Nationaal Lucht- en Ruimtevaartlaboratorium
National Aerospace Laborator y NLR
NLR-TP-98148
Full-scale fuselage panel tests
R.W.A. Vercammen and H.H. Ottens
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217-02
DOCUMENT CONTROL SHEET
ORIGINATOR'S REF. SECURITY CLASS.
TP 98148 Unclassified
ORIGINATOR National Aerospace Laboratory NLR, Amsterdam, The Netherlands
TITLE Full-scale fuselage panel tests
PRESENTED ATthe 21st ICAS Congress, Melbourne, 13-18 September 1998.
AUTHORS DATE pp refR.W.A. Vercammen and H.H. Ottens
980323 11 6
DESCRIPTORS Aerodynamic loads Fatigue life Residual strengthAxial loads Fatigue tests Structural designBending Fiber reinforced composites Test facilitiesCurved panels Full scale tests Weight reductionCyclic loads FuselagesDamage assessment Laminates
ABSTRACTIn fuselage design studies there will always be the necessity to testcomponents in a realistic way. The fuselage panel test facility at NLRoffers the possibility to subject fuselage skin sections to residualstrength and fatigue tests. The fatigue test loads simulate cabinpressurization in radial and axial direction and axial loadsrepresentative of fuselage bending. To verify the test methodology andthe specifications of the test set-up, the performance of the test set-upis evaluated using a GLARE panel designed by Fokker Aircraft. The testresults indicate the suitability of the biaxial load introduction systemsto load a panel comparable to a panel in a pressurized fuselage subjectedto bending. In addition, the tests show that compared to conventionalfuselages, the designed GLARE Panel combines a substantial weightreduction with an excellent fatigue behaviour and sufficient staticstrength. Test on fuselage panels designed by Shorts, DASA and Aleniawill be used to demonstrate the technological feasibility of GLAREfuselages also with window cut-outs, to study the growth of multiple sitedamage in stiffened lap-joints of aluminum curved panels and to determinethe effect of multiple site damage on the residual strength of suchpanels.
-2-NLR-TP-98148
Contents
Abstract 3
Introduction 3
Test facility 4
GLARE panel test 5
Current tests 6
Conclusions 6
References 7
11 Figures
(11 pages in total)
-3-NLR-TP-98148
FULL-SCALE FUSELAGE PANEL TESTS
Roland W.A. Vercammen and Harold H. Ottens
National Aerospace Laboratory NLR
Anthony Fokkerweg 2
NL-1059 CM Amsterdam, The Netherlands
Abstract
In fuselage design studies there will always be the
necessity to test components in a realistic way. The
fuselage panel test facility at NLR offers the possibility
to subject fuselage skin sections to residual strength and
fatigue tests. The fatigue test loads simulate cabin
pressurization in radial and axial direction and axial loads
representative of fuselage bending. To verify the test
methodology and the specifications of the test set-up, the
performance of the test set-up is evaluated using a
GLARE panel designed by Fokker Aircraft. The test
results indicate the suitability of the biaxial load
introduction systems to load a panel comparable to a
panel in a pressurized fuselage subjected to bending. In
addition, the tests show that compared to conventional
fuselages, the designed GLARE panel combines a
substantial weight reduction with an excellent fatigue
behaviour and sufficient static strength. Tests on fuselage
panels designed by Shorts, DASA and Alenia will be
used to demonstrate the technological feasibility of
GLARE fuselages also with window cut-outs, to study the
growth of multiple site damage in stiffened lap-joints of
aluminum curved panels and to determine the effect of
multiple site damage on the residual strength of such
panels.
Introduction
In service, fuselages are subjected to the combined
loading of cabin pressure and fuselage bending. It is
therefore highly desirable that in case of fuselage design
studies, curved structures are tested under those biaxial
loading conditions. For this purpose barrel test set-ups are
generally used. These set-ups, however, are less attractive
for generic studies which are not directly related to a
particular aircraft design: the radius of curvature is fixed,
a large number of panels has to be tested simultaneously
and the test frequency is rather low. In addition, barrel
tests are expensive due to the large number of panels
required and the long testing duration. The panel test
facility at NLR (Fig. 1) was developed to avoid these
disadvantages. In this facility, which is flexible in panel
diameter, panel width and panel length, a single fuselage
panel can be tested in a relatively short time(1).
