,eries 07
Aerospace Materials 05
Residual Thermal Stresses around
Bonded Fibre Metal Laminate Repair
Patches on an Aircraft Fuselage
Residual Thermal Stresses around
Bonded Fibre Metal Laminate Repair
Patches on an Aircraft Fuselage
Bibliotheek TU DelH
11 lil
11111
~
C 312.121912.112392
325
Series 07: Aerospace Materials
05
\
.
Residual Thermal Stresses around
Bonded Fibre Metal Laminate Repair
Patches on an Aircraft Fuselage
A. V/otIT. Soerjantol/. YerilJ.A. Schelling
Published and distributed by:
Delft University Press Mekelweg 4 2628 CD Delft The Netherlands Telephone + 31 (0) 15 278 32 54 Fax +31 (0)152781661 e-mail: DUP@DUP.TUDelft.NL by order of:
Faculty of Aerospace Engineering Delft University of Technology Kluyverweg 1 P.O. Box 5058 2600 GB Delft The Netherlands Telephone +31 (0)15278 1455 Fax +31 (0)152781822 e-mail: Secretariaat@LR.TUDelft.NL website: http://www.lr.tudelft.nl
Cover: Aerospace Design Studio, 66.5 x 45.5 cm, by:
Fer Hakkaart, Dullenbakkersteeg 3, 2312 HP Leiden, The Netherlands Tel. + 31 (0)71 51267 25
90-407-1591-2
Copyright © 1998 by Faculty of Aerospace Engineering
All rights reserved.
No part of the material protected by this copyright notice may be reproduced or utilized in any form or by any means, electronic or
mechanical, including photocopying, recording or by any information storage and retrieval system, without written permission from the publisher: Delft University Press.
CONTENTS
1. Introduction 3
2. Test Set-up and Procedures 5 3. Results 8
4. Discussion 9 5. Conclusions 12 6. References 13
1. INTRODUCTION
The continued use of ageing transport aircraft stimulates the need of safe, damage tolerant and co st effective repair techniques. Bonded patches restore the strength and fatigue properties of the structure without creating new fatigue critical and difficult to inspeet areas as in the case of conventional riveted repairs. The optimal patch material must have a high strength and excellent fatigue properties. Further, the patch extensional stiffness may be not too high in order to minimize the disturbance of the load path as weil as the shear stresses in the adhesive. In addition, the thermal expansion mismatch with the aluminium structure must be small to minimize residual stress es after curing the adhesive, as weil as the induced thermal stresses caused by the ground-to-air cycle of the aircraft. Fibre Metal Laminates combine these properties. Patch materials with a low coefficient of thermal expansion, like carbon/epoxy and boron/epoxy, wil! develop high shear stresses in the adhesive layer, especially at cruising altitude (1). Fibre Metal Laminates are a family of fatigue resistant, damage tolerant aerospace structural materiais. They consist of thin aluminium alloy sheets adhesively bonded in alternating layers with fibre/ epoxy prepregs. A 3/2 lay-up for example, consists of three layers of aluminium bonded together by two layers of prepreg. ARALL is the name of Fibre Metal Laminates with aramid fibres, whereas the variant with S2-glass fibres is called GLARE. The US Air Force and Delft University of Technology are working together on a joint research program to evaluate the application of elevated temperature bonded Fibre Metal Laminate repair patches on the fuselage of ageing transport aircraft (1,2). The thermal effects of arealistic repaired fuselage are rather complex. The effective coefficient of thermal expansion (CTE) of the skin in the heated zone is smaller than of the aluminium alloy, because the area around the patch is only locally heated by a heat blanket. The surrounding cool structure will restrain the expansion of the skin in the heated zone. During curing of the adhesive, the patch is win expand but compressive stresses will be induced in the skin by the cool surrounding of the heated zone. When the skin and the patch have an equal CTE, the subsequent cooling to room temperature win create a tension in the patch which is balanced by a favourable crack elosing compression in the skin. When patch and skin have different CTE's, the sign of the stresses may be reversed. For most composite patches the effective CTE of the metallic skin win be higher than of the patch and unfavourable tension win be present at the crack in the metallic skin, balanced by compression in the patch after curing. Because of the small CTE mis match between aluminium and GLARE, the effective CTE of the skin becomes similar or even lower than the CTE of the patch. A favourable crack elosing stress may be present in the skin after curing. The cure cyele induces a static intern al stress which may lead to a mean stress fatigue effect, stress corrosion, distortion, and creep in the adhesive (3). At cruising altitude the whole fuselage including the patch will cool down to a temperature of approximately -55
o
e.
