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TRANSCRIPT
NASA Contractor Report 165839
NASA~CR~165839 19820018589
Develop, Delllonstrate, and Verify Large Area COlllposite Structural Bondirlg With Polyilllide Adhesives
Bashir D. Bhombal Donald H. Wykes Keith C. Hong Arnold A. Stenersen Rockwell International Downey, CA 90241
Contract NAS 1 .. 15843 May 1982
NI\S/\ National Aeronautics and Space Administration
Langley Research Center Hampton. Virginia 23665
111111111111111111111111111111111111111111111 NF01350
https://ntrs.nasa.gov/search.jsp?R=19820018589 2020-03-12T03:28:10+00:00Z
NASA Contractor Report 165839
Develop, Detnonstrate, and Verify Large Area Cotnp()site Structural Bonding With Polyirnide Adhesives
Bashir D. Bhomhal Donald H. Wykes Keith C. Hong Arnold A. Stenersen Rockwell International Downey, CA 90241
Contract NAS 1 .. 15843 May 1982
NI\SI\ National Aeronautics and Space Administration
Langley Research Center Hampton Virginia 23665
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FOREWARD
This is the final report on an effort to develop, demonstrate, and verify
large area composite structural bonding with polyimide adhesive. This
program was conducted by the Advanced Manufacturing Technology of Space
Transportation and Systems Group of Rockwell International, under Contract
NASl-15843, for. the Materials Division of NASA's Langley Research Center,
Hampton, Virginia. Robert M. Baucom was the NASA Technical Monitor.
The Program Manager during the Process Development phase was Arnold A.
Stenersen. Donald He Wykes assumed program management responsibilities
of Technology Demonstrator Segment and final report phases of the Contract.
Bashir D. Bhombal was the Technical Project Manager.
The following Rockwell personnel were principal contributors to this program.
J. D. Leahy Materials Technology
R. L. Crabb Materials Technology
D. W. Houston Chemical Testing
R. L. Long Specification Development
S. R. Graves Advanced System
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SECTIO~
1.0
2.0
3.0
CONTENTS
INTRODUCTION. • • • • lit • • • • • • • •
SCREENING OF ADHESIVE SYSTEMS •
2. I Selection Criteria for the Adhesives . . . 2.2 Industry Survey •• . . · . . ... . . . 2.3 Polyimide Resin Chemistry
Condensation vs. Addition Type PI Resins •
2.5 Adhesive Screening Tests • •
2.6 Lap Shear Test • • • • • • . . . . . 2.7 Flatwise Tensile Test
2.8 Climbing Drum Peel Test. . . . 2.9 Glass Transition Temperature •
2.10 Out-Time Study. · . . .
2
2
2
t" .)
6
7
B
2. II Mid-Plane Bonded Large Area Panels • 9
2.12 Evaluation of Primers and Surface Treatments. 9
2. 13 Selection of Three Adhesive Systems
PROCESS DEVELOPMENT . . . . . . . . . . . . . . 3.1
3.2
3.3
3.5
Celion/PMR-15 and Celion/LaRC-160 Adherends ••
Surface Treatment Study.
Primer Evaluation · . . . . . . . . . . . Processing of Adhesive • . . . Process for Mid-Plane Bonding Large Area • Panels
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1 I
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SECTION
4.0
5.0
6.0
7.0
CONTENTS
3.6 Process for Bonding Honeycomb Sandwich Panels.
3.7 Core-Coating of Paste Adhesives •••••
3.8 Aging Stability of Selected Adhesives ••
3.9 Out-time of Selected Adhesives •
3. JO Effect of Water Immersion on Lap Shear · , · . 3. 11 Final Selection of Adhesives · . . QUALIFICATION OF SELECTED ADHESIVES
4. 1
4.2
4.3
4.4
4.5
Process for Fabrication of Honeycomb Sandwich. Panels
Process for Mid-Plane Bonding Large Area Panels
· . Effect of Bond-Line Thickness on Lap Shear • • Strength
Effect of Out-Time on Filleting Properties • • of LaRC-I3
Evaluation of Catalysts to Reduce Cure • • • • Temperature
SUMMARY • • • • • • . . . . . . TECHNOLOGY DEMONSTRATOR SEGMENT • •
FABRICATION OF STRUCTURAL TEST ASSEMBLY · . . · · · · 7. 1 Prepreg Tape . · • 7.2 Leading Edge Cover Panel . . . . · · · · 7.2. 1 Skin . . • . . . • . . · · • · 7.2.2 Closeout Channel · · · • 7.2.3 Assembly of Leading Edge - Bonding Tool.
7.2.4 Prefit of Details
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CONTENTS
SECTION
7.2.5 Cleaning and Priming of Details •••••
7.2.6 Assembly of Details on Tool. • • • • • • • •• 26
7.3 Stability Rib 27
7.3.1 Stability Rib Sandwich Panel. • 27
7.3.2 Pi Cap Element . . . . . . .. " . . . . . . . . 28
7.3.3 Bonding of Closeout Channel and Pi Cap to Web. 28 Sandwich Panel
7.4 Lower and Upper Cover Panels. 29
7.4.1 Honeycomb Core 29
7.4.2 Skin ••• 29
7.4.3 Close-Out Edge Member • 30
7.4.4 Honeycomb Sandwich Assembly. 30
8.0 DEMONSTRATOR SEGMENT TEST PROGRAM 31
8. 1 Test Approach 32
8.2 Test Constraints. . .. . . . 32
8.3 Technology Demonstrator Segment NASTRAN Model. 32
8.4 Room Temperature and 5330 K Mechanical Tests 33
8.5 400 Cycle Simulated Fatigue Test. • 34
8.6 Thermal Cycling Test 34
8.7 Acoustic Fatigue Test • 35
8.8 Internal Pressure Test 3,5
8.9 Summary • . . . 36
DOCUMENTATION • . . . . . . . . . . . . . . 37
APPENDIX A • . . . . . . . . . . . . . . . . . . 115
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FIGURE
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ILLUSTRATIONS
LARC-13 Monomer Reactants ••••• . . . LARC-13 Polyimide Reaction Sequence
Adhesive Prepreg Mill • • • • • • • • •
Reaction Scheme of NR150B2G Adhesive . . . . . . . Flatwise Tensile Test Specimen • • •
Flatwise Tensile Test Specimen After Testing •••
Climbing Drum Peel Test Apparatus • • • • • •
C-Scan of 76 x 46 cm (30"x 18") Panel. 8 Ply, Unidirectional
C-Scan of 76 x 46 cm (31" x 18") Panel of Celion/ • PMR-15 Composite Panel. 8-ply, Unidirectional
C-Scan of Honeycomb Sandwich Panel with FM34B-18 Adhesive
C-Scan of Honeycomb Sandwich Panel with LARC-13 •••• Adhesive
Comparison of Primers for Bonding Celion 6000/PMR-15 •• Panels with LARC-13
C-Scan of 30.5 by 30.5 cm (12 x 12 in.) Mid-Plane Bonded Panel with LARC-13 Adhesive on Style 104 carrier
. . .
C-Scan of 30.5 cm (12 x 12 in.), Mid-Plane Bonded ••• Panel with M-LARC-13E Adhesive on Style 104 Carrier
C-Scan of 30.5 cm by 30.5 cm (12 x 12 in.) Mid-Plane Bonded Panel with FM34B-18 Adhesive on Style 104 Carrier
CUre Cycle for the Fabrication of Honeycomb Sandwich Panels with LARC-13 M-LARC-13E and FM34B-18 Adheives
ix
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FIGURE
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ILLUSTRATIONS
Technology Demonstrator Segment Completed Structure •
Staging Lay-up for Flat Laminates •
Autoclave Cure for Flat Laminates • . . . . . . . . . "u" Channel Lay-up Mold . . . . . . . . . . . . Staging Lay-up for "u" Shape Laminate • . . . . Curing Lay-up for "u" Shape Laminate . . . . . . . Alternate curing Lay-up for "u" Shape Laminate
Curing Lay-up for Honeycomb Sandwich Panel •• . . . . Cure Cycle for Cover Panels • • • • • • • • • •
Tool Configuration for Stability Rib Close-out. Channel
. . . Curved "pi" Joint Element Design . . . . . . . . . . Cure Cycle for Skins . . . . . Oven Post Cure Cycle for Skins ••••
C-Scan of 76.2 cm x 157.5 cm (30" x 62" Skin for Leading. Edge
Skin for the Leading Edge • • • • • • • • • •
Vacuum Forming of the Close-out Channels
Leading Edge Panel Bonding Jig. . . Leading Edge Bonding Assembly-Cross Section View. •
Leading Edge Cover Panel on Bonding Jig •
Leading Edge/Cover Panel
Primer Being Applied on Rib Web for Bonding of "U". Channel
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FIGURE
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ILLUSTRATIONS
Primer Being Applied to the Inner Section "U" Channel
"U" Channel Being Inserted to Web Sandwich Panel • • . . . "U" Channel Complete for Bonding. • • . . " . . . . . Drilling Fixture for Bonding Close-out Channel and Pi. • • Cap to Web Sandwich Panel
CloBe up View of Forward Section of the Fixture ••
GrlPI Composite Body Flap Technology Demonstrator •• Segment (TDS)
GriP! Technology Verification TDS (All Bonded Structure) •
Demonstrator Segment Test Concept Constraints ••••
Technology Demonstrator Segment NASTRAN Model •••
R.T. Mechanical Load Test . . . . . . . . . . . . . . . ~
Typical Cover Panel Stress Contours . . . . . . . . . High Temperature Mechanical Load Test. • . . . . . . Simulated Fatigue Test • . . . . . . . . . . . Thermal Cycle Test • • • • • • • • • . . . . . . . . . Acoustic Fatigue Test . . . . . . . . . . . . . . . . Internal Pressure Test • . . , . . . . . . . . . .
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TABLE
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IS
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TABLES
Lap-Shear Strength of Polyimide Adhesives • • • • •
Flatwise Tensile Strength of Polyimide Adhesives
Climbing Drum Peel Strength of Polyimide Adhesives •••
Glass Transition Temperature of Polyimide Adhesives
Effect of Out-Time on Lap-Shear Strength of Polyimide • Adhesives
Effect of Out-Time in Humid Environment of Lap-Shear Strength
Comparison of Primers for Bonding Celion 6000/PMR-15 Panels with LARC-13 Adhesive
Comparison of Surface Treatment for Bonding Calion/ PMR-15 Laminates with LARC-13 Adhesives
Comparison of Polyimide Adhesives • • • • • • • • • • •
Comparison of FM34B-18, LARC-13 and LARC-13E Adhesives-
Properties of Celion 6000/PMR-15 Prepreg and Cured •• -Laminates
Flatwise Tensile Strength of FM34B-18 LARC-13 and ••• M-LARC-13E Adhesives
Lap-Shear Strength of LARC-13 Adhesive Prepared by •• Anhydride and Ester Methods
Short Beam Shear Strength of LARC-13 and FM34B-18 •••
Slotted Shear Strength of Test Specimens from Mid- •• Plane Bonded ·30.5 by 30.5 em Panels
Aging Stability of Polyimide Adhesives Wrapped in • • • Aluminum Foil
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TABLE
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TABLES
Aging Stability of Polyimide Adhesives Not Wrapped •••• in Aluminum Foil
Aging Stability of Selected Adhesives . . . . . . . Effect of Water Immersion on Lap-Shear Strength of •
LARC-13 and FM34B-18
Effect of Bondline Thickness on Lap-Shear Strength •
Out of Refrigeration Effect of Filleting Properties. of LARC-13
Use of Catalyst in LARC-13 to Reduce Cure Temperature •
Orbiter Body Flap/Demo Segment Stress Comparison
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1.0 INTRODUCTION
Contract NASI-15843 is a NASA La.RC Program under Project CASTS
(Composites for Advanced Space Transportation Systems). This project
was established to develop and demonstrate the technology required to
produce graphite/polyimide structural components with operational capa
bility at 5890 K (6000F).
The CASTS project includes several current research contracts to
develop and demonstrate processes for fabricating graphite/polyimide com
posites materials and nondestructive inspection (NDI) of structures. The
fabrication of structural elements in these studies includes flat laminates,
honeycomb panels, skin stringer panels and chopped fiber moldings. These
elements are a.ssembled and joined primarily by adhesive bonding. Techniques
for assembling and bonding the structural elements into a completed structure
were developed in this program.
The objective of this contractual study program was to develop pro
cesses for joining large area structural subcomponents by bonding with
polyimide adhesives, to verify the reliability of the adhesive bonds by
mechanical and nondestructive testing and to demonstrate the quality of
the large area bonding technique by fabrication of a large completed stru
cture. The original plan to fabricate a box beam as a demonstration com
ponent was changed to the fabrication of components for a deliverable struc
tural test assembly representative of the Space Shuttle aft body flap.
Identification of commercial products in this report is to adequately
describe the materials and does not constitute official endorsement, express
ed or implified, of such products or manufacturers by the National Aero
nautics and Space Administration.
The NASA "A" standard to ultrasonic evaluation of composites referenced
on p. 10 is a sensitivity level that is utilized at NASA as an indication
1
of high quality, voidfree composites. Rockwell used the NASA "A" standard
as a point of reference only in this report since this standard is not
universal in the art.
2.0 SCREENING OF ADHESIVE SYSTEMS
2.1 S~lection Criteria for Adhesives
The criteria used in selecting adhesive systems suitable for
(I) bonding large area Celion graphite reinforced PMR-15 polyimide
laminates; and (2) fabricating large area honeycomb and skin/stringer
panels with Celion 6000/PMR-15 or Celion 6000/LARC-160 face sheets
were:
o Commercial availability
o Existing data base
o Processability
Resin processability
Processability of supported film adhesive (adhesive
prepreg)
Processing and cure of primer and adhesive 5890 K (600oF)
and 1.379 MPa (200 psi) nominal upper processing limits).
o Handleability
o Stability under normal storage conditions
o Filleting capability
o Low areal weight
o Mechanical properties
Lap shear strength
Climbing drum peel strength
Flatwise tension strength
o Ability to form low porosity bond lines in mid-plane bonded
panels.
2.2 Industry Survey
2
An industry survey was made to assess the status of development
of polyimide adhesive with operational capability at 5890
K (600oF).
The data obtained in this survey indicated that the LARC-13 adhesive,
developed by NASA-LARe, was the most promising polyimide adhesive avail
able for large area bonding. Other potential polyimide adhesives with
operational capability at 5890 K (6000 F) included several systems based
on dUPont's NR-150B2 polyimide resins, the FM34B-18 adhesive and
adhesive systems based on the PMR-15 and LARC-160 polyimide resins.
The FM34B-18 adhesive is commercially available from American Cyanamid
Company both in paste and scrim-supported film form. The NR-150
polyimide resins are available in solution form at resin solids con
centrations of about SO percent. The LARC-13, LARC-160, and PMR-15
polyimide resins are produced from monomer reactants. Most of these
monomer reactants are readily available from several commercial sup
pliers.
For production of scrim-supported film adhesives, based on these
polyimide resin systems, a hot-melt process and a solvent evaporation
process had been evaluated at Rockwell. The hot-melt process was found
to be the most suitable of these methods in regard to production rate,
film quality, film uniformity, and reproducibility. As a result of
the industry survey, the hot-melt polyimide adhesive systems listed
below were selected for initial screening.
o FM34B-18
o LARC-13
o Mod LARC-13
o Mod LARC~ 13
o NR-lS0B2
o PMR-15
o Mod PMR-15
American Cyanamide Company
NASA Process
Use of methyl and ethyl esters of BTDA and
NA used in place of anhydrides, and flex
ibilizers such as silicon and rubber materials.
Use of amine substitutes by replacing part or
all of 3,3' MDA compared with the armomatic
diemines such as 4,4' MDA and 3,3' and 4,4'
DADS (Diamino di Sulphone).
Linear and cross-linked systems duPont
NASA formulation prepared by Rockwell
Use of amine substitutes by replacing part or
all of 3,3' MDA compared with other aromatic
drainides such as 4,4' MDA and 3,3' and 4,4'
DADS.
3
o LARC-160
o Mod LARC-160
NASA formulation, prepared by Rockwell
Use of flexibilizers such as silicon and
rubber materials.
2.3 Polyimide Resin Chemistry
4
The LARC-13 addition-type polyimide polymer is formed by reactions
of benzophenone tetracarboxylic dianhydride (BTDA), nadic anhydride
(NA), and 3;3'-methylenedianiline (3,3'MDA). This polymer also can
be formed by reacting the methyl or ethyl esters of BTDA and NA with
3,3'-MDA. The chemical structures of these monomer reactants are
shown in Figure I.
When the anhydrides, BTDA and NA are used, the sequence of the
polymer reactions generally is reported to take place in a simplified
manner as shown in Figure 2. At first, 3,3'-MDA reacts with PMDA
nadic groups. These reactions are followed by the imidization inter
mediate. This imidization reaction is generally carried out on
staging of the adhesive prepreg at normal pressure and temperatures
from 422oK(300oF) to 4490 K (350°F). On application of higher tempera
tures, S330 K (500°F) to S890 K (600°F), and increased pressure, addition
type reactions occur at the norbornene groups. These addition-type
reactions, reverse Diels-Alder reactions, result in chain extension
and cross-linking.
