CACRC DEPOT BONDED REPAIR INVESTIGATION – ROUND … · CACRC DEPOT BONDED REPAIR INVESTIGATION...

FAA Joint Advanced Materials & Structures (JAMS) Center of Excellence 8th

Annual Technical Review Meeting April 4th, 2012

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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



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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])



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• 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



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



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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










TABLE 4. CONTAMINATION INVESTIGATION PANEL ID



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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



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FIGURE 5. CORE POTTING USING CORFILL 658

FIGURE 6. CORE POTTING, VACUUM APPLICATION

FIGURE 7. MASKING TAPE REMOVAL



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FIGURE 8. POTTED PANEL

FIGURE 9. POTTED CORE TRANSFER ONTO FACESHEET 1 (ASSEMBLY 1)

FIGURE 10. RELEASE FILM AND FAIRING BAR APPLICATION



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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.




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FIGURE 14. FINAL ASSEMBLY BAGGING

FIGURE 15. FINAL BAGGING



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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



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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



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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








105



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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



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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.



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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.



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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


Contaminant 1: CACRC-R1 Compression 3 3 3 3

Pre-Bond Moisture - WA75 CACRC-R2 Compression 3 3



Pre-Bond Moisture - Drying Cycles CACRC-R2 Compression 3 3



Skydrol + Water CACRC-R2 Compression 3 3


Cure Cycle Deviation 1 CACRC-R1 Compression 3 3 3 3

CACRC-R2 Compression 3 3


Cure Cycle Deviation 2 CACRC-R1 Compression 3 3 3 3

CACRC-R2 Compression 3 3

315

Static Fatigue



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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



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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

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