Structural Testing of Homebuilt Aircraft

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STRUCTURAL TESTINGOF HOMEBUILTSNote: Alex Strojnik'shavein Sport Aviation manyin the past decade. A nativeY u g o s l a v i a , Alex has veryacademic credentials.degree in electrical engi-a Ph. D. in aerodynamics . .

Arizonaas a physic ist ! Though itbecame his profession, Alexcontinued to carry a torch forthroughout his life. Whilea young engineer and gliderY u g o s l a v i a , heand built an all-wood, tail-glider, the S-1 . . . the crash ofhe fortunately survived.into physics by his gov-t, he was not able to getpersonal flying until comingultimately discover-homebuilding.by Bruce Carmichael'sand September 1976 Sport

Aviation articles on laminar flow inlightplane design, Alex designedand built a very low drag poweredsailplane, the S-2 (Sport Aviation,April 1982), which would becomethe first homebuilt motorglider inwhich International FAI Silver, Goldand Diamond badges would beearned. More recently, he hasdesigned and built the S-4 LaminarMagic (Sport Aviation, January1990), a tiny 30 h. p. machine thatheld the world's Class C-1.A/O 3kilometer speed record for a time.Alex is also the author of severalbooks on the design of aircraft withlaminar flow characteristics (seehis classified ad in this issue under"Books/Films, etc.").As a veteran EAA TechnicalCounselor, Alex often recommendsthat builders load test their home-builts before f lying them. Hebelieves this is advisable whenevermodifications have been made toan existing design, and, of course,

in all cases of new designs. Healso believes load testing may be inorder in a number of instancesinvolving composite airframes.While there has been no history ofstructural failure in compositehomebuilts that have been con-structed according to thedesigner's instructions . . . andwhile designers of composite air-craft normally make allowances forbuilder variances, still there may bethose who have a nagging uneasi-ness about the integrity of thestructures they have built. To thosepeople, Alex says that load testingtheir a i r f rames is so stra ight for-ward that there is little reason notto do i t . . . if for no other reasonthan for peace of mind. In his arti-cle, he tells builders of compositeaircraft why a load test is desirable,how to conduct it and, very impor-tantly, how to intrepret the results .. . some of which may be both unex-pected and quite surprising.

tructural proof loading, whilet supplying al l answers to allnevertheless tell thea great deal about thestructural integrityhis aircraft. It will also contribuepeace of mind.Let us start by pausing briefly topeculiar mechanical prop-of composite materials. Figure 1the relationship between thestress the material is subjectedrelativeor strain for several materi-aircraft construction.is the stretching load per unitsection in pounds/square inchwith 1,000 psi = 1strain is the change in length of aof a specimen of constant cross1 inch long, usuallyin inch/inch or in percent.)kinds of car-composite (HM, HT), of Kevlar 49is nicely "straight", meaning

in proportion with the stresspointat the ends of respective

By ALEX STROJNIKEAA 610062337 E. Manhattan Dr.Tempe, AZ 85282

straight lines). Twice the stress, twicethe strain. The behavio r of thestretched aluminum alloy (20 24 -T3),spruce, or 4130 steel (not shown) is,how ev er , quite different. Take, forexample, aluminum. As the tensileforce is applied, the specimen (forexample, in the testing machine) ini-tially elongates in proportion with thestress. Here the aluminum obeys this" law of proportionality" quite well.However, at a certain point - let uscall it a "yield point" - on thestress/strain line, the original nice lin-earity ends. In Figure 1 this occurs atan approximate stress of 42 ksi (=42,000 psi). Any further stressincrease results in a disproportionallylarge increase in strain. Aluminumbegins to "yield" - an experiencefamiliar to all homebuilders, by theway - with large strain increases forrelatively small stress increases.Eventually, somewhere around 64 ksi(64,000 psi) this specimen will fracture

