|
|
We help Organizations quickly Conceptualize, Design, Develop, and Implement Innovative Cost-Effective Solutions for Avionics, Aircraft, and Flight Guidance-Navigation-Control Systems. |
|---|---|
|
|
|
Home
YOU
US
What
We Can
Do
What
We Have
Done -
On
Learjets
Why
ASTECH
ASTECH's
Founder
Expert's
Corner-
FAA FAR Part 23 C
Links
In
Closing
|
23.301 Loads. 23.321 General. Control Surface and System Loads 23.391 Control surface loads. Horizontal Stabilizing and Balancing Surfaces 23.421 Balancing loads. 23.441 Maneuvering loads. Ailerons, Wing Flaps, and Special Devices 23.455 Ailerons. 23.471 General. 23.521 Water load conditions. 23.561 General. 23.571 Metallic pressurized cabin structures. Authority: 49 U.S.C. 106(g), 40113, 44701-44702,
44704. |
|
(a) Strength requirements are specified in terms of limit loads (the maximum loads to be expected in service) and ultimate loads (limit loads multiplied by prescribed factors of safety). Unless otherwise provided, prescribed loads are limit loads. (b) Unless otherwise provided, the air, ground, and water loads must be placed in equilibrium with inertia forces, considering each item of mass in the airplane. These loads must be distributed to conservatively approximate or closely represent actual conditions. Methods used to determine load intensities and distribution on canard and tandem wing configurations must be validated by flight test measurement unless the methods used for determining those loading conditions are shown to be reliable or conservative on the configuration under consideration. (c) If deflections under load would significantly change the distribution of external or internal loads, this redistribution must be taken into account. (d) Simplified structural design criteria may be used if they result in design loads not less than those prescribed in Secs. 23.331 through 23.521. For conventional, single-engine airplanes with design weights of 6,000 pounds or less, the design criteria of Appendix A of this part are an approved equivalent of Secs. 23.321 through 23.459. If Appendix A is used, the entire Appendix must be substituted for the corresponding sections of this part. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23-28, 47 FR 13315, Mar. 29, 1982; Amdt. 23-42, 56 FR 352, Jan. 3, 1991] Sec. 23.302 Canard or tandem wing configurations. The forward structure of a canard or tandem wing configuration must: (a) Meet all requirements of subpart C and subpart D of this part applicable to a wing; and (b) Meet all requirements applicable to the function performed by these surfaces. [Doc. No. 25811, 56 FR 352, Jan. 3, 1991] Sec. 23.303 Factor of safety. Unless otherwise provided, a factor of safety of 1.5 must be used. Sec. 23.305 Strength and deformation. (a) The structure must be able to support limit loads without detrimental, permanent deformation. At any load up to limit loads, the deformation may not interfere with safe operation. (b) The structure must be able to support ultimate loads without failure for at least three seconds, except local failures or structural instabilities between limit and ultimate load are acceptable only if the structure can sustain the required ultimate load for at least three seconds. However when proof of strength is shown by dynamic tests simulating actual load conditions, the three second limit does not apply. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. No. 23-45, 58 FR 42160, Aug. 6, 1993] Sec. 23.307 Proof of structure. (a) Compliance with the strength and deformation requirements of Sec. 23.305 must be shown for each critical load condition. Structural analysis may be used only if the structure conforms to those for which experience has shown this method to be reliable. In other cases, substantiating load tests must be made. Dynamic tests, including structural flight tests, are acceptable if the design load conditions have been simulated. (b) Certain parts of the structure must be tested as
specified in Subpart D of this part. |
|
|
(a) Flight load factors 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 one in which the aerodynamic force acts upward, with respect to the airplane. (b) Compliance with the flight load requirements of this subpart must be shown-- (1) At each critical altitude within the range in which the airplane may be expected to operate; (2) At each weight from the design minimum weight to the design maximum weight; and (3) For each required altitude and weight, for any practicable distribution of disposable load within the operating limitations specified in Secs. 23.1583 through 23.1589. (c) When significant, the effects of compressibility must be taken into account. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended Amdt. No. 23-45, 58 FR 42160, Aug. 6, 1993] Sec. 23.331 Symmetrical flight conditions. (a) The appropriate balancing horizontal tail load must be accounted for in a rational or conservative manner when determining the wing loads and linear inertia loads corresponding to any of the symmetrical flight conditions specified in Secs. 23.333 through 23.341. (b) The incremental horizontal tail loads due to maneuvering and gusts must be reacted by the angular inertia of the airplane in a rational or conservative manner. (c) Mutual influence of the aerodynamic surfaces must be taken into account when determining flight loads. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-42, 56 FR 352, Jan. 3, 1991] Sec. 23.333 Flight envelope. (a) General. Compliance with the strength requirements of this subpart must be shown at any combination of airspeed and load factor on and within the boundaries of a flight envelope (similar to the one in paragraph (d) of this section) that represents the envelope of the flight loading conditions specified by the maneuvering and gust criteria of paragraphs (b) and (c) of this section respectively. (b) Maneuvering envelope. Except where limited by maximum (static) lift coefficients, the airplane is assumed to be subjected to symmetrical maneuvers resulting in the following limit load factors: (1) The positive maneuvering load factor specified in Sec. 23.337 at speeds up to VD; (2) The negative maneuvering load factor specified in Sec. 23.337 at VC; and (3) Factors varying linearly with speed from the specified value at VC to 0.0 at VD for the normal and commuter category, and --1.0 at VD for the acrobatic and utility categories. (c) Gust envelope. (1) The airplane is assumed to be subjected to symmetrical vertical gusts in level flight. The resulting limit load factors must correspond to the conditions determined as follows: (i) Positive (up) and negative (down) gusts of 50 f.p.s. at VC must be considered at altitudes between sea level and 20,000 feet. The gust velocity may be reduced linearly from 50 f.p.s. at 20,000 feet to 25 f.p.s. at 50,000 feet. (ii) Positive and negative gusts of 25 f.p.s. at VD must be considered at altitudes between sea level and 20,000 feet. The gust velocity may be reduced linearly from 25 f.p.s. at 20,000 feet to 12.5 f.p.s. at 50,000 feet. (iii) In addition, for commuter category airplanes, positive (up) and negative (down) rough air gusts of 66 f.p.s. at VB must be considered at altitudes between sea level and 20,000 feet. The gust velocity may be reduced linearly from 66 f.p.s. at 20,000 feet to 38 f.p.s. at 50,000 feet. (2) The following assumptions must be made: (i) The shape of the gust is-- Ude 2(Pi)s U ---- (1 - cos ---- ) 2 25C Where-- s Distance penetrated into gust (ft.); C Mean geometric chord of wing (ft.); and Ude Þrived gust velocity referred to in subparagraph (1) of this section. (ii) Gust load factors vary linearly with speed between VC and VD . (d) Flight envelope. [ ...Illustration appears here... ] Flight Envelope [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR 13087, Aug. 13, 1969; Amdt. 23-34, 52 FR 1829, Jan. 15, 1987] Sec. 23.335 Design airspeeds. Except as provided in paragraph (a) (4) of this section, the selected design airspeeds are equivalent airspeeds (EAS). (a) Design cruising speed, VC. For VC the following apply: (1) VC (in knots) may not be less than-- (i) 33 W/S (for normal, utility, and commuter category airplanes); and (ii) 36<radical>W/S (for acrobatic category airplanes). (2) For values of W/S more than 20, the multiplying factors may be decreased linearly with W/S to a value of 28.6 where W/S 0. (3) VC need not be more than 0.9 VH at sea level. (4) At altitudes where an MD is established, a cruising speed MC limited by compressibility may be selected. (b) Design dive speed VD. For VD, the following apply: (1) VD/MD may not be less than 1.25 VC/MC; and (2) With VC min, the required minimum design cruising speed, VD (in knots) may not be less than-- (i) 1.40 Vc min (for normal and commuter category airplanes); (ii) 1.50 VC min (for utility category airplanes); and (iii) 1.55 VC min (for acrobatic category airplanes). (3) For values of W/S more than 20, the multiplying factors in paragraph (b)(2) of this section may be decreased linearly with W/S to a value of 1.35 where W/S 0. (4) Compliance with paragraphs (b) (1) and (2) of this section need not be shown if VD/MD is selected so that the minimum speed margin between VC/MC and VD/MD is the greater of the following: (i) The speed increase resulting when, from the initial condition of stabilized flight at VC/MC, the airplane is assumed to be upset, flown for 20 seconds along a flight path 7.5 deg. below the initial path, and then pulled up with a load factor of 1.5 (0.5 g. acceleration increment). At least 75 percent maximum continuous power for reciprocating engines, and maximum cruising power for turbines, or, if less, the power required for VC/MC for both kinds of engines, must be assumed until the pullup is initiated, at which point power reduction and pilot-controlled drag devices may be used.(ii) Mach 0.05 (at altitudes where an MD is established). (c) Design maneuvering speed VA. For VA, the following applies: (1) VA may not be less than VS<radical>n where-- (i) VS is a computed stalling speed with flaps retracted at the design weight, normally based on the maximum airplane normal force coefficients, CNA; and (ii) n is the limit maneuvering load factor used in design (2) The value of VA need not exceed the value of VC used in design. (d) Simplified structural design criteria may be used if they result in design loads not less than those prescribed in Secs. 23.331 through 23.521. For airplane configurations described in appendix A, Sec. 23.1, the design criteria of appendix A of this part are an approved equivalent of Secs. 23.321 through 23.459. If appendix A of this part is used, the entire appendix must be substituted for the corresponding sections of this part. