NASA Contractor Report 178272 

Time-Temperature-Stress Capabilities of Composite Materials 

for Advanced Supersonic Technology Application 

J.R. Kerr and J.F. Haskhs. 

6enetal Dynamics Corporation 

Convair Division 

San Diego, CA 92138 

Cantract NAS7 - 72308 

May 1987 

NASA 

National Aeronautics and 

Space Administration 

Langley Aetsearch Center 

Hampton. Virginia 23665-5225 
FOR€WO RD 

The work presented in this report was performed at General Dynamics Convair 

Division, San Diego, California, under Contract NASI-12308 for the National 

Aeronautics and Space Administration, Langley Research Center, Hampton, 

Virginia. Mr. Bland A. Stein and Dr. Donald R. Rder served as project 

monitors. 

Mr. James F. Haskins and Dr. James R.. Kerr served as General Dynamics program managers. Dr. Dick J. Wlkin.cl formerly with the General Dynamics Fort 

Worth Division provided the random load spectrum and assisted in the wearout 

theory analysis. Other key contributors to the program were: 

Composite Fabrication 

Specimen Preparation 

Quality Assurance 

Testing and Fixture Design 

Thermal Aging 

Flight Simulation Exposure 

and Fixture Design 

SEM and Metallographic 

Evaluation 

J.E. Hilzinger 

Chuck May 

Dr. Michael Miller 

Mike Varias 

Charlie Maikish 

N.R. Martin 

Gerald O’Barr 

Jack Christian 

Dr. Ray Adsit 

E.S. Harrison 

Curt Tanner 

Ron Torgerson 

Dr. John Waszczak 

BU Witzell 

Roscoe Anderson 

Mal Campbell 

Don Bowers 

Charlie Dunkel 

Dick Gilbert 

A.A. McClain 

Dr. Mike Featherby 

Bob Pettyjohn 

Dr. K.L. Agarwd 

- ~~XEDWG PAGE BLANK NOT FILMED 

iii 
TABLE OF CONTENTS 

Sadon . Pogo 

1 INTRODUCTION ............................................... 1-1 

2 DESCRIPTION OF PROGRAM ................................... 2-1 

3 MATERIAL SELECTION AND ACQUISITION ..................... 3-1 

Metal Matrix Composite Material ............................ 3-1 

Resin Matrix Com@te Ma- ............................ 3-2 

3.2 MATERIAL ACQUISITION ................................ 3-7 

Metal Matrix Composite Material ............................ 3-7 

Resin Matrix Composite Material ............................ 3-8 

PREPREG QUALITY ASSURANCE TESTING ..................... 4-1 

4.2 GRAPHITEEPOXI ...................................... 4-2 

. BORON POLYIMIDE . : ................................... 4-3 

4.4 GRAPHITE POLYIMIDE .................................. 4-4 

4.4.2 Celion 6000/LARC-l60 ..................................... 4-5 

LAMINATE PROCESSING AND SPECIMEN FABRICATION ........ 5-1 

3.1 MATERIALSELEX"I0N .................................. 3-1 

3.1.1 

3.1.2 

3.2.1 

3.2.2 

4 

4.1 BORONEPOXY .......................................... 4-1 

4.3 

4.4.1 HT.SI710 ................................................. 4-4 

5 

5.1 

5.1.1 

5.1.2 

5.1.3 

' 5.1.4 

5.2 

5.3 

5.3.1 

5.3.2 

5.3.3 

LAMINATE PROCESSING ................................ 5-1 

BoronlPolyimide .......................................... 5-3 

GraphitdPOiyimide ........................................ 5-5 

BORONlALUMINUM TUBE FABRICATION ................. 5-6 

SPECIMEN FABRICATION ................................ 5-8 

Panel Layouts and Specimen Identification .................... 5-8 

MallhiningOperatiom ...................................... 5-8 

BoronlEm ............................................. 5-1 

Grap hite/Epoxy ........................................... 5-1 

DoublerBonding ......................................... 5-14 

6 LAMINATE QUALITY ASSURANCE TESTING .................... 6-1 

METAL MATRIX COMPOSITE MATERIAL ................. 6-1 

RESIN MATRIX COMPOSITE MATERIALS ................. 6-7 

6.2.1 B~ronlEpoxy ............................................. 6-7 

6.2.2 GraphitdEpo~ ........................................... 6-9 

6.2.4 GraphitdPolyimide ....................................... 6-17 

6.1 

6.2 

6.2.3 BorodPolyimide ......................................... 6-17 
TABLE OF CONTENTS. Contd 

kction Page 

7 BASELINE TESTING ........................................... 7-1 

7.1 

7.1.1 

7.12 

7.2 

7.2.1 

7.2.2 

7.3 

7.3.1 

7.3.2 

7.4 

7.4.1 

7.4.2 

TENSILE TES"XNG ...................................... 7-2 

Specimen Design and 'IBst MW .......................... 7-2 

Wt Results and Discussion ................................. 7-9 

SHEAR TESTING ....................................... 7-46 

Swm D- and ?Bst pmced~ ......................... 7-48 

Tbst Results and Discussion ................................ 7-54 

FRACTURE TESTING ................................... 7-61 

SpecimenD esi~ ......................................... 7-63 

%st procedure, Results and Discussion ...................... 7-63 

COMPRESSIVE TESTING ................................ 7-81 

Specimen De andlktt procedure ......................... 7-86 

'Ikst Result3 and Discussion ................................ 7-88 

8 THERMAL AGING ............................................. 8-1 

8.1 

8.2 

8.2.1 

8.2.2 

8.23 

8.3 

8.3.1 

8.3.2 

TEST EQUIPMENT AND TEST PROCEDURE ............... 8-2 

THERMAL AGING/TENSION . TEST RESULTS AND 

DISCUSSION ............................................ 8-4 

Boron/Ep~~y and GraphitdEpoxy ............................ 8-4 

GraphiteIPoiyimide ....................................... 8-39 

BodAluminum ......................................... 8-67 

THERMAL AGINGCOMPRESSION . TEST RESULTS 

AND DISCUSSION ...................................... 8-93 

Grap hite/Epoxy .......................................... 8-93 

Celion 6000/LARC-160 GraphitdPolyimide ................... 8-99 

9 ENVIRONMENTALAGING ...................................... 9-1 

10 

9.1 

9.1.1 

RESIN MATRIX COMPOSITE SYSTEMS ................... 9-1 

Specimen Desi- and 'ht Fmcedures ......................... 9-1 

9.1.2 'ht Results and Discussion ................................. 9-3 

9.2.1 Specimen Design and Tbst Procedure ......................... 9-11 

9.2.2 %t Results and Discussion ................................ 9-13 

9.2 METAL MATRIX COMPOSITE SYSTEM ................... 9-11 

CREEP TESTING .............................................. 10-1 

10.1 SpecimenDe sign ......................................... 10-1 

10.2 %st Equipment and 'Ikst Procedure .......................... 10-2 

10.3 'ht Results and Discussion ................................ 10-4 

10.3.1 BorodEpoxy ............................................ 10-7 

10.3.2 GraphitelEpoxy ......................................... 10-12 

10.3.3 GmphiWPolyimide ...................................... 10-14 
CDC-cRAD-85-105 

TABLE OF CONTENTS. Contd 

Pug. 

10.3.4 Boron/Aluminum ........................................ 10-19 

Section 

11 BASELINE FATIGUE TESTING ................................. 11-1 

11.1 SPECIMEN DESIGN .................................... 11-1 

TEST EQUIPMENT AND TEST PROCEDURE .............. 11-2 

TEST RESULTS AND DISCUSSION ....................... 11-5 

11.3.1 Boron/Epoq and Graphite/Epo~~ ........................... 11-5 

11.3.2 G&~WPolpimide ...................................... 11-12 

11.3.3 Boron/Alumimrm ........................................ 11-16 

11.3.4 Stacking sequence^ Evaluation ............................. 11-19 

Poisson's Ratio Evaluation ................................ 11-37 

Compression Study ...................................... 11-38 

FLIGHT SIMULATION TESTING ............................... 12-1 

121 INTRODUCTION ........................................ 12-1 

DESCRIPTION OF SIMULATOR .......................... 12-3 

122.1 '1Bst Fixtures ............................................ 124 

122.2 Load Programmer and Controllers ........................... 12-4 

12.2.3 Programmer Safety Features ' ........................ 12-4 

12.2.4 Hydraulic System ......................................... 12-6 

12.2.5 HeatingSy~tem .......................................... 12-6 

11.2 

11.3 

11.3.5 

11.3.6 Baseline Fatigue Data for the Thermal Aging/ 

12 

122 

....... 

1226 C00lingSystem .......................................... 12-8 

12.27 ControiCo nsole .......................................... 12-8 

12.2.8 DataRecordingSystem .................................. 12-10 

12.2.9 MonitorlSafety System ................................... 12-10 

12.210Checkout ............................................... 12-11 

12.2.11 Automatic Operation ..................................... 12-11 

12.3 SPECIMEN CONFIGURATION .......................... 12-12 

RANDOM LOAD SPECTRUM ............................ 12-15 

Spectrum Generation for the Long%rm %sts ................ 12-17 

Spectrum for the Short-h Tzsts .......................... 12-18 

WEAROUT MODEL .................................... 12-20 

SHORT-TERM FLIGHT SIMULATION .................... 12-27 

'&t Plan and Ptocedures ................................. 12-28 

Sh0rt-h %St Results .................................. 12-36 

12.6.3 Short-%= Analysis ..................................... 12-57 

LONG-TERM FLIGHT SIMULATION ..................... 12-67 

%st Plan and Procedures ................................. 12-67 

Phase I L0ngTbx-m %st Results ............................ 12-70 

Phase I1 Longh Qst Results .......................... 12-128 

12.4 

12.4.1 

12.4.2 

12.5 . 12.6 

12.6.1 

12.6.2 

12.7 

12.7.1 

12.7.2 

12.7.3 

13 THERMAL EXPOSURE TEST PROGRAM SUMMARY ............. 13-1 

vii 
Goc-cRAD-85-105 

TABLE OF CONTENTS, Contd 

Sodion Pagm 

14 CONCLUSIONS ................................................ 14-1 

15 REFERENCES ................................................ 15-1 

APPENDIX 

A AN ANALYSIS OF THE B/PlOBAC BORON POLYIMIDE SYSTEM 

B QUALITY ASSURANCE ?'EST PROCEDURES 

viii 
TlME-TEMPERATURE-STRESS CAPABILITIES 

OF COMPOSITE MATERlALS FOR ADVANCED 

SUPERSONIC TECHNOLOGY APPLICATION 

J.R. Kerr and J.F. Haskinr 

Convair Division of General Dynamics 

SUMMARY 

High temperature composite materials have the potential of significantly reducing the weight of future supersonic aircraft structures. However, implementation of the material into vehicle usage is contingent upon demonstrating its airwortlliness when exposed to the long-time supersonic flight service environment. The primary objective of this research was to establish the timetemperature-stress characteristics and capabilities of five classes of high temperature 

composite materials to determine their suitability for advanced supersonic technology consideration. The systems studied were BIE, GIE, BIPI, GIPI, and BlAL 

The tests that were conducted on the composite systems were baseline tensile, 

notched tensile, compressive, shear, and fracture; environmental and thermal 

aging; constant amplitude fatigue, creep and creep rupture; and both accelerated 

and real-time flight simulation testing using totally rzindom fatigue loading. 

Phase I of the two-phase program included the baseline, fatigue, and creep determinations and all the aging and flight simulation studies up through 10,000 

hours of cumulative exposure. Phase I1 continued this effort up to 50,000 

cumulative hours of simulated flight service. The baseline, constant amplitude 

fatigue, and creep tests characterized the composite materials and formed a 

basis for comparison with the post-exposure test results. The baseline tests were 

also used to set the test conditions for the short-term exposures (100 and 200 

hours, random fatigue at temperature). The short-term service history tests 

were performed to generate wearout analysis model data, and to set the stress 

levels for the long-term service history tests. Finally, the long-term tests were to 

provide data on the effect of 10,000-, 25,000-, and 50,000-hour simulated supersonic flight service on residual properties of the composites. The thermal aging 

and the environmental aging exposures were concurrent with the service history 

tests. The considerable quantity of data that have been obtained from the 

baseline tests, the 50,000 hour thermal and environmental aging studies, and 

the flight simulation exposures of almost 40,000 cumulative hours are presented 

in this report. 

ix 
SECTION 1 

INTROOUCTlON 

Approximately one year after the cancellation of the United States Supersonic Transport Pro= 

gram, the National Aeronautics and Space Administration initiated the Advanced Supersonic 

Technology (AST) Program. The name of the AST Program, which started in 1972, was subs+ 

quently changed to Supersonic Cruise Aircraft Research (SCAR), and then shortened to Supersonic Cruise Research (SCR). The overall objectives of this program were: 

a. To provide an expanded technology base for future civil and military supersonic aircraft. 

b. To provide the data needed to assess the environmental and economic impacts on the 

United States of present, and in particular, future foreign supersonic transport aircraft. 

c. To provide a good technical basis for any future consideration that may be given by the 

Unites States to the development of an environmentally acceptable and economically 

viable commercial supersonic transport. 

A block diagram showing the structure of the program is given in Figure 1-1. An extensive 

review of the activities associated with the SCR Program from 1972 to 1983 is presented in the 

bibliographies found in NASA Reference Publications 1003, 1063, and 1117 (ref. 1 to 3). The 

scope of the program included the development base, system studies, and the following 

disciplines: 

SCR PROPULSION 

SCR STRATOSPHERID EMISSIONS IMPACT 

SCR STRUCTURES AND MATERIALS 

SCR AERODYNAMIC PERFORMANCE 

SCR STABILITY AND CONTROL 

The work being reported on here was done under the Structures and Materials discipline of the 

SCR program. One of the primary goals in the Structures and Materials area was to advance 

the technology and establish a data base for advanced composites so that sound technical decisions may be made in the future regarding the use of these materials in supersonic cruise aircraft structures. Primary emphasis was placed on composites because of their high potential 

for reducing both weight and cost of future supersonic cruise aircraft structures. Information 

on the effects of long-time cyciic exposure to the environments and loads representative of 

longtime supersonic cruise aircraft service for composite materials of interest was not available in 1972, and, as a result, this program to generate such information was initiated in 1973. 

