FAA Joint Advanced Materials & Structures (JAMS) Center of Excellence 8th
Annual Technical Review Meeting April 4th, 2012
1
CACRC DEPOT BONDED REPAIR INVESTIGATION – ROUND ROBIN TESTING (Program Status)
John Tomblin and Lamia Salah
National Institute for Aviation Research, Wichita, KS
Curtis Davies, Lynn Pham FAA William J. Hughes Technical Center, Atlantic City International Airport, NJ
ABSTRACT
The use of fiber reinforced composites in aircraft structural components has significantly
increased in the last few decades due to their improved specific strength and stiffness and
superior corrosion resistance and fatigue performance with respect to their metal counterparts.
These materials also improve airline profitability in terms of lower operating and maintenance
costs. With the migration of these new materials from secondary components to primary flight
critical structural elements, new challenges associated with the repair of these structures are
continually arising and increasing in complexity. Rigorous material, structural and process
substantiation and validation is crucial to ensure the structural integrity of these bonded
structures.
The main objective of this research program is to evaluate the static and residual strength
after cyclic loading of OEM vs field bonded repairs applied to composite sandwich structures,
performed at different operator depots using existing CACRC (Commercial Aircraft Composite
Repair Committee) standards for composite repair. A second task will evaluate process
parameters such as pre-bond contamination and ineffective heat application during cure on the
residual strength of these repairs. The variability due to repair implementation at various
operator depots as well as the fatigue performance under severe temperature and moisture
environments are investigated and key elements in process and implementation of bonded repairs
are reported.
INTRODUCTION
Composites have many advantages for use as aircraft structural materials including their high
specific strength and stiffness, resistance to damage by fatigue loading and resistance to
corrosion. Thus, extensive use of composites should reduce the high maintenance costs
associated with repair of corrosion damage normally associated with conventional aluminum
alloys. Similarly, costs associated with the repair of damage due to fatigue should also be
substantially reduced, since composites do not, in general, suffer from the cracking encountered
with metallic structures.
As more composites are increasingly used on aircraft components, new challenges associated
with the use of these new materials are continually arising. These challenges are primarily
focused towards the migration of composite repairs, the majority of which was previously in
control surfaces and fairings, to the fuselage, wings and other safety critical primary structure.
As most repair depots and maintenance facilities prepare for this migration, the philosophy and
training necessary to ensure the structural integrity and durability of these repairs will continue to
FAA Joint Advanced Materials & Structures (JAMS) Center of Excellence 8th
Annual Technical Review Meeting April 4th, 2012
2
increase. These repairs will affect the new general aviation business jet aircraft and smaller
piston driven planes as well as large commercial transport aircraft. Numerous studies have
demonstrated the importance of rigorous robust processes to ensure the structural integrity of
bonded repairs to composite structures [1].
The proposed project will investigate the effects of several bonded repair variables and
characterize the strength of the repairs using various experimental methods to determine the
effectiveness of these repairs. The repairs will be representative of typical OEM and field repairs
in an attempt to characterize the quality of the repair and if any deviations in the processing and
repair techniques at the depot can result in poor repair performance. The methods and repair
procedures proposed by the Commercial Aircraft Composite Repair Committee (CACRC) will
be utilized whenever possible and input will be provided to the FAA which can be used in
general guidelines for bonded repair and also placed into training curriculum for courses on
composite repair.
Furthermore, the current NDI methods cannot provide absolute assurance of bond integrity (i.e.,
may fail to detect a weak bond due to poor surface preparation, pre-bond moisture, under or
over-cure, surface contamination, etc…). As a consequence, a substandard repair is not detected
until it actually disbonds, leading to a possible failure of the repaired part. It is therefore
essential to quantify the performance of these weak joints and draw attention to the need for
appropriate training in the composite repair community. This will help identify the degree of
criticality of the different steps within a bonded repair and subsequently lead to more rigorous
repair procedures.