in tension - but not until the strainreaches a huge value of some 0.12 or12%. This value of the strain is so farout to the right it does not even showin Figure 1. Similar behavior can beobserved on a tensile specimen of acertified spruce (yield point around5.3 ksi, ultimate strength about 9.4ksi), or a low carbon steel, or CrMosteel, among others.Experiments in the testing device,as well as our own daily experience,show that as long as the material isstressed to less than its "yield point",it has the ability to return to almostits original shape, as soon as theload disappears. (To be precise, thiselasticity of the material does nothave its upper limit at exactly theyield point; the difference, however,at least for materials shown in Figure1, is too small to be considered anyfurther.) This "elastic limit" is a veryimportant point. All structural parts

of an aircraft (or the kitchen table, orthe chair, or the children's swing, orthe bridge, or. . .) must be designedwell under this limit. No structuralpart may ever, as long as it is usedwithin its design purpose, be stressedSPORT AVIATION 33


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Figure 2L Figure 2R

of the fuselage.designer or the plans vendor usu-indicates either in the technicaliption or right in the plans to"load f ac to r " he designed hisUnless explained otherwise,factor refers to the "limit"as expected during the normalof the aircraft when pulling

speed. The FAA hasload factors as follows:(normal category), 4.4 (utility), 5.3motorgliders), and 6aerobatic aircraft. The designer ofusually goes higher thanload factors, real iz ing themay not have the quali-of a supervised aircraft factorymust also be understoodthese (minimum!) load f ac to r sto the entire aircraft, not just theDuring that sudden pull-up, thegenerates a lift that is by a "loadlarger than the weight of theresult is that the reactionthis suddenly increased lifting forcepilot onto his seat withamount equal to that "load factor"his weight. A pilot weighing 200pulling up at n = 4.4 will pressthe seat at a "weight" of 880 Ibs.

pound battery will "weigh" 132the 180 pound engine will reactsuddenly weighs 792 Ibs. Hasconstructed the pilot'sfor continuously and repeatedlyelastically, an 880 lb. pilot?find someanswers.Will the seat of the aircraf t to beelastically hold 880 lbs.?w many builders will allow theto load thewith 880 Ibs.(each!)? Maybe ato the designer/plans vendor canthe matter. Maybe only thehas been designed for n = 4.4the rest of the aircraft remains atmercy of n = 2 or 3. These and

the builder, the de-

signer/plans vendor and the EAATechnical Counselor before the struc-tural testing begins. It should also bementioned that, while most oftenquoted, the maneuvering load factoris not the only load factor important inflight. Just as important, especially insailplanes, which must seek vert icalgusts, is the so-called gust load fac-tor. In a majority of light amateur builtsport planes, the gust load factor is,however, often less than the maneu-vering load factor. In any case, thedesigner/plans vendor can easily clar-ify the situation.

TESTING THE WINGExperience shows that the part ofan experimental aircraft most likely tofail structurally is the wing. While anengineering type homebuilder mightconsider an extens ive, detailed andcomplex wing testing project, a greatdeal can be accomplished with verysimple means. A structural testing ofthe wing as will be described hererequires so little money, time and hotcof fee that we will really have diff i-culty finding an excuse for NOTperforming it.The complete wing - meaning with

ailerons and flaps firmly installed - willbe tested inverted. If the wing is com-posed of two or more parts, we firstinspect the spar junctions. In a can-tilever wing we pay special attentionto metal fittings joining the spar-sparor spar-fuselage and we concentrateon the bolts and main pins. Later,when the wing loading is over, we willagain inspect these parts to find out ifthere have been some changes intheir appearance (hole elongations?bolt/nut looseness?). Figure 2L showsa typical wooden spar central fitting(please note the bolts are NOT situ-ated along one single line), and Figure2R a metal spar (Monerai). During thewing bending, these fittings and boltswill be subject to particularly highstresses - this is why FAA prescribes

