SAE AIR 4566A-2010 Crashworthy Landing Gear Design《防撞起落架设计》.pdf

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5、eaffirmed 2015-11 Superseding AIR4566 Crashworthy Landing Gear Design RATIONALE AIR4566A has been reaffirmed to comply with the SAE five-year review policy. TABLE OF CONTENTS 1.SCOPE 32.REFERENCES 32.1FAA Publications . 32.2EASA Publications 32.3U.S. Government Publications 32.4Other Publications .

6、33.REQUIREMENTS . 43.114 CFR 25.721 (a) 43.2CS 25.721 (a) 44.REASONS FOR CRASHWORTHY LANDING GEAR DESIGN . 55.APPROACHES TO CRASHWORTHY LANDING GEAR . 55.1Fuse Pin Design and Analysis 56.GENERAL DESCRIPTIONS . 66.1Boeing 757 Main Landing Gear 66.2Boeing 737 Classic Main Landing Gear . 106.3Boeing 76

7、7 Main Landing Gear 106.3.1Fuse Pins 106.4Boeing 747 Wing Landing Gear 106.4.1Fuse Pins 106.5Airbus A330/340 Wing Landing Gear . 156.5.1Loading Conditions . 156.6Lockheed L-1011 Main Landing Gear 156.7Bombardier Dash-8 Main Landing Gear Structural Fuse Design . 156.7.1Fuse Pin Design 196.8Saab 340 M

8、ain Landing Gear . 196.8.1Landing Gear Overload . 196.8.2Landing Gear Separation 196.9Bell-Boeing V-22 Main Gear Structural Fuse Design . 226.10McDonnell Douglas AV-8B Structural Fuse Design Data . 226.10.1Construction 226.10.2Drag Failure Strength 226.10.3Side Load Failure 226.11Sikorsky Design 226

9、.11.1Crashworthy Landing Gear History . 226.11.2Energy Absorption Devices . 286.12Boeing Helicopter UTTAS Prototype YUH 109 . 296.13Bell Helicopter YAH-63 . 296.14McDonnell Douglas AH-64 Helicopter 296.15Bombardier Challenger and Regional Jets Main Landing Gear . 316.16NATO Helicopter Industries (NH

10、I) NH90 Tactical Transport Helicopter (TTH) 346.16.1MLG Crashworthiness Concept 346.16.2NLG Crashworthiness concept . 357.NOTES 36FIGURE 1 - BOEING 757 MAIN LANDING GEAR . 7FIGURE 2 BOEING 757 MAIN LANDING GEAR STRUCTURAL FUSES 8FIGURE 3 BOEING 757 MAIN LANDING GEAR STATIC POSITION 9FIGURE 4 - BOEIN

11、G 737 CLASSIC MAIN LANDING GEAR 11FIGURE 5 - BOEING 767 MAIN LANDING GEAR . 12FIGURE 6 - BOEING 747 WING LANDING GEAR 13FIGURE 7 - STRUCTURAL FUSE 747 WING LANDING GEAR 14FIGURE 8 AIRBUS A330/340 WING LANDING GEAR 16FIGURE 9 LOCKHEED L-1011 BREAKAWAY MAIN LANDING GEAR FAILURE SEQUENCE . 17FIGURE 10

12、BOMBARDIER DASH 8-100, -200, -300 SERIES STRUCTURAL FUSE DESIGN 18FIGURE 11 SAAB 340 MAIN LANDING GEAR ASSEMBLY 20FIGURE 12 SAAB 340 MLG DOOR MECHANISM AND UPLOCK INSTALLATION (SHOWN IN “UP“ POSITION) . 21FIGURE 13 BELL-BOEING V-22 MAIN LANDING GEAR STRUCTURAL FUSE 23FIGURE 14 - SIKORSKY AIRCRAFT CR

13、ASHWORTHY GEAR DESIGNS 24FIGURE 15 SIKORSKY UH-60A ENERGY ABSORPTION SYSTEM 26FIGURE 16 SIKORSKY ACAP S-75 LANDING GEAR . 27FIGURE 17 - MCDONNELL DOUGLAS AH-64 HELICOPTER LANDING GEAR . 30FIGURE 18 BOMBARDIER CHALLENGER ARTICULATED MAIN LANDING GEAR . 32FIGURE 19 BOMBARDIER CRJ-700/900/1000 CANTILEV

14、ER MAIN LANDING GEAR . 33FIGURE 20 BOMBARDIER AIRCRAFT MLG TRUNNION FITTING SIDE VIEW . 33FIGURE 21 BOMBARDIER AIRCRAFT MLG TRUNNION FITTING - ISO. 34FIGURE 22 NH90 TTH MAIN LANDING GEAR 35FIGURE 23 NH90 TTH NOSE LANDING GEAR . 36FIGURE 24 NH90 TTH NOSE LANDING GEAR BREAKAWAY . 36TABLE 1 - SIKORSKY