RESEARCH OBJECTIVE/ METHODOLOGY
The main objective of the proposed research program is to evaluate the ultimate strength and
durability (mechanical loading) of OEM vs field bonded repairs applied to composite sandwich
structures, performed at different operator depots using OEM and CACRC standards for
composite repair implementation and technician training. The primary goal is to investigate the
effectiveness of OEM vs field repairs and the variability due to repair implementation at various
operator depots, to identify key elements in the implementation of bonded repairs that ensure
repeatability and structural integrity of these repairs and to provide recommendations pertaining
to repair technician training and repair process control. This program will also investigate the
static strength and fatigue/ durability of damaged repairs and “substandard” repairs bonded to a
contaminated surface or subjected to a nonstandard cure cycle.
The following is a list of all OEMs/ airline depots and the POCs that are participating in the
CACRC round robin and process parameter investigation:
• Northwest/ Delta Airlines (Ray Kaiser, [email protected])
• United Airlines (Eric Chesmar, [email protected])
• US airways (Mike Tallarico, [email protected])
• Aviation Technology Associates (Marc G Felice, [email protected])
• Spirit Aerosystems (Mike Borgman, [email protected]; John Welch,
[email protected]; Brian Kitt, [email protected])
• Airbus (Francois Museux, [email protected])
• Boeing (Rusty Keller, [email protected])
FAA Joint Advanced Materials & Structures (JAMS) Center of Excellence 8th
Annual Technical Review Meeting April 4th, 2012
3
• Hexcel (Justin Hamilton, [email protected])
All parent materials are supplied by the OEM and all panel fabrication is conducted at NIAR/
NCAT facility using OEM approved processes.
Panel Manufacturing Procedure
A total of 50 large panels 36”x48” are manufactured for the purpose of this investigation as
shown in Figure 1 below. The parent substrate is 4-ply sandwich with 3/16” core cell size, 2”
thick for the large beams and 1” thick for the small beams. The parent material is T300/934 PW
graphite epoxy prepreg with FM377U adhesive. Parent and repair materials and corresponding
specifications are summarized in Table 1 below. Panel stacking sequence and identification are
summarized in Tables 2, 3 and 4 below.
The panels are cured in two operations: a first operation where facesheet 1 is cured with the core
(including the potting compound) and a second operation where facesheet 2 is bonded to
assembly 1. The part is cured at 45 psi at 355 +/- 10°F for two hours. A representative cure
cycle is illustrated in Figure 2 below. All panels were manufactured according to OEM
specifications. A facility audit and process review and approval was conducted by the OEM
prior to building the panels.
CL Symmetric
Potting Material
0°
90°
45°
Gauge Section
7.00
4.00
36.00
48.00
6.00
CACRC-002-0102-RTA-01 CACRC-002-0102-RTA-02 CACRC-002-0102-RTA-03
FIGURE 1. TEST PANEL GEOMETRY
FAA Joint Advanced Materials & Structures (JAMS) Center of Excellence 8th
Annual Technical Review Meeting April 4th, 2012
4
0
50
100
150
200
250
300
350
400
0 50 100 150 200 250 300 350 400 450 500
Te
mp
era
ture
(°F
)
Time (min)
Leading Thermocouple vs. Lagging Thermocouple
Leading TC
Lagging TC
FIGURE 2. CURE CYCLE FOR CACRC-034-1702 FACESHEET 1
Lay-up Procedure
All articles are fabricated with T300/934 prepreg FM377S adhesive, potting compound Cytec
Corfill 658 and HRP -3/16-8 core with 2 layers of film adhesive.
Uncured facesheet 1 is assembled with film adhesive and core as shown in Figures 3 through 11
below. This is called assembly 1. Corfill 658 is subsequently used to fill the core cells and
assembly 1 is then bagged and cured.