When the methyl esters of pMDA and NA are used in place of the
anhydrides, these esters at first react with 3,3'-MDA to form amide
acid and methanol. These reactions are followed by imidization, chain
extension, and cross-linking reactions, as shown in Figure 2.
The PMR-IS and LARC-J60 resins are similar to LARC-J3 in chem-
ical structure and polymer reactions. Several other polyimide resins
evaluated also contained PMDA, and NA as monomer reactants along with
different aromatic diamines such as diaminodiphenyl sulfone (DADS).
Initially, the LARC-13, LARC-160 and PMR-15 resins used for the
polyimide adhesives were processed with methyl esters of BTDA and
NA and an aromatic diamine used as monomer reactants. Later the
anhydrides, BTDA, and NA were used as monomer reactant,s with armomatic
diamines.
The methyl esters of BTDA and NA were prepared by refluxing a
mixture of these andydrides with methanol for two hours. The aromatic
amine then was added in several portions and mixed for one-half hour.
Filler and other additives then were added and mixed at room tempera
ture.
The adhesive prepreg, or film adhesive was processed on a labor
atory scale with use of a coating device as shown in Figure 3. Glass
fabrics were used as carrier materials for the film adhesives.
2.4 ,Condensation vs. Addition Type P.I. Resins
The PMR-15 and LARC-160 resins are similar to LARC-13 in chem
ical structure and polymer reactions. Several other polyimide resins
evaluated also contained PMDA and NA as monomer reactants along with
different aromatic diamines such as diaminodiphenyl sulfone (DADS).
The NR-150B2G and NR-056X polyimide resins are condensation
type polyimide systems structured as shown in Figure 4. The poly
mers are formed by reacting 4,4' (Hexafloroisopropy lidene)-bis
(phthalic acide) ("6F tetra acid) with 4,4' phenylene diamine and
3,3' -phenylene diamine. Water is split off as a condensation product.
The PMR-15 (II) on the other hand can be classified as an
addition type resin. It contains 4,4' (Hexafloroisopropy lidene)-
bis (phthalicacide) and phenylenediamine, as does NR1S0-B2G but in
addition, it contains NA and end-capper. The reaction sequence for
this monomer reaction is similar to that of LARC-13. At first amic
acids is formed which on heating to a temperature such as 4500
K (350o
F)
5
is converted into a po1yimide prepo1ymer. On further heating at
higher temperature. such as 5890
K (600oF) and application of pressure
such as 100 psi, chain extension and cross-linking takes place at the
norbornene groups as for LARC-13. The chain extension and cross-link
ing reactions are addition type reactions by which no condensation
products are formed. Therefore, the PMR-15 (II) resin solution also
should be suitable for large area bonding.
2.5 Adhesive Screening Tests
6
The tests used for screening of the selected po1yimide adhesives
are shown below. The initial tests included a lap shear test (Table
1) and a flatwise tensile test (Table 2) at R.T. and 5890 K (600°F)
and a climbing drum peel test (Table 3) at R.T. The glass transition
temperature (Tg). the storage stability or out-time, and ability to
form low porosity, large area bonded also were evaluated for the most
promising adhesives.
Lap-shear strength at R.T.
and 5890K (600°F)
F1atwise tensile strength
o ° at R.T. and 589 K (600 F)
Glass transition temperature
Tg.
12.7 mm (0.5 in.) overlap specimen
of unidirectional, 8 ply, 1.02 to
1.115 mm (0.040 to 0.045 inch) thick
composite laminates; Fed. spec.
MMM-132A.
5.1 by 5.1 cm (2 by 2 in.) sandwich
panels of HRH 327-3/16-4 T = 1.27
(0.5) honeycomb with 8 ply composite
face-sheet MIL-A-25463A.
6.25 mm diameter 1.02. to 1.14 mm
(0.040 to 0.045 in.) thick adhesive
panels.
Storage stability of adhesive
film prior to use (out-time)
by lap-shear tests.
12.77 mm (0.5 in.) overlap spE~cimen
of unidirectional, 8 ply, 1.02 to
1.14 mm (0.040 to 0.045 inch) thick
composite laminates. Fed Spec.
MMM-132A.
Ability to form low poro.sity 15.2 by 1.52 cm (6 by 6 in.) com-
bond lines in Mid-plane bondedposite panels and steel panels,
panels. T = 0.12 (0.5).
2.6 La}, Shear, Test
The adhesive systems screened for lap-shear strength to Celion/
PMR-15 laminates and the test data obtained are shown in Table I.
A unidirectional, 8 ply laminate of Celion 6000/PMR-15 composite of
1.02 to 1.14 mm (0.040 to 0.045 inch) thickness was selected for this
test to provide strain compatibility (equal strain under a given load)
with one inch wide titanium adherends. (Titanium finger panels, 25.4
mm wide and 1.27 mm thick (I inch wide and 0.050 inch thick), of 6AL-
4V titanium, had been used in several research and development pro
grams at Rockwell for evaluations of polyimide adhesives for use in
the Space Shuttle). The selected thickness of the Celion/PMR-15
composite was obtained from the following relationship:
Acomp x Ecomp :: ATi x ETi
Acomp :: ATi x ETi
Ecomp
Acomp = ( I x 0.05) (15 x J06)
18 x 106
Acomp = 0.042
x T = 0.042
'I~ = .042 :: 1.07 mm
7
8
Where
Acomp = (I x T) = Cross sectional area of one inch wide
composite lap shear specimen.
ATi = ( I x 0.05) = Cross sectional area of one inch wide by
0.05" thick titanium lap shear specimen.
Ecomp = 18 x 106 = Modulus of Celion/PMR-15 composite.
ETi = 15 x 106 = Modulus of 6AL4V titanium.
T = Thickness of one inch wide Celion/PMR-15 composite.
2.7 Flatwise Tension Test
A series of adhesives that appeared promising in the lap-shear
test were evaluated further by a flatwise tensile test (Table 2)
using test specimens as illustrated in Figure 5 and 6.
2.8 Climbing Drum Peel Test
A climbing drum peel test, according to specification MIL-A-
2546A, Figure 7 was selected as a measure of the peel strength of
the polyimide adhesives. This test was used primarily to provide
data for design and handleability of honeycomb sandwich components
to be fabricated. The target value set for the climbing drum peel
test of the polymide adhesives was 900 g/cm width (5 in. lb/in.) at
R.T. (Table 3).
2.9 Glass Transition Temperature
Measurements of the glass transition Tg. were made for a series
of adhesives using a Dupont 941 TMA-900TA analyzer. The adhesive
samples were fabricated into laminates of about 1.14 mm (0.045 in.)
thickness. The samples were cured for two hours at 5890 K (600oF)
and postcured for ten hours at 5890 K (600oF). The test data is
shown in Table 4.
2.10 Out~Time StudX
A series of six adhesives were evaluated for storage stability
over an II-day period. The relative humidity during this period
varied between 6 and 40 percent and the temperature varied between
291 and 2990 K (65 and 80°F). The test data obtained as shown in
Table 5 shows that the lap-shear strength of the six adhesives was
not affected by exposure to this dry environment. A similar study
was conducted when the relative humidity in the storage area was 40
90 percent. The lap shear data obtained as shown in Table 6 shows
that the strength values of FM34B-18 adhesive are significantly
reduced after the three and seven days out-time periods. In com
parison, the lap-shear properties of the LARC-13 and M-LARC-13
adhesives were not detrimentally affected by these exposures. The
results show that exposure of the FM34B-18 to a humid environment
must be avoided prior to application.
2.11 Mid-Plane Bonded Large Area Pan!.!.:!
Several of the selected polyimide adhesives that looked prom
ising by the lap-shear climbing drum peel and the flatwise tensile
tests were screened for ability to produce low porosity bond-line
in mid-plane bonded panels. The dimensions of the panels used were
t.52 by 15.2 cm (6 by 6 in.). The porosity in the bond-line was
inspected by ultrasonic c-scanning, using the NASA "A" sensitivity
standard.
2.12 Evaluation of Primers and Surface Treatments
The primers evaluated in the initial experiment consisted of an
alcohol/diglyme solution of condensation and addition type polyimide
adhesive. The use of primers resulted in increased bond strength,
apparently because their polymer chains are linear, more flexible,
and tougher than the cross-linked addition type polymers. The
result is increased load carrying capacity; as shown in Table 7 and
Table 8.
2.13 Selection of Three Polyimide Adhesive Sxstems
Table 9 shows a comparison of availability and physical prop
erties for six of the most promising adhesive systems. Based on
this data, the following three adhesive systems were selected for
9
further study and evaluation:
LARC-13 - Produced by NASA process
Modified LARC-13 - Partial substitution. of 3,3' MDAwith
4,4' MDA
FM34B-18 - American Cyanamid Company
A more in-depth comparison of the three selected systems is shown in
Table 10.
The laminate surface treatment selected was light abrading follow
ed by a diglyme wash. BR34B-18 was chosen as the laminate primer
through lap shear tests as shown in Table 7.
3.0 PROCESS DEVELOPMENT
3.1 Celion/PMR-IS and Celion/LARC-160 Adherend
10
All of the Celion/PMR-IS test panels used for evaluation of the
polyimide adhesives were fabricated using procedures developed in
Contract NASI-ISI83 (Design, Fabrication, and Test of Graphite Poly
imide Structural Elements and Specimens).
The composite materials fabricated for the lap shear tests were
unidirectional eight-ply panels of Celion 6K/PMR-IS prepreg tape.
The prepreg tape was furnished by U.S. Polymeric and Fiberite Corp
oration.
The quality control tests used for these lap shear panels in
cluded specific gravity, fiber volume, fiber weight, and void fraction.
The test data obtained for a series of 76.2 by 4S.7 cm (30 by 18 in.)
panels, fabricated for this test, are shown in Table II. These panels
also were inspected by ultrasonic c-scanning. The c-scan recordings
show that all of the panels met the NASA "A" sensitivity standard with
few detectable defects as indicated for the panels in Figures 8 and
9.
The flatwise tensile tests panels were fabricated using eight
ply prepreg with symmetrical (0, +4S, -45, 90)s construction. The
test data for these panels are included in Table 12. The c-scans
of these panels also met the NASA "A" sensitivity standard as
illustrated in Figure 10 and 11.
Two composite panels used for the climbing drum peel tests were
fabricated using two plies of Celion 6K/PMR-15 in a (O,90-degree)
lay-up constructions. The cured thicknesses of these laminates were
about 0.2 rom (O.OOS in.). A four-ply laminate of symmetrical con
struction (0, ! 45, 90) of Celion 3K/LARC-160 tape also was fab
ricated for this test. After cure the two ply laminates curled up
into a cylindrical form while the symmetrical laminates retained
their flat shape.
3.2 Surface Treatment Studz
A series of seven surface treatments were evaluated by lap
shear tests for bonding Celion 6000/PMR-15 composite laminates:
with LARC-13 adhesive. The BR34B-18 material was used as primer for
all adherends in this study. The surface treatments used and the
test results obtained are shown in Table 8. The highest bond
strength values were obtained with light grit blasting folloWE~d by
a diglyme wipe.
3.3 Primer Evaluation
The standard lap shear test used for screening of the adhesives
was used for evaluation of a series of primers. The BR34B-18 primer,
used in all of the screening tests, was used for comparison. Table
7 shows the primer systems evaluated and the test results obtained
at RT and 5890 K (600°F). A bar chart for comparison of the test
values is shown in Figure 12.
The data shows that the highest bond strength values at RT
were obtained with the linear polyimide primers, BR34B-18, NR-150B2,
and 050X, respectively. At 5890 K (6000 F), the highest bond values
were obtained with NR-150B2, PMR-15E, and BR34B-18 primers, respect
ively. Since BR34B-18 was among the best primers evaluated and is
11
12
a comercially available product; it was selected for use.
3.4 Processing of Adhesive
The LARC-13 adhesive can be produced either by the standard
anhydride method developed by NASA or by an ester method. In the
anhydride method, the amine component, 3.3' MDA, is first dissolved
in the DMF solvent. Then the anhydride components, BTDA and NA, are
added gradually to the amine solution and stirred until they are
dissolved.
By the ester method; the BTDA and NA components are first con
verted in methyl or ethyl esters. Then the amine component is added
and mixed until it is dissolved.
A study to evaluate and compare the safety and economics of
these methods was conducted both at Hexcel and Rockwell. Hexcel
found the ester method to have the following advantages over the
anhydride method for large scale production: (1) better control of
the advancement of the resin mix, (2) a shorter mixing period, and
(3) less toxic solvents. Therefore, tests were conducted to compare
the lap shear strength of the LARC-13 adhesive prepared by the two
processes. The results in Table 13 show that there is not sig
nificant difference in lap shear strength resulting from the two
processing methods. The only advantage observed for the anhydride
process was slightly better flow of the adhesive in cure.
3.5 Process for Mid-Plane Bonding Large Area Panels
Processes for mid-plane bonding 30.5 by 30.5 em (12 by 12 inch)
composite panels were developed. The porosity in the bond lines
was examined by c-scanning, using NASA standard "A" sensitivity.
Void free or near void free bond lines were obtained using several
bonding me~hods for LARC-13 and M-LARC-13E film adhesives with style
112 carrier. The simplest of these processes had the following cure
cycle:
1 • Pull full vacuum and apply 100 psi pressure
2. Heat to 6000 F at 50 F/min.
3. Cure at 6000 F for 1 hour
4. Cool to 1500 F before r.eleasing pressure and vaCUum
The c-scans obtained for the LARC-13 and M-LARC-13E adhesives
with thi.s cure cycle, Figures 13 and 14, show practically void free
bond lines. The best c-scans obtained in mid-plane bonded 30.5 by
30.5 cm (12 by 12 inch) composite panel with FM34B-18 is shown in
Figure 15.
A study of shear beam and slotted shear was done, which was
prepared from mid-plane bonded panels. They are shown in Tables
14 and 15.
3.6 Process for Bondins Honeycomb Sandwich Panels
The use of scrim-supported film adhesive resulted in good
filleting and high flatwise tensile strength values for each of the
three selected adhesives with the porous HRH 327-3/16-4 core, as
shown in Table 12. In this process, the core and the Celion 6000/
PMR-15 face sheets were at first cleaned and primed with the 13R34B-18
primer. Then a layer of the adhesive film, having an areal weight
of 293 g/m2 (0.06 lb/sq. ft.), was applied on each side of the core,
permitting the panel to be cured in one operation. A suitable cure
cycle for the FM34B-18, LARC-13 and the M-LARC-13E adhesive is
shown in Figure 16. In this cUre cycle, a maximum pressure of
2.75MPa (40 psi) and a maximum temperature of 5890 K (600oF) were used.
The application of lower pressures around 1.72 MPa (25 psi) produced
good panels while the use of 3.45 MPa (50 psi) pressure resulted
in slight damage to the honeycomb core.
3.7 Core Coating of Paste Adhesives
The pl.rpose of applying the polyimide paste adhesive by a core
coating process is to optimize distribution of the adhesive at the
core-face sheet interface, thereby reducing the required adhesive
weight. This process requires proper surface tension and thixotropic
13
14
properties for the adhesive. To evaluate the core-coating ability
of the LARC-13 and M-LARC-13E adhesives, (FM34B-18 was available
as a supported film only) experiments were conducted with rubber
paint rollers and a laboratory coater. It was apparent that the
surface tension and the viscosity of the adhesives were too low to
produce sufficient quantity and uniform distribution of the adhesives
in one application, as required for an automated core-coating process.
However, good distribution of the adhesives was obtained using a
rubber paint roller and three or four layers of the adhesive.
3.8 Aging Stability of Selected Adhes.ives
A study was undertaken to evaluate the aging stability of
LARC-13 and FM34B-18 adhesives. Honeycomb sandwich panels were
fabricated and flatwise tensile specimens 5.08 em (2") x 5.08 em
(2") were machined. Half of these specimens were wrapped in alumi
num foil and aged at 5890 K at (600oF) for 125 hours in an air circu
lating oven. The other half of the specimens were exposed directly
to the hot air stream in the oven. It was observed that the flat
wise tensile strength decreased on aging, but the decrease was much
smaller for the specimens wrapped in aluminum foil than for those
subjected directly to the air stream. The results are shown in
Table 16 and 17.
An evaluation of the physical properties of the three adhesives
was continued with an aging study of lap shear specimens at 5890 K
(600oF) using Celion 6000/PMR-15 adherends. The test data obtained
after exposure of the lap shear specimens in an air circulating oven
at this temperature is shown in Table 18. The values for the three
adhesives were reduced significantly by exposure at 5890 K (600oF)
for periods of 125 and 250 hours. However, the lap shear specimens
failed partly in the laminates, indicating that the aging stability
of the adh~sives is as good as that of the laminates.
3.9 Out-Time of the Selected Adhesives
An out-time study was made of the FM34B-18, LARC-13, and the
M-LARC-13E adhesives for periods up to II days in a very dry atmos
phere (6-40 percent RH). The effect of this storage environment was
evaluated, using lap shear tests with Celion 6000/PMR-15 adherends.