A range of baseline properties was determined for representatives of five composite materials 

classes: borodepoxy, graphitelepoxy, borodpolyimide, graphitelpolyimide. and borodaluminum. Long term exposures out to 10,000: 25,000; and 50,000 hours were attempted in static 

thermal environments (thermal aging) and in ones that simultaneously combined programmed 

thermal histories and mechanical load histories (flight simulation). Material behavior during 

these exposures and the post-exposure residual property tests provided exposure effects and 

assisted in understanding material degradation mechanisms. 

1-1 
Id 

I 

I 

1 I 
Because of the goal of 50,000 hours of real-time exposure for both the thermal aging and the 

flight simulation tests the program waa, of necessity, a very long om. The commercially pm 

duced composite materials selectsd at the beginning of the contract were thought to be ones 

that would be readily available for many years. In fact, one of the criteria for selecting the corn 

posites was that each material shall be reasonably well established, Le., currently available and 

with a finn manufacturing maturity. As the program progressed it became apparent that, in 

general, the composites did not possess these requirements. Material problems that were 

encountered included: low tensile properti- and lack of elevated temperature stability for one 

of the systems, sigmficant changes in mechanical properties for a second material when repur 

chased several years after the initial purchase, change in designation with some mechanical 

property changes for a tbird material, and unavailability of a fourth material seved years into 

the program. This made it very difficult, when adding tests later in the study, to correlate 

results from two batches of supposedIy identical materiai. 

Another problem, experienced at the begking of the program, was overly optimistic estimates 

of the maximum use temperatures (for extended service) of each of the composite systems. For 

graphitelpolyimide and borodduminum pre- thermal aging data assisted in the selection of reaiistic flight simulation temperature profiles, which were considerably lower than 

those currently being proposed by the composite community. For the two epoxy classes, 

however, selection of excessive temperatures for the flight simulation tests resulted in numerous specimen failures and considerable overall property degradation prior to the first 10,000 

hours of exposure. As discussed later, these tests were restarted with replacement specimens 

using a considerably modified thermal profile. 

In spite of.these problems, thermal aging exposures of up to 50,000 hours and flight simulation 

exposures of almost 40,000 hours were accomplished. 

1-3 
SECTION 2 

DESCRlPTlON OF PROGRAM 

The general objective of this study was to assess the suitability of advanced filamentary reinforced composite materiais for future supersonic cruise aircraft structures. The study had two 

phases. The first included all material property determinations and aging simulation exposures 

up through 10,000 hours. The second continued these tests toward a goal of 50,000 hours. 

Representative materials from each of the five composite classes listed in Figure 2-1 were 

selected for evaluation in Phase I. These were: 

a. Rigidite 5505/4 borodepoxy. 

b. A-S/3501 graphitelepoxy. 

C. 5.6-d borod6061 borodaluminum. 

d. HT-SI710 graphitdpolyimide. 

e. 4.0-mil BIPIOSAC borodpolyimide. 

Early in Phase I the borodpolyimide was removed from test because of low tensile properties 

and extensive thermal degradation during elevated temperature exposure. During Phase I1 a 

second graphitelpolyimide, Celion 60001LARC-160, was added to the program for evaluation 

up through 10,000 hours. 

MATE RIALS 

BEING 

ASSESSED 

B/E 

GIE 

B/AI 

G/Pl 

B/PI THERMAL & ENVIRONMENTAL AGING 

ANALYSIS - LOAD GENERATION 

- CREEP - 

PHASE I 10,000 HA 

POST- 

+ EXPOSURE 

50.000 HR TESTING 

MOISTURE PHASE I I i\\\\\\\\\\

A 

- 6532132 

Figure 2- 1 Time-Ternperaturo-Stross Capabilities of Cornposit. Materials - Program Outlino 

2.1 
Figure 2-1 is a schematic diagram that shows the principal elements of this work. The baseline 

properties are typical mechanical properties for each of the composites, and these properties 

were generated in sufficient detail so that any changes in them due to repeated loading and 

thermal or environmental aging could be measured and reported. Fatigue life was measured 

before and after various load and temperature histories. The materials’ creep behavior was 

measured early in the program in order to fully evaluate filamentary composites of this type. 

The aging characteristics of the materials were measured at various temperatures and pressures. A limited study of moisture exposure effects was included for the resin matrix materials: 

However, with this exception, moisture effects have been deliberately avoided by use of standard bakeout procedures prior to the various tests conducted during the program. Inclusion of 

moisture effects to each of the test areas was beyond the scope of the overall program. 

The changes in baseline tensile and fatigue, notched tensile and fatigue, compressive, and 

shear properties that ocmrred during times out to 50,000 hours were measured for environmental and thermal aging conditions, as well as, random cyclic loading with cyclic temperature 

variations. 

These latter tests were intended to simulate conditions experienced in a lower wing surface 

during supersonic flight and, hence, consisted almost entirely of tensile loads. For the same 

reason the majority of the residual property specimens were intended to be tested in tension. 

As will be discussed later, the failure modes and degradation mechanisms for the various exposum tests were found to be primarily matrix dominated. Because of these observations, residual compressive and shear tests were included with the tensile and fatigue tests already scheduled to increase the usefulness of the data 

One other change to the original program occurred as a result of the matrix dominated failure 

modes observed during Phase I. A study to determine the effects of thermal aging on the compressive properties of graphitelepoxy and graphitdpolyimide was added during Phase I1 to 

supplement the thermal agingltensile program conducted on all of the composites in Phase I. 

In order to provide a device capable of simulating supersonic flight conditions of load and 

temperature, a machine was built that was capable of not only random loading and temperature variation but could test a large number of specimens at the same time. It was desirable 

that the apparatus be able to accommodate a large number of test specimens because of the 

expected data scatter associated with advanced composites. The machine was also capable of 

operating for long periods of time with little maintenance and monitoring and was designed in 

such a manner as to possess adequate safeguards against either overloading or overheating 

the specimens. 

Several assumptions were made to define the loading spectrum. The mission profile and loads 

were obtained from reference 4. These data were also used for the study described in reference 5. 

The life of a supersonic airplane was assumed to consist of 25,000 flights of 2 hours duration 

each For this work, each flight was assumed to be an “operational flight” as defined in reference 4. Each flight consisted of a 10-minute climb, a 90-minute cruise, a 20-minute descent, 

and a singie landing load. The number of loads per flight and the number of loads in each of the 

segments were modeled after the reference 1 load sequence. However, the loads were randomized such that all the loads expected to occur in 25,000 flights were included. Since the 90-minute cruise was scheduled to include only about eight loads per flight, the mean load was held 

for 44 minutes before and after the imposition of a 2-minute cruise loading period. The climb 

and descent loads were distributed over the entire climb and descent time. 

2-2 
The job of setting load levels for each of the composites materials for the random load spectrum so that some but not all specimens would fail during the 50,000-hour program was indeed 

a difficult task. This had to be done by running accelerated testa in which many loads could be 

applied at elevated temperature in a fairly short time. We chose to apply a onehalf lifetime of 

loads in 200 hours, which worked out to be about one load a second. The real-time flight spectrum is more like 20 loads per hour. The accelerated tests were run for both 100 and 200 hours 

at load levels such that we would have 80% survival if the tests were run to 400 hours. 

The residual strength was measured after specimens had been subjected to onefourth and one 

half lifetime of loads at elevated temperature. These strengths were then used to calculate, using the Hdpin wearout model (ref. 6). the load levels for real-time testing, again using 80% survival at the end of 50,000 hours. 

The calculated mean and peak load levels set by the 80% survival for the 50,000 hours or one 

lifetime were used to set the real-time load levels for the flight simulator. These differed for 

each composite system and aIso differed for the notched and unnotched specimens of the same 

composite. Thus each composite system was to be subjected to 50,000 hours of real-time flight 

simulation testing using random loading and temperature described by the climb, cruise, and 

descent portions of the flight. 

Following completion of the 10,000-hour Phase I portion of flight simulation exposure, half of 

the specimens were removed from test. An extensive examination of the specimens was conducted to determine the extent of damage and, for those specimens that had failed, to identify 

the degradation mechanisms responsible for these early failures. The specimens that survived 

the 10,000 hours of flight simulation exposure were used to determine residual tensile and 

fatigue, notched tensile and fatigue, compressive, and shear properties. At the end of Phase I, 

two modifications were made to the test plan. Because of the high failure rate and extensive 

visual degradation of the epoxy specimens, d the remaining epoxy specimens were removed 

and new specimens were fabricated to replace them. A significantly reduced cruise temperature (calculated from thermal aging data) was used for subsequent flight simulation testing of 

these replacement specimens. 

The second modification was the addition of three more flight simulation specimens of each 

material to be exposed for 25,000 hours. The specimens remnining in the simulator after 

Phase I were intended to complete 25,000 or 50,000 hours of exposure in Phase I1 and provide 

residual strength data similar to that obtained in Phase I. As will be shown, this goal of 50,000 

hours was not realized. However, considerable residual strength data was obtained from 

specimens that survived exposures up to 36,000 hours, as well as, from the undamaged portions of specimens that failed after times up through 40,000 hours. The results of these longterm fight simulation tests, both Phase I and Phase 11, were then compared with the wearout 

model predictions based on accelerated flight simulation exposure. 

2-3 
SECT10N 3 

MATERIAL SELECTION AND ACQUlSiTlON 

The “Structural Design Guide for Advanced Composites Applications” (ref. 7) has limited the 

term “advanced composites” to those systems characterized by high-strength, high-modulus 

fibers. For this study, advanced composites are those systems whose reinforcements have a 

filament tensile strength greater than 2100 MNIm2 (300 ksi) and a filament modulus greater 

than 210 GNIm2 (30 million psi). 

At the time this contract was initiated, a multitude of advanced composite material systems 

had been developed and evaluated to various degrees. In general, five classes of material systems had been used by numerous workers in this fiekk borodepoxy (BIE), graphitelepoxy (G/E), 

borodpolyimide (BIPI), graphite/polyimide (GIPI), and bomdaluminum (BIAl). Within each of 

these five classes there were several types of matrix materials and different types of fibers. 

The following sections discuss each of the five classes of materials and Convair’s choices as to 

the specific advanced composite systems that were evaluated. It should be pointed out that 

this program was started in the early 1970s. and selection of all but one of the composite mate 

rials was based on the best available systems at that time. The exception was the second GIPI, 

Celion 60001LARC-160, that was added after the completion of Phase I in 1980. 

3.1 MATERIAL SELECTION 

3.1.1 METAL MATRlX COMPOSITE MATERIAL. Table’3-1 shows the large variety of filaments and 

matrix materials that were available when the program was initiated. Table 3-2 lists a number 

of the most promising and widely used processing methods for the primary fabrication of 

sheet, plate, and structural shapes of metal-matrix composite materials. 

tabk 3-1. Mdol-Mutrix Composito Raw Matorials 

Fiborr Mahix 

Boron (4 mil) 

Boron (5.6 mil) 

Borsic (4.2 miU 

Borsic (5.7 mil) 

Nitridecoated Boron 

Silicon Carbide 

Stainless Steel 

Graphite 

Aluminum 

1100 

2024 

2219 

606 1 

7002 

7075 

713 

718 

KO 1 

Magnesium 

Titanium 

Nickel 

7178 

3- 1 
Tablo 3-2. Primary Procasing Mothods 

Processing M&od Typical Forms Producod 

Diffusion Bonding 

Braze Bonding 

Mono- or multilayer tape. sheet, plate, 

structural shapes 

Mono- or multilayer tape, sheet, plate, 

shapes 

Eutectic Bonding Sheek plate, structural shapes 

Plasma spraying Monolayer tape 

Molten-metal Infiltration 

(Casting) 

Mono- or multilayer tape, sheet, plate, 

structural shapes 

It has been found that no one metal-matrix composite material is superior for all possible applications. However, data based on filament-matrix compatibility, comprehensive physical and 

mechanical property tests, primary and secondary fabrication studies, raw material and processing costs, and extensive experience in working with composite materials indicated that 

diffusion-bonded BlAL in particular the large diameter (5.6-mil) boron with 6061 aluminum 

matrix, offered the most promise for widespread structural applications in various aircraft 

systems. BlAl composite material had also been found to be highiy resistant to various environmental exposures, with excellent retention of properties. ab cryogenic and elevated 

temperatures, and withlittle or no effect of space exposures on mechanical or physical pioperties. Another attribute of BlAL and metal-matrix composite material in general, was its very 

high transverse modulus, about 140 GN/m2 (20 million psi) for B/AL This enables the use of 

unidirectional layups in a number of structural applications that would not be possible with 

organic-matrix materials with their low transverse modulii. Because of these considerations 

(ref. 8 to 14) diffusion bonded, 5.6 mii borod6061 aluminum was chosen for evaluation on this 

program. 

Singlestep hot pressing was the technique selected for fabricating the BlAl sheet material 

because of the ability to produce large panel sizes in a single pressing operation, and because of 

lower processing costs. In this method, boron filaments are wound over a mandrel onto thingage sheet material. These layers are then cut and stacked to the desired thickness and the 

material consolidated in a hot press. Because of more economical production of large sheet 

material in pressing operations, this method was chosen over the isostatic pressing technique 

that required consolidation of prefabricated laminates (tape material). At the beginning of the 

program two qualified suppliers were available. They had been using the singlestep hot pressing technique to supply highquality metal-matrix composites on a production basis for 

numerous hardware programs at Convair and other laboratories. 

3.1.2 RESIN MATRIX COMPOSITE MATERIAL 

3.1.2.1 BorodEpoxy. At the time that this program started B/E was the most highly 

chcterized of the advanced composite materials. As a result of approximately 7 years of 

evaluation, the industry had generally settled on two resin systems (Avco 5505 and 3M 
SP-272). 3M also provided the same resin on uncoated boron (SP-292) and maintained the same 

basic properties. Other 3M systems such as SP-290 had some better high-temperature properties than the SP-292 but were not as thoroughly characterized. General Dynamics had 

generated extensive design-allowable data on the Avco 5505 under Contracts 

F33615-68-C-1474 and AF33(615)-5257 (ref. 15 and 16). Convair preferred the Avco 5505 

system because of its somewhat higher temperature resistance and its wide application at that 

time. 

Most early work, conducted by Convair and elsewhere, used the 4.0-mil boron filaments. Some 

work had been conducted using 5.6- and 6.0-milaeter boron filaments, but selection of the 

4.0-mil filament was made to avoid an extensive program necessary to characterize the larger 

diarneter material. Considerable engineering data already existed on the coated 4.0-mil 

diameter boron including static and fatigue characteristics for many configurations with and 

without imperfections. IITRI (ref. 17) generated extensive data on the effects of environmental exposure on the mechanical and physical properties of the 4.0-mil Avco 5505. Since the purpose of this program was to evaluate long-term effects, selection of a fully characterized 

system was desirable. In addition, the long-term effects were thought to be primarily matrix 

dependent and independent of filament diameter. 