TABLE 1. PARENT AND REPAIR MATERIAL SPECIFICATIONS
Parent Materials Vendor Specification Alternate
Specification
Prepreg Cycom 934 PW T300 3K N/A
Film Adhesive FM 377U Adhesive Film 0.055 psf N/A
Core HRP-3/16-8.0 N/A
Potting Compound Cytec Corefil 658 SMS-7
Repair Materials Vendor Specification Alternate
Specification
FAA Joint Advanced Materials & Structures (JAMS) Center of Excellence 8th
Annual Technical Review Meeting April 4th, 2012
5
Prepreg Cycom 934 PW T300 3K N/A
Film Adhesive FM 377S Adhesive Film N/A
Cytec Fabric T300 3K PW N/A
Paste Adhesive EA9396C2 N/A
Film Adhesive EA 9696 N/A
Laminating Resin Huntsman Epocast 52 A/B SAE AMS 2980
Hexcel Fabric G0904 D1070 TCT SAE AMS 2980
Prepreg Hexply M20/G904 SAE AMS 3970
Film Adhesive EA9695 SAE AMS 3970
Specification
TABLE 2. STACKING SEQUENCE, CACRC PANELS CACRC 0101 THRU 2002
Stacking Sequence
P1 0°/90° PW
P2 ±45° PW
P3 ±45° PW
P4 0°/90° PW
TABLE 3. CACRC ROUND ROBIN PANEL ID
Panel ID Core ID Geometry
CACRC-001 thru 40 0101 thru 2002 Figure 1
Panel List, Large Beams 2” thick Core
CACRC-001-0101 CACRC-011-0601 CACRC-021-1101 CACRC-031-1601
CACRC-002-0102 CACRC-012-0602 CACRC-022-1102 CACRC-032-1602
CACRC-003-0201 CACRC-013-0701 CACRC-023-1201 CACRC-033-1701
CACRC-004-0202 CACRC-014-0702 CACRC-024-1202 CACRC-034-1702
CACRC-005-0301 CACRC-015-0801 CACRC-025-1301 CACRC-035-1801
CACRC-006-0302 CACRC-016-0802 CACRC-026-1302 CACRC-036-1802
CACRC-007-0401 CACRC-017-0901 CACRC-027-1401 CACRC-037-1901
CACRC-008-0402 CACRC-018-0902 CACRC-028-1401 CACRC-038-1902
CACRC-009-0501 CACRC-019-1001 CACRC-029-1501 CACRC-039-2001
CACRC-010-0502 CACRC-020-1002 CACRC-030-1502 CACRC-040-2002
TABLE 4. CONTAMINATION INVESTIGATION PANEL ID
FAA Joint Advanced Materials & Structures (JAMS) Center of Excellence 8th
Annual Technical Review Meeting April 4th, 2012
6
Panel ID Core ID Geometry
CACRC-041 thru 50 0101 thru 0502 Figure 1
Panel List, Small
Beams 1” thick Core
CACRC-041-0101
CACRC-042-0102
CACRC-043-0201
CACRC-044-0202
CACRC-045-0301
CACRC-046-0302
CACRC-047-0401
CACRC-048-0402
CACRC-049-0501
CACRC-050-0502
FIGURE 3. FACESHEET 1 LAY-UP
FIGURE 4. FACESHEET 1 ADHESIVE APPLICATION
FAA Joint Advanced Materials & Structures (JAMS) Center of Excellence 8th
Annual Technical Review Meeting April 4th, 2012
7
FIGURE 5. CORE POTTING USING CORFILL 658
FIGURE 6. CORE POTTING, VACUUM APPLICATION
FIGURE 7. MASKING TAPE REMOVAL
FAA Joint Advanced Materials & Structures (JAMS) Center of Excellence 8th
Annual Technical Review Meeting April 4th, 2012
8
FIGURE 8. POTTED PANEL
FIGURE 9. POTTED CORE TRANSFER ONTO FACESHEET 1 (ASSEMBLY 1)
FIGURE 10. RELEASE FILM AND FAIRING BAR APPLICATION
FAA Joint Advanced Materials & Structures (JAMS) Center of Excellence 8th
Annual Technical Review Meeting April 4th, 2012
9
FIGURE 11. FINAL ASSEMBLY
Uncured facesheet 2 is assembled with uncured film adhesive onto assembly 1 as shown in
Figures 12 through 16 below. This is the final panel assembly. The final assembly is also be
cured with facesheet 2 against the tool side using the same cure cycle.