Additional Safety Factors fo r them.Even so, fractures often appear in thisarea (did the designer forget aboutthese Additional Safety Factors?).During the proof loading, we will care-fully LISTEN to possible "crackling"sounds close to these junctions.Many modern two-part wings use theso-called fork-and-tongue spar junc-tions. Figure 3(a) shows th e metal(French Cricket) and 3(b) the compos-ite (wood?) solution. Wings of thiskind should preferably be tested whileinstalled in the fuselage because theanti-torsional fittings and pins, Figure3(c), and holes require a tight fit and acertain degree of f reedom in move-ment, which only the fuselage itselfcan guarantee. The arrow in Figure 3and an X indicate the area where thetop spar flange (cap) is often crusheddue to insuff icient attention to localstresses. Experienced designers liketo over-dimension this area of themain spar (and the local reinforcedrib). Othe rs become experienceddesigners after building - and testing -two or three wings. The one-piecewing experiences the highest bendingmoment at the point of the engage-ment of its main fitting with thecorresponding fitting in the fuselagewa lls. It is worth remembering thatthis highest bending moment remainsconstant throughout the cross-sec-tion of the fuselage. Swept-backwings should also be tested whilef irmly installed in the fuselage, as itmay be difficult to construct a simple,yet reliable fixture for such wings. Inan y case, the wing supporting fixtureshould provide a sturdy support notjust for the spar bending test but alsofor the following torsional test.Figure 4 shows, schematically,how a fixture for a light aircraft couldlook. A square box made of house-building quality plywood, some 3/4"thick, fastened together at the cornersby 2" x 2" boards, using Elmer's car-penter's glue and some nails to speedit up, and a couple of boards, say 3/4"

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Figure 3A

Figure 3B

TORSIONANTI-TORSIONFITTING

REINFORCEDROOT RIB

ANTI-TORSION FITTING

Figure 3Cx 6-8" extending way back (you canuse the same 3/4"plywood with grainin proper direction) and glued thesame way to the box is all we need.We join the boards at the far (back ofthe aircraft) end,say 5 ft. behind thebox, if these boards are 8 ft. long,with another board (and a couple of2" x 2" boards) in order to providespace for additional loads we willapply when testing the wing for tor-sional stiffness, as indicated by thetwo vertical arrows. It makes a lot ofsense if we select the distancebetween the two front vertical 2" x 2"boards such that they can serve assupports for the fittings which will36 MARCH 1992

accept the wing central fittings. Thisdistance is indicated by an X in Figure4. Using the same reasoning we mayselect the length of the box in such away that the two back 2" x 2" and theback plywood supports the rearattach points of the wing. The buildercan, of course, use his own ideas inconstructing the wing supporting fix-ture, as long as he keeps in mind thatthe wing may be severely damaged ifthe fixture collapses.Fabric covered wings with internalwire bracing must be tested after thefabric has been applied. The diagonalwires - or equivalent - must be ten-sioned as per instructions and never

Figure 4touched during the wing testing. Thebuilder is, howeve r , urged to takenotes of any looseness in these wiresas it may appear during the loading.Furthermore, those internally bracedwings must be simultaneously testedfor their resistance against the airdrag. FAAsuggests that the wing beinclined some 10-13 degrees nosedown (when inverted) depending onthe airfoil characteristics (Figure 5). Ifno information is available from thedesigner/plans vendor, 12 degrees isa good starting point. Fiberglass/ply-wood covered wings do not need thatinclination.In order to protect the wing and theobservers, two aux i l ia ry supports(jacks?) should be positioned close tothe wing tips well below the bottom ofthe wing. This is to catch the wing ifsomething goes wrong - i.e., sliding ofthe loading bags, unsymmetricalloads, failure of the spar.When choosing the height of thewing supporting f i x tu re , it is worthremembering that an aluminum orcarbon spar bends little, a woodenspar quite a bit more and a fiberglassspar - unless designed for stiffnessand not for strength - very much.Sometimes the central, short partof the wing is built as a permanent,integral part of the fuselage. In such acase the fuselage itself, just as in theaforementioned case of a swept-backwing or a tongue-and-fork main sparwing, will have to act as the wing sup-porting fixture. Remembering thatforce acting on the fixture/fuselage islikely to amount to thousands ofpounds, we may ask the designer todesignate proper points of support ofthe fuselage. A strong bulkhead is

