed in Sec. 23.345(b); and (2) Not more than VF established under Sec. 23.345(a), (c), and (d).(i) VF established under Sec. 23.345; or (ii) VF established under Sec. 23.457. (b) Additional combinations of flap setting, airspeed, and engine power may be established if the structure has been proven for the corresponding design conditions. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-50, 61 FR 5192, Feb. 9, 1996] Sec. 23.1513 Minimum control speed. The minimum control speed VMC, determined under Sec. 23.149, must be established as an operating limitation. Sec. 23.1519 Weight and center of gravity. The weight and center of gravity limitations determined under Sec. 23.23 must be established as operating limitations. Sec. 23.1521 Powerplant limitations. (a) General. The powerplant limitations prescribed in this section must be established so that they do not exceed the corresponding limits for which the engines or propellers are type certificated. In addition, other powerplant limitations used in determining compliance with this part must be established. (b) Takeoff operation. The powerplant takeoff operation must be limited by-- (1) The maximum rotational speed (rpm); (2) The maximum allowable manifold pressure (for reciprocating engines); (3) The maximum allowable gas temperature (for turbine engines); (4) The time limit for the use of the power or thrust corresponding to the limitations established in paragraphs (b) (1) through (3) of this section; and (5) The maximum allowable cylinder head (as applicable), liquid coolant and oil temperatures. (c) Continuous operation. The continuous operation must be limited by-- (1) The maximum rotational speed; (2) The maximum allowable manifold pressure (for reciprocating engines); (3) The maximum allowable gas temperature (for turbine engines); and (4) The maximum allowable cylinder head, oil, and liquid coolant temperatures. (d) Fuel grade or designation. The minimum fuel grade (for reciprocating engines), or fuel designation (for turbine engines), must be established so that it is not less than that required for the operation of the engines within the limitations in paragraphs (b) and (c) of this section. (e) Ambient temperature. For all airplanes except reciprocating engine- powered airplanes of 6,000 pounds or less maximum weight, ambient temperature limitations (including limitations for winterization installations if applicable) must be established as the maximum ambient atmospheric temperature at which compliance with the cooling provisions of Secs. 23.1041 through 23.1047 is shown. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23-21, 43 FR 2319, Jan. 16, 1978; Amdt. 23-45, 58 FR 42165, Aug. 6, 1993; Amdt. 23-50, 61 FR 5192, Feb. 9, 1996] Sec. 23.1522 Auxiliary power unit limitations. If an auxiliary power unit is installed, the limitations established for the auxiliary power must be specified in the operating limitations for the airplane. [Amdt. No. 23-45, 58 FR 42166, Aug. 6, 1993; 58 FR 51970, Oct. 5, 1993] Sec. 23.1523 Minimum flight crew. The minimum flight crew must be established so that it is sufficient for safe operation considering-- (a) The workload on individual crewmembers and, in addition for commuter category airplanes, each crewmember workload determination must consider the following: (1) Flight path control, (2) Collision avoidance, (3) Navigation, (4) Communications, (5) Operation and monitoring of all essential airplane systems, (6) Command decisions, and (7) The accessibility and ease of operation of necessary controls by the appropriate crewmember during all normal and emergency operations when at the crewmember flight station; (b) The accessibility and ease of operation of necessary controls by the appropriate crewmember; and (c) The kinds of operation authorized under Sec. 23.1525. [Amdt. 23-21, 43 FR 2319, Jan. 16, 1978, as amended by Amdt. 23-34, 52 FR 1834, Jan. 15, 1987] Sec. 23.1524 Maximum passenger seating configuration. The maximum passenger seating configuration must be established. [Amdt. 23-10, 36 FR 2864, Feb. 11, 1971] Sec. 23.1525 Kinds of operation. The kinds of operation authorized (e.g. VFR, IFR, day or night) and the meteorological conditions (e.g. icing) to which the operation of the airplane is limited of from which it is prohibited, mut be established appropriate to the installed equipment. [Amdt. No. 23-45, 58 FR 42166, Aug. 6, 1993] Sec. 23.1527 Maximum operating altitude. (a) The maximum altitude up to which operation is allowed, as limited by flight, structural, powerplant, functional or equipment characteristics, must be established. (b) A maximum operating altitude limitation of not more than 25,000 feet must be established for pressurized airplanes unless compliance with Sec. 23.775(e) is shown. [Amdt. 23-45, 45 FR 42166, Aug. 6, 1993] Sec. 23.1529 Instructions for Continued Airworthiness. The applicant must prepare Instructions for Continued Airworthiness in accordance with Appendix G to this part that are acceptable to the Administrator. The instructions may be incomplete at type certification if a program exists to ensure their completion prior to delivery of the first airplane or issuance of a standard certificate of airworthiness, whichever occurs later. [Amdt. 23-26, 45 FR 60171, Sept. 11, 1980] |
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(a) The airplane must contain-- (1) The markings and placards specified in Secs. 23.1545 through 23.1567; and (2) Any additional information, instrument markings, and placards required for the safe operation if it has unusual design, operating, or handling characteristics. (b) Each marking and placard prescribed in paragraph (a) of this section-- (1) Must be displayed in a conspicuous place; and (2) May not be easily erased, disfigured, or obscured. (c) For airplanes which are to be certificated in more than one category-- (1) The applicant must select one category upon which the placards and markings are to be based; and (2) The placards and marking information for all categories in which the airplane is to be certificated must be furnished in the Airplane Flight Manual. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23-21, 43 FR 2319, Jan. 16, 1978] Sec. 23.1543 Instrument markings: general. For each instrument-- (a) When markings are on the cover glass of the instrument, there must be means to maintain the correct alignment of the glass cover with the face of the dial; and (b) Each arc and line must be wide enough and located to be clearly visible to the pilot. (c) All related instruments must be calibrated in compatible units. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-50, 61 FR 5192, Feb. 9, 1996] Sec. 23.1545 Airspeed indicator. (a) Each airspeed indicator must be marked as specified in paragraph (b) of this section, with the marks located at the corresponding indicated airspeeds. (b) The following markings must be made: (1) For the never-exceed speed VNE, a radial red line. (2) For the caution range, a yellow arc extending from the red line specified in paragraph (b)(1) of this section to the upper limit of the green arc specified in paragraph (b)(3) of this section. (3) For the normal operating range, a green arc with the lower limit at VS1 with maximum weight and with landing gear and wing flaps retracted, and the upper limit at the maximum structural cruising speed VNO established under Sec. 23.1505(b). (4) For the flap operating range, a white arc with the lower limit at VS0 at the maximum weight, and the upper limit at the flaps-extended speed VFE established under Sec. 23.1511. (5) For reciprocating multiengine-powered airplanes of 6,000 pounds or less maximum weight, for the speed at which compliance has been shown with Sec. 23.69(b) relating to rate of climb at maximum weight and at sea level, a blue radial line. (6) For reciprocating multiengine-powered airplanes of 6,000 pounds or less maximum weight, for the maximum value of minimum control speed, VMC, (one- engine-inoperative) determined under Sec. 23.149(b), a red radial line. (c) If VNE or VNO vary with altitude, there must be means to indicate to the pilot the appropriate limitations throughout the operating altitude range. (d) Paragraphs (b)(1) through (b)(3) and paragraph (c) of this section do not apply to aircraft for which a maximum operating speed VMO/MMO is established under Sec. 23.1505(c). For those aircraft there must either be a maximum allowable airspeed indication showing the variation of VMO/MMO with altitude or compressibility limitations (as appropriate), or a radial red line marking for VMO/MMO must be made at lowest value of VMO/MMO established for any altitude up to the maximum operating altitude for the airplane. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-3, 30 FR 14240, Nov. 13, 1965; Amdt. 23-7, 34 FR 13097, Aug. 13, 1969; Amdt. 23-23, 43 FR 50593, Oct. 30, 1978; Amdt. 23-50, 61 FR 5193, Feb. 9, 1996] Sec. 23.1547 Magnetic direction indicator. (a) A placard meeting the requirements of this section must be installed on or near the magnetic direction indicator. (b) The placard must show the calibration of the instrument in level flight with the engines operating. (c) The placard must state whether the calibration was made with radio receivers on or off. (d) Each calibration reading must be in terms of magnetic headings in not more than 30 degree increments. (e) If a magnetic nonstabilized direction indicator can have a deviation of more than 10 degrees caused by the operation of electrical equipment, the placard must state which electrical loads, or combination of loads, would cause a deviation of more than 10 degrees when turned on. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23-20, 42 FR 36969, July 18, 1977] Sec. 23.1549 Powerplant and auxiliary power unit instruments. For each required powerplant and auxiliary power unit instrument, as appropriate to the type of instruments-- (a) Each maximum and, if applicable, minimum safe operating limit must be marked with a red radial or a red line; (b) Each normal operating range must be marked with a green arc or green line, not extending beyond the maximum and minimum safe limits; (c) Each takeoff and precautionary range must be marked with a yellow arc or a yellow line; and (d) Each engine, auxiliary power unit, or propeller range that is restricted because of excessive vibration stresses must be marked with red arcs or red lines. [Amdt. 23-12, 41 FR 55466, Dec. 20, 1976, as amended by Amdt. 23-28, 47 FR 13315, Mar. 29, 1982; Amdt. No. 23-45, 58 FR 42166, Aug. 6, 1993] Sec. 23.1551 Oil quantity indicator. Each oil quantity indicator must be marked in sufficient increments to indicate readily and accurately the quantity of oil. Sec. 23.1553 Fuel quantity indicator. A red radial line must be marked on each indicator at the calibrated zero reading, as specified in Sec. 23.1337(b)(1). [Amdt. 23-50, 61 FR 5193, Feb. 9, 1996] Sec. 23.1555 Control markings. (a) Each cockpit control, other than primary flight controls and simple push button type starter switches, must be plainly marked as to its function and method of operation. (b) Each secondary control must be suitably marked. (c) For powerplant fuel controls-- (1) Each fuel tank selector control must be marked to indicate the position corresponding to each tank and to each existing cross feed position; (2) If safe operation requires the use of any tanks in a specific sequence, that sequence must be marked on or near the selector for those tanks; (3) The conditions under which the full amount of usable fuel in any restricted usage fuel tank can safely be used must be stated on a placard adjacent to the selector valve for that tank; and (4) Each valve control for any engine of a multiengine airplane must be marked to indicate the position corresponding to each engine controlled. (d) Usable fuel capacity must be marked as follows: (1) For fuel systems having no selector controls, the usable fuel capacity of the system must be indicated at the fuel quantity indicator. (2) For fuel systems having selector controls, the usable fuel capacity available at each selector control position must be indicated near the selector control. (e) For accessory, auxiliary, and emergency controls-- (1) If retractable landing gear is used, the indicator required by Sec. 23.729 must be marked so that the pilot can, at any time, ascertain that the wheels are secured in the extreme positions; and (2) Each emergency control must be red and must be marked as to method of operation. No control other than an emergency control, or a control that serves an emergency function in addition to its other functions, shall be this color. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23-21, 43 FR 2319, Jan. 16, 1978; Amdt. 23-50, 61 FR 5193, Feb. 9, 1996] Sec. 23.1557 Miscellaneous markings and placards. (a) Baggage and cargo compartments, and ballast location. Each baggage and cargo compartment, and each ballast location, must have a placard stating any limitations on contents, including weight, that are necessary under the loading requirements. (b) Seats. If the maximum allowable weight to be carried in a seat is less than 170 pounds, a placard stating the lesser weight must be permanently attached to the seat structure. (c) Fuel, oil, and coolant filler openings. The following apply: (1) Fuel filter openings must be marked at or near the filler cover with-- (i) For reciprocating engine-powered airplanes-- (A) The word "Avgas"; and (B) The minimum fuel grade.(ii) For turbine engine-powered airplanes-- (A) The words "Jet Fuel"; and (B) The permissible fuel designations, or references to the Airplane Flight Manual (AFM) for permissible fuel designations.(iii) For pressure fueling systems, the maximum permissible fueling supply pressure and the maximum permissible defueling pressure. (2) Oil filler openings must be marked at or near the filler cover with the word "Oil" and the permissible oil designations, or references to the Airplane Flight Manual (AFM) for permissible oil designations. (3) Coolant filler openings must be marked at or near the filler cover with the word "Coolant". (d) Emergency exit placards. Each placard and operating control for each emergency exit must be red. A placard must be near each emergency exit control and must clearly indicate the location of that exit and its method of operation. (e) The system voltage of each direct current installation must be clearly marked adjacent to its exernal power connection. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; as amended by Amdt. 23-21, 42 FR 15042, Mar. 17, 1977; Amdt. 23-23, 43 FR 50594, Oct. 30, 1978; Amdt. No. 23-45, 58 FR 42166, Aug. 6, 1993] Sec. 23.1559 Operating limitations placard. (a) There must be a placard in clear view of the pilot stating-- (1) That the airplane must be operated in accordance with the Airplane Flight Manual; and (2) The certification category of the airplane to which the placards apply. (b) For airplanes certificated in more than one category, there must be a placard in clear view of the pilot stating that other limitations are contained in the Airplane Flight Manual. (c) There must be a placard in clear view of the pilot that specifies the kind of operations to which the operation of the airplane is limited or from which it is prohibited under Sec. 23.1525. [Amdt. 23-50, 61 FR 5193, Feb. 9, 1996] Sec. 23.1561 Safety equipment. (a) Safety equipment must be plainly marked as to method of operation. (b) Stowage provisions for required safety equipment must be marked for the benefit of occupants. Sec. 23.1563 Airspeed placards. There must be an airspeed placard in clear view of the pilot and as close as practicable to the airspeed indicator. This placard must list-- (a) The operating maneuvering speed, Vo; and (b) The maximum landing gear operating speed VLO. (c) For reciprocating multiengine-powered airplanes of more than 6,000 pounds maximum weight, and turbine engine-powered airplanes, the maximum value of the minimum control speed, VMC (one-engine-inoperative) determined under Sec. 23.149(b). [Amdt. 23-7, 34 FR 13097, Aug. 13, 1969, as amended by Amdt. No. 23-45, 58 FR 42166, Aug. 6, 1993; Amdt. 23-50, 61 FR 5193, Feb. 9, 1996] Sec. 23.1567 Flight maneuver placard. (a) For normal category airplanes, there must be a placard in front of and in clear view of the pilot stating: "No acrobatic maneuvers, including spins, approved." (b) For utility category airplanes, there must be-- (1) A placard in clear view of the pilot stating: "Acrobatic maneuvers are limited to the following ------" (list approved maneuvers and the recommended entry speed for each); and (2) For those airplanes that do not meet the spin requirements for acrobatic category airplanes, an additional placard in clear view of the pilot stating: "Spins Prohibited." (c) For acrobatic category airplanes, there must be a placard in clear view of the pilot listing the approved acrobatic maneuvers and the recommended entry airspeed for each. If inverted flight maneuvers are not approved, the placard must bear a notation to this effect. (d) For acrobatic category airplanes and utility category airplanes approved for spinning, there must be a placard in clear view of the pilot-- (1) Listing the control actions for recovery from spinning maneuvers; and (2) Stating that recovery must be initiated when spiral characteristics appear, or after not more than six turns or not more than any greater number of turns for which the airplane has been certificated. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR
258, Jan. 9, 1965, as amended by Amdt. 23-13, 37 FR 20023,
Sept. 23, 1972; Amdt. 23-21, 43 FR 2319, Jan. 16, 1978;
Amdt. 23-50, 61 FR 5193, Feb. 9, 1996] |
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(a) Furnishing information. An Airplane Flight Manual must be furnished with each airplane, and it must contain the following: (1) Information required by Secs. 23.1583 through 23.1589. (2) Other information that is necessary for safe operation because of design, operating, or handling characteristics. (3) Further information necessary to comply with the relevant operating rules. (b) Approved information. (1) Except as provided in paragraph (b)(2) of this section, each part of the Airplane Flight Manual containing information prescribed in Secs. 23.1583 through 23.1589 must be approved, segregated, identified and clearly distinguished from each unapproved part of that Airplane Flight Manual. (2) The requirements of paragraph (b)(1) of this section do not apply to reciprocating engine-powered airplanes of 6,000 pounds or less maximum weight, if the following is met: (i) Each part of the Airplane Flight Manual containing information prescribed in Sec. 23.1583 must be limited to such information, and must be approved, identified, and clearly distinguished from each other part of the Airplane Flight Manual.