Fabrications of extruded aluminum alloy structures are reviewed. The design criteria and the fabrication of the main structure of a sailplane wing made of a few extruded profiles longitudinally connected one to the other are illustrated. Structural tests recently carried out are reported upon

Morelli, P.

Romeo, G.

English

NASA Technical Reports Server

i
 79-27077
EXPER_AL INVESTIGATION YNTO THE FFASItlLT_y
OF AN "EXTRUDED" WING
Piero Morelli and Giulio Romeo
Politecnico di Torin-, iLaly
SUMMARY
Research work in the Politecnico di Torino and realizations (fabrications)
of extruded aluminium alloy structures during the past years is briefly reviewed.
The design criteria and the realization of the main structure of a sailplane
wing made of a few extruded profiles longitudinally connected one to the other
are then illustrated. Structural tests recently carried out are reported upon.
INTRODUCTION
Early research work and the first realizations t r_ the M-300 sailplane pro-
totypes were reported upon in reference 1. Figure 1 illustrates the cross sec-
tion _f the M-300 extruded stmlctures: first and second realization of %he
ailerons (a,b), tailplane (c), wir_ spar (d). An aluminium alloy _i TA16
(A.A. 6063-T6) was employed for the ext_sion except for the spar, which was of
A.A.7075.
In more recent years the same structural concept was adopted by the firm
C_proni Vizzola Costruzioni Aeronautiche, manufacturer of the two-seater sail-
plane Calif A-21S (ref.2 and 3). Figure 2 illustrates some of the parts of this
glider which were realized by extruslon using the same aluminium alloy men_oned
above: airbrake (_), flap (b), aileron (c), elevator (d), and leading edge of
the wing central part (e). The aileron and elevator extruded profiles incorpo-
rate the hinge (A). In the aileron leading edge lodgir_ is provided (B) for the
counterweight, uniformly distributed along the span for static and dynaLtcbalance.
In the M-3CO and Calif extruded structures the original wall thickness of
1.8 to 2.0 mm was reduced to design values of .5 to .8 mm by chemical milling
of the outer surface.
All these structures are basically ribless. They proved light and largely
adequate in str_n_th and stiffness.
One of the M-300 prototypes is still active. The Calif two-seater has been
series produced with the extruded parts mentioned here since 1975, except the
extruded airbrake which was introduced in 1978.
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Advantages and limitations of the extruded structures were discussed in
references I, 2, 3. They are briefly summarized here.
Main advantages are:
1. Reduction of manhours required to realize the structure, mainly during the
assembling stage.
2. Roduetion of cost in a series production men the cost of the expensive ex-
trusion _ios can be distributed over _ high number of pieces.
3. Correct reproduction of section contours with consequent aerodynamic benefit.
i.
2,
The practical limitations are principally the following:
The extruded profile has necessarily a constant cross section. Through suit-
able mechanical and chemical operations, however, it is possible to achieve
a certain degree of cross section variation along the beam axis.
The maximum linear dimension and net area of the profile section are limited
by the power of the available extrusion press.
3. The difficulty of extruding increases with high strength aluminium alloy
such as 2024 or 7075.
. A mlnimumwall thickness is imposed by the extrusion process, which is some-
tin,es excessive in relation to the strength and veight/strengtn ratio re-
_ired.
A wide field of possible applications seem to exist notwithstanding these
limitations, particularly for gliders and l_bt powered aircraft.
A glider has been conceived, which is similar to the M-300 from which it is
derived and is suitable for a wide use of extruded structures, whose locations
are indicated by the shadowed areas in figure 3.
The realization of the central part of the wing of this glider is the aim
of the research work started a few years ago at the Politecnico di Torino, after
the completion of the first stage which led to the realization of the above de-
scribed M-300 extruded parts.
%
THE DESIGN OF AN "EXTPUDED" WING
The wing illustrated in figure 3 is 15 m span with rectangular-trapezoid
planfcrm, the central rectangular part being extended over _ m.
It is a three-plece wing: the central part is a flat sinfle piece connected
to the Fuselage by a h-polnt attachment; the outer trapezoid pauels are a*tachod
to the ends of the central part and give the wing the required dih_drn_ r,ncle.
