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23.601 General. 23.641 Proof of strength. 23.651 Proof of strength. 23.671 General. 23.721 General. 23.751 Main float buoyancy. Personnel and Cargo Accommodations 23.771 Pilot compartment. 23.841 Pressurized cabins. 23.851 Fire extinguishers. Electrical Bonding and Lighting Protection 23.867 Electrical bonding and protection against lighting and static electricity. 23.871 Leveling means. Authority: 49 U.S.C. 106(g), 40113, 44701-44702,
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The suitability of each questionable design detail and part having an important bearing on safety in operations, must be established by tests. Sec. 23.603 Materials and workmanship. (a) The suitability and durability of materials used for parts, the failure of which could adversely affect safety, must-- (1) Be established by experience or tests; (2) Meet approved specifications that ensure their having the strength and other properties assumed in the design data; and (3) Take into account the effects of environmental conditions, such as temperature and humidity, expected in service. (b) Workmanship must be of a high standard. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-17, 41 FR 55464, Dec. 20, 1976; Amdt. 23-23, 43 FR 50592, Oct. 10, 1978] Sec. 23.605 Fabrication methods. (a) The methods of fabrication used must produce consistently sound structures. If a fabrication process (such as gluing, spot welding, or heat- treating) requires close control to reach this objective, the process must be performed under an approved process specification. (b) Each new aircraft fabrication method must be substantiated by a test program. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23-23, 43 FR 50592, Oct. 10, 1978] Sec. 23.607 Fasteners. (a) Each removable fastener must incorporate two retaining devices if the loss of such fastener would preclude continued safe flight and landing. (b) Fasteners and their locking devices must not be adversely affected by the environmental conditions associated with the particular installation. (c) No self-locking nut may be used on any bolt subject to rotation in operation unless a non-friction locking device is used in addition to the self-locking device. $Amdt. 23-48, 61 FR 5148, Feb. 9, 1996$ Sec. 23.609 Protection of structure. Each part of the structure must-- (a) Be suitably protected against deterioration or loss of strength in service due to any cause, including-- (1) Weathering; (2) Corrosion; and (3) Abrasion; and (b) Have adequate provisions for ventilation and drainage. Sec. 23.611 Accessibility provisions. For each part that requires maintenance, inspection, or other servicing, appropriate means must be incorporated into the aircraft design to allow such servicing to be accomplished. $Amdt. 23-48, 61 FR 5148, Feb. 9, 1996$ Sec. 23.613 Material strength properties and design values. (a) Material strength properties must be based on enough tests of material meeting specifications to establish design values on a statistical basis. (b) Design values must be chosen to minimize the probability of structural failure due to material variability. Except as provided in paragraph (e) of this section, compliance with this paragraph must be shown by selecting design values that ensure material strength with the following probability: (1) Where applied loads are eventually distributed through a single member within an assembly, the failure of which would result in loss of structural integrity of the component; 99 percent probability with 95 percent confidence. (2) For redundant structure, in which the failure of individual elements would result in applied loads being safely distributed to other load carrying members; 90 percent probability with 95 percent confidence. (c) The effects of temperature on allowable stresses used for design in an essential component or structure must be considered where thermal effects are significant under normal operating conditions. (d) The design of the structure must minimize the probability of catastrophic fatigue failure, particularly at points of stress concentration. (e) Design values greater than the guaranteed minimums required by this section may be used where only guaranteed minimum values are normally allowed if a "premium selection" of the material is made in which a specimen of each individual item is tested before use to determine that the actual strength properties of that particular item will equal or exceed those used in design. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23-23, 43 FR 50592, Oct. 30, 1978; Amdt. No. 23-45, 58 FR 42163, Aug. 6, 1993] Sec. 23.615 [Removed. Amdt. No. 23-45, 58 FR 42164, Aug. 6, 1993] Sec. 23.619 Special factors. The factor of safety prescribed in Sec. 23.303 must be multiplied by the highest pertinent special factors of safety prescribed in Secs. 23.621 through 23.625 for each part of the structure whose strength is-- (a) Uncertain; (b) Likely to deteriorate in service before normal replacement; or (c) Subject to appreciable variability because of uncertainties in manufacturing processes or inspection methods. [Amdt. 23-7, 34 FR 13091, Aug. 13, 1969] Sec. 23.621 Casting factors. (a) General. The factors, tests, and inspections specified in paragraphs (b) through (d) of this section must be applied in addition to those necessary to establish foundry quality control. The inspections must meet approved specifications. Paragraphs (c) and (d) of this section apply to any structural castings except castings that are pressure tested as parts of hydraulic or other fluid systems and do not support structural loads. (b) Bearing stresses and surfaces. The casting factors specified in paragraphs (c) and (d) of this section-- (1) Need not exceed 1.25 with respect to bearing stresses regardless of the method of inspection used; and (2) Need not be used with respect to the bearing surfaces of a part whose bearing factor is larger than the applicable casting factor. (c) Critical castings. For each casting whose failure would preclude continued safe flight and landing of the airplane or result in serious injury to occupants, the following apply: (1) Each critical casting must either-- (i) Have a casting factor of not less than 1.25 and receive 100 percent inspection by visual, radiographic, and either magnetic particle, penetrant or other approved equivalent non-destructive inspection method; or (ii) Have a casting factor of not less than 2.0 and receive 100 percent visual inspection and 100 percent approved non-destructive inspection. When an approved quality control procedure is established and an acceptable statistical analysis supports reduction, non-destructive inspection may be reduced from 100 percent, and applied on a sampling basis. (2) For each critical casting with a casting factor less than 1.50, three sample castings must be static tested and shown to meet-- (i) The strength requirements of Sec. 23.305 at an ultimate load corresponding to a casting factor of 1.25; and (ii) The deformation requirements of Sec. 23.305 at a load of 1.15 times the limit load. (3) Examples of these castings are structural attachment fittings, parts of flight control systems, control surface hinges and balance weight attachments, seat, berth, safety belt, and fuel and oil tank supports and attachments, and cabin pressure valves. (d) Non-critical castings. For each casting other than those specified in paragraph (c) or (e) of this section, the following apply: (1) Except as provided in paragraphs (d) (2) and (3) of this section, the casting factors and corresponding inspections must meet the following table: Casting factor Inspection 2.0 or more 100 percent visual. Less than 2.0 but more than 1.5 100 percent visual, and magnetic particle or penetrant or equivalent nondestructive inspection methods. 1.25 through 1.50 100 percent visual, magnetic particle or penetrant, and radiographic, or approved equivalent nondestructive inspection methods. (2) The percentage of castings inspected by nonvisual methods may be reduced below that specified in subparagraph (d)(1) of this section when an approved quality control procedure is established. (3) For castings procured to a specification that guarantees the mechanical properties of the material in the casting and provides for demonstration of these properties by test of coupons cut from the castings on a sampling basis-- (i) A casting factor of 1.0 may be used; and (ii) The castings must be inspected as provided in paragraph (d)(1) of this section for casting factors of "1.25 through 1.50" and tested under paragraph (c)(2) of this section. (e) Non-structural castings. Castings used for non-structural purposes do not require evaluation, testing or close inspection. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965; Amdt. No. 23-45, 58 FR 42164, Aug. 6, 1993] Sec. 23.623 Bearing factors. (a) Each part that has clearance (free fit), and that is subject to pounding or vibration, must have a bearing factor large enough to provide for the effects of normal relative motion. (b) For control surface hinges and control system joints, compliance with the factors prescribed in Secs. 23.657 and 23.693, respectively, meets paragraph (a) of this section. [Amdt. 23-7, 34 FR 13091, Aug. 13, 1969] Sec. 23.625 Fitting factors. For each fitting (a part or terminal used to join one structural member to another), the following apply: (a) For each fitting whose strength is not proven by limit and ultimate load tests in which actual stress conditions are simulated in the fitting and surrounding structures, a fitting factor of at least 1.15 must be applied to each part of-- (1) The fitting; (2) The means of attachment; and (3) The bearing on the joined members. (b) No fitting factor need be used for joint designs based on comprehensive test data (such as continuous joints in metal plating, welded joints, and scarf joints in wood). (c) For each integral fitting, the part must be treated as a fitting up to the point at which the section properties become typical of the member. (d) For each seat, berth, safety belt, and harness, its attachment to the structure must be shown, by analysis, tests, or both, to be able to withstand the inertia forces prescribed in Sec. 23.561 multiplied by a fitting factor of 1.33. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR 13091, Aug. 13, 1969] Sec. 23.627 Fatigue strength. The structure must be designed, as far as practicable, to avoid points of stress concentration where variable stresses above the fatigue limit are likely to occur in normal service. Sec. 23.629 Flutter. (a) It must be shown by the methods of paragraph (b) and either paragraph (c) or (d) of this section, that the airplane is free from flutter, control reversal, and divergence for any condition of operation within the limit V-n envelope and at all speeds up to the speed specified for the selected method. In addition-- (1) Adequate tolerances must be established for quantities which affect flutter, including speed, damping, mass balance, and control system stiffness; and (2) The natural frequencies of main structural components must be determined by vibration tests or other approved methods. (b) Flight flutter tests must be made to show that the airplane is free from flutter, control reversal and divergence and to show that-- (1) Proper and adequate attempts to induce flutter have been made within the speed range up to VD; (2) The vibratory response of the structure during the test indicates freedom from flutter; (3) A proper margin of damping exists at VD; and (4) There is no large and rapid reduction in damping as VD is approached. (c) Any rational analysis used to predict freedom from flutter, control reversal and divergence must cover all speeds up to 1.2 VD. (d) Compliance with the rigidity and mass balance criteria (pages 4-12), in Airframe and Equipment Engineering Report No. 45 (as corrected) "Simplified Flutter Prevention Criteria" (published by the Federal Aviation Administration) may be accomplished to show that the airplane is free from flutter, control reversal, or divergence if-- (1) VD/MD for the airplane is less than 260 knots (EAS) and less than Mach 0.5, (2) The wing and aileron flutter prevention criteria, as represented by the wing torsional stiffness and aileron balance criteria, are limited in use to airplanes without large mass concentrations (such as engines, floats, or fuel tanks in outer wing panels) along the wing span, and (3) The airplane-- (i) Does not have a T-tail or other unconventional tail configurations; (ii) Does not have unusual mass distributions or other unconventional design features that affect the applicability of the criteria, and (iii) Has fixed-fin and fixed-stabilizer surfaces. (e) For turbopropeller-powered airplanes, the dynamic evaluation must include-- (1) Whirl mode degree of freedom which takes into account the stability of the plane of rotation of the propeller and significant elastic, inertial, and aerodynamic forces, and (2) Propeller, engine, engine mount, and airplane structure stiffness and damping variations appropriate to the particular configuration. (f) Freedom from flutter, control reversal, and divergence up to VD/MD must be shown as follows: (1) For airplanes that meet the criteria of paragraphs (d)(1) through (d)(3) of this section, after the failure, malfunction, or disconnection of any single element in any tab control system. (2) For airplanes other than those described in paragraph (f)(1) of this section, after the failure, malfunction, or disconnection of any single element in the primary flight control system, any tab control system, or any flutter damper. (g) For airplanes showing compliance with the fail-safe criteria of Secs. 23.571 and 23.572, the airplane must be shown by analysis to be free from flutter up to VD/MD after fatigue failure, or obvious partial failure, of a principal structural element. (h) For airplanes showing compliance with the damage tolerance criteria of Sec. 23.573, the airplane must be shown by analysis to be free from flutter up to VD/MD with the extent of damage for which residual strength is demonstrated. (i) For modifications to the type design that could affect the flutter characteristics, compliance with paragraph (a) of this section must be shown, except that analysis based on previously approved data may be used alone to show freedom from flutter, control reversal and divergence, for all speeds up to the speed specified for the selected method. [Amdt. 23-23, 43 FR 50592, Oct. 30, 1978, as amended
by Amdt. 23-31, 49 FR 46867, Nov. 28, 1984; Amdt. No. 23-45,
58 FR 42164, Aug. 6, 1993; 58 FR 51970, Oct. 5, 1993; Amdt.
23-48, 61 FR 5148, Feb. 9, 1996$ |
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Sec. 23.641 Proof of strength. The strength of stressed-skin wings must be proven by
load tests or by combined structural analysis and load
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(a) Limit load tests of control surfaces are required. These tests must include the horn or fitting to which the control system is attached. (b) In structural analyses, rigging loads due to wire bracing must be accounted for in a rational or conservative manner. Sec. 23.655 Installation. (a) Movable surfaces must be installed so that there is no interference between any surfaces, their bracing, or adjacent fixed structure, when one surface is held in its most critical clearance positions and the others are operated through their full movement. (b) If an adjustable stabilizer is used, it must have stops that will limit its range of travel to that allowing safe flight and landing. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. No. 23-45, 58 FR 42164, Aug. 6, 1993] Sec. 23.657 Hinges. (a) Control surface hinges, except ball and roller bearing hinges, must have a factor of safety of not less than 6.67 with respect to the ultimate bearing strength of the softest material used as a bearing. (b) For ball or roller bearing hinges, the approved rating of the bearing may not be exceeded. [Docket No. 4080, 29 FR 17955, Dec. 18. 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23-48, 61 FR 5148, Feb. 9, 1996$ Sec. 23.659 Mass balance. The supporting structure and the attachment of concentrated mass balance weights used on control surfaces must be designed for-- (a) 24 g normal to the plane of the control surface; (b) 12 g fore and aft; and (c) 12 g parallel to the hinge line. |
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(a) Each control must operate easily, smoothly, and positively enough to allow proper performance of its functions. (b) Controls must be arranged and identified to provide for convenience in operation and to prevent the possibility of confusion and subsequent inadvertent operation. Sec. 23.672 Stability augmentation and automatic and power-operated systems. If the functioning of stability augmentation or other automatic or power- operated systems is necessary to show compliance with the flight characteristics requirements of this part, such systems must comply with Sec. 23.671 and the following: (a) A warning, which is clearly distinguishable to the pilot under expected flight conditions without requiring the pilot's attention, must be provided for any failure in the stability augmentation system or in any other automatic or power-operated system that could result in an unsafe condition if the pilot was not aware of the failure. Warning systems must not activate the control system. (b) The design of the stability augmentation system or of any other automatic or power-operated system must permit initial counteraction of failures without requiring exceptional pilot skill or strength, by either the deactivation of the system or a failed portion thereof, or by overriding the failure by movement of the flight controls in the normal sense. (c) It must be shown that, after any single failure of the stability augmentation system or any other automatic or power-operated system-- (1) The airplane is safely controllable when the failure or malfunction occurs at any speed or altitude within the approved operating limitations that is critical for the type of failure being considered; (2) The controllability and maneuverability requirements of this part are met within a practical operational flight envelope (for example, speed, altitude, normal acceleration, and airplane configuration) that is described in the Airplane Flight Manual (AFM); and (3) The trim, stability, and stall characteristics are not impaired below a level needed to permit continued safe flight and landing. [Amdt. No. 23-45, 58 FR 42164, Aug. 6, 1993] Sec. 23.673 Primary flight controls. Primary flight controls are those used by the pilot for the immediate control of pitch, roll, and yaw. [Docket No. 4080, 29 FR 17955, Dec. 18. 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23-48, 61 FR 5148, Feb. 9, 1996$ Sec. 23.675 Stops. (a) Each control system must have stops that positively limit the range of motion of each movable aerodynamic surface controlled by the system. (b) Each stop must be located so that wear, slackness, or takeup adjustments will not adversely affect the control characteristics of the airplane because of a change in the range of surface travel. (c) Each stop must be able to withstand any loads corresponding to the design conditions for the control system. [Amdt. 23-17, 41 FR 55464, Dec. 20, 1976] Sec. 23.677 Trim systems. (a) Proper precautions must be taken to prevent inadvertent, improper, or abrupt trim tab operation. There must be means near the trim control to indicate to the pilot the direction of trim control movement relative to airplane motion. In addition, there must be means to indicate to the pilot the position of the trim device with respect to both the range of adjustment and, in the case of lateral and directional trim, the neutral position. This means must be visible to the pilot and must be located and designed to prevent confusion. The pitch trim indicator must be clearly marked with a position or range within which it has been demonstrated that take-off is safe for all center of gravity positions and each flap position approved for takeoff. (b) Trimming devices must be designed so that, when any one connecting or transmitting element in the primary flight control system fails, adequate control for safe flight and landing is available with-- (1) For single-engine airplanes, the longitudinal trimming devices; or (2) For multiengine airplanes, the longitudinal and directional trimming devices. (c) Tab controls must be irreversible unless the tab is properly balanced and has no unsafe flutter characteristics. Irreversible tab systems must have adequate rigidity and reliability in the portion of the system from the tab to the attachment of the irreversible unit to the airplane structure. (d) It must be demonstrated that the airplane is safely controllable and that the pilot can perform all maneuvers and operations necessary to effect a safe landing following any probable powered trim system runaway that reasonably might be expected in service, allowing for appropriate time delay after pilot recognition of the trim system runaway. The demonstration must be conducted at critical airplane weights and center of gravity positions. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR 13091, Aug. 13, 1969; Amdt. 23-34, 52 FR 1830, Jan. 15, 1987; Amdt. 23-42, 56 FR 353, Jan. 3, 1991; Amdt. 23-49, 61 FR 5165, Feb. 9, 1996$ Sec. 23.679 Control system locks. If there is a device to lock the control system on the ground or water: (a) There must be a means to-- (1) Give unmistakable warning to the pilot when lock is engaged; or (2) Automatically disengage the device when the pilot operates the primary flight controls in a normal manner. (b) The device must be installed to limit the operation of the airplane so that, when the device is engaged, the pilot receives unmistakable warning at the start of the takeoff. (c) The device must have a means to preclude the possibility of it becoming inadvertently engaged in flight. [Amdt. No. 23-45, 58 FR 42164, Aug. 6, 1993] Sec. 23.681 Limit load static tests. (a) Compliance with the limit load requirements of this part must be shown by tests in which-- (1) The direction of the test loads produces the most severe loading in the control system; and (2) Each fitting, pulley, and bracket used in attaching the system to the main structure is included. (b) Compliance must be shown (by analyses or individual load tests) with the special factor requirements for control system joints subject to angular motion. Sec. 23.683 Operation tests. (a) It must be shown by operation tests that, when the controls are operated from the pilot compartment with the system loaded as prescribed in paragraph (b) of this section, the system is free from-- (1) Jamming; (2) Excessive friction; and (3) Excessive deflection. (b) The prescribed test loads are-- (1) For the entire system, loads corresponding to the limit airloads on the appropriate surface, or the limit pilot forces in Sec. 23.397(b), whichever are less; and (2) For secondary controls, loads not less than those corresponding to the maximum pilot effort established under Sec. 23.405. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR 13091, Aug. 13, 1969] Sec. 23.685 Control system details. (a) Each detail of each control system must be designed and installed to prevent jamming, chafing, and interference from cargo, passengers, loose objects, or the freezing of moisture. (b) There must be means in the cockpit to prevent the entry of foreign objects into places where they would jam the system. (c) There must be means to prevent the slapping of cables or tubes against other parts. (d) Each element of the flight control system must have design features, or must be distinctively and permanently marked, to minimize the possibility of incorrect assembly that could result in malfunctioning of the control system. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-17, 41 FR 55464, Dec. 20, 1976] Sec. 23.687 Spring devices. The reliability of any spring device used in the control system must be established by tests simulating service conditions unless failure of the spring will not cause flutter or unsafe flight characteristics. Sec. 23.689 Cable systems. (a) Each cable, cable fitting, turnbuckle, splice, and pulley used must meet approved specifications. In addition-- (1) No cable smaller than 1/8 inch diameter may be used in primary control systems; (2) Each cable system must be designed so that there will be no hazardous change in cable tension throughout the range of travel under operating conditions and temperature variations; and (3) There must be means for visual inspection at each fairlead, pulley, terminal, and turnbuckle. (b) Each kind and size of pulley must correspond to the cable with which it is used. Each pulley must have closely fitted guards to prevent the cables from being misplaced or fouled, even when slack. Each pulley must lie in the plane passing through the cable so that the cable does not rub against the pulley flange. (c) Fairleads must be installed so that they do not cause a change in cable direction of more than three degrees. (d) Clevis pins subject to load or motion and retained only by cotter pins may not be used in the control system. (e) Turnbuckles must be attached to parts having angular motion in a manner that will positively prevent binding throughout the range of travel. (f) Tab control cables are not part of the primary control system and may be less than 1/8 inch diameter in airplanes that are safely controllable with the tabs in the most adverse positions. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR 13091, Aug. 13, 1969] Sec. 23.691 Artificial stall barrier system. If the function of an artificial stall barrier, for example, stick pusher, is used to show compliance with Sec. 23.201(c), the system must comply with the following: (a) With the system adjusted for operation, the plus and minus airspeeds at which downward pitching control will be provided must be established. (b) Considering the plus and minus airspeed tolerances established by paragraph (a) of this section, an airspeed must be selected for the activation of the downward pitching control that provides a safe margin above any airspeed at which any unsatisfactory stall characteristics occur. (c) In addition to the stall warning required Sec. 23.07, a warning that is clearly distinguishable to the pilot under all expected flight conditions without requiring the pilot's attention, must be provided for faults that would prevent the system from providing the required pitching motion. (d) Each system must be designed so that the artificial stall barrier can be quickly and positively disengaged by the pilots to prevent unwanted downward pitching of the airplane by a quick release (emergency) control that meets the requirements of Sec. 23.1329(b). (e) A preflight check of the complete system must be established and the procedure for this check made available in the Airplane Flight Manual (AFM). Preflight checks that are critical to the safety of the airplane must be included in the limitations section of the AFM. (f) For those airplanes whose design includes an autopilot system: (1) A quick release (emergency) control installed in accordance with Sec. 23.1329(b) may be used to meet the requirements of paragraph (d), of this section, and (2) The pitch servo for that system may be used to provide the stall downward pitching motion. (g) In showing compliance with Sec. 23.1309, the system must be evaluated to determine the effect that any announced or unannounced failure may have on the continued safe flight and landing of the airplane or the ability of the crew to cope with any adverse conditions that may result from such failures. This evaluation must consider the hazards that would result from the airplane's flight characteristics if the system was not provided, and the hazard that may result from unwanted downward pitching motion, which could result from a failure at airspeeds above the selected stall speed. $Amdt. 23-49, 61 FR 5165, Feb. 9, 1996$ Sec. 23.693 Joints. Control system joints (in push-pull systems) that are subject to angular motion, except those in ball and roller bearing systems, must have a special factor of safety of not less than 3.33 with respect to the ultimate bearing strength of the softest material used as a bearing. This factor may be reduced to 2.0 for joints in cable control systems. For ball or roller bearings, the approved ratings may not be exceeded. Sec. 23.697 Wing flap controls. (a) Each wing flap control must be designed so that, when the flap has been placed in any position upon which compliance with the performance requirements of this part is based, the flap will not move from that position unless the control is adjusted or is moved by the automatic operation of a flap load limiting device. (b) The rate of movement of the flaps in response to the operation of the pilot's control or automatic device must give satisfactory flight and performance characteristics under steady or changing conditions of airspeed, engine power, and attitude. (c) If compliance with Sec. 23.145(b)(3) necessitates wing flap retraction to positions that are not fully retracted, the wing flap control lever settings corresponding to those positions must be positively located such that a definite change of direction of movement of the lever is necessary to select settings beyond those settings. [Docket No. 4080, 29 FR 17955, Dec. 18. 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23-49, 61 FR 5165, Feb. 9, 1996$ Sec. 23.699 Wing flap position indicator. There must be a wing flap position indicator for-- (a) Flap installations with only the retracted and fully extended position, unless-- (1) A direct operating mechanism provides a sense of "feel" and position (such as when a mechanical linkage is employed); or (2) The flap position is readily determined without seriously detracting from other piloting duties under any flight condition, day or night; and (b) Flap installation with intermediate flap positions if-- (1) Any flap position other than retracted or fully extended is used to show compliance with the performance requirements of this part; and (2) The flap installation does not meet the requirements of paragraph (a)(1) of this section. Sec. 23.701 Flap interconnection. (a) The main wing flaps and related movable surfaces as a system must-- (1) Be synchronized by a mechanical interconnection between the movable flap surfaces that is independent of the flap drive system; or by an approved equivalent means; or (2) Be designed so that the occurrence of any failure of the flap system that would result in an unsafe flight characteristic of the airplane is extremely improbable; or (b) The airplane must be shown to have safe flight characteristics with any combination of extreme positions of individual movable surfaces (mechanically interconnected surfaces are to be considered as a single surface). (c) If an interconnection is used in multiengine airplanes, it must be designed to account for the unsummetrical loads resulting from flight with the engines on one side of the plane of symmetry inoperative and the remaining engines at takeoff power. For single-engine airplanes, and multiengine airplanes with no slipstream effects on the flaps, it may be assumed that 100 percent of the critical air load acts on one side and 70 percent on the other. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-14, 38 FR 31821, Nov. 19, 1973; Amdt. 23-42, 56 FR 353, Jan. 3, 1991; Amdt. 23-42, 56 FR 5455, Feb. 11, 1991; Amdt. 23-49, 61 FR 5165, Feb. 9, 1996$ Sec. 23.703 Takeoff warning system. For commuter category airplanes, unless it can be shown that a lift or longitudinal trim device that affects the takeoff performance of the aircraft would not give an unsafe takeoff configuration when selection out of an approved takeoff position, a takeoff warning system must be installed and meet the following requirements: (a) The system must provide to the pilots an aural warning that is automatically activated during the initial portion of the takeoff role if the airplane is in a configuration that would not allow a safe takeoff. The warning must continue until-- (1) The configuration is changed to allow safe takeoff, or (2) Action is taken by the pilot to abandon the takeoff roll. (b) The means used to activate the system must function properly for all authorized takeoff power settings and procedures and throughout the ranges of takeoff weights, altitudes, and temperatures for which certification is requested. $Amdt. 23-49, 61 FR 5165, Feb. 9, 1996$ |