3.1.2.2 GmphitdEpoxy. Convair had worked on high-strength, high-modulus graphite fibers as 

a reinforcement for epoxy and polyimide resina since 1965. In this period there had been a prcl 

liferation of graphite fibers. EarIy fibers based on visose rayon, such as Thornel 25, Thornel 40, 

and HITCO HMG30, had been surpassed and were no longer seriously considered for structural aircraft applications. Newer fibers, primarily based on polyarcrylonitrile precursors, had 

become the serious contenders for present and future aircraft applications. These fibers could 

be generally divided into four categories: 

a. Low moduiudultra-high strength. 

b. Medium moduludhigh strength 

c. High moduludmedium strength 

d. Ultra-high moduludmedium strength 

The first category of low modulus, 207 to 221 GNlm2 (30 to 32 million psi), coupled with ultrahigh strength, greater than 2760 MN/m2 (400 hi), was finding most usage in current aircraft 

applications because of its high strength and very low cost. Fibers that fell in this category 

were Hercules Type A-S and Stackpole Panex 301A. Others having the same modulus but 

somewhat lower strength were Thornel 300, Fortafil30-T, and Morganite 111. Type A-S had 

had the most usage at that time. General Dynamics has used this type of fiber extensively in 

the F-111 pivot fitting fairing program and the F-5 fuselage program. 

The second category of fibers, those having medium modulus, 262 to 276 GNlm2 (38 to 40 

million psi), and high strength, greater than 2410 MNlm2 (350 ksi), included such products as 

HT-S, Modmor Type 11, and Fortafil4-T and had been used extensively in the preceding years. 

They were inherently costlier than the 207 GNlm2 (30 million psi) fibers and were expected to 

have lower utilization in the future than they had in the past. 

The third category of fibers, those having high modulus, 345 to 414 GN/m2 (50 to 60 million 

psi), and medium strength, 2070 to 2760 MNlm2 (300 to 400 ksi), included such products as 

3-3 
HM-S, Modmor Type I, and Fortafil6-T. These fibers had not been used as extensively a3 the 

first two categories because of higher cost, lower strength, and poorer resin compatibility 

(lower shear strength in composites). 

The fourth category of fibers, those having ultra-high” modulus, greater than 483 GNlm2 (70 

million psi), and medium strength, 1520 to 2240 MNlm2 (220 to 325 ksi), included GY-70 and 

Thornel 75 materials. They had the same general problems as the third category described 

above. 

Based on the above review of available graphite fibers, Convair selected the epoxy resinimpregnated Type A-S fiber. 

In the area of high-temperature 450 K (350 O F) epoxy resins, there had also been a profileration 

of candidate systems. Up to 1971, the prime resins evaluated and used for structural aerospace 

applications were Fiberite X-904, Hercules 3002, Whittaker 5205 and 5206. The 3002 was 

extensively evaluated for design properties by IITRI (ref. 17) and North American Rockwell 

(ref. 18). However, the 3002 was adversely affected by moisture, was costly, and was difficult 

to process into prepreg the availability of this resin for future programs was in doubt at this 

time period. 

In 1972 and 1973 a number of new high-temperature epoxy resins had been introduced, such as 

Ferro E-350 and E-450; Fiberite X-911, X-915, and X-934; Whittaker 5208: and Hercules 3501. 

Of these, the E-350, X-934,5208, and 3501 were seeing the most use in evaluations and in pro 

duction of aircraft parts. 

Convair conducted an extensive screening of in-house data, vendor information, and outside 

literature relating to the newer high-temperature epoxy resins. The leading candidates were 

Ferro E-350 and Hercules 3501 with the E-350 appearing to be slightly better. However, at 

that time Ferro Corporation had recently moved its plant, lost some key personnel, and was 

leaving the graphite prepreg business. In addition to the favorable rating in the screening 

evaluation, the A413501 system was currently being used in large commercial applications, 

was being evaluated extensively for missile applications, and was believed to be the system 

that would be the standard G/E composite for aircraft applications. Based on these considerations and an excellent cost picture, the A-S/3501 system was selected for the program. 

3.1.2.3 BorodPolyirnide. Early studies of BlPI composites were conducted with the P13N, 

PlOP, and 703 polyimide resin systems (ref. 19 to 23). Initially, attention was paid strictly to 

the fabrication of low-void laminates. As such, the prepegs were tack-free, highly imidized 

materials that were acceptable for making flat laminates but unacceptable for large parts. 

Later work led to modifications in the P13N and PlOP resin systems to produce P105A, which 

wag the first step toward developing a tacky prepreg that would yield low-void polyimide 

laminates. At about this time Convair under contracts NAS 8-26198 “Development of Boron 

and Graphite Poiyimide Design Data,” and NAS 8-28201, “Polyimide Compression Panel,” 

was investigating the development and use of BlPI advanced composites. Some of the candidate resin systems evaluated were: 700, 703, 709, 710, P13N, PlOP, P105A, Gemon L, 4707, 

and BPI-373. Two of these, P105A and 703, were selected for additional processing studies, 

and some preliminary test data were developed for each system using both the 5.6- and 4.0-d 

diameter boron fiber. Test results showed that 1) the data for the BIP105A composite system 

was significantly better than the B/703 composite system, and 2) the 4.0-mil boron fiber was 

3-4 
superior to the 5.6-mil boron fiber in that resin contents and laminate shear properties were 

higher (ref. 24). For these reasons the 4.0-d boron fiber and the P105A polyimide resin 

system were selected for the program. At about the time that the B/PI composite was being 

selected, Ciba-Geigy, the manufacturer of the resin, replaced the P105A with a newer version 

designated PlO5AC. It was this polyimide that was evaluated on the test program. 

3.1.2.4 Gmphite/Polyimido. Unlike the BlPI composite systems, a fully developed and 

characterized G/PI system was available at the time that the program was initiated. The GlPI 

system (HT-S/710) was developed by Convair under Contract NAS 8-26198 (ref. 24). Substantial quantities of mechanical property, thermal aging (short time), moisture resistance, and 

creep data were available on this system (ref. 25). At the same time, under Contract NAS 

8-28201 (ref. 261, Convair was determhing baseline design properties for A-S/710 GiPI composite systems. These data were used to design, fabricate, and analyze a large compression 

panel for short-term applications at 589 K (600" F). Based on the results of these two contracts, Convair selected the HT-S/710 system. 

The HT-S graphite fiber was selected because of its high strength and lower weight loss at 589 

K (600 O F) than the Type A-S graphite fiber. Data developed and reported by Morgan Hanson 

of the NASA Lewis Research Center at the National SAMPE Conference in Huntsville, 

Alabama in 1971 indicated that the Type A-S fiber oxidized severely at 589 K (600 O F). After 

1000 hours, the uncoated A-S fiber had weight losses approaching 100 % (ash). What happens 

in a composite was unknown. However, for this program Convair believed the extra $5 to $10 

per pound for the HT-S fiber was justified since it had higher oxidative stability than the Type 

A-S fiber. The A-S fiber appeared to be fully acceptable for use at 450 K (350" F) with epoxy 

resin systems but not for use at 589 K (600" F) with polyimide systems. 

The 710 resin system was selected because it was believed to be fully characteriied, compatible 

with graphite fibers, commercially available as prepreg from three material suppliers, and 

showed higher temperature stability than any of the other then available polyimide resins. 

One of the goals of the oved program was to follow the progress of advanced composite systems development and to add new materials if sufficiently promising ones appeared. Prior to 

the completion of Phase I, but several years into the program, several high temperature 

graphite fiber reinforced polyimide systems were beginning to become commercially available, 

and a decision was made to add one of the these to the program. To select the best of these 

emerging systems a screening test was planned. The selection criteria were: 

a. Availability and cost of prepreg material 

b. Ability to fabricate both full-size 12-ply panels and a simple structural shape. 

c. Moisture resistance. 

d. High temperature thermal stability and oxidation resistance. 

Three resin systems with the HT-S graphite fiber were to be evaluated. The HT-S fiber was 

chosen because it was the reinforcement used in the system already in the program (HT-S/710). 

The potential candidates at that time are listed in Table 3-3. Of those only the NR-150B2, 

PMR-15, and HR-600 were considered likely candidates, and only the first two were available 

from a commercial prepregger. The remaining resins were eliminated because of insufficient 
Tabla 3-3. bndidotms for a Sc-ing Evaluation to Selod an Additional Gaphito Polyimide Systom 

Maidol Systom Pofmniial Vendors 

HT-S/NR-l50B 2 

HT-S/PMR-15 US. Polymeric Fern 

HT-SIHR-600 Fiberite 

HT-SI5230 Narmco 

€3"-SIF-178 Hexcei 

ma4397 Hercules 

Narmco, Fiberite, US. Polymeric 

high temperature thermal stability and oxidation resistance. A screening test was conducted 

on HT-SINR-15OB2 and HT-SIPMR-15 with the current system, HT-S/710, included as a base 

line comparison. A major portion of the screening evaluation was devoted to thermal aging 

exposures. 

Drastic decreases in weight and flexural strength were observed for ail three systems. The 

poor results were traced to a lack of elevated temperature stability, with very high weight loss, 

of the HT-S graphite fiber (ref. 27). Modifications in the fiber processing conditions at the pre 

ducer during this time period (several years after the purchase of the initial HT-Sl710) 

appeared to have been responsible for the changes. Because of the unacceptable thermal aging 

results, testing of the three systems was discontinued. Since General Dynamics did not believe 

that HT-S fiber with good elevated temperature stability could be obtained at that time, it was 

necessary to select another graphite fiber. The most promising choices were Modmor 11, 

Celion, and HT-$2, a low-sodium modification of HT-S. At about this time Morganite Ltd. 

discontinued the manufacture of Modmor I1 graphite fiber. For the two remaining fibers. good 

results during high temperature aging studies (on both bare fibers and panels) had been 

reported for both. Availability, current and future, as well as a somewhat larger data base 

favored Celion fiber at that time, and it was selected. 

For the renewed screening test program, batches of Celion 6000l710, Celion 6000lNR-150B2, 

and Celion 6000/PMR-15 were obtained. They were ordered with an epoxy sizing on the Celion 

6000 graphite fiier. The sizing was necessary to facilitate handling of the graphite during the 

prepregging operation. An epoxy rather than a polyimide sizing was selected because at the 

time of placing the order it was the only material available. In addition, previous studies at 

Convair had shown that CeiiodNR-15OB2 GlPI panels fabricated from both epoxy sized and 

polyimide sized fiber were essentially identical with regard to fabricability and short time 

mechanical properties. The small amount of sizing (approximately 0.2 to 0.4%) is uncured and 

dissolves into the matrix during fabrication. Even though this small amount of epoxy resin is 

probably not detectable after cure and post cure of a panel, the question arose as to its effect 

after long time thermal exposure (Le., 50,000 hours at 450 OF). Because no data were available 

on long-time effects, the decision was made that for the SCR program only polyimide sized 

graphite would be used. Consequently, revised purchase orders were prepared for prepreg with 

polyimide sized graphite. 

The screening evaluation was becoming a much longer program than ori@y planned. Several important events occurred during this time period (1980-1981). First, the 710 resin that 

3-6 
had been removed from production by Monsanto somewhat earlier was now no longer available 

as all the stack had been depleted. Second, the NR-150B2 resin producer, DuPont, terminated 

sales because current and projected profitability did not justify continued production. Third, a 

very promising polyimide resin, LARC-160, developed at NASA Langely Research Center was 

becoming commercially available. 

Because of the lengthy delays that had occurred it became evident that if a new material was 

to be added to the program it must be done very soon. Further delays could prevent comple 

tion of the thermal aging and flight simulation exposures of this material by the time that the 

Phase I1 tests on the original materials were finished. 

Consequently, the planned screening tests were canceled, and selection of the new material 

was made based on current data and prepreg availability. Because of availability, the choice 

was limited to Celion 60001PMR-15 and Celion 60001LARC-160. The LARC-160 system was 

selected because of superior fabricability. At that time the PMR-15 system required a press 

cure for panel fabrication. In addition, the Celion 60001PMR-15 prepreg tended to be quite 

body, and fabricated panel3 had shown a greater tendency to microcrack. 

3.2 MATERIAL ACQUISITION 

Because of changes in the scope of the Phase I1 test program, the addition of a new material, 

and the added thermal agingkompression study, it was necessary to purchase material at 

three times during the perfommnce of the contract. The initial order was made in 1973 for the 

five composite systems selected at the beginning of Phase I. The second order was made in 

1976 for BE, GIE, HT-S/710 GIPI, and BlAl systems to be used in the new material screening 

evaluation. The final order was placed in 1980 for Celion 60001LARC-160 GiPI and GIE to 

satisfy the requirements of the new material addition and the thermal aginglcompression 

evaluation. 

The following sections describe the procedures involved in purchasing the advanced composite 

systems selected for evaluation on this program. For three of the materials, BIE, BIAL and 

Celion 6000/LARC-160 GIPI, little or no difficulties were encountered in procurement or during quality assurance testing. For three of the materials, however, considerable problems 

occurred either in acquiring the prepregs or during acceptance testing. 

3.2.1 METAL-MATRIX COMPOSITE MATERIAL. The 5.6-mil borod6O6l aluminUm composite, 

referred to as BIa was obtained from AVCO Corporation, Systems Division, Lowell, Massachusetts. The materid, ordered to General Dynamics Specification 0-00854, was received in 

the form of finished sheets, 0.3 m by 0.6 m (12 in. by 24 in) in dimensions. The requirements for 

the program were: 

Phase I Phase I1 

6 ply, [Oo f 45OIs layup - 11 sheets 6 ply, [Oo A 45'1, lapp - 1 sheet 

6 ply, [OoIs layup - 2 sheets 

3-7 
Delivery of the B/Al was on schedule, and. with the exception of one 6-ply unidirectional panel 

ail were of acceptable quality. The one panei showed indications of poor bonding during 

ultrasonic Gscan testing and was returned to AVCO for replacement. 

3.2.2 RESIN MAtwX COMPOSITE MATERIAL 

3.2.2.1 BorodEpoxy. The Rigidite 5605/4 borodepoxy composite, referred to as BIE, was also 

obtained from AVCO Corporation, Systems Division, Lowell, Massachusetts. The material 

was ordered to meet the requiremenb of General Dynamics Specification FMS-2001B. A total 

of 12.7 kg (28 Ib) in the form of 0.08 m (3 in.) wide prepreg tape was required for Phase I. An 

additional 1.8 kg (4 Ib) was required for Phase I1 of the program. The material was delivered on 

schedule and successfully passed all the quality assurance requirements for both the prepreg 

and fabricated laminates. 