In the full-scale fuselage panel test facility at NLR,
fuselage panels with curvatures between those of
relatively small aircraft such as the Fokker 50 and those
of large aircraft such as the Airbus A300, can be tested
under flight loading conditions or can be subjected to
residual strength tests. The loading sequences will be
derived from the aircraft flight loading conditions. This
results in circumferential load sequences caused by
internal air pressure, synchronized with an axial load
spectrum representative of both cabin pressure and
fuselage bending due to taxiing or gust loading. With an
average number of 30 axial load cycles and one cabin
pressure load cycle within one flight, the testing
frequency is about 10,000 flights per 24 hours. Next to
the fatigue tests static strength and residual strength tests
can be performed. During these tests large cracks such as
a two-bay crack can be allowed.
In order to verify the methodology to load a panel
comparable to a panel in a pressurized fuselage subjected
to bending, the radial expansion, the axial load
introduction and the circumferential load introduction
obtained in the test facility are evaluated using a GLARE•
fuselage panel(2). This GLARE panel was designed and
built by Fokker Aircraft within the Brite Euram IMT
• GLARE = Glass fibre reinforced aluminum laminate
-4-NLR-TP-98148
2040 project "Fibre reinforced metal laminates and CFRP
fuselage concepts".
Fokker Aircraft designed the GLARE panel to verify the
applicability of GLARE as a fuselage skin material in the
crown section of a Fokker 100. Therefore, the panel was
subjected to a fatigue test in which 180.000 flights (two
lifetimes) were simulated. The fatigue test was followed
by static tests to Limit Load, Ultimate Load and failure.
In the framework of the current Brite Euram programmes
"Advanced concepts for large primary metallic aircraft
structures" and " Structural maintenance of ageing
aircraft" NLR performs durability tests, crack propagation
tests and residual strength tests on several fuselage (side)
panels of Shorts, DASA and Alenia. The tests are to
demonstrate the technological feasibility of GLARE
panels with window cut-outs, to study the growth of
multiple site damage in stiffened lap-joints of aluminum
curved panels and to determine the effect of multiple site
damage on the residual strength of such panels.
Test facility
The major components of the test facility are the main
frame, the pressure chamber and the load introduction
systems (Fig. 1). The main frame is a very stiff steel
structure. It consists of heavy bottom and top beams and
two vertical main columns. A triangular shaped frame,
which houses the hydraulic actuator, is mounted above
the top beam and two auxiliary vertical columns. The
panel is mounted in the frame such that the centre of
gravity of the panel cross-section is in the working line
of the actuator. The height of the test facility is about
7.2 m, the width is about 4.5 m. The pressure chamber is
formed by a seal and base structure. The base structure of
the pressure chamber is formed by a stiffened base plate
and two support beams which are bolted to the vertical
columns of the main structure. The base plate has a large
central hole for air supply. At the front side the base plate
has curved wooden blocks around the edges which form
the side walls of the pressure chamber. An inflatable
inner seal is mounted on the wooden blocks and the
pressure chamber is closed by the panel. In order to
accomplish an air-tight seal without net radial force acting
on the panel edges, an inflatable outer seal is mounted at
the outside of the panel just opposite the inner seal. The
outer seal is bonded on the reaction frame, which consists
of an open rectangular steel frame with curved wooden
blocks. With a transport system the base structure of the
pressure chamber can easily be shifted aside during the
test, which significantly enhances the inspectability of the
inside of the test panel. The chamber pressurization, axial
loads and seals pressurization are regulated by a control
system.
The philosophy of the full-scale fuselage panel test
facility is that a curved test panel should be loaded
comparable to a panel in a pressurized fuselage subjected
to bending. Therefore, a radial expansion of the fuselage
panel due to cabin pressure should be accommodated
(Fig. 2). Pressurization of a fuselage results in the
development of circumferential stresses in the skin and
the frames. The stress levels in the frames will largely
depend on the stiffness of the frame-skin connection. To
obtain an appropriate radial expansion for all kind of
frame-skin connections, the ratio of stresses in the skin
and frames must be correct and adjustable. Therefore,
NLR chose to use separate loading mechanisms for the
skin and frames. The circumferential stresses in the skin
are reacted by bonded unidirectional glass fibres. The
loads in the frames are transferred to steel rods.
To transfer the loads in the frames to the steel rods, the
ends of the panel frames are locally reinforced. In
addition, the wooden side walls of the pressure chamber
have several holes through which the panel frame tensile
rods are guided. The openings between the panel frame
tensile rods and the hole edges are sealed air-tight with
silicone rubber collars. The steel frame loading rods bend
due to the applied cabin pressure and the axial panel
loads (Fig. 3). To keep these undesirable bending
moments as small as possible, the diameter of the rods
are minimized and the length is maximized. In case the
undesirable bending loads become too high, pre-stressing
of the frames, by displacing the frame loading rods,
reduces the frame bending loads.