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For commuter category airplanes that have a passenger seating configuration, excluding pilot seats, of 10 or more, the following general requirements for the landing gear apply: (a) The main landing-gear system must be designed so that if it fails due to overloads during takeoff and landing (assuming the overloads to act in the upward and aft directions), the failure mode is not likely to cause the spillage of enough fuel from any part of the fuel system to consitute a fire hazard. (b) Each airplane must be designed so that, with the airplane under control, it can be landed on a paved runway with any one or more landing-gear legs not extended without sustaining a structural component failure that is likely to cause the spillage of enough fuel to consitute a fire hazard. (c) Compliance with the provisions of this section may be shown by analysis or tests, or both. [Amdt. 23-34, 52 FR 1830, Jan. 15, 1987] Sec. 23.723 Shock absorption tests. (a) It must be shown that the limit load factors selected for design in accordance with Sec. 23.473 for takeoff and landing weights, respectively, will not be exceeded. This must be shown by energy absorption tests except that analysis based on tests conducted on a landing gear system with identical energy absorption characteristics may be used for increases in previously approved takeoff and landing weights. (b) The landing gear may not fail, but may yield, in a test showing its reserve energy absorption capacity, simulating a descent velocity of 1.2 times the limit descent velocity, assuming wing lift equal to the weight of the airplane. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23-23, 43 FR 50593, Oct. 30, 1978; Amdt. 23-49, 61 FR 5165, Feb. 9, 1996$ Sec. 23.725 Limit drop tests. (a) If compliance with Sec. 23.723(a) is shown by free drop tests, these tests must be made on the complete airplane, or on units consisting of wheel, tire, and shock absorber, in their proper relation, from free drop heights not less than those determined by the following formula: h (inches) 3.6 (W/S) 1/2 However, the free drop height may not be less than 9.2 inches and need not be more than 18.7 inches. (b) If the effect of wing lift is provided for in free drop tests, the landing gear must be dropped with an effective weight equal to [h+(1-L)d] We W ------------ (h+d) where-- We the effective weight to be used in the drop test (lbs.); h specified free drop height (inches); d deflection under impact of the tire (at the approved inflation pressure) plus the vertical component of the axle travel relative to the drop mass (inches); WWM for main gear units (lbs), equal to the static weight on that unit with the airplane in the level attitude (with the nose wheel clear in the case of nose wheel type airplanes); WWT for tail gear units (lbs.), equal to the static weight on the tail unit with the airplane in the tail-down attitude; WWN for nose wheel units lbs.), equal to the vertical component of the static reaction that would exist at the nose wheel, assuming that the mass of the airplane acts at the center of gravity and exerts a force of 1.0 g downward and 0.33 g forward; and L the ratio of the assumed wing lift to the airplane weight, but not more than 0.667. (c) The limit inertia load factor must be determined in a rational or conservative manner, during the drop test, using a landing gear unit attitude, and applied drag loads, that represent the landing conditions. (d) The value of d used in the computation of We in paragraph (b) of this section may not exceed the value actually obtained in the drop test. (e) The limit inertia load factor must be determined from the drop test in paragraph (b) of this section according to the following formula: We n nj ---- + L W where-- njthe load factor developed in the drop test (that is, the acceleration (dv/ dt) in g's recorded in the drop test) plus 1.0; and We, W, and L are the same as in the drop test computation. (f) The value of n determined in accordance with paragraph (e) may not be more than the limit inertia load factor used in the landing conditions in Sec. 23.473. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR 13091, Aug. 13, 1969; Amdt. 23-48, 61 FR 5148, Feb. 9, 1996$ Sec. 23.726 Ground load dynamic tests. (a) If compliance with the ground load requirements of Secs. 23.479 through 23.483 is shown dynamically by drop test, one drop test must be conducted that meets Sec. 23.725 except that the drop height must be-- (1) 2.25 times the drop height prescribed in Sec. 23.725(a); or (2) Sufficient to develop 1.5 times the limit load factor. (b) The critical landing condition for each of the design conditions specified in Secs. 23.479 through 23.483 must be used for proof of strength. [Amdt. 23-7, 34 FR 13091, Aug. 13, 1969] Sec. 23.727 Reserve energy absorption drop test. (a) If compliance with the reserve energy absorption requirement in Sec. 23.723(b) is shown by free drop tests, the drop height may not be less than 1.44 times that specified in Sec. 23.725. (b) If the effect of wing lift is provided for, the units must be dropped with an effective mass equal to WeWh/(h+d), when the symbols and other details are the same as in Sec. 23.725. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR 13091, Aug. 13, 1969] Sec. 23.729 Landing gear extension and retraction system. (a) General. For airplanes with retractable landing gear, the following apply: (1) Each landing gear retracting mechanism and its supporting structure must be designed for maximum flight load factors with the gear retracted and must be designed for the combination of friction, inertia, brake torque, and air loads, occurring during retraction at any airspeed up to 1.6 VS1 with flaps retracted, and for any load factor up to those specified in Sec. 23.345 for the flaps-extended condition. (2) The landing gear and retracting mechanism, including the wheel well doors, must withstand flight loads, including loads resulting from all yawing conditions specified in Sec. 23.351, with the landing gear extended at any speed up to at least 1.6 VS1 with the flaps retracted. (b) Landing gear lock. There must be positive means (other than the use of hydraulic pressure) to keep the landing gear extended. (c) Emergency operation. For a landplane having retractable landing gear that cannot be extended manually, there must be means to extend the landing gear in the event of either-- (1) Any reasonably probable failure in the normal landing gear operation system; or (2) Any reasonably probable failure in a power source that would prevent the operation of the normal landing gear operation system. (d) Operation test. The proper functioning of the retracting mechanism must be shown by operation tests. (e) Position indicator. If a retractable landing gear is used, there must be a landing gear position indicator (as well as necessary switches to actuate the indicator) or other means to inform the pilot that each gear is secured in the extended (or retracted) position. If switches are used, they must be located and coupled to the landing gear mechanical system in a manner that prevents an erroneous indication of either "down and locked" if each gear is not in the fully extended position, or "up and locked" if each landing gear is not in the fully retracted position. (f) Landing gear warning. For landplanes, the following aural or equally effective landing gear warning devices must be provided: (1) A device that functions continuously when one or more throttles are closed beyond the power settings normally used for landing approach if the landing gear is not fully extended and locked. A throttle stop may not be used in place of an aural device. If there is a manual shutoff for the warning device prescribed in this paragraph, the warning system must be designed so that when the warning has been suspended after one or more throttles are closed, subsequent retardation of any throttle to, or beyond, the position for normal landing approach will activate the warning device. (2) A device that functions continuously when the wing flaps are extended beyond the maximum approach flap position, using a normal landing procedure, if the landing gear is not fully extended and locked. There may not be a manual shutoff for this warning device. The flap position sensing unit may be installed at any suitable location. The system for this device may use any part of the system (including the aural warning device) for the device required in paragraph (f)(1) of this section. (g) Equipment located in the landing gear bay. If the landing gear bay is used as the location for equipment other than the landing gear, that equipment must be designed and installed to minimize damage from items such as a tire burst, or rocks, water, and slush that may enter the landing gear bay. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR 13091, Aug. 13, 1969; Amdt. 23-21, 43 FR 2318, Jan. 1978; Amdt. 23-26, 45 FR 60171, Sept. 11, 1980; Amdt. No. 23-45, 58 FR 42164, Aug. 6, 1993; Amdt. 23-49, 61 FR 5165, Feb. 9, 1996$ Sec. 23.731 Wheels. (a) The maximum static load rating of each wheel may not be less than the corresponding static ground reaction with-- (1) Design maximum weight; and (2) Critical center of gravity. (b) The maximum limit load rating of each wheel must equal or exceed the maximum radial limit load determined under the applicable ground load requirements of this part. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. No. 23-45, 58 FR 42165, Aug. 6, 1993] Sec. 23.733 Tires. (a) Each landing gear wheel must have a tire whose approved tire ratings (static and dynamic) are not exceeded-- (1) By a load on each main wheel tire) to be compared to the static rating approved for such tires) equal to the corresponding static ground reaction under the design maximum weight and critical center of gravity; and (2) By a load on nose wheel tires (to be compared with the dynamic rating approved for such tires) equal to the reaction obtained at the nose wheel, assuming the mass of the airplane to be concentrated at the most critical center of gravity and exerting a force of 1.0 W downward and 0.31 W forward (where W is the design maximum weight), with the reactions distributed to the nose and main wheels by the principles of statics and with the drag reaction at the ground applied only at wheels with brakes. (b) If specially constructed tires are used, the wheels must be plainly and conspicuously marked to that effect. The markings must include the make, size, number of plies, and identification marking of the proper tire. (c) Each tire installed on a retractable landing gear system must, at the maximum size of the tire type expected in service, have a clearance to surrounding structure and systems that is adequate to prevent contact between the tire and any part of the structure of systems. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR 13092, Aug. 13, 1969; Amdt. 23-17, 41 FR 55464, Dec. 20, 1976; Amdt. No. 23-45, 58 FR 42165, Aug. 6, 1993] Sec. 23.735 Brakes. (a) Brakes must be provided. The landing brake kinetic energy capacity rating of each main wheel brake assembly must not be less than the kinetic energy absorption requirements determined under either of the following methods: (1) The brake kinetic energy absorption requirements must be based on a conservative rational analysis of the sequence of events expected during landing at the design landing weight. (2) Instead of a rational analysis, the kinetic energy absorption requirements for each main wheel brake assembly may be derived from the following formula: KE0.0443 WV**2/N where-- KEKinetic energy per wheel (ft.-lb.); WÞsign landing weight (lb.); VAirplane speed in knots. V must be not less than Vs<radical>, the poweroff stalling speed of the airplane at sea level, at the design landing weight, and in the landing configuration; and NNumber of main wheels with brakes. (b) Brakes must be able to prevent the wheels from rolling on a paved runway with takeoff power on the critical engine, but need not prevent movement of the airplane with wheels locked. (c) During the landing distance determination required by Sec. 23.75, the pressure on the wheel braking system must not exceed the pressure specified by the brake manufacturer. (d) If antiskid devices are installed, the devices and associated systems must be designed so that no single probable malfunction or failure will result in a hazardous loss of braking ability or directional control of the airplane. (e) In addition, for commuter category airplanes, the rejected takeoff brake kinetic energy capacity rating of each main wheel brake assembly must not be less than the kinetic energy absorption requirements determined under either of the following methods-- (1) The brake kinetic energy absorption requirements must be based on a conservative rational analysis of the sequence of events expected during a rejected takeoff at the design takeoff weight. (2) Instead of a rational analysis, the kinetic energy absorption requirements for each main wheel brake assembly may be derived from the following formula-- KE0.0443 WV/2/N where, KEKinetic energy per wheel (ft.-lbs.); WÞsign takeoff weight (lbs.); VGround speed, in knots, associated with the maximum value of V1 selected in accordance with Sec. 23.51(c)(1); NNumber of main wheels with brakes. [Amdt. 23-7, 34 FR 13092, Aug. 13, 1969, as amended by Amdt. 23-24, 44 FR 68742, Nov. 29, 1979; Amdt. 23-42, 56 FR 354, Jan. 3, 1991; Amdt. 23-49, 61 FR 5166, Feb. 9, 1996$ Sec. 23.737 Skis. The maximum limit load rating of each ski must equal or exceed the maximum limit load determined under the applicable ground load requirements of this part. [Amdt. No. 73-45, 58 FR 42165, Aug. 6, 1993] Sec. 23.745 Nose/tail wheel steering. (a) If nose/tail wheel steering is installed, it must be demonstrated that its use does not require exceptional pilot skill during takeoff and landing, in crosswinds, or in the event of an engine failure; or its use must be limited to low speed maneuvering. (b) Movement of the pilot's steering control must not interfere with the retraction or extension of the landing gear. [Amdt. 23-49, 61 FR 5166, Feb. 12, 1996] |
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(a) Each main float must have-- (1) A buoyancy of 80 percent in excess of the buoyancy required by that float to support its portion of the maximum weight of the seaplane or amphibian in fresh water; and (2) Enough watertight compartments to provide reasonable assurance that the seaplane or amphibian will stay afloat without capsizing if any two compartments of any main float are flooded. (b) Each main float must contain at least four watertight compartments approximately equal in volume. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. No. 23-45, 58 FR 42165, Aug. 6, 1993] Sec. 23.753 Main float design. Each seaplane main float must meet the requirements of Sec. 23.521. [Amdt. No. 23-45, 58 FR 42165, Aug. 6, 1993] Sec. 23.755 Hulls. (a) The hull of a hull seaplane or amphibian of 1,500 pounds or more maximum weight must have watertight compartments designed and arranged so that the hull auxiliary floats, and tires (if used), will keep the airplane afloat without capsizing in fresh water when-- (1) For airplanes of 5,000 pounds or more maximum weight, any two adjacent compartments are flooded; and (2) For airplanes of 1,500 pounds up to, but not including, 5,000 pounds maximum weight, any single compartment is flooded. (b) Watertight doors in bulkheads may be used for communication between compartments. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. No. 23-45, 58 FR 42165, Aug. 6, 1993; Amdt. 23-48, 61 FR 5148, Feb. 9, 1996$ Sec. 23.757 Auxiliary floats. Auxiliary floats must be arranged so that, when
completely submerged in fresh water, they provide a righting
moment of at least 1.5 times the upsetting moment caused by
the seaplane or amphibian being tilted. |
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For each pilot compartment-- (a) The compartment and its equipment must allow each pilot to perform his duties without unreasonable concentration or fatigue; (b) Where the flight crew are separated from the passengers by a partition, an opening or openable window or door must be provided to facilitate communication between flight crew and the passengers; and (c) The aerodynamic controls listed in Sec. 23.779, excluding cables and control rods, must be located with respect to the propellers so that no part of the pilot or the controls lies in the region between the plane of rotation of any inboard propeller and the surface generated by a line passing through the center of the propeller hub making an angle of 5 degrees forward or aft of the plane of rotation of the propeller. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-14, 38 FR 31821, Nov. 19, 1973] Sec. 23.773 Pilot compartment view. (a) Each pilot compartment must be-- (1) Arranged with sufficiently extensive, clear and undistorted view to enable the pilot to safely taxi, takeoff, approach, land, and perform any maneuvers within the operating limitations of the airplane. (2) Free from glare and reflections that could interfere with the pilot's vision. Compliance must be shown in all operations for which certification is requested; and (3) Designed so that each pilot is protected from the elements so that moderate rain conditions do not unduly impair the pilot's view of the flight path in normal flight and while landing. (b) Each pilot compartment must have a means to either remove or prevent the formation of fog or frost on an area of the internal portion of the windshield and side windows sufficiently large to provide the view specified in paragraph (a)(1) of this section. Compliance must be shown under all expected external and internal ambient operating conditions, unless it can be shown that the windshield and side windows can be easily cleared by the pilot without interruption of moral pilot duties. [Amdt. 23-45, 58 FR 42165, Aug. 6, 1993] Sec. 23.775 Windshields and windows. (a) The internal panels of windshields and windows must be constructed of a nonsplintering material, such as nonsplintering safety glass. (b) The design of windshields, windows, and canopies in pressurized airplanes must be based on factors peculiar to high altitude operation, including-- (1) The effects of continuous and cyclic pressurization loadings; (2) The inherent characteristics of the material used; and (3) The effects of temperatures and temperature gradients. (c) On pressurized airplanes, if certification for operation up to and including 25,000 feet is requested, an enclosure canopy including a representative part of the installation must be subjected to special tests to account for the combined effects of continuous and cyclic pressurization loadings and flight loads, or compliance with the fail-safe requirements of paragraph (d) of this section must be shown. (d) If certification for operation above 25,000 feet is requested the windshields, window panels, and canopies must be strong enough to withstand the maximum cabin pressure differential loads combined with critical aerodynamic pressure and temperature effects, after failure of any load- carrying element of the windshield, window panel, or canopy. (e) The windshield and side windows forward of the pilot's back when the pilot is seated in the normal flight position must have a luminous transmittance value of not less than 70 percent. (f) Unless operation in known or forecast icing conditions is prohibited by operating limitations, a means must be provided to prevent or to clear accumulations of ice from the windshield so that the pilot has adequate view for taxi, takeoff, approach, landing, and to perform any maneuvers within the operating limitations of the airplane. (g) In the event of any probable single failure, a transparency heating system must be incapable of raising the temperature of any windshield or window to a point where there would be-- (1) Structural failure that adversely affects the integrity of the cabin; or (2) There would be a danger of fire. (h) In addition, for commuter category airplanes, the following applies: (1) Windshield panes directly in front of the pilots in the normal conduct of their duties, and the supporting structures for these panes, must withstand, without penetration, the impact of a two-pound bird when the velocity of the airplane (relative to the bird along the airplane's flight path) is equal to the airplane's maximum approach flap speed. (2) The windshield panels in front of the pilots must be arranged so that, assuming the loss of vision through any one panel, one or more panels remain available for use by a pilot seated at a pilot station to permit continued safe flight and landing. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR 13092, Aug. 13, 1969; Amdt. No. 23-45, 58 FR 42165, Aug. 6, 1993; 58 FR 51970, Oct. 5, 1993; Amdt. 23-49, 61 FR 5165, Feb. 9, 1996] Sec. 23.777 Cockpit controls. (a) Each cockpit control must be located and (except where its function is obvious) identified to provide convenient operation and to prevent confusion and inadvertent operation. (b) The controls must be located and arranged so that the pilot, when seated, has full and unrestricted movement of each control without interference from either his clothing or the cockpit structure. (c) Powerplant controls must be located-- (1) For multiengine airplanes, on the pedestal or overhead at or near the center of the cockpit; (2) For single and tandem seated single-engine airplanes, on the left side console or instrument panel; (3) For other single-engine airplanes at or near the center of the cockpit, on the pedestal, instrument panel, or overhead; and (4) For airplanes, with side-by-side pilot seats and with two sets of powerplant controls, on left and right consoles. (d) The control location order from left to right must be power (thrust) lever, propeller (rpm control), and mixture control (condition lever and fuel cutoff for turbine-powered airplanes). Power (thrust) levers must be at least one inch higher or longer to make them more prominent than propeller (rpm control) or mixture controls. Carburetor heat or alternate air control must be to the left of the throttle or at least eight inches from the mixture control when located other than on a pedestal. Carburetor heat or alternate air control, when located on a pedestal must be aft or below the power (thrust) lever. Supercharger controls must be located below or aft of the propeller controls. Airplanes with tandem seating or single-place airplanes may utilize control locations on the left side of the cabin compartment; however, location order from left to right must be power (thrust) lever, propeller (rpm control) and mixture control. (e) Identical powerplant controls for each engine must be located to prevent confusion as to the engines they control. (1) Conventional multiengine powerplant controls must be located so that the left control(s) operates the left engines(s) and the right control(s) operates the right engine(s). (2) On twin-engine airplanes with front and rear engine locations (tandem), the left powerplant controls must operate the front engine and the right powerplant controls must operate the rear engine. (f) Wing flap and auxiliary lift device controls must be located-- (1) Centrally, or to the right of the pedestal or powerplant throttle control centerline; and (2) Far enough away from the landing gear control to avoid confusion. (g) The landing gear control must be located to the left of the throttle centerline or pedestal centerline. (h) Each fuel feed selector control must comply with Sec. 23.995 and be located and arranged so that the pilot can see and reach it without moving any seat or primary flight control when his seat is at any position in which it can be placed. (1) For a mechanical fuel selector: (i) The indication of the selected fuel valve position must be by means of a pointer and must provide positive identification and feel (detent, etc.) of the selected position.(ii) The position indicator pointer must be located at the part of the handle that is the maximum dimension of the handle measured from the center of rotation. (2) For electrical or electronic fuel selector: (i) Digital controls or electrical switches must be properly labelled.(ii) Means must be provided to indicate to the flight crew the tank or function selected. Selector switch position is not acceptable as a means of indication. The "off" or "closed" position must be indicated in red. (3) If the fuel valve selector handle or electrical or digital selection is also a fuel shut-off selector, the off position marking must be colored red. If a separate emergency shut-off means is provided, it also must be colored red. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR 13092, Aug. 13, 1969; Amdt. 23-33, 51 FR 26656, July 24, 1986; Amdt. 23-51, 61 FR 5136, Feb. 9, 1996] Sec. 23.779 Motion and effect of cockpit controls. Cockpit controls must be designed so that they operate in accordance with the following movement and actuation: (a) Aerodynamic controls: Motion and effect (1) Primary controls: Aileron Right (clockwise) for right wing down. Elevator Rearward for nose up. Rudder Right pedal forward for nose right. (2) Secondary controls: Flaps (or auxiliary lift devices) Forward or up for flaps up or auxiliary device stowed; rearward or down for flaps down or auxiliary device deployed. Trim tabs (or equivalent) Switch motion or mechanical rotation of control to produce similar rotation of the airplane about an axis parallel to the axis control. Axis of roll trim control may be displaced to accommodate comfortable actuation by the pilot. For single- engine airplanes, direction of pilot's hand movement must be in the same sense as airplane response for rudder trim if only a portion of a rotational element is accessible. (b) Powerplant and auxiliary controls: Motion and effect (1) Powerplant controls: Power (thrust) lever Forward to increase forward thrust and rearward to increase rearward thrust. Propellers Forward to increase rpm. Mixture Forward or upward for rich. Fuel Forward for open. Carburetor, air heat or alternate Forward or upward for cold. air Supercharger Forward or upward for low blower. Turbosuperchargers Forward, upward, or clockwise to increase pressure. Rotary controls Clockwise from off to full on. (2) Auxiliary controls: Fuel tank selector Right for right tanks, left for left tanks. Landing gear Down to extend. Speed brakes Aft to extend. [Amdt. 23-33, 51 FR 26656, July 24, 1986, as amended by Amdt. 23-51, 61 FR 5136, Feb. 9, 1996] Sec. 23.781 Cockpit control knob shape. (a) Flap and landing gear control knobs must conform to the general shapes (but not necessarily the exact sizes or specific proportions) in the following figure: [ ...Illustration appears here... ] Flap Control Knob [ ...Illustration appears here... ] Landing Gear Control Knob (b) Powerplant control knobs must conform to the general shapes (but not necessarily the exact sizes or specific proportions) in the following figure: [ ...Illustration appears here... ] Power (Thrust) Control Knob [ ...Illustration appears here... ] RPM Control Knob [ ...Illustration appears here... ] Mixture Control Knob [ ...Illustration appears here... ] Carb Heat or Alternate Air Control Knob [ ...Illustration appears here... ] Supercharger Control Knob [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23-33, 51 FR 26657, July 24, 1986] Sec. 23.783 Doors. (a) Each closed cabin with passenger accommodations must have at least one adequate and easily accessible external door. (b) Passenger doors must not be located with respect to any propeller disk or any other potential hazard so as to endanger persons using the door. (c) Each external passenger or crew door must comply with the following requirements: (1) There must be a means to lock and safeguard the door against inadvertent opening during flight by persons, by cargo, or as a result of mechanical failure. (2) The door must be openable from the inside and the outside when the internal locking mechanism is in the locked position. (3) There must be a means of opening which is simple and obvious and is arranged and marked inside and outside so that the door can be readily located, unlocked, and opened, even in darkness. (4) The door must meet the marking requirements of Sec. 23.811 of this part. (5) The door must be reasonably free from jamming as a result of fuselage deformation in an emergency landing. (6) Auxiliary locking devices that are actuated externally to the airplane may be used but such devices must be overridden by the normal internal opening means. (d) In addition, each external passenger or crew door, for a commuter category airplane, must comply with the following requirements: (1) Each door must be openable from both the inside and outside, even though persons may be crowded against the door on the inside of the airplane. (2) If inward opening doors are used, there must be a means to prevent occupants from crowding against the door to the extent that would interfere with opening the door. (3) Auxiliary locking devices may be used. (e) Each external door on a commuter category airplane, each external door forward of any engine or propeller on a normal, utility, or acrobatic category airplane, and each door of the pressure vessel on a pressurized airplane must comply with the following requirements: (1) There must be a means to lock and safeguard each external door, including cargo and service type doors, against inadvertent opening in flight, by persons, by cargo, or as a result of mechanical failure or failure of a single structural element, either during or after closure. (2) There must be a provision for direct visual inspection of the locking mechanism to determine if the external door, for which the initial opening movement is not inward, is fully closed and locked. The provisions must be discernible, under operating lighting conditions, by a crewmember using a flashlight or an equivalent lighting source. (3) There must be a visual warning means to signal a flight crewmember if the external door is not fully closed and locked. The means must be designed so that any failure, or combination of failures, that would result in an erroneous closed and locked indication is improbable for doors for which the initial opening movement is not inward. (f) In addition, for commuter category airplanes, the following requirements apply: (1) Each passenger entry door must qualify as a floor level emergency exit. This exit must have a rectangular opening of not less than 24 inches wide by 48 inches high, with corner radii not greater than one-third the width of the exit. (2) If an integral stair is installed at a passenger entry door, the stair must be designed so that, when subjected to the inertia loads resulting from the ultimate static load factors in Sec. 23.561(b)(2) and following the collapse of one or more legs of the landing gear, it will not reduce the effectiveness of emergency egress through the passenger entry door. (g) If lavatory doors are installed, they must be designed to preclude an occupant from becoming trapped inside the lavatory. If a locking mechanism is installed, it must be capable of being unlocked from outside of the lavatory. [Docket No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23-36, 53 FR 30813, Aug. 15, 1988; Amdt. 23-46, 59 FR 25772, May 17, 1994; Amdt. 23-49, 61 FR 5166, Feb. 12, 1996] Sec. 23.785 Seats, berths, litters, safety belts, and shoulder harnesses. There must be a seat or berth for each occupant that meets the following: (a) Each seat/restraint system and the supporting structure must be designed to support occupants weighing at least 215 pounds when subjected to the maximum load factors corresponding to the specified flight and ground load conditions, as defined in the approved operating envelope of the airplane. In addition, these loads must be multiplied by a factor of 1.33 in determining the strength of all fittings and the attachment of-- (1) Each seat to the structure; and (2) Each safety belt and shoulder harness to the seat or structure. (b) Each forward-facing or aft-facing seat/restraint system in normal, utility, or acrobatic category airplanes must consist of a seat, a safety belt, and a shoulder harness, with a metal-to-metal latching device, that are designed to provide the occupant protection provisions required in Sec. 23.562. Other seat orientations must provide the same level of occupant protection as a forward-facing or aft-facing seat with a safety belt and a shoulder harness, and must provide the protection provisions of Sec. 23.562. (c) For commuter category airplanes, each seat and the supporting structure must be designed for occupants weighing at least 170 pounds when subjected to the inertia loads resulting from the ultimate static load factors prescribed in Sec. 23.561(b)(2) of this part. Each occupant must be protected from serious head injury when subjected to the inertia loads resulting