Figure 5


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6avai lable in that area of thebeing part of the structuret rans fers wing forces into theWe do not want to unnec-this bulkhead, so weseek the designer's advice.complications insituation is to support the wingoutside the fuse lage at thewhere the wing "enters" the(Figure 6). We realize this ist exactly the point of the maxi-bending moment in real flight;if we are careful the errorbe small. The thing we have toabout is a gentle distribu-of those thousands of poundsa tiny area of contact betweensupporting fixture.contact area is, of course, thatspar flange - it isw the bottom flange - and thefixture. Since we havepossible to thewe may have a problemcontact area largeaccept the heavy con-load without pressingthe already heavily compres-loaded flange. Not surprisingly

sparat this very point as it cannotthe (bending) compres-AND the compressive bearingby the fixture, a 1/4" thickgood plywood properly

placed may alleviate this problem.No metal plates, please, and defi-nitely no foam here!One final thought - if our aircraftis heavy, really heavy, and/or thefloor is not very rigid, and/or the fix-ture is not very stiff, we mayexperience difficulties in measuringwing deflection during the proofloading. We can, of course, performthe proof loading of the wing with-out simultaneously measuring wingdeflection at, say, half wing spanand at the tip - but hanging those 5wooden yardsticks (20 cents each)surely does not represent such a bigeffort. A simple way to make mea-surements independent of outsidedisturbances is to construct a sim-ple double triangle (Figure 7) madeof 1x1 inch wooden sticks, heldtogether with plywood and carpen-ter's glue and fasten it rigidly in thecenter of the wing in such a waythat measuring sticks freely hangdown the wing leading edge. Now,with the unpleasant part of our pro-ject behind us, we can sit back,enjoy another cup of coffee and dosome simple calculations.DISTRIBUTING THE LOADS

We must load our wing in such away that our load distribution equalsthe lift distribution along the wing

span at a high angle of attack (AOA)as experienced for example duringthat powerful pull up. As the top ofFigure 8 shows, the lift distributionon reasonably "normal" wings andcanards somewhat resembles a halfellipse. Different wing planformsgenerate slightly differing lift distribu-tions, as Figure 8 (Table 1) shows,however, the differences are surpris-ingly small - unless one chooses atriangular wing planform (a no-no!) ora strongly swept wing. The Tableshows lift distribution for 3 differentaspect ratios (AR = 6,10,20) and for4 taper (tip chord/root chord) values,for a constant chord wing (taper = 1),for two trapezoidal wings (taper 0.5and 0.1), and for the elliptical wing.This last, elliptical distribution isquite useful, as it applies to all semi-tapered and double-tapered wings,examples of which are shown inFigure 9. Except for constant chordwings, which for obvious reasonsusually do not employ any twist,wings of present day aircraft rarelyuse more than a few degrees of twist.At the high AOA occurring duringthat intense pull up, this smallamount of twist has almost no influ-ence on the lift distribution and canbe neglected here. By the way, byneglecting it we err on the safe side.Returning to Figure 8, we see thatthe half-wing has been subdividedinto 6 wing elements, with the widthof each of the 4 inner elementsamounting to 20% of the wing semi-span and the remaining twoamounting to 10%. This apparentlystrange division of the half-wing willspeed up our wing loading. It is,however, not binding and the buildercan find his own system in subdivid-ing the wing. Each wing elementgenerates a certain amount of lift,expressed in percent of the total liftwhich, of course, adds to 100%.Each wing planform has its own liftcontributions appearing in the "win-dows" of the Table.