None of the three adhesives were significantly affected by storage
in the dry atmosphere prior to the bonding operation. To determine
effects of out-time exposure in a humid environment, a similar study
was conducted during the month of December, when the relative
humidity in the storage area was 40 - 90 percent. The lap shear
data obtained (Table 6) shows that strength values of the FM34B-18
adhesive are significantly reduced after the three and seven day
out-time periods. In comparison. the lap shear strength properties
of the LARC-13 andM-LARC-13E adhesives were not detrimentally
affected by these exposures. The results show that exposure of the
FM34B-18 adhesives to a humid environment must be avoided prior to
application.
3.10 Effect of Water Immersion or Lap-Shear Strength
The data in Table 19 shows that the lap-shear strength of
LARC-13 and FM34B-18 adhesives are only slightly affected by water
immersion at RT for 7 days.
3.11 Final Selection of Adhesives
The selection of an adhesive for fabrication of a structural
test assembly was made by a comparison of 14 factors as sho~~ in
Table 10 for the FM34B-18, LARC-13 and the M-LARC-13 adhesives.
These factors included commercial availability, material cost,
out-time, tack and drape, filleting, tooling and processing cost,
mechanical and adhesive properties. Each factor was given a
significance rating of I, 2, or 3. By this comparison, the LARC-13
adhesive had the highest point rating and was selected for qual
ification tests.
4.0 QUALIFICATION OF SELECTED ADHESIVES
4.1 Process for Fabrication of Honeycomb Sandwich Panels
15
A method for fabrication of honeycomb sandwich panels with the
LARC-13 adhesive was developed. The materials to be used for .the
fabrication of honeycomb sandwich panels are HRH 3/16-4 core of
1.27 cms (0.5") thickness and eight plies of Celion 6000/PMR-15 or
Celion 6000/LARC-160.
4.2 Process for Mid-Plane Bonding Large Area Panels
Composite laminates of Celion 6000/PMR-15 or Celion 6000/LARC-160
were used for fabrication of mid-plane bonded panels.
The 30.5 cms (12") x 30.5 ems (12") mid-plane bonded panel fab
ricated with LARC-13 and FM34B-18 adhesives were c-scanned as shown in
Figure 13 and IS and subsequently cut into slotted shear and short
beam shear specimens to evaluate bond properties. Test results are
shown in Tables 14 and 15.
4.3 Effect of Bond Line Thickness on Lap-Shear ,Strength
Several lap shear specimens with varying numbers of LARC-13
adhesive layers 0.034 gm/cm2 (0.07 Ibs/ft2 per layer) were fabricated.
These specimens were subjected to destructive testing after the bond
line thickness had been determined. It was concluded from the data
obtained, as shown in Table 20 that the lap-shear strength decreases
as the bond-line thickness increases. These lower bond strength
values, with increased bond-line thickness, is apparently due to high
porosity in the thick bond lines.
4.4 Effect of Out-Time on Filleting Properties of LARC-13
16
A study to observe the effects of out of refrigerator time vs
filleting was completed. A series of 17.8xll.4xl.27 cms (7"x4i"xi")
thick sandwich panels were bonded with a single batch of LaRC-J3
adhesive that was exposed to normal atmospheric temperature for up to
8 days. 2"x2" flatwise tensile specimens were cut from each panel
and tested at R.T. and 5890 K (600oF) after measuring fillet thickness.
Results of these tests are shown in Table 21. It is concluded through
the data obtained that the adhesive LaRC-13 is not significantly
affected and there is no remarkable difference in filleting as well.
4.5 Evaluation of Catalxst
A study to evaluate catalysts to reduce the required cure tem
perature for LARC-13 adhesive was undertaken. Initial studies, (with
organic peroxides like USP-138 which is a new cyclic peroxyketal
recommended for broad high temperature applications as a polymeri
zation catalyst and cross linking agent and Experox-IO which is
tertiery butyl perbenzoate recommended for the same reasons from
WITCO chemicals), showed promising results when the adhesive was
cured at 5330 K while maintaining the same 5890 K (600oF) post oven
cure cycle. Preliminary specimen tests resulted in lap-shear
strength in excess of 13.6MPa (2000 psi). Catalyst used is in the
amount of 2% by weight.
Further work indicated that if the cure temperature is reduced
to as low as SOSoK (4S0oF), poor bonding results and the laminates
debond while being post cured. Therefore, based on the study and
the data shown in Table 22, it was concluded that by using a catalyst
we can reduce the cure temperature at S330 K (SOOoF) and still main
tain the lap-shear strength above 13.6MPa (2000 psi).
Thus, our initial efforts with catalized product indicated that
a drastic change from the conventional LARC-13 cure temperature of
589°K (600°F) to below 5330 K (500°F) could not be made by the addi
tion of the organic peroxide and still achieve adequate cure. From
this evidence, it was concluded that the material cannot be cured as
are current aerospace grade epoxies.
5.0 SUMMARY
During the course of this program, a series of polyimide adhe
sives were screened for mechanical and physical properties and for
processibility in fabrication of large mid-plane bonded panels and
honeycomb sandwich panels. The mid-plane bonded panels were fabri-
17
18
cated using Celion 6000/PMR-15, composite sheets and honeycomb panels
of HRH 327 3/16-4 core with face-sheets of Celion 6000/PMR-15. From
a series of 41 adhesive formulations, the LaRC-13, FM34B-18; and a
modified LaRC-13 adhesive designated as M-LaRC-13E were selected for
process evaluation studies. On the basis of the data obtained during
these studies as shown in Table 10, LaRC-13 adhesive was rated as the
best of the three adhesives in terms of availability, cost, process
ability, properties and ability to produce void free large area
( 12"x 12") midplane bonds.
An evaluation also was made of surface treatment and primers
for the graphite/polyimide adhesives to be used in bonding studies.
The selected surface treatment consisted of light-abrading with
Behr-Tex pads followed by diglyme wash. One selected primer was
BR34B-18. Processes also were developed for fabrication of honey
comb sandwich panels of very good quality for the LaRC-13, M-LaRC-13E,
and the FM34B-18 adhesives. The good quality was evidenced by rup
ture in the honeycomb core rather than in the honeycomb face sheet
bands on flatwise tensile strength testing.
In future studies on polyimide adhesives, it is recommended
that the following areas be addressed:
o Polyimide adhesives investigated to date tend to be brittle
and fail in peel mode more readily than lower temperature,
higher strength epoxy systems. To overcome this disadvantage,
plasticiers must be studied to determine if a more flexible
PI adhesive can be produced with improved peel strength,
increased handleability, and minimum effect on lap shear
and other adhesive qualities.
o A low cost substitute for 3,3' MDA should be located and
evaluated to study the overall affect on adhesive cost,
mechanical properties, and processing characteristics.
6.0 TECHNOLOGY DEMONSTRATOR SEGMENT
The initial intent of this task was to fabricate a represent
ative demonstration component of the Space Shuttle aft body flap.
It was intended that the component design would incorporate all
developed processes and structural configurations to demonstrate
manufacturing feasibility of graphite/LARC-160 to full-scale
structures.
The requirement "to fabricate a representative demonstration
component' was later changed "to fabricate a representative struc
tural test component". This structural test component has been
given the designation of Technology Demonstrator Segment (TDS). The
completed TDS, ready for installation of instrumentation for ground
testing, is shown in Figure 17.
In changing from a demonstration to a structural test component,
the complexity of the task changed correspondingly: All aspects of
design, tooling, NDI, fabrication, and assembly became more critical.
To implement this change, the scope of Contract NASI-15843 (Develop,
Demonstrate, and Verify Large Area Composite Structural Bond with
Polyimide Adhesive) was amended to fabricate the cover panels, ribs,
and leading edge covers of the TDS. Fabrication of the remaining
TDS components and final assembly was accomplished under Contract
NAS1-15371, Development and Demonstration of Manufacturing Processes
for Fabricating Graphite/LARC-160 Polyimide Structural Elements.
Design details of the TDS test component as shown in Appendix
drawings, all fabricated elements of the TDS i.e., solid laminate
structures, laminate skins. and bonded honeycomb panels were non
destructively inspected to assure that high quality structure within
the state-of-the-art was fabricated. Typical NDI records for the
TDS are included in the report to indicate the high quality of the
laminates and sandwich assemblies produced.
19
7.0 FABRICATION OF STRUCTURAL TEST ASSEMBLY
7.1 Prepreg Tape
Unidirectional prepreg tape used in fabrication of all details
was supplied by U. S. Polymeric, Inc., Santa Ana, California. The
tape consisted of 3000 graphite fiber filaments in the basic tows
impregnated with high temperature polyimide resin. The filaments
were high modulus (typical 234 GPa or 34 x J06 psi) and high
strength (typical 2.758 MFa or 400 x 103 psi) PAN-based carboni
graphite fibers manufactured by Celanese, Inc. (under the trade name
of Celion 3000) and sized with polyimide resin NRI50-B2, had a
typical fiber density of 1.76 g/cm3 (0.064 Ib/in3). The polyimide
resin (originally developed by NASA-Langley Research Center with
designation LARC-160) was prepared via addition-type reaction or
reverse Diel-Alder reaction and had a upper operating temperature
limit of 5890 K (6000 F).
The composite tape was stored inside sealed bag at 2550 K (OoF)
in a freezer until it was ready to be used. The sealed bag was
removed from freezer and thawed to room temperature before opening
. to avoid any deposit of moisture on the tape. They were inspected
visually to examine overall quality, such as drape, tack, amount of
tow splice, fibers collimation, uniformity of resin on the tape and
thickness of tape. The nominal cured ply thickness was 0.0064 cm
<0.0025"). The supplier provided samples of resin and fiber for
reference. Several feet of the tape from each roll underwent quality
control tests to confirm the reported properties from supplier. The
cover consisted of three major components, honeycomb core, upper and
lower skins, and forward and aft closeout edge members as shown on
Appendix A drawings.
7.2 Leading Edge Cover Panel
20
7.2.1 Skin
Prepreg tape was removed from freezer and allowed to thaw to
room temperature before any work would start. It was cut according
to its specific orientation with aid of templates. The lay up was
done with great care and once i.t was completed the skins were weighed
and arranged under vacuum bag as shown in Figure 18. It was imidized
following the cycle below:
I. Apply 5.1 cm (2") Hg vacuum and maintain throughout the cycle.
2. Raise temperature to 4910 K (425°F) at a rate of 2.20 K (4°F)
per minute.
3. Stage the skin at 4910 K (425°F) for 30 minutes.
4. Cool to 3380 K (150°F) before removing the part from oven.
After imidization, the laminate was weighed and examined "isually
to determine any abnormali ty being committed to further work. The
staged laminates were prepared as in Figure 19 to be autoclave cured
with the following cycle (Figure 28):
I. Apply full vacuum and 1.4 MPa (200 psi) autoclave pressure
and maintain throughout cure.
2. Raise temperature to 5190 K (475°F) at a rate of 4°F per
minute.
3. Hold 5140 K (466°F) for 42 minutes.
4. Increase temperature to 6020 K (625°F) at a rate of 4°F
per minute.
5. Hold at 602°K (625°F) for 2 hours and 45 minutes.
6. Lower temperature to 4220 K (300°F at a rate of 20 K (3.70 F)
per minute.
7. Below 4130 K (284°F), cooling rate may be increased to 4.SoK
(8°F) per minute.
21
8. Release autoclave pressure at 338°K (IS00 F) and remove part
from autoclave.
The curing temperature of 6020 K (62SoF) has substantially de
creased the post cure time. The cured laminate was free standing post
cured in the oven at S890 K (600oF) for 2 hours. The heating and cool
ing rate for this process were not closely monitored. The heating
rate was at 2.SoK (SoF) to 4.SoK (8°F) per minute and cooling rate
varied even more. After the assembly was completed, the skin was post
cured for an additional 4 hours (Figure 29). Several parameters were
used to assess the condition of the skins after autoclave cure. These
included total resin loss, fiber volume (calculated through several
methods), skin thickness, skin final weight, glass transition tempera
ture (Tg) and overall cosmetic appearance. A computer program was
developed earlier to relate most of these variables.
Ultrasonic transmission c-scan tests were performed as part of
quality control effort on skins and all other assemblies in the cured
and post-cured conditions. Established,NASA-LARC "A" sensitivity
standard specimens were used as a control in setting equipment sen
sitivity for quality verification. This particular nondestructive
testing has been a very useful and time-saving method to detect any
voids or delaminations. C-scan recording of both skins were made
and no defects were detected. One of the recordings is shown in
Figure 30.
Because of its unsymmetrical orientation, the skins curled up
after autoclave cure (Figure 31). These skins were thin (0.031 cm
Or 0.021 inch) and very fragile; to minimize damage during handling
for ultrasonic c-scanning and clearing for water-broken surface,
aluminum picture frame was designed and fabricated to the correct 'size
to hold skin stretched out and lay flat.
7.2.2 Closeout Channel
22
. Same material for the fabrication of skins was used to fabricate
the closeout channels. These are closeout "U" shape edge members on
the honeycomb sandwich cover assemblies. One of the cover panels
will be mechanically fastened to the final body flap segment. A
16-ply internal doubler was incorporated to one of the flanges.
The difference in thickness on both flanges required a specially
designed tool for layup, staging and curing. The layup was a time
consuming and tedious process. Four plies of prepreg tapes were laid
up, debulked and consolidated under full vacuum at room temperature
for 2 hours, then vacUUm formed onto the tool (on the tape that was
already laid up) at room temperature, heat was applied to facilitate
the transfer. Layup 0 f more than 4-ply debulked tape would make the
vacuum form difficult and increase the formation of fiber twisting
and wrinkle. Transparent vacuum bagging material was used to provide
maximum visibility (Figure 32).
The arrangement for staging and curing these channels are shown
in Figures 21 and 22. respectively. For staging, autoclave cure and
post cure, same cycles as those for the skins were used on closeout
channels. During the bagging for autoclave cure, the channel should
be placed on the correct side on the layup die as the flanges had
different thickness; wrinkles could be introduced very easily and
more care should be exercised to avoid, or at least to minimize,
any wrinkles.
Several modifications in the staging cycle were needed to obtain
void free and targeted-fiber-volume or targeted-thickness channels.
These are primarily higher staging temperature (up to 497°K or 4350 F),
longer staging time (up to 1-1/2 hour) and different numbers of
bleeder/breather or combination of these.
Because of unbalanced channel design (thicker on one leg) it:
assumed the expected moderately bowed condition when removed from
mold lay-up die (LUD). This bow caused some difficulty during next
assembly bonding. In a production mode either the part or the tool
should be revised to produce a straight part at ambient temperature.
Cured channels were submitted for c-scan inspection and con'
firmed to be void free before trimming was done.
23
7.2.3 Assembly of Leading Edge - Bonding Tool
Because of its favorable thermal expansion coefficient, stain
less steel was used for fabrication of the hinged cover bonding tool.
The tool was roll formed to the desired curvature as close as possible.
However, it was not stiff enough and two additional flat stainless
steel sheets were bolted to the inside of the curved sheet to avoid
rocking and stabilize the bonding tool. Two L-shape steel angle
members were fastened to the bonding tool along its length to protect
the sandwich assembly from crushing during bonding operation. The
angle members were located so that no additional positioning on the
width was needed. The completed bonding tool was sprayed with a thin
coat releasing agent FREKOTE-33, the agent was dried in the oven at
3660 K (2000 F) minimum for 30 minutes or more (Figure 33).
7.2.4 Prefit of Details
The relatively small value of radius over of radius over core
thickness ratio of sandwich panel indicated a potential difficulty
in fitting the core to the curvature of "the tool. A plain core seg
ment was forced to follow the contour of the bonding tool with vac-
24
"uum bag pressure employed at room temperature. A vacuum level of
approximately 500 mm Hg (20 inches) was sufficient to have the core
fitting nicely to the tool, but upon removing the vacuum bag pressure,
the core sprang back to its original shape. Another approach was
tried. A segment of this 1.91 em (0.75 inch) thick core was heated
in oven (free standing) at 5890 K (6000 F) for about an hour. Immedi
ately following its removal from the oven, the core was tied down to
the bonding tool by heat-sensitive shrink tape and dead weight for
about 30 minutes. However, once the tapes were removed, the core
again returned to its flat shape. Preforming this cured honeycomb
core to the bonding tool before assembly for sandwich panel bonding
was impossible.
In order to determine the exact contour of cured panels fab
ricated from this tool and at the same time develop the bonding
process, several smaller sandwich panels were fabricated at different
locations on the bonding tool. The arrangement for the autoclave
bonding was similar to that shown in Figure 34. Three of these small
panels were fabricated and the exact number of 0.0076 cm (0.0030 inch)
porous separators and 0.013 cm (0.005 inch) non-porous separators to
be used as tooling shims to make panel fit the stability rib fr()nt
curved segment better was determined.
7.2.5 Cleaning and Priming of Details
Some sort of surface treatment of the bonding areas must be
carried out to achieve optimum strength.
For skin and closeout channel bonding surface, heavy duty
"Behrtex" type scrubbing pad and "Comet" cleaning agent were used
to remOVE~ all grease and other contaminants. Parts were rinsed with
deionized water to check the waterbreak condition. Then they were
dried in oven at 394°K (250°F) for about 30 minutes and the details
were ready to be primed and bonded.