from these load factors by a safety belt and shoulder harness, with a metal-to-metal latching device, for the front seats and a safety belt, or a safety belt and shoulder harness, with a metal-to-metal latching device, for each seat other than the front seats. (d) Each restraint system must have a single-point release for occupant evacuation. (e) The restraint system for each crewmember must allow the crewmember, when seated with the safety belt and shoulder harness fastened, to perform all functions necessary for flight operations. (f) Each pilot seat must be designed for the reactions resulting from the application of pilot forces to the primary flight controls as prescribed in Sec. 23.395 of this part. (g) There must be a means to secure each safety belt and shoulder harness, when not in use, to prevent interference with the operation of the airplane and with rapid occupant egress in an emergency. (h) Unless otherwise placarded, each seat in a utility or acrobatic category airplane must be designed to accommodate an occupant wearing a parachute. (i) The cabin area surrounding each seat, including the structure, interior walls, instrument panel, control wheel, pedals, and seats within striking distance of the occupant's head or torso (with the restraint system fastened) must be free of potentially injurious objects, sharp edges, protuberances, and hard surfaces. If energy absorbing designs or devices are used to meet this requirement, they must protect the occupant from serious injury when the occupant is subjected to the inertia loads resulting from the ultimate static load factors prescribed in Sec. 23.561(b)(2) of this part, or they must comply with the occupant protection provisions of Sec. 23.562 of this part, as required in paragraphs (b) and (c) of this section. (j) Each seat track must be fitted with stops to prevent the seat from sliding off the track. (k) Each seat/restraint system may use design features, such as crushing or separation of certain components, to reduce occupant loads when showing compliance with the requirements of Sec. 23.562 of this part; otherwise, the system must remain intact. (l) For the purposes of this section, a front seat is a seat located at a flight crewmember station or any seat located alongside such a seat. (m) Each berth, or provisions for a litter, installed parallel to the longitudinal axis of the airplane, must be designed so that the forward part has a padded end-board, canvas diaphragm, or equivalent means that can withstand the load reactions from a 215-pound occupant when subjected to the inertia loads resulting from the ultimate static load factors of Sec. 23.561(b)(2) of this part. In addition-- (1) Each berth or litter must have an occupant restraint system and may not have corners or other parts likely to cause serious injury to a person occupying it during emergency landing conditions; and (2) Occupant restraint system attachments for the berth or litter must withstand the inertia loads resulting from the ultimate static load factors of Sec. 23.561(b)(2) of this part. (n) Proof of compliance with the static strength requirements of this section for seats and berths approved as part of the type design and for seat and berth installations may be shown by-- (1) Structural analysis, if the structure conforms to conventional airplane types for which existing methods of analysis are known to be reliable; (2) A combination of structural analysis and static load tests to limit load; or (3) Static load tests to ultimate loads. [Amdt. 23-36, 53 FR 30813, Aug. 15, 1988; Amdt. 23-36, 54 FR 50737, Dec. 11, 1989; Amdt. 23-49, 61 FR 5166, Feb. 9, 1996] Sec. 23.787 Baggage and cargo compartments. (a) Each baggage and cargo compartment must: (1) Be designed for its placarded maximum weight of contents and for the critical load distributions at the appropriate maximum load factors corresponding to the flight and ground load conditions of this part. (2) Have means to prevent the contents of any compartment from becoming a hazard by shifting, and to protect any controls, wiring, lines, equipment or accessories whose damage or failure would affect safe operations. (3) Have a means to protect occupants from injury by the contents of any compartment, located aft of the occupants and separated by structure, when the ultimate forward inertial load factor is 9g and assuming the maximum allowed baggage or cargo weight for the compartment. (b) Designs that provide for baggage or cargo to be carried in the same compartment as passengers must have a means to protect the occupants from injury when the baggage or cargo is subjected to the inertial loads resulting from the ultimate static load factors of Sec. 23.561(b)(3), assuming the maximum allowed baggage or cargo weight for the compartment. (c) For airplanes that are used only for the carriage of cargo, the flightcrew emergency exits must meet the requirements of Sec. 23.807 under any cargo loading conditions. [Docket No. 4080, 29 FR 17955, Dec. 18. 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23-49, 61 FR 5167, Feb. 9, 1996] Sec. 23.791 Passenger information signs. For those airplanes in which the flightcrew members cannot observe the other occupants' seats or where the flightcrew members' compartment is separated from the passenger compartment, there must be at least one illuminated sign (using either letters or symbols) notifying all passengers when seat belts should be fastened. Signs that notify when seat belts should be fastened must: (a) When illuminated, be legible to each person seated in the passenger compartment under all probable lighting conditions; and (b) Be installed so that a flightcrew member can, when seated at the flightcrew member's station, turn the illumination on and off. [Amdt. 23-49, 61 FR 5167, Feb. 9, 1996] Sec. 23.803 Emergency evacuation. (a) For commuter category airplanes, an evacuation demonstration must be conducted utilizing the maximum number of occupants for which certification is desired. The demonstration must be conducted under simulated night conditions using only the emergency exits on the most critical side of the airplane. The participants must be representative of average airline passengers with no prior practice or rehearsal for the demonstration. Evacuation must be completed within 90 seconds. (b) In addition, when certification to the emergency exit provisions of Sec. 23.807(d)(4) is requested, only the emergency lighting system required by Sec. 23.812 may be used to provide cabin interior illumination during the evacuation demonstration required in paragraph (a) of this section. [Amdt. 23-34, 52 FR 1831, Jan. 15, 1987, as amended by Amdt. 23-46, 59 FR 25773, May 17, 1994] Sec. 23.805 Flightcrew emergency exits. For airplanes where the proximity of the passenger emergency exits to the flightcrew area does not offer a convenient and readily accessible means of evacuation for the flightcrew, the following apply: (a) There must be either one emergency exit on each side of the airplane, or a top hatch emergency exit, in the flightcrew area; (b) Each emergency exit must be located to allow rapid evacuation of the crew and have a size and shape of at least a 19- by 20-inch unobstructed rectangular opening; and (c) For each emergency exit that is not less than six feet from the ground, an assisting means must be provided. The assisting means may be a rope or any other means demonstrated to be suitable for the purpose. If the assisting means is a rope, or an approved device equivalent to a rope, it must be-- (1) Attached to the fuselage structure at or above the top of the emergency exit opening or, for a device at a pilot's emergency exit window, at another approved location if the stowed device, or its attachment, would reduce the pilot's view; and (2) Able (with its attachment) to withstand a 400-pound static load. [59 FR 25773, May 17, 1994] Sec. 23.807 Emergency exits. (a) Number and location. Emergency exits must be located to allow escape without crowding in any probable crash attitude. The airplane must have at least the following emergency exits: (1) For all airplanes with a seating capacity of two or more, excluding airplanes with canopies, at least one emergency exit on the opposite side of the cabin from the main door specified in Sec. 23.783 of this part. (2) [Reserved] (3) If the pilot compartment is separated from the cabin by a door that is likely to block the pilot's escape in a minor crash, there must be an exit in the pilot's compartment. The number of exits required by paragraph (a)(1) of this section must then be separately determined for the passenger compartment, using the seating capacity of that compartment. (4) Emergency exits must not be located with respect to any propeller disk or any other potential hazard so as to endanger persons using that exit. (b) Type and operation. Emergency exits must be movable windows, panels, canopies, or external doors, openable from both inside and outside the airplane, that provide a clear and unobstructed opening large enough to admit a 19-by-26-inch ellipse. Auxiliary locking devices used to secure the airplane must be designed to be overridden by the normal internal opening means. The inside handles of emergency exits that open outward must be adequately protected against inadvertent operation. In addition, each emergency exit must-- (1) Be readily accessible, requiring no exceptional agility to be used in emergencies; (2) Have a method of opening that is simple and obvious; (3) Be arranged and marked for easy location and operation, even in darkness; (4) Have reasonable provisions against jamming by fuselage deformation; and (5) In the case of acrobatic category airplanes, allow each occupant to abandon the airplane at any speed between VSO and VD; and (6) In the case of utility category airplanes certificated for spinning, allow each occupant to abandon the airplane at the highest speed likely to be achieved in the maneuver for which the airplane is certificated. (c) Tests. The proper functioning of each emergency exit must be shown by tests. (d) Doors and exits. In addition, for commuter category airplanes, the following requirements apply: (1) In addition to the passenger entry door-- (i) For an airplane with a total passenger seating capacity of 15 or fewer, an emergency exit, as defined in paragraph (b) of this section, is required on each side of the cabin; and (ii) For an airplane with a total passenger seating capacity of 16 through 19, three emergency exits, as defined in paragraph (b) of this section, are required with one on the same side as the passenger entry door and two on the side opposite the door. (2) A means must be provided to lock each emergency exit and to safeguard against its opening in flight, either inadvertently by persons or as a result of mechanical failure. In addition, a means for direct visual inspection of the locking mechanism must be provided to determine that each emergency exit for which the initial