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR 13088, Aug. 13, 1969; Amdt. 23-16, 40 FR 2577, Jan. 14, 1975; Amdt. 23-34, 52 FR 1829, Jan. 15, 1987; Amdt. 23-24, 52 FR 34745, Sept. 14, 1987; Amdt. 23-48, 61 FR 5143, Feb. 9, 1996] Sec. 23.337 Limit maneuvering load factors. (a) The positive limit maneuvering load factor n may not be less than-- (1) 2.1+(24,000**(W+10,000)) for normal and commuter category airplanes, where WÞsign maximum takeoff weight, except that n need not be more than 3.8; (2) 4.4 for utility category airplanes; or (3) 6.0 for acrobatic category airplanes. (b) The negative limit maneuvering load factor may not be less than-- (1) 0.4 times the positive load factor for the normal utility and commuter categories; or (2) 0.5 times the positive load factor for the acrobatic category. (c) Maneuvering load factors lower than those specified in this section may be used if the airplane has design features that make it impossible to exceed these values in flight. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR 13088, Aug. 13, 1969; Amdt. 23-34, 52 FR 1829, Jan. 15, 1987; Amdt. 23-48, 61 FR 5144, Feb. 9, 1996] Sec. 23.341 Gust loads factors. (a) Each airplane must be designed to withstand loads on each lifting surface resulting from gusts specified in Sec. 23.333(c). (b) The gust load for a canard or tandem wing configuration must be computed using a rational analysis, or may be computed in accordance with paragraph (c) of this section, provided that the resulting net loads are shown to be conservative with respect to the gust criteria of Sec. 23.333(c). (c) In the absence of a more rational analysis, the gust load factors must be computed as follows-- KgUdeVa n 1 + ------------ 498(W/S) Where-- Kg0.88micro-g/5.3+micro-ggust alleviation factor; micro-g2(W/S)/<rho>Cagairplane mass ratio; UdeÞrived gust velocities referred to in Sec. 23.333(c) (f.p.s.); <rho>Þnsity of air (slugs/cu.ft.); W/SWing loading (p.s.f.) due to the applicable weight of the airplane in the particular load case. C Mean geometric chord (ft.); g ¨celeration due to gravity (ft./sec.**2) V Airplane equivalent speed (knots); and a Slope of the airplane normal force coefficient curve CNA per radian if the gust loads are applied to the wings and horizontal tail surfaces simultaneously by a rational method. The wing lift curve slope CL per radian may be used when the gust load is applied to the wings only and the horizontal tail gust loads are treated as a separate condition. [Amdt. 23-7, 34 FR 13088, Aug. 13, 1969, as amended by Amdt. 23-42, 56 FR 352, Jan. 3, 1991; Amdt. 23-48, 61 FR 5144, Feb. 9, 1996] Sec. 23.343 Design fuel loads. (a) The disposable load combinations must include each fuel load in the range from zero fuel to the selected maximum fuel load. (b) If fuel is carried in the wings, the maximum allowable weight of the airplane without any fuel in the wing tank(s) must be established as "maximum zero wing fuel weight," if it is less than the maximum weight. (c) For commuter category airplanes, a structural reserve fuel condition, not exceeding fuel necessary for 45 minutes of operation at maximum continuous power, may be selected. If a structural reserve fuel condition is selected, it must be used as the minimum fuel weight condition for showing compliance with the flight load requirements prescribed in this part and-- (1) The structure must be designed to withstand a condition of zero fuel in the wing at limit loads corresponding to: (i) Ninety percent of the maneuvering load factors defined in Sec. 23.337, and (ii) Gust velocities equal to 85 percent of the values prescribed in Sec. 23.333(c). (2) The fatigue evaluation of the structure must account for any increase in operating stresses resulting from the design condition of paragraph (c)(1) of this section. (3) The flutter, deformation, and vibration requirements must also be met with zero fuel in the wings. [Amdt. 23-48, 61 FR 5144, Feb. 9, 1996] Sec. 23.345 High lift devices. (a) If flaps or similar high lift devices are to be used for takeoff, approach or landing, the airplane, with the flaps fully extended at VF, is assumed to be subjected to symmetrical maneuvers and gusts within the range determined by-- (1) Maneuvering, to a positive limit load factor of 2.0; and (2) Positive and negative gust of 25 feet per second acting normal to the flight path in level flight. (b) VF must be assumed to be not less than 1.4 VS or 1.8 VSF, whichever is greater, where-- (1) VS is the computed stalling speed with flaps retracted at the design weight; and (2) VSF is the computed stalling speed with flaps fully extended at the design weight. (3) If an automatic flap load limiting device is used, the airplane may be designed for the critical combinations of airspeed and flap position allowed by that device. (c) In determining external loads on the airplane as a whole, thrust, slipstream, and pitching acceleration may be assumed to be zero. (d) The flaps, their operating mechanism, and their supporting structures, must be designed to withstand the conditions prescribed in paragraph (a) of this section. In addition, with the flaps fully extended at VF, the following conditions, taken separately, must be accounted for: (1) A head-on gust having a velocity of 25 feet per second (EAS), combined with propeller slipstream corresponding to 75 percent of maximum continuous power; and (2) The effects of propeller slipstream corresponding to maximum takeoff power. [Amdt. 23-48, 61 FR 5144, FEb. 9, 1996] Sec. 23.347 Unsymmetrical flight conditions. (a) The airplane is assumed to be subjected to the unsymmetrical flight conditions of Secs. 23.349 and 23.351. Unbalanced aerodynamic moments about the center of gravity must be reacted in a rational or conservative manner, considering the principal masses furnishing the reacting inertia forces. (b) Acrobatic category airplanes certified for flick maneuvers (snap roll) must be designed for additional asymmetric loads acting on the wing and the horizontal tail. [Docket No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-48, 61 FR 5144, Feb. 9, 1996] Sec. 23.349 Rolling conditions. The wing and wing bracing must be designed for the following loading conditions: (a) Unsymmetrical wing loads appropriate to the category. Unless the following values result in unrealistic loads, the rolling accelerations may be obtained by modifying the symmetrical flight conditions in Sec. 23.333(d) as follows: (1) For the acrobatic category, in conditions A and F, assume that 100 percent of the semispan wing airload acts on one side of the plane of symmetry and 60 percent of this load acts on the other side. (2) For normal, utility, and commuter categories, in Condition A, assume that 100 percent of the semispan wing airload acts on one side of the airplane and 75 percent of this load acts on the other side. (b) The loads resulting from the aileron deflections and speeds specified in Sec. 23.455, in combination with an airplane load factor of at least two thirds of the positive maneuvering load factor used for design. Unless the following values result in unrealistic loads, the effect of aileron displacement on wing torsion may be accounted for by adding the following increment to the basic airfoil moment coefficient over the aileron portion of the span in the critical condition determined in Sec. 23.333(d): <Delta>cm--0.01<delta> where-- <Delta>cm is the moment coefficient increment; and <delta> is the down aileron deflection in degrees in the critical condition. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR 13088, Aug. 13, 1969; Amdt. 23-34, 52 FR 1829, Jan. 15, 1987; Amdt. 23-48, 61 FR 5144, Feb. 9, 1996] Sec. 23.351 Yawing conditions. The airplane must be designed for yawing loads on the vertical surfaces resulting from the loads specified in Secs. 23.441 through 23.445. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-42, 56 FR 352, Jan. 3, 1991] Sec. 23.361 Engine torque. (a) Each engine mount and its supporting structure must be designed for the effects of-- (1) A limit engine torque corresponding to takeoff power and propeller speed acting simultaneously with 75 percent of the limit loads from flight condition A of Sec. 23.333(d); (2) A limit engine torque corresponding to maximum continuous power and propeller spped acting simultaneously with the limit loads from flight condition A of Sec. 23.333(d); and (3) For turbopropeller installations, in addition to the conditions specified in paragraphs (a)(1) and (a)(2) of this section, a limit engine torque corresponding to takeoff power and propeller speed, multiplied by a factor accounting for propeller control system malfunction, including quick feathering, acting simultaneously with lg level flight loads. In the absence of a rational analysis, a factor of 1.6 must be used. (b) For turbine engine installations, the engine mounts and supporting structure must be designed to withstand each of the following: (1) A limit engine torque load imposed by sudden engine stoppage due to malfunction or structural failure (such as compressor jamming). (2) A limit engine torque load imposed by the maximum acceleration of the engine. (c) The limit engine torque to be considered under paragraph (a) of this section must be obtained by multiplying the mean torque by a factor of-- (1) 1.25 for turbopropeller installations; (2) 1.33 for engines with five or more cylinders; and (3) Two, three, or four, for engines with four, three, or two cylinders, respectively. [Amdt. 23-26, 45 FR 60171, Sept. 11, 1980, as amended by Amdt. No. 23-45, 58 FR 42160, Aug. 6, 1993] Sec. 23.363 Side load on engine mount. (a) Each engine mount and its supporting structure must be designed for a limit load factor in a lateral direction, for the side load on the engine mount, of not less than-- (1) 1.33, or (2) One-third of the limit load factor for flight condition A. (b) The side load prescribed in paragraph (a) of this section may be assumed to be independent of other flight conditions. Sec. 23.365 Pressurized cabin loads. For each pressurized compartment, the following apply: (a) The airplane structure must be strong enough to withstand the flight loads combined with pressure differential loads from zero up to the maximum relief valve setting. (b) The external pressure distribution in flight, and any stress concentrations, must be accounted for. (c) If landings may be made with the cabin pressurized, landing loads must be combined with pressure differential loads from zero up to the maximum allowed during landing. (d) The airplane structure must be strong enough to withstand the pressure differential loads corresponding to the maximum relief valve setting multiplied by a factor of 1.33, omitting other loads. (e) If a