The honeycomb core used for the fabrication of hinged covers
were supplied by Hexcel Corporation. It was HRH-327 high tempera
ture glass-fabric reinforced polyimide honeycomb core with 0.47 cm
(3/16 inch) cell size, 1.91 cm (0.75 inch) thickness and 0.048 g/cc
(3 Ib/ft3)density. This core did not need to be degreased. Methyl
ethyl ketone was sprayed into core cells to remove dUst and any
contaminants.
The liquid paste adhesive counterpart of FM34B-18 film adhesive,
BR34B-18, was used for priming all bonding surfaces. This primer has
an aluminum-powder filled adhesive with 90% solid content. (Alu
minum powder + organic polymer). It was diluted into 35% solid con
tent with ketone/alcohol thinner, this diluted primer was sprayed
on skins and closeout channels. A thin coat of 25.4 cm or 0.001 inch
thick prime~ solution on the surface was sufficient. Roller was used
to apply a thicker coat of primer solution onto the cores. The
primer was air dried at room temperature for 30 minutes and oven
dried at 366°K (2000 F) for 2 hours.
25
7.2.6 Assembly of petails On Tool
26
Up to 5 plies of 0.0076 cm (0.0030 inch) 3TLL porous teflon
coated glass fabric separator were used on the bonding tool at
different locations. The separators were stacked with different
width and length to avoid any sudden change on tool surfaces; up to
3 plies of thicker 0.013 cm (0.005 inch) 5TB non-porous teflon
coated glass fabric separators were used for the same purpose of
shimming. One ply of 5TB separator was used to cover the entire
bonding tool for the panel assembly.
The adhesive used for the bonding was supplied by American
Cyanamid Company_ It was scrim-cloth-carrier polyimide adhesive
FM34B-18, aerial density 0.04 g/cm2 (0.09 Ib/ft2). In order to
achieve the desired ribbon direction on the core, 3 pieces of core
were spliced together with FM34B-IS. The splicing adhesive was
cured together with the sandwich panel assembly.
As mentioned earlier the cores would not form to the contour
of the bonding tool. Great care and effort were needed to ensure
that the cores were positioned correctly. After the cores were in
their proper positions, several openings existed between cores and
the closeout channels. Beads and stacks of FM34B-18 adhesive were
used to fill up and bridge these areas.
One layer of FM34B-18 adhesive was put on top and bottom of
the cores to be bonded to the skins. The top skin was covered
with one layer of porous separator and stainless steel caul plate
to obtain smooth outer surface. The lay-up die (with straight sides)
of the closeout channels were used to avoid channel crushing and
maintain proper alignment. Pieces of wood wrapped with teflon were
taped down to bonding tool to protect sides of sandwich panel. The
whole assembly was arranged as shown in Figure 34 and cured in
autoclave with the following cycle (Figure 25).
I. Apply full vacuum and 17.6 KPa (25 psi) autoclave pressure
and maintain throughout cure cycle.
2. Raise temperature to 4630 K (3750 F) at a rate of 2.SoK
(SoF) per minute.
3. Cure adhesive at 456°K (3620 F for 2 hours.
4. Cool to 33SoK (1500 F) at a rate of 20 K (3.7°F) per
minute before releasing autoclave pressure.
The assembly was free standing oven post cured with the same cycle
as the skins and closeout channels.
The leading edge cover was visually inspected for its bonding
quality. Adhesive filleting on the honeycomb core cell wall in··
dicated good flow was achieved and good quality bonding was expected.
The cover had uniform thickness and retained its curved configuration
(Figure 35 and 36). It was checked against the stability ribs and
some change on the contour would give a better fit (Figure 36).
During the fabrication on the second leading edge cover, two
additional plies of STB separators were put on the straight portion
of the bonding tool on top of the shimming that was used for the
first cover panel. The fabrication procedure for the second cover
. was identical to the first one. With this extra shimming, the second
hinged cover fitted the stability ribs better and was used as the
bonded, lower hinged cover of TDS. The first cover was used as the
bolted, upper hinged cover.
7.3 STABILITY RIB , W'
The stability ribs have honeycomb sandwich panel webs with
secondard bonded pi shaped closeout out channels top and bottom
completing the "I" beam configuration. A "u" channel closed out
the front edge with a Pi shaped channel at rear.
7 .3.1 Stability Rib Sa:n,d:vich Panel
The skins for these ribs were rather thin, the nominal thickness
was 0.020 cm (O.OOS") and had unsymmetrical ply orientation of
(90, ~ 45). All skins were fabricated as in section 7.2.1.
27
7.3.2
7.3.3
28
The honeycomb core was the same material as used in leading edge
panels except they were 1.27 cm (0.50") thick and no splicing was
needed to obtain the desired ribbon direction. The panel was bonded
as in section 7.2.6.
The front U shape closeout channels were fabricated from woven
fabric broadgoods, Thornell 300/LARC 160. The nominal thickness of
cured laminate was 0.018 cm (0.007"). The channel was laid up on
the tooling shown in Figure 26 and held in place by heat-sensitive
plastic tapes, the channels were staged and cured with cycles used
in section 7.2.1.
Undiluted 80% solid BR34B-18 primer was used in place of
adhesive FM34B-18 for bonding of closeout channels to stability ribs,
a thick coat of primer was applied on the bonding surface using a
paint brush.
Pi Cap Element
The configuration of the Pi joints is shown in Figure 27,
special tooling had been designed and machinEd for the fabrication
of these joints. These joints were autoclave cured with cycle
similar to those described in section 7.2.1 except that lower cure
temperature of 5600 K (5500 F) was used to prevent degradation to the
silicone rubber tool details. To obtain the desired 6250 F Tg the
Pi joints were subsequently post cured in an oven for four (4) hours.
Bonding of Closeout Channel and pi Cap to Web Sandwich Panel
All faying surface of the web sandwich panel, closeout channel
and Pi cap were cleaned and oven dried with the same procedure as
above. Similar priming process used in Section 7.2.5 for leading
edge cover was carried out. The flanges of Pi cap were clamped
down to an aluminum tool that would increase the width of center
opening to ease the fitting of Pi cap on web panel.
For the forward closeout channel to web bonding area, FM34B-18
film adhesive was not used because it was so tacky that installing
the channel to web would produce non-uniform bond line. Undiluted
80% primer solution of BR34B-18 was applied by paint brush to all
areas (Figures 37 and 38). The channel was installed relatively
easy (Figures 39 and 40). After FM34B-18 adhesive was wrapped on
the upper and lower bonding areas on the web, additional undiluted
primer solution was brushed on. This lubricated the entire arE~as.
Graphite pin passing through both pi cap and web after the assembly
was posi.tioned and indexed in the drilling fixture which was :~hown
in Figures 41 and 42. This fixture was modified to allow to be used
as bondi.ng fixture for stability rib assembly fabrication.
NDI technique was performed to verify bonding quality.
7.4 Lo;wer and Upper C.o,ver Peane
l1s
7.4.1 Honeycomb Cor,e
High temperature glass-fabric reinforcedpo1yimide honeycomb
core HRH-327 (supplied by Hexce1 Corp.) was used-cell size 0.47 cm
(3/16"), thickness 1.91 em (0.75"), density 0.048 g/cc (3 1b/ft3).
Several pieces of honeycomb core were spliced together with high
temperature polyimide adhesive (FM34B-18) to obtain the size required
in the ribbon direction. The splice lines added a mere 45 gm to the
total weight of 3.9Kg of full size honeycomb core.
The aft edge of the spliced core was recessed to fit the outer
skin spar recess joggle.
7.4.2 Skin -The skin for cover panel was laid up, staged and cured following
the same procedure used in fabricating skin leading edge cover in
Section 7.2.1. During autoclave cure, a stainless steel metal strip
was placed under the aft edge of laminate to mold into the skin to
form a joggle of 0.051 cm (0.020") thick; this would allow for the
flush fit of the rear spar to cover panels.
29
7.4.3 Closeout Edge Member
A typical edge member layup die (LUD) is shown in Figure 20.
These molds were made of aluminum. For production use steel would
be more appropriate considering the 6020 K (6250 F) cure temperature.
Eight ply 0.051 cm (0.020" thick) forward and aft channels
were laid up and bagged as shown in Figure 21, (for oven staging).
The staging cycle was similar to the one used for cover skins. The
staged channels were prepared as in Figure 22 and autoclave cured
using c9ver skins cure cycle. They underwent similar post cure and
NDI procedures.
In order to get a completely wrinkle-free outer surface on the
channels, a different bagging system was used as shown in Figure 23.
The high-temperature-resistant silicone rubber bag improved the
cosmetics quality of channels due to its wiping action as vacuum
was applied. One pair of forward and aft channels were fabricated
using this preferred method.
7.4.4 J:ion,eycomb S~ndwich Asse1ll;bly
30
To achieve optimum bonding strength the faying surfaces of
skins, core, and fore and aft edge members were cleaned and primed
following the procedure used in section 7.2.5. All the detail parts
were assembled without adhesive and a vinyl impression check was
carried out in the autoclave under 25 psi autoclave pressure and
full vacuum at 4220 K (3000 F) for 30 minutes. The impression check
verified fit up of details.
The detail parts were assembled with the same po1yimide adhesive
used for leading edge cover fabrication and autoclave cured under the
same cycle used in Section 7.2.6. The cover was a large flat honey
comb sandwich panel and no bonding fixture other than a modified
picture frame was used as arm around the whole panel. After cure
NDI ultrasonic c-scanning was also used to check bonding quality~
8.0 DEMONSTRATOR SEGMENT TEST PROGRAM .. -
A preliminary test program has been defined for the graphite/
polyimide body flap Technology Demonstrator Segment (TDS) shown in
Figure 43. The test objective is to verify the structural adequacy
of a typical composite body flap section to sustain acoustic, mechan
ical, and thermal load cycles representative of 100 Space Shuttle
missions. The proposed TDS test program is outlined below and in
volves mechanical testing at room temperature and S330 K (SOOoF) and
thermal cycling and acoustic testing under overall sound pressure
level (OASPL), resources permitting, The S330 K (SOOoF) maximum
temperature is based on Orbiter entry and Terminal Area Energy
Management (TAEM) environment. The fabrication of a large, all
bonded graphite/polyimide structure is advanced state-of-the-art,
and as a result the integrity of the major bonded joints is open to
question. Loading conditions have been determined to subject these
joints to stress levels equivalent to the compos:i.te body flap stress
levels. The TDS will also be subjected to internal pressure to test
the adhesive bonds in flatwise tension.
The TDS is stiffness. rather than strength, due to the structure/
TPS deflection requirements. Therefore, failure under mechanical
loads is not anticipated. Due to the location of the body flap
relative to the Orbiter main engines, acoustic fatigue is the pre
dicted failure mode;
The primary purpose for testing the TDS is to develop confidence
in the structural capabilities of graphite/polyimide. The confidence
is high that a static test will verify the reliability of the NASTRAN
model and other TDS static analysis techniques. The confidence level
associated with the thermal models is relatively high. However, the
confidenc·e in the acoustic model is low, and acoustic fatigue is the
predicted failure mode. These relative confidence levels are illus
trated in Figure 44. If the TDS can survive all of the tests, the
confidence level in GR/Pi structure will be high, and the all bonded
GR/pi structure technology will be verified.
31
8. 1 Test Approach
The body flap is loaded and supported through four-hinge ribs.
However. lack of a hinge rib on the TDS precludes actual body flap
loads. Test conditions to adequately simulate the stress state in
the region of the body flap stability ribs were input to theTDS
NASTRAN model for several loading and support conditions. The tech
nology demonstration segment test approach is outlined below:
MECHANICAL LOAD TEST
o Simulated Load at Room Temperature (RT) and 5330 K
(5000 F).
o 400 Cycle Simulated Fatigue 5330 K (5000 F)
o Internal Pressure at RT
THERMAL CYCLING TEST
o 100 Cycles (222K and 5890 F) for 100 missions
ACOUSTIC FATIGUE TEST
o 165 dB OASPL for 34 Minutes (lift Off)
o 161.5 dB OASPL for 38 Minutes (Aerodynamic)
0157 dB OASPL for 60 Minutes (Aero/Shock)
8.2 Test Constraints
Test constraints for the TDS are summarized in Figure 45.
Acoustic test constraints require attachment to all three stability
ribs. Thermal test constraints make application of distributed loads
at elevated temperature complex. Mechanical test constraints require
the TDS to be contilevered from the outer ribs, while loading is to
be applied at the trailing edge. Funding constraints preclude spec
trum fatigue and thermal cycling tests.
8.3 Technology Demonstrator Segment NASTRAN Model
32
The TDS NASTRAN model was used to define the loading .condition
for the mechanical load test and is presented in Figure 46. The
374 node model was extracted from the 1000+ node body flap model.
The cover panels are represented by QUAD 1 sandwich elements which
model the honeycomb core and laminate face-sheets. The spars and
ribs are modeled by SHEAR panels stabilized by ROD elements. Spar
and rib caps are represented by offset BAR elements.
8.4 Room Temperature an~5330K M~chanical Load Tests
Lack of a hinge rib (through which the body flap is loaded
and supported) on the Technology Demonstration Segment makes appli
cation of actual body flap loads impossible. The objective of the
mechanical load test is to simulate the body flap stress state and
verify the structural adequacy of the TDS to sustain Orbiter static
stress levels. A combination of supports and loads that would re
produce the stress state in thd vicinity of the body flap stability
ribs were determined from several test conditions input to the TDS
NASTRAN model. Test program loading support concept is described
in the following paragraph.
As shown in Figure 47. the TDS is cantilever supported at the
outer ribs, has the middle rib displaced to stress the front spar,
and has concentrated loads at the trailing edge. These concen
trated loads do not require the attachment of a special loading
fixture. A TDS/body flap cover panel stress contour comparison
is presented in Figure 48 and the stress levels in the caps, webs,
and cover panels are compared in Table 23. The stress state of
the load concepts adequately compare with the body flap stress
state. The stability rib web shear, and stability rib cap com
pressive stress are somewhat larger than the body flap stress
levels, but a large factor of safety is still apparent.
The loads applied at the rear spar cause a twisting downward
moment, and flex the rear spar. These loads simulate the ultimate
stress levels, and will also be applied at 5330K as shown in
Figure 49.
33
8.5 400 Cycle Simulated Fatigue Test
The objective of the fatigue test is to verify the TDS struc
tural integrity to sustain mechanical loads simulating 100 missions.
Because of the high cost, spectrum loads will not be applied. In
stead four lifetimes (400 cycles) of limit loads will be applied
(R = + 0.10) at 5330 K (Figure 50). Spectrum loads would have only
a few cycles approaching limit loads, making the test condition more
severe than an actual service environment. Due to test fixture
limitations, however, load reversals will not be applied.
8.6 Thermal Cycling Test
34
The body flap is subjected to a thermal environment with a
temperature range from 2220 K (_600 F) to 5890 K (6000 F). The maximum ~
applied mechanical loads occur at body flap temperatures between
1660 K and 5330K, with maximum temperature 5890 K (6000 F) resulting
from heat soak back after the orbiter is on the ground. The objectives
of the thermal cycling test are: (1) verification of the TDS struc
ture under 125 thermal load cycles, and (2) verification of the
analytically predicted thermal strains.
The selected thermal cycle is shown in Figure 51. The cycle
represents the temperatures and thermal gradients obtained during
an abort-once-around (AOA) thermal trajectory for the upper 5890 K
(6000 F) temperatures, and the cold temperatures 1660 K (-160.80 F)
are from a severe cold on orbit mission. The structure would not
be subjected to these extremes on a single mission.
The baseline aluminum body flap was not tested under thermal
cycling. However, some of the body flap joints and other components
were subjected to thermal cycling. The components were exposed to
a temperature range of 1990 K (101.40 F) to 4490 K (348.SoF) for ISO
cycles. The baseline graphite/epoxy OMS Pods were subjected to 400
mechanical load cycles and 100 thermal cycles (1660 K to 4490 K) to
simulate 100 space missions. It therefore appears that a similar
test program is required for the composite body flap structure.
8.7 Acoustic Fatigue Te~t
Due to the location of the body flap relative to the orbiter
main engines, acoustic fatigue is potential failure mode. The TDS
should be exposed to acoustically induced random vibration simulating
overall sound pressure levels as high as 165 dB anticipated du:ring
Space Shuttle Vehicle (SSV) ascent. The objectives of the acoustic
test are as follows: verify the primary structure acoustic fatigue
life (100 Space Shuttle missions). demonstrate the integrity of
selected regions of the direct· .. bond TPS. and verify the analytically
predicted cover panel RMS strain and frequency.
ThE! body flap acoustic environment is shown in Figure 52.
Upon successful completion of the mechanical and thermal cycling
tests, the TDS will be modified for acoustic test_ funding permitting.
Modifications will include adding direct-bond fibrous reflects the
composite insulation (FReI) tiles, and closing out the TDS open
bays. Testing is to be conducted by NASA.