(ii) The information prescribed in Secs. 23.1585 through 23.1589 must be determined in accordance with the applicable requirements of this part and presented in its entirety in a manner acceptable to the Administrator. (3) Each page of the Airplane Flight Manual containing information prescribed in this section must be of a type that is not easily erased, disfigured, or misplaced, and is capable of being inserted in a manual provided by the applicant, or in a folder, or in any other permanent binder. (c) The units used in the Airplane Flight Manual must be the same as those marked on the appropriate instruments and placards. (d) All Airplane Flight Manual operational airspeeds, unless otherwise specified, must be presented as indicated airspeeds. (e) Provision must be made for stowing the Airplane Flight Manual in a suitable fixed container which is readily accessible to the pilot. (f) Revisions and amendments. Each Airplane Flight Manual (AFM) must contain a means for recording the incorporation of revisions and amendments. [Amdt. 23-21, 43 FR 2319, Jan. 16, 1978, as amended by Amdt. 23-34, 52 FR 1834, Jan. 15, 1987; Amdt. No. 23-45, 58 FR 42166, Aug. 6, 1993; Amdt. 23-50, 61 FR 5193, Feb. 9, 1996] Sec. 23.1583 Operating limitations. The Airplane Flight Manual must contain operating limitations determined under this part 23, including the following-- (a) Airspeed limitations. The following information must be furnished: (1) Information necessary for the marking of the airspeed limits on the indicator as required in Sec. 23.1545, and the significance of each of those limits and of the color coding used on the indicator. (2) The speeds VMC, VO, VLE, and VLO, if established, and their significance. (3) In addition, for turbine powered commuter category airplanes-- (i) The maximum operating limit speed, VMO/MMO and a statement that this speed must not be deliberately exceeded in any regime of flight (climb, cruise or descent) unless a higher speed is authorized for flight test or pilot training; (ii) If an airspeed limitation is based upon compressibility effects, a statement to this effect and information as to any symptoms, the probable behavior of the airplane, and the recommended recovery procedures; and (iii) The airspeed limits must be shown in terms of VMO/MMO instead of VNO and VNE. (b) Powerplant limitations. The following information must be furnished: (1) Limitations required by Sec. 23.1521. (2) Explanation of the limitations, when appropriate. (3) Information necessary for marking the instruments required by Sec. 23.1549 through Sec. 23.1553. (c) Weight. The airplane flight manual must include-- (1) The maximum weight; and (2) The maximum landing weight, if the design landing weight selected by the applicant is less than the maximum weight. (3) For normal, utility, and acrobatic category reciprocating engine- powered airplanes of more than 6,000 pounds maximum weight and for turbine engine-powered airplanes in the normal, utility, and acrobatic category, performance operating limitations as follows-- (i) The maximum takeoff weight for each airport altitude and ambient temperature within the range selected by the applicant at which the airplane complies with the climb requirements of Sec. 23.63(c)(1).(ii) The maximum landing weight for each airport altitude and ambient temperature within the range selected by the applicant at which the airplane complies with the climb requirements of Sec. 23.63(c)(2). (4) For commuter category airplanes, the maximum takeoff weight for each airport altitude and ambient temperature within the range selected by the applicant at which-- (i) The airplane complies with the climb requirements of Sec. 23.63(d)(1); and (ii) The accelerate-stop distance determined under Sec. 23.55 is equal to the available runway length plus the length of any stopway, if utilized; and either: (iii) The takeoff distance determined under Sec. 23.59(a) is equal to the available runway length; or (iv) At the option of the applicant, the takeoff distance determined under Sec. 23.59(a) is equal to the available runway length plus the length of any clearway and the takeoff run determined under Sec. 23.59(b) is equal to the available runway length. (5) For commuter category airplanes, the maximum landing weight for each airport altitude within the range selected by the applicant at which-- (i) The airplane complies with the climb requirements of Sec. 23.63(d)(2) for ambient temperatures within the range selected by the applicant; and (ii) The landing distance determined under Sec. 23.75 for standard temperatures is equal to the available runway length. (6) The maximum zero wing fuel weight, where relevant, as established in accordance with Sec. 23.343. (d) Center of gravity. The established center of gravity limits. (e) Maneuvers. The following authorized maneuvers, appropriate airspeed limitations, and unauthorized maneuvers, as prescribed in this section. (1) Normal category airplanes. No acrobatic maneuvers, including spins, are authorized. (2) Utility category airplanes. A list of authorized maneuvers demonstrated in the type flight tests, together with recommended entry speeds and any other associated limitations. No other maneuver is authorized. (3) Acrobatic category airplanes. A list of approved flight maneuvers demonstrated in the type flight tests, together with recommended entry speeds and any other associated limitations. (4) Acrobatic category airplanes and utility category airplanes approved for spinning. Spin recovery procedure established to show compliance with Sec. 23.221(c). (5) Commuter category airplanes. Maneuvers are limited to any maneuver incident to normal flying, stalls, (except whip stalls) and steep turns in which the angle of bank is not more than 60 degrees. (f) Maneuver load factor. The positive limit load factors in g's, and, in addition, the negative limit load factor for acrobatic category airplanes. (g) Minimum flight crew. The number and functions of the minimum flight crew determined under Sec. 23.1523. (h) Kinds of operation. A list of the kinds of operation to which the airplane is limited or from which it is prohibited under Sec. 23.1525, and also a list of installed equipment that affects any operating limitation and identification as to the equipment's required operational status for the kinds of operation for which approval has been given.(i)--(j) [Reserved] (i) Maximum operating altitude. The maximum altitude established under Sec. 23.1527. (j) Maximum passenger seating configuration. The maximum passenger seating configuration. (k) Allowable lateral fuel loading. The maximum allowable lateral fuel loading differential, if less than the maximum possible. (l) Baggage and cargo loading. The following information for each baggage and cargo compartment or zone-- (1) The maximum allowable load; and (2) The maximum intensity of loading. (m) Systems. Any limitations on the use of airplane systems and equipment. (n) Ambient temperatures. Where appropriate, maximum and minimum ambient air temperatures for operation. (o) Smoking. Any restrictions on smoking in the airplane. (p) Types of surface. A statement of the types of surface on which operations may be conducted.(See Sec. 23.45(g) and Sec. 23.1587 (a)(4), (c)(2), and (d)(4)). [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR 13097, Aug. 13, 1969; Amdt. 23-10, 36 FR 2864, Feb. 11, 1971; Amdt. 23-21, 43 FR 2320, Jan. 16, 1978; Amdt. 23-23, 43 FR 50594, Oct. 30, 1978; Amdt. 23-34, 52 FR 1834, Jan. 15, 1987; Amdt. No. 23-45, 58 FR 42166, Aug. 6, 1993; Amdt. 23-50, 61 FR 5193, Feb. 9, 1996] Sec. 23.1585 Operating procedures. (a) For all airplanes, information concerning normal, abnormal (if applicable), and emergency procedures and other pertinent information necessary for safe operation and the achievement of the scheduled performance must be furnished, including-- (1) An explanation of significant or unusual flight or ground handling characteristics; (2) The maximum demonstrated values of crosswind for takeoff and landing, and procedures and information pertinent to operations in crosswinds; (3) A recommended speed for flight in rough air. This speed must be chosen to protect against the occurrence, as a result of gusts, of structural damage to the airplane and loss of control (for example, stalling); (4) Procedures for restarting any turbine engine in flight, including the effects of altitude; and (5) Procedures, speeds, and configuration(s) for making a normal approach and landing, in accordance with Secs. 23.73 and 23.75, and a transition to the balked landing condition. (6) For seaplanes and amphibians, water handling procedures and the demonstrated wave height. (b) In addition to paragraph (a) of this section, for all single-engine airplanes, the procedures, speeds, and configuration(s) for a glide following engine failure, in accordance with Sec. 23.71 and the subsequent forced landing, must be furnished. (c) In addition to paragraph (a) of this section, for all multiengine airplanes, the following information must be furnished: (1) Procedures, speeds, and configuration(s) for making an approach and landing with one engine inoperative; (2) Procedures, speeds, and configuration(s) for making a balked landing with one engine inoperative and the conditions under which a balked landing can be performed safely, or a warning against attempting a balked landing; (3) The VSSE determined in Sec. 23.149; and (4) Procedures for restarting any engine in flight including the effects of altitude. (d) In addition to paragraphs (a) and either (b) or (c) of this section, as appropriate, for all normal, utility, and acrobatic category airplanes, the following information must be furnished: (1) Procedures, speeds, and configuration(s) for making a normal takeoff, in accordance with Sec. 23.51 (a) and (b), and Sec. 23.53 (a) and (b), and the subsequent climb, in accordance with Sec. 23.65 and Sec. 23.69(a). (2) Procedures for abandoning a takeoff due to engine failure or other cause. (e) In addition to paragraphs (a), (c), and (d) of this section, for all normal, utility, and acrobatic category multiengine airplanes, the information must include the following: (1) Procedures and speeds for continuing a takeoff following engine failure and the conditions under which takeoff can safely be continued, or a warning against attempting to continue the takeoff. (2) Procedures, speeds, and configurations for continuing a climb following engine failure, after takeoff, in accordance with Sec. 23.67, or enroute, in accordance with Sec. 23.69(b). (f) In addition to paragraphs (a) and (c) of this section, for commuter category airplanes, the information must include the following: (1) Procedures, speeds, and configuration(s) for making a normal takeoff. (2) Procedures and speeds for carrying out an accelerate-stop in accordance with Sec. 