FIGURE 12. FINAL ASSEMBLY
FAA Joint Advanced Materials & Structures (JAMS) Center of Excellence 8th
Annual Technical Review Meeting April 4th, 2012
10
FIGURE 13. FINAL ASSEMBLY
FIGURE 14. FINAL ASSEMBLY BAGGING
FIGURE 15. FINAL BAGGING
FAA Joint Advanced Materials & Structures (JAMS) Center of Excellence 8th
Annual Technical Review Meeting April 4th, 2012
11
FIGURE 16. CACRC -001-0101 PANEL
Specimen Design Validation
CACRC-002-0102 panel was used for specimen design validation/ verification. Three sandwich
elements were machined and tested and the experimental results are summarized in table 5
below. The sandwich beams were instrumented using 7 strain gages to monitor strain
distribution during loading and compression failures were observed for all three elements with a
corresponding average ultimate strain of 9328 microstrain (strain gage 6) consistent with
predictions.
TABLE 5. CACRC ROUND ROBIN TESTING PANEL LIST
Specimen and Test Set-Up Conformity
All panels manufactured per Table 3, are machined into 3 large beams 11.5”x 48” as illustrated
in Figures 17 and 18. Specimen conformity will be conducted at NIAR by an OEM delegated
DMIR. The element dimensions will be conformed according to Figure 17. The dimensions will
be measured and verified by OEM QA department.
Test Set-Up and all lab equipment that is used for these tests will be conformed. This includes
all load cells, LVDTs, displacement transducers, environmental chambers and thermocouples.
Specimen # T W L Span Ultimate Actuator Deflection S1 S2 S3 S4 S5 S6 S7(in) (in) (in) (in) Load (lb) (in) (in) µεµεµεµε µεµεµεµε µεµεµεµε µεµεµεµε µεµεµεµε µεµεµεµε µεµεµεµε
FAA-CACRC-002-0102-RTA-1 2.1 11.6 48.0 42 7269 1.47 1.79 -9699 8801 -9428 8662 -9480 -9653 -9915
FAA-CACRC-002-0102-RTA-2 2.1 11.9 48.0 42 7349 1.44 1.77 -9277 8669 -9485 8520 -9914 -9653 -9293
FAA-CACRC-002-0102-RTA-3 2.1 11.3 48.0 42 6510 1.35 1.65 -8532 7992 -8650 8309 -8812 -8677 -8620
Average 7043 1.42 1.74 -9169 8487 -9188 8497 -9402 -9328 -9276
Standard Deviation 463 0.06 0.07 591 434 467 178 555 564 648
COV% 6.6 4.4 4.2 6.4 5.1 5.1 2.1 5.9 6.0 7.0
FAA Joint Advanced Materials & Structures (JAMS) Center of Excellence 8th
Annual Technical Review Meeting April 4th, 2012
12
4.0
11.5+0.1-0.0
8.0 5.0
48.0+0.0-0.5
-A-
// A 0.1
FIGURE 17. LARGE BEAM DRAWING
FIGURE 18. LARGE BEAM MACHINING
FAA Joint Advanced Materials & Structures (JAMS) Center of Excellence 8th
Annual Technical Review Meeting April 4th, 2012
13
CACRC Round Robin Testing Investigative Plan
A total of 105 elements will be used for the CACRC Round robin repair investigation. The
proposed test matrix is summarized in Table 6 below. Four repair systems are considered, an
OEM repair system using the parent material and adhesive for repair (T300/934 PW with FM
377 adhesive, 350°F cure repair, labeled as OEM-R1), a wet lay-up repair system using Tenax
HTA 5131 200tex f3000t0 fabric with EA9396 C2 laminating resin and EA9696 adhesive
(labeled as OEM-R2) and two CACRC field repair systems using Hexcel M20/G904 prepreg
(250°F cure repair, labeled as CACRC-R1) and using Tenax HTA 5131 200tex f3000t0 fabric
with Epocast 52A/B (200°F cure repair, wet lay-up) (labeled as CACRC-R2). The two
dimensional repairs, illustrated in Figure 19, are typical of either OEM or field repairs. These
repairs will be quantitatively compared to baseline pristine unrepaired coupons, unrepaired
coupons with a 2.5” hole diameter and OEM repaired coupons.