8.8 Internal. Pressure Test
The TDS may be tested under internal pressure (approximately
.02 MPa (3 psi» to assess the integrity of the major adhesive bonds
by placing them in flatwise tension. The TDS torque box would be
filled with styrofoam to reduce danger due to the pressure. The
test article would be sealed by replacing the front spar access
panels with metal panels equipped with pressure seals and fittings.
The front spar attachment "T" members were designed for shear
loads. Concern has been expressed over the integrity of these
members under internal pressure loads (see Figure 53). Analytical
models predict failure at .03 MPa (5 psi), but confidence in the
analysis is low. It is recommended that this test be conducted
last, predicated on successful completion of a subelement test.
35
36
8.9 Summary
The test objective is to verify the structural adequacy of
a typical composite body flap section to sustain acoustic, mechan
ical, and thermal load cycles representative of 100 Space Shuttle
missions. The lack of a hinge rib on the TDS precludes actual
body flap loads, therefore, loading conditions to simulate the
body flap stress state were determined by using a NASTRAN model of
the TDS.
Since it would be very difficult to perform mechanical and
thermal cycling simultaneously, an alternate method is suggested:
First, perform the mechanical testing; Second, perform the thermal
cycling and acoustic testing; and finally, the mechanical test
could be performed again to assess the structural degradation
from the previous testing. It is assumed that the acoustic test
ing would take place in a NASA facility -- JSC.
Because the TDS is stiffness rather than strength critical,
failure under mechanical loads is not anticipated. However, due to
the extreme sonic environment (165 dB OASPL), acoustic fatigue is
a significant failure mode. The body flap must also be able to
withstand the Orbiter thermal environment 166°K to 589°K (-160 to
6000 F for 100 space missions.
The TDS design was updated to provide for cantilever support
during testing. The stability rib caps were beefed up in the lead
ing edge area, and potting and inserts were added to the bolt holes.
When testing is complete, analytical predictions for stress
levels, deflections, and acoustic fatigue will be correlated with
test data. The analytical models for the GR/Pi Orbiter Body Flap
can then be updated.
DOCUMENTATION
1. AS-EMR-SO-053, "Technology Demonstration Segment NASTRAN Model", dated
April 14, 1980.
2. AS-EMR-80-052, "GR/pI Technology Demonstration Segment Design Update",
dated April 14, 1980.
3. AS-EMR-SO-,On, "Technology Demonstration Segment Test Definition",
dated May 2, 1980.
4. SSV80-22, "GR/pI Technology Demonstration Segment Test Concepts:,
dated April 14, 19S0.
5. AS-EMR-80··164, "Composite Body Flap Segment Test Program", dated
September 2, 1980.
37
This Page Intentionally Left Blank
0 0 0 H 0
" " " I " O'C:O-C~C'O ... c c, H2N -0- ~ -0-NH2 IO:g:O
" " H " 0 0 0
BTOA ~~ NA
BENZOPHENONE 3,3' ·METHYLENE 5·NORBORNENE TETRACARBOXYLIC OIANILINE 2·3-OICARBOXYLIC ACID ANHYDRIDE
ACID ANHYDRIDE
0 0 0 H 0
" , , ,
" HO-C:o'C"():C -OH H2N -o-~-o-NH2 IO:C- OH H3COC C - OCH3 C-OCH3 , " H " 0 0 0
.!!TDE 3,3' ·MDA 1!!.
DIMETHYL ESTER 3,3' ·METHYLENE METHYL ESTER OF OF 3,3', 4,4'· DIANILINE 5·NORBORNENE BENZOPHENONE 2·3 ·DICARBOXYLIC TETRACARBOXYLIC ACID ACID
Figure 1. LARC 13 Monomer Reactants
A .. PRESSURE
MONOMER REACTANTS
AMIDE ACID
POLYIMIDE
'/
0 f 0 0 OJ 0(3 II II II " II
~ c..... c c c.... ,e THERMALLY ~- N"O"CH2"QTN~:or IOC~, N1O""CH2-o- N,~ CROSS LINKED
o 0 0 n 0
Figure 2. LARC 13 Po1yimide Reaction Sequence
39
Figure 3. Adhesive Prepreg Mill
o CF3 0 II I n
HO-C C C-OH
Ho-clOrc~PC-OH II II o 0
+
HEAT 4,4' ·PHENYLENE DIAMINE SOLVENT
9- NH2
NH2 3,3' ·PHENYLENE DIAMINE
4,4' ·(HEXAFLUOROISOPROPYLIDENE) • BIS(PHTALIC ACID)
Figure 4. Reaction Scheme of NR150B2G
41
Figure 5. Flatwise Tensile Test Specimen
42
Figure 6. Flatwise Tensile Test Specimen After Testing
44
LOADING BAR
BOLT HOLE
A ~~ ,<»1 • 1 ~
E._ STRAP
CLAMP 8
SECTION A-A
(1.995-2.005) 5.067-5.092
I---+-- (1 • 50!,0 . 01) 3.8!,0.03
CLAMP A
r i PLUS (0.500+0.005) 1.27+0.01
COUNTER-- . -WEIGHT
DIMENSIONS IN INCHES. NOT COMPLETELY DIMENSIONED r i • RADIUS OF DRUM TO MID-DEPTH OF SPECIMEN FACING r • RADIUS OF MID-DEPTH OF STRAP o
Figure 7. Climbing Drum Peel Test Apparatus
Figure 8. C-Scan of 76 x 46 cm (30" x 18") Panel of Celion 6000/PMR-15 Composite Panel, 8 Ply, Unidirectional
Figure 9. C-Scan of 76 x 46 cm (30" x 18") Panel of Celion 6000/PMR-15 Composite Panel, 8 Ply, Unidirectional
NASA STANDARD 'WI SENSITIVITY
Figure 10. C-Scan of Honeycomb Sandwich Panel With FM34B-18 Adhesive
47
NASA STANDARD "A" SENSITIVITY
Figure 11. C-Scan of Honeycomb Sandwich Panel With LARC 13 Adhesive
48
::c I-~ z w a:: l-V)
a:: « w ::c V)
Q. « ..J
PS I MPa 1 8
2500 r-
15
2000 r-
1500 I-10
1000 I-
5
500 l-
CJ - RT
r-- ~ 589°K (600°F) ~
r--I- ,...-
I"""'"" ., ,.....
~
~
~ ~ ~ ~ P')c
~ P')c ~ ~ I>c)< ~ ~
~ >< ~ ~ X< x >Y »« ~ ><x" >< x;< ~ ~ ><x" - X<
~ >« "- >0- X)< >xX 0x »c >0- X)< X<x >0- >« »c >xX X'X >< ~ X)< X<x >0. >0 >< >xX ~ -h ~ >< ><>< »« ~ X" »< xx >xX Y>< ~ -h X" ~ )eX >xX >0- >0c X" X'>< Yx )c>< ><x Y>< >0- X< X'>< Yx >-y >x< ><x Y>< >0c ~ >9 >0c Yx X)< »« ><x Y>< '10< X< yy. ~ ><;t- X)< »c ~ X< Yx >Y ><y. »c X<x '10< X" >Y
Y>< X)< XX< »c ~ >0c X< ><x" -- X<x ~ ~ >0c X)< ><;t- »c '10< X" >Y D0c ><;t-
~ )0 ><y X" >Y D0c >y ;0 ><y .;x >« ><x" Y>< ;0 Yx D0c ~ Y>< >y ~ X< X" X< :>c:>< X< X< ~ X)' :>0< )c:>< :? y. ~ ~ X" )
~ y X)' :>:>< ~ ~ Y)c ) ) >y ~ :>>c ~ >< Y) :xy. ~ »>< ~ )< ~ >c Y> ~ ~ 'I >c t>? >- t>? ~ Y> ~ ~ ~
)c t>? Y) DY< >y 'I Y. t>? ~ ~ D<)< ~> 'I )c )c
BR34B-18 LARC-13 LARC-160 PMR-15E NR-150B2 052X 050X AI1130L
Figure 12. Comparison of Primers For Bonding Ce1ion 6000/PMR-15 Panels With LARe 13 Adhesive
r--
~ >c >0 >0< Yx ~ ~ ~ ~ »c -h X" X"
X< Y>< ~ ><>c >c>< ><>c
»c>< »c>< .~ ~ »c>< ~ ~ >xc >xc
AI830
HBS-S6. NASA STANDARD "A" SENSITIVITY
Figure 13. C-Sean of 30.5 by 30.5 em (12 x 12 in.) Mid-Plane Bonded Panel With LARC 13 Adhesive on Style 104 Carrier
50
MBS-4B. NASA STANDARD "A" SENSITIVITY
Figure 14. C-Sean of 30.5 em by 30.5 em (12 x 12 in.), Mid-Plane Bonded Panel With M-LARC-13E Adhesive on Style 104 Carrier
51
MBS-16 NASA STANDARD "A" SENSITIVITY
Figure 15. C-Sean of 30.5 em by 30.5 em (12 x 12 in.) Mid-Plane Bonded Panel With FM 34B-18 Adhesive on Style 104 Carrier
52
Figure 16. Cure Cycle for the Fabrication of Honeycomb Sandwich Panels with FM34B-18 LARC-13 and M-LARC-13E
A810828 F-3C
Figure 17. Technology Demonstrator Segment Completed Structure
2" WIDE KAPTON TAPE
GS43
KAPTON FILM
~~~~~~~~~~~~~~~~~~~~~120 G~SS CLOTH
------ , .. PERFORATED KAPTON
--f I .. ____ ...... LA ..... V-_U .... P_T ... O __ B .. E _I.M .... I D ... I ... Z_E .. D ____ ..... I > 3TLL POROUS SEPARATOR
- - - - - -- - - --s ___ 5TB NON-POROUS
~~~==========================================1~~ SEPARATOR STAGING TOOL
Figure 18. Oven Imidization For Flat Lamir~te
PICTURE FRAME
BOLT
A-Boo SEALANT
./ KAPTON FILM
,-----------------------------------~ ______________________________________ f~120 GLASS CLOTH
I=--=-~-:~~:--=--=-Z)3TLL POROUS SEPARATOR
-----------------------YS/ KAPTON FILM
CURl NG TOOL
Figure 19. Autoclave Cure For Flat Laminate
A810430 A-8e
Figure 20. "u" Channel Lay-Up Mold
57
KAPTON FILM
STAGING TOOL
NOTE: TOOL SECTION HAS BEEN COATED WITH FREKOTE 33 OR SIMILAR TYPE OF RELEASING AGENT.
58
Figure 21. Staging Lay-Up For "u" Shape Laminate (In Oven)
A-SOO SEALANT
GS:-43 SEALANT
Figure 22.
KAPTON ENVELOPE BAG
120 GLASS CLOTH
- 3TLL POROUS SEPARATOR
(KAPTON ENVELOPE BAG)
Curing Lay-Up For "u" Shape Laminate (End View)
59
SILICONE RUBBER BAG (BEFORE VACUUM DRAWN)
A-Boo SEALANT
CUR ING TOOL
SILICONE RUBBER BAG (AFTER VACUUM DRAWN)
120 GLASS CLOTH
~ 3TLL POROUS SEPARATOR
MECHANICAL FASTENER (CLAMP OR RING)
Figure 23. Alternate Curing Lay-Up For "U" Shape Laminate
60
WRIGHTON 8400
GS-43 SEALANT
________________ 1162 GLASS CLOTH
~THIN CAUL PLATE
Pl-----------~V ............. 3TLL POROUS ________ ~ SEPARATOR
~====-~I=!~.:I =H=O=NE=Y=CO=M=B===PA=N=EL==~I ~ :..... ~ (PICTURE FRAME WITH
l DOUBLE BACK TAPE
~-------------------------------------------TOOL
5TB NON-POROUS SEPARATOR
Figure 24. Curing Lay-Up For Honeycomb Sandwich Panel
LLI a: ~
589
533
~ 478 LLI 0.. ~ LLI I-
LLI 422 a: ~ u
366
311 (100)
CURE TIME (HOURS)
Figure 25. Cure Cycle For Cover Panels
11.359 r-- (4.427 11
)----..
-~-
Figure 26. Tool Configuration For Stability Rib Closeout Channel
31.8 (12.5")
63
~ 10.16R (19.911 + 0.025R)
+(4.00R) ,(7 .839 ~.010R)
2.54 11.00)
(.25)TI~", ____ ------ (49.18)------------i-It-.. -
6.35 0.64 --, 124.92 r A
lip III ANGLE
0.175 CM (0.069)
"PI" CHANNEL 90 0 PLY DIRECTION
J--l.349-+.010 I ( .531) -.000
+ 0.292CM (1.25) (0.115) ~.18 0.475R (0.187R)
~~~~~.-l t.:-~ (3. 00 )I __ "':---_ .. ~I' lip I" CAP
90 0 PLY DIRECTION 7.62 SECTION A-A 00 FILLET, 2 PlCS
(2.50)
I LA 10.16R (4.00R) 2 PLCS.
.f- I .. (12.06) 30.63
Figure 27. Curved "PI" Joint Element Design 8879-00250-007 (and -011)
!ill
UJ 533 a:::: :::> .-;i! UJ 0.. 478 x: UJ .-UJ a:::: B 422
366
311
(600) P
(500)
(400)
(300)
(200)
(100)
I:': ,,;n ' !'~: m: !::~ [::i ::i Ii:: iii: ,~;:i mi:::::: m: ';: 1::.m'T,:t;:',- " ,,: ,: J/I:I':::;:" WL :T I::; In" , , I;:::",',::::' ;:,,' , ':"I"H" ":';i1i:+:it;;i';;,;H>:'l::,;;,,',,';;':::; :';: i\;d,:·:i:L: i:; ::', . :' I":: , ,: ::': i:; l::T'J;;' :;!H:;;::
hir ::: I!!H ll!i!jfljjil'i[1: i;. , ~;i ':1 ,:::: U gil:: lii~ ,Hi ' iii' .ji;; I,::: i:ll h ' Ii; llEi 1:1 UiP miilWllH::: ~:, ii: g:, g: :1:./!:;:I"::: : I::: ,':' i> I , !n:I,,::!':: .:;:T:: ,H I:",
.~ g: ::: :,", ::;ii,;:: ,ii' iii!;';l:: :::;-:: .. , I:::: ::i; " , , 'iTi!:UiI!!li:i; iii, '!: j~E'ijiJ!!r:: Ill' ,;',,1: :\i!ididEIT ;:' I::: • !:li m .' :l: H, rit::'!:::l:H, 'Il! :i:
TEMPERATURE 'I'::
. !:::'
. I:;;:
:iii ;': , [::;: IE
,PF",:I::,I',: .. +,:
bi m , Ii!!: !iHi ,: ,liE H ' lilHHT;;:;:, • ,;\': ' :' : IIItHII iHlIlIllla: 1'1' HHIIIU; 11111 ,,:: '"''
1 2 3 4 5 6 7 8 9
CURE TIME - HOURS
Figure 28. Cure Cycle For Skins
'" '"
u.! rr. ::::> I-
~ u.! 0... ::E: u.! I-
OK
589
533
478
422
366
311
• of
~ (GOO)
(SOOf
(400~ • (300)- •
R' , ' (200)
(100
. '-t .
