C-105 STRUCTURE ) Class!licofion cancelled / chan(Jed to ..... \i.l!f.1~.!'!.!'!20.~~ By authority ol... .... !?:Y.g().J,'.:.'.: 0 .:'; .Pe~.~.~.~.~.~f.: .... ~.?.::!E~ ..... Date .... .... .Z.S ... J:.::J ... ~.7... .. .. ......... .. Signature ....... y:;/~ ...... _ ..... 1 ... gCJ::.9.~'!.~EP_~.!'.' 80n Unit I Ronk / Appointm/nt .............. .D..s.:rn ... .3 ......... _._ ... _ C-1O5 ST h UC TUR E THIS REPORT COW: RS IN A BP.OAD MANNER THE MAJOR STRUCTURAL ASPECTS OF THE C-1O5 AIRCRAFT UNDER THE FOLLOWING HEADil'{;S: - 1. LOAD ANALYSIS 2. SHAPES J. IN'IERNAL PHF.SSURES 4. LANDING GEAR 5. HEAT 6. FATIGUE 7. NEW MATERIALS t ,,,,,,,.,.~~ .. 1 ~ 1956 ANALYZED - 1 - INTRODUCTION The aerodynamic requirerrents as applied to the C-105 has re~ulted in a very difficult structure problem, To meet these very rigid geometric requirements and produce a light weight structure has required an almost complete change in the concept of stressinr, and structural arran,·emer,t. There is no place for conservative assumptions to short cut analysis. In fact, in most cases it is impossible to • ~ake conservative assumptiors, for in many cases extra material can be as detrimental to the structure as insufficient material, The work is character isej by rational analysis with close control on distorted shapes. ':,odd An.:.lysis The problem of analysing complex low aspect ratio wings and fuselage conbination must be solved for on a rational basis. The fuselage or centre section can no longer be assumed to be a rigid support but bends under load, thereby inn uenci ng the load distribution to a large extent. Fig, 1 illustrates this point. It is therefore eviden~ that the fuselage wing combination must be analysed as a corr.plete complex redundant structure. The multi-spar, stressed skin arran1errent with a large cut-out for the landing gear was realised to be IJighly redundant an:i an analysis was developed, built around the use of high speed digital computers. It was decided to calculate directly for stresses ard to compute dis ~•lacerrents from known stresses, The wing was considerr-.d to act as· a plate (a two dimensional problem) rather than as a beam (a one-dimmsional problem) as is customary for straight or swept wings with relatively small taper 0 /continued ............. . II 0 .. ~ '/tf, -I- u. '/11/, g <. I0,000 LBjJN 2 \ -- \ \ I \ \ \ \ \ \ \ \ \ \ \ ' ' ' ' ' ' ' ' /. '✓ FIG. 5 . \ 14000 ~ l3bo0 ~~ 1BoO ~ \ 1000 ~ ~ B~o · ~ 11700 ~~ loeoo ~~ 300 ~ \1~~ a..: ~~~4qo: /0~~ 4crR1>. 2 '.oe.rLEc-r10Ns _,,._; Fuse.L"-O..._ s,.,._I 1 _J_ ,_ l..~ t -+-- l·O ·B \NliH FR~ME 14SS + ~ - -,-1- W\TUQ\Ji f'RP-.Ml;. 48 " 1 --+ · p • I I + -1 -r + ~ 2 ~ -0 ::i: 5 2 :::J ~___;::::_----- - --- =------ / I 1,/2 /+\ r '- f J. \,.~/ ~~v I I' ¥ -r / .,.-7 7 DISTORTION OF LANDING GEAR AND F,c.. ID STRUCTURE STN 165 403 RIB 10 - - - - LANDING GEA.R DISPLACEMENT FUNCTION l ·S DISPL~CEMENT FUNCTION TIME (SECONDS) · ICo ·20 2.4- ·2.8 ·32 ·44 FIG H - 8 - estimate the allowable aircraft accelerations at higher Mach, numbere 0 Analytical work as well as flight test probing will be necessary to obtain this goal, Fig, 12 is a graph of ultimate tensile strength plotted against temperature for 75ST6 plate arrl 24ST3 plate. From this curve it is clear that 75ST6 material for a given aircraft weight will allow considerably higher aircraft accelerations for all flight conditions having a temperature of 270° For lower. Above this temperature I 24ST3 will allow higher aircraft accelerations, It has been agreed that even with possible aircraft performance ' improvements, the great majority of flying hours will be at flight conditions below 270° F. For this reason 75ST6 has been used extensively. In order to safely probe for higher temperature conditions the effects of transient temperature conditions as well as creep will have to be carefully analysed. The determination of transient temperatures and thermal stresses in ~ings proceeds as follows. The aeroplane is