1、IMIL-A-8866C(AS)20 May 1987SUPERSEDINGMIL-A-8866(ASG)18 May 1960(See 6.4)IMILITARY SPECIFICATIONAIRPLANE STRENGTH AND RIGIOITY RELIABILITYREQUIREMENTS, REPEATED LOADS, FATIGUE AND DAMAGE TOLERANCEThis specification is approved for use within the NavalAir Systems Command, Department of the Navy, and
2、isavailable for use by alI Departments and Agencies of theDepartment of Defense.1. SCOPE1.1 Scope. This specification defines the strength and rigidityrequirements for repeated loading condition applicable to Navy procuredairplanes. -2. APPLICABLE DOCUMENTS ”K.2.1 Government documents.2.1.1 Specific
3、ations and Standards: The following specifications form apart of this specification to the extent specified herein. Unless otherwisespecified, the issues of these documents shall be those listed in the issue ofthe Department of Oefense Index of Specifications and Standards (DDDISS) andsupplement the
4、reto, cited in the solicitation.SPECIFICATIONSMILITARYHIL-D-B708 - Demonstration Requirements for Airplanes.HIL-A-B860 - Airplane Strength and Rigidity, GeneralSpecification for.NIL-A-8861 - Airplane Strength and Rigidity F1ight Loads.MIL-A-8863 - Airplane Strength and Rigldity Ground Loads forNavy
5、Procured Airplanes.HIL-A-B867 - Airplane Strength and Rigidity Ground Tests.Beneficial comments (recommendations,additions, deletions) and any pertinentdata which may be of use in improving this document should be addressed to:Naval Air Engineering Center, Systems Engineeringand StandardizationDepar
6、tment (Code 93), Lakehurst. NJ 08733-5100, by usin9 the self-addressedStandardization OocurnentImprovement Proposal (DD Form 1426) appearing at theend of this document, or by letter.AMSC N/A FSC 1510DISTRIBUTION STATEMENT A. Approved for public release; distribution is unlimited.Provided by IHSNot f
7、or ResaleNo reproduction or networking permitted without license from IHS-,-,-MIL-A-8866C(AS)SPECIFICATIONSMILITARY (centd)MIL-A-8868 -MIL-A-8870 -MIL-L-22589 -STANDARDMILITARYMIL-STD-2066 -,-Airplane Strength and Rigidlty Data and Reports.Airplane Strength and Rigidity, Vbratton,Flutter, and Diverg
8、ence.Launching System, Nose Gear Type, Aircraft.Catapulting and Arresting Gear Forcing Functions forAircraft Structural Design.(Copiesof specificationsand other Government documents (publications)required by contractors in connection with specific acquisition functionsshould be obtained from the con
9、tracting actlvlty or as directed by thecontractingofficer.)2.1.2 Other publications. The fol16wing document forms a part of thisspecificationto the extent specifiedh%reln. Unless otherwise specified, theIssuesof the document which are DOD adopted shall be those listed in the issueof the CXIDISSspeci
10、fied in the solicitation. Unless otherwise specified, theIssuesof documents not 1isted in the 00DISS shal1 be the issue of the non-governmentdocuments which 1s current on the date of the sollcitatlon.AMERICAN SOCIETY FOR TESTING AND K4TERIALS (ASTM)ASTM Test Method E399-83 - Test for Plane Strain Fr
11、acture Toughness ofMetalllc Materials.(Applicationfor copies of ASTM publications should be addressed to theAmerican Society for Testing and Materials, 1916 Race Street, Philadelphia,Pennsylvania 19103.)2.2 Order of precedence. In the event of a conflict between the text of thisspecificationand the
12、references cited herein (except for associated detai1specifications,specification sheets or MS standards), the text of this speci-fIcation shal1 take precedence. Nothing in this specification, however, shallsupersedeapplicable laws and regulations unless a specific exemption has beenobtained.3. REQU
13、IREMENTS3.1 General. The structural design of the airplane shall be such thatrepeated loads shal1 not cause failure or permanent deformation of any part ofthe airplane, interferewith its mechanical operation, or affect 1ts aerodynamiccharacteristics. Further, the design shall not require repair, ins
14、pection, orreplacementof components other than as specificallY approved by the contractingactivity. The above requirements apply to the planned service 1Ifeoftheairplane for the repeated loads environment resultlng from ground and f1lght2L,Provided by IHSNot for ResaleNo reproduction or networking p
15、ermitted without license from IHS-,-,-.MIL-A-8B66C(AS)operations. lncludin9 loads and load cqnbinatlons associated w!th maneuvers,field and carrier arrested landings , gusts, buffeting, dynamic response,pressurization (fuel and cockpit), aeroacoustlcs, vibration, store Installationand release, catap
16、ulting, taxilng, operation of devices, and exposure to achemical or thermal environment.3.2 Service life. The service life of the airplane shall not be less thanthat specified by the contractIng actlvlty in terms of the followlng (asapplicable):a. F1ight hoursb. XGround-air-ground cycle.(flights)Fie