Wh en the patch has a lower CTE than the skin material (as for composite patches), the patch tends to open the crack and the effectiveness of the patch is reduced. This ground-to-air cycles impose a cyelic loading in the skin and patch and especially in the adhesive layer, which may lead to thermal fatigue problems. Minimizing the difference in CTE and lowering of the cure temperature between patch and skin material are therefore favourable, both for the internal stresses present after curing as wen as for the induced thermal cyeling of the repaired structure.Table 1. Typical properties of repair materials material CTEL (1O"6
/°C) CTET (10.6 IOC) EL (GPa) ET (GPa) 2024-T3 22.7 22.7 72.4 72.4 GLARE-2 3/2 0.2 16.3 24.5 69.0 54.0 GLARE-3 3/2 0.2 19.8 19.8 57.0 57.0 boronl epoxy 4.5 20.0 210 25 (unidirectional) carbonl epoxy -0.9 26 186 12 (unidirectional)
According to Baker (3) the following measures can be taken to mlDlmlze the residual
stresses: minimizing the size of the heated area, use of low temperature curing adhesives, pre-cu ring at a low temperature followed by post-curing of the adhesive, pre-stressing of the repair region, and a minimum thickness of the pateh. Pre-stressing into compression of the repair zone is not realistic for a fuselage skin. Missing in the list of Baker is a patch material with a small difference in CTE with the aluminium skin as can be achieved with GLARE! Table 1 shows typical values of the Young's moduli and the CTE of the aluminium skin and potential repair materiais.
This paper describes the results of experiments on a Fokker F28 fuselage section using the new material GLARE. The section was instrumented with strain gauges and thermocouples and tested in the laboratory. The internal strains which are present during and after cu ring were determined. Also the out-of-plane deflection of the skin during curing was recorded. The fuselage section was not pressurized.
2. TEST SET-UP AND PROCEDURES
A cylindrical F28 fuselage section between the wing and the cockpit was tested, see Figure 1. The length of the section is 3 mand its diameter 3.5 m. The skin is Al 2024-T3 c1ad with a thickness of 1.0 mmo The frame distance of the section is 500 mm and the stiffener pitch of the repaired zone is 170 mmo The stiffeners are both bonded and riveted to the skin.
The patches were bonded at the outside of the fuselage with 3M's AF-163-2K elevated temperature curing adhesive with a thickness of 0.13 mm, using the following surface preparatlon:
-removing paint with sandpaper, - c1eaning with Acetone,
- c1eaning primed patch surface with Acetone,
- lay-up of adhesive film and patch, using a roller to remove air bubbles.
A vacuum bag technique was used, with a pressure of 250 to 350 mbar, to create pressure on the adhesive. A thermocouple was installed at the cent re of the patch to
control the temperature during curing. Glass cloth was used over the patch as a breather. The square 300 x 300 mm heat blanket was spread over the cloth. The vacuum bag was sealed with tape. The pressure and temperature were monitored with a control unit. A heating rate of 10 °C/min. was applied. Insulation blankets were only used at the outside of the fuselage to cover the repaired zone.
Elliptical patches were bonded to three locations:
#1 on an uncracked skin in the cent re between two stringers and frames,
#2 on the same spot but with a crack (saw cut) in the skin, and #3 over a riveted/bonded lap joint (no crack).
Repairs #1 and #2 are the most unfavourable positions with respect to possible out- of-plane deformation. The position between stringers and frames is sensitive for dents due to accidental impact which need repair. Repair #3 is on a spot which is prone to fatigue damage. Thermocouples and strain gauges were installed on the skin at several positions in both longitudinal and circumferential direction of the fuselage. The thermocouples were attached with heat resistant tape. Several strain gauges and thermocouples were also installed at the outside of the patch.
Figure 2 shows the location of the thermocouples and strain gauges for repair #2 relative to the position of the patch. The strains gauges mounted on a line in the longitudinal direction of the fuselage we re also oriented in the longitudinal direction. Also gauges were bonded perpendicular to this direction, measuring the strains in circumferential direction. The exact location of the gauges and thermocouples for the three repairs are shown in the figures with the measured results. Figure 3 gives the cross section of the area of repair #3. A Al 2024-T3 shim was used under the patch to bridge the gap caused by the lap joint. The paint of the fuselage was removed with sandpaper at the position of the gauges and thermocouples before instrumentation. All data was stored and processed by computer.