3.2.2.2 GraphitdEpoxy. The A-S/3501-5 graphite/epoxy, referred to as G/E, was obtained from 

Hercules Incorporated, Bacchus Works, Magna, Utah. Three batches were required to meet 

the needs of the program. At the beginning of Phase I 17 kg (37 Ib) of 0.08 m (3 h) wide 

prepreg tape was purchased to the following Hercules Specifications: fiber properties per HDSG-2-6001; prepreg properties per HD-SG-24006A except flow to be mn at 436 K (325 O F), 100 

psi, 15 minutes: composite properties per HD-SG-6002A. This batch of material was received 

on schedule and checked out perfectly in the prepreg quality assurance tests. The flexural and 

short beam shear properties determined from the quality assurance panels were also satisfactory. Based on these quality assurance results, 12 full size panels were fabricated for the test 

program. The baseline tensile data, however, were widely scattered and significantly lower 

than nod The transverse tensile strengths were particularly bad, with an average value of 

22.8 MNlmz (3.3 ksi) at room temperature. To check that the panels were unacceptable, resin 

and fiber content determinations and three tensile tests were made on each of the 6-ply 

unidirectional and crossply panels. The results are given in Table 3-4. The wide scatter and 

generally low values are clearly evident. Expected values of F, are about 1400 MNIm2 (200 

ksi) and 55 MNlm2 (8 ksi) for the longitudinal and transverse directions of the unidirectional 

layups and about 550 MNIm2 (80 ksi) for the crossply layups. The resin and fiber contents 

shown in Table 3-4 are all acceptable, although Panel No. 10 has a somewhat low fiber content. 

A check with the vendor (Hercules) revealed that the material was among the first to come off 

the line from a new prepregging process. Although the resin content was acceptable (determined during quality assurance testing), the resin distribution through the material was poor 

because of a lack of penetration into the graphite tows. Similar results were reported by others 

from material obtained during the same time period. Hercules corrected the production problem and replaced the material at no charge, but a considerable time delay was incurred because 

of the rejected material. 

Visual examination of the replacement batch of GE prepreg revealed none of the resin poor 

areas that were typical of the previously rejected batch. The material successfully passed all of 

the quality assurance requirements for both the prepreg and fabricated laminates. As an added 

check on the acceptability of the panels, tensile coupons were prepared from unused portions 

of the unidirectional quality assurance panel and the first full size crossply paneL The values 

obtained, 1655 MNlm2 (240 ksi) and 579 MNlmz (84 ksi), were typical of good GIE material. 

Transverse tensile tests on the unidirectional panel gave an average ultimate strength of 59 

MNIrnZ (8.6 ksi). This value is again typical of good G/E material 

3-8 
Tablo 3-4. Quality Asrumnco Data for WE Panels from Rojd Batch of A-S/3501-5 Proprog 

Rosin Flbu 

Pond Content Content Tonsilo Strongth 

No. Orionfation Plios (wt %) (vol. %) MN/& (W 

10 

11 10 0i6 

6 36.2 55.5 1310 

752 

190 

109 

696 101 

6 32.6 59.4 1080 157 

1350 196 

1380 200 

13 [0oi6 6 32.2 

16 [O O112 12 27.8 

17 [O" f 45 "Is2 12 28.5 

22 [O" 2 4501, 6 32.3 

23 [O" 2 4501, 6 31.7 

24 [O" f 4501, 6 30.9 

25 [0" 2 45'1, 6 29.2 

26 [O" 2 4501, 6 30.6 

aFlewral strength 

59.8 1903 

1510 

1030 

64.6 a1640 

a1730 

a1780 

63.8 a 786 

a793 

a848 

59.7 331 

379 

352 

60.4 337 

317 

483 

61.2 414 

434 

538 

63.2 414 

414 

455 

61.6 345 

538 

276 

219 

150 

a238 

a25 1 

a258 

a114 

a115 

a123 

48 

55 

51 

49 

46 

70 

60 

63 

78 

60 

60 

66 

50 

78 

379 55 

The second batch, 2.3 kg (5 Ib) of a 0.08 m (3 in.) wide prepreg tape, was purchased at the completion of the Phase I 10.000 hour flight simulation tests. The third batch, 5.4 kg (12 lb) of 0.3 m 

(12 in.) wide prepreg tape, was ordered midway through Phase I1 to provide test specimens for 

the add-on effort to study the effect of thermal aging on compression properties. Both of these 

orders were delivered on schedule and passed all the quality assurance requirements for both 

the prepreg and fabricated laminates. At the time of the last purchase the designation had 

changed from A-S/3501-5 to A-S/3501-5A. Data sheets from Hercules Incorporated for the two 

designations showed them to be very similar with only a few slight differences in properties. 

3-9 
3.2.2.3 Boron/Poiyimido. The 4.0-mil B/PlO5AC borodpolyimide, referred to as B/PI, was 

obtained from AVCO Corporation, Systems Division, Lowell, Massachusetts. As no Genera3 

Dynamic3 procurement specification existed, the material was ordered to the following 

requirements: 

a. Fiber Strength: 500,000 psi minimum 

b. Resin Content: 38 to 48% 

c. Volatiles: 10 to 20% 

d. Flow: 450 K (350' F) 30 to 42% 

A total of 28 kg (62 Ib) of prepreg in 0.3 m by 1.8 m (12 in. by 72 in.) sheets was required for the 

program. Delivery of the prepreg was delayed approximately 3 months because of difficulties 

encountered by AVCO in obtaining the polyimide resin from the manufacturer, Ciba-Geigy. 

Upon receipt of the prepreg, quality assurance testing was performed. The material was found 

to be acceptable and two quality assurance panels were fabricated. Quality assurance testing 

of the laminates was completed with the material being found acceptable. While the BlPI 

material passed the prepreg and laminate quality assurance requirements, it wag later removed 

from the test program because of low tensile properties and extensive thermal degradation 

obsented during thermal aging and short-term flight simulation testing. Complete details of 

the problems encountered with this material system and the reasons for the poor results are 

found in Appendix A. 

3.2.2.4 Graphito/Polyimido. The graphitdpolyimide (G/PI) system selected at the beginning of 

the contract was HT-S/710. In like manner to the G/E material, three batches were required: 

one for Phases I and 11, one for the Phase I1 25,000 hour.fligfit simulation tests and the G/PI 

new material screening evaluation. and the third for the add-on effort to study the effects of 

thermal aging on compression properties. A second G/PI system, Celion 6000/LARC-160, was 

added midway in the contract for testing to the Phase I requirements including 10,000 hours 

of thermal aging and flight simulation exposure and for evaluation in the add-on thermal 

aginglcompression study. 

HT-SnlO G/PI. This composite system was the most difficult material to procure. More than a 

year passed from the time of placing the initial order until completion of quality assurance 

testing and acceptance of the test material. One of the reasons for selecting HT-SI710 was the 

belief that the prepreg was readily available from at least three material suppliers. As it turned 

out one source was not interested in the small volume required for the program while the other 

two sources experienced considerable problems in fabricating acceptable prepreg. Because no 

General Dynamics procurement specification existed the material was ordered to the following 

requirements: 

a. Fiber Strength: 350,000 psi minimum 

b. Resin Content: 35-45% 

c. Volatiles: 10-20% 

d. Flow 15.25% 

e. Acid Equivalent No.: 520 to 540 

3-10 
The following is a chronological account of how the initial material was obtained. 

Jull973 

Aug 1973 

Sep 1973 

Oct 1973 

Dec 1973 

Jan 1974 

Feb 1974 

Mar 1974 

Apr 1974 

May 1974 

An order for 17 kg (37 lb) of prepreg tape was placed with E.I. DuPont De 

Nernom and Company, Fabrics and Finishes Department, Saugus, California. 

The prepreg was delivered on schedule but was rejected because of low fiber 

volume and low acid equivalent number. 

A second batch was delivered. The prepreg properties were within specification 

requirements; however, a panel fabricated from the prepreg showed precipitation and waviness. At this time a second source of the prepreg was being 

checked. Hercules Incorporated quoted a 50% higher price and an 8 to 10 week 

delivery schedule. Wont reported that they discovered that the resin used to 

fabricate the first two batches of prepreg was bad and that a third batch, made 

with new resin, would be prepared. 

The third batch of prepreg delivered by DuPont failed to produce acceptable 

panels and was also rejected. A new order was placed with Hercules for the 17 kg 

(37 lb). We the price was about 50% higher, Hercules had had considerable experience with the prepreg, and no further difficulties were anticipated. Delivery, 

however, was not expected until December 1973. 

Delivery was delayed to allow the acid equivalent number of the polyimide resin 

to reach an acceptable value prior to manufacturing the prepreg. 

Delivery was further delayed when it was determined that the HT-S graphite 

fibers were not within specifications. A new batch of fibers was obtained and 

manufacture of the prepreg material was scheduled for February 1974. 

The first batch of prepreg manufactured by Hercules failed to meet specifications. A second batch was prepared from the remaining fiber and resin. This 

material was found to be acceptable, but the quantity was less than that 

ordered, 11 kg (24 lb) of an order of 17 kg (37 Ib). Because of the high costs incurred and the small size of the order Hercules was not willing to complete the 

Order. 

DuPont agreed to prepare a demonstration sample of prepreg. If the material 

was found to be satisfactory an order for the 17 kg (37 lb) was to be given to 

them. 

DuPont reported a delay in preparation of the demonstration sample because of 

out of date resin that was unacceptable for manufacturing prepreg. A new batch 

of resin was ordered from Monsanto. 

The small demonstration sample of prepreg was received from DuPont. The first 

panel fabricated from this batch was acceptable, but later panels tended to precipitate and blister. Because of these inconsistencies no order was placed with 

DuPont. As a result of the numerous delays encountered in obtaining the mate 

- rial it was decided that the 11 kg (24 lb) batch produced by Hercules would be 

purchased for the test program if satisfactory test panels could be fabricated. It 

was thought that sufficient panels could be made from the 11 kg (24 lb) to satisfy 

the program requirements. A small test sample was, therefore, ordered from 

Hercules. 

3.11 
Jun 1974 The sample submitted by Henrules was evaluated and found to produce an excellent laminate. To check the consistency of the prepreg a second panel was fabricated, and, again, an acceptable laminate was obtained. Based on these results 

the remaining 10 kg (22 Ib) of prepreg were purchased from Hercules. 

The material was delivered and successfully passed all of the quality assurance 

requirements for both the prepreg and fabricated laminates. 

JUU 

Aug 1974 

The second batch of HT-SI710,4.1 kg (9 lb) of 0.08 m (3 in.) wide prepreg tape, was purchased 

from E.I. DuPont De Nemours and .Company. The matexial was delivered on schedule and 

passed the prepreg requirements. Quality assurance results from a fabricated laminate were 

comparable to those obtained from the Phase I batch and the mat& was accepted. 

The third batch of HT-S/710 GIPI, required for the thermal aginglcompression study, could 

not be obtained. As discussed in Section 3.1.2.4, problems existed with both the fiber and the 

resin system. 

Celfon 6000/uRC= 160. The Celion 6000/LARC-160 graphitdpolyimide composite was obtained 

from Fiberite Corporation, Wmona, Minnesota. Twenty-three kg (50.7 lb) of 0.3 m (12 in.) pre 

preg tape were purchased to a Fiberite specification for hyE1678E. The material was received 

on schedule and passed all the quality assurance requirements for both the prepreg and fabricated laminates. 

3-12 
SECT10N 4 

PREPREG QUALITY ASSURANCE TESTlNG 

The organic matrix composite materials were purchased to specific requirements on the purchase orders or, when available, to existing General Dynamics or industry specifications. To 

ensure that these requirements were satisfied the following properties were determined upon 

receipt of the prepreg: 

a. Fiber content (percentage by weight). 

b. Resin solids (percentage by weight). 

c. Volatiles (percentage by weight). 

d. Resin flow (percentage by weight). 

e. Process geL 

A description of the test methods that were used is given in Appendix B. 

4.1 BORON EPOXY 

BE was purchased from AVCO Corporation to General Dynamics Specification FMS-2001B. 

Two batches of material were obtained during the program. 

The FMS2001B specification requirements, the values from the AVCO Certified Test Reports 

received with the shipments, and the results of the Convair quality assurance tests are given in 

Tables 41 and 4-2. Both the AVCO and Convair results for batch one satisfied the procurement specification requirements and the material was considered to be acceptable. For batch 

two, the Convair results for resin solids and resin flow were slightly above the procurement 

specification requirements but the material was considered acceptable contingent on satisfying the bminate quality assurance tests. 

Tablo 4-1. Pnpng Roquiremontr and Proporfios for Rigidito 5505 WE - Batch One 

Prop.* 

Roquiromonts of AVCO Certifiod Convair Qualify 

Spoceif ication Tost Roport Assurance Test 

FMS-20018, % Data, % Results, % 

Fiber Content (wt %) - - 

Resin Solids Content (wt So) 29-34 31.7 

Volatile Content (wt %) 2% max. 1.08 

Resin Flow (wt %) 10-20 13.5 

68.1 

31.0 

0.9 

15.7 

4- 1 
Tablo 4-2. Pnpmg Roquiromonts and Proportlor for Rigidit. 550s WE - Batch Two 

Rmquirmmonts of AVCO Cortlfiod Comair Quality 

Spocclf icotion lost Roport Assuranco lest 

FMS-20018, % Data, % Rorults, % 

~~~~ ~~~~ 

Fiber Content (wt %) 

Resin Solids Content (wt %) 

Voiatile Content (wt %) 

Resin Flow (wt %) 

- 

29-34 

2% max. 

10-20 

32.7 

0.6 

17.0 

~~~ 

64.5 

34.8 

0.7 

24.9 

4.2 GRAPHITE EPOXY 

The G/E prepreg was purchased from Hercules Incorporated to Hercules Specification HDSG-26006A. Three batches of material were obtained during the program. 

The HD-SG-2-6006A specification requirements, the values from the Hercules Quality Assurance tests received with the shipments, and the results of the Convair Quality Assurance tests 

are given in Tables 4-3 through 45. Both the Hercules and Convair results satisfied the prcl 

curernent specification requirement3 and the materials were considered to be acceptable. 

Tablo 4-3. Pnprog Roquiromonts and Proportlor for A-S/3501-5 G/E - Batch On. 