II •
2 3 4 5 6 7 8 9 10 11 12
POST CURE TIME-HOURS
Figure 29. Oven Post Cure Cycle For Skins
Figure 30. C-Sean of 76.2 em by 157.5 em (30" x 62") Skin For Leading Edge
g; A801023 G-7
Figure 31. Skin For the Leading Edge
A810430 A-7C
Figure 32. Vacuum Forming of the Close-Out Channels
69
-...J o
A810918 G-IC
Figure 33. Leading Edge Panel Bonding Jig
STAINLESS
120 GLASS CLOTH
WRIGHTON 8400 VAC PAK
STEEL CAUL PLATE
3 TLL POROUS SEPARATOR
LEADING EDGE PANEL
CLOSE-OUT CHANNEL
SEALANT
Figure 34. Leading Edge Bonding Assembly-Cross Sectional View
A810803 A-22C
Figure 35. Leading Edge Cover Panel on Bonding Tool
72
A810803 A-19C
Figure 36. Leading Edge Cover Panel
73
A810420 C-2
Figure 37. Primer Being Applied on Rib Web For Bonding "U" Channel
74
A810420 C-3
Figure 38. Primer Being Applied to the Inner Sec tion of "U" Channel
7S
A810420 C-4
Figure 39. "U" Channel Being Inserted to Web Sandwich Panel
76
A810420 C-S
Figure 40. "U" Channel Complete For Bonding
77
A810126 F-IO
Figure 41. Drilling Fixture For Close-Out Channel & Pi Cap to Web Sandwich Panel
A810126 F-8
Figure 42. Close-Up View of Forward Section of the Fixture
00 0
COVER PANELS TDS CONFIGURATION
(HONEYCOMB SANDWICH) • THREE STABILITY RIBS • NO HINGE RIB
LEADING·EDGE ACCESS PANEL • NO TRAILING EDGE
TDS OBJECTIVE • CONFIRM GRIPI COMPOSITE
TECHNOLOGY
~
HINGE RIB ">-• '~H OUTBOARD SEGMENT
STABILITY I!IBS LHINBDARD
HINGE RIB SEGMENT FRONT SPAR WEB RH INBOARD
/ ~EGMENT < PLANE OF TORQUE BOX LEADING·EDGE SYMMETRY ~~ '-.... FAIRING RH OUTBOARD ./ ~ •• ~ SEGMENT
LEADING EDGE ~
BODY FLAP ....... ~PLANEOF
SYMMETRY
TDS STATUS • PRELIMINARY DESIGN COMPLETED OCT 1980 • FABRICATION DEVELOPMENT
- TOOLING INITIATED FEB 1980 - COMPONENT FAB INITIATED APRIL 1980
• FABRICATION - COMPLETION AUGUST 1981
Figure 43. GRIp! Composite Body Flap Technology Demonstrator Segment (TDS)
CONFIDENCE
THERMAL STRUCTURAL
ANALYSIS
PREDICTED FAILURE MODE
ACOUSTIC FATIGUE
ANALYSIS
Figure 44. GRip! Technology Verification (TDS - All Bonded Structure)
00 N
DEMONSTRATION SEGMENT
BODY FLAP
COVER PANELS (HONEYCOMB SANDWICH)
• MECHANICAL TEST CONSTf~AlfH
-LACK OF UlNGE RIB PRECLUDES ACTUAL BOllY FLAP LOAllS
-CAN SIMULATE LOCAL STRESS fIELDS
• THERMAL TEST CONSTRAINT
- APPLICATION Of DISTRIBUTED LOADS AT EL(VATED TEMPERATURE (~5000f) 528°K COMPLEX
- TlIERMAL & LOAD CYCLING RESOURCE R£QUIHLHlNTS ARE HIGH
• ACOUSTIC TEST CONSTRAINT
-LACK OF IUNG( RIB REQUIRES DEMO SEGMUn TO UE SUPPORTED ON SlAIHllTY RIB
-TEST FACILITY AVAILABILITY LIMITED
Figure 45. Demonstrator Segment Test Concept Constraints
00 w
TUREE STABILITY
374 NODES ---
NASTRAN BODY FLAP MODEL
1000 NODES ..
END RIB
FORWARD SPAR
TRAJlIN(i EDGE
Figure 46. Technology Demonstrator Segment - NASTRAN MODEL
STABILITY RIBS
APPLIED AT ROOM TEMPERATURE
(2520 LB) 1145 KG
255 KG (560 LB)
1018 KG (2240 LB)
ANCHOR (TEST Fl XTURE)
TEST OBJECTIVES
-SIMULATE BODY FLAP STRESS STATE & VERIFY GRIPI STRUCTURAL INTEGRITY UNDER ORBITER ULTIMATE STATIC STRESS LEVELS
-VERIFY ANALYTICALLY PREDICTED STRAIN LEVELS
TEST CONSTRAINTS
- LACK OF A HINGE RIB PRECLUDES ACTUAL BODY FLAP LOADS -- LOCAL STRESS FIELDS IN THE VICINITY OF THE BODY FLAP STABILITY RIBS SIMULATED
APPROACH
- LOADING CONDITIONS DEFINED TO REPRODUCE THE BODY FLAP STRESS LEVELS AS DETERMINED FROM TDS NASTRAN MODEL
-APPLY CONCENTRATED LOADS AT TDS TRAILING EDGE
-DISPLACE CENTER SUPPORT -CANTILEVER FROM OUTER RIBS
LEADING EDGE
Figure 47. R. T. Mechanical Load Test
00 VI
.. . ... 0::: X
*STAB IL lTV RIB **HINGE RIB
----------- ----------------,
TECHNOLOGY DE~10NSTRAT I ON SEGMENT
• DEFINE LOAD CONCEPTS TO SIMULATE STRESS FIELD BETWEEN BODY FLAP STABiLiTY RiBS
Figure 48. Typical Cover Panel Stress Contours
255 KG (560 LB)
1018 KG (2240 LB)
(2520 LB) 1145 KG
I 1510 KG (3324 LB)
TEST OBJ ECTI VES
• SIMULATE BODY FLAP STRESS STATE & VERIFY GR/PI STRUCTURAL INTEGRITY UNDER ORBITER ULTIMATE STATIC STRESS LEVELS AT (500°F) 528° K
• VERIFY ANALYTICALLY PREDICTED THERMAL & MECHANICAL STRAINS
TEST CONSTRAINTS
• LACK OF A HINGE RIB PRECLUDES ACTUAL BODY FLAP LOADS
APPROACH
-STRESS FIELD IN THE VICINITY OF THE BODY FLAP STABILITY RIBS SIMULATED
• LOADING CONOITIONS DEFINED TO REPRODUCE THE BODY FLAP STRESS LEVELS AS DETERMINED FROM TDS NASTRAN MODEL
• APPLY CONCENTRATED LOADS AT THE TRAILING EDGE
• DISPLACE CENTER SUPPORT AT LEADING EDGE
• CANTILEVER SUPPORT FROM OUTER RIBS AT' LEADING EDGE
• APPLY 528°K (500°F) TO THE LOWER COVER & 444°K (340°F) TO THE UPPER COVER. INTERNAL TEMPERATURE WILL BE 485°K (41S0F)
Figure 49. High Temperature Mechanical Load Test
400 CYCLES OF LIMiT LOAD APPLIED AT 528°K (500°F)
727 KG
c t- C ZO 161-'
~t-161-o..~
-'
(1800 LB) 818 KG .
181 KG (400 LB) I (1600 LB) t
100
R = (0.10) 0.25
10
ANCHOR (TEST Ff XTU"E)
NUMBER OF CYCLES 400
... TEST OBJECTIVtS
• VERIFY Tbs STftUCTUAAl INTEGRITY TO SUSTAIN MECHANICAL LOADS SIMULATING 100 MISSIONS
• DEVELOP CONFIDENCE IN ALL-BONDED STRUCTURE SUBJECTED TO A HIGH-TEMP SIMULATED FATIGUE ENVIRONMENT
APPROACH
o APPLY 4 LIFETIMES (400 CYCLES) OF R = (+0.10) 0.25
o MAINTAIN UPPER COVER AT 528°K (50QoF), LOWER COVER AT 444°K (340°F) FOR INDUCED THERMAL STRESSES
RATIONALE
o 400 LIMIT CYCLES AT 528 8 K (500°F) IS HIGHER LOADING CONDITION THAN SPECTRUM LOADS WOULD APPLY
~ Figure 50. Simulated Fatigue Test
00
00 SELECTED THERMAL CYCLE OK of 589 600
533 500
478 400
422 300
-589°i< (600°F) -.
~ I~ - --~£8~
--- -- r'\ - - r-- - . - -533°K (500 0 F)-t .~ --\ -,-- ---------
1--
~ - . ·-1---1----
~ I- - - -11 -- ·-I/l - - ,-11'1 i- - ,
1- - - I--
ILl a::
366 ::J 200 I-
~~ ,I-\-LOWER COYE~ FACE" SH-EE! .~ _\
I-- - r-- \- - 1-:\
~ ILl a... x:
311 ILl
100 I-
'" 1\1\. \ -
255 0 \ I\. _\ -'-\ ,
"\ 200 -100 \ \ , -\
..... " I. ........
"0 0.2
I 1-,-I-- J I 1-/- - -1
I II I-
--7- 1-
1\ "\
-- \-\ , -
'" -
~ - Vi I\. !.dr"'" -- -
UPPER COVER PANEL (0)
I~ -:---:f-I- -'" " -UPPER COYER FACE SHEET -L
" - £. - --,- -..I
~ - 1-- ..... - - ---..I ./ ,/
k!: ".. .... t-
0.4 0.6 0.8 TIME (~OURS)
Figure 51. Thermal Cycle Test
--
1.0
t- LOWER COYER PANEL (0)
00 \0
ACOUSTICALLY DRIVEN AREAS (LIFT-OFF CRITICAL)
153 .dB
157 dB
• PANEL FLUTTER ISSUE
~ SONIC FATIGUE ISSUE
[g BASIC STRENGTH AREA
• 165 dB OASPL FOR 34 MINUTES (LIFT-OFF) • 161.5 dB OASPL FOR 38 MINUTES (AERODYNAMIC) • 157 dB OASPl FOR 60 MINUTES (AEROSHOCK)
160 dJJ 163 dB
164 dB
165 dB
BODY FLAP SIZED TO ACCOUNT FOR SONIC FATIGUE
• BODY FLAP IS SUBJECTED TO 165 dB SOUND PRESSURE LEVELS
• ACOUSTIC FATIGUE IS THE PREDICTED FAILURE MODE
• TEST IS PROPOSED TO BE CONDUCTED AT NASA/JSC
Figure 52. Acoustic Fatigue Test
\0 o
FILL TORQUE BOX WITH STYRQFOAM TO REDUCE DANGER DUE TO PRESSURE
REPLACE FRONT SPAR ACCESS PANELS WITH METAL PANELS EQUIPPED WITH PRESSURE SEALS & FITTINGS
COYER PANEL (BONDED)
CRITICAL ~ AREA
P
•
, ..
LEADINl1 EDGE ACCESS PANEL (BOLTED) \
. .,
rt I , I I I
" "TEE"
I t __ !"\
~- FRONT SPAR '. rd- .---WEB (BOLTED) ~I-, . II-
• I NTERNAL PRESSURE MAY DAMAGE FRON SPAR "T"
• PRESSURIZE TO APPROXIMATELY 0.02 MPa (3.0 PSI) • MAX LOAD ON "T" AT 0.02 MPa (3 PS I) = (20 LB/IN.) 3579 GMS/CM
• PREDICTED FAILURE AT 5 PSI
• LITTLE CONFIDENCE IN ANALYSIS
• SUBELEMENT TEST REQUIRED
• RECOMMEND TEST BE DEFERRED PENDING SUBELEMENT TEST
Figure 53. Internal Pressure Test
Table 1. Lap-Shear Strength of Polyimide Adhesives
AVerage Lap Shear Strength, NPa (psi)
Type Mode of 58yOK M,)J" uf Adhesive Scrim Basic P.1. Resin Sy.stc·m In Failur .. (bOO°F) Failur~
FH34-1B No. I 104 FM34B-IB 19.3 (2800) 80/;L 20;.C 11.9 (J 730) - 100%C FM34-18 No. 2 104 FM34B-IB lS.1 (2190) - 100%C 11.6 (1680) - 100% LARC 13A No. 1 104 LARC 13 NASA process 19.6 (2BSO) 35%L 6S%C lS.6 (2260) 10%L 90%C LARC 13A No. 2 104 l.ARC 13 NASA process, DNF 17 .7 (2450) 40%L 60%C 12.7 (1B45) - 100%C LARC 13A 108 LARC 13 NASA process, DM!' 16.6 (2410) 40%L 60%C 13.0 (lBBO) - 100%C LARC 13A 112 LARC 13 NASA process, DM!' lS.4 (2240) 30%L 70%C 12.S (1BI0) - 100%C LARC BE 104 LARC 13 ester, diglyme* 19.2 (27BO) 40%L 60%C IS.2 (2200) 20%L 80/.C I.ARC BE 108 LARC 13 ester, diglyme 17.1 (2480) 2S%L 7S%C 12.7 (1840) 50%L 95%C LARC 13E 112 !.ARC J3 ester, diglyme 17.1 (2480) SO%L SO%C 12.7 (1840) 10%L 90%C 17-51:: 104 BTDE/NE, 3,3'/4,4' - ~WA~100/0 lS.4 (2240) 80%L 20%C 12.B (1860 ) 15%L 8S%C 17-6E 104 BTDE/NE, 3,3'/4,4' - MDA~BO/20 17 .4 (2S30) 30%L 70%C lS.4 (2240) 9S%L 5i.:C 17-7E 104 BTDE/NE, 3,3'/4.4' - MDA~60/40 IS.B (2300) 20%L 80%C 13.2 (19 20) S%L 9S%C 17-8E 104 BTDE/NE, 3,3'/4,4' - }UlA=40/60 lS.7 (2270) 70%L 30%C lS.7 (22BO) 100%L -17-9E 104 BTDE/NE, 3,3'/4,4' - }WA=20/BO 14.2 (2060) 100%L - 14.6 (2120) 100%L -17-10E 104 BTDE/NE, 3,3'/4,4' - MDA=0/100 16.4 (2380) 30%L BO%C 14.1 (2140) 30%L 70%C 18-6 104 BTDE/NE, 3,3'/4,4' - MDA=25/75 17.1 (2480) 95%L 5%C 16.0 (2330) - 100%C 1B-6A 104 BTDE/NE, 3,3'/4,4' - HDA=25/75 16.0 (23 2S) 90%L 1O;;C 15.8 (2290) - 100%C 1B-6B 104 BTDE/NE, 3,3'/4,4' - }IDA=25/7S lS.5 (2250) 30%L 70/;C 11.6 (16BO) 30%L 70%C 18-6C 104 BTDE/NE, 3,3'/4,4' - HDA=25/75 lS.3 (2220) 95%L 5%C lS.0 (2185) 60%L 40%C 21-5 104 BTDE/NE,H,M' - phenylenediamine 12.0 (1750) 30%L 70%C 13.7 (1990) 60%L 40%C IB-l 112 BTDE/NE, DADS, DMF 12.3 (1790) - 100%C lS.2 (2200) - 100%e 18-2 112 BTDE/NE, DADS/4, 4' - NDA 13.S (1960) - 100%C 11.4 (1660) - 100%C IB-3 104 BTUE/NE, DADS/3,3' - MDA 11.4 (1660) - 100%C 18-4 104 BTDE/NE, DADS, 4,4' - MDA 13.4 (1940) 10%1. 90%C 14.2 (2070) - 100%C 15-1 104 BTDA/NA, DADS, om' 9.9 (1430) lO%L 90%C 15-2 104 BTDA/NA, DADS, diglyme B.9 (1260) - 100%C lS0-1 104 OS6X (NR-1S0B2) 17/0 (2470) 20%L 80%C 13.S (1970) - 100,;C 150-2 No. 1 104 050X (NR-lS0B2) 12.S (1820) - 100%C 12.8 (lB60) - 100%C lS0-2 No. 2 108 050X (NR-1S0B2) 13.5 (1910) 25%L 75%C 12.7 (lB50) 30%1. 70%C 150-3 No. 1 104 OS2X (NR-1S0B2) 15.B (2300) 10%L 90%C 15.2 (2200) 10%L 90%C 150-3 No. 2 lOB 052X (NR-150E2) 15.B (2290) 30%L 70i;C 14.3 (2080) 30%L 70%C 150-4 104 051X (NR-150B2) 15.B (2300) 20%L BO%C 12.7 (1840) 15%L 85%C IB-IB 104 PMR-15. diglyme 16.9 (2450) 70%L 30%C 14.6 (2120) 40%L 60%C 20-1 104 PHR-15, tackifier, dlglyme 14.1 (2050) 30%L 70%C 15.5 (2250) 30%L 70%C 17-11B 104 PHR-15, NR-150B2, UHF 12.7 (1B40) 50%L 95%C 12.1 (1750) - 100%C 17-11C 104 PHR-15, AI-1l30L, DHF 11.2 (1625) - 100%C B.O (1160) - 100%C 1B-12C 104 LARC 160*, dig1yme 12.3 (1790) - 100%C 12.0 (1745) - 100%C 18-8 104 LARC 160, DADS, tackifier 15.2 (2210) 60%L 40%C 15.2 (2200) SO%L SO%e
diglynte 18-12B 104 tARC 160, AI325 17.0 ( 2460) - 100/.C J3 .9 (2020) - 100%C IB-1211 104 tARC 1 bO, A1ll301. 11.4 (J650) - IOOZC 7.7 (J 260) - 100%C 21-4 104 LARC 160,4,4' - HilA, dlglyme 13 .8 (ZOOO) 501;i. 95%C j() .0 (j 560) lOO:tc -
Unidirectional eight-ply laminates Average vf four tests; all adhesives cured two hour>: at 58yOK (600 0 n and 0.6895 MPa (100 psi} pressure plus 10 hours at 5;l9°K (600 0 n All aJherends primed with BR34B-18 after light abrading and diglyme wipe L: Failure in composite laminate, C: Cohesive fa i lure in bond line
\0 N
Table 2 - Flatwise Tensile Strength of Polyimide Adhesives
Flatwise Tensile Fillet, mm (in.) Strength, MFa (psi) (1)
Base Polyimide 5890 K (600°F) Adhesive/Scrim Resin System Top Bottom RT
FM34B-18/104 FM-34 3.0 (0.12) 1.2 (0.05) ].93 (280) 2.34 (340)
LARC- ] 3A/ 104 (Hexcel) LARC-13A 0.50(0.02) 0.50(0.02) 2.51 (365) 2.07 (300)
LARC-13E/112 (Hexcel) LARC-13E 0.50(0.02) 0.50(0.02) 1.86 (270) 2.48 (215)
21-1/104 PMR-15 207 (0.11) 0.50(0.02) 2.72 (395) 2.96 (430)
21-4/ ]04 MOD. LARC-] 60 ] .5 (0.06) 1.2 (0.05) 2.90 (420) 2.31 (335)
]8-6/104 MOD. LARC-13 2.5 (0.10) 0.50(0.02) 2.48 (360) 2.41 (350) 3,3'/4,4'-MDA
18-7/104 MOD. LARC-13 3.8 (0.15) 300 (0012) 2e58 (375) 2.55 (370) 3,3'/4,4'-MDA
18-4/104 BTDE/NE/DADS 3.8 (0.15) 0.7 (0.03) 2.44 (355) 2.44 (355)
18-2/104 BTDE/NE/DADS 1.2 (0.05) 0025(0.0]) 1.72 (250) 1093 (280)
18-12C/I04 LARC-160 3.3 (0013) 2.54(0.10) 2.0 (295) 1.86 (270)
150-3/104 NR-150B(052X) 0.38(0.015) 0.25(0001) 1051 (220) -150-2/104 NR-150B(050X) 3.8 (0.15) 002 (0.0)) 1.37 (200) -
(1) Average of three tests. All adhesive cured (in autoclave) for two hours at 5890K (600°F) under full vacuum and 0.1724 MFa (25 psi) pressure and postcured for 10 hours at 5890 K (600°F) and normal pressure.