The out-of-plane deflection of the skin was measured at the centre of the repair at the inside of the fuselage with an L VDT displacement transducer which was fixed to the floor of the fuselage section.
Table 2. Repair characteristics
no. patch material patch size cunng process location (mm) #1 GLARE-3 3/2 0.2 175 x 110 5 hrs. at 90
oe
- no crack [2024/0°/90°/2024/90% °/2024] 1 hr. at 120oe
- between t = 1.1 mm stringers and frames #2 GLARE-2 3/2 0.2 205 x 140 5 hrs. at 95oe
-over 42 mm [2024/0% °/2024/0% °/2024] saw cut t = 1.1 mm -bet ween stringers and frames -#3 - over longitudinal riveted/bonded jointThe procedure for the three repairs was, see also Table 2:
Repair #1 The patch material was cross-plied GLARE-3 in a 3/2 lay-up with 0.2 mm A12024-T3 layers and a total thickness of 1.1 mmo The repair position was in the centre
of two stringers and frames. The major axis of 175 mm of the elliptical patch was oriented in the circurnferential direction. No crack was present under the patch.
First a cure cycle was done up to 120
oe
-
i.e., the specified curing temperature of AF-163-2K - without adhesive to check the behaviour of the structure. Thermal bucklingaccompanied by a loud 'bang' occurred at 105
oe.
A deflection of 5.3 mm towards theoutside of the fuselage was measured at 120
o
e.
It was decided to pre-cure the adhesive at a lower temperature. The repair was cured under steady state conditions for 5 hrs. at 90
oe,
followed by a post-cure for 1 hr. at 120oe.
No sound was audible anymore at 105oe.
Repair #2 The patch material was GLARE-2 3/2 0.2 (see Table 2). This is a unidirectional GLARE variant in a 3/2 lay-up with 0.2 mm Al 2024-T3 layers. A saw cut was made in the skin in longitudinal direction with a length of 42 mmo The major axis of 205 mm of the elliptical patch was oriented in the circumferential direction and perpendicular to the saw cut. The rep air was cured for 5 hrs. at 95
oe
to avoidtemperatures lower than 90
o
e
in the adhesive.Repair #3 For this repair the same material, geometry and curing cyele was used as for repair #2. The patch and the shim were bonded in one cyele.
3. RESULTS
The following observations were made:Repair #1 Significant out-of-plane bending of the skin towards the outside of the fuselage occurred during curing at 90
oe,
accompanied by torsion of the adjacent stringers. At 90oe
3.4 mrn displacement of the skin and a stringer displacement sidewards of 3.0 mrn were measured. After cooling down to room temperature a 0.73 mm deflection of the skin appeared to be permanent.No difference was found between the residual strain after the pre-cure at 90
o
e
and after the post-cure at 120o
e.
Therefore the post-cure was not done for repairs #2 and #3.Repair #2 At 95
o
e
4.0 mrn displacement of the skin at the repair cent re was observed. After cooling down to room temperature 1.14 mm out-of-plane deflection remained.Repair #3 Both the riveted/bonded skins at the lap joint and the stringer which was located under th is joint (see Figure 3) cause a higher bending stiffness than the single skin of rep air #1 and #2. As aresult, the deflection for this repair was significantly smaller. A deflection of 2.64 mm was measured at 95
o
e
of which only 0.14 mm remained permanent at room temperature.The measured temperature distributions in longitudinal and circumferential direction are shown in Figure 4. A comparison bet ween both directions reveals that the temperature distribution is steeper in the circumferential direction than in the longitudinal direction for all three repairs. This is ascribed to the heat transfer caused by
the stringers. The difference in temperature distribution between repair #1 and #2 is probably due to a small difference in the position of the insulation blanket over the repair. Apparently the two joined skins of repair #3 cause a steeper heat transfer than of the single skin of rep air #1 and #2, as is reflected by the steeper temperature drop of repair #3, again especially in circumferential direction. Figure 4 also illustrates that the temperature of the skin underneath the heat blanket and therefore also under the patch in the adhesive is not constant. A larger blanket is needed for actual repairs of lap joints to create a more homogeneous temperature distribution in the adhesive. The long curing time used for the present tests assured a sufficient cure of the adhesive for our purpose. A larger blanket wiU probably cause out-of-plane deformation in a larger area.