23.55. (3) Procedures and speeds for continuing a takeoff following engine failure in accordance with Sec. 23.59(a)(1) and for following the flight path determined under Sec. 23.57 and Sec. 23.61(a). (g) For multiengine airplanes, information identifying each operating condition in which the fuel system independence prescribed in Sec. 23.953 is necessary for safety must be furnished, together with instructions for placing the fuel system in a configuration used to show compliance with that section. (h) For each airplane showing compliance with Sec. 23.1353 (g)(2) or (g)(3), the operating procedures for disconnecting the battery from its charging source must be furnished. (i) Information on the total quantity of usable fuel for each fuel tank, and the effect on the usable fuel quantity, as a result of a failure of any pump, must be furnished. (j) Procedures for the safe operation of the airplane's systems and equipment, both in normal use and in the event of malfunction, must be furnished. [Amdt. 23-50, 61 FR 5194, Feb. 9, 1996] Sec. 23.1587 Performance information. Unless otherwise prescribed, performance information must be provided over the altitude and temperature ranges required by Sec. 23.45(b). (a) For all airplanes, the following information must be furnished-- (1) The stalling speeds VSO and VS1 with the landing gear and wing flaps retracted, determined at maximum weight under Sec. 23.49, and the effect on these stalling speeds of angles of bank up to 60 degrees; (2) The steady rate and gradient of climb with all engines operating, determined under Sec. 23.69(a); (3) The landing distance, determined under Sec. 23.75 for each airport altitude and standard temperature, and the type of surface for which it is valid; (4) The effect on landing distances of operation on other than smooth hard surfaces, when dry, determined under Sec. 23.45(g); and (5) The effect on landing distances of runway slope and 50 percent of the headwind component and 150 percent of the tailwind component. (b) In addition to paragraph (a) of this section, for all normal, utility, and acrobatic category reciprocating engine-powered airplanes of 6,000 pounds or less maximum weight, the steady angle of climb/descent, determined under Sec. 23.77(a), must be furnished. (c) In addition to paragraphs (a) and (b) of this section, if appropriate, for normal, utility, and acrobatic category airplanes, the following information must be furnished-- (1) The takeoff distance, determined under Sec. 23.53 and the type of surface for which it is valid. (2) The effect on takeoff distance of operation on other than smooth hard surfaces, when dry, determined under Sec. 23.45(g); (3) The effect on takeoff distance of runway slope and 50 percent of the headwind component and 150 percent of the tailwind component; (4) For multiengine reciprocating engine-powered airplanes of more than 6,000 pounds maximum weight and multiengine turbine powered airplanes, the one-engine-inoperative takeoff climb/descent gradient, determined under Sec. 23.66; (5) For multiengine airplanes, the enroute rate and gradient of climb/ descent with one engine inoperative, determined under Sec. 23.69(b); and (6) For single-engine airplanes, the glide performance determined under Sec. 23.71. (d) In addition to paragraph (a) of this section, for commuter category airplanes, the following information must be furnished-- (1) The accelerate-stop distance determined under Sec. 23.55; (2) The takeoff distance determined under Sec. 23.59(a); (3) At the option of the applicant, the takeoff run determined under Sec. 23.59(b); (4) The effect on accelerate-stop distance, takeoff distance and, if determined, takeoff run, of operation on other than smooth hard surfaces, when dry, determined under Sec. 23.45(g); (5) The effect on accelerate-stop distance, takeoff distance, and if determined, takeoff run, of runway slope and 50 percent of the headwind component and 150 percent of the tailwind component; (6) The net takeoff flight path determined under Sec. 23.61(b); (7) The enroute gradient of climb/descent with one engine inoperative, determined under Sec. 23.69(b); (8) The effect, on the net takeoff flight path and on the enroute gradient of climb/descent with one engine inoperative, of 50 percent of the headwind component and 150 percent of the tailwind component; (9) Overweight landing performance information (determined by extrapolation and computed for the range of weights between the maximum landing and maximum takeoff weights) as follows-- (i) The maximum weight for each airport altitude and ambient temperature at which the airplane complies with the climb requirements of Sec. 23.63(d)(2); and (ii) The landing distance determined under Sec. 23.75 for each airport altitude and standard temperature. (10) The relationship between IAS and CAS determined in accordance with Sec. 23.1323 (b) and (c). (11) The altimeter system calibration required by Sec. 23.1325(e). [Amdt. 23-50, 61 FR 5194, Feb. 9, 1996] Sec. 23.1589 Loading information. The following loading information must be furnished: (a) The weight and location of each item of equipment that can be easily removed, relocated, or replaced and that is installed when the airplane was weighed under the requirement of Sec. 23.25. (b) Appropriate loading instructions for each possible loading condition between the maximum and minimum weights established under Sec. 23.25, to facilitate the center of gravity remaining within the limits established under Sec. 23.23. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR
258, Jan. 9, 1965, as amended by Amdt. No. 23-45, 58 FR
42167, Aug. 6, 1993; Amdt. 23-50, 61 FR 5195, Feb. 9,
1996] |
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(a) The design load criteria in this appendix are an approved equivalent of those in Secs. 23.321 through 23.459 of this subchapter for an airplane having a maximum weight of 6,000 pounds or less and the following configuration: (1) A single engine excluding turbine powerplants; (2) A main wing located closer to the airplane's center of gravity than to the aft, fuselage-mounted, empennage; (3) A main wing that contains a quarter-chord sweep angle of not more than 15 degrees fore or aft; (4) A main wing that is equipped with trailing-edge controls (ailerons or flaps, or both); (5) A main wing aspect ratio not greater than 7; (6) A horizontal tail aspect ratio not greater than 4; (7) A horizontal tail volume coefficient not less than 0.34; (8) A vertical tail aspect ratio not greater than 2; (9) A vertical tail platform area not greater than 10 percent of the wing platform area; and (10) Symmetrical airfoils must be used in both the horizontal and vertical tail designs. (b) Appendix A criteria may not be used on any airplane configuration that contains any of the following design features: (1) Canard, tandem-wing, close-coupled, or tailless arrangements of the lifting surfaces; (2) Biplane or multiplane wing arrangements; (3) T-tail, V-tail, or cruciform-tail (+) arrangements; (4) Highly-swept wing platform (more than 15-degrees of sweep at the quarter-chord), delta planforms, or slatted lifting surfaces; or (5) Winglets or other wing tip devices, or outboard fins. A23.3 Special symbols. n1 Airplane Positive Maneuvering Limit Load Factor. n2 Airplane Negative Maneuvering Limit Load Factor. n3 Airplane Positive Gust Limit Load Factor at VC. n4 Airplane Negative Gust Limit Load Factor at VC. nflap Airplane Positive Limit Load Factor With Flaps Fully Extended at VF. * V Minimum Design Flap Speed F min 11.0 <radical> n1 W/S [kts] * V Mimimum Design Manuevering A min Speed 15.0 <radical> n1 W/S [kts] * V Minimum Design Cruising Speed C min 17.0 <radical> n1 W/S [kts] * V Minimum Design Dive Speed D min 24.0 <radical> n1 W/S [kts] A23.5 Certification in more than one category. The criteria in this appendix may be used for certification in the normal, utility, and acrobatic categories, or in any combination of these categories. If certification in more than one category is desired, the design category weights must be selected to make the term " n1W " constant for all categories or greater for one desired category than for others. The wings and control surfaces (including wing flaps and tabs) need only be investigated for the maximum value of " n1W ", or for the category corresponding to the maximum design weight, where " n1W " is constant. If the acrobatic category is selected, a special unsymmetrical flight load investigation in accordance with paragraphs A23.9(c)(2) and A23.11(c)(2) of this appendix must be completed. The wing, wing carrythrough, and the horizontal tail structures must be checked for this condition. The basic fuselage structure need only be investigated for the highest load factor design category selected. The local supporting structure for dead weight items need only be designed for the highest load factor imposed when the particular items are installed in the airplane. The engine mount, however, must be designed for a higher side load factor, if certification in the acrobatic category is desired, than that required for certification in the normal and utility categories. When designing for landing loads, the landing gear and the airplane as a whole need only be investigated for the category corresponding to the maximum design weight. These simplifications apply to single-engine aircraft of conventional types for which experience is available, and the Administrator may require additional investigations for aircraft with unusual design features. A23.7 Flight loads. (a) Each flight load may be considered independent of altitude and, except for the local supporting structure for dead weight items, only the maximum design weight conditions must be investigated. (b) Table 1 and figures 3 and 4 of this Appendix must be used to determine values of n1, n2, n3, and n4, corresponding to the maximum design weights in the desired