The advantages of using unidirectional glass fibres is that
the loads are introduced very evenly over the length of
the panel(2)(3). Therefore, in the panel hardly any distance
is required for stress redistribution. In addition, the use of
-5-NLR-TP-98148
unidirectional fibres does not result in local stiffening of
the panel edges in axial direction. The length of the glass
fibre sheets can be chosen sufficiently large to limit the
rotation of the fibres at the upper side of the panel due to
the axial elongation of the panel. The glass fibres are
bonded to steel tangential plates. Because of their large
width, the tangential plates act more or less as hinges.
The outward movement of a panel due to pressurization
is therefore nearly radial and does not result in a
significant change in the shape of the panel from circular
to oval. The tangential clamping system is designed such
that the angle between the tangential plates and the
vertical column was adjustable to allow for a large range
of panel diameters to be tested.
In axial direction tensile loads simulate the fuselage
bending. To introduce these loads into the skin and
stringers a single load concept was chosen: At the axial
panel-ends the stringers are ended and the stringers loads
are carried by a gradually increased skin thickness. This
reinforced skin is loaded by tensile rods using steel
brackets (Fig. 4). The termination of the stringers makes
it possible to seal directly on the skin.
GLARE panel test
In order to verify the applicability of GLARE A• as a
fuselage skin material, Fokker Aircraft designed and built
a Fokker 100 fuselage panel, representative of the crown
section just in front of the wing, with a GLARE A skin
and GLARE N stringers. The panel (1210 mm x
3030 mm, R = 1650 mm) consisted of a stringer stiffened
skin, attached to aluminum frames by means of rigid
aluminum stringer attachments (cleats) and a longitudinal
riveted lap-joint in the panel centre (Fig. 4). A complete
GLARE Fokker 100 crown section would have a weight
that is 63 % of the current design in aluminum.
Before subjecting the GLARE panel to a durability and
• GLARE A=GLARE 3-2/1-0.3: 2*(0.3 mm 2024 sheet)+
(0.25 mm cross-ply glass prepreg)
GLARE N=GLARE 1-3/2-0.3: 3*(0.3 mm 7475 sheet)+
2*(0.25 mm UD glass prepreg)
a residual strength test, the GLARE panel is used for
evaluation of the test facility. The evaluation tests showed
a radial expansion of the panel and a uniform load
introduction in axial and circumferential direction: By
using glass fibres hardly any distance is required for
stress redistribution, i.e. the tangential strain distribution
is uniform after one stringer pitch from the panel edge
(Fig. 5). In addition, the use of unidirectional fibres did
not result in local stiffening of the panel edges in axial
direction (Fig. 6). The chosen axial load introduction
concept achieved a smooth distribution of the axial strain
levels in the middle of the panel (Fig. 6, 7) and the
stringer run-outs, skin reinforcements showed a negligible
influence after the first frame (Fig. 8). The fact, that the
GLARE panel did not expand radially (Fig. 9) is caused
by the presence of the lap-joint and different radii of the
panel halves.
After the verification tests a fatigue test was performed in
which 180,000 flights (two lifetimes) were simulated. The
axial loading sequences of the fatigue spectrum are
derived from the spectrum applied in the Fokker 100 full
scale test(4). The axial load is written as:
Faxial = a1*M y+a2*∆p with
Faxial = axial load in fuselage panel (N)
My = bending moment at particular
Fuselage Station (Nm)
∆p = differential cabin pressure (N/m2)
a1,2 = Fuselage Station dependent constants.
The spectrum consists of 36 repeating testblocks of 5000
flights. Each testblock of 5000 flights is subdivided in
four subblocks of 1250 flights. Three subblocks are
exactly equal, the fourth block is equal but for one
severest flight. Within this spectrum eight flight types
with different gust loading severity have been defined.
Figure 10 shows the axial loading and frequency per 5000
flights for these eight flight types. Each flight has five
segments: ground, initial climb, climb/descent, cruise and
approach. During the ground segments the cabin pressure
∆p=zero, during the cruise segment∆p=∆pmax. In case of
climb/descent the cabin pressure variates between zero
and∆pmax.