17、ld taxi runs: F1eld takeoffse. Catapult launchesf. Landings(1) Field(2) FCLP (Field Carrier Landing Practice)(3) Carrier arrested(4) Carrier touch-and-go3.3 Fatigue. -.L3.3.1 Spectra. The airplane usage spectra for analysis and test shallInclude repeated loads from al1 types of ground and f1lght ope
18、rations as noted in3.1 and shall be supplemented, as required, to ensure that each airplane compon-ent is designed and tested to the proper repeated loadings. Consideration shallbe given to the effects of load sequence, load truncation, load induced residualstress, and other factors as appropriate t
19、o assure that usage spectra foranalysis and test provide the most conservative fatigue 1lfe. Ordering andfrequency of loads W1thin the usage spectra shal1 be random, conslstent withflight-by-flightairplane operation, load exceedance and occurrence rates, andplanned service llfe values.3.3.2 Complian
20、ce. Compliance with 3.1 shal1 be demonstrated by the fatigueanalysls of MIL-A-8868 and the fatjgue tests of MIL-A-8867 utl1izing crackInitiationas the primary failure crlterlon. Specifically, the structural designof the airplane shalI be such that the usage spectra will not cause structuraldefects (
21、cracks, deformations, loss of modulus, delaminatlons, dlsbonds, etc.) orfailure, within 4 times service life based upon analysis and 2 times service lifebased upon full scale tests. If any part of the airplane should fai1 to demon-strate compliance WIth the above requirements, that part shal1 be red
22、esigned andthen shown by analysis and test to be compliant. No inspections shalI be re-quired as a function of design within two service lifetimes. For both fatigueanalysis and tests, the use of fatigue 1Ife-enhancingmechanical processes (suchas shot peenlng, roller burnlshlng, etc.), other than spl
23、it sleeve cold workingand interferencefit, are prohiblted 1n demonstrating compliance.3.3.2.1 Anal.vsis. For analysls purposes, substantiationof fatigue lifeshal1 be fn accordance With prediction methods as approved by the contractingactivity. For approved interferencefit and/or cold working enhance
24、ments, fatigueanalysis shal1 tndlcate the airplane Wi 11 be free from structural defect for atleast 1 service 1ife WIthout the benef1t of Interferencefit and/or cold working.3Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-MIL-A-8866C(AS) 3.3.2.2 . T
25、esting shall continue beyond 2 times service life untilcatastrophic failure, or until4 times service life has been sustained by thetest article. At test conclusion, the test article shall be subjected to acomplete destructive teardown inspection, including fractographlc examination, totdentlfy all f
26、ailures. All repairs made to the test article prior to 2 timesservice llfe, and alI repairs for cracks or failures concluded to have beeninitiated prior to 2 times service Iife, shall be incorporated into all forward-and retrofit-production airplanes, provided the repairs”have denmstratedcompliance
27、by analysis and test. For fatigue life certification of compositestructure by test, accounting for variability and environmental factors shall berelatable.to and confirmable by the results of the design development pre-production component and static tests of MIL-A-8867.3.3.3 Low frequency vibratory
28、 loads. The airframe structure shal1 haveunlImlted 1ife due to low frequency vibratory loadings. When these low frequencyvibratory loadlngs are combined with the other various airplane loadlng condi-tions (i.e., PU1lUPS, banks, high angle of attack (AOA), gusts, etc.), thevibratory loadinqs shal1 no
29、t cause the stuctural fatique life to be dearaded fromthat whlh resuls when separately applying the other-various loadlng onditlonsto the airframe structure.,-.-3.4 Damage tolerance. The design.aj.constructlon of the alrframe structure,and the selection of materials to be used shall include provisio
30、n for damagetolerance. Damage tolerant material shall be chosen on the basis of available,confident data. Damage tolerance shall be in addition to, rather than in lleuof, provision for adequate structural fatigue characterlstlcs, and shall serve asa means for preventing catastrophic structural failu
31、re loss of control of theaircraft after a predefined 1imit of structural damage has occurred.3.4.1 Analysis. Al1 areas of structural components established as primaryor critical shall be analyzed using the methods of linear elastic fracturemechanics, as a mlntmum, to tdentify the character and dimen
32、sions of defectswhich could grow to critical size in time periods as limited by aircraftinspection periods or wing required lifetime, as applicable. These analysesshal1 assume the presence of crack-like defects and/or delamlnations placed Inthe most unfavorable orientation WIth respect to the materi
33、al properties andapplied stress consistent with the aircraft loads environment, and shall predictthe growth behavior of the chemical, thermal, and sustained and repeated-loadsenvironment to which the component will be subjected.3.4.2 Compliance.3.4.2.1 Metals. For all primary or critical structures,