4.00
CL Symmetric
Core Fill (4 places)
11.50
Ø7.50
8.005.00
1/8" cell, 3 pcf 3/16" cell, 7 pcf
Ø2.50
FIGURE 19. LARGE BEAM CONFIGURATION
TABLE 6. CACRC ROUND ROBIN TEST MATRIX
Repair
Station
Coupon
ConfigurationRepair Material Loading Mode
Static
RTA
Static
ETW
Fatigue
ETWN/A Pristine/ Undamaged N/A Compression 3 3 3
N/A 2.5" hole N/A- Open Hole Compression 3 3
OEM/ NIAR Repair/ 2.5" hole OEM-R1 Compression 3 3
OEM/ NIAR Repair/ 2.5" hole OEM-R2 Compression 3 3
OEM/ NIAR Repair/ 2.5" hole OEM-R2 Tension 3 3
OEM/ NIAR Repair/ 2.5" hole CACRC-R1 Compression 3 3
OEM/ NIAR Repair/ 2.5" hole CACRC-R1 Tension 3 3
OEM/ NIAR Repair/ 2.5" hole CACRC-R2 Compression 3 3
OEM/ NIAR Repair/ 2.5" hole CACRC-R2 Tension 3 3
Field Station 1 Repair/ 2.5" hole CACRC-R1 Compression 3 3
Field Station 1 Repair/ 2.5" hole CACRC-R2 Compression 3 3
Field Station 2 Repair/ 2.5" hole CACRC-R1 Compression 3 3
Field Station 2 Repair/ 2.5" hole CACRC-R2 Compression 3 3
Field Station 3 Repair/ 2.5" hole CACRC-R1 Compression 3 3
Field Station 3 Repair/ 2.5" hole CACRC-R2 Compression 3 3
Field Station 4 Repair/ 2.5" hole CACRC-R1 Compression 3 3
Field Station 4 Repair/ 2.5" hole CACRC-R2 Compression 3 3
105
FAA Joint Advanced Materials & Structures (JAMS) Center of Excellence 8th
Annual Technical Review Meeting April 4th, 2012
14
Detailed repair procedures are being drafted and will be forwarded to the OEMs, participating
airline depots/ operators and FAA POCs for review. Upon approval, these specific repair
instructions will follow the panels to the repair stations. All coupons, upon repair, will be
characterized using non-destructive inspection using various techniques. The coupons will be
sent to SNL (Sandia National Laboratories) for inspection and detection of potential weak bonds.
Mechanical testing will be used to validate the NDI data.
Coupon Moisture Conditioning All ‘wet’ conditioned samples will be exposed to elevated temperature and humidity conditions
to establish moisture equilibrium of the material. Specimens will be exposed to 85 ± 5 %
relative humidity and 145 ± 5 °F until an equilibrium moisture weight gain of traveler, or
witness coupons is achieved. ASTM D5229 procedure C [2] is used as a guideline for
environmental conditioning and moisture absorption.
Effective moisture equilibrium is achieved when the average moisture content of the traveler
specimen changes by less than 0.02% for two consecutive readings within a span of 7 ± 0.5 days
and is expressed by:
Wi = weight at current time
Wi-1 = weight at previous time
Wb = baseline weight prior to conditioning
Coupon Instrumentation Strain gages will be installed using AE-10 adhesive per NIAR CP5401. Strain gages shall be
applied in all repair elements in seven locations as shown in Figure 20 below. Strain gages 1, 3,
5 and 6 are installed in the compression surface (repair surface) whereas strain gages 2 and 4 are
installed in the tension surface. A deflection transducer will be used at the center of the beam to
monitor beam deflection.
11.50
5.00 7.00 4.00
CL SymmetricPot Core (4 places)
Gauge Section
2.00
5.00
1,2
3,4
6 5
FIGURE 20. CACRC COUPON STRAIN GAGE LAYOUT
%2 0.0 <
W
W - W
b
1 - ii
FAA Joint Advanced Materials & Structures (JAMS) Center of Excellence 8th
Annual Technical Review Meeting April 4th, 2012
15
Mechanical Testing
The purpose of this task is to validate existing repair processes and demonstrate the
effectiveness, capability and repeatability of field versus OEM repairs. Six specimens from each
repair configuration will be repaired at the OEM or a given airline depot. All specimens will be
tested at elevated temperature wet (defined as 180°F) for ultimate strength and residual strength
after fatigue. Fatigue strain will be derived from the static testing and the sandwich elements
will be cycled for 165000 cycles followed by residual strength evaluation.