Table 3 - Climbing Drum Peel Strength of Polyimide Adhesives
Adhesive Peel Strength at RT(I)
Form No. Resin System g/cm Width in. lb/in.
FM34B-18 FM34 630
LARC-13A LARC-13, anhydride process 756
LARC-13E LAR,C-13, ester process 486
18-12C MOdified LARC-160 90
17-6E MOd. LARC-13. 3.3'/4.4'MDA=80/20 594
17-9E MOd. LARC-13. 3.3~4.4'MDA==20/80 360
17-10E MOd. LARC-13. 3.3'/4.4'MDA=0/100 486
18-6B MOd. LARC-13. 3,3~4,4'MDA=25/75 594
18-4 BTDE/NE.DADS,4,4'MDA 300
(I) HRH' 327-3/16-4 honeycomb core. Celion~OOO/PMR-15 face sheets in
two-ply (0,90) construction
3.5
4.2
2.7
0.5
3.3
2.0
2.7
3.3
1.7
Test specimens, 30.5 em (12 in.) long and 7.6 cm (3 in.) wide, were
tested according to Spec. MIL-A-25463A.
The average skin load was 11.24 Kg (25 lb)
93
94
Table 4. Glass Transition Temperature of Polyimide Adhesives(l)
Adhesive Glass Transition
Form No. Resin System Temperature, Tg, oK (OF) (1)
FM34B-lS FM34B-18 Not measurable
LARC-13 on 112 glass IARC-13 588°K (599°F) scrim NASA Process
l8-6A Mod LARC-13 639°K (690°F) amine substitute
lS-6B Mod LARC-13 629°K (672°F) amine substitute
l7-11B PMR-15 modified 60SoK (634°F) with NR-lSOB2
17-11C PMR-15 modified with Not measurable AI l130L
l8-12J Mod LARC-13 62SoK (671°F) amine substitute
IS-12K Mod LARC-13 59SoK (617°F) amine substitute
(l)Tg measured on a duPont 941 TMA-900TA analyzer at a heating rate of 5°K (9°F) per min. The expansion probe used had a 0.25 cm (O.l-inch) diameter. The adhesives cured for two hours at 5S9°K (600°F) under 0.6895 MFa (100 psi) pressure and postcured for ten hours and 589°K (600°F) under normal pressure.
Table 5. Effect of Out Time in Dry Environment on Lap-Shear Strength
Adhesive
FM34B-18
LARC-13A/104
LARC-13E/104
18-6A (MOD. LARC··13E)
lS-6C
OUt-Time Days
1 5
11
1 5
11
1 5
11
1 5
11
1 5
11
Lap Shear Strength (2) , MPa (psi)
RT
19.3 (2800) 18.4 (2670) 15.9 (2320)
16.6 (2415) 16.6 (2350) 14.7 (2130)
16.0 (2335) 15.7 (2250) 11. 7 (1706)
14.2 (2050) 16.2 (2360) 15.1 (2200)
15.0 (2185) 13.6 (1980) 15.2 (2210)
11.9 (1730) 15.6 (2270) 15.3 (2225)
12.7 (1845) 1~.3 (2085) 14.5 (2100)
14.3 (2080)
14.3 (2075)
13.9 (2010) 16.0 (2330) 13.6 (1980)
15.3 (2225) 13.4 (1950) 15.6 (2270)
(1) Storage condition: 6-40% RH, 291-299°K (65-80 0 F) Adherends: Celion 6000/PMR-15, eight-ply, unidirectional Primer: BR34B-18
(2) Surfaee treatment: light abrading, diglyrnewipe Cure: Two hours at 589°K (6000 F) and 0.6895 MPa (100 psi) pressure Postcure: 10 hours at 589°K, (600o
F) normal pressure
95
Table 6
EFFECT OF OUT-TIME IN HUMID ENVIRONMENT ON LAP SHEAR STRENGTH
Lap Shear Strength, MPa (psi) (1)
Out-Time Mode of (2 Adhesive Days RT Failure (%) 5890K (6000F)
Hode of (2) Failure (%)
FM34B";18
LARC-l3
o 3 7
o 3 7
15.8 (2300) 11.4 (1660) 11. 7 (1700)
15.0 (2175) 16.1 (2340) 15.0 (2190)
L40, C60 LIS, C85 L5, C95
L20, C80 L90, CI0 L30, C70
l3.4 (1950) 12.2 (1770) 11.2 (1620)
13.2 (1910) 16.1 (2330) 14.8 (2140)
L40, C60 LI0, C90 L45, C55
LIS, C8S L80, C20 L50, C50
M-LARC-l3E o 3 7
14.1 (2055) 16.1 (2195) 14.2 (2080)
LIS, C85 L30, C70 L55, C45
12.3 (1790) 1l.8 (1710) 14.0 (2030)
L30, C70 L50, C50 L60, C40
96
(1) Environment: 40-90% RH, 288-294 (60 - 70°F)
Adherends: Ce1ion 6000/PMR-15, 8-p1y, unidirectional
Surface Treatment: Light abrading, dig1yme wipe
Primer: BR 34B-18
(2) Hode of Failure: L - Laminate failure, C - Cohesive failure in adhesive
Table 7 - Comparison of Primers for Bonding Celion 6000/PMR-l5 Panels with LARC-13 Adhesive
Primer Lap Shear Strength MPa (psi) _. __ .. _-----
Form No. Batch No. RT 600°F
BR34B-18 B-673 16.6 (2410) 12.9 (1880)
LARC-13A 8-23-1 14.3 (2085) 11.1 (1623)
LARC-160 8-23-2 14.6 (212S) 11.7 (1705)
PMR-1SE 8-23-3 13.7 (1995) 14.8 (2,150)
NR1S0-B2 8-23-4 IS.3 (223S) 12.9 (1880)
052X 8-23-5 13.5 (1970) 9.6 (1400)
050X 8-23-6 15.6 (2265) 121 (1760)
AI Il30L 8-23-7 13.9 (2030) 11.9 (172S)
AI 830 8-23-8 13.5 (1960) 11.9 (1740)
Adherends: Celion 6000/PMR-·15 composites; eight-ply, unidirectional construction
Adhesive: LARC-13 (NASA process) with style 108 scrim Surface Treatment: Light abrading (Behr Tex) diglyme wipe Average of three or four test specimens Cure: Two hours at 589°K (600°F) and 0.6895 MPa (100 psi)
pressure + 10 hours at 5890 K (600°F) under normal pressure
97
\0 00
Table 8 - Comparison of Surface Treatments for Bonding Celion/PMR-15 Laminates with LARC-13 Adhesive
Surface Bond Line Test Lap Shear Treatment mm (mils) Temperature Strength, MPa
Behr Tex 0.15 (6) RT 20.3 (2950) MEK 589K ( 600°F) 9.4 (1360)
NMP 0.15 (6) RT 17 .9 (2600) 589K (600°F) 15.6 (2260)
MEK 0.19 (8) RT 14.9 (2160) 589K (600°F) 13.4 (1950)
Grit blasting, 0.15 (6) RT 20.5 (2970) diglyme 589K (600°F) 19~3 (2800)
Alkaline 0.15 (6) RT 16.7 (2430) etch 589K (600°F) 15.9 (2310)
Acid etch 0.15 (6) RT 13.6 (2010) 589K (600°F) 15.0 (2180)
Grit blasting, 0.15 (6) RT 17 .6 (2550) (sand)-MEK 589K (600°F) 19.8 (2870)
Adherends: Celion/PMR-15 composite laminates, eight-Plr' unidirectional Cure: 2 hours at 5890 K (600°F) and 0.6895 MPa (100 psi pressure Postcure: Ten hours at 5890 K (60QoF) and normal pressure
Mode (psi) of Failure
laminate cohesive
laminate cohesive
cohesive cohesive
laminate laminate
cohesive cohesive
cohesive cohesive
laminate laminate
Table 9 - Comparison Availabilities and Properties of Po1yimide Adhesives
Ava llabi 11 ties LARC-13 Mod. LARC-13 Mod. PMR-15 and Properties FM34B-18 NASA Process 18-6 20-1
Commercially available Yes No No No Production methOd Yes Yes Yes
developed or under development
Can be supplied by Yes Yes Yes Hexcel on request
Resin processability Fair Good Good Hot melt processability Good Good Good Tack and drape Good Good Good Good Flow on cure Good Good Good Good Filleting ability Good Good Good Good
Lap shear strength 1 Composite-to-composite( ) toTa (psi), RT 15.1-19.3 17.7-19.6 19.2 (2780) 15.3 (2225)
(2190-2800) (2450-2850) 589°K (600°F) 11.6-11.9 12.7 ..... 15.6 15.2 (2200) ,
13.9 (2010) (1680-1730) (1845-2260)
F1atwise tensile strength(l) MFa (psi), RT 1.93 (280) 2.52 (365) 2.48 (360) 2.92 (395)
589°K (600°F) 2.34 (340) 2.07 (300) 2.41 (350) 2.96 (430)
Climbing drum peel strength at RT, '1.1 cm width (2) (in.lb/in.) 630 (3.5) 756 (4.2) 59!, (3.]) 360 (2.0) Glass transition temp., Not 315 (599Op) 366 (6900 F) -
(Tg)OC Measurable Ability to form low porosity Poor Good Good (;ood
large area bonds
(1) Adherends: Ce1ion 6000/PMR-15 composite laminates, eight-ply unidirectional Primer: BR34B-18; surface treatment: light ahrading, dig1yme wipe
(2) Adherends: Celion 6000/PMR-15 composite, two-ply (0.90) construction Primer: BR34B-18; surface treatment: light abrading, diglyme wipe
Mod. LARC-160 NR-150B 18-8 150-3
No No Yes No
Yes I No
Good Good Poor Good Fair Good Fair Good Poor
15.2 (2210) 16.9 (2450)
15.2 (2200)
2.03 (295) 1. 52 (220) 1.86 (270) -
90 (0.5) i -- , -
Good 1 Poor
I-' o o
I.
2.
1.
4.
5.
6.
7.
8.
9.
10.
u. 12.
13.
14 ..
el) (2)
Table 10 - Comparison of PM 34B-18, LARC-13 and LARC-13E Adhesives
- -FM348-1H J.ARC-Il M-LAltC- I )E
. 'ACTOa RATING PTS. -~~. 1.2)' 1U\'l'INC ~f .. (iY
-W'lNG l·T~. fi£ ~-Commercial Availability V. Good 1 2 6 Good 2 2 4 Good 2
Hatedal Coat High 2 2 " lIigh 2 2 4 lIigh 2
Out-TiM Fair I 2 2 V.Good 1 2 6 V.Good 3
Tack and Orape Good 2 2 " Good 2 2 " Good 2
,illetiDI Good 2 1 6 Good 2 2 " Good 2
Too1iaa , .roc.aaiD8 Coat Excel. " 1 12 Fair I 1 1 Fair I
Ability to provide void-free Poor 0 1 0 V.Good 3 3 9 V.Good 1 1arle araa boDda. add-plane bo_iDI
• Ability to provide larae area Fair I 3 3 V.Good 1 1 9 V.Good 1 booded Hie .andwich panela
Lap shear atrenath vs. over- Fair I 3 3 V.Good 3 3 9 V.Good 3 lap shear area
Load-carryin8 capacity in Poor 0 1 0 V.Good 1 1 9 V.Good 3 shear of aid-plane bonded panela
Peel streogth (climbing drum) 1-'air I 2 2 Fair 1 2 2 Fair 1
Ftatviae tensile Str. Hlc Good 2 1 6 Good 2 3 6 Good 2 core to GrlPi composite
ASiDS .tability @5890 K(600oF) Fair I 3 3 Fair 1 3 3 Fair 1
Amount of Industry Experience V.Good 3 :J 9 Good 2 J I) Fair 1 S.,. : Significance Factor -_. __ h. __ . __ ._._. ___________ 1--____ ._
TOt'ALS 6 I 78 T: Total
S.L \~)
2 4
2 4
2 6
2 4
2 " 1 3
1 9
1 9
1 9
3 9
'l 2 .. 3 6
J 3
3 3
15
.... o ....
Panel No.
Prepreg Lot No.
Supplier
Prepreg Physical Properties
1. Fiber areal weight, g/m2
2. Resin solids (%)
3. Calculated thickness per ply, mm (mils)
Composite Physical Properties
1. Specific gravity
2. Fiber volume (%)
3. Void fraction (%)
4. Cured thickness mm (miles)
5. Weight loss on cure (%)
6. Tg OK (OF)
7. C-scan transmission (%)
Table 11. Properties of Ce1ion 6000/PMR-15 Prepreg and Cured Laminates
CPSU-A1 CPSU-A2 CPSU-A3 CPSU-A4 CPSU-A5 CPSU-A6 CPSU-A17 CPSU-A1S CPSU-A19 CPSU-A20 CPSU-A21 CPSU-A22 CP8U-A23 CP8U-A24
2W4474 2W4474 2W4474 2W4474 2W4474 2W4474 CO-0l5 CO-015 CO-015 CO-015 CO-015 CO-015 CO-015 CO-015
USP USP USP USP USP USP Fiberite Fiberite Fiberite Fiberite Fiberite Fiberite Fiberite Fiberite
14S.5 148.5 148.5 14S.5 14S.5 14S.5 156.2 156.2 156.2 156.2 156.2 156.2 156.2 156.2
36.5 36.5 36.5 36.5 36.5 36.5 33.S 33.S 33.S 33.S 33.S 33.S 33.S 33.S
0.127(5) 0.127(5) 0.127(5) 0.127(5) 0.127(5) 0.127 (5) 0.127(5) 0.127(5) 0.127(5) 0.127(5) 0.127(5) 0.127(5) 0.127(5) 0.127(5)
1.5S 1.59 1.59 1.56 1.59 1.59 1.5S loSS 1.5S 1.57 1.59 1.60 1.60 1.60
61.6 63.3 63.6 56.6 63.7 62.9 60.1 61.4 60.3 5S.1 64.3 64.S 64.5 65.4
0.4 0.2 0.3 0.3 0.3 0.0 -0.1 0.3 -0.1 0.0 0.5 -0.1 -0.2 0.1
1. 06 (42) 1.09(43) 1.06(42) 1.19(47) 1.09(43) 1.09(43) 1.14(45) 1.09(43) 1.06(42) 1. 09 (43) 1.43(45) 1.11(44) 1.09(43) 1.11(44)
17.S lS.0 17.9 17.S lS.2 13.0 12.7 12.0 12.0 12.7 11.9 11.S 13.2 13.6
6000K 599°K 5SSoK 5S7°K 599°K 592°K 615°K 616°K 615°K 611 0K 613°K 614°K 60SoK 613°K (620°F) (61S0F) (599°F) (597°F) (61S0F) (606°F) (647°F) (649°F) (64S0F) (640°F) (644°F) (645°F) (635°F) (643°F)
100 100 100 995 100 100 100 100 100 100 100 100 100 100
TABLE 12 - FLATWISE TENSILE STRENGTH OF FM34B-18 LARC-13 AND M-LARC-13E ADHESIVES
Fillets, mm(in.) F1atwise Tensile Test Temp Strength
Adhesive/Scrim Top Bottom OK(oP) MPa(psi) (1)
FM34B-18/104 2.0 (0.10) . 1.0 (0.05) RT 3.17 (460)
2.0 (0.10) 1.0 (0 .. 05) 589 (600) 1.72 ( 250)
LARC-13/104 2.0 (0.10) 1.0 (0.05) RT 3.17 (460)
2.0 (0.10) 1.0 (0.05) 589 (600) 2.76 (400)
M-LARC-13E/l04 1.0 (0.05) 1.0 (0.05) RT 2.76 (400)
1.0 (0.05) 1.0 (0.05) 589 (600) 1.86 (270)
(1) Face Sheets: Ce1ion 6000/PMR-15, 8-p1y, crossp1ied construction
Primer: BR34B-18
Core: HRH327-3/16-4. T=0.5
(2) Core: Rupture in the honeycomb core
C: Cohesive failure in the adhesive
Mode of Failure (%) (2)
Core 5, C95
Core 20, C80
Core 10, C90
Core 50, C50
ClOO
Core 50, C50
TABLE 13 - LAP SHEAR STRENGTH OF LARC-13 ADHESIVE PREPARED BY ANHYDRIDE AND ESTER METHODS
Mode of Test Temp.· Lap Shear Strength Method oK (OF) MPa (psi) Failure (%) (1)
Anhydride RT 17.4 (2525) L95, C5
Anhydride 5890 K (600°F) 16.6 (2410) L80, C20
Ester RT 16.9 (2445) L70, C30
Ester 5890 K (600°F) 18.3 (2660) L80, C20
(1) L: Laminate failure; C: cohesive failure in adhesive
Adherends: Celion 6000/PMR-15, eight-ply, unidirectional
Surface Treatment: Light abrading, diglyme wipe
Primer: BR34B-18
103
I-' ~ Table 14 - Short Beam Shear Strength of LARC-13 and FM34B-18
SHORT BEAM STRENGTH -MPa(psi)
ADHESIVE TEMP 1 i 3 4 5 AVE
LARC-13/l08 R.T. 0.09(14.3) 0.09( 14.4) 0.10(15.5) 0.1005.9) 0.10(15.5) 0.10(15.1) 0.07 Lb/ft2
589 0 K 0.04(6.9) 0.04(6.7) 0.04(6.5) 0.04(6.4) 0.03(5.0) 0.04(6.3) (600°F) .