4. DISCUSSION
Temperature distribution The measured temperature distribution was compared with the analytical solutions derived by Rose (I, 4). Rose modeIs a section between two stringers and frames by a circular plate with radius R which is heated in the centre with radius d
(=
heat blanket radius) at a constant temperature ~ (see Figure 8). For the temperature at the edge of the plate it can be derived:[1]
with T.mbienl as the ambient temperature (in our case 29°C, a hot Dutch summerl). The
constant m depends on the thermal conductivity of the skin and the natural convection
of the air. For an aluminium skin m is approximately:
[2]
with tp as the skin thickness in meters (also R en d should be substituted in Equation [1]
in meters). The temperature Tas function of the coordinate ris:
[3] fOT R~n:.d
fOT T:sd
The longitudinal temperature distribution was ca!culated with tp = 0.001 m, T.mbienl =
29 °c, ~ = 95°C (repair #2), das the heat blanket radius = 0.15 m, and R as the frame
di stance = 0.5 m. The measured distribution is below that of the model, see Fig.4 ..
Three reasons can be mentioned for this difference. The model only takes into account the heat transport in the plane of the skin. In reality a part <:>f the heat will be transferred to the surrounding air. Also the heat transfer by the stringers and frames is not accounted for in the model, and we already concluded that the effect of the stringers is significant. Finally the temperature distribution under the heat blanket is assumed to
be constant which is not realist ic. Despite this, the correspondence between the theoretica! model and the measured results is satisfactory.
model of Rose (1, 4), we will measure a strain distribution during curing which reflects the measured temperature field. This is clearly not the case. As can be seen in Figures 5 to 7, the stringers and frames act like hinges or nodes with respect to the bending deformation of the skin. The bending deformation is more pronounced in circumferential direction than in longitudinal direction because the frame distanee is larger than the stiffener pitch.
The bending deformation is mainly limited to the area under the heat blanket. The stringer und~rneath the zone of repair #3 is not able to prevent bending. After curing, a part of the bending remains present in the patch and the skin.
The residu al stresses after curing in the centre of the patch are calculated from the strains and are shown in Table 3. The experimental stresses are calculated using the biaxial Hooke's law, using the two strain gauges the cent re of the patch in both directions. This table shows that the patch stresses for #1 and #2 are approximately equal, whereas the skin stresses differ. This is caused by the influence of the saw cut in the skin which causes higher stresses. The smaller ben ding which occurs in repair #3 results in higher and more negative stresses in the skin and the pateh.
A comparison was made with a simple analytical model which is used to calculate the internal stresses of Fibre Metal Laminates based on the compatibility and equilibrium equations (5)~ This model will be close to reality if the effect of the cool surrounding is small and consequently the effective CTE of the skin is equal to the CTE of the aluminium alloy. The properties given in Table 1 and 2 are used to calculate the values of Table 3.
Table 3. Skin and patch residual stresses and strains at the centre of the repair
# locati expen- model
on mental
€longitudin~l €circumferential Ulongitudial o circumferentÎal Ulongitudinal a circumferenti:LJ (MP a) (MPa) (MP a) (MPa) #1 patch 0.000146 0.000017 9.5 3.8 -7.5 -7.5 skin -0.000007 0.000074 1.2 5.7 +8.3 +8.3 #2 patch 0.000145 -0.000011 10.7 1.9 -0.1 -14.2 skin -0.00023 0.00021 -13 11 +0.1 + 15.7 #3 patch -0.000008 -0.00078 -18.3 -28.6 skin -0.000034 -0.00026 -8.9 -21
which wiB be introduced by the internal pressurization of the fuselage. The sign of the experimental stresses is not consistent to those indicated by the model. The skin stress es in the skin of repair #2 are in the vicinity of the crack, and are therefore different from those of repair #1. The measured stresses of repair #3 over the lap joint are compressive stresses wheras repair #1 and #2 indicated mainly tensile stresses.
The bending deformation makes the situation rather complicated as can be illustrated by Figure 9. More research is needed to describe the thermal stresses accurately. Three contributions to the thermal stresses can be identified:
1. Permanent curvature introduced during curing. During heating up of the repair zone the skin will bend outwards. The patch will be pressed to this bended skin by the vacuum bag. The adhesive cures and during the cooling phase 'spring back' occurs and a part of the ben ding disappears. A part of the bending will be fixed· by the bonded patch and consequently a part of the corresponding stress es is permanent.