A custom-made four-point bending fixture, as shown in Figure 21, will used for mechanical
testing. Load will be applied using two cylindrical upper steel bearings which are in contact with
the upper facesheet of the coupon such that the load applied is uniformly distributed along the
areas of contact of the bearings with the specimens. The lower steel bearings act as simple
supports for the large beam elements. The four-point bending fixture is set-up with an 18”
loading span and 44” support span as shown in Figure 22 and uses a 55-kip servohydraulic
actuator for loading. The test machine is calibrated periodically according to the ASTM E4 [3]
standard to ensure the accuracy of load and displacement readings. Data acquisition will be
performed using the Basic Testware software. The data acquired corresponds to actuator load,
displacement, deflectometer, and strain gage readings. The element alignment will be checked
prior to testing. Structural tests will be conducted when all strain gage readings are within 10%.
All static tests will be conducted, following the guidelines of ASTM D7249-06 [4], under
displacement control at a rate of 0.2-0.25in/min in order to reach the maximum load between
three to six minutes. A deflectometer will be used to monitor the bending deformation at the
centerline of the coupons.
FIGURE 21. ISOMETRIC VIEW OF FOUR-POINT BENDING TEST FIXTURE.
FAA Joint Advanced Materials & Structures (JAMS) Center of Excellence 8th
Annual Technical Review Meeting April 4th, 2012
16
FIGURE 22. LARGE BEAM TEST SET-UP
CACRC Process Parameter Investigation
The purpose of this task is to investigate the effects of process deviations on the strength and
durability of bonded OEM and field repairs. Defective repairs will be created by contaminating
the scarfed surfaces prepared for bonding and deviating from the recommended cure cycles. The
defective repairs/ weak bonds will be sent to Sandia National Laboratories (SNL) for inspection.
NDI results will be validated with subsequent mechanical static and residual strength testing.
A total of 315 coupons will be manufactured for the purpose of this investigation. The parent
substrate is 4-ply sandwich with 1/8” core cell size, 1” thick. The parent material is T300/934
graphite epoxy prepreg with FM377 adhesive. Three repair systems are considered, an OEM
repair system (labeled as OEM) using the parent material and adhesive for repair (350°F cure
repair, prepreg) and two field repair systems using Hexcel T300/M20 prepreg (250°F cure repair,
prepreg) (labeled as R1) and using Epocast 52A/B (200°F cure repair, wet lay-up) (labeled as
R2). Other alternate wet lay-up resins include EA 9396 C2 laminating resin. The proposed
coupon configuration is shown in Figure 23 below. The coupon is a small beam 4” wide by 24”
long with a 1-D repair as shown in the Figure.
Process variables considered include exposure to various contaminants and cure cycle deviations.
Contaminants considered are water (WA75), Hydraulic fluid (HF) and a mixture of skydrol and
water. Water panels are exposed to moisture at 85%RH at a temperature of 145°F until moisture
equilibrium is achieved. Once moisture equilibrium is achieved, panels are subsequently dried to
achieve %saturation levels of 75% (corresponding to a equivalent 0.75% moisture weight
percent). Panels contaminated with hydraulic fluid will be soaked in the fluid for 3 months at
elevated temperature and contaminant uptake will be monitored. Panels contaminated with
skydrol-water mixture will be soaked in the fluid for 3 months and contaminant uptake will be
monitored.
FAA Joint Advanced Materials & Structures (JAMS) Center of Excellence 8th
Annual Technical Review Meeting April 4th, 2012
17
The effects of cure cycle deviations will also be investigated. Two potential cure cycles
deviations will be simulated; an over-cure scenario where selected repairs will be subjected to
multiple cures and/or temperatures exceeding the target cure temperature and an interrupted cure
scenario where repairs will be subjected to a heat blanket and vacuum failure during cure.
Detailed test matrix is shown in Table 7 below.