The inquisitive reader may w onderhow we arrived at this Table. Well ,years ago NACA exactly calculated liftdistributions of a large number ofstraight tapered wings and includedthe influence of the wing twist. Theresults are neatly presented in, amongother sources, the well known bookby I. H. Abbott and A. E. vonDoenhoff, THEORY OF WING SEC-TIONS, which is available from EAA.The Table is the result of using NACAdata and adding up (integrating) con-tributions of each single wing element.The reader may also want to checkthe influence of the small amount oftwist to convince himself of non-importance of the twist at high AOA.If our wing has a taper between thevalues given in the Table, we simply

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Figure 8CENTERLINE

WING SHAPE

CONSTANT CHORD

TAPER 0.5

TAPER 0.1

ELLIPSE

I I0 20 % 40| WING SEMI

AR ?0 % 20 %6

10?06

102061020

23.3 %22.721.725.825.825.729.130.131.125.5

22.9 %22.521.624.223.923.526,026.426.624.4

1% 60

20 %21 .9 %21 .921 .421 .521.220.921.521.32121 .8

?% 80

20 %19.8 %19.920.917.817.817.815.715.114.418.0

? 9 \% 90 100

10*7.7%8.29.26.97.27.55.34.94.76.8

10%4.4%4.85.83.84.14.62.42.22.13.5

TABLE 1

%

take values between the adjacent val-ues of taper. For example, for an AR=6wing with a taper of 0.6 we take valuesof the constant chord wing (taper = 1)and those of the taper = 0.5, remem-bering, of course, that our taper = 0.6is much closer to taper = 0.5 than totaper = 1. We could take somethinglike 25.3% for the first 20% w ide wingelement, 23.9% for the next, etc. Weuse the same interpolation method ifour wing has an AR betwe en thoseprinted in the Table. We now have 6"packages" representing the entirehalf-wing l i f t - all 100% of it. All wehave to do is transform these percent-ages, sitting on their respective wingelements, into pounds.Suppose our aircraft weighs, fullyloaded, 2,000 lbs., 300 Ibs. of which isthe weight of the wooden wing. W ehave to know the weight of the wingbecause in flight the win g will more orless support its own weight and thewing spar can forget about that part of

the aircraft weight. It will have to carryonly the rest of the weight, in our case,2,000 Ibs. - 300 Ibs. = 1,700 Ibs. Tounderstand this, let us consider a long,slender wooden board. When sup-ported in the middle, it visibly sags atboth ends. However, placed in waterit immediately straightens, as thebuoyance exerts its "lift" along theentire board's length. There may becomplications here if, for example, wehave heavy fuel tanks sitting at thewing tips, or wing-mounted engines,but these cases are beyond our pre-sent discussion. The de-signer hasprescribed fo r this aircraft, say, a n =4.4 limit load factor. This means thatin a violent pull up (or upon entering avery strong upwards gust) the wing isexpected to suddenly generate a 4.4times stronger lift. The spar(s) in bothwing halves will have to withstand -elastically - 4.4 x 1,700 Ibs. = 7,480lbs., or 3,740 Ibs. on each half wing.This 3,740 Ibs. now corresponds to

TOTALwing element, % 20% 20% 20% 20% 10% 10% 100%wing element, ft 2 2 2 2 1 1 10element load, % 23.3% 22.9% 21.9% 19.8% 7.7% 4.4% 100%element load, Ibs 871.4 856.5 819.1 740.5 288.0 164.6 3740element load, Ibs 870 860 820 740 290 160 3740