2. Thermal stress es due to cooling down after curing. Due to the mis match in CTE thermal stresses will be present after curing. The CTE of the aluminium skin is larger than of GLARE in fibre direction. Therefore we expect a compressive stress in the patch and tension in the skin. If the effective CTE of the skin is smaller due to constraints of the cooler zone, the stress state may be reversed. However, when a thermal bending as described under 1. occurs, then this restraint will be significantly smaller and the effective CTE of the skin will be close to the free expansion value of aluminium. In this case the situation shown in Figure 9 (2.) might be expected.
3. Bending as a consequence of the stresses described under 2. The distribution shown under 2. is not in equilibrium and will induce a second type of bending as indicated in Figure 9 (3.). Also this bending will cause tensile stresses at the outside of the patch and compressive stresses at the inside of the skin.
The simple model only takes the biaxial stress situation of 2. into account. The actual thermal stresses wiB be a summation of the three described stress states. From Figure 9, the two types of out-of-plane bending tends to relieve the in-plane thermal stress es caused by the CTE mismatch resulting in smaller thermal stress es than for a plane stress situation (2.).
5. CONCLUSIONS
1. Thermal buckling of the fuselage skin was observed at a cure temperature of 105°C of the F28 fuselage skin. Considerable bending deformation of the skin towards the outside of the. fuselage was observed accompanied with torsion of the adjacent stringers also at lower temperatures. An initial cure of the adhesive at a lower temperature followed by a post-cure at the prescribed temperature is recommended. Especially a rep air location with a thin skin without a stringer or frame underneath should be carefully treated. Significant out-of-plane deflection occurred for the repair zone on a joint with a stiffener underneath. Although for the last case most of this disappeared after cooling down to room temperature, it will certainly have an influence on the residu al stresses.
2. The temperature distribution predicted by the Rose model somewhat overestimates the measured distribution (max. 20%). A faster temperature drop was measured in circumferential than in longitudinal direction of the fuselage, which is due to the heat transfer by the stringers. The temperature distribution was not constant under the heat blanket which means that a larger heat blanket is required for complete curing.
3. The bending deformation causes a complex stress state in the rep air zone. Bending occurs during the heating phase and a second bending mode wiU be caused by the mismatch of the coefficients of thermal expansion of patch and skin. Both out-of-plane bending modes tend to relieve the thermal stresses compared to a two-dimensional stress situation. The resulting thermal stresses are small compared to the stresses induced by the flight loads (pressurization) of the fuselage.
4. Designers using the Rose model to analyze bonded repairs must take care, especially wh en such repairs are intended on a flexible region of the fuselage. In such case an assumption of more lirnited constraint may be appropriate.
6.REFERENCES
1. R.S. Fredell, Damage Tolerant Repair Techniques for Pressurized Aircraft Fuselages,
Dissertation Delft University of Technology, 1994
2. R.S. Fredell, W. van Barneveld and A. Vlot, Analysis of Composite Crack Patching of Fuselage Structures: High Patch Elastic Modulus Isn't the Whole Story, SAMPE
International Symposium 39, April 1994.
3. A.A. Baker, Fibre Composite Repair of Cracked Metallic Aircraft Components
-Practical and Basic Aspects, Composites, 18, (4), pp.293-308, 1987.
4. L.R.F. Rose, Residual Thermal Stresses, in: Bonded Repair of Aircraft Structures, A.A.
Baker and R. Jones (eds.), Dordrecht: Kluwer Academic Publishers, 1988.
5. C.H. Honselaar, Application of ARALL in the A320 Fuselage Shell Structure, MSc-Thesis, Delft University of Technology, 1988.
_.
-.
I-.
I Fra 6805 Fra 6305 Str 33=::i======ll=========t==\
Sir 32-t
I~-~.=====:I==========f=
iStr31-t
l
-.
~"#
-
'
-
-
=;=====+========r==
Str 30•
-i
-
LongitudinaI .' :-~ •• L .... ~._ .. _.+. ;' ___ • ___ • . ___ ._. ___ • ___ • __ -.,
... .
:h:;;:=:::~===J=======f=~=
Str 29 fuside CircumferemiaI frame STA 6805 Str30 Longitudinal ~ Straingages • Thermocouple OutsideFigure 2. Location of the strain gauges and thermocouples for repair #2
Patch
t=
1.1mm
205mm
103mm
FilIer
t=lmm
rivet (pitch
= 20 mm)
Skin (t= 1 mm)
120 100 ;
..