pressurized cabin has two or more compartments separated by bulkheads or a floor, the primary structure must be designed for the effects of sudden release of pressure in any compartment with external doors or windows. This condition must be investigated for the effects of failure of the largest opening in the compartment. The effects of intercompartmental venting may be considered. Sec. 23.367 Unsymmetrical loads due to engine failure. (a) Turbopropeller airplanes must be designed for the unsymmetrical loads resulting from the failure of the critical engine including the following conditions in combination with a single malfunction of the propeller drag limiting system, considering the probable pilot corrective action on the flight controls: (1) At speeds between VMC and VD, the loads resulting from power failure because of fuel flow interruption are considered to be limit loads. (2) At speeds between VMC and VC, the loads resulting from the disconnection of the engine compressor from the turbine or from loss of the turbine blades are considered to be ultimate loads. (3) The time history of the thrust decay and drag buildup occurring as a result of the prescribed engine failures must be substantiated by test or other data applicable to the particular engine-propeller combination. (4) The timing and magnitude of the probable pilot corrective action must be conservatively estimated, considering the characteristics of the particular engine-propeller-airplane combination. (b) Pilot corrective action may be assumed to be initiated at the time maximum yawing velocity is reached, but not earlier than 2 seconds after the engine failure. The magnitude of the corrective action may be based on the limit pilot forces specified in Sec. 23.397 except that lower forces may be assumed where it is shown by analysis or test that these forces can control the yaw and roll resulting from the prescribed engine failure conditions. [Amdt. 23-7, 34 FR 13089, Aug. 13, 1969] Sec. 23.369 Rear lift truss. (a) If a rear lift truss is used, it must be designed to withstand conditions of reversed airflow at a design speed of-- V 8.7 <radical</(W/S) + 8.7 (knots), where W/S wing loading at design maximum takeoff weight. (b) Either aerodynamic data for the particular wing section used, or a value of CL equalling -0.8 with a chordwise distribution that is triangular between a peak at the trailing edge and zero at the leading edge, must be used. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR 13089, Aug. 13, 1969; 34 FR 17509, Oct. 30, 1969; Amdt. No. 23-45, 58 FR 42160, Aug. 6, 1993; Amdt. 23-48, 61 FR 5145, Feb. 9, 1996] Sec. 23.371 Gyroscopic and aerodynamic loads. (a) Each engine mount and its supporting structure must be designed for the gyroscopic, inertial, and aerodynamic loads that result, with the engine(s) and propeller(s), if applicable, at maximum continuous r.p.m., under either: (1) The conditions prescribed in Sec. 23.351 and Sec. 23.423; or (2) All possible combinations of the following-- (i) A yaw velocity of 2.5 radians per second; (ii) A pitch velocity of 1.0 radian per second; (iii) A normal load factor of 2.5; and (iv) Maximum continuous thrust. (b) For airplanes approved for aerobatic maneuvers, each engine mount and its supporting structure must meet the requirements of paragraph (a) of this section and be designed to withstand the load factors expected during combined maximum yaw and pitch velocities. (c) For airplanes certificated in the commuter category, each engine mount and its supporting structure must meet the requirements of paragraph (a) of this section and the gust conditions specified in Sec. 23.341 of this part. [Amdt. 23-48, 61 FR 5145, Feb. 9, 1996] Sec. 23.373 Speed control devices. If speed control devices (such as spoilers and drag flaps) are incorporated for use in enroute conditions-- (a) The airplane must be designed for the symmetrical maneuvers and gusts prescribed in Secs. 23.333, 23.337, and 23.341, and the yawing maneuvers and lateral gusts in Secs. 23.441 and 23.443, with the device extended at speeds up to the placard device extended speed; and (b) If the device has automatic operating or load limiting features, the airplane must be designed for the maneuver and gust conditions prescribed in paragraph (a) of this section at the speeds and corresponding device positions that the mechanism allows. [Amdt. 23-7, 34 FR 13089, Aug. 13, 1969] |
|
|
The control surface loads specified in Secs. 23.397 through 23.459 are assumed to occur in the conditions described in Secs. 23.331 through 23.351. [Docket No. 4080, 29 FR 17955, Dec. 18. 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23-48, 61 FR 5145, Feb. 9, 1996] Sec. 23.393 Loads parallel to hinge line. (a) Control surfaces and supporting hinge brackets must be designed to withstand inertial loads acting parallel to the hinge line. (b) In the absence of more rational data, the inertial loads may be assumed to be equal to KW, where-- (1) K 24 for vertical surfaces; (2) K 12 for horizontal surfaces; and (3) W weight of the movable surfaces. [Amdt. 23-48, 61 FR 5145, Feb. 9, 1996] Sec. 23.395 Control system loads. (a) Each flight control system and its supporting structure must be designed for loads corresponding to at least 125 percent of the computed hinge moments of the movable control surface in the conditions prescribed in Secs. 23.391 through 23.459. In addition, the following apply: (1) The system limit loads need not exceed the higher of the loads that can be produced by the pilot and automatic devices operating the controls. However, autopilot forces need not be added to pilot forces. The system must be designed for the maximum effort of the pilot or autopilot, whichever is higher. In addition, if the pilot and the autopilot act in opposition, the part of the system between them may be designed for the maximum effort of the one that imposes the lesser load. Pilot forces used for design need not exceed the maximum forces prescribed in Sec. 23.397(b). (2) The design must, in any case, provide a rugged system for service use, considering jamming, ground gusts, taxiing downwind, control inertia, and friction. Compliance with this subparagraph may be shown by designing for loads resulting from application of the minimum forces prescribed in Sec. 23.397(b). (b) A 125 percent factor on computed hinge moments must be used to design elevator, aileron, and rudder systems. However, a factor as low as 1.0 may be used if hinge moments are based on accurate flight test data, the exact reduction depending upon the accuracy and reliability of the data. (c) Pilot forces used for design are assumed to act at the appropriate control grips or pads as they would in flight, and to react at the attachments of the control system to the control surface horns. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR 13089, Aug. 13, 1969] Sec. 23.397 Limit control forces and torques. (a) In the control surface flight loading condition, the airloads on movable surfaces and the corresponding deflections need not exceed those that would result in flight from the application of any pilot force within the ranges specified in paragraph (b) of this section. In applying this criterion, the effects of control system boost and servo-mechanisms, and the effects of tabs must be considered. The automatic pilot effort must be used for design if it alone can produce higher control surface loads than the human pilot. (b) The limit pilot forces and torques are as follows: Maximum forces or torques for design weight, weight equal to or less than Minimum forces Control 5,000 pounds /1/ or torques /2/ Aileron: Stick 67 lbs 40 lbs. Wheel /3/ 50 D in.-lbs /4/ 40 D in.-lbs./4/ Elevator: Stick 167 lbs 100 lbs. Wheel (symmetrical) 200 lbs 100 lbs. Wheel (unsymmetrical) /5/ 100 lbs. Rudder 200 lbs 150 lbs. /1/ For design weight (W) more than 5,000 pounds, the specified maximum values must be increased linearly with weight to 1.18 times the specified values at a design weight of 12,500 pounds and for commuter category airplanes, the specified values must be increased linearly with weight to 1.35 times the specified values at a design weight of 19,000 pounds. /2/ If the design of any individual set of control systems or surfaces makes these specified minimum forces or torques inapplicable, values corresponding to the present hinge moments obtained under Sec. 23.415, but not less than 0.6 of the specified minimum forces or torques, may be used. /3/ The critical parts of the aileron control system must also be designed for a single tangential force with a limit value of 1.25 times the couple force determined from the above criteria. /4/ Dwheel diameter (inches). /5/ The unsymmetrical force must be applied at one of the normal handgrip points on the control wheel. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR 13089, Aug. 13, 1969; Amdt. 23-17, 41 FR 55464, Dec. 20, 1976; Amdt. 23-34, 52 FR 1829, Jan. 15, 1987; Amdt. No. 23-45, 58 FR 42160, Aug. 6, 1993] Sec. 23.399 Dual control system. (a) Each dual control system must be designed to withstand the force of the pilots operating in opposition, using individual pilot forces not less than the greater of-- (1) 0.75 times those obtained under Sec. 23.395; or (2) The minimum forces specified in Sec. 23.397(b). (b) Each dual control system must be designed to withstand the force of the pilots applied together, in the same direction, using individual pilot forces not less than 0.75 times those obtained under Sec. 23.395. [Amdt. 23-48, 61 FR 5145, Feb. 9, 1996] Sec. 23.405 Secondary control system. Secondary controls, such as wheel brakes, spoilers, and tab controls, must be designed for the maximum forces that a pilot is likely to apply to those controls. Sec. 23.407 Trim tab effects. The effects of trim tabs on the control surface design conditions must be accounted for only where the surface loads are limited by maximum pilot effort. In these cases, the tabs are considered to be deflected in the direction that would assist the pilot. These deflections must correspond to the maximum degree of "out of trim" expected at the speed for the condition under consideration. Sec. 23.409 Tabs. Control surface tabs must be designed for the most severe combination of airspeed and tab deflection likely to be obtained within the flight envelope for any usable loading condition. Sec. 23.415 Ground gust conditions. (a) The control system must be investigated as follows for control surface loads due to ground gusts and taxiing downwind: (1) If an investigation of the control system for ground gust loads is not required by paragraph (a)(2) of this section, but the applicant elects to design a part of the control system of these loads, these loads need only be carried from control surface horns through the nearest stops or gust locks and their supporting structures. (2) If pilot forces less than the minimums specified in Sec. 23.397(b) are used for design, the effects of surface loads due to ground gusts and taxiing downwind must be investigated for the entire control system according to the formula: H K c S q where-- H limit hinge moment (ft.