(rounded-off)TABLE 2

those 100% of the total "lift" Figure 8s h o w s . If our 40" chord c o n s t a n tchord wing has a wing span of, say, 20ft., the wing semi-span will be 10 ft.Those 6 wing elements along the wingsemi-span will amount to what you seein Table 2 and will have to withstandlimit loads, obtained by multiplying thepercentage element loads wi th thetotal limit load of 3,740 Ibs. For exam-ple, the innermost wing element will beloaded to 23.3% x 3,740 lbs., or 0.233x 3,740 = 871.4 Ibs. Table 2 alsoshows the rest of the loads for onehalf-wing of our example aircraft. Aswe will see shortly, it makes sense toround off those loads down to 20 or 10Ibs. We have now determined theloads to be put on respective wing ele-ments on each side of the wing. If thewing has been properly constructedand if the designer did not goof, thewing should easily support this totalload and will, upon the removal of theload, nicely "recover".

Strictly speaking, the lift and loaddistributions along the wing span, asshown here, hold only for an isolatedwing or perhaps a parasol or a high-wing ar rangemen t . Wha t about amid-wing or a low wing position - asthe central part of the wing "disap-pears" into the fuselage. Theory saysthat the loss of lift due to the fuselagedoes not amount to as much as onemight expect. The lift appears todevelop in the region between the leftand right half of the wing to almost fullextent. Numerous wind tunnel investi-gations tend to confirm this, especiallywhen proper wing fairings have beeninstalled. We will simply assume thatno mistake is made if we extend thewing loading across the fuselage as ifthe fuselage were the central part ofthe wing. (What about those parasolwings with huge cut-outs right in thecenter and above the cockpi t . . . sothe poor pilot can squeeze himself inand out?)

LOADING SCHEDULEAs to the loading material, we canuse bags of lead shot (expensive)water in 1, 2 or 5 gallon cans (bulky),iron (or lead or gold!) bars, or sand-bags (best). Let us assume we aregoing to use sandbags. Dependingon total load (7,480 Ibs. in our exam-ple), we need bags in an assortmentof 100, 50, 20 and 10 Ibs. There arelikely sand and gravel suppliers in thearea that will be happy to help outand it will cost next to nothing. Theentire wing testing can be performedat varying engineering levels. The

very least we do is slowly and gentlyload the wing up to its full load, wait afew minutes and unload. Nothingwrong with that, except that now wedo not know whether the wing has38 MARCH 1992


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Element, %load, Ibs

off, Ibsconsisting

20174.3

170100+50+20170

20171.3

170100

+50+20170

20163.8

160100+50+10160

TABLE 3

20148.1

150100

+50-

150

1057.6

6050

+10-

60

1032.9

3020

+10-

30

TOTAL100%

748 Ibs.740 Ibs.

740 Ibs

excessive "permanent set."other words, we are not sure thewhile fully loaded, stayed withinregion, as it should. Withlittle more effort we can recorddeflections by reading off thosethat have been suggestedimportant ones are at theand in the wing center, how-yardsticks installed midwaythe wing root and the tipsinterest for the more engi-oriented builder.Three people can comfortably han-loading and unloading and theof loads and deflections. Ita very good idea to invite an EAAFigure 9CENTER LINE

100 *

80 %

Technical Counselor. He can do allthe recording on the prepared form,check the weights of the sandbags asthey are lowered onto the wing andtake pictures of the wing under load .. . and of the builder under stress. Hecan also keep assuring the builderthat a properly designed and consci-entiously built wing is not beingtortured at all.First we have to rem ove slack in thesupporting fixture and in fittings, pins,bolts, rivets, etc. We note positions ofthe yardsticks and then carefully loadthe wing to some 10-20% of the fullload. It does not matter what percent-age we take. What matters is that weload all wing elements exactly corre-sponding to the selected percentage.If we decide to use 20% in our exam-ple wing, the wing elements will beloaded with the rounded-off loads asshown in Table 3.This rounding-off has resulted in aslightly lower percentage (19.8%)which we duly note in our records.We now read the new positions of theyardsticks and note them. The differ-ence between the f i rs t reading andthe second is the deflection of thewing tip. There should be no differ-ence at the wing center. If there is, itmeans the wing supporting fixture - orthe fuselage, as the case may be -also "settled" a little. This settling inthe middle must be subtracted fromthe wing tip deflections. On the finaldiagram, Figure 10, which we willcompose after all measurements havebeen completed, the position of the"virgin" wing is shown as (0) and theposition at the 10%-20%, in ourexample wing 19.8%, is shown aspoint (1). On the horizontal axis weplot wing deflections and the distance(0)-(1) indicates the wing deflectiondue to 19.8% load which is plottedalong the vertical axis.This first loading has no subse-quent value - it was only necessary tomake sure the internal disorders in thewing structure have settled and that