The test frequency was 10,000 flights per 24 hours. After
-6-NLR-TP-98148
prescribed numbers of flights the panel is inspected by
eddy current at the axial lap-joint and stringer run-outs.
The panel is also checked visually at the stringers,
frames, stringer-frame connections and load introduction
points of the frames. In accordance with the lifetime
predictions(5) no cracks or damages were found.
The fatigue test is followed by static tests. The GLARE
panel is subjected to one Limit Load case, two Ultimate
Load cases and a failure strength test. These static load
cases are intended to demonstrate that after two times the
design life (2*90,000 flights) and possible undetectable
cracks in the GLARE skin the residual strength is still
sufficient to carry Limit and Ultimate Load. During the
Limit Load case (∆p=∆pmax, My=My,Limit Load) and the
cabin pressure Ultimate Load case (∆p=2*∆pmax, My=0)
pillowing occurred and the panel expanded linearly
without permanent deformations. After the cabin pressure
Ultimate Load case the second Ultimate Load test
(∆p=1.5*∆pmax, My=1.5*My,Limit Load) was applied. No
final failure occurred and the strain distribution was,
except for plasticity which occurred above Limit Load,
conform the distribution during the Limit Load test. The
second Ultimate Load case is followed by the failure
strength test at∆p=∆pmax by increasing the axial load
until failure of the panel. Failure occurred at an axial load
equal to 1.32* axial Ultimate Load(6). This axial load is
15 % higher than the axial failure load expected
according to the theory(5).
Conclusion. The results of the GLARE fuselage panel
tests proved that the use of GLARE A in the crown
section of a Fokker 100 leads to a substantial weight
reduction without affecting the fatigue or static strength.
The full-scale fuselage panel test facility has shown to
induce a satisfactory radial expansion of the panel and a
uniform load introduction in axial and circumferential
direction.
Current tests
In the framework of the Brite Euram programmes
"Advanced concept for large primary metallic aircraft
structures" and " Structural maintenance of ageing
aircraft" NLR performs durability tests, crack propagation
tests and residual strength tests on several fuselage (side)
panels of Shorts, DASA and Alenia.
To demonstrate the technological feasibility of complex
GLARE panels with window cutouts, NLR will tests two
RJ130 GLARE side panels (1215 mm x 3032 mm,
R=1346 mm), designed and built by Shorts. The first
panel has GLARE stringers, the second panel has
extruded AL7150 stringers. During the fatigue test, at
least two lifetimes will be simulated by applying internal
air pressure, synchronized with an axial load spectrum. In
the course of these durability tests artificial damages will
be introduced. Finally, residual strength tests will be
done.
The objective of the tests on full-scale aluminum fuselage
panels of DASA and Alenia is to study the growth of
multiple site damage (MSD) in complex stiffened
structural joints in curved panels and to determine their
effect on the residual static strength. The test results will
be used to assess and to improve predictive models.
The aluminum ATR42 fuselage panel of Alenia
(1249 mm x 3032 mm, R=1432 mm) will have a lead
crack in the longitudinal lap-joint, with additional MSD
in adjacent fastener holes. After a limited number of
flights, to achieve sharp cracks tips at the artificial cracks,
residual strength tests will be performed. The applied
loads will be a combination of pressurization and axial
loads.
The two aluminum fuselage panels of DASA (1128 mm x
3030 mm, R=2820 mm) are cut from old A300 aircrafts.
One panel has an artificial lead crack with additional
MSD in the longitudinal lap-joint. The other panel has
only a lead crack. Like the ATR42 panel, the A300
panels will be subjected to a limited number of flights, to
extend the artificial cracks. During these flights only
cabin pressure will be simulated. Finally, residual strength
tests will be done.
Conclusions
The biaxial load introduction system of the full-scale
fuselage panel test facility showed to apply uniform loads
-7-NLR-TP-98148
in axial and circumferential direction.
Because of the excellent fatigue behaviour and sufficient
static strength combined with a weight reduction of 37 %
the designed GLARE panel proved the feasibility of
GLARE as fuselage material.
The full-scale fuselage panel test facility is developed to
have in case of fuselage design studies a test set-up
which offers the possibility to test single panels with
variable radius of curvature, panel width and panel length
at a high test frequency. The test performed on a Fokker
fuselage panel and the current tests on panels of Shorts,
DASA and Alenia indicate the interest of the aircraft
industry in testing curved structures under biaxial loading
conditions to evaluate fuselage design concepts, to
demonstrate the feasibility of fuselage materials and to
improve crack growth predictive models.