FM34B-18/104 R.T. 0.0903.0) 0.09(13.3) 0.07 Lb/ft2
0.09( 13 .5) 0.09(13.5) 0.0903.2) 0.09(13.3)
5890 K (600°F) 0.04(6.1) 0.04(6.3) 0.04(6.1) 0.04(6.1) 0.04(6.2) 0.04(6.2)
0) Adherends: Ce1ion 6000/PMR-15~ 8 Ply. unidirectional
Surface Treatment: Light Abrading~ dig1yme wipe
Primer: BR34B-18
Test Methods: MMM-A-132
...... o \J1
Table 15. Slotted Shear Strength of Test Specimens From Mid-Plane Bonded 30.5 em (12".0 in.) by 30.5 em (12.0 in.) Panels
Ultimate Load Lap Shear Strength Adhesive Cm (in.) Test Tempo MPa (psi) MPa (psi)
. FM34B-18 1.27 (0.50) RT 109 275 3 08 (550)
5.18 (2000) 3.2 460 1.6 (230)
1027 (.50) 5890 K (600°F) 1055 225 3.1 (450) 5.18 (200) 504 770 2.7 (385)
LARC-13 1.27 (.50) RT 8.05 1170 16.1 2340 5018 (2.0) 19.6 2840 908 1420
1.27 (.50) 5890 K (600°F) 7.25 1055 14.5 2110 5.lS (2.0) I 1S.2 2640 9.1 1320
Adherends: Celion 6000/PMR-15, 8 ply, unidirectional
Surface Treatment: Light abrading, diglyme wipe
Primer: BR34B-18
Test Method: MMM-A-132
Mode of Failure
ClOO% ClOO%
ClOO% ClOO%
L40, C60 L50, C50
L35, C65 L35, C65
Table 16. Aging (1) Stability of Polyimide Adhesives Wrapped in Aluminum Foil
FLATWISE TENSILE STRENGTH MPa (psi) ( 2)
R.T. 5890K(6000F) Adhes~ve 1 2 3 AVE 1 2 3 AVE
FM34B-18 1.41(203.8) 0.86025.0) 1.38( 200.0) 1.22(176.3) .95032.5) 0.97(140.0) 1.0(146.3) 0.97(139.6)
LARC-13 1.57(227.5) 1.59(2300) 1.17(170.0) 1.44(209.1) 1.10(160.0) 1.55(225.0) 0.88(127.0) 1.18(170.7)
(1) Specimens were aged for 125 Hr. at 589~ (600°F) ~n air circulated oven. All specimens were wrapped in aluminum foil.
(2) All adhesives cured in autoclave for two hours at 5890 K under full vacuum and 0.1724 MPa (40 psi) pressure and post cured for 10 hours at 5890K (600°F) free standing in oven.
Table 17. Aging (1) Stability of Polyimide Adhesives Not Wrapped in Aluminum Foil
FLATWISE TENSILE STRENGTH, MPa (psi) (2) RT 5890 K (6000 F)
i
ADHESIVE 1 2 AVE 1 2 AVE
FM34B-18 0.70(102.5) 0.64(92.5) 0.67(97.5) 0.84(122.5 0.55(80.0) 0.69(101.2)
LARC-13 1.10(152.5) 1.38(200.0) 1. 24 (176 • 3 ) 1.34 (195.0) 1.26(182.5) 1.30(188.8)
(1) Specimens were aged for 125 hours at 589~ (600°F) in air circulated oven w All specimens were no t wrapped in aluminum fo il.
(2) All adhesives were cured in autoclave for two hours at 5890K (600°F) under full vacuum and 0.1724 MPa (40 psi) pressure and post cured for 10 hours at 5890K (6000F) free standing in oven.
I-' 0 00
Table 18 - Aging Stability of Selected Adhesives at 5890K (600°F)
AGING LAP SHEAR STRENGTH, MPa (psi) (1)
PERIODS MODE OF MODE OF FAILURE ADHESIVE HOURS RT FAILURE AT RT(2) 5890K(6000F) AT 5890K (6000F(2)
FM34B-18/l04 0 15.5(2250) L20%, C80% 13.4(1945) L40%, C60% 0.06 lb/ft2 125 8.7(1255) L15, C85 7.9(1145) L5, C95
250 7. 7( l1l5) LlO, C90 8.1(l180) L15, C85
LARC-13/l08 0 14.0(2030) L70, C30 14.5(2100) L35, C65 0.07 lb/ft2 125 9.2(1330) L5, C95 8.3(1210) L5, C95
250 7.0(1020 L5, C95 8.1( l180) ClOO
M-LARC-13E~108 0 15.8(2295) L65. C35 12.7(1845) L25, C75 0.07 lb/ft 125 8.9(290) L15, C85 9.7(1400) LlO, C90
250 6.7(965) L20. C80 8.1(1180 ) LlO, C90
(1) Aging Environment: Lap shear specimens exposed directly to air current in circulating oven at 5890K( 600°F)
Adherends: Celion 6000/PMR-15, 8 ply, unidirectional
Surface Treatment: Light abrading, diglyme wipe
Primer: BR34B-18
(2) L: Laminate failure, C: Cohesive failure
I-' o \0
Table 19 - Effect of Water Immersion on Lap Shear Strength of LARC-13 and FM34B-18Adhesives
TEST LAP SHEAR WATER-IMMERSION TEMP STRENGTH MODE OF
ADHESIVE PERIODS, DAYS °K(Op) MPa (psi) (1) FAILURE
FM34B-18!I04 0 RT 16.5 (2390) L80%, C20% 0.06 Ib/ft2 5890K (600°F) 1104 (1660) L5 , C95 .
7 RT 5890K (6000F) 12.2 (1770) L5 C95
LARC-13/108 0 RT 15.4 (2240) L25 C75 0.07 Ib/ft2 5890K (6000F) 14.9 2165 L5 C95
7 RT 14.4 (2095) L5, C95 5890K(6000F) 12.0 (1745) ClOO
I I (1) Adherends: Celion 6000/PMR-15
8-ply, unidirectional Surface Treatment: Light abrading, diglyme wipe Primer: BR34B-18
(2) Laminate failure, C-Cohesive failure
(2)
Adhesive
LARC-13/l08
LARC-13 I 108
LARC-13 I 108
LARC-13/108
LARC,.... 13 I 108
LARC-13/108
LARC-13 I 108
LARC-13/108
LARC-13 I 108
LARC-13/108
LARC-13 I 108
LARC-13 I 108
LARC-13 I 108
LARC-13/108
LARC-13 I 108
LARC-13/108
No. of Layers
1
1
2
2
3
3
4
4
5
5
6
6
7
7
8
8
Table 20. Effect of Bondline Thickness on LaplShear Strength(l)
Test Temp.
oK (OF)
R.To 589°K (600°F)
R.T. . 589°K (600°F)
R.T. 589°K (6-()OOF)
R.T. 589°K (600°F)
R.T" 589°K (600°F)
RoTo 589°K (600°F)
R.T. 589°K (600°F)
R.T. 589°K (600°F)
Bond line Thickness
After Cure
0.0020
0.0015
0 0 0050
0.0040
0 0 0055
0.0050
0.008
0.008
0.010
0.011
0.012
0.012
Lap Shear Strength-PSI 1 Z 3;
i AVE
2590
2400
2790 I I
2300
2700
I 2600
2693
2433
2830 2880 3150 2955
2020 2550 2240 2270
2760 2630 2790 2725
1740 2580 1760 2026
2340 2870 2460 2556
1700 1740 1990 1810
2800 3280 3140 3073
1920 1910 1910 1913
2900 2530 1960 1796
1460 1560 1660 1560
1680 1350 1580 1536
1120 1250 980 1116
1590 1600 1530 1573
1340 1450 1300 1363
Mode of Failure % (2)
L95 C5
L75 C25
L95 C5
L50 C50
L90 C10
L5 C95
L85 CIS
L25 C75
L45 C55
L10 C90
L50 C50
LlO C90
L80 C20
LlO C90
L60 C40
L5 C95
(1) Adherend: Ce1ion 6000/PMR-15, 8 ply unidirectional Primer: BR-34B-18 Cure: 2 hours at 5890K (600°F) and 0.6895 MPa (100 psi) pressure, Postcure: 10 hours at 5890 K (600°F) and normal pressure.
(2) L: Laminate failure, C: Cohesive failure in the adhesive
Table 2l. Out of Refrigeration Effect on Filleting Properties of LARC 13(2)
,
Test Fillet Flatwise Tensile Spec Out of Refr igera tion Temp. Thickness Strength (1) Mode of No. Time oK (OF) mm (Inch) MPa (psi) Failure
HCS104 1 Hl: • R.T. 0.8(0.035") 2 0 34(340) Core to Adhesive HCS104 1 n 589°K II 2.24 (326) II II II
(600°F) HCS105 2 Hrs. R.T. 0.70(0.030") 2.48(360) Core to Adhesive HCS105 " II 589°K II 2.20(320 II II II
(600°F) HCS106 8 Hrs. R.T~ 0.5(0.020") 4.06(590) Core to Adhesive HCS106 " II 589°K " 2.27(330) " " II
(600°F) HCSI09 8 Days R.T. 0.2(0.010") 3.37(490) Core to Adhesive HCS109 11 II 589°K II 2.58(375) II u !!
(600°F)
(1) Face-Sheets: Ce1ion 6000/PMR-15 8-p1y, crossp1ied Construction
Primer: BR34B-18
Core: HRH 327-3/16-4, T = 0.5
(2) LARC-13/108, 0.07 1b/ft2
I-' I-' N
Table 22. Use of Catalyst in LARC 13 to Reduce Cure Temperature
ADHESIVE CURE TEST LAP SHEAR STRENGTH-PSI~~ MODE OF FAILURE TEl1P TEMP. I 2 "3 AVE %
LARC-J3( I) 5280 K e;oOoF) R.'l. I
2260 2320 2460 2346 L25 C75 5280 K (500°F) 600°F 2340 2040 2420 2266 L5 C95
LARC-13(1) 5050 K (450°F) R.Ta 1380 1530 1740 1550 L65 C35 600 F 830 1320 1600 1050 - CIOO"
LARC_I/ 2) 5280 K (500°F) R.T. 1980 1860 2160 2000 LIOO C-1 600°F 1640 1820 1840 1766 L20 cso I
. LARC-13(2) 5050 K (450°F) DID NOT GET CURED TOO RITTLE. PANEL SEPARATED
(I) Catalyst: USP-138, 2% by Weight
(2) Catalyst: Experox - 10, 2% by Weight
(3) Adherend: Celion/PMR-15, 8 Ply Unidirectional Primer: BR34B-18
Table 23. Orbiter Body Flap/Demo Segment Stress Comparison
BODY FLAP (ULT}MPA (KSD LOAD CONCEPT NO. 8 MPA (KSI)
MAX MAX STRESS MIN STRESS MAX SHEAR
RT RT MAX MIN SHEAR STRESS STRESS STRESS
241 -73.0 STABILITY RIB CAP (35.0) (-10.6)
FRONT SPAR CAP 43.4 -48.2 (6.3) (-7.0)
REAR SPAR CAP 9.6 -31.7 (1.4) (-4.6)
STABILITY RIB WEB 37.9 (5.5)
FRONT SPAR WEB 43.4 (6.3)
REAR SPAR WEB 33.7 (4.9)
HIC COVER PANELS 91.7 -82.7 53.8 (13.3) (-12.0) ( 7.8)
r:... I-" \.oJ
This Page Intentionally Left Blank
APPENDIX A
This appendix contains the Engineering Design drawings which define the
TDS. The drawing list is as follows:
SS79-00249 2 Sheets
SS79-00250
SS79-00251 2 Sheets
SS79-00252
SS79-o0253 3 Sheets
Technology Demonstration Segment -Graphite/Polyimide, Shuttle Body Flap
Cover - Body Flap Segment
Rib-Stability Technology Demonstration Segment, Graphite/Polyimide
Leading Edge Panels - Body Flap Segment
Spar Webs - Body Flap Segment
115
--I 14 13 4 12 --11------, 10 ------, 9 I =h ___ a ._1 u ____ 7 6
<:. SvMMETRY
I I
. .
-'S~7'7-00,~OO4 , Z"lEQ'O / ,
/
SS79-C~3-002.
/
' 2 R£Q' 0 ,
/ / I' -' . i -
I ' I •• I
I ,. I. 1 ' .
r @ NEA~ !!oIOE. IqO~ ~~ :::J ,A HO\.E :OAR. :;10:: w .... ::. r:581C-3T-I3' ROI!>.!)-IOOZ-OOO~ MDIIA-loa, -1004 I!) PL...ACE,S
56 _I
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"'"0
,
~
_...:.~_-;;_;;;;;;;:;; _~ __ ~~_-='~:~'9f------t ---I I I ,I I
II 9(aUA~ 5~
I 2. PLACES
I
.:.0,.---- -----
6 ECLlA.L 5PACE~ • !>T~E~ED
Sl79-OOZ.:l2--oQ
Q9
I x.=.!IOOQ
I"'~"'" V ~~r' r r5S1"OD' - I I / 3 R ~5I'OOI -'- \ I - oo~n
EQO I-~' J.. _ 1!tl IZP,Aa'S -
- I -, 6.9Z
<t 5'<'MMETRY
. ,
" I r=t=-t-t+1Hi+t--+-t4--H-lH-!..-' J -~ .. CT.'POSm I r .... 1
.t;~~~~~~~~~g~' ~I £~i' ~ SS7<rOOz::.=l-OO;:)
r---- ------,(, .5YWIl-~
1.50 ~ I ~~O~1J I I I ~--:
L~S79 .. DOZ;:)!l~ "cbo\ClPPOSITt 2. RECI' D '-='
~ SS79-ClOZ!IO-009
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1. R.port No
NASA CR 165839 T 2. Governmtnt ACCllSlon No. 3. Recipient's ~t.alog No
4 Tltl. and Subtltl. Develop, Demonstrate and Verify Large Area Composite
5. Repon Dlte
May 1982 Structural Bonding with Polyimide Adhesives.
7 Author(s)
Bashir D. Bhombal, Donald H. Wykes, Keith C. Hong, Arnold A. Stenersen
9 Performing Organization Name and Address
Rockwell International 12214 Lakewood Blvd. Downey, Ca. 90241
6. Performing OrganlZltlon Code
8 Performing Organization Repon No
10 Work Unit No
11 Contract or Grant No
NASI-15843 r.::-::----:---:~-----------------------1'3 Type of Repon and Period Covered
12 Sponsoring Agency Nam. and Address
National Aeronautics and Space Administration -Washington, D.C. 20546
Contractor Report 14 Sponsoring Agency Code
15 Supplementary Notes Langley Technical Monitor: Robert M. Baucom Final Report
16 Abstract
This program was conducted to develop and demonstrate processes for large area structural bonding with polyimide adhesives. The program consisted of two parts: Process Development for PI resin systems and fabrication of Demonstration components. Process development included establishing quality assurance of the basic PI adhesives, non-destructive inspection of fabricated components, developing processes for specific structural forms and qualification of processes through mechanical testing. In the second part of the program, demonstration components were fabricated using the processes developed in part one. The demonstration components consisted of adhesively bonded honeycomb sandwich cover panels, ribs, and leading edge covers of Celion graphite/LARC-160 polyimide laminates, and a Technology Demonstration Segment (TDS) representative of the Space Shuttle aft body flap.
17 Key Words (Suggested by Author(s)) 18 Distribution Stl,ement
Polyimide Resin Systems, Celion/LARC-160, Graphite/Polyimide Polyimide Adhesive Bonding Non-Destructive Inspection
19 Secunty oasslf (of thIS report)
Unclassified
20 Security elaUlf (of thIS PIli')
Unclassified
Unclassified - Unlimited
21 No of Pages
134 22 Price
H-305 For sale by the National Technical Information Service. Springfield. Virginia 22161
End of Document