-lbs.); c mean chord of the control surface aft of the hinge line (ft.); S area of control surface aft of the hinge line (sq. ft.); q dynamic pressure (p.s.f.) based on a design speed not less than 14.6 <radical</(W/S) + 14.6 (f.p.s.) where W/S wing loading at design maximum weight, except that the design speed need not exceed 88 (f.p.s.); K limit hinge moment factor for ground gusts derived in paragraph (b) of this section.(For ailerons and elevators, a positive value of K indicates a moment tending to depress the surface and a negative value of K indicates a moment tending to raise the surface). (b) The limit hinge moment factor K for ground gusts must be derived as follows: Surface K Position of controls (a) Aileron 0.75 Control column locked lashed in mid-position. (b) Aileron +/-0.50 Ailerons at full throw; + moment on one aileron, - moment on the other. (c) Elevator +/-0.75 (c) Elevator full up (-). (d) Elevator (d) Elevator full down (+). (e) Rudder +/-0.75 (e) Rudder in neutral. (f) Rudder (f) Rudder at full throw. (c) At all weights between the empty weight and the maximum weight declared for tie-down stated in the appropriate manual, any declared tie-down points and surrounding structure, control system, surfaces and associated gust locks, must be designed to withstand the limit load conditions that exist when the airplane is tied down and that result from wind speeds of up to 65 knots horizontally from any direction. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR 13089, Aug. 13, 1969; Amdt. No. 23-45, 58 FR 42160, Aug. 6, 1993; Amdt. 23-48, 61 FR 5145, Feb. 9, 1996$ |
|
|
(a) A horizontal surface balancing load is a load necessary to maintain equilibrium in any specified flight condition with no pitching acceleration. (b) Horizontal balancing surfaces must be designed for the balancing loads occurring at any point on the limit maneuvering envelope and in the flap conditions specified in Sec. 23.345. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR 13089, Aug. 13, 1969; Amdt. 23-42, 56 FR 352, Jan. 3, 1991] Sec. 23.423 Maneuvering loads. Each horizontal surface and its supporting structure, and the main wing of a canard or tandem wing configuration, if that surface has pitch control, must be designed for the maneuvering loads imposed by the following conditions: (a) A sudden movement of the pitching control, at the speed VA, to the maximum aft movement, and the maximum forward movement, as limited by the control stops, or pilot effort, whichever is critical. (b) A sudden aft movement of the pitching control at speeds above VA, followed by a forward movement of the pitching control resulting in the following combinations of normal and angular acceleration: Normal Angular acceleration acceleration Condition (n) (radian/sec**2) Nose-up pitching 1.0 +39nm.Vx(nm-1.5) Nose-down ptiching nm -39nm.Vx(nm-1.5) where-- (1) nmpositive limit maneuvering load factor used in the design of the airplane; and (2) Vinitial speed in knots. The conditions in this paragraph involve loads corresponding to the loads that may occur in a "checked maneuver" (a maneuver in which the pitching control is suddenly displaced in one direction and then suddenly moved in the opposite direction). The deflections and timing of the "checked maneuver" must avoid exceeding the limit maneuvering load factor. The total horizontal surface load for both nose-up and nose-down pitching conditions is the sum of the balancing loads at V and the specified value of the normal load factor n, plus the maneuvering load increment due to the specified value of the angular acceleration. [Doc. No. 25811, 56 FR 353, Jan. 3, 1991; Amdt. 23-42, 56 FR 5455, Feb. 11, 1991] Sec. 23.425 Gust loads. (a) Each horizontal surface, other than a main wing, must be designed for loads resulting from-- (1) Gust velocities specified in Sec. 23.333(c) with flaps retracted; and (2) Positive and negative gusts of 25 f.p.s. nominal intensity at VF corresponding to the flight conditions specified in Sec. 23.345(a)(2). (b) [Reserved] (c) When determining the total load on the horizontal surfaces for the conditions specified in paragraph (a) of this section, the initial balancing loads for steady unaccelerated flight at the pertinent design speeds VF, VC, and VD must first be determined. The incremental load resulting from the gusts must be added to the initial balancing load to obtain the total load. (d) In the absence of a more rational analysis, the incremental load due to the gust must be computed as follows only on airplane configurations with aft-mounted, horizontal surfaces, unless its use elsewhere is shown to be conservative: Kg Ude V<alpha>ht Sht de <Delta>Lht --------------------- (1- -- ) 498 da where-- <Delta>LhtIncremental horizontal tailload (lbs.); KgGust alleviation factor defined in Sec. 23.341; UdeÞrived gust velocity (f.p.s.); VAirplane equivalent speed (knots); <alpha>htSlope of aft horizontal lift curve (per radian); ShtArea of aft horizontal lift surface (ft**2); and de (1- -- ) Downwash factor da [Doc. No. 4080, 20 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR 13089 Aug. 13, 1969; Amdt. 23-42, 56 FR 353, Jan. 3, 1991] Sec. 23.427 Unsymmetrical loads. (a) Horizontal surfaces other than main wing and their supporting structure must be designed for unsymmetrical loads arising from yawing and slipstream effects, in combination with the loads prescribed for the flight conditions set forth in Secs. 23.421 through 23.425. (b) In the absence of more rational data for airplanes that are conventional in regard to location of engines, wings, horizontal surfaces other than main wing, and fuselage shape: (1) 100 percent of the maximum loading from the symmetrical flight conditions may be assumed on the surface on one side of the plane of symmetry; and (2) The following percentage of that loading must be applied to the opposite side: Percent 0-10 (n-1), where n is the specified positive maneuvering load factor, but this value may not be more than 80 percent. (c) For airplanes that are not conventional (such as airplanes with horizontal surfaces other than main wing having appreciable dihedral or supported by the vertical tail surfaces) the surfaces and supporting structures must be designed for combined vertical and horizontal surface loads resulting from each prescribed flight condition taken separately. [Amdt. 23-14, 38 FR 31820, Nov. 19, 1973, as amended
by Amdt. 23-42, 56 FR 353, Jan. 3, 1991] |
|
|
(a) At speeds up to VA, the vertical surfaces must be designed to withstand the following conditions. In computing the loads, the yawing velocity may be assumed to be zero: (1) With the airplane in unaccelerated flight at zero yaw, it is assumed that the rudder control is suddenly displaced to the maximum deflection, as limited by the control stops or by limit pilot forces. (2) With the rudder deflected as specified in paragraph (a)(1) of this section, it is assumed that the airplane yaws to the overswing sideslip angle. In lieu of a rational analysis, an overswing angle equal to 1.5 times the static sideslip angle of paragraph (a)(3) of this section may be assumed. (3) A yaw angle of 15 degrees with the rudder control maintained in the neutral position (except as limited by pilot strength). (b) For commuter category airplanes, the loads imposed by the following additional maneuver must be substantiated at speeds from VA to VD/MD. When computing the tail loads-- (1) The airplane must be yawed to the largest attainable steady state sideslip angle, with the rudder at maximum deflection caused by any one of the following: (i) Control surface stops; (ii) Maximum available booster effort; (iii) Maximum pilot rudder force as shown below: BILLING CODE 4910-13-M [ ...Illustration appears here... ] BILLING CODE 4910-13-C (2) The rudder must be suddenly displaced from the maximum deflection to the neutral position. (c) The yaw angles specified in paragraph (a)(3) of this section may be reduced if the yaw angle chosen for a particular speed cannot be exceeded in-- (1) Steady slip conditions; (2) Uncoordinated rolls from steep banks; or (3) Sudden failure of the critical engine with delayed corrective action. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR 13090, Aug. 13, 1969; Amdt. 23-14, 38 FR 31821, Nov. 19, 1973; Amdt. 23-28, 47 FR 13315, Mar. 29, 1982; Amdt. 23-42, 56 FR 353, Jan. 3, 1991; Amdt. 23-48, 61 FR 5145, Feb. 9, 1996$ Sec. 23.443 Gust loads. (a) Vertical surfaces must be designed to withstand, in unaccelerated flight at speed VC, lateral gusts of the values prescribed for VC in Sec. 23.333(c). (b) In addition, for commuter category airplanes, the airplane is assumed to encounter derived gusts normal to the plane of symmetry while in unaccelerated flight at VB, VC, VD, and VF. The derived gusts and airplane speeds corresponding to these conditions, as determined by Secs. 23.341 and 23.345, must be investigated. The shape of the gust must be as specified in Sec. 23.333(c)(2)(i). (c) In the absence of a more rational analysis, the gust load must be computed as follows: [ ...Illustration appears here... ] Where-- LvtVertical surface loads (lbs.); [ ...Illustration appears here... ] UdeÞrived gust velocity (f.p.s.); rAir density (slugs/cu.ft.); Wthe applicable weight of the airplane in the particular load case (lbs.); SvtArea of vertical surface (ft.2); ctMean geometric chord of vertical surface (ft.); avtLift curve slope of vertical surface (per radian); KRadius of gyration in yaw (ft.); lvtDistance from airplane c.g. to lift center of vertical surface (ft.); g¨celeration due to gravity (ft./sec.2); and VEquivalent airspeed (knots). [Amdt. 23-7, 34 FR 13090, Aug. 13, 1969, as amended by Amdt. 23-34, 52 FR 1830, Jan. 15, 1987; 52 FR 7262, Mar. 9, 1987; Amdt. 23-24, 52 FR 34745, Sept. 14, 1987; Amdt. 23-42, 56 FR 353, Jan. 3, 1991; Amdt. 23-48, 61 FR 5147, Feb. 9, 1996$ Sec. 23.445 Outboard fins or winglets. (a) If outboard fins or winglets are included on the horizontal surfaces or wings, the horizontal surfaces or wings must be designed for their maximum load in combination with loads induced by the fins or winglets and moments or forces exerted on the horizontal surfaces or wings by the fins or winglets. (b) If outboard fins or winglets extend above and below the horizontal surface, the critical vertical surface loading (the load per unit area as determined under Secs. 23.441 and 23.443) must be applied to-- (1) The part of the vertical surfaces above the horizontal surface with 80 percent of that loading applied to the part below the horizontal surface; and (2) The part of the vertical surfaces below the horizontal surface with 80 percent of that loading applied to the part above the horizontal surface. (c) The end plate effects of outboard fins or winglets must be taken into account in applying the yawing conditions of Secs. 23.441 and 23.443 to the vertical surfaces in paragraph (b) of this section. (d) When rational methods are used for computing loads, the maneuvering loads of Sec. 23.441 on the vertical surfaces and the one-g horizontal surface load, including induced loads on the horizontal surface and moments or forces exerted on the horizontal surfaces by the vertical surfaces, must be applied simultaneously for the structural loading condition. [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] |