References
(1) De Jong, G.T., Elbertsen, G.A., Hersbach, H.J.C.
and Van der Hoeven, W., "Development of a full-
scale fuselage panel test methodology", NLR
contract report CR 95361 C, May 1995 (Restricted).
(2) Vercammen, R.W.A., Ottens, H.H., "Full-scale
GLARE fuselage panel tests", NLR report TP 97278
U, May 1997.
(3) De Jong, G.T., Botma,J. and Ottens, H.H., "Stress
analysis of the load introduction concept for the
fuselage panel test facility", NLR report CR 95142
C, May 1995 (Restricted).
(4) Jongebreur, A.A., Louwaard, E.P., and Van der
Velden, R.V., "Damage tolerance test program of the
Fokker 100", ICAF Doc. 1490, Pisa, May 1985.
(5) Wit, G.P., "Fuselage top panels tests (GLARE A)",
Stress Office Technical Data Sheet T.D.No avb7226,
March 1995.
(6) Vercammen, R.W.A., "Full-scale GLARE fuselage
top-panel test", NLR contract report CR 96301 C,
May 1996.
-8-NLR-TP-98148
Figu
re 1
Fus
elag
e pa
nel t
est f
acili
ty a
t NL
R
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78/3
3
botto
mbe
amte
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Avi
ew A
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view
B -
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stru
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top
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ator
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ator
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tsy
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nel
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whi
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-9-NLR-TP-98148
Figu
re 4
GL
AR
E p
anel
Figu
re 3
Ben
ding
of f
ram
e lo
adin
g ro
ds
coupling between frame and tangential loading system
strin
ger
1
1st b
ay
fram
e 1
Figu
re 2
Idea
l bou
ndar
y co
nditi
ons
∆R
6 ×
500
= 30
00
∆pan
el
∆1∆2
∆3∆4
∆5
due
to c
abin
pre
ssur
e
due
to a
xial
load
s
-10-NLR-TP-98148
ε (µ
stra
in)
1000 80
0
600
400
200 0
-200 -6
00-4
00-2
000
200
400
600
posi
tion
of s
train
gau
ges
(mm
)
R04
aR
07a
R10
aR
13a
R16
aR
19a
R22
aR
25a
Figu
re 6
Axi
al s
trai
ns d
ue to
axi
al lo
ads
Fax=
Fax(
bend
ing)
+Fax
(∆p)
ε (µ
stra
in)
3000
2500
2000
1500
1000 50
0 0 -600
-400
-200
020
040
060
0po
sitio
n of
stra
in g
auge
s (m
m)
R05
cR
08c
R11
c
R14
cR
17c
R20
cR
23c
R26
c
Figu
re 5
Tang
entia
l str
ains
dur
ing
1st U
ltim
ate
Loa
d te
st,
in th
e m
iddl
e of
the
pane
l
Fax=
Fax(
bend
ing)
+Fax
(∆p)
-600
600
Figu
re 7
Axi
al s
trai
ns d
urin
g th
e L
imit
Loa
d te
st, i
n th
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iddl
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the
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l
-400
-200
020
040
0po
sitio
n of
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in g
auge
s (m
m)
R04
aR
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R10
aR
13a
R16
aR
19a
R22
aR
25a
ε (µ
stra
in)
3500
3000
2500
2000
1500
1000 50
0
∆p=1
.0 b
arFa
x=0+
89 k
N∆p
=0.9
bar
Fax=
0+80
kN
∆p=0
.8 b
arFa
x=0+
71 k
N∆p
=0.6
bar
Fax=
0+54
kN
∆p=0
.4 b
arFa
x=0+
36 k
N∆p
=0.1
bar
Fax=
0+9
kN
∆p=0
.0 b
arFa
x=10
0 kN
∆p=0
.0 b
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x=80
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∆p=0
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x=60
kN
∆p=0
.0 b
arFa
x=40
kN
∆p=0
.0 b
arFa
x=20
kN
∆p=0
.0 b
arFa
x=0
kN
∆p=0
.534
bar
Fax=
290
kN∆p
=0.5
34 b
arFa
x=19
0 kN
∆p=0
.534
bar
Fax=
100+
48 k
N∆p
=0.3
0 ba
rFa
x=10
0+27
kN
∆p=0
.0 b
arFa
x=10
0 kN
∆p=0
.534
bar
Fax=
390
kN
Fram