15、CRASHWORTHY LANDING GEARS 25SAE INTERNATIONAL AIR4566A 2 OF 361. SCOPE The intent of this SAE Aerospace Information Report (AIR) is to document the design requirements and approaches for the crashworthy design of aircraft landing gear. This document covers the field of commercial and military airpla

16、nes and helicopters. This summary of crashworthy landing gear design requirements and approaches may be used as a reference for future aircraft. 2. REFERENCES The following publications for a part of this document to the extent specified herein. The latest issue of SAE publications shall apply. The

17、applicable issue of the other publications shall be the issue in effect on the date of the purchase order. In the event of conflict between the text of this document and references cited herein, the text of this document takes precedence. Nothing in this document, however, supersedes applicable laws

18、 and regulations unless a specific exemption has been obtained. 2.1 FAA Publications Available from Federal Aviation Administration, 800 Independence Avenue, SW, Washington, DC 20591, Tel: 866-835-5322, www.faa.gov.14 CFR 23.561, 23.721, 25.561, 25.721, 27.561, 29.561 2.2 EASA Publications Available

19、 from European Aviation Safety Agency, Postfach 10 12 53, D-50452 Koeln, Germany, Tel: +49-221-8999-000, www.easa.eu.intCS 23.561, 23.721, 25.561, 25.721, 27.561, 29.561 2.3 U.S. Government Publications Available from the Document Automation and Production Service (DAPS), Building 4D, 700 Robbins Av

20、enue, Philadelphia, PA 19111-5094, Tel: 215-697-6257, http:/assist.daps.dla.mil/quicksearch/AFSC DH 2-1 Design Note 2A2, paragraph 4 SD-24L, Volume 1, paragraph 3.8.3 MIL-STD-1290 (AV) JSSG-2009, Joint Services Specification Guide, Air Vehicle Subsystems, 30 October 1998, Appendix A, A.3.4.1.3.1.3 2

21、.4 Other Publications Light Fixed-and Rotary-Wing Aircraft Crash Resistance, MIL-STD-1290A (AV) (March 1986), Dept. of Defense, Washington, DC Army Helicopter Crashworthiness - Carper, C. H., Burrows, L. T. and Smith, K. F., Applied Technology Laboratory, US Army Research and Technology Laboratories

22、, Fort Eustis, Virginia 23604. Presented in May 1983 Helicopter Crashworthiness - Fox, R. G., Bell Helicopter Textron Inc. Presented at Flight Safety Foundation Corporate Aviation Safety Seminar (April 1989) Tilt Rotor Crashworthiness - Cronkhite, J. D., Bell Helicopter Textron Inc. and Tanner, A. E

23、., Boeing Vertol Co. Presented at 41st Annual Forum of the American Helicopter Society, Fort Worth (May 1985) SAE INTERNATIONAL AIR4566A 3 OF 36Structural Design of a Crashworthy Landing Gear for the AH-64 Attack Helicopter - McDermott, J. M., Hughes Helicopters Inc. Presented at 38th Annual Forum o

24、f the American Helicopter Society (May 1982) Crashworthiness Versus Cost Based on a Study of Severe Army Helicopter Accidents During 1970 and 1971, Haley, J. L. and Hicks, J. E., U.S. Army. Published in April 1980 issue of Journal of American Helicopter Society Advanced Technology Helicopter Landing

25、 Gear Preliminary Design Investigation - Sen, J. K. Votaw, M. W., Weber, D. C., Hughes Helicopters, Inc. (July 1985), AD-A158 816; USAAVSCOM TR-84-D-20 Helicopter Landing Gear Design and Test Criteria Investigation - David Crist, L. H. Symes, Bell Helicopter Textron Inc. (Aug. 81) USAAVRADCOM-TR-81-

26、D-15 KRASH Analysis Correlation with Full Scale YAH-63 Helicopter Crash Test - Berry, V. L. and 2. For airplanes that have a passenger seating configuration, excluding pilots seats, of 10 seats or more, the spillage of enough fuel from any part of the fuel system to constitute a fire hazard. 3.2 CS

27、25.721 (a) a. The landing gear system must be designed so that when it fails due to overloads during takeoff and landing, the failure mode is not likely to cause spillage of enough fuel to constitute a fire hazard. The overloads must be assumed to act in the upward and aft directions in combination

28、with side loads acting inboard and outboard. In the absence of a more rational analysis, the side loads must be assumed to be up to 20% of the vertical load or 20% of the drag load, whichever is greater. SAE INTERNATIONAL AIR4566A 4 OF 364. REASONS FOR CRASHWORTHY LANDING GEAR DESIGN The reason for

29、crashworthy landing gear design is to contribute to the overall aircraft design goals in the event of a crash. These goals are to prevent occupant fatalities and injuries, damage to the aircraft, and damage or injury to adjacent equipment or persons. On fixed wing commercial and navy aircraft it is