categories. (c) Figures 1 and 2 of this Appendix must be used to determine values of n3 and n4 corresponding to the minimum flying weights in the desired categories, and, if these load factors are greater than the load factors at the design weight, the supporting structure for dead weight items must be substantiated for the resulting higher load factors. (d) Each specified wing and tail loading is independent of the center of gravity range. The applicant, however, must select a c.g. range, and the basic fuselage structure must be investigated for the most adverse dead weight loading conditions for the c.g. range selected. (e) The following loads and loading conditions are the minimums for which strength must be provided in the structure: (1) Airplane equilibrium. The aerodynamic wing loads may be considered to act normal to the relative wind, and to have a magnitude of 1.05 times the airplane normal loads (as determined from paragraphs A23.9 (b) and (c) of this appendix) for the positive flight conditions and a magnitude equal to the airplane normal loads for the negative conditions. Each chordwise and normal component of this wing load must be considered. (2) Minimum design airspeeds. The minimum design airspeeds may be chosen by the applicant except that they may not be less than the minimum speeds found by using figure 3 of this Appendix. In addition, VCmin need not exceed values of 0.9 VH actually obtained at sea level for the lowest design weight category for which certification is desired. In computing these minimum design airspeeds, n1 may not be less than 3.8. (3) Flight load factor. The limit flight load factors specified in Table 1 of this Appendix 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 flight load factor is an aerodynamic force acting upward, with respect to the airplane. A23.9 Flight conditions. (a) General. Each design condition in paragraphs (b) and (c) of this section must be used to assure sufficient strength for each condition of speed and load factor on or within the boundary of a V-n diagram for the airplane similar to the diagram in figure 4 of this Appendix. This diagram must also be used to determine the airplane structural operating limitations as specified in Secs. 23.1501(c) through 23.1513 and Sec. 23.1519. (b) Symmetrical flight conditions. The airplane must be designed for symmetrical flight conditions as follows: (1) The airplane must be designed for at least the four basic flight conditions, "A", "D", "E", and "G" as noted on the flight envelope of figure 4 of this Appendix. In addition, the following requirements apply: (i) The design limit flight load factors corresponding to conditions "D" and "E" of figure 4 must be at least as great as those specified in Table 1 and figure 4 of this Appendix, and the design speed for these conditions must be at least equal to the value of VD found from figure 3 of this Appendix.(ii) For conditions "A" and "G" of figure 4, the load factors must correspond to those specified in Table 1 of this Appendix, and the design speeds must be computed using these load factors with the maximum static lift coefficient CNA determined by the applicant. However, in the absence of more precise computations, these latter conditions may be based on a value of CNA +/-1.35 and the design speed for condition "A" may be less than VAmin.(iii) Conditions "C" and "F" of figure 4 need only be investigated when n3W/S or n4W/S are greater than n1W/S or n2W/S of this Appendix, respectively. (2) If flaps or other high lift devices intended for use at the relatively low airspeed of approach, landing, and takeoff, are installed, the airplane must be designed for the two flight conditions corresponding to the values of limit flap-down factors specified in Table 1 of this Appendix with the flaps fully extended at not less than the design flap speed VFmin from figure 3 of this Appendix. (c) Unsymmetrical flight conditions. Each affected structure must be designed for unsymmetrical loadings as follows: (1) The aft fuselage-to-wing attachment must be designed for the critical vertical surface load determined in accordance with paragraph SA23.11(c) (1) and (2) of this Appendix. (2) The wing and wing carry-through structures must be designed for 100 percent of condition "A" loading on one side of the plane of symmetry and 70 percent on the opposite side for certification in the normal and utility categories, or 60 percent on the opposite side for certification in the acrobatic category. (3) The wing and wing carry-through structures must be designed for the loads resulting from a combination of 75 percent of the positive maneuvering wing loading on both sides of the plane of symmetry and the maximum wing torsion resulting from aileron displacement. The effect of aileron displacement on wing torsion at VC or VA using the basic airfoil moment coefficient modified over the aileron portion of the span, must be computed as follows: (i) CmCm +0.01<delta><mu> (up aileron side) wing basic airfoil.(ii) CmCm -0.01<delta><mu>(down aileron side) wing basic airfoil, where <delta><mu> is the up aileron deflection and <delta>d is the down aileron deflection. (4) <Delta> critical, which is the sum of <delta><mu>+<delta>d must be computed as follows: (i) Compute <Delta><alpha> and <Delta>b from the formulas: VA D a -- x D p and VC VA Db 0.5 -- x Dp VD Where Dp the maximum total deflection (sum of both aileron deflections) at VA with VA, VC, and VD described in subparagraph (2) of Sec. 23.7(e) of this Appendix. (ii) Compute K from the formula: (Cm-0.01db) VD2 K ------------------ (Cm-0.01da) VC2 where <delta><alpha> is the down aileron deflection corresponding to <Delta><alpha>, and <delta>b is the down aileron deflection corresponding to <Delta>b as computed in step (i).(iii) If K is less than 1.0, <Delta><alpha> is <Delta> critical and must be used to determine <delta>u and <delta>d. In this case, VC is the critical speed which must be used in computing the wing torsion loads over the aileron span.(iv) If K is equal to or greater than 1.0, <Delta>b is <Delta> critical and must be used to determine <delta>u and <delta>d. In this case, Vd is the critical speed which must be used in computing the wing torsion loads over the aileron span. (d) Supplementary conditions; rear lift truss; engine torque; side load on engine mount. Each of the following supplementary conditions must be investigated: (1) In designing the rear lift truss, the special condition specified in Sec. 23.369 may be investigated instead of condition "G" of figure 4 of this Appendix. If this is done, and if certification in more than one category is desired, the value of W/S used in the formula appearing in Sec. 23.369 must be that for the category corresponding to the maximum gross weight. (2) Each engine mount and its supporting structures must be designed for the maximum limit torque corresponding to METO power and propeller speed acting simultaneously with the limit loads resulting from the maximum positive maneuvering flight load factor n1. The limit torque must be obtained by multiplying the mean torque by a factor of 1.33 for engines with five or more cylinders. For 4, 3, and 2 cylinder engines, the factor must be 2, 3, and 4, respectively. (3) Each engine mount and its supporting structure must be designed for the loads resulting from a lateral limit load factor of not less than 1.47 for the normal and utility categories, or 2.0 for the acrobatic category. A23.11 Control surface loads. (a) General. Each control surface load must be determined using the criteria of paragraph (b) of this section and must lie within the simplified loadings of paragraph (c) of this section. (b) Limit pilot forces. In each control surface loading condition described in paragraphs (c) through (e) of this section, the airloads on the movable surfaces and the corresponding deflections need not exceed those which could be obtained in flight by employing the maximum limit pilot forces specified in the table in Sec. 23.397(b). If the surface loads are limited by these maximum limit pilot forces, the tabs must either be considered to be deflected to their maximum travel in the direction which would assist the pilot or the deflection must correspond to the maximum degree of "out of trim" expected at the speed for the condition under consideration. The tab load, however, need not exceed the value specified in Table 2 of this Appendix. (c) Surface loading conditions. Each surface loading condition must be investigated as follows: (1) Simplified limit surface loadings for the horizontal tail, vertical tail, aileron, wing flaps, and trim tabs are specified in figures 5 and 6 of this appendix.(i) The distribution of load along the span of the surface, irrespective of the chordwise load distribution, must be assumed proportional to the total chord, except on horn balance surfaces.(ii) The load on the stabilizer and elevator, and the load on fin and rudder, must be distributed chordwise as shown in figure 7 of this appendix.(iii) In order to ensure adequate torsional strength and to account for maneuvers and gusts, the most severe loads must be considered in association with every center of pressure position between the leading edge and the half chord of the mean chord of the surface (stabilizer and elevator, or fin and rudder).(iv) To ensure adequate strength under high leading edge loads, the most severe stabilizer and fin loads must be further considered as being increased by 50 percent over the leading 10 percent of the chord with the loads aft of this appropriately decreased to retain the same total load.