e 1
posi
tion
ofst
rain
gau
ges
R05
cR
08c
R11
cR
14c
R26
cR
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R20
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Strin
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Faxi
al
Fram
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tion
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R04
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R10
aR
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R25
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R19
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Strin
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Faxi
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Fram
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tion
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aR
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R10
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R25
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R19
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Strin
ger 1
Faxi
al
Figu
re 8
Axi
al s
trai
ns d
urin
g L
imit
Loa
d te
st, b
etw
een
botto
m a
nd 4
th fr
ame
2000
500
1000
1500
posi
tion
of s
trai
n ga
uges
(m
m)
R43
aR
37a
R31
a131
123
110
R10
a
ε (µ
stra
in)
3500
3000
2500
2000
1500
1000 0
∆p=0
.534
bar
Fax=
290
kN∆p
=0.5
34 b
arFa
x=19
0 kN
∆p=0
.534
bar
Fax=
100+
48 k
N∆p
=0.3
0 ba
rFa
x=10
0+27
kN
∆p=0
.0 b
arFa
x=10
0 kN
∆p=0
.534
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Fax=
390
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500
R52
a
Fram
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posi
tion
ofst
rain
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ges
R10
a11
012
313
1R
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R37
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R52
a
Strin
ger 1
Faxi
al
-11-NLR-TP-98148
-600
-400
-200
020
040
0po
sitio
n of
LV
DT
's (
mm
)
LVD
T3
LVD
T6
LVD
T7
LVD
T9
LVD
T11
5 4 3 2 1 0
∆r (
mm
)
600
Figu
re 9
Rad
ial e
xpan
sion
of t
he G
LA
RE
pan
el
∆p=0
.4 b
arFa
x=0+
36 k
N∆p
=0.3
bar
Fax=
0+27
kN
∆p=0
.2 b
arFa
x=0+
18 k
N∆p
=0.1
bar
Fax=
0+9
kN∆p
=0.0
bar
Fax=
0 kN
∆p=0
.5 b
arFa
x=0+
45 k
N
Faxi
al
posi
tion
of L
VD
T's
LVD
T3
LVD
T5
LVD
T7
LVD
T11
LVD
T9
Strin
ger 7
+ ∆
r
- ∆r
Fram
e 1
Figu
re 1
0A
xial
load
on
the
GL
AR
E p
anel
for t
he e
ight
flig
ht ty
pes
Figu
re 1
1A
xial
stra
ins
durin
g 2n
d U
ltim
ate
Loa
d te
st, b
etw
een
botto
man
d 4t
h fr
ame
posi
tion
of s
trai
n ga
uges
(m
m)
1000
0
8000
6000
4000
2000
∆p=0
.534
bar
Fax=
670
kN∆p
=0.5
34 b
arFa
x=57
0 kN
∆p=0
.534
bar
Fax=
470
kN∆p
=0.5
34 b
arFa
x=37
0 kN
∆p=0
.534
bar
Fax=
270
kN∆p
=0.5
34 b
arFa
x=76
0 kN
ε (µ
stra
in)
2000
500
1000
1500
R43
aR
37a
R31
a13
112
311
R10
a
0
R52
a
Fram
e 5
posi
tion
ofst
rain
gau
ges
R10
a11
012
313
1R
31a
R37
aR
43a
R52
a
Strin
ger 1
Faxi
al
Fax=
Fax(
bend
ing)
+Fax
(∆p)
Load
(kN
)
300
250
200
150
100 50 0
Flig
ht ty
pe 1
, fre
quen
cy p
er 5
000
fligh
ts=1
250
200
150
100 50 0
Flig
ht ty
pe 2
, fre
quen
cy p
er 5
000
fligh
ts=3
Load
(kN
)
250
200
150
100 50 0
Flig
ht ty
pe 3
, fre
quen
cy p
er 5
000
fligh
ts=8
Load
(kN
)
250
200
150
100 50 0
Flig
ht ty
pe 4
, fre
quen
cy p
er 5
000
fligh
ts=2
8
Load
(kN
)
250
200
150
100 50 0
Flig
ht ty
pe 5
, fre
quen
cy p
er 5
000
flight
s=14
8
Flig
ht ty
pe 8
, fre
quen
cy p
er 5
000
flight
s=25
4820
0
150
100 50 0
Flig
ht ty
pe 6
, fre
quen
cy p
er 5
000
flight
s=67
620
0
150
100 50 0
Flig
ht ty
pe 7
, fre
quen
cy p
er 5
000
flight
s=15
8820
0
150
100 50 0