420
fi
The central part i_ conceived as a combination of extruded profiles: a
possible typical cross section is illustrated in figure h (airfoil FX 67-K-
-170]17 ), which is purely indicative of the basic idea. Corresponding to the
airfoil maximum thickness a box _tructure can be seen which carries practically
all bending loads and a good portion of the shear/torslon loads. The other two
thin walled extruded profiles are riveted to the central box and contribute to
the shear/torsion strength and stiffness of the whole structure. The extruded
profile at the trailing edge is a flap.
The w_ng structure is intended as basically ribless, as far as tests will
confirm tl_a_ ribs can be eliminated.
In order to provide the central box with the required bending strength and
stiffness under the prescribed loading condltlo_ (according to the OSTIV Air-
worthines_ Requirements, ref.h) a cellular structure was adovted for the dorsal
and ventral panels. This multi-cell structure was tentatively designed to pre-
vent _eneral and local elastic instability.
The central box is made of two profiles _oined by riv,_tin_ the two hnlfw_
alone %he span (A.__ _061-T6).
The large bendin6 d_formation, tyT,ical of a h_<h asp,*ct ratio sailplane
wing, combined with the absence of ribc makes the problem of res[s+.,[ng the
"crushing " loads a basic one. One of uhe main eb_iectives of tb,_ testlng prc-
6ram is to ascertain how far the webs alone are capable of w;_thstandin_, _he
crushing loads.
The central box section is reduced along the svan bv chemically etchln_:
the outer surface of each of the two profiles so that the ori#:inal wall thick-
ness of the skin panels is decreased from 2._ mm down to .8 m_ at a spanwise
station about 2.65 m from the w_ng centerline. ',"hlsthicknc-;_: iz then kepf con-
stant over the rest of the wing.
Figure 5 shows the reduction of skin thick:less along th_ span in two po_-
ible ways. A step reduction (above) or a continuous tapering (below) can be real-
ized, the latter requlrin_',,however, additional equipment for ehemlcal milling:
at variable time of immersion.
Figures 6 and 7 show the central box cro_;s seeti,_n_-=at the wii_::root and
at a spanwise station from ,?.t,5m on. REPRODUCIBILITY OF "_""
ORIGINAL PAGE I_ I .
t.:XPt'_ INb:NTAL PtIoG._A5_ A.';PRT.f,]'I;:'S
%
;',,,v..ra[probb-m:: ure to !,,,faced in the realization of a wink as described
in tlu" prccodin,! u:tral'r:i!_t:.
A pi','l im{nar'y ,,XDi, l'ir._.lit:t] {'iV, .:I iC;lt[t"il i,'fii_ C. :1::_ i,'l',' ! .';,',','.'::'!V:," ;,*l ,'!" '."
_0 ch,'c'k l.tie f'ollowiill! r,_int_::
• ,h,, capability ,_f the i,/,];lll:tl" Drill')|-: .'_ Wi __:'___tl i ill' ' I [ i I'i CO_PII*'''''I [ : iI ,, , . .i.,.:i.-_:
stresses without tncurrinK local instability phenomena at low load factors.
2. The capability of the box structure to withstand the deslgn bending moments
and, in particular, the crushing loads due to bending deformation.
3. The capability to obtain the centre/ box profiles bY extrusion of a suitable
material at an acceptable degree of accuracy and reasonable cost.
h. The feasibility of satisfactory chemical mtllin8 in relation to the par-
ticular aluminium alloy adopted for the extrusion.
With reference to points 1 and 2, it was decided to check the general de-
sign of the _entral box structure, and of the cellular panels in particular, by
realizing a 'slmula,ed extruded structure and submitting it to pure compression
and pure bendln_ tr.sts. The cross section of the simulated structure in figure 8
shows its conventional construction through Z-stringers and metal sheet, both of
dural, connected by rlvets.
Notwithstanding the difference in material and some geometrical features
these tests gave some valuable indications (ref.2) so that the realization of
the expensive extrusion dies could be undertaken with reasonable confidence•
The two extruded profiles were then obtained, having the cross section
shown in figure 6.
Several attempts were necessary, with modification of the die, before an
acceptable degree of accuracy of the section co**tour was achieved.
The aluminium alloy employed on the first extruded profiles was not satls-
factory (inadequate values of the yield and rupture stress). A different alumin-
ium alloy was then used of higher strength but, perhaps, rather poor plastic
characteri st ics.