(v) The most severe elevator and rudder loads should be further considered as being distributed parabolically from three times the mean loading of the surface (stabilizer and elevator, or fin and rudder) at the leading edge of the elevator and rudder, respectively, to zero at the trailing edge according to the equation: [ ...Illustration appears here... ] BILLING CODE 4910-13-M [ ...Illustration appears here... ] BILLING CODE 4910-13-C Where-- P(x)local pressure at the chordwise stations x, cchord length of the tail surface, cfchord length of the elevator and rudder respectively, and waverage surface loading as specified in Figure A5. (vi) The chordwise loading distribution for ailerons, wing flaps, and trim tabs are specified in Table 2 of this appendix. (2) If certification in the acrobatic category is desired, the horizontal tail must be investigated for an unsymmetrical load of 100 percent w on one side of the airplane centerline and 50 percent on the other side of the airplane centerline. (d) Outboard fins. Outboard fins must meet the requirements of Sec. 23.445. (e) Special devices. Special devices must meet the requirements of Sec. 23.459. A23.13 Control system loads. (a) Primary flight controls and systems. Each primary flight control and system must be designed as follows: (1) The flight control system and its supporting structure must be designed for loads corresponding to 125 percent of the computed hinge moments of the movable control surface in the conditions prescribed in A23.11 of this Appendix. In addition-- (i) The system limit loads need not exceed those that could be produced by the pilot and automatic devices operating the controls; and (ii) The design must provide a rugged system for service use, including jamming, ground gusts, taxiing downwind, control inertia, and friction. (2) Acceptable maximum and minimum limit pilot forces for elevator, aileron, and rudder controls are shown in the table in Sec. 23.397(b). These pilots loads must be assumed to act at the appropriate control grips or pads as they would under flight conditions, and to be reacted at the attachments of the control system to the control surface horn. (b) Dual controls. If there are dual controls, the systems must be designed for pilots operating in opposition, using individual pilot loads equal to 75 percent of those obtained in accordance with paragraph (a) of this section, except that individual pilot loads may not be less than the minimum limit pilot forces shown in the table in Sec. 23.397(b). (c) Ground gust conditions. Ground gust conditions must meet the requirements of Sec. 23.415. (d) Secondary controls and systems. Secondary controls and systems must meet the requirements of Sec. 23.405. Table 1--Limit Flight Load Factors [Limit flight load factors] Flight load Normal Utility Acrobatic factors category category category Flaps up: n1 3.8 4.4 6.0 n2 -0.5 n1 n3 (/1/) n4 (/2/) Flaps down: n flap 0.5 n1 n flap /3/ Zero /1/ Find n3 from Fig. 1 /2/ Find n4 from Fig. 2 /3/ Vertical wing load may be assumed equal to zero and only the flap part of the wing need be checked for this condition. [ ...Illustration appears here... ] [ ...Illustration appears here... ] [ ...Illustration appears here... ] Figure A7.--Chordwise Load Distribution for Stabilizer and Elevator or Fin and Rudder [ ...Illustration appears here... ] BILLING CODE 4910-13-C [ ...Illustration appears here... ] where: waverage surface loading (as specified in figure A.5) Eratio of elevator (or rudder) chord to total stabilizer and elevator (or fin and rudder) chord. d'ratio of distance of center of pressure of a unit spanwise length of combined stabilizer and elevator (or fin and rudder) measured from stabilizer (or fin) leading edge to the local chord. Sign convention is positive when center of pressure is behind leading edge. clocal chord. Note: Positive values of w, P1 and P2 are all measured in the same direction. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR 13097, Aug. 13, 1969; 34 FR 14727, Sept. 24, 1969; Amdt. 23-16, 40 FR 2577, Jan. 14, 1975; Amdt. 23-28, 47 FR 13315, Mar. 29, 1982; Amdt. 23-48, 61 FR 5149, Feb. 9, 1996$ |
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Tail wheel type Tail-down Condition Level landing landing Reference section 23.479(a)(1) 23.481(a)(1) Vertical component at c. g nW nW Fore and aft component at c. g KnW 0 Lateral component in either direction at c. g 0 0 Shock absorber extension (hydraulic shock absorber) Note (2) Note (2) Shock absorber deflection (rubber or spring shock absorber), percent 100 100 Tire deflection Static Static Main wheel loads (both wheels) (Vr) (n-L)W (n-L)W b/d Main wheel loads (both wheels) (Dr) KnW 0 Tail (nose) wheel loads (Vf) 0 (n-L)W a/d Tail (nose) wheel loads (Df) 0 0 Notes (1), (3), and (4) (4) [ ...Table continues... ] Nose wheel type Level landing Level landing with nose wheel with inclined just clear of Condition reactions ground Tail-down landing Reference section 23.479(a)(2)(i) 23.479(a)(2)(ii) 23.481(a)(2) and (b). Vertical component nW nW nW. at c. g Fore and aft KnW KnW 0. component at c. g Lateral component 0 0 0. in either direction at c. g Shock absorber Note (2) Note (2) Note (2). extension (hydraulic shock absorber) Shock absorber 100 100 100. deflection (rubber or spring shock absorber), percent Tire deflection Static Static Static. Main wheel loads (n-L)W a'/d' (n-L)W (n-L)W.(both wheels) (Vr) Main wheel loads KnW a'/d' KnW 0.(both wheels) (Dr) Tail (nose) wheel (n-L)W b'/d' 0 0. loads (Vf) Tail (nose) wheel KnW b'/d' 0 0. loads (Df) Notes (1) (1), (3), and (4) (3) and (4). Note (1). K may be determined as follows: K0.25 for W3,000 pounds or less; K0.33 for W6,000 pounds or greater, with linear variation of K between these weights. Note (2). For the purpose of design, the maximum load factor is assumed to occur throughout the shock absorber stroke from 25 percent deflection to 100 percent deflection unless otherwise shown and the load factor must be used with whatever shock absorber extension is most critical for each element of the landing gear. Note (3). Unbalanced moments must be balanced by a rational or conservative method. Note (4). L is defined in Sec. 23.735(b). Note (5). n is the limit inertia load factor, at the c.g. of the airplane, selected under Sec. 23.473 (d), (f), and (g). Basic Landing Conditions [ ...Illustration appears here... ] Level Landing [ ...Illustration appears here... ] Tail Down Landing [ ...Illustration appears here... ] Level Landing with Inclined Reactions [ ...Illustration appears here... ] Level Landing with Nose Wheel Just Clear of Ground [ ...Illustration appears here... ] Tail Down Landing [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as
amended by Amdt. 23-7, 34 FR 13099, Aug. 13, 1969] |
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(a) The following method for determining wheel spin-up loads for landing conditions is based on NACA T.N. 863. However, the drag component used for design may not be less than the drag load prescribed in Sec. 23.479(b). FHmax 1/re <radical> 2Iw(VH--Vc)nFVmax/tS where-- FHmaxmaximum rearward horizontal force acting on the wheel (in pounds); reÔfective rolling radius of wheel under impact based on recommended operating tire pressure (which may be assumed to be equal to the rolling radius under a static load of njWe) in feet; Iwrotational mass moment of inertia of rolling assembly (in slug feet); VHlinear velocity of airplane parallel to ground at instant of contact (assumed to be 1.2 VS0, in feet per second); Vcperipheral speed of tire, if prerotation is used (in feet per second) (there must be a positive means of pre-rotation before pre-rotation may be considered); nequals effective coefficient of friction (0.80 may be used); FVmaxmaximum vertical force on wheel (pounds) njWe, where We and nj are defined in Sec. 23.725; tstime interval between ground contact and attainment of maximum vertical force on wheel (seconds).(However, if the value of FVmax, from the above equation exceeds 0.8 FVmax, the latter value must be used for FHmax.) (b) The equation assumes a linear variation of load factor with time until the peak load is reached and under this assumption, the equation determines the drag force at the time that the wheel peripheral velocity at radius re equals the airplane velocity. Most shock absorbers do not exactly follow a linear variation of load factor with time. Therefore, rational or conservative allowances must be made to compensate for these variations. On most landing gears, the time for wheel spin-up will be less than the time required to develop maximum vertical load factor for the specified rate of descent and forward velocity. For exceptionally large wheels, a wheel peripheral velocity equal to the ground speed may not have been attained at the time of maximum vertical gear load. However, as stated above, the drag spin-up load need not exceed 0.8 of the maximum vertical loads. (c) Dynamic spring-back of the landing gear and adjacent structure at the instant just after the wheels come up to speed may result in dynamic forward acting loads of considerable magnitude. This effect must be determined, in the level landing condition, by assuming that the wheel spin-up loads calculated by the methods of this appendix are reversed. Dynamic spring-back is likely to become critical for landing gear units having wheels of large mass or high landing speeds. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. No. 23-45, 58 FR 42167, Aug. 6, 1993] |