Roquinmonts of Horculos Quality Convair Quality 

Spuificotion . Assuranco lest Assurance Tost 

prop* HO-SE24006A. 7% Rorults, % Rosultr, 7% 

Fiber Content (wt %) - 58.3 

Resin Solids Content (wt %) 39-45 40.5 

Volatile Content (wt %) 2% max. 1.2 

Resin Flow (wt %) 15-35 18 

57.2 

41.6 

1.2 

21.7 

4-2 
Table 44. Pnpng Requirements and Propertior for A-S/3501-5 G/E - Batch Two 

Roquinrnonh of Hercules Quality Comair Quality 

Spufffcotion Assuranco T& Assurance Test 

Prow HD-SG2-6006A. Ye ROSIJ~~S, Ya Results, % 

~~ ~ 

Fiber Content (wt %) - - 58.8 

Resin Solids Content (wt So) 39-45 41 40.7 

Volatile Content (wt So) 2% max. 0.61 0.49 

Resin Flow (wt %) 15-35 - 21.1 

Tablo 1-5. Pnpng Roquinmentr and Properties for A-S/3501-5A WE - Batch Threo 

Roquiremonh of Herculos Quality Convair Quality 

Spdficotion Assurance Test Assurance Test 

Prop** , HO-SG2-6006A. Ye ResultS. % Results. % 

Fiber Content (wt %) - 

Resin Solids Content (wt So) 39-45 

Volatile Contept (wt SO)’ 2% max. 

Fiber Areal Weight (gm/m2) - 

Resin Flow (wt %) 15-35 

- 

43.6 

- 

- 

146 

56.1 

43.2 

0.7 

24.5 

151 

4.3 BORON POLYIMIDE 

The B/PI prepreg was purchased from AVCO Corporation with a list of Convair requirements. 

No procurement specification or industry specification was available. Table 4-6 lists these Convair requirements, the values from the AVCO Certificate of Conformance received with the 

shipment, and the results of the Convair quality assurance tests. AlI of the AVCO and Convair 

results except for the slightly high Convair value for volatile content satisfied the Purchase 

Order requirements. The small difference, 0.7% high, was not considered to be sufficient cause 

for rejection of the shipment. 

4-3 
Tablo 4-6. Propreg Roquinmonts and Proportios for B/PlOSAC B/Pl - Rosin Batch Numbor 5-73, 

Shoot Numbors AX-PI OSA-38 to AX-P 1 05A- 105 

Convair Quality 

Roquinmonts of Comair AVCO Cortlficato of Assumnco Tost 

Pro# Purehaso Ordor, % Conformance Data. % Rosults, % 

Fiber Content (wt %) - 

Resin Solids Content (wt %) 38-48 

Volatile Content (wt 9%) 10-20 

Resin Flow (wt %) 30-42 

- 

38-46 

14-20 

- 

35.7 

43.6 

20.7 

38.3 

4.4 GRAPHITE POLYIMIDE 

4.4.1 HT-S/7lO. The HT-S/710 GlPI prepreg was purchased from Hercules Incorporated with a 

list of Convair specified requirements. No procurement specification or industry specification 

was available. For the two batches of material required on the program, Tables 4-7 and 4-8 list 

these Convair requirements and the results of the Convair quality assurance tests. No quality 

assurance test results from Hercules were received with the shipments. As shown in the tables, 

the Convair results satisfied the Purchase Order requirements and the material was considered 

to be acceptable. 

Tablo 4-7. Proprog Roquiromonts and ProporHos for HT-SDlO G/PI - Batch Ono 

Roquinmonts of 

Convair Purehaso 

Ordor. % 

Convair Quality 

Assurance Test 

Results. % 

Fiber Content (wt %) 

Resin Solids Content (wt 70) 

Volatile Content (wt So) 

Resin Flow (wt So) 

- 

35-40 

10-20 

15-25 

44.4 

40.0 

15.6 

17.2 

4-4 
Tablo 4-8. Pnpreg Roquinmonts and Proportior for HtW7lO G/Pl - Batch Two 

Roquinmonk of 

Convair Purehaso 

Odor, % 

Convair Qualily 

Assuranco Test 

Rosults, % 

Fiber Content (wt %) 

Resin Soiids Content (wt %) 

Volatile Content (wt %) 

Resin Flow (wt %) 

- 

35-40 

10-20 

15-25 

48.5 

35.5 

16.0 

18.2 

4.4.2 CEUON 6000/LARC-160. The Celion 60001LARC-160 prepreg was purchased from Fiberite 

Corporation to the Fiberite Specification for hy-E 16783. The results of the Convair quality 

assurance tests and the values from the Fiberite Certification received with the shipment of 

prepreg are given in Table 49. 

Tablo 4-9. Pnpreg Proportior for Cdion 6000/LARC-160 

Fiborih Codifid 

Tat RopoH 

Data, % 

Convair Quality 

Assurance Test 

Rosults, % 

Fiber Content (wt %) 

Resin Solids Content (wt So) 

Volatile Content (wt %) 

Resin Flow (wt so) 

Fiber heal Weight Cgm/m2) 

47.0 

40.4 

12.6 

- 

140.9 

54.6 

37.5 

7.9 

27.6 

145 

4-5 
SECTION 5 

LAMINATE PROCESSlNG AND SPECIMEN FABRICATION 

5.1 LAMINATE PROCESSING 

All of the composite fabrication3 on this program were processed using autoclave curing 

methods. This included flat panels that ranged in size up to 0.93 m2 (10 ftz), and 0.025 m (1 in.) 

diameter by 0.3 m (12 in.) long tubes. Wo of the materials, GIE and B/PI, were fabricated into 

both flat panels and the torsion tube configuration. In addition, precompaction and/or “B” 

staging operations were included in several of the ovefall processes and are described in detail. 

The G/E tubes and the BlPI tubes both included the precompaction operations prior to final 

cure. HT-9710 G/PI pan& also required a precompaction and “B” staging of the condensation type polyimide system in order to achieve satisfactory densification during cure. The B/E, * GIE, and Celion 60001LARC-160 GlPI systems, however, were processed using simple rapid 

curing schedules. 

5.1.1 BORON/EPOXY. The fabrication of the BE panels from prepreg consisting of 4-mil 

diameter boron and AVCO 5505 resin was accomplished using the relatively straight-forward 

vacuum bag and autoclave cure cycle shown below. 

1. With full vacuum in the bag, heat at 1.7 to 2.8 K (3 to 5 O F) per minute to 450 K (350’ F). 

2. Apply 0.59 MNlm2 (85 psi), venting vacuum bag to atmosphere when autoclave pressure 

reaches 0.34 MNh2 (50 psi). 

3. Hold 2 hours at 450 K (350’ F‘) and 0.34 MNlm2 (50 psi). 

4. Cool under pressure to 353 K (175 O F). 

Laminates produced by this process exhibited good densification with measured specific 

gravities exceeding 2 and having resin contents of 26 to 32 weight percent. The layup procedum used for the 6 and 12 ply laminates (Figure 5-1) was identical except for the number of 

bleeder plies used. In all cases the prepreg was laid up with the scrim side down and with 0.001 

m (0.04 in.) holes on 0.025 m (1 in.) centers perforating the Teflon film applied above the 120 

glass bleeder. 

5.1.2 GWPHITE/EPOXY. A-S/3501-5 prepreg produced by Hercules Incorporated was the G/E 

system used to fabricate the flat panels and the 0.025 m (1 in.) diameter by 0.3 m (12 in.) long 

tubes required for Phases I and 11. Both the 6 and 12 ply laminate fabrications used identical 

cure schedules and layup procedures with one exception: the thinner panels used two 120 glass 

bleeders while the thicker panels used four plies of the same glass fabric for bleeders. The cure 

cycle used a constant heatup rate of 1.1 to 1.7 K (2 to 3 O F) per minute to 380 K (225 O F) with 

full vacuum starting at room temperature. After applying 0.69 MNlrn2 (100 psi) the 

temperature was held 1 hour at 380 K (225 O F) before heating to 450 K (350 O F). A hold of 1 hour 

at 450 K (350 O F) was followed by cooldown to below 350 K (150 O F) under pressure. Details of 

the layup material stacking sequence are shown in Figure 5-2. 

5- 1 
MOOIFIEO NYLON FILM 

VACUUM BAG (VACPAK) -- - - - STYLE 1534 GLASS CLOTH 

SEPA RAT0 R (TEF LON F1 LM) 

TEFLON COATED 

GLASS (ARMALONI 

-- - -- 

----------- 

............................................... .... ............................................. <I ----------- ****-* 

*1 PLY FOR 6 PLY EIE LAYUP 2 PLIES FOR 12 PLY BiE LAYUP 

Figun 5-1 Schomatic of B/E Cur0 Layup 

MODIFIED NYLON FILM 

VACUUM BAG (VACPAK) 

SEPARATOR (TEFLON FILM) 

STYLE 120 GLASS CLOTH' 

HIE LAYUP 

SEPARATOR (TEFLON FILM) 

61321 7-3 

-- - -- STYLE 1534GlASS CLOTH 

-- - 

- - -!STYLE 120 GLASS CLOTH' -- --I 

................................................... ........................... ........................ SEPARATOR (TEFLON FILM] 

'2 PLIES FOR 6 PLY GIE LAYUP 4 PLIES FOR 12 PLY GIE LAYUP 

GIE LAYUP 

SEPARATOR (TEFLON FILM) 

6632134 

Figure 5-2 Schomatic of G/E Cure Layup 

Fabrication of the GE tubes was accomplished with an aluminum mandrel covered with a silicone rubber bladder. Two plies of G/E prepreg, 0.31 m (12.25 in.) by 0.30 m (12 in.) were cut and 

laid up with a 0.013 m (0.5 in.) offset. These plies were precompacted (Figure 5-3) in an oven at 

338 K (150" F) to 344 K (160" K) for 20 minutes with full vacuum in the bag and 0.10 MNlm2 

(15 psi) external pressure before wrapping on the mandrel. A spiral wrap of paper bleeder 0.038 

m (1.5 h) wide was then applied over the separator covering the prepreg. Vent material addition and bagging were followed by the following short cure cycle. 

5-2 
MOOlFlED NYLON FILM 

VACUUM BAG (VACPAK) -- - -- 

SEPARATOR (TEFLON FILM) ---- -- 

/-------- --- .............................. TEFLON COATED ...................................... ............. GLASS (ARMALON) \------ - 

STY LE 1534 G LASS CLOTH 

SFYLE 120 GLASS CLOTH 

GIE LAYUP 

SEPARATOR (TEFLON FILM) 

65321 7-6 

Figuro 5-3 Schomatk of G/E Tub. Pmompadion Layup 

1. With full vacuum in the bag and 0.10 MNlrn2 (15 psi) in the autoclave, heat to 450 K (350 O F) 

at 1.7 to 2.8 K (3 to 5" F) per minute. 

2. Hold 1 hour and cool to 338 K (150" F) with vacuum in the bag and 0.10 iMNlm2 (15 psi) 

pressure in the autoclave. 

This procedure produced composite tubes having a specific gravity of 1.56, resin contents of 31 

to 33 weight percent, and uniform wall thicknesses ranging from 0.00097 to 0.00102 m (0.038 

to 0.040 in.). 

The A-SI3501-5A prepreg purchased for the thermal aging/compression study was processed 

using the same cure schedules and lamp procedures as the A-S/3501-5 prepreg except that for 

the 24 ply laminate, eight plies (or equivalent) of 120 glass bleeder were used. 

5.1.3 BORON/POLYlMIDE. AVCO Corporation's 4.0-mil BIP105AC prepreg was the system 

used to fabricate the B/PI pan& and tubes. The flat laminates used a top caul plate and bleeding from both sides, as shown in Figure 5-4. A single layer of 104 scrim cloth was added to the 

prepreg stack to balance the layup. Full vacuum in the bag was applied at room temperature 

and the layup was heated in the autoclave to 464 K (375 O F) at 2.2 K (4 O F) per minute and held 

there 2 hours. Pressure of 0.69 MNlm2 (100 psi) was applied and the temperature was raised to 

561 K (550' F), again at 2.2 K (4" F) per minute. After a 2-hour hold, the layup was cooled to 

353 K (175 O F) at 1.1 K (2" F) per minute under autoclave pressure. Post cure consisted of a free 

standing heating cyde of 4 hours at 589 K (600" F). 

Torsional tubes were fabricated using the 4.0-d B/PlOSAC prepreg in a manner similar to the 

technique used for the G/E tubes. Aluminum mandrels covered with silicone shrink tubing 

were rolled with a two-ply 0.311 m (12.25 in.) by 0.308 m (12.125 in.) staggered prepreg that had 

been precompacted for 15-20 minutes at 383 K (230 O F) with full vacuum in a bag using the layup shown in Figure 5-5. As before, bleeder paper 0.038 m (1.5 in.) wide was spirally wrapped, 

but two layers of bleeder were used for the more volatile polyimide resin system. One ply of a 

polyimide film was applied over the bleeder, followed by a venting system consisting of one 

layer of 1534 glass cloth, a layer of mat, and a final layer of 1534 glass fabric. The following 

cure and post cure cycles were used. 

.. 
MODIFIED NYLON FILM 

VACUUM BAG (VACPAK) 

-- --A 

STYLE 181 GLASS CLOTH POLYIMIDE FILM (KAPTON), 

SPLIT AT 2 IN. INTERVALS A -- 

----------- 

-- 

- 

- 

--H 

- 1 

----------- 

.......................................... r.-:.-:.- ..-.-;........................................ 1 B/PI LAYUP 

----------- 

----------- < - - STYLE 181 GLASS CLOTH 

TEFLON COATED 

G LASS (AR MA LO N) 

-- 

SEPARATOR (TEFLON FILM) 

653213.6 

Figurm 5-4 Schomutic of B/PI Cum Layup 

MODIFIED NYLON FILM 

-- - -I VACUUM BAG (VAC-PAK) 

-- - -- I STY LE 1534 G LASS CLOTH 

SEPARATOR (TEFLON FILM) 

1.- .................................................. TEFLON COATED .................................................. 1 B/PI LAYUP 

GLASS (ARMALONI 

-- - 

-- - -- -1 STY LE 120 GLASS CLOTH 

SEPARATOR (TEFLON FILM) 

6532 7 7-7 

Figuro 5-5 Schomatic of B/PI Tube Precompadion Layup 
a. cure 

1. With full vacuum in the bag and 0.17 MN/d (25 psi) autoclave pressure, heat to 464 K 

(375" F) f 6 K (10" F) at 2.2 K (4" F) per minute. 