from now on any increase of the loadwi l l rea l is t ica l ly demonstrate wingbending. After a minute or two,wegent ly remove these f i rst bags andmuch to our surprise we note that thewing does NOT return to its original(0) position. The new wing tip posi-tion is now found at (2). This is notthe so much feared "permanent set".It is simply the settling of the wing.(Note: horizontal distances on thediagram are exaggerated for clarity!)The real "proof testing" of the wingnow begins. We exactly repeat theprevious loading (in the example wing19.8%) and read new deflections.Then we move to 60%, 80%, 90%,95% and 100%. The reason w e makeshorter intervals as we are approach-ing 100% is not that we feel insecureabout our wing. If we did not expectthe wing to perform flawlessly weshould not have brought it to the test-ing in the first place.We make shorter steps becausewe wan t to find out how close thewing's elasticity is to the 100% oad.The expected "straight" line betweenthe 20% load and 100% load willmost likely be slightly bent. Nothingwrong with that. However, if it curvesexcessively, as indicated in Figure 10with a broken curve, we know w e areapproaching some kind of yield pointand we better discuss this with thedesigner, or kit seller or FAA. The realtest of the amount of elasticity willcome, however, at the very end whenwe definitely unload the wing. As wekeep increasing the loads we l istencarefully for any screeching sounds -or, for that matter, for any soundscoming from the wing. As soon assomething is heard we stop andinspect the suspect area. Fortunatelyfo r us, the top of the wing, which isloaded in compression, is at the bot-tom now and easily accessible. Evenif we do not hear anything, we watchfor skin buckling. It may start appear-ing at some 50% load and does notnecessarily represent a weak spot. Itmay be purely elastic and it will disap-pear after the wing has beenunloaded. However, we make notesof the appearance.On the way down to final unloadingwe will watch the same places andwait for the buckling to disappear.This buckling, if it appears over theglued areas (over the main and auxil-iary spar, ribs), can signal poorbonding. We make notes about that,hold a brief conference and decidewhat to do. Do we stop here, unloadthe wing, take it into the workshop,open it and exam ine the suspect area,or should we go on with loading andsee what develops? Both approacheshave advantages. Just make sure wedo not close our eyes in the hope thatthe ugly delamination will go away. It

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20 - -

( 0 ) (2') ( 1 4 ) WING T IP DE FL E CT I ON in i n . -Figure 10

will not.Some readers may wonder why wedo not perform measurements atsmall intervals, say 10% intervals.W hy push the load immediately to60%? Well, even this 60% is perhapstoo low. Maybe we should proceedimmediately to 80%. The reason issimple. If the wing shows any kind ofweakness at that miserly 60%, it isgood for nothing and we should throwit away. We just have to learn to livewith the fact that the so-called "limitloads" will exist when we fly and thatthe air is much less forgiving than thesandbags. Of course, one can alwaystry to locate the weakness, and try torebuild the wing. Then proof-test itagain. But, defects at 60%? Hmmm.As we gently progress towards100%, listening and watching forpossible visible signs - with points (3)-(8) indicating our progress on Figure10, we carefully maintain an absolutesymmetry of loading on both halvesof the wing (each man working hishalf wing) and we strictly follow theloading distribution, as explained indetail for the 20% load. At each step(20, 60, 80, 90, 95, 100%), we wait acouple of minutes for the wing to"settle", before taking readings. Afterthat we begin to gently and symmetri-cally unload the wing, stopping at thesame loads as we did on the way up,and take readings again, again eachtime waiting a minute or two for thewing to settle. Finally, we arrive at theinitial 19.8% load (point 13 on the dia-gram) and at the zero-load wing