34、 crack growth undersustainedand repeated loads shall not occur at a rate such that initial flawscan reach critical size at the residual strength requirement load in one lifetimeof expected service usage. For purposes of these analyses, the lnitlal flaw sizeshall be .010 inches minimum in metals. and
35、 at failure not smaller than 0.25inches (surface length). Critical flaw sizes shall be determined using theappropriate critical fracture toughness values determined on avalidstatisticalbasls in accordance with the procedures of the American Society for Testing andMaterials Test Method E399-83 entlti
36、ed “Test for Plane Strain Fracture Toughness L4Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-MIL-A-8866C(AS)of Metallic Materials.” The analysls shall identify plane strain, planestress, or mixed mode conditions at the onset of rapid crack propagat
37、ion, andshall include all crack growth rate and critical crack length data on whichthe analysis was based. The effect of sheet thickness on fracture resistanceshall be proposed by the contractor and accepted by the contracting activity.3.4.2.2 Composites. Damaged (e.g., delamination) structure, at o
38、r belowthe threshold of being clearly visible, shall be capable of fully compensatedultimate load statically with no damage growth under sustained and repeatedloads for oneservice 1ifetime . Demonstration of requirement shalI be bytest, or as proposed by the contractor and speclfically approved by t
39、hecontracting activity, accounting for the effects of variability andenvironment as relatable to and confirmable by the results of the designdevelopment, pre-production component and static tests of MIL-A-8867.3.5 Maneuver loads. For determination of loads, the airplane shall be atthe critical speed
40、 and altitude which results in the minimum life on thecomponent being considered. The spectrum of loads shall include symmetricpull-ups and push-overs, and asymetrlc rol1ing pull-outs, rolI reversals, andlevel rolls. Except for level rolls, the maneuvers shal1 have the number ofpositive and negative
41、 exceedances.ofvertical load factor as specified by thecontracting activity. The percentageof asymmetric maneuvers at each loadlevel as well as the total numbe “6flevel rol1s shal1 be proposed by thecontractor and approved by the contracting activlty.3.6 Gust loads. The gust ,loadspectra shall encom
42、pass the anticipatedmlsslon usage of the airplane and be approved by the contracting activity.The usage shall include at least 4 percent of the airplanes life at V. atsea level. The fatique analysis shalI include at least the dynamic responsein the rigid-body modes of pitch and translation, for elas
43、tic modes asappropriate to the structural character stlcs and configuration, and the modesof any autopilot or artificial stabiIIty devices. .The dynamic response shallbe determined for the power spectral density of atmospheric turbulence inaccordance WIth MIL-A-8861,continuous turbulence aproach of
44、3.5.2.3.7 Ground loads.3.7.1 Catapult takeoff. The weight shall be the maximum design grossweight. The off-center spotting eccentricity of the airplane at release shallbe half of that resulting from the criteria of MIL-L-22589. Half of thetakeoffs shal1 be to the right and half to the left. For each
45、 takeoff thesequence and magnttude of load application shal1 be as follows:a. Two cycles of buffing load. Each cycle shalI be from zero to 80percent of the release load of MIL-A-8863. At the end of each cyclethe load shal1 drop to zero.b. Tenslonlng load of MIL-A-8863.Provided by IHSNot for ResaleNo
46、 reproduction or networking permitted without license from IHS-,-,-MIL-A-8866C(AS)c. Release load lncludlng the effects off-center spotting.d. External and Internal loads that result during the catapult run, Includlngbending moments that result from eccentricities of the launch bar withrespect to th
47、e shuttle as welI as loads due to tire and steering frictionin off-center takeoffs. The applied tow force shal1 be the maximumcatapult tow force of MIL-A-8863, except that for ten percent of thetakeoffs the tow force shal1 be equal to the upper 90-90 envelope boundaryof MIL-STO-2066.3.7.2 Landings.
48、For all landings the applicable design gross weight shallapply.3.7.2.1 Landing impact loads. The alrplanerational combination of sinking speeds of Tableattitudes (pitch and roll) and engaging speeds,landinq loads. For loads on the nose sears forloads shall be determined by aI with variations in land
49、ingincludinq arrestina and driftcarrier rrested lndings,tenth ;f all landings at each sinking peed shall be considered free-flightoneengagements unless it can be shown by analysis that free-fIight engagementscannot occur within the specified distribution of landng conditions.I.!-3.7.2.2 Arresting loads. The horizdntai component of the arresting hookforce shall be equal to the limit load%; 90 percent of the arrestments. For 10percent of the arrestmen