2. Hold 2 hours and apply 0.69 MNIm2 (100 psi). 

3. Heat at 2.2 K (4" F) per minute to 561.K (550" F) and hold 2 hours. 

4. Cool to 353 K (175" F') or below at 1.1 K (2" F) per minute under autoclave pressure. 

b. Postcure 

1. Heat to 589 K (600 O F) at 2.2 K (4 " F) per minute and hold 4 hours. 

2. Cool below 366 K (200 " F) at 1.7 K (3 " F) per minute maximSeveral of the tubes fabricated in this manner used lightweight glass sleeving as bleeder in lieu 

of the paper. 

5.1.4 GRAPHITE/POLYIMIDE 

5.1.4.1 HT-S/7lO. Processing of the HT-SI710 GIPI system required a precompaction cycle for 

"B" staging and volatile removal before cure. This procedure consisted of sandwiching the prepreg layup using six plies of 120 &s cloth bleeder covered by separator applied to both sides. 

The layup was vacuum bagged and heated to 353 K (175" F) at 2.8.K (5" F) per minute. An 

external pressure of 0.17 MNImz (25 psir was then applied and held 15 minutes before cooling 

to 325 K (125" F). The layup was then bagged and fresh bleeder applied (Figure 5-6) before 

applying the following cycle. 

1. With full vacuum in the bag, heat to 353 K (175 " F) at 1.7 K (3 F) to 2.8 K (5 O F) per 

minute. 

2. Hold 30 minutes at 353 K (175 " F), heat to 400 K (260" F). 

3. Hold 25 minutes at 400 K (260" F), apply 0.69 MN/m2 (100 psi) autoclave pressure. 

4. Heat to 450 K (350" F') and hold 2 hours. 

5. Cool at 1.1 K (2 " F) per minute under autoclave pressure to 353 K (175 " F). 

A slow 1.1 K (2" F) per minute, postcure in air to 644 K (700" F) completed the processing 

cycle. 

5.1.4.2 Celion 6000/LARC-l60. The procedure used for fabrication of the Celion 6000ILARC-160 

laminates was the following: 

A partial vacuum of 0.010 MNIm2 (1.5 psi) was applied to a laminate, and the temperature was 

increased to 366 K (200°F). The laminate was held at this temperature for 15 minutes to allow 

most of the solvent to be removed. The temperature was then increased to 435 K (325°F) and 

held for 1 hour to allow the laminate to imidize and to remove the iniidization by-products. The 

laminate was then heated to 603 K (625°F) and held for 2 hours. During the heat-up to 603 K 
MOOlFlED NYLON FILM 

VACUUM BAG (VACPAKI 

SEPARATOR (TEFLON FILM) 

STY LE 181 GLASS CLOTH 

STY LE 1 534 G LASS CLOTH 

STY LE 181 GLASS CLOTH 

9 STYLE 120 GLASS CLOTH 

I--- - -- - - - - - ................................................... ................................................... GiPl UYUP TEFLON COATED 

GLASS (ARMALON) 

'STYLE 120 GLASS CLOTH 

Figure 54 Schodc of HTW710 C/Pl Cur. Layup 

(625 OF), a full vacuum and 1.4 MNlm2 (200 psi) were applied to the bminate when the laminate 

temperature reached 546 K (525 OF). After the %our hold at 603 K (625 OF) the temperature 

was decreased to 338 K (150 OF), the pressure and vacuum were released, and the laminate was 

removed from the autoclave. Details of the layup material stacking sequence are shown in 

Figure 5-7. 

5.2 BORON/ALUMINUM TUBE FABRlCAflON 

The unidirectional BlAl tubes were specially designed and fabricated for the torsion tests 

(baseline shear) using a variation in the General Dynamics Convair BlAl tube making process. 

The tubular specimen was 0.15 m (6 in.) long, 0.025 m (1 in.) in diameter, and 0.001 m (0.004 in) 

(6 ply) in thickness. The boron filaments were 5.6 mils in diameter and oriented in the long-axis 

direction of the tube. The composite was 48 volume percent boron in a 6061 aluminum matrix. 

A sketch of the tube is shown in Figure 5-8. 

The BlAl tube with steel reinforced ends was made by winding unidirectional, singlelayer tape 

on a thin-wall, annealed, mild steel tube (inner mandrel). The assembly was then inserted into a 

thick-wall steel tool (outer mandrel). The assembly was next sealed by welding rings to the 
VACALLOYBAG 

-- - 

-- - - - I STYLE 1534 GLASS CLOTH 

-- - 

- I STYLE 181 GLASS CLOTH 

-- - - STYLE 120 GLASS CLOTH 

-- - - 

............................................. ........................................... ........................................... G/Pl LAYUP 

6138 SILJCONE 

COATED PEELPLY 

-- - 

-- - - - 1 STYLE 181 GLASS CLOTH 

653217-315 

Figure 5-7 Schematic of Collon 6000/LARC-160 G/Pl Cur. Layup 

DnAlL: @ 

0 

a001 m ((1040 in) THICKNESS (6 PLY) B/AI TUBE 

(LO024 m ((1094 in.) THICKNESS STEEL REINFORCEMENT 

(LO0071 m ((1028 id THICKNESS STEEL REINFORCEMENT 

653217-9 

Figure 5-8 Unidiredional B/AI tub. 

5-7 
ends of the inner and outer mandrels. An air evacuation tube was installed through the outer 

mandrel. The assembly was then helium leak tested, evacuated, heated to outgas any contaminants, and sealed (at the evacuation tube) while still under vacuum. 

High temperature, high pressure autoclave bonding followed. The ductile, hin-wded inner 

mandrel yields during the cycle, transferring the applied pressure to the composite and its 

near-molten aluminum matrix. The result is complete bonding of the B/N layers together and 

the bonding of the inner and outer steeI tubes to the composite. 

The bonded assembly wag masked prior to etching by dipping each end in a molten plastic 

maskant to a depth of 0.038 m (1.5 in.), thus leaving the center portion free. The assembly was 

then submerged in a nitric acid bath that selectively dissolved both the inner and out steel 

jackets, leaving a BlAl tube with steel reinforced ends. 

5.3 SPECIMEN FABRICATION 

5.3.1 PANEL LAYOUTS AND SPECIMEN IDENTlFICATION. Cutting plana were prepared for each 

composite laminate showing the location and identification number of each specimen. These 

plans accompanied &ch laminate through the cutting and machining operations and then were 

retained to keep a permanent record as to which panel ana the specific location from which 

each specimen was obtained. A typical cutting plan for a BlAl laminate is shown in Figure 5-9 

and for a G/E laminate in Figure 5-10. The BlAl panels had good edges and could be used without trimming. For the resin matrix composites, a 0.013 m (0.5 in.) strip was cut from the perimeter of each panel because of edge thinning. A cutting allowance of 0.0032 m (0.125 in.) was 

used in preparing the cutting plans. Specific gravity and resin content samples were cut from 

scrap pieces (not from the perimeter) for each resin matrix laminate. For the B/Al panels a 

0.013 m (0.5 in.) strip in the longitudinal direction was cut from one edge of each panel for quality assurance testing (fiber content, tensile properties, and metallographic examination). 

The numerous test specimens used during this program were identified according to the 

system described in Table 5-1. The ID numbers wereplaced on the resin matrix specimens with 

either a white pencil on the composite itself or with permanent black ink on the end doublers. 

For the B/Al specimens either permanent black ink on the BlAl or vibra-tool marking of the 

metal end doublers was used. No difficulties with unreadable ID numbers were experienced 

during any of the moisture or thermal exposures. Data books with a record of each specimen 

were prepared for the six material systems. All test results were filed in these logs, thus ensuring ready access to any of the data developed during the lengthy program. 

5.3.2 MACHINING OPERAllONS. The machining operations required for the fabrication of the 

composite specimens included sawing, hole drilling, and slotting. Cutting of the flat laminates 

and torsion tubes into specimen blank.9 was accomplished with the Iarge gantry-type diamond 

radial saw pictured in Figures 5-11 and 5-12. The resin matrix materials were cut using a 0.25 

m (10 h) diameter diamond plated blade. The diamond size was grit number 60 (270 micron). 

Sawing was done without coolant at a cutting rate of 5600 sfm with powered feed rates of 6 to 

10 ipm. The specimens were cut to final size with no edge finishing required. For B/Al a 0.25 m 

(10 in.) diameter diamond impregnated (grit number 46 (350 micron) metal bond) saw blade was 

used. Cutting rates were the same as those used for the resin matrix composites. The BlAl 

specimens were cut wet using a 24 to 1 ratio water soluble oil with flood application. Again, 

good edges were obtained without additional finishing operations. 

5-a 
c 

I 

- 

H 1 

.. 

z 

Y 

- 

0 

C 

0 

Q 

C 

0 

Q 

- 
T 

I 

c' .- 

14 - 0 

E 

*- -c 

I 

d 

c .- 

- 2 

E 

x 

c 

c 

e 

0 

L 

- 
Tablo 5-1. Spocimon Idontificotion Cod. 

"A BIE 

B GIE 

C BlPI 

E BlAI 

F GIPI (Celion 

D GlPI (HT-S/710) 

60001LARC-160) U unidirectional 

c Crossply 

aF Notched Crossply 

0 Baseline 

1. Tensile 

2. NotchedTensile 

5. Fracture 

6. Shear 

1 ThermalAging 

1. Tensile 

2. Shear 

2 Ambient Aging 

1. Control 

2. 20 weeks 

3. 52 Weeks 

3 MoistureExposure 

1. 24 hr H,O Boil 

2. 6WeekHumidity 

4 Atmospheric 

Exposure 

5 creep 

6 Stacking Sequence 

7 Fatigue 

8 Short-Term Flight 

Simulation 

Simulation 

10 Thermal Agingl 

Compression 

9 Long-Term Flight 

1. ResinMatrix 

2. MetalMatrix 

'1. [450, -450, OO], 

2. 100, 450, -457, 

3. [45", 0" -45"], 

1. Notched 

2. Unnotched 

1. Compression 

2. Sheer 

3. Fatigue 

4. Weight Change 

Transverse specimens are marked with "'I"' before the regular ID No. 

For example: E-C-7-1-24 is BlAL Crossply. Notched Fatigue Specimen No. 24 

aUsed on notched flight simulation specimens only. 

5-1 1 
Figure 5-1 1 Gantry-Type Diamond Rodial Saw 

Holes were required in the notched tensile, fatigue, and flight simulation (both short- and longterm) specimens of the resin matrix materials and in the notched tensile and fatigue specimens 

of B/AL Carbide and diamond coated core drills were used in a No. 2 Moore jig bore machine for 

fabricating the 0.0064 m (0.250 in.) holes in the resin matrix composites. No coolant was used. 

Electredischarge machining (EDM) was used to produce the 0.0032 m (0.125 in.) holes in the 

BlAl specimens. The equipment used was an Elox HRP 104 EDM machine with a graphite 

electrode. 

The other machining operations were cutting the center and edge slots in the B/Al laminate 

posite materials. The B/A1 shear specimens were slotted with the EDM method in a manner 

similar to that used for producing the holes in the notched tensile and fatigue specimens. For 

the fracture specimens, a rather narrow trapezoidal center notch with tip radii of approximately 0.00005 m (0.002 in.) was desired. Cutting of these notches was accomplished with a 

Sheffield ultrasonic impact grinding machine. The machine was operated at 20 kHz using 600 

grit (17 micron) boron carbide abrasive. Optical comparator measurements of the notch radii 

gave values of 0.000038 m (0.0015 in.) to 0.000064 m (0.0025 in.) for both the resin matrix and 

the metal matrix specimens. 

interfiber shear specimens and the center notches in the fracture specimens of all the com- I 
Figure 5-12 Closeup View of Diamond Radial Saw 

5-13 
In addition to the machining operations required on the composite materials, it was necessary 

to drill loading fixture attachment pinholes in the titanium end doublers of the creep and flight 

simulation specimens. These holes were produced using standard drilling practices with the 

aid of a drill fixture that was fabricated for this purpose. The holes were drilled on the centerline of the test specimens to within f 0.000064 m (0.0025 in.). 

5.3.3 DOUBLER BONDING. With the exception of 1) short be.- shear, 2) flexural, 3) fracture, 

and 4) laminate shear, all specimexw required tapered doublers bonded to the ends to prevent 

failures in the gripping fixtures during testing. The major criteria considered in the selection of 

the doubler materials were: 

a. Temperature capability. 

b. Thed expansion matching of doubler and specimen. 

c. Strength. 

d. cost. 

Most of the resin matrix specimens were fabricated with either epoxy-glass laminate or 

polyimidequartz laminate doublers depending on the test temperature. Early in the program, 

two epoxy-glass laminate types were used. One, Scotchply Type 1002, was used at and below 

mom temperature and the second, Scotchply Type 1007, a high temperature epoxy, was used 

for elevated temperature tests up to 450 K (350" F). To avoid the possibility of mistakenly 

using the Type 1002 material at elevated temperature, where its strength is quite low, a 

change was made midway in the program, and the Type 1007 epoxy was used for both the low 

and high temperature test specimens. For testing of resin matrix specimens at temperatures 

above 450 K (350" F) a polyimidequartz (Skybond 703lAstroquartz) doubler material was 

used. Both the epoxy and the polyimide doublers were 0.0015 m (0.060 in.) in thickness. 

Doublers for the majority of B/Al specimens were made from 6061-T4, aluminum alloy of the 

same nominal thickness as the composite being tested. 

The exceptions to the above described specimens were the creep and flight simulation specimens of both the resin and metal matrix systems. These specimens were pin loaded in contrast 

to the others, which were either clamped or loaded through a wedge action type of gripping fixture. The creep and flight simulation specimens also had the requirement of being able to sustain a longtime elevated temperature test environment. For a pin loaded application, particularly at elevated temperature, a metal doubler is far superior to either the epoxy or polyimide 

types discussed above. Titanium was chosen as having the best combination of elevated 

temperature strength and thermal expansion match with the composite systems. For the [0 O f 

45 'Is crossplied specimens, annealed A-70 titanium alloy (MIL-T-9046F, Type I, Composition 

B) was used. Because of the requirement of much higher test loads, solution treated and aged 

Ti-6Al-4V titanium alloy (MIL-T-9046F. Type 111, Composition C) was selected for the 

unidirectional BlAl specimens. The thickness of the titanium alloy doublers was 0.0018 m 

(0.070 in.). 