|
|
(a) The ailerons must be designed for the loads to which they are subjected-- (1) In the neutral position during symmetrical flight conditions; and (2) By the following deflections (except as limited by pilot effort), during unsymmetrical flight conditions: (i) Sudden maximum displacement of the aileron control at VA. Suitable allowance may be made for control system deflections. (ii) Sufficient deflection at VC, where VC is more than VA, to produce a rate of roll not less than obtained in paragraph (a)(2)(i) of this section. (iii) Sufficient deflection at VD to produce a rate of roll not less than one-third of that obtained in paragraph (a)(2)(i) of this section. (b) [Reserved] [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR 13090, Aug. 13, 1969; Amdt. 23-42, 56 FR 353, Jan. 3, 1991] Sec. 23.457 [Removed. Amdt. 23-48, 61 FR 5147, Feb. 9, 1996] Sec. 23.459 Special devices. The loading for special devices using aerodynamic
surfaces (such as slots and spoilers) must be determined
from test data. |
|
|
The limit ground loads specified in this subpart are considered to be external loads and inertia forces that act upon an airplane structure. In each specified ground load condition, the external reactions must be placed in equilibrium with the linear and angular inertia forces in a rational or conservative manner. Sec. 23.473 Ground load conditions and assumptions. (a) The ground load requirements of this subpart must be complied with at the design maximum weight except that Secs. 23.479, 23.481, and 23.483 may be complied with at a design landing weight (the highest weight for landing conditions at the maximum descent velocity) allowed under paragraphs (b) and (c) of this section. (b) The design landing weight may be as low as-- (1) 95 percent of the maximum weight if the minimum fuel capacity is enough for at least one-half hour of operation at maximum continuous power plus a capacity equal to a fuel weight which is the difference between the design maximum weight and the design landing weight; or (2) The design maximum weight less the weight of 25 percent of the total fuel capacity. (c) The design landing weight of a multiengine airplane may be less than that allowed under paragraph (b) of this section if-- (1) The airplane meets the one-engine-inoperative climb requirements of Sec. 23.67(b)(1) or (c); and (2) Compliance is shown with the fuel jettisoning system requirements of Sec. 23.1001. (d) The selected limit vertical inertia load factor at the center of gravity of the airplane for the ground load conditions prescribed in this subpart may not be less than that which would be obtained when landing with a descent velocity (V), in feet per second, equal to 4.4 (W/S) 1/4, except that this velocity need not be more than 10 feet per second and may not be less than seven feet per second. (e) Wing lift not exceeding two-thirds of the weight of the airplane may be assumed to exist throughout the landing impact and to act through the center of gravity. The ground reaction load factor may be equal to the inertia load factor minus the ratio of the above assumed wing lift to the airplane weight. (f) If energy absorption tests are made to determine the limit load factor corresponding to the required limit descent velocities, these tests must be made under Sec. 23.723(a). (g) No inertia load factor used for design purposes may be less than 2.67, nor may the limit ground reaction load factor be less than 2.0 at design maximum weight, unless these lower values will not be exceeded in taxiing at speeds up to takeoff speed over terrain as rough as that expected in service. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR 13090, Aug. 13, 1969; Amdt. 23-28, 47 FR 13315, Mar. 29, 1982; Amdt. No. 23-45, 58 FR 42160, Aug. 6, 1993; Amdt. 23-48, 61 FR 5147, Feb. 9, 1996$ Sec. 23.477 Landing gear arrangement. Sections 23.479 through 23.483, or the conditions in Appendix C, apply to airplanes with conventional arrangements of main and nose gear, or main and tail gear. Sec. 23.479 Level landing conditions. (a) For a level landing, the airplane is assumed to be in the following attitudes: (1) For airplanes with tail wheels, a normal level flight attitude. (2) For airplanes with nose wheels, attitudes in which-- (i) The nose and main wheels contact the ground simultaneously; and (ii) The main wheels contact the ground and the nose wheel is just clear of the ground. The attitude used in paragraph (a)(2)(i) of this section may be used in the analysis required under paragraph (a)(2)(ii) of this section. (b) When investigating landing conditions, the drag components simulating the forces required to accelerate the tires and wheels up to the landing speed (spin-up) must be properly combined with the corresponding instantaneous vertical ground reactions, and the forward-acting horizontal loads resulting from rapid reduction of the spin-up drag loads (spring-back) must be combined with vertical ground reactions at the instant of the peak forward load, assuming wing lift and a tire-sliding coefficient of friction of 0.8. However, the drag loads may not be less than 25 percent of the maximum vertical ground reactions (neglecting wing lift). (c) In the absence of specific tests or a more rational analysis for determining the wheel spin-up and spring-back loads for landing conditions, the method set forth in appendix D of this part must be used. If appendix D of this part is used, the drag components used for design must not be less than those given by appendix C of this part. (d) For airplanes with tip tanks or large overhung masses (such as turbo- propeller or jet engines) supported by the wing, the tip tanks and the structure supporting the tanks or overhung masses must be designed for the effects of dynamic responses under the level landing conditions of either paragraph (a)(1) or (a)(2)(ii) of this section. In evaluating the effects of dynamic response, an airplane lift equal to the weight of the airplane may be assumed. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-17, 41 FR 55464, Dec. 20, 1976; Amdt. No. 23-45, 58 FR 42160, Aug. 6, 1993] Sec. 23.481 Tail down landing conditions. (a) For a tail down landing, the airplane is assumed to be in the following attitudes: (1) For airplanes with tail wheels, an attitude in which the main and tail wheels contact the ground simultaneously. (2) For airplanes with nose wheels, a stalling attitude, or the maximum angle allowing ground clearance by each part of the airplane, whichever is less. (b) For airplanes with either tail or nose wheels, ground reactions are assumed to be vertical, with the wheels up to speed before the maximum vertical load is attained. Sec. 23.483 One-wheel landing conditions. For the one-wheel landing condition, the airplane is assumed to be in the level attitude and to contact the ground on one side of the main landing gear. In this attitude, the ground reactions must be the same as those obtained on that side under Sec. 23.479. Sec. 23.485 Side load conditions. (a) For the side load condition, the airplane is assumed to be in a level attitude with only the main wheels contacting the ground and with the shock absorbers and tires in their static positions. (b) The limit vertical load factor must be 1.33, with the vertical ground reaction divided equally between the main wheels. (c) The limit side inertia factor must be 0.83, with the side ground reaction divided between the main wheels so that-- (1) 0.5 (W) is acting inboard on one side; and (2) 0.33 (W) is acting outboard on the other side. (d) The side loads prescribed in paragraph (c) of this section are assumed to be applied at the ground contact point and the drag loads may be assumed to be zero. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23-45, 58 FR 42160, Aug. 6, 1993] Sec. 23.493 Braked roll conditions. Under braked roll conditions, with the shock absorbers and tires in their static positions, the following apply: (a) The limit vertical load factor must be 1.33. (b) The attitudes and ground contacts must be those described in Sec. 23.479 for level landings. (c) A drag reaction equal to the vertical reaction at the wheel multiplied by a coefficient of friction of 0.8 must be applied at the ground contact point of each wheel with brakes, except that the drag reaction need not exceed the maximum value based on limiting brake torque. Sec. 23.497 Supplementary conditions for tail wheels. In determining the ground loads on the tail wheel and affected supporting structures, the following apply: (a) For the obstruction load, the limit ground reaction obtained in the tail down landing condition is assumed to act up and aft through the axle at 45 degrees. The shock absorber and tire may be assumed to be in their static positions. (b) For the side load, a limit vertical ground reaction equal to the static load on the tail wheel, in combination with a side component of equal magnitude, is assumed. In addition-- (1) If a swivel is used, the tail wheel is assumed to be swiveled 90 degrees to the airplane longitudinal axis with the resultant ground load passing through the axle; (2) If a lock, steering device, or shimmy damper is used, the tail wheel is also assumed to be in the trailing position with the side load acting at the ground contact point; and (3) The shock absorber and tire are assumed to be in their static positions. (c) If a tail wheel, bumper, or an energy absorption device is provided to show compliance with Sec. 23.925(b), the following apply: (1) Suitable design loads must be established for the tail wheel, bumper, or energy absorption device; and (2) The supporting structure of the tail wheel, bumper, or energy absorption device must be designed to withstand the loads established in paragraph (c)(1) of this section. [Docket No. 4080, 29 FR 17955, Dec. 18. 