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(a) Conditioning. Specimens must be conditioned to 70 degrees F, plus or minus 5 degrees, and at 50 percent plus or minus 5 percent relative humidity until moisture equilibrium is reached or for 24 hours. Only one specimen at a time may be removed from the conditioning environment immediately before subjecting it to the flame. (b) Specimen configuration. Except as provided for materials used in electrical wire and cable insulation and in small parts, materials must be tested either as a section cut from a fabricated part as installed in the airplane or as a specimen simulating a cut section, such as: a specimen cut from a flat sheet of the material or a model of the fabricated part. The specimen may be cut from any location in a fabricated part; however, fabricated units, such as sandwich panels, may not be separated for a test. The specimen thickness must be no thicker than the minimum thickness to be qualified for use in the airplane, except that: (1) Thick foam parts, such as seat cushions, must be tested in 1/2 inch thickness; (2) when showing compliance with Sec. 23.853(d)(3)(v) for materials used in small parts that must be tested, the materials must be tested in no more than 1/8 inch thickness; (3) when showing compliance with Sec. 23.1359(c) for materials used in electrical wire and cable insulation, the wire and cable specimens must be the same size as used in the airplane. In the case of fabrics, both the warp and fill direction of the weave must be tested to determine the most critical flammability conditions. When performing the tests prescribed in paragraphs (d) and (e) of this appendix, the specimen must be mounted in a metal frame so that (1) in the vertical tests of paragraph (d) of this appendix, the two long edges and the upper edge are held securely; (2) in the horizontal test of paragraph (e) of this appendix, the two long edges and the edge away from the flame are held securely; (3) the exposed area of the specimen is at least 2 inches wide and 12 inches long, unless the actual size used in the airplane is smaller; and (4) the edge to which the burner flame is applied must not consist of the finished or protected edge of the specimen but must be representative of the actual cross section of the material or part installed in the airplane. When performing the test prescribed in paragraph (f) of this appendix, the specimen must be mounted in metal frame so that all four edges are held securely and the exposed area of the specimen is at least 8 inches by 8 inches. (c) Apparatus. Except as provided in paragraph (g) of this Appendix, tests must be conducted in a draft-free cabinet in accordance with Federal Test Method Standard 191 Method 5903 (revised Method 5902) which is available from the General Services Administration, Business Service Center, Region 3, Seventh and D Streets SW., Washington, D.C. 20407, or with some other approved equivalent method. Specimens which are too large for the cabinet must be tested in similar draft-free conditions. (d) Vertical test. A minimum of three specimens must be tested and the results averaged. For fabrics, the direction of weave corresponding to the most critical flammability conditions must be parallel to the longest dimension. Each specimen must be supported vertically. The specimen must be exposed to a Bunsen or Tirrill burner with a nominal 3/8 -inch I.D. tube adjusted to give a flame of 1 1/2 inches in height. The minimum flame temperature measured by a calibrated thermocouple pryometer in the center of the flame must be 1550 deg. F. The lower edge of the specimen must be three- fourths inch above the top edge of the burner. The flame must be applied to the center line of the lower edge of the specimen. For materials covered by Secs. 23.853(d)(3)(i) and 23.853(f), the flame must be applied for 60 seconds and then removed. For materials covered by Sec. 23.853(d)(3)(ii), the flame must be applied for 12 seconds and then removed. Flame time, burn length, and flaming time of drippings, if any, must be recorded. The burn length determined in accordance with paragraph (h) of this Appendix must be measured to the nearest one-tenth inch. (e) Horizontal test. A minimum of three specimens must be tested and the results averaged. Each specimen must be supported horizontally. The exposed surface when installed in the airplane must be face down for the test. The specimen must be exposed to a Bunsen burner or Tirrill burner with a nominal 3/8 -inch I.D. tube adjusted to give a flame of 1 1/2 inches in height. The minimum flame temperature measured by a calibrated thermocouple pyrometer in the center of the flame must be 1550 deg. F. The specimen must be positioned so that the edge being tested is three-fourths of an inch above the top of, and on the center line of, the burner. The flame must be applied for 15 seconds and then removed. A minimum of 10 inches of the specimen must be used for timing purposes, approximately 1 1/2 inches must burn before the burning front reaches the timing zone, and the average burn rate must be recorded. (f) Forty-five degree test. A minimum of three specimens must be tested and the results averaged. The specimens must be supported at an angle of 45 degrees to a horizontal surface. The exposed surface when installed in the aircraft must be face down for the test. The specimens must be exposed to a Bunsen or Tirrill burner with a nominal 3/8 inch I.D. tube adjusted to give a flame of 1 1/2 inches in height. The minimum flame temperature measured by a calibrated thermocouple pyrometer in the center of the flame must be 1550 deg.F. Suitable precautions must be taken to avoid drafts. The flame must be applied for 30 seconds with one-third contacting the material at the center of the specimen and then removed. Flame time, glow time, and whether the flame penetrates (passes through) the specimen must be recorded. (g) Sixty-degree test. A minimum of three specimens of each wire specification (make and size) must be tested. The specimen of wire or cable (including insulation) must be placed at an angle of 60 degrees with the horizontal in the cabinet specified in paragraph (c) of this appendix, with the cabinet door open during the test or placed within a chamber approximately 2 feet high x 1 foot x 1 foot, open at the top and at one vertical side (front), that allows sufficient flow of air for complete combustion but is free from drafts. The specimen must be parallel to and approximately 6 inches from the front of the chamber. The lower end of the specimen must be held rigidly clamped. The upper end of the specimen must pass over a pulley or rod and must have an appropriate weight attached to it so that the specimen is held tautly throughout the flammability test. The test specimen span between lower clamp and upper pulley or rod must be 24 inches and must be marked 8 inches from the lower end to indicate the central point for flame application. A flame from a Bunsen or Tirrill burner must be applied for 30 seconds at the test mark. The burner must be mounted underneath the test mark on the specimen, perpendicular to the specimen and at an angle of 30 degrees to the vertical plane of the specimen. The burner must have a nominal bore of three-eighths inch, and must be adjusted to provide a three-inch-high flame with an inner cone approximately one-third of the flame height. The minimum temperature of the hottest portion of the flame, as measured with a calibrated thermocouple pyrometer, may not be less than 1,750 deg.F. The burner must be positioned so that the hottest portion of the flame is applied to the test mark on the wire. Flame time, burn length, and flaming time drippings, if any, must be recorded. The burn length determined in accordance with paragraph (h) of this appendix must be measured to the nearest one-tenth inch. Breaking of the wire specimen is not considered a failure. (h) Burn length. Burn length is the distance from the original edge to the farthest evidence of damage to the test specimen due to flame impingement, including areas of partial or complete consumption, charring, or embrittlement, but not including areas sooted, stained, warped, or discolored, nor areas where material has shrunk or melted away from the heat source. [Amdt. 23-23, 43 FR 50594, Oct. 30, 1978, as amended
by Amdt. 23-34, 52 FR 1835, Jan. 15, 1987; 52 FR 34745,
Sept. 14, 1987; Amdt. 23-49, 61 FR 5170, Feb. 9,
1996] |
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(a) This appendix specifies requirements for the preparation of Instructions for Continued Airworthiness as required by Sec. 23.1529. (b) The Instructions for Continued Airworthiness for each airplane must include the Instructions for Continued Airworthiness for each engine and propeller (hereinafter designated 'products'), for each appliance required by this chapter, and any required information relating to the interface of those appliances and products with the airplane. If Instructions for Continued Airworthiness are not supplied by the manufacturer of an appliance or product installed in the airplane, the Instructions for Continued Airworthiness for the airplane must include the information essential to the continued airworthiness of the airplane. (c) The applicant must submit to the FAA a program to show how changes to the Instructions for Continued Airworthiness made by the applicant or by the manufacturers of products and appliances installed in the airplane will be distributed. G23.2 Format. (a) The Instructions for Continued Airworthiness must be in the form of a manual or manuals as appropriate for the quantity of data to be provided. (b) The format of the manual or manuals must provide for a practical arrangement. G23.3 Content. The contents of the manual or manuals must be prepared in the English language. The Instructions for Continued Airworthiness must contain the following manuals or sections, as appropriate, and information: (a) Airplane maintenance manual or section. (1) Introduction information that includes an explanation of the airplane's features and data to the extent necessary for maintenance or preventive maintenance. (2) A description of the airplane and its systems and installations including its engines, propellers, and appliances. (3) Basic control and operation information describing how the airplane components and systems are controlled and how they operate, including any special procedures and limitations that apply. (4) Servicing information that covers details regarding servicing points, capacities of tanks, reservoirs, types of fluids to be used, pressures