Of equal importance to selection of doubler material was the choice of adhesives used in bonding them to the specimen ends. The major considerations were strength at temperature and, 

for the creep and flight simulation specimens, long-time thermal stability. At room temperature and below, EA-9309(l) was selected €or the resin matrix specimens and EC-2216 for the 

5-14 
B/Al specimens. EA-9309 is a twepart epoxy paste adhesive manufactured by the Hysol Division of the Dexter Corporation. EC-2216 is a modified amine-cured filled epoxy manufactured 

by the 3M Company. Both wem cured at room temperature for at least 72 hours prior to use. 

The adhesive that was considered to have had the best combination of elevated temperature 

strength and stability and the one selected for all elevated test specimens was HT-424. This is 

a modified epoxy-phenolic film adhesive with an aluminum filler on a giass carrier manufactured by the Sloomingdale Department of American Cyanamid Company. The cure cycle used 

for HT-424 was 1 hour at 450 K (350O F). The specimens were vacuum bagged and cured in an 

autockve at 0.14 MN/m2 (20 psi]. Prior to application of the adhesive the specimens and 

doublers are prepared as follows: 

a Resin-matrix specimens and doublers: lightly polished with A-1 very fine Scotch-Brite and 

solvent wiped. 

b. B/Al specimens and 6061-T4 Al doublers: degreased, lightly polished with A-1 very fie 

Scotch-Brite, solvent wiped, acid etched with Pasa-Jell 105, and rinsed with deionized 

water. 

c. Titanium alloy doublers: degreased, acid etched in nitric plus hydrofluoric solution, and 

rinsed with deionized water. 

(l)Midway through the program EA-9309 m replaced by EA934. also a room temperature curing two-part epoxy 

paste adhesive manufactured by Hysol Division. 

5-15 
SECTION 6 

LAMINATE QUALITY ASSURANCE TESTING 

Quality assurance testing was performed on alI purchased (B/U or Convair fabricated (Bm, 

GIE, BIPI, and G/PI) laminates to ensure good quality and, where applicable, concurrence with 

all specification requirements. Unacceptable material was rejected and replaced. In general, 

quality assurance testing consisted of visual examination, ultrasonic C-scan, volume or weight 

percent determinations, mechanical property testing, thickness measurement, and, for the 

BIAl panels, metallographic examination. Metallographic examinations were also made on all 

laminates fabricated for the thermal agingkompression study. The following sections describe 

the tests and present the results for the materials included in the program. 

. 

6.1 METAL MATRIX COMPOSITE MATERIAL 

The BIAl quality assurance test program consisted of nondestructive testing, volume percent 

determinations, thickness measurements, tensile tests, and metallographic examinations. All 

the composite panels were nondestructively evaluated by ultrasonic testing. The ultrasonic 

technique was pulsereflection through-trammission with a single short-focused, 5 MHz 

lithium sulfate transducer. The results are recorded on a C-scan recording wherein shades of 

gray lighter than some maximum are relatable to acoustic transmission losses within the test 

panel The overall integrity of a test panel is described by an arbitrary rating system with 

numerical values from 0 to 5 assigned on the basis of Convair's experience in ultrasonic testing' 

of hundreds of composite pan&. The rating reflects variations from nod conditions. Normal does not necessarily mean perfect For example, a few widely scattered stray boron filament fragments have no measurable effect on the structural performance of a given component. Although undesired from a workmanship standpoint, if small isolated defects cause no 

structural effects, they are judged to be nod Degrees of variation from normalcy are subjectively determined by engineers with wide experience in the evaluation of composite 

materials. Ratings of 3, 4, and 5 represent severe or widespread defects judged to have 

detrimental effects upon the structural performance of the component. Ratings of 1 or 2 relate 

to scattered or isolated defects caused by faulty workmanship or minor loss of process controL 

These ratings apply to defects that should not adversely affect the structural performance of 

the component. Table 6-1 summarkes the ultrasonic test results of both the unidirectional and 

crossply panels. Overall panel integrity was excellent. 

- 

. 

6- 1 
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0 

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Volume percent (do) determinations were performed on each B/Al panel to determine the 

amount of filaments and matrix present in the composite test materials. Broken tensile 

specimeps were used as test samples. Each specimen was washed, dried, and weighed. The filaments were leached from the samples in a 50% solution of sodium hydroxide at approximately 

339 K (150 O F). The filaments were collected, washed, dried, and weighed. Assuming the density of aluminum is 2700 k@m3 and the density of the filaments is 2510 kdm3 for 5.6 mil 

boron, the volume percentages of filaments were calculated using the following formula: 

Filament Volume Percent (do) = 

W 

4 

W W 

Q e 

F 

F 

x 100 

-+, A F 

A F 

WF = Weight of filaments 

QF = Density of filaments 

WA = Weight of Aluminum 

eA = Density of Aluminum 

The results of the vlo determinations are given in Table 6-2. For the 6 ply laminates the variation ranged from a low of 46.1 to a high of 51.4 vlo of boron with an oved average of 48.8 v/o 

boron. These results are typical of a nominal 50 vlo B/Al composite material. The 12 ply 

laminates were somewhat lower with values of 43.3 and 42.5 v/o boron. 

Thickness measurements were made on the B/Al panels using micrometers. The results (Table 

6-1) show relatively smaU differences in thickness within each panel or between the unidirectional and [O O f 45 OlS cmssplied BlAl panels. 

Tensile tests were performed on each of the BlAl panels. The configuration of the tensile 

specimen used was a simple, straight-sided specimen 0.15 m (6 in.) long and 0.013 m (0.5 in.) 

wide cut in the longitudinal direction. Bonded doublers were utilized to minimize stress concentrations in the gripping areas. The doublers were made from 6061 T-4 bare aluminum and 

were of the same nominal thickness as the composite being tested. The doublers were bonded 

to the composite specimens with EC-2216 epoxy, a modified amine-cured filled epoxy that 

cures at mom temperature. Tensile specimens were tested in an Instron testing machine using 

a head rate of 0.00025 m per minute (0.01 in. per minute). Stress-strain curves and strain 

measurements were made with Baldwin microfonners. Modulus values were calculated from 

the stress-strain curves. Results of the tensile tests on the 6 ply laminates, shown in Table 6-2, 

indicate an average tensile strength of 1450 MN/m2 (211 hi) for the unidirectional BlAl 

material and 581 MN/m2 (84.2 ksi) for the [0 O k 45 'Is crossplied layup. This is in good agree 

ment with early work at Convair on 5.6-mil BlAl composite material The elastic modulus 

values of 200 GN/m2 (29.0 million psi) for the unidirectional and 127 GNlm2 (18.4 million psi) 

for the crossplied layup, however, are about 6% less than typical of previous results. No reason 

could be found for the lower modulus properties. Because of the lower fiber contents, the tensile strengths of the 12 ply panels were somewhat lower than those obtained from the 6 ply 

panels. 

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Metallographic examinations were also made of each panel to substantiate the nondestructive 

evaluation (NDE) tests and to observe the microstructure, layup, and filament array. Samples 

cut from the broken tensile specimens were metallographically mounted and polished. Several 

of the samples showed some &bonding, which probably occurred during tensile testing or 

specimen preparation. Additional samples taken from the center of the B/AI panels instead of 

the edge where the quality assurance (QA) test specimens were taken showed none of the disbonding present in the broken tensile specimens. Photomicrographs of typical microstructures 

for the unidirectional and [O O 2 45 OlS crossplied material are shown in Figures 6-1 and 6-2. 

6.2 RESIN MATRIX COMPOSITE MATERIAIS 

The quaiity assurance test program for the resin-matrix fabricated laminates consisted of 

ultrasonic C-scan, resin and fiber content, and specific gravity, and for the 12 and 24 ply 

panels, fled and short beam shear strength. For the Phase I1 laminates and the new 

material addition, Celion 6000/LARG160, room and elevated temperature tensile tests and 

metallographic examinations were also performed. Metallographic examinations and glass 

transition tempera& (Tg) deterrmna * tions were also included for the laminates fabricated for 

the thermal aginglcompression study. The following sections present the resuits for the four 

resin-matrix composite systems. 

6.2.1 BORON/EPOXY. Ultrasonic C-scan testing was performed on all fabricated panels to 

detect possible delamination or voids in the laminates. Among the most useful nondestructive 

test methods for locating defects aligned parallel to pMcipal surfaces is the ultrasonic 

through-transmission method. A variation of this method, and the one employed for this program, is the singletransducer reflector plate method shown in Figure 6-3. 

Ultrasonic energy transmki through the specimen is returned to the transducer from a 

smooth reflector. If disbonds (or voids) are present, reflection of the sound at the interfaces 

results in less transmitted energy. The resultant energy drop is ready detected as a decrease 

in the strength (amplitude) of the return signal. The gated si& input to a recorder amplifier, 

provides an output voltage to the recorder that is proportional to the signal amplitude. 

The C-scan recorder is-mechanically linked to scan and index motions of the transducer and is 

electrically connected to the recorder amplifier. With a signal in the gate, the recorder receives 

the amplified signal and “writes” during that time. A series of the line scans synchronized with 

the movement of the test transducer (or test specimen) produces a full scale plan of the test 

object and shows defects in either black and white or variable shades. 

No rating system similar to that used for the B/Al C-scan evaluation was available for the 

resin-matrix materiais. The C-scan recordings were studied by an experienced NDE engineer 

who rated the panels as being either acceptable or unacceptable. Areas of suspected delaations or voids that were not sufficiently large enough to reject a laminate were marked on the 

paneL and these areas were avoided during specimen machining. For the BIE system all fabricated laminates were judged to be acceptable based on the ultrasonic c-scan tests. 

Resin and fiber content and specific gravity determinations were made on all of the BIE panels 

using the test methods described in Appendix B. The results are listed in Table 6-3 dong with 

the dimensions and intended use for each paneL 

6-7 
Figure 6- 1 Photomicrograph Showing Typical Microstructure of Unidirectional B/Al 

Figure 6-2 Photomicrograph Showing Typical Microstructure of [Oo f 4S0], Crossplied B/Al 

6-8 
REFLECTOR 

TRANSOUCER RECoRoER 

1 

L - 

GATED 1 

INOEX 

MOTION 

SIGNAL w 

TEST OBJECT 

BRIDGE 

*,em 

/ 

4 

I RECORDING q PAPER 

Flguro 6-3 Singlo-Transducor Ultrasonic Tlchniquo 

In addition to the nondestructive examination and the weight percent determinations conducted on each panel, a series of mechanical property tests was performed. Flexural strengths 

at 297 K (75 OF) and 450 K (350 OF) and short beam shear strength at 297 K (75 OF) were obtained 

for the two 12 ply QA panels. Details of the test procedures are described in Appendix B. Table 

6-4 lists the results. Tensile tests at 297 K (75°F) and 450 K (350°F) were performed on 

material from the 6 ply Phase I1 panel, LRC77-1. Test procedures are described in Section 7, 

Baseline Testing. Table 6-5 lists the results. Based on the results of the QA tests, all of the B/E 

fabricated laminates were considered to be satisfactory. 

6.2.2 GRAPHITE/EPOXY. Using the same technique described for the B/E material, all G/E fabricated laminates were ultrasonic C-scan tested for delaminations and voids. All panels were 

judged to be acceptable based on this evaluation. 

The results of the remaining quality assurance tests on the Phase I and 11 GIE laminates are 

presented in Table 6-6 for resin and fiber contents and specific gravity values, Table 6-7 for 

room and elevated temperature fled and room temperature short beam shear strengths 

(Phase I material only), and Table 6-5 for room and elevated temperature tensile strengths 

(Phase 11 material only). 

Specific gravity and fiber and resin contents for the three panels fabricated for the thermal 

agingkompression study are listed in Table 6-8. Also included is the glass transition temperature (Tg) for the 6 ply panel. The Tg measurements were made using a Perkin Elmer 3600 Thermal Analysis System utilizing a TMS-2 Thermomechanical Analyzer Module. Short beam 

shear and flexural strength data are presented in Table 6-9 for the 12 and 24 ply panels. The 

metallographic examinations of the full size panels are discussed in the thermal aging section 

of the report. 

In addition to the G/E panels, four G/E 8 ply unidirectional tubes were fabricated for the base 

line shear tests. From each 0.30 m (12 in.) long tube, three specimens and one QA sample were 

obtained. Resin and fiber contents and specific gravity determinations were made on each 

quality assurance sample. The results are listed in Table 6-10. 

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6-16 
Tablo 6-10. Quality Assumnc. Data for G/E Tubor 

Fibor bniont Rosin Contomt 

Tub. No. (w %.) (w, %) Spocffic Gravity 

66.6 

68.6 

67.5 

68.2 

33.4 

31.4 

32.5 

31.8 

1.56 

1.55 

1.56 

1.56 

Based on the results of the QA tests, aU of the G/E fabricated laminates and tubes were considered to be satisfactory. 

6.2.3 BORON/POLYIMIDE. Using the same tecuque described for the B/E material, all B/PI 

fabricated laminates were ultrasonic C-scan tested for delaminations and voids. All panels 

were judged to be acceptable based on this evaluation. 

The results of the remllining QA tests on the B/PI laminates are presented in Table 6-11 for 

resin and fiber contents and specific gravity values and Table 6-12 for room and elevated temp 

erature flexural and room temperature short beam shear strengths. 

In addition to the BlPI panels, five BlPI 8 ply unidirectional tubes were fabricated for the 

baseline shear tests. Resin and fiber contents and specific gravity determinations were made 

on two of the tubes. The other three tubes were not tested as a result of the decision to drop the 

B/PI system born the program because of the very low tensile properties and the extensive 

thermal degradation observed during thermal aging and short-term flight simulation testing. 

Data for the two tubes that were tested are listed in Table 6-13. 

Based on the results of the quality assurance tests, all of the B/PI fabricated laminates and 

tubes were considered to be satisfactory. 

6.2.4 GRAPHlTE/POLYIMIDE 

6.2.4.1 HTW7lO. Using the same technique described for the B/E material, all HT-S/710 G/PI 

fabricated laminates were ultrasonic C-scan tested for delaminations and voids. All panels 

were judged to be acceptable based on this evaluation. 

The results of the remaining QA tests on the HT-S/710 GlPI laminates are presented in Table 

6-14 for resin and fiber contents and specific gravity values, Table 6-15 for room and elevated 

temperature flexural and room temperature short beam shear strengths (Phase I material 

only), and Table 6-5 for room and elevated temperatures tensile strengths (Phase I1 material 

only). 

Based on the results of the QA tests all of the HT-SI710 GlPI fabricated laminates were considered to be satisfactory. 