assumed to fly initially at M = 1.0 which, in the stratosphere, produces an equilibrium temperature of 0° F approximately, One of the conditions which cause the highest thermal stresses is a dive in which the aeroplane accelerates to M • 2,0 within 34.5 seconds (1.1 G acceler~tion) ' and continues at that speed, The adiabatic wall temperature is 2500 F, Assuming a heat transfer coefficient h =• Btu/hr. ft. 2° F from boundary layer to skin, neglecting the effect of finite conductivity of tre skin, and assuming a fairly good ther~~ bond between skin md web (h • 1000), curves of temperature variation / conti nued ..... ... . . .. • • - 9 - • with time at various points of the web and curves of temperature dis tribution in the web at various times have been calculated (Fig. 13 & 14) . The maximum thermal stress occurs when the temperature drop in the structure is maximum, i.e. at 178 seconds after beginning of acceleration. The maximum tt:iermal stress is approximately 26,000 psi (tensile) at the web centre line. Manufacture of test specimen is under way, EYtensive testing will be done using a typical wing box with internal fuel and pressures. The flew .,,r•irv ' • • ... of heat and the stresses will be measured with the box under varying applied external loads. Heat will be applied under a very close time control. Fatigue The presence of heat and high internal pressures aggravated the fatigue problem considerably. It is generally agreed that one cannot afford a ·weight penalty to improve the fatigue life of a fighter aircraft. It is however, more important than ever that the best fatigue life to weight ratio be obtained and that the structure has a uniformity of fatigue life, i.e. there should be no point where the stress concPntration factors greatly exceed the average factor. With this philosphy in mind, specimen testing of joints, fittings etc. has been inaugurated in the early design stage of the aircraft. A considerable amount of development testing was found necessary and the results incorporated into the aircraft. A particularly interesting example is in relation to joints in 1/Jick wing skins. A low aspect ratio wing is subject to quite high chordwise berrling stresses. The typical spanwise la~ joint proved to be particularly bad from a fatigue point of view. The offset of load from the skin to the joint plate causes very high local stresses in both plates. /continued ........••••• • • OF TRA.NSIENT TEMPERATURES IN SKIN A.ND SPAR WEB . No F'UEL. fv = 40 BtU,/h..~ ft2 °F -h, 1 = IOQQ II II ,--- ---- 1---2-so BOUND L,e,...YER,.. I I I ~ \ \\ t L \ \ \ \ \ -, \ \ \ \ \ \ \ \ \ i--- ---i-se I ~- ,- UJ ~ :::i I- <{ ~ lu ft:--tea ~ UJ I- ---se \ - __ "£. ____ , I HEP.: INPL)T X l;, -8" -+~~tl:u~,...:_!-IGG,GS'" ' I I -j- ;+ I - -, - - \ \ \ SKIN ,__ ____ -'-----+--- --+-- WEB FIG . 13. IN -----f <&. 1 WEB/ I - 1 • • 10 - Fig. 15 shows the result of this investigation, Type "A" is a normal lap joint. Type "B" uses a high heat treat steel strap just below the skin. This method did not improve the offset but did put the steel in the high stressed area. Since steel has a much hi~her endurance limit than aluminum there was a very significant improvement in the fatigue life of the joint. Type "C" is however the optimum arrangement. A very thin steel strap is used externally by machining a step in the skin. This arrangement reduces the offset to a minimum and provides very satisfactory fatigue characteristics. From these S.N. curves, design allowables were obtained which when compared to the actual loads, dictated the type of joint required. In fact all three types of joints are used in the wing design • The decision to do component fatigue testing will be made folladng static testing. If the load distribution is found to be satisfactory arrl if the specimen test results are satisfactory, expensive component fatigue testing may not be justified. However, if high concentration of stresses do occur in certain areas, it may