It shouid be remarked here that, in our I+_alian situation, the choice of
materials for extrusion is extremely limited. In fact, since our factories are
not furnishers of the aircraft industry, the supply of a small quantity of
extrusions such as required for research cm, only be made of a material of
current use, i.e. having rather low strength characteristics.
Pure bending tests were planned and carried out on several specimen,lO00 mm
long, of the real extruded structure using the bending test machine of the Istl-
tuto d[ Progetto di Aeromobili - Politecnico di Torino.
Figure 9 shows the testin_ equipment. Figures I0 and ii show the deflection
curve_: measured on sueclmenswith wall thickness of 2.3 and .8 mm, i.e. h_vln_;
the cross sections i_llustr ated in figure 6 and 7, respectively.
Figure i? shows the typical failure in compression due to bendin_ which
occurred on one of the tin.8 m_n specimen.
The results of these tests were encouraging, although of still l[mitedv_i-
422
dlty for two _ain reasons:
1. Since %be ends of the specimen are rigi_ attached to the test machine, only
a rather short central portion of the structure is free from their restrain-
ing influence. Therefore, the capability of the structure to withstand the
crushing loads cannot be fully evaluated.
2. Shear is not present.
Testing on a full scale structure was therefore planned.
A test structure was prepared corresponding to the central box of the rec-
tangular part of the sailplane wing illustrated in figure 3.
The span of this test specimen was 7.67 m, less than the 9.0 m span of the
wing rectangular part, due to limitations of the available equipment for chem-
ical milling.
The skin thickness was reduced spanwise through chemical milling by .3 mm
steps from 2.3 down to .8 mm as shown on the upper part of fig. 5. Two exten-
sions were added at both ends of the structure to allow the appllcation of con-
centrated loads corresponding to the actual distributed load carried by the
outer portions of the wing (see figure 13).
The spanwise wing llft and mass distributions were evaluated and then re-
placed by ten concentrated loads, giving a good approximation of the bending
moment and shear distribution (see figure 13).
Figure lh shows the structure under the load corresponding to lo_d factor
ns8.
The i_crementsl load was 2,h13 N corresponding to a unity load factor in-
crement. The ultimate load was 2_ ,07_ H corresponding to a rather high ultimate
load factor of 9.975.
T_e structural failure occurred at a load factor n=8.72.
As shown by figure 15, the dorsal cellulsr panel between the fittings, simu-
lating the wing-fuselage attachments, collapsed under the combined effects of
compression and crushing loads. In this area both webs were larsely cut out to
allow the connection of the fittings to the structure.
Figure 16 shows the deflection curves of the whole structure at load fac-
tors of 2, 4, 6 and 8. It c_m be seen that, at high load factors, the deflec-
tion of the left wink becomes a little higher if compared with the other wing,.
This is presumably due to the growing elastic buckling of the dorsal panel
caused by the large cut-outs of the wing central part where the failure
finally occurred (fig. 15_.
The deflections at different stations are plotted _ersus load factor in
flgure 17.
Jt q )DUCtBP.J OP
ORIG.'At _' %(;8 |S P" ""
423
Strain gage measurements shoved: a) a slight elastic buckling of both
webs in their lov_itudinally compressed part at load factors above n = h ; b) no
buckling whatever of the 4orsal panels along the span; and c) a max_ local
normal stress of 235 B/_ 2 at nz8 on both dor_ and ventral panals.
CONCLUSIONS
The failure under bending having occurred at a very high load factor
(nl8.72) and in the central part of the structure where the webs can be easily
reinforced, the result of this first static test ¢,_n be considered maccessful.
There is a reasonable confidence that, after rein._orcement of the web cut-outs
in the central portion of the structure, the residual load factor increment
A n-9.975-8.T20" 1.255 will be attained.
Although torsion static strength and fatigue life are to be demonstrated
before a structure of this type can be assessed to be adequate for a sailplane
wing, the result of the actual shear/bending test should probably be considered
of basic importance as it practically demonstrates the feasibility of a ribless
structure made of a few extruded profiles longitudinally connected one to the
other.
424
REFImE_CES
i. Morelli, P. : Extruded Light Alloy AircrefZ Structures. Proceedings of the
._ ist International Symposium on the Teehmology and Science of Motorless
Flight. M.I.T., Cambridge, U.S.A., October 1972.