FIGURE 23. SMALL BEAM CONFIGURATION
TABLE 7. PROCESS PARAMETER/ DAMAGE TOLERANCE INVESTIGATION TEST MATRIX
Variables Repair Loading Mode CTD RTA 180W RTF 180WF
OEM-R1 Compression 3 3 3 3 3
Baseline Repair CACRC-R1 Compression 3 3 3 3 3
E parent = E repair CACRC-R2 Compression 3 3 3 3 3
OEM-R1 Tension 3 3 3 3 3
Baseline Repair CACRC-R1 Tension 3 3 3 3 3
E parent = E repair CACRC-R2 Tension 3 3 3 3 3
OEM-R1 Compression 3 3 3 3 3
Parent/ Repair Stiffness Mismatch CACRC-R1 Compression 3 3 3 3 3
CACRC-R2 Compression 3 3 3 3 3
OEM-R1 Compression 3 3 3 3
Impact (BVID) CACRC-R1 Compression 3 3 3 3
Inclusions CACRC-R2 Compression 3 3
OEM-R1 Compression 3 3 3 3
Contaminant 1: CACRC-R1 Compression 3 3 3 3
Pre-Bond Moisture - WA75 CACRC-R2 Compression 3 3
OEM-R1 Compression 3 3 3 3
Contaminant 2: CACRC-R1 Compression 3 3 3 3
Pre-Bond Moisture - Drying Cycles CACRC-R2 Compression 3 3
OEM-R1 Compression 3 3 3 3
Contaminant 3: CACRC-R1 Compression 3 3 3 3
Skydrol + Water CACRC-R2 Compression 3 3
OEM-R1 Compression 3 3 3 3
Cure Cycle Deviation 1 CACRC-R1 Compression 3 3 3 3
CACRC-R2 Compression 3 3
OEM-R1 Compression 3 3 3 3
Cure Cycle Deviation 2 CACRC-R1 Compression 3 3 3 3
CACRC-R2 Compression 3 3
315
Static Fatigue
FAA Joint Advanced Materials & Structures (JAMS) Center of Excellence 8th
Annual Technical Review Meeting April 4th, 2012
18
Coupon Instrumentation
Strain gages will be installed using AE-10 adhesive per NIAR CP5401. Strain gages shall be
applied in all repair elements in eight locations as shown in Figure 24 below. Strain gages 1, 3, 5
and 7 are installed in the compression surface (repair surface) whereas strain gages 2 ,4,6 and 8
are installed in the tension surface. A deflection transducer will be used at the center of the beam
to monitor beam deflection.
1.50
1.50
0.75
1, 2
3, 4
5, 6
7, 8
5.00
Compression Side: 1, 3, 5, 7
Tension Side: 2, 4, 6, 8
7.00
Repair Scarf Edge
3.004.00
11.0012.00
FIGURE 24. CACRC PROCESS PARAMETER AND WEAK BOND EVALUATION COUPON STRAIN
GAGE LAYOUT
Mechanical Testing
Specimens will be tested at room temperature (RTA) and, -65°F (CTD) elevated temperature
(180°F) for ultimate strength and durability. Fatigue strain will be derived from the static testing
and the sandwich elements will be cycled for 165000 cycles followed by residual strength
evaluation.
The test fixture used for all testing is shown in Figure 25. The test machine shall be verified in
accordance with ASTM E4 to an accuracy of ±1% within the test loading range. Mechanical
testing will be conducted in accordance to ASTM D7249, in displacement control, at a constant
rate of 0.25 in/min with an 8” loading span and a 22” support span. The test article should be
loaded to a load equivalent to 500 microstrain and all strain gage readings should be verified
before conducting the tests.
FIGURE 25. CACRC PROCESS PARAMETER AND WEAK BOND EVALUATION TEST SET-UP
FAA Joint Advanced Materials & Structures (JAMS) Center of Excellence 8th
Annual Technical Review Meeting April 4th, 2012
19
REFERENCES
1 Tomblin, J. et. al.,"CACRC Depot Bonded Repair Investigation," FAA Joint Advanced
Materials and Structures Center of Excellence 7th annual review meeting, 2011.
2 ASTM D5229 Standard Test Method for Moisture Absorption Properties and Equilibrium
Conditioning of Polymer Matrix Composite Materials
3 ASTM E4-03 Standard Practices for Force Verification of Testing Machines
4 ASTM D7249-06 Standard Test Method for Facing Properties of Sandwich Constructions
by Long Beam Flexure