11
-4
-
-
--
::-
-
-
-repair#1 :: heat blanket repair #28
80 ~ ::I 60 iiî .... repair #3 Q) C-E 40 Q) I-20 ~ : : ... patch #1 :.
-
......
..---frame :patch #2 and #3 ..o
10 20 30 40 50 60 70 80 Longitudinal coordinate (cm) 120 I :.
.
I : 100 I: repair#18
80 ~.a
~ 60 / / / repair #2 / / ' repair#3 / ' //
/'
/
Q) c-E Q) 40 I-20 - heat blanket 0 , 0 10 20 30 40 50 60 70 80 Circumferential coordinate (cm)0.0002
I -__
~-~---;:======;-Ipatch (atter ~uring)
I
repair #10.0001
o
c: -0.0001 .~û5
-0.0002 . -0.0003 ~ patchskin (atter curing)
patch (during curing)
'.e ~---.
.
.
,. '.
-
.
I:' .:1 :' I I frame.
.
..
. ...•...
·~kin
(during curing'j'..•... ' .
-0.0004 ..1--_ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ - - - '
o
2.5 5 7.5 10 15 20 27 37.5 57.5Longitudinal coordinate (cm)
0.0006
patch (during curing) ! " - -repair #1 - ---' !
0.0004 ~/ patch (atter cu ring)
•
skin (during curing)0.0002 . ....
-'
-
,
....
/
/
~kin
(atter curing)c: 0 ïii ~ I·:
û5
-0.0002.
:-, "I : . '.
-0.0004 I ' i ~ patch -0.0006.
..
•
I stringer istringer -0.0008 0 2.5 5 11.5 15.5 20.5 25 33 45 58 Circumferential coordinate (cm)0.003
1 - - - = = = = = = : - - 1
I
repair#2
1
0.0025 0.002 patch frame c 0.0015.~ skin (after curing)
. I
û5
0.001 ' .. " skin (during curing)",
0.0005 patch (during curing) ,
... /
~.~atch (a~er
curing)~
o
r'Y'.
'
~."
;"
'~
'
...
.
.
_:"
"
.
"
'
....
~
...
~
"'
.
"""
i""
'"
-0.0005 -1.-_ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ ----1 -3.5 0 2.5 5 7.5 10 15 20 30 50 70 Longitudinal coordinate (cm) 0.0004 .'.
. . 0.0003 I: 0.0002 -0.0001 . :'.~
I:: ::-I stringerO
·
c .~ -0.0001û5
-0.0002 -0.0003 -0.0004,
-
:...
..•....... ~ ...•....'skin (after curing)
I '
.
.
"
'-patch (after curing) ....•.. ~ skin (during curing)
,
.
. I , . : I•
; :patch i stringer-0.0005 ~ patch (during curing)
-0.0006
0 2.5 5 11 15 20 25 35 49.5 58.5
0.003 , . - - - , c '(ij .... (jj c ...• 0.0025 '
~
0.002 . 0.0015 .. 0.001 0.0005 0 -0.0005 -0.001 0 i,
\ patch (during curing)
patch (after curing)
:patch
...
...
.
.
!~-frame I
skin (after cu ring) skin (d~ring curing)!
~ -w. I
,;
... . 4 8 14 20 Î' .• -.' I 28 35 Longitudinal coordinate (cm) 43 51 0.0008 .r - - - : - - - _ 0.0006 0.0004 0.0002o
I repair #3 i ", _ _ _ skin (during curing)skin (after curing)
•
..:....:....:....:....:...~--..
.. :--...•..~ -0.0002 " .• I ; stringer
I • 4" ~
'::J.i.-.L---
patch (after curing)~ :' !.~ patch (during cu ring)
/.... . U) -0,0004 . -0.0006 -0,0008 ..
....