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23-48, 61 FR 5147, Feb. 9, 1996$ Sec. 23.499 Supplementary conditions for nose wheels. In determining the ground loads on nose wheels and affected supporting structures, and assuming that the shock absorbers and tires are in their static positions, the following conditions must be met: (a) For aft loads, the limit force components at the axle must be-- (1) A vertical component of 2.25 times the static load on the wheel; and (2) A drag component of 0.8 times the vertical load. (b) For forward loads, the limit force components at the axle must be-- (1) A vertical component of 2.25 times the static load on the wheel; and (2) A forward component of 0.4 times the vertical load. (c) For side loads, the limit force components at ground contact must be-- (1) A vertical component of 2.25 times the static load on the wheel; and (2) A side component of 0.7 times the vertical load. (d) For airplanes with a steerable nose wheel that is controlled by hydraulic or other power, at design takeoff weight with the nose wheel in any steerable position, the application of 1.33 times the full steering torque combined with a vertical reaction equal to 1.33 times the maximum static reaction on the nose gear must be assumed. However, if a torque limiting device is installed, the steering torque can be reduced to the maximum value allowed by that device. (e) For airplanes with a steerable nose wheel that has a direct mechanical connection to the rudder pedals, the mechanism must be designed to withstand the steering torque for the maximum pilot forces specified in Sec. 23.397(b). [Docket No. 4080, 29 FR 17955, Dec. 18. 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23-48, 61 FR 5147, Feb. 9, 1996$ Sec. 23.505 Supplementary conditions for skiplanes. In determining ground loads for skiplanes, and assuming that the airplane is resting on the ground with one main ski frozen at rest and the other skis free to slide, a limit side force equal to 0.036 times the design maximum weight must be applied near the tail assembly, with a factor of safety of 1. [Amdt. 23-7, 34 FR 13090, Aug. 13, 1969] Sec. 23.507 Jacking loads. (a) The airplane must be designed for the loads developed when the aircraft is supported on jacks at the design maximum weight assuming the following load factors for landing gear jacking points at a three-point attitude and for primary flight structure jacking points in the level attitude: (1) Vertical-load factor of 1.35 times the static reactions. (2) Fore, aft, and lateral load factors of 0.4 times the vertical static reactions. (b) The horizontal loads at the jack points must be reacted by inertia forces so as to result in no change in the direction of the resultant loads at the jack points. (c) The horizontal loads must be considered in all combinations with the vertical load. [Amdt. 23-14, 38 FR 31821, Nov. 19, 1973] Sec. 23.509 Towing loads. The towing loads of this section must be applied to the design of tow fittings and their immediate attaching structure. (a) The towing loads specified in paragraph (d) of this section must be considered separately. These loads must be applied at the towing fittings and must act parallel to the ground. In addition: (1) A vertical load factor equal to 1.0 must be considered acting at the center of gravity; and (2) The shock struts and tires must be in there static positions. (b) For towing points not on the landing gear but near the plane of symmetry of the airplane, the drag and side tow load components specified for the auxiliary gear apply. For towing points located outboard of the main gear, the drag and side tow load components specified for the main gear apply. Where the specified angle of swivel cannot be reached, the maximum obtainable angle must be used. (c) The towing loads specified in paragraph (d) of this section must be reacted as follows: (1) The side component of the towing load at the main gear must be reacted by a side force at the static ground line of the wheel to which the load is applied. (2) The towing loads at the auxiliary gear and the drag components of the towing loads at the main gear must be reacted as follows: (i) A reaction with a maximum value equal to the vertical reaction must be applied at the axle of the wheel to which the load is applied. Enough airplane inertia to achieve equilibrium must be applied.(ii) The loads must be reacted by airplane inertia. (d) The prescribed towing loads are as follows, where W is the design maximum weight: Load Tow point Position Magnitude No. Direction Main gear 0.225W 1 Forward, 2 parallel to 3 drag axis. 4 Forward, at 30 deg. to drag axis. Aft, parallel to drag axis. Aft, at 30 deg. to drag axis. Auxiliary gear Swiveled 0.3W 5 Forward. forward 6 Aft. Swiveled aft 0.3W 7 Forward. 8 Aft. Swiveled 45 0.15W 9 Forward, in deg. from 10 plane of forward wheel. Aft, in plane of wheel. Swiveled 45 0.15W 11 Forward, in deg. from aft 12 plane of wheel. Aft, in plane of wheel. [Amdt. 23-14, 38 FR 31821, Nov. 19, 1973] Sec. 23.511 Ground load; unsymmetrical loads on multiple-wheel units. (a) Pivoting loads. The airplane is assumed to pivot about on side of the main gear with-- (1) The brakes on the pivoting unit locked; and (2) Loads corresponding to a limit vertical load factor of 1, and coefficient of friction of 0.8 applied to the main gear and its supporting structure. (b) Unequal tire loads. The loads established under Secs. 23.471 through 23.483 must be applied in turn, in a 60/40 percent distribution, to the dual wheels and tires in each dual wheel landing gear unit. (c) Deflated tire loads. For the deflated tire condition-- (1) 60 percent of the loads established under Secs. 23.471 through 23.483 must be applied in turn to each wheel in a landing gear unit; and (2) 60 percent of the limit drag and side loads, and 100 percent of the limit vertical load established under Secs. 23.485 and 23.493 or lesser vertical load obtained under paragraph (c)(1) of this section, must be applied in turn to each wheel in the dual wheel landing gear unit. [Amdt. 23-7, 34 FR 13090, Aug. 13, 1969] |
|
|
(a) The structure of seaplanes and amphibians must be designed for water loads developed during takeoff and landing with the seaplane in any attitude likely to occur in normal operation at appropriate forward and sinking velocities under the most severe sea conditions likely to be encountered. (b) Unless the applicant makes a rational analysis of the water loads, Secs. 23.523 through 23.537 apply. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. No. 23-45, 58 FR 42160, Aug. 6, 1993; Amdt. 23-48, 61 FR 5147, Feb. 9, 1996$ Sec. 23.523 Design weights and center of gravity positions. (a) Design weights. The water load requirements must be met at each operating weight up to the design landing weight except that, for the takeoff condition prescribed in Sec. 23.531, the design water takeoff weight (the maximum weight for water taxi and takeoff run) must be used. (b) Center of gravity positions. The critical centers of gravity within the limits for which certification is requested must be considered to reach maximum design loads for each part of the seaplane structure. [Amdt. No. 23-45, 58 FR 42160, Aug. 6, 1993] Sec. 23.525 Application of loads. (a) Unless otherwise prescribed, the seaplane as a whole is assumed to be subjected to the loads corresponding to the load factors specified in Sec. 23.527. (b) In applying the loads resulting from the load factors prescribed in Sec. 23.527, the loads may be distributed over the hull or main float bottom (in order to avoid excessive local shear loads and bending moments at the location of water load application) using pressures not less than those prescribed in Sec. 23.533(c). (c) For twin float seaplanes, each float must be treated as an equivalent hull on a fictitious seaplane with a weight equal to one-half the weight of the twin float seaplane. (d) Except in the takeoff condition of Sec. 23.531, the aerodynamic lift on the seaplane during the impact is assumed to be 2/3 of the weight of the seaplane. [Amdt. No. 23-45, 58 FR 42161, Aug. 6, 1993; 58 FR 51970, Oct. 5, 1993] Sec. 23.527 Hull and main float load factors. (a) Water reaction load factors nw must be computed in the following manner: (1) For the step landing case C1(VSO)**2 nw ------------------------- (Tan**(2/3)b)W**(1/3) (2) For the bow and stern landing cases C1(VSO)**2 K1 nw ------------------------- x ------------------ (Tan**(2/3)b)W**(1/3) (1+(rx)**2)**(2/3) (b) The following values are used: (1) nwwater reaction load factor (that is, the water reaction divided by seaplane weight). (2) C1empirical seaplane operations factor equal to 0.012 (except that this factor may not be less than that necessary to obtain the minimum value of step load factor of 2.33). (3) VSOseaplane stalling speed in knots with flaps extended in the appropriate landing position and with no slipstream effect. (4) bAngle of dead rise at the longitudinal station at which the load factor is being determined in accordance with figure 1 of appendix I of this part. (5) Wseaplane landing weight in pounds. (6) K1empirical hull station weighing factor, in accordance with figure 2 of appendix I of this part. (7) rxratio of distance, measured parallel to hull reference axis, from the center of gravity of the seaplane to the hull longitudinal station at which the load factor is being computed to the radius of gyration in pitch of the seaplane, the hull reference axis being a straight line, in the plane of symmetry, tangential to the keel at the main step. (c) For a twin float seaplane, because of the effect of flexibility of the attachment of the floats to the seaplane, the factor K1 may be reduced at the bow and stern to 0.8 of the value shown in figure 2 of appendix I of this part. This reduction applies only to the design of the carrythrough and seaplane structure. [Amdt. No. 23-45, 58 FR 42161, Aug. 6, 1993; 58 FR 51970, Oct. 5, 1993] Sec. 23.529 Hull and main float landing conditions. (a) Symmetrical step, bow, and stern landing. For symmetrical step, bow, and stern landings, the limit water reaction load factors are those computed under Sec. 23.527. In addition-- (1) For symmetrical step landings, the resultant water load must be applied at the keel, through the center of gravity, and must be directed perpendicularly to the keel line; (2) For symmetrical bow landings, the resultant water load must be applied at the keel, one-fifth of the longitudinal distance from the bow to the step, and must be directed perpendicularly to the keel line; and (3) For symmetrical stern landings, the