opening movement is outward is fully locked. (3) Each required emergency exit, except floor level exits, must be located over the wing or, if not less than six feet from the ground, must be provided with an acceptable means to assist the occupants to descend to the ground. Emergency exits must be distributed as uniformly as practical, taking into account passenger seating configuration. (4) Unless the applicant has complied with paragraph (d)(1) of this section, there must be an emergency exit on the side of the cabin opposite the passenger entry door, provided that-- (i) For an airplane having a passenger seating configuration of nine or fewer, the emergency exit has a rectangular opening measuring not less than 19 inches by 26 inches high with corner radii not greater than one-third the width of the exit, located over the wing, with a step up inside the airplane of not more than 29 inches and a step down outside the airplane of not more than 36 inches; (ii) For an airplane having a passenger seating configuration of 10 to 19 passengers, the emergency exit has a rectangular opening measuring not less than 20 inches wide by 36 inches high, with corner radii not greater than one-third the width of the exit, and with a step up inside the airplane of not more than 20 inches. If the exit is located over the wing, the step down outside the airplane may not exceed 27 inches; and (iii) The airplane complies with the additional requirements of Secs. 23.561(b)(2)(iv), 23.803(b), 23.811(c), 23.812, 23.813(b), and 23.815. (e) For multiengine airplanes, ditching emergency exits must be provided in accordance with the following requirements, unless the emergency exits required by paragraph (a) or (d) of this section already comply with them: (1) One exit above the waterline on each side of the airplane having the dimensions specified in paragraph (b) or (d) of this section, as applicable; and (2) If side exits cannot be above the waterline, there must be a readily accessible overhead hatch emergency exit that has a rectangular opening measuring not less than 20 inches wide by 36 inches long, with corner radii not greater than one-third the width of the exit. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR 13092, Aug. 13, 1969; Amdt. 23-10, 36 FR 2864, Feb. 11, 1971; Amdt. 23-34, 52 FR 1831, Jan. 15, 1987; Amdt. 23-36, 53 FR 30814, Aug. 15, 1988; 53 FR 34194, Sept. 2, 1988; Amdt. 23-46, 59 FR 25773, May 17, 1994; Amdt. 23-49, 61 FR 5167, Feb. 9, 1996] Sec. 23.811 Emergency exit marking. (a) Each emergency exit and external door in the passenger compartment must be externally marked and readily identifiable from outside the airplane by-- (1) A conspicuous visual identification scheme; and (2) A permanent decal or placard on or adjacent to the emergency exit which shows the means of opening the emergency exit, including any special instructions, if applicable. (b) In addition, for commuter category airplanes, these exits and doors must be internally marked with the word "exit" by a sign which has white letters 1 inch high on a red background 2 inches high, be self-illuminated or independently, internally electrically illuminated, and have a minimum brightness of at least 160 microlamberts. The color may be reversed if the passenger compartment illumination is essentially the same. (c) In addition, when certification to the emergency exit provisions of Sec. 23.807(d)(4) is requested, the following apply: (1) Each emergency exit, its means of access, and its means of opening, must be conspicuously marked; (2) The identity and location of each emergency exit must be recognizable from a distance equal to the width of the cabin; (3) Means must be provided to assist occupants in locating the emergency exits in conditions of dense smoke; (4) The location of the operating handle and instructions for opening each emergency exit from inside the airplane must be shown by marking that is readable from a distance of 30 inches; (5) Each passenger entry door operating handle must-- (i) Be self-illuminated with an initial brightness of at least 160 microlamberts; or (ii) Be conspicuously located and well illuminated by the emergency lighting even in conditions of occupant crowding at the door; (6) Each passenger entry door with a locking mechanism that is released by rotary motion of the handle must be marked-- (i) With a red arrow, with a shaft of at least three-fourths of an inch wide and a head twice the width of the shaft, extending along at least 70 degrees of arc at a radius approximately equal to three-fourths of the handle length; (ii) So that the center line of the exit handle is within +/- one inch of the projected point of the arrow when the handle has reached full travel and has released the locking mechanism; (iii) With the word "open" in red letters, one inch high, placed horizontally near the head of the arrow; and (7) In addition to the requirements of paragraph (a) of this section, the external marking of each emergency exit must-- (i) Include a 2-inch colorband outlining the exit; and (ii) Have a color contrast that is readily distinguishable from the surrounding fuselage surface. The contrast must be such that if the reflectance of the darker color is 15 percent or less, the reflectance of the lighter color must be at least 45 percent. "Reflectance" is the ratio of the luminous flux reflected by a body to the luminous flux it receives. When the reflectance of the darker color is greater than 15 percent, at least a 30 percent difference between its reflectance and the reflectance of the lighter color must be provided. [Amdt. 23-36, 53 FR 30814, Aug. 15, 1988; 53 FR 34194, Sept. 2, 1988, as amended by Amdt. 23-46, 59 FR 25773, May 17, 1994] Sec. 23.812 Emergency lighting. When certification to the emergency exit provisions of Sec. 23.807(d)(4) is requested, the following apply: (a) An emergency lighting system, independent of the main cabin lighting system, must be installed. However, the source of general cabin illumination may be common to both the emergency and main lighting systems if the power supply to the emergency lighting system is independent of the power supply to the main lighting system. (b) There must be a crew warning light that illuminates in the cockpit when power is on in the airplane and the emergency lighting control device is not armed. (c) The emergency lights must be operable manually from the flightcrew station and be provided with automatic activation. The cockpit control device must have "on," "off," and "armed" positions so that, when armed in the cockpit, the lights will operate by automatic activation. (d) There must be a means to safeguard against inadvertent operation of the cockpit control device from the "armed" or "on" positions. (e) The cockpit control device must have provisions to allow the emergency lighting system to be armed or activated at any time that it may be needed. (f) When armed, the emergency lighting system must activate and remain lighted when-- (1) The normal electrical power of the airplane is lost; or (2) The airplane is subjected to an impact that results in a deceleration in excess of 2g and a velocity change in excess of 3.5 feet-per-second, acting along the longitudinal axis of the airplane; or (3) Any other emergency condition exists where automatic activation of the emergency lighting is necessary to aid with occupant evacuation. (g) The emergency lighting system must be capable of being turned off and reset by the flightcrew after automatic activation. (h) The emergency lighting system must provide internal lighting, including-- (1) Illuminated emergency exit marking and locating signs, including those required in Sec. 23.811(b); (2) Sources of general illumination in the cabin that provide an average illumination of not less than 0.05 foot-candle and an illumination at any point of not less than 0.01 foot-candle when measured along the center line of the main passenger aisle(s) and at the seat armrest height; and (3) Floor proximity emergency escape path marking that provides emergency evacuation guidance for the airplane occupants when all sources of illumination more than 4 feet above the cabin aisle floor are totally obscured. (i) The energy supply to each emergency lighting unit must provide the required level of illumination for at least 10 minutes at the critical ambient conditions after activation of the emergency lighting system. (j) If rechargeable batteries are used as the energy supply for the emergency lighting system, they may be recharged from the main electrical power system of the airplane provided the charging circuit is designed to preclude inadvertent battery discharge into the charging circuit faults. If the emergency lighting system does not include a charging circuit, battery condition monitors are required. (k) Components of the emergency lighting system, including batteries, wiring, relays, lamps, and switches, must be capable of normal operation after being subjected to the inertia forces resulting from the ultimate load factors prescribed in Sec. 23.561(b)(2). (l) The emergency lighting system must be designed so that after any single transverse vertical separation of the fuselage during a crash landing: (1) At least 75 percent of all electrically illuminated emergency lights required by this section remain operative; and (2) Each electrically illuminated exit sign required by Sec. 23.811 (b) and (c) remains operative, except those that are directly damaged by the fuselage separation. [T.D. 23-46, 59 FR 25774, May 17, 1994] Sec. 23.813 Emergency exit access. (a) For commuter category airplanes, access to window-type emergency exits may not be obstructed by seats or seat backs. (b) In addition, when certification to the emergency exit provisions of Sec. 23.807(d)(4) is requested, the following emergency exit access must be provided: (1) The passageway leading from the aisle to the passenger entry door must be unobstructed and at least 20 inches wide. (2) There must be enough space next to the passenger entry door to allow assistance in evacuation of passengers without reducing the unobstructed width of the passageway below 20 inches. (3) If it is necessary to pass through a passageway between passenger compartments to reach a required emergency exit from any seat in the passenger cabin, the passageway must be unobstructed; however, curtains may be used if they allow free entry through the passageway. (4) No door may be installed in any partition between passenger compartments unless that door has a means to latch it in the open position. The latching means must be able to withstand the loads imposed upon it by the door when the door is subjected to the inertia loads resulting from the ultimate static load factors prescribed in Sec. 23.561(b)(2). (5) If it is necessary to pass through a doorway separating the passenger cabin from other areas to reach a required emergency exit from any passenger seat, the door must have a means to latch i |