2. Romeo, G.: Realizzazione per estrusione di strutture aeronautiche - Ricer-
ca su struttura s/are estrusa. Prestampa AIDAA n.60, IV Congresso Nazio-
nale dell'Associazione Italiana di Aeronautica e Astronautica, Milano,
September 1977.
3. Romeo, G.: Prog, ess on Extruded Structures. 16th OSTIV Congress,Chateauroux,
France, July 1978.
h. OSTIV Airworthiness Requirements for Sailplanes, September 1976.
425
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FIG.3- GLIDER DESIGNED FOR WIDE
USE OF EXTRUDED STRUCTURES
FIG.4-TYPICAL CROSS SECTION OF AN "EXTRUDED" WING
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FIG.5- SKIN THICKNESS REDUCTION ALONG THE SPAN
FIG_6- CENTRAL BOX ROOT SECTION
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¢FIG.7- CENTRAL BOX OUTER SECTION
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FIG,8- CR06S SECTION OF "SIMULATED" STUCTURE
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FIG. 10- DEFLECTION CURVES OF THE EXTRUDED
SPECIMEN t:2.3aa UNDER BENDING TEST
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DEFLECTION 4
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i M b- 6700 NmJ _ _ "..
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l,_e_.___r, .05,..., 200 ..,,,. 200 ..., 105,,. . I_
f__t ..... 7..o.ooo. ...... ,,.o .... . .... ° ...... o..-o°---- Mb
.... __ .--; .... -----:: ::z _-,.-- '-_-_li I
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FIG.11- DEFLECTION CURVES OF THE EXTRUDED
SPECIMEN t,.Sm_, UNDER BENDING TEST
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BENDING MOMENT S H EA R
Calculated Calculated -
Test ------ Test ._
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FIG.14,-STRUCTURE AT LOAD FACTOR n-8
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FIG.$11- ITRUCTURE FAILURE
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DEFLECTION
FIG.16- DEFLECTION CURVES
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DEFLECTION
FI0.17-DEFLECTION VS. LOAD FACTOR
N79-27078
Tt.mk2t_n' _ _ _L ImCHNIZNI8 OF LZGHT kTIULKJtHB8
IN 'flHZ FM_L'tlR _ I_AOCBS8
l[Lmar J, Br•lthmch*
Langley lteNareh Center
mY
ltecantlyw it ha8 become sore and sore evident that many difficulties
encountered in the course of aircraft flutte_ analyses can be traced to strong
localised nonl/_e•rities in the control mechanism. To cope with these prob-
lems, wore reliable mathematical nvdels paying special attention to control
system no_line•ritles nay be estsbltahed by neans of nodified ground vibration
test procedures in combination with suitably adapted modal synthesis approaches.
Three different concepts are presented in de_ail.
INTNODUCTIOti
At first glance the flutter clear•rice of soaring and light airplanes does
not seen to r•ise any serious problems which cannot be solved by means of
today's asroelastic tools. This is true even for the determination of the
unsteady serodynanic loac_s as long as cases with larg_ aspect r•tios at coapa-
rably low speeds •re considered. The elastodynaatc•l characteristics can be
deternined by using common experinental or analytic•l methods if structural
linearity can be 888uned to be • proper approximation. However, as experience
hu shown, the control mechanism8 of light airplanes 1 are generally nonlinear
to such a large extent that setting up • dependable sothemetioal nodal requires
spatial attention, including mdifications to standard linearised procedures.
In the first part of this paper some of the east frequently occurring types
of _ontrol-systen nonlinearities are described. To get an ida• of the influence
of scan typical ha•line•rifles on the •eroel•stic stability the results of wind
tunnel flutter tests on • nonlinear wing aileron model are presented. After
that, it is shown in detail how the aeroelastic equations of light airplanes
with locslizod nonline•rities may be formulated bF using various suitably modi-
fied ground vibration test (G3;T) proc•dures all based on the well-known mod81
synthesis approach. The shortcomings as well as the usefulness of the different
concepts are di_ussed.
*IAC-HASA Senior Resident Research Associate.
1Light airplanes as sued in this paper include both powered and unpowered
vehicles where the power to the flight control system is supplied by the pilot
without eleotrioal or hydraulic boost through a system of cables, pulleyse push-
rods, bellcrankn, or other Ichanic•l linkages.
43?
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Experimental investigation into the feasibility of an extruded wing

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