applicable to the various systems, location of access panels for inspection and servicing, locations of lubrication points, lubricants to be used, equipment required for servicing, tow instructions and limitations, mooring, jacking, and leveling information. (b) Maintenance instructions. (1) Scheduling information for each part of the airplane and its engines, auxiliary power units, propellers, accessories, instruments, and equipment that provides the recommended periods at which they should be cleaned, inspected, adjusted, tested, and lubricated, and the degree of inspection, the applicable wear tolerances, and work recommended at these periods. However, the applicant may refer to an accessory, instrument, or equipment manufacturer as the source of this information if the applicant shows that the item has an exceptionally high degree of complexity requiring specialized maintenance techniques, test equipment, or expertise. The recommended overhaul periods and necessary cross reference to the Airworthiness Limitations section of the manual must also be included. In addition, the applicant must include an inspection program that includes the frequency and extent of the inspections necessary to provide for the continued airworthiness of the airplane. (2) Troubleshooting information describing probable malfunctions, how to recognize those malfunctions, and the remedial action for those malfunctions. (3) Information describing the order and method of removing and replacing products and parts with any necessary precautions to be taken. (4) Other general procedural instructions including procedures for system testing during ground running, symmetry checks, weighing and determining the center of gravity, lifting and shoring, and storage limitations. (c) Diagrams of structural access plates and information needed to gain access for inspections when access plates are not provided. (d) Details for the application of special inspection techniques including radiographic and ultrasonic testing where such processes are specified. (e) Information needed to apply protective treatments to the structure after inspection. (f) All data relative to structural fasteners such as identification, discard recommendations, and torque values. (g) A list of special tools needed. (h) In addition, for commuter category airplanes, the following information must be furnished: (1) Electrical loads applicable to the various systems; (2) Methods of balancing control surfaces; (3) Identification of primary and secondary structures; and (4) Special repair methods applicable to the airplane. G23.4 Airworthiness Limitations section. The Instructions for Continued Airworthiness must contain a section titled Airworthiness Limitations that is segregated and clearly distinguishable from the rest of the document. This section must set forth each mandatory replacement time, structural inspection interval, and related structural inspection procedure required for type certification. If the Instructions for Continued Airworthiness consist of multiple documents, the section required by this paragraph must be included in the principal manual. This section must contain a legible statement in a prominent location that reads: "The Airworthiness Limitations section is FAA approved and specifies maintenance required under Secs. 43.16 and 91.403 of the Federal Aviation Regulations unless an alternative program has been FAA approved." [Amdt. 23-26, 45 FR 60171, Sept. 11, 1980, as amended
by Amdt. 23-34, 52 FR 1835, Jan. 15, 1987; 52 FR 34745,
Sept. 14, 1987; Amdt. 23-37, 54 FR 34329, Aug. 18,
1989] |
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(a) This appendix specifies requirements for installation of an APR engine power control system that automatically advances power or thrust on the operating engine(s) in the event any engine fails during takeoff. (b) With the APR system and associated systems functioning normally, all applicable requirements (except as provided in this appendix) must be met without requiring any action by the crew to increase power or thrust. H23.2, Definitions. (a) Automatic power reserve system means the entire automatic system used only during takeoff, including all devices both mechanical and electrical that sense engine failure, transmit signals, actuate fuel controls or power levers on operating engines, including power sources, to achieve the scheduled power increase and furnish cockpit information on system operation. (b) Selected takeoff power, notwithstanding the definition of "Takeoff Power" in part 1 of the Federal Aviation Regulations, means the power obtained from each initial power setting approved for takeoff. (c) Critical Time Interval, as illustrated in figure H1, means that period starting at V1 minus one second and ending at the intersection of the engine and APR failure flight path line with the minimum performance all engine flight path line. The engine and APR failure flight path line intersects the one-engine-inoperative flight path line at 400 feet above the takeoff surface. The engine and APR failure flight path is based on the airplane's performance and must have a positive gradient of at least 0.5 percent at 400 feet above the takeoff surface. Figure H1--Critical Time Interval Illustration [INSERT: Line graph plotting engine and APR failure flight path against minimum performace all engine flight path] H23.3, Reliability and performance requirements. (a) It must be shown that, during the critical time interval, an APR failure that increases or does not affect power on either engine will not create a hazard to the airplane, or it must be shown that such failures are improbable. (b) It must be shown that, during the critical time interval, there are no failure modes of the APR system that would result in a failure that will decrease the power on either engine or it must be shown that such failures are extremely improbable. (c) It must be shown that, during the critical time interval, there will be no failure of the APR system in combination with an engine failure or it must be shown that such failures are extremely improbable. (d) All applicable performance requirements must be met with an engine failure occurring at the most critical point during takeoff with the APR system functioning normally. H23.4, Power setting. The selected takeoff power set on each engine at the beginning of the takeoff roll may not be less than-- (a) The power necessary to attain, at V1, 90 percent of the maximum takeoff power approved for the airplane for the existing conditions; (b) That required to permit normal operation of all safety-related systems and equipment that are dependent upon engine power or power lever position; and (c) That shown to be free of hazardous engine response characteristics when power is advanced from the selected takeoff power level to the maximum approved takeoff power. H23.5, Powerplant controls--general. (a) In addition to the requirements of Sec. 23.1141, no single failure or malfunction (or probable combination thereof) of the APR, including associated systems, may cause the failure of any powerplant function necessary for safety. (b) The APR must be designed to-- (1) Provide a means to verify to the flight crew before takeoff that the APR is in an operating condition to perform its intended function; (2) Automatically advance power on the operating engines following an engine failure during takeoff to achieve the maximum attainable takeoff power without exceeding engine operating limits; (3) Prevent deactivation of the APR by manual adjustment of the power levers following an engine failure; (4) Provide a means for the flight crew to deactivate the automatic function. This means must be designed to prevent inadvertent deactivation; and (5) Allow normal manual decrease or increase in power up to the maximum takeoff power approved for the airplane under the existing conditions through the use of power levers, as stated in Sec. 23.1141(c), except as provided under paragraph (c) of H23.5 of this appendix. (c) For airplanes equipped with limiters that automatically prevent engine operating limits from being exceeded, other means may be used to increase the maximum level of power controlled by the power levers in the event of an APR failure. The means must be located on or forward of the power levers, must be easily identified and operated under all operating conditions by a single action of any pilot with the hand that is normally used to actuate the power levers, and must meet the requirements of Sec. 23.777 (a), (b), and (c). H23.6, Powerplant instruments. In addition to the requirements of Sec. 23.1305: (a) A means must be provided to indicate when the APR is in the armed or ready condition. (b) If the inherent flight characteristics of the airplane do not provide warning that an engine has failed, a warning system independent of the APR must be provided to give the pilot a clear warning of any engine failure during takeoff. (c) Following an engine failure at V1 or above, there must be means for the crew to readily and quickly verify that the APR has operated satisfactorily. [Amdt. 23-43, 58 FR 18979, Apr. 9, 1993] |
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Figure 1. Pictorial definition of angles, dimensions, and directions on a seaplane [INSERT: Diagrams] Figure 2. Hull station weighing factor [INSERT: Diagrams] Figure 3. Transverse pressure distributions [INSERT: Diagrams] [Amdt. No. 23-45, 58 FR 42167, Aug. 6, 1993; 58 FR 51970, 51971, Oct. 5, 1993] 14 CFR Part 23 * Amendment 23-51 * Feb. 9, 1996 |
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For more information on how ASTECH Engineering may be able to help you, please contact Jeff Wilson at astech@cox.net or call 316-304-6157. © Copyright 1996 ASTECH Engineering. All rights reserved. No part of this document may be reproduced in any form without the expressed written consent of the author. |