30、a requirement that the landing gear enables the failure modes to be such as to prevent damage, which can cause spillage of fuel sufficient to cause a fire hazard. In cases where centerline gears are used, consideration is given to the amount of damage that can be caused should the centerline gear fa

31、il so as to penetrate the fuselage, causing injury to persons, damage to structure and in the case of body fuel tanks spillage of fuel. For helicopters the landing gear is part of a crash energy attenuation system to limit loads to the occupants and to minimize aircraft damage. The most stringent re

32、quirements are for helicopters designed to MIL-STD-1290 (AV) which requires maintenance of the occupied area and transmittal of noninjurious accelerative loadings to occupants for up to a combined high angle impact of a vertical velocity of 42 ft/s and a longitudinal velocity of 27 ft/s onto a rigid

33、 surface at specified roll and/or pitch angle attitudes. The other components of the energy absorption system are a crushable subfloor structure, a stroking seat, which is primarily a load limiter, and other structural yielding and strain during the crash.The most significant difference in fixed win

34、g and rotary wing crashes is in the forward velocity at impact. The primary concern for fixed wing is to have the landing gear break away from the aircraft without rupturing the fuel tanks. For a rotary wing at near zero forward speed, the “crash“ is excessive vertical velocity and the landing gear

35、should function in its crash mode to absorb energy and limit the “g“ level of the occupants. 5. APPROACHES TO CRASHWORTHY LANDING GEAR The design approaches applied to crashworthy landing gear design include: a. Control of energy absorbed and dissipated by the landing gear so as to limit the loads a

36、pplied to the airframe and its occupantsb. Inclusion of structural fuses and fuse pins to control the mode of failure of the landing gear when excessive loads are appliedFor commercial and some military fixed wing aircraft, these requirements apply during a crash or emergency situation involving the

37、 landing gear during landing, takeoff, taxiing, or ground maneuvering. Also, for military helicopters, to reduce damage, MIL-STD-1290 (AV) states that as a minimum the landing gear shall be capable of decelerating the aircraft with 1 DGW rotor wing lift and from a vertical impact velocity of 20 ft/s

38、 onto a level, rigid surface without allowing the fuselage to contact the ground. Plastic deformation and damage of the landing gear is acceptable; however, the remainder of the aircraft structure should be flightworthy after the impact. The aircraft shall be capable of meeting these criteria in acc

39、idents including a simultaneous fuselage angular alignment of 10 degree roll and +15 to -5 degree pitch. The landing gear shall be designed so that gear failure does not increase danger to occupants, either by gear penetration of the occupied areas or by rupturing flammable fluid containers or by da

40、maging onboard stores such as missiles, rockets, and ammunition. It is desirable that the landing gear continue to absorb energy even after fuselage contact has been made, to maximize the protection afforded by the gear. 5.1 Fuse Pin Design and Analysis The analysis of fuse pins is very complex and

41、must be verified through test as discussed below. Typically, as shown in the following sketches (Figure 6, 8, 10 and 13), fuse pins are notched at the inside diameter at the shear face of the pin to assure a clean failure at the desired location. While the maximum shear stress will drive the failure

42、, the state of stress is complex and is oftentimes analyzed using finite element analysis (FEA) to best resolve the state of stress. Even using FEA techniques this is not a simple solution due to the plasticity of high strength steels. The analyses may need to be iterative by varying the inside diam

43、eter of the pin and the configuration of the notch until the FEA analysis will no longer converge. This approach will assure a relatively constant state of stress through the wall at the notch around the circumference of the pin. SAE INTERNATIONAL AIR4566A 5 OF 36In order to minimize the weight impa

44、ct and to protect the back-up structure, the tolerance on the static shear strength due to material and manufacturing variability of each heat treat lot of fuse pins should be held to a narrow range. It is suggested that this range be between 5% and 10%. The backup structure must then be strength ch

45、ecked to some load factor above the maximum fuse pin strength to ensure that the fuse pin is the weak point and not the backup structure.The value of the backup load factor used depends on the accuracy to predict the actual load of the fuse pins and the accuracy of the structural analysis. On some o

46、lder models, this load factor has been as high as 25% but with current advanced modeling and analytical techniques, 10% is a more reasonable number. In addition to the static stress, the fatigue strength of the pin is normally considered. The notch created to assure static failure could reduce the f

47、use pin fatigue life depending on the geometric configuration of the notch, resulting in a reduced life limit on the pin. Care should be taken to minimize the fatigue notch effects. While the most care and thought may be put into the analysis, the analysis does not account for variation in tensile s

48、trength of the pins from various heat treat lots. Therefore, each heat treat lot is normally tested after heat treat and the final inside diameter of the pin determined by test. For example, it may be required to test three (3) pins after heat treat todetermine the required inside diameter for final machining. After final machining, a number of pins from each heat treat batch should be tested in the final configuration to ensure the proper inside diameter. The number of pins