..
r---patch
-0.001 j ... t - + -- -stringer -0.0012 - " - - - -_ _ _ _ _ _ _ _ _ _ _ _ - - 1 2.5 7 13 18 23 28.5 37 44 54 Circumferential coordinate (cm)Figure 7. Measured strains for repair #3
frame stringer y fram
tV
r
~x R,1
heated region1<
R,>1
Figure 8. Geometry of the Rose model to calculate temperature distribution (1)
1
2
3
adhesive
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+
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+
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+
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276 pages / ISBN 90-407-1577-702. C. Bill / F. van Dalen / A. Rothwell, 'Aircraft Design and Analysis System (ADAS)'
1997 / X + 222 pages / ISBN 90-407-1578-5
03. E. Torenbeek, 'Optimum Cruise Performance of Subsonic Transport Air-craft'
1998 / X + 66 pages / ISBN 90-407-1579-3
Series 03: Control and Simulation
01. J.C. Gibson, 'The Definition, Understanding and Design of Aircraft Handling Qualities'
1997/ X + 162 pages / ISBN 90-407-1580-7
02. E.A. Lomonova, 'A System Look at Electromechanical Actuation for Primary Flight Control'
1997 / XIV
+
110 pages / ISBN 90-407-1581-503. C.A.A.M. van der Linden, 'DASMAT-Delft University Aircraft Simulation Model and Analysis TooI. A Matlab/Simulink Environment for Flight Dyna-mi cs and Control Analysis'
1998 / XII + 220 pages / ISBN 90-407-1582-3
Series 05: Aerospace Structures and
Computional Mechanics
01. A.J. van Eekelen, 'Review and Selection of Methods for Structural
Reliabili-ty Analysis'
1997 / XIV + 50 pages / ISBN 90-407-1583-1
02. M.E. Heerschap, 'User's Manual for the Computer Program Cufus. Ouick Design Procedure for a CUt-out in a FUSelage version 1.0'
1997 / VIII
+
144 pages / ISBN 90-407-1584-X03. C. Wohlever, 'A Preliminary Evaluation of the B2000 Nonlinear Shell Element 08N.SM'
1998/ IV + 44 pages / ISBN.90-407-1585-8
04. L. Gunawan, 'Imperfections Measurements of a Perfect Shell with Specially Designed Equipment (UNIVIMP)
Series 07: Aerospace Materials
01. A. Vasek / J. Schijve, 'Residual Strenght of Cracked 7075 T6 AI-alloy Sheets under High Loading Rates'
1997 / VI
+
70 pages / ISBN 90-407-1587-402. I. Kunes, 'FEM Modelling of Elastoplastic Stress and Strain Field in Centre-cracked Plate'
1997 / IV + 32 pages / ISBN 90-407-1588-2
03. K. Verolme, 'The Initial Buckling Behavior of Flat and Curved Fiber Metal Laminate Panels'
1998/ VIII + 60 pages / ISBN 90-407-1589-0
04. P.W.C. Provó Kluit, 'A New Method of Impregnating PEl Sheets for the
In-Situ Foaming of Sandwiches'
1998 / IV
+ 28 pages / ISBN 90-407-1590-4
05. A. Vlot / T. Soerjanto / I. Yeri / J.A. Schelling, 'Residual Thermal Stresses around Bonded Fibre Metal Laminate Repair Patches on an Aircraft Fusela-ge'
1998 / IV + 24 pages / ISBN 90-407-1591-2
06. A. Vlot, 'High Strain Rate Tests on Fibre Metal Laminates' 1998/ IV + 44 pages / ISBN 90-407-1592-0
07. S. Fawaz, 'Application of the Virtual Crack Closure Technique to Calculate Stress Intensity Factors for Through Cracks with an Oblique Elliptical Crack Front'
1998 / VIII + 56 pages / ISBN 90-407-1593-9
08. J. Schijve, 'Fatigue Specimens for Sheet and Plate Material' 1998/ VI + 18 pages / ISBN 90-407-1594-7
Series 08: Astrodynamics and Satellite Systems
01. E. Mooij. 'The Motion of a Vehicle in a Planetary Atmosphere' 1997 / XVI
+
156 pages I ISBN 90-407-1595-502. G.A. Bartels, 'GPS-Antenna Phase Center Measurements Performed in an Anechoic Chamber'
1997 I X
+
70 pages I ISBN 90-407-1596-303. E. Mooij, 'Linear Quadratic Regulator Design for an Unpowered, Winged Re-entry Vehicle'
Residual stresses are present after elevated temperature cure of adhesively bonded patches to cracked aircraft structures. These residual stresses will affect the performance and the durability of the repair. Strain and temperature measurements were taken on a real aircraft fuselage section during and after repair at three locations. Thermal buckling of the skin was observed and therefore a precuring at a lower temperature was employed. Also at a lower temperature a significant outward bending of the skin was observed which influences the residual stresses. The measured temperature field corresponded weil with the theoretical model of Rose.