position (14). With this we have com-pleted the physical part of the wingtesting. It is time now to take a lookat the results and interpret them.HOW GOOD IS OUR WING

After drawing both left and rightload/deflection diagrams and takingcare of the possible displacement ofthe center of the wing at each loadchange, we are first surprised that thetwo diagrams are not absolutelyequal. There is nothing wrong withthat, unless we notice some ve rystriking differences. No two men orwomen are absolutely equal, no twowings are absolutely equal and yourleft hand is not absolutely equal toyour right hand. Major differences,however, require some investigation.Often we make mistakes when takingreadings. If the discrepancy is local-ized, only one test point out of order,we may disregard it but only after athorough discussion. We will alsonotice that more often than not thetest points on Figure 10 are scattered,they do not follow exactly that idealline. Nothing wrong with that either,as long as discrepancies are notridiculously large. The more carefulwe were during the testing and thebetter tools we were using, the moreconsistent these test points willappear on the diagram.Now, how do we know how "good"our wing is? The fact that it did notbreak is, of course, not enough. Wedemand that after each unloading

from the limit load, be it at the prooftesting or in flight, the wing returns toalmost its original state. An idealwing would be represented by astraight line and points (7)-(9), (6)-(10),(5)-(11),(4H12),(3)-(13)and(2H14)would come together. Distance (2)-(14) , representing that importantparameter "Permanent set", would bezero. Real wings display a certainpermanent set, very small comparedwith the maximum wing tip deflection- distance (14)-(15) in Figure 10. Andjust how small is "very small"? Sincethe return line (8) . . . (14) is practicallya straight line we see that the "perma-nent set" depends on the curvaturethat begins to s how as we areapproaching the 100% load. Per-manent set therefore depends on howfar below that earlier mentioned "yieldpoint" the wing is when fully loaded. Ifwe had a steel or aluminum specimenin a testing machine it would be a sim-ple matter of finding out w here thepermanent set appears. Engineeringagreement (please note that I am notusing terms such as "Law of Nature"or "Law of Physics") says that we candefine the permanent set or perma-nent deformation in a tensile test whenthe permanent strain is less than 0.002inch per inch length of the sampletested. (This is strictly valid for thedefinition of the yield stress, but wehave earlier pointed out that the elas-ticity limit appears close to the yieldpoint.) A wing loaded in bending is amuch more complex structure, so wemay find it difficult to apply the above"engineering agreement" to our wingtesting. FA A goes to the very root ofthe problem by stating: "The primarystructure should be capable of sup-porting, without detrimental permanentdeformation, the limit loads, if theloads are properly distributed andapplied. In addition, temporary defor-mations that occur before the limitload is reached should be such thatrepeated occurrence would notweaken or damage the primary struc-ture." FAA obv ious l y lets thedesigner/builder/tester determine howmuch that distance (2)-(14) shouldamount to - compared to the maxi-mum wing tip deflection. Just to givean example of the designer/builderresponsibility - as 100% of the limitload is reached, do we find it difficultto activate flaps or ailerons? Thinkabout it. In our experience, a goodwing should show a permanent set (2)-(14) that is less than 1% of themaximum deflection (14)-(15).With that we can disassemble thetesting set up and examine the f i t-tings as mentioned earlier. If nobodyfinds anything out of order, we arehappy, for while before we thoughtwe had a structurally sound wing . . .now we know. *

Structural Testing of Homebuilt Aircraft

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