resultant water load must be applied at the keel, at a point 85 percent of the longitudinal distance from the step to the stern post, and must be directed perpendicularly to the keel line. (b) Unsymmetrical landing for hull and single float seaplanes. Unsymmetrical step, bow, and stern landing conditions must be investigated. In addition-- (1) The loading for each condition consists of an upward component and a side component equal, respectively, to 0.75 and 0.25 tan b times the resultant load in the corresponding symmetrical landing condition; and (2) The point of application and direction of the upward component of the load is the same as that in the symmetrical condition, and the point of application of the side component is at the same longitudinal station as the upward component but is directed inward perpendicularly to the plane of symmetry at a point midway between the keel and chine lines. (c) Unsymmetrical landing; twin float seaplanes. The unsymmetrical loading consists of an upward load at the step of each float of 0.75 and a side load of 0.25 tan b at one float times the step landing load reached under Sec. 23.527. The side load is directed inboard, perpendicularly to the plane of symmetry midway between the keel and chine lines of the float, at the same longitudinal station as the upward load. [Amdt. No. 23-45, 58 FR 42161, Aug. 6, 1993] Sec. 23.531 Hull and main float takeoff condition. For the wing and its attachment to the hull or main float-- (a) The aerodynamic wing lift is assumed to be zero; and (b) A downward inertia load, corresponding to a load factor computed from the following formula, must be applied: CTO(VSI)**2 n --------------------- (Tan**(2/3)b)W**(1/3) Where-- ninertia load factor; CTOempirical seaplane operations factor equal to 0.004; VS1seaplane stalling speed (knots) at the design takeoff weight with the flaps extended in the appropriate takeoff position; bangle of dead rise at the main step (degrees); and WÞsign water takeoff weight in pounds. [Amdt. No. 23-45, 58 FR 42161, Aug. 6, 1993] Sec. 23.533 Hull and main float bottom pressures. (a) General. The hull and main float structure, including frames and bulkheads, stringers, and bottom plating, must be designed under this section. (b) Local pressures. For the design of the bottom plating and stringers and their attachments to the supporting structure, the following pressure distributions must be applied: (1) For an unflared bottom, the pressure at the chine is 0.75 times the pressure at the keel, and the pressures between the keel and chine vary linearly, in accordance with figure 3 of appendix I of this part. The pressure at the keel (p.s.i.) is computed as follows: C2K2(VSI)**2 PK ---------------- Tan(bk) where-- Pkpressure (p.s.i.) at the keel; C20.00213; K2hull station weighing factor, in accordance with figure 2 of appendix I of this part; VS1seaplane stalling speed (knots) at the design water takeoff weight with flaps extended in the appropriate takeoff position; and bKangle of dead rise at keel, in accordance with figure 1 of appendix I of this part. (2) For a flared bottom, the pressure at the beginning of the flare is the same as that for an unflared bottom, and the pressure between the chine and the beginning of the flare varies linearly, in accordance with figure 3 of appendix I of this part. The pressure distribution is the same as that prescribed in paragraph (b)(1) of this section for an unflared bottom except that the pressure at the chine is computed as follows: C3K2(VSI)**2 Pch ---------------- Tan(b) where-- Pchpressure (p.s.i.) at the chine; C30.0016; K2hull station weighing factor, in accordance with figure 2 of appendix I of this part; VS1seaplane stalling speed (knots) at the design water takeoff weight with flaps extended in the appropriate takeoff position; and bangle of dead rise at appropriate station. The area over which these pressures are applied must simulate pressures occurring during high localized impacts on the hull or float, but need not extend over an area that would induce critical stresses in the frames or in the overall structure. (c) Distributed pressures. For the design of the frames, keel, and chine structure, the following pressure distributions apply: (1) Symmetrical pressures are computed as follows: C4K2(VSO)**2 P ---------------- Tan(b) where-- Ppressure (p.s.i.); C40.078 C1 (with C1 computed under Sec. 23.527); K2hull station weighing factor, determined in accordance with figure 2 of appendix I of this part; VS0seaplane stalling speed (knots) with landing flaps extended in the appropriate position and with no slipstream effect; and bangle of dead rise at appropriate station. (2) The unsymmetrical pressure distribution consists of the pressures prescribed in paragraph (c)(1) of this section on one side of the hull or main float centerline and one-half of that pressure on the other side of the hull or main float centerline, in accordance with figure 3 of appendix I of this part. (3) These pressures are uniform and must be applied simultaneously over the entire hull or main float bottom. The loads obtained must be carried into the sidewall structure of the hull proper, but need not be transmitted in a fore and aft direction as shear and bending loads. [Amdt. No. 23-45, 58 FR 42161, Aug. 6, 1993; 58 FR 51970, Oct. 5, 1993] Sec. 23.535 Auxiliary float loads. (a) General. Auxiliary floats and their attachments and supporting structures must be designed for the conditions prescribed in this section. In the cases specified in paragraphs (b) through (e) of this section, the prescribed water loads may be distributed over the float bottom to avoid excessive local loads, using bottom pressures not less than those prescribed in paragraph (g) of this section. (b) Step loading. The resultant water load must be applied in the plane of symmetry of the float at a point three-fourths of the distance from the bow to the step and must be perpendicular to the keel. The resultant limit load is computed as follows, except that the value of L need not exceed three times the weight of the displaced water when the float is completely submerged: C5(VSO)**2(W**(2/3)) L ------------------------------ Tan**(2/3)bs(1+(ry)**2)**(2/3) where-- Llimit load (lbs.); C50.0053; VS0seaplane stalling speed (knots) with landing flaps extended in the appropriate position and with no slipstream effect; Wseaplane design landing weight in pounds; bsangle of dead rise at a station 3/4 of the distance from the bow to the step, but need not be less than 15 degrees; and ryratio of the lateral distance between the center of gravity and the plane of symmetry of the float to the radius of gyration in roll. (c) Bow loading. The resultant limit load must be applied in the plane of symmetry of the float at a point one-fourth of the distance from the bow to the step and must be perpendicular to the tangent to the keel line at that point. The magnitude of the resultant load is that specified in paragraph (b) of this section. (d) Unsymmetrical step loading. The resultant water load consists of a component equal to 0.75 times the load specified in paragraph (a) of this section and a side component equal to 0.025 tan b times the load specified in paragraph (b) of this section. The side load must be applied perpendicularly to the plane of symmetry of the float at a point midway between the keel and the chine. (e) Unsymmetrical bow loading. The resultant water load consists of a component equal to 0.75 times the load specified in paragraph (b) of this section and a side component equal to 0.25 tan b times the load specified in paragraph (c) of this section. The side load must be applied perpendicularly to the plane of symmetry at a point midway between the keel and the chine. (f) Immersed float condition. The resultant load must be applied at the centroid of the cross section of the float at a point one-third of the distance from the bow to the step. The limit load components are as follows: vertical PgV CXP(V**(2/3))(K VSO)**2 aft ----------------------- 2 CYP(V**(2/3))(K VSO)**2 side ----------------------- 2 where-- Pmass density of water (slugs/ft. 3 ) Vvolume of float (ft. 3 ); CXcoefficient of drag force, equal to 0.133; CYcoefficient of side force, equal to 0.106; K0.8, except that lower values may be used if it is shown that the floats are incapable of submerging at a speed of 0.8 Vso in normal operations; Vsoseaplane stalling speed (knots) with landing flaps extended in the appropriate position and with no slipstream effect; and g¨celeration due to gravity (ft/sec**2). (g) Float bottom pressures. The float bottom pressures must be established under Sec. 23.533, except that the value of K2 in the formulae may be taken as 1.0. The angle of dead rise to be used in determining the float bottom pressures is set forth in paragraph (b) of this section. [Amdt. No. 23-45, 58 FR 42162, Aug. 6, 1993; 58 FR 51970, Oct. 5, 1993] Sec. 23.537 Seawing loads. Seawing design loads must be based on applicable test data. [Amdt. No. 23-45, 58 FR 42163, Aug. 6, 1993] |
|
|
(a) The airplane, although it may be damaged in emergency landing conditions, must be designed as prescribed in this section to protect each occupant under those conditions. (b) The structure must be designed to give each occupant every reasonable chance of escaping serious injury when-- (1) Proper use is made of the seats, safety belts, and shoulder harnesses provided for in the design; (2) The occupant experiences the static inertia loads corresponding to the following ultimate load factors-- (i) Upward, 3.0g for normal, utility, and commuter category airplanes, or 4.5g for acrobatic category airplanes; (ii) Forward, 9.0g; (iii) Sideward, 1.5g; and (iv) Downward, 6.0g when certification to the emergency exit provisions of Sec. 23.807(d)(4) is requested; and (3) The items of mass within the cabin, that could injure an occupant, experience the static inertia loads corresponding to the following ultimate load factors-- (i) Upward, 3.0g; (ii) Forward, 18.0g; and (iii) Sideward, 4.5g. (c) Each airplane with retractable landing gear must be designed to protect each occupant in a landing-- (1) With the wheels retracted; (2) With moderate descent velocity; and (3) Assuming, in the absence of a more rational analysis-- (i) A downward ultimate inertia force of 3 g; and (ii) A coefficient of friction of 0.5 at the ground. (d) If it is not established that a turnover is unlikely during an emergency landing, the structure must be designed to protect the occupants in a complete turnover as follows: (1) The likelihood of a turnover may be shown by an analysis assuming the following conditions-- (i) The most adverse combination of weight and center of gravity position; (ii) Longitudinal load factor of 9.0g; (iii) Vertical load factor of 1.0g; and (iv) For airplanes with tricycle landing gear, the nose wheel strut failed with the nose contacting the ground. (2) For determining the loads to be applied to the inverted airplane after a turnover, an upward ultimate inertia load factor of 3.0g and a coefficient of friction with the ground of 0.5 must be used. (e) Except as provided in Sec. 23.787(c), the supporting structure must be designed to restrain, under loads up to those specified in paragraph (b)(3) of this section, each item of mass that could injure an occupant if it came loose in a minor crash landing. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-7, 34 FR 13090, Aug. 13, 1969; Amdt. 23-24, 52 FR 34745, Sept. 14, 1987; Amdt. 23-36, 53 FR 30812, Aug. 15, 1988; Amdt. 23-46, 59 FR 25772, May 17, 1994; Amdt. 23-48, 61 FR 5147, Feb. 9, 1996$ Sec. 23.562 Emergency landing dynamic conditions. (a) Each seat/restraint system for use in a normal, utility, or acrobatic category airplane must be designed to protect each occupant during an emergency landing when-- (1) Proper use is made of seats, safety belts, and shoulder harnesses provided for in the design; and (2) The occupant is exposed to the loads resulting from the conditions prescribed in this section. (b) Except for those seat/restraint systems that are required to meet paragraph (d) of this section, each seat/restraint system for crew or passenger occupancy in a normal, utility, or acrobatic category airplane, must successfully complete dynamic tests or be demonstrated by rational analysis supported by dynamic tests, in accordance with each of the following conditions. These tests must be conducted with an occupant simulated by an anthropomorphic test dummy (ATD) defined by 49 CFR Part 572, Subpart B, or an FAA-approved equivalent, with a nominal weight of 170 pounds and seated in the normal upright position. (1) For the first test, the change in velocity may not be less than 31 feet per second. The seat/restraint system must be oriented in its nominal position with respect to the airplane and with the horizontal plane of the airplane pitched up 60 degrees, with no yaw, relative to the impact vector. For seat/restraint systems to be installed in the first row of the airplane, peak deceleration must occur in not more than 0.05 seconds after impact and must reach a minimum of 19g. For all other seat/restraint systems, peak deceleration must occur in not more than 0.06 seconds after impact and must reach a minimum of 15g. (2) For the second test, the change in velocity may not be less than 42 feet per second. The seat/restraint system must be oriented in its nominal position with respect to the airplane and with the vertical plane of the airplane yawed 10 degrees, with no pitch, relative to the impact vector in a direction that results in the greatest load on the shoulder harness. For seat/restraint systems to be installed in the first row of the airplane, peak deceleration must occur in not more than 0.05 seconds after impact and must reach a minimum of 26g. For all other seat/restraint systems, peak deceleration must occur in not more than 0.06 seconds after impact and must reach a minimum of 21g. (3) To account for floor warpage, the floor rails or attachment devices used to attach the seat/restraint system to the airframe structure must be preloaded to misalign with respect to each other by at least 10 degrees vertically (i.e., pitch out of parallel) and one of the rails or attachment devices must be preloaded to misalign by 10 degrees in roll prior to conducting the test defined by paragraph (b)(2) of this section. (c) Compliance with the following requirements must be shown during the dynamic tests conducted in accordance with paragraph (b) of this section: (1) The seat/restraint system must restrain the ATD although seat/restraint system components may experience deformation, elongation, displacement, or crushing intended as part of the design. (2) The attachment between the seat/restraint system and the test fixture must remain intact, although the seat structure may have deformed. (3) Each shoulder harness strap must remain on the ATD's shoulder during the impact. (4) The safety belt must remain on the ATD's pelvis during the impact. (5) The results of the dynamic tests must show that the occupant is protected from serious head injury.(i) When contact with adjacent seats, structure, or other items in the cabin can occur, protection must be provided so that the head impact does not exceed a head injury criteria (HIC) of 1,000.(ii) The value of HIC is defined as-- 1 t2 2.5 HIC { (t2-t1) [---------- S a(t)dt ] } (t2-t1) t1 Max Where: t1 is the initial integration time, expressed in seconds, t2 is the final integration time, expressed in seconds, (t2-t1) is the time duration of the major head impact, expressed in seconds, and a(t) is the resultant deceleration at the center of gravity of the head form expressed as a multiple of g (units of gravity). (iii) Compliance with the HIC limit must be demonstrated by measuring the head impact during dynamic testing as prescribed in paragraphs (b)(1) and (b)(2) of this section or by a separate showing of compliance with the head injury criteria using test or analysis procedures. (6) Loads in individual shoulder harness straps may not exceed 1,750 pounds. If dual straps are used for retaining the upper torso, the total strap loads may not exceed 2,000 pounds. (7) The compression load measured between the pelvis and the lumbar spine of the ATD may not exceed 1,500 pounds. (d) For all single-engine airplanes with a VSO of more than 61 knots at maximum weight, and those multiengine airplanes of 6,000 pounds or less maximum weight with a VSO of more than 61 knots at maximum weight that do not comply with Sec. 23.67(a)(1); (1) The ultimate load factors of Sec. 23.561(b) must be increased by multiplying the load factors by the square of the ratio of the increased stall speed to 61 knots. The increased ultimate load factors need not exceed the values reached at a VS0 of 79 knots. The upward ultimate load factor for acrobatic category airplanes need not exceed 5.0g. (2) The seat/restraint system test required by paragraph (b)(1) of this section must be conducted in accordance with the following criteria: (i) The change in velocity may not be less than 31 feet per second.(ii)(A) The peak deceleration (gp) of 19g and 15g must be increased and multiplied by the square of the ratio of the increased stall speed to 61 knots: gp .0 (VS0/61)**2 or gp .0 (VS0/61)**2 (B) The peak deceleration need not exceed the value reached at a VS0 of 79 knots.(iii) The peak deceleration must occur in not more than time (tr), which must be computed as follows: 31 .96 tr -------- --- 32.2(gp) gp where-- gpThe peak deceleration calculated in accordance with paragraph (d)(2)(ii) of this section trThe rise time (in seconds) to the peak deceleration. (e) An alternate approach that achieves an equivalent, or greater, level of occupant protection to that required by this section may be used if substantiated on a rational basis. [Amdt. 23-36, 53 FR 30812, Aug. 15, 1988, as amended by Amdt. 23-44, 58 FR 38639, July 19, 1993; Amdt. 23-50, 61 FR 5192, Feb. 9, 1996] |
|
|
For normal, utility, and acrobatic category airplanes, the strength, detail design, and fabrication of the metallic structure of the pressure cabin must be evaluated under one of the following: (a) A fatigue strength investigation in which the structure is shown by tests, or by analysis supported by test evidence, to be able to withstand the repeated loads of variable magnitude expected in service; or (b) A fail safe strength investigation, in which it is shown by analysis, tests, or both that catastrophic failure of the structure is not probable after fatigue failure, or obvious partial failure, of a principal structural element, and that the remaining structures are able to withstand a static ultimate load factor of 75 percent of the limit load factor at VC, considering the combined effects of normal operating pressures, expected external aerodynamic pressures, and flight loads. These loads must be multiplied by a factor of 1.15 unless the dynamic effects of failure under static load are otherwise considered. (c) The damage tolerance evaluation of Sec. 23.573(b). [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23-14, 38 FR 31821, Nov. 19, 1973; Amdt. No. 23-45, 58 FR 42163, Aug. 6, 1993; Amdt. 23-48, 61 FR 5147, Feb. 9, 1996$ Sec. 23.572 Metallic wing, empennage, and associated structures. (a) For normal, utility, and acrobatic category airplanes, the strength, detail design, and fabrication of those parts of the airframe structure whose failure would be catastrophic must be evaluated under one of the following unless it is shown that the structure, operating stress level, materials and expected uses are comparable, from a fatigue standpoint, to a similar design that has had extensive satisfactory service experience: (1) A fatigue strength investigation in which the structure is shown by tests, or by analysis supported by test evidence, to be able to withstand the repeated loads of variable magnitude expected in service; or (2) A fail-safe strength investigation in which it is shown by analysis, tests, or both, that catastrophic failure of the structure is not probably after fatigue failure, or obvious partial failure, of a principal structural element, and that the remaining structure is able to withstand a static ultimate load factor of 75 percent of the critical limit load factor at Vc. These loads must be multiplied by a factor of 1.15 unless the dynamic effects of failure under static load are otherwise considered. (3) The damage tolerance evaluation of Sec. 23.573(b). (b) Each evaluation required by this section must-- (1) Include typical loading spectra (e.g. taxi, ground-air-ground cycles, maneuver, gust); (2) Account for any significant effects due to the mutual influence of aerodynamic surfaces; and (3) Consider any significant effects from propeller slipstream loading, and buffet from vortex impingements. [Amdt. 23-7, 34 FR 13090, Aug. 13, 1969, as amended by Amdt. 23-14, 38 FR 31821, Nov. 19, 1973; Amdt. 23-34, 52 FR 1830, Jan. 15, 1987; Amdt. 23-38, 54 FR 39511, Sept. 26, 1989; Amdt. No. 23-45, 58 FR 42163, Aug. 6, 1993; Amdt. 23-48, 61 FR 5147, Feb. 9, 1996$ Sec. 23.573 Damage tolerance and fatigue evaluation of structure. (a) Composite airframe structure. Composite airframe structure must be evaluated under this paragraph instead of Secs. 23.571 and 23.572. The applicant must evaluate the composite airframe structure, the failure of which would result in catastrophic loss of the airplane, i |