be necessary to fatigue test that particular component. (Load distribution will be detennined from the static test by means of a multitude of strain gauges). At the present, good fatigue life characteristics are dependent on good load analysis and specimen fatigue testing. New Materials New alloys are being used extensively to improve performance and safety :. of the aircraft. The following lists four relatively new alloys being used. New and better materials will be vigorously pursuedo /continued • •·•••• •· ••· ' - 11 - New Materials L Magnesium The large area fuselage an:! relatively low loads allowed the advantageous use of a low density high strength material. Magnesium was the obvious choice. High allowables at elevated temperatures have been obtained and small production orders have been placed. Very small rivets are used to eliminate the need for dimpling. This combination produces an excellent aerodynamic surface and with a much higher point of initial buckling. Buckling occurs in the magnesium sheets at 2,75 g. Fig. 15 is a photograph of a magnesium panel under high shear loads producing tension field effects. Proi:;-er precautions are taken against electrolytic action, Fig. 16 shows the general use of magnesium on the C-105 aircraft. 2. Aluminum Forgings The necessary use of high s t r ength aluminum forgings has produced a serious problem. High strength aluminum forgings have particularly poor elongation properties in the short transverse grain as well as high residual stresses. The large scale use of X?079 ST has improved this situation considerably. This material has a guaranteed minimum elongation in the short transverse grain direction of 4% with the average elongation between 6% and 7%. This is considerably better than the 1% guaranteed for the 7075ST6. Howeve'r, careful control of stresses in the short transverse grain is still required. 3. High Heat Treat Steels Steels in the heat treat range between 240,000 psi and 260,000 psi are being used, Space requirements and weight requirements have /continued ........•. - 12 - New Materials (contd.) dictated the use of this material in some areas. Extensive research testing is under way to develop satisfactory manufacturing processes, t 4. Titanium Titanium is used in engine compartments where temperatures require something better than aluminum. In these cases titanium is considerably lighter than steel. • ~ 1800 IGOO 1400 1200 LCP..D 1000 1 PER BOLT [ LB] 800 600 400 200 0 10 ' • REPE-"'TEO LC,.._OING DF 1/4" DI..._ , N f:>-.5 c 'sK BOLTED JOINTS \ SPECIMEN .. ,,:_ t:7 ._--, f-=--i ............ \ _j_ SPECIMEN.B ·Oio2 R?[ 'n ! ONTCRS.L Sra>L S>RJ " ' +- C T ·\~. ', LE GEND' - A P..,/:>,,. B BB :- C :- cc :- , SPECIMEN•;,; (al 10° F • " 2so°F "B' • 70°F • • 2so' F ·c· I 10°F , • 2so'F \ ~ ,__ c:.c /:>,,. ~~~ • -~ ----....::::.. .... ___ ' < ~ --':::---,. , I '· - i"-- t---...... SPE.CIMEN° c" S031~ E:T_:~NP..L STE:L STR"'P I •- -.___ " .' ., , ; < ~ . ~ H--- ---------, . BB ' 101 10• 10 4 - t 1 NUMBER OF LOl>.DING CYCLES FIG. 15 10' ~-·---~ ~ ~--- - . ~ ~ ! ... - ~ - 10• 10' A Y R 0 C A N A D A CF-105 USE OF MAGNESIUM SI AVRO AIRCRAFT SPECIFICATION M-7-6 PROTECTIVE TREATMENT - I 4 EORl! I I I I I I I I I ~'~ F-105 USE OF MAGNESIUM SHEET ZH-62 CIFICATION M-7-6 PROTECTIVE TREATMENT - DOW 17 AND HEAT RESISTING ENAMEL C:::=J VISIBLE } .OAO IN. MAGNESIUM CJHIDDEN [==:J VISIBLE } .051 IN. MAGNESIUM c_]HIDDEN · Ei@fMH .032 IN. MAGNESIUM C=3 .102 IN . MAGNESIUM ~~ I ( ~~ J} ,t r 11 , l / Clossificolion caiWJed / chani ed to . .. Y.n.C:.J.!J,.S..s..:l.f.i.e.d. By aut~ l/~ ~1-· .. 8 7 . . ?. .. . ~r..r..~."!. ...... :.~.~.~~.8. \ ~.: ..... B..?..~.Ig··••U•· Date ·········· ····~···•······............. . .••••.•••.•• • ~ Co-Cha · Signature....... . .. . . . .... . ... t .·-··········-···-l-.1:Il.e.r.s on Unit / Rank / Appointme .......... ..... P..~.I~ ... ) .... ,-···--····· ....... ~. ---,,,- 7 -r ~ -,,-

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