6-17 
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6-19 
Tablo 6-13. Quality Assumnca Data for WPI Tuba 

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Fib cOnt.nt Rosin ConTub No. (wt %) (wt %) Spulfk Gravity 

2 

3 

64.6 

66.6 

26 

24 

1.98 

2.02 

6.2.4.2 Celion 6000/LARC-160. Using the same technique described for the BIE material, all 

Celion 6000ILARG160 GIPI fabricated laminates were ultrasonic C-scan tested for delaminations and voids. All panels were judged to be acceptable based on this evaluation. 

The results of the remaining quality assurance tests on the Celion 6000/LARC-160 G/PI lamimates fabricated for the Phase I series of tests are pmsented in Table 6-16 for resin and fiber 

contents and specific gravity values, Table 6-17 for room and elevated temperature fled 

and short beam shear strengths, and Table 6-18 for mom and elevated temperature tensile 

strengths of the 6 ply panels. 

Specific gravity and fiber and resin contents for the two paneh fabricated for the thermal 

agingkompression study are listed in Table 6-19. Also included is the Tg value for the 6 ply 

panel. Short beam shear and flexural strength data for the 24 ply panel have been included in 

Table 6-17. Again, the metallographic examinations of the two panels are discussed in the thermal aging section of the report. 

Based on the results of the QA tests all of the Celion 6OOOLARC-160 GiPI fabricated laminates 

were considered to be satisfactory. 

6-20 
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6-25 
Tablo 6.18. Quality Arsumnco tonsilo Data for 6 Ply [OO f 45O], Colion 6000/LARC-160 C/PI 

Average 

111-4 450 

Average 

111-6 297 

111-4 297 75 643 

669 

506 

606 

680 

717 

745 

714 

758 

752 

703 

Average 

111-5 450 

Average 

111-7 297 

Average 

111-7 450 

Average 

111-8 297 

350 

75 

350 

75 

350 

75 

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738 

738 

73 1 

665 

711 

47.4 

527 

554 

518 

629 

540 

518 

562 

745 

710 

643 

Average 

I118 450 

Average 

111-9 297 

Average 

350 

75 

699 

724 

625 

595 

648 

703 

745 

667 

705 

6-26 

93.3 

97.1 

73.4 

87.9 

98.6 

- 

104 

108 

57 

52 

49 

53 

62 

74 

52 

104 

110 

109 

102 

107 

107 

106 

- 

- 96.5 

103 

68.8 

76.5 

80.4 

75.2 

91.3 

78.3 

75.2 

81.6 

- 

- 

108 

103 

- 93.2 

101 

105 

90.6 

86.3 

94.0 

- 

102 

108 

102 

- 96.7 

~ 

63 

57 

52 

54 

54 

73 

49 

61 

50 

62 

57 

56 

56 

60 

59 

- 

58 

60 

57 

53 

57 

52 

54 

- 

53 

63 

62 

66 

64 

8.3 

7.6 

7.1 

7.7 

9.0 

10.8 

7.5 

9.1 

8.3 

7.6 

7.9 

7.9 

10.6 

7.1 

8.8 

7.2 * 

9.0 

8.3 

8.2 

8.1 

8.7 

8.6 

- 

- 

- 

- 

- 

- 

8.5 

8.7 

8.3 

7.7 

8.2 

- 

7.6 I 

7.8 

7.7 

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9.0 

9.5 

9.2 

- 
Tablo 6-18. Quality Assurana Tonsilo Data for 6 Ply [OO 2 4501, Cellon 6OOO/LARC-160 C/PI - Condudod 

111-9 

AveraG 

111-10 

Average 

111-10 

Average 

111-11 

Average 

111-11 

Average 

450 350 678 

786 

738 

734 

297 75 745 

' 745 

752 

747 

450 350 848 

703 

710 

757 

297 75 

450 350 

643 

546 

629 

606 

731 

731 

690 

717 

98.4 

114 

107 

106 

108 

108 

109 

108 

- 

- 

123 

102 

103 

109 

- 

93.3 

79.2 

91.2 

87.9 

- 

1 0'6 

106 

100 

104 

- 

67 

59 

49 

58 

65 

59 

63 

9.7 

8.6 

7.1 

8.5 

9.4 

8.6 

9.2 

- 

62 

61 

52 

51 

55 

55 

60 

59 

58 

55 

54 

57 

9.1 

8.8 

7.6 

7.4 

7.9 

7.9 

8.7 

8.5 

8.4 

8.0 

7.9 

8.3 

- 

- 

55 8.1 

6-27 
6-28 
SECTION 7 

BASELINE TESTING 

The purpose of the baseline tests was threefold. First, these data served to characterize the 

composite materials and add to any existing data bases. Second, the baseline tests provided 

the scale and shape parameters necessary to define the statistical distribution of the ultimate 

tensile strengths for each of the material systems. These, in turn, were used to set the loads for 

the short-term tests, and with the short-term results, were used in a wearout analysis model to 

relate static and fatigue strengths. Finally, the baseline tests provided a rational starting 

point against which the varioua environmental effects may be measured. 

Tests that were conducted included ultimate tensile, tensile modulus, Poisson's ratio, notched 

tensile (K, = 2.43). transverse tensile (unidirectional laminates only), ultimate shear strength, 

shear modulus, and fracture. In addition to those tests that were originally planned, a limited 

number of compressive testa were also performed. These were added late in Phase I of the program because of indicatiom during thermal aging and flight simulation testing that matrixdominated failure modes were operating. Testing was performed over the temperature range 

from 218 K (-67 " F) to 450 K (350 O F) for the epoxy specimens, to 533 K 1500 O F) for the Celion 

6000/LARC-160 polyimide Specimens, to 616 K (650 OF) for the €3"-SI710 polyimide specimens, 

and to 700 K (800" F) for the B/Al specimens. Because of the large amount of data already 

available for the BIE system, only.limited testing was cqnducted on this material. Baseline 

testing of the B/PI system had not been completed when the mated was removed from the 

test program because of low tensile properties and extensive thermal degradation during thermal aging and short-term flight simulation testing. The HT-S/710 G/PI torsion tube tests were 

not run because of a lack of material with which to fabricate the test specimens. Some shear 

data were available at General Dynamics Convair, however, from a then current program 

evaluating the same HT-S/710 system. For the Celion 6OOOILARC-160 GIPI that was added 

after completion of Phase I, several changes were made to the baseline test program. As stated 

earlier, no unidirectional panela were fabricated for this material, and. consequently, all the 

unidirectional baseline tests were eliminated. Also eliminated were the fracture tests on the $0 O 

t 45'1, crossplied material, Based on the Phase I results it was thought that the limited 

amount of fracture data obtained would not be worth the rather high cost of specimen fabrication and testing. Two tests were added to the baseline evaluation of Celion 6000/LARC-160 

GIPI. These were compressive and interlaminar shear strength tests on specimens made by 

bonding together several pieces of the 6 ply [0 " 2 45 "Is laminates. Again, these tests were 

inciuded because of the matrixdominated failure modes observed during Phase I. 

The testing procedures used for this study were those that had been adopted by General 

Dynamics Convair and the composites industry at the time this contract was initiated. In 

general, there were no standards other than those called out in the Structural Design Guide for 

Advanced Composite Applications (ref. 7). 

All tests were conducted on calibrated test equipment. Temperature measurements were ah0 

made with calibrated thermocouples and recording devices. The General Dynamics Standards 

Laboratory utilizes a fully documented system that meets all the requirements of NASA and 

DOD. The system has reports to establish traceability to the National Bureau of Standards or 

equivalent. 

7- 1 
The basehe tests, which are discussed below, have been grouped into four categories: 1) tensile, 2) shear, 3) fracture, and 4) compressive. 

7.1 TENSlLE TESTING 

Included in this section are the unnotched and notched tensile tests, elastic modulus determinations, and the Poisson's ratio measurements. 

7.1.1 Sf€UMEN DESlGN AND TEST PROCEDURE. The tensile specimen configuration was rectangular with adhesively bonded end tabs to reduce stress concentrations at the gripping fixtures. 

Unidirectional and crossplied longitudinal specimens, both notched and unnotched, were pre 

pared from 6 ply laminates while the unidirectional transverse specimens were 6 ply for BlAl 

and 12 ply for the epoxy and polyimide systems. The widths and lengths were dependent on 

the type of composite, the test direction, and the configuration, Le., notched or unnotched. 

Details of the specimen geometry for the various materials and configurations are listed in 

Table 7-1. Based on results from numerous test programs conducted previously at General 

Dynamics, a width of 0.013 m (0.5 in.) was selected for the longitudinal tensile specimens of 

both the resin and metal matrix composites. The same width was suitable for the transverse 

BlAl specimens but was increased to 0.025 m (1.0 in.) for the transverse epoxy and polyimide 

specimens because of their low strength. The wider specimen reduced the likelihood of accidental failures during handling in the machining, doubler bonding, or testing operations. The 

notched configuration for the resin matrix specimens was the same as that used for the shortterm flight simulation specimens. It consisted of a 0.025 m (1.0 in.) wide specimen with a 

0.0064 m (0.25 in.) hole in the center. The theoretical stress concentration (KJ for this configuration is 2.43 (ref. 28). The original plan for the BlAl system called for a side notched 

specimen, 0.013 m (0.5 in.) in width; with a theoretical notch acuity (K3 of 6.3. After cutting the 

blanks but before machining the side notches, a decision was made to use a configuration 

similar to that chosen for the resin matrix specimens. To keep the holetespecimen width ratio 

the same, a 0.00318 m (0.125 in.) diameter hole was machined in the center of the B/Al 

specimens. 

The tensile specimens used on the program are shown in Figure 7-1 for the longitudinal resin 

matrix and metal matrix specimens and in Figure 7-2 for the transverse resin matrix 

specimens. The transverse B/Al specimen is identical in dimensions to the longitudinal 

specimen of Figure 7-1. The notched tensile configurations are shown in Figure 7-3 for the resin 

and metal matrix composites. 

Tensile testing was conducted in both Instron and Tinius-Olsen testing machines using 

Instron capstan grips with ball or universal joints to ensure axial loading of the specimens. A 

crosshead rate of 0.0025 m (0.01 in.) per minute was used for all tests. Strain was measured in 

two ways. For the specimens where Poisson's ratio was required, four strain gages were bonded 

to the test section. Elastic modulus values could also be determined from these specimens. A 

few of the G/PI and B/Al specimens were also strain gaged (two gages only) to determine 

modulus values at the higher temperatures where conventional extensometer measurements 

were not always reliable. The second method of strain measurement was with a clamp-on strain 

gage extensometer. The extensometer was used from 218 K (-6'7' F) to 561 K (550" F) with 

reasonably good results. At the higher temperatures, however, the bonded strain gage method 

was superior. An elevated temperature tensile specimen mounted in the capstan grips with a 

strain gage extensometer and thermocouple attached ready for test is shown in Figure 7-4. 

' 

7-2 
r) 

00- 

??9 

ro 

PI 

131 

9 

4 r) lq 

0 

rouaz 

999 

cvN 

00co 

7-3 
0.064 T (2.5) 

T 

t., 0.051 (2) + 0.051 (2) 

1. 

0.11 

0.06 

- 

c_( 0.013 (0.5) 

i 

0.013 (0.5) 

t 

1 

0.013 (0.5) 

OOUBLERS -4 

t 

PLACES 

DIMENSIONS IN m (IN.) 

.-__ . 

6532 17-1 7 

Flguro 7-1 Tonsilo Tort Specimon for (a) Rosin Matrix Comporitos and (b) Mota1 Matrix Comporitos 

7-4 
1 

0.064 

- 

O.l( 

0.064 (2.5 

I 

~ 

0.013 (0.5) 

-r 

I 

0.013 (0.5) 

t 

D~MENSIONS IN m (IN.) 

BONDED DOUBLERS - 4 PLACES 

65321 7-1 8 

Figuro 7.2 Tronsvooo Tonsil. Test Spocimen for Resin Matrix Compositor 

7-5 
0.064 (2.5) 

L 

0.025 (1)- 

0.013 (05)- 

0.1 14 

0.013 (0.5)- 

0.00318 (0.125) 

'0.0064 (0.25) 

DIA 

HOLE 

- 

I 

0.076 (3) 

I 

c\ BONDED DOUBLERS -4 PLACES 

DIMENSIONS IN rn (IN.) 

(b) 

65321 7-19 

Figun 7-3 Notchod Tonsilo Tort Spoeimon for (a) Resin Matrix Comporitos and (b) Metal Matrix 

Compos itos 

7-6 
ORIGINS PAGE IS 

OF POOR QUFigur8 7-4 Elevated Temperature tensile Test Setup 

Elevated temperature tests were conducted with either a Conrad-Missher chamber or a small 

ring furnace. The advantage of the ring furnace was that the grips were not directly heated and 

doubler bond failures were reduced. The ring furnace was particularly beneficial at temperatures above about 561 K (550 ' F). All the tests at 218 K (-67 ' F) were performed in the 

Conrad-Missimer chamber. Temperature was measured with a thermocouple clipped to the 

specimen such that the bead was in contact with the surface. Specimens were held at temperature for a minimum of 10 minutes prior to loading. 

For the specimens that were to be strain gaged the following procedure was used. Prior to 

application of the gages, the specimens were lightly sanded in the gage section and precoated 

with M-bond 900 adhesive (manufactured by Micro-Measurement) to fill any void areas in the 

surface. The precoat was cured at 366 K (200" F) for two hours, and the specimens were again 

lightly sanded. The strain gages were then installed on the clean surfaces. Poisson's ratio 

spechens were instrumented with four strain gages, one longitudinal and one transverse gage 

on each side. For determination of elastic modulus only two gages were required, one 

longitudinal gage on each side. 

Table 7-2 lists the type of gages (manufactured by BLH Electronics, Inc.), adhesives, and cure 

cycles selected for the different test specimens. 

7-7 
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rb 

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0 

0 

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PI 

0000 

ssss 

7-8 
Testing of the strain gaged specimens was conducted in a manner similar to that used for the 

specimens tested with a ciamp=on extensometer except for the method of loading. These specimens were loaded incrementally to failure with strain measurements recorded automatically 

with a B & F Instruments, Inc Muitichrrnnel Digital Strain Indicator, 161 Mini-System, at 

each load leveL The output was both a typewritten recording of load versus strain and a 

punched paper tape that could bet processed by a computer, programmed to produce a stressstrain plot. A typical computer generated plot is shown in Figure 7-5. Poisson's ratio and 

elastic modulus values were determined from these curves. Figure 7-6 shows the Mini-System 

in operation during a mom temperature tensile test. 

Prior to tensile testing of the resin matrix material, all elevated temperature specimens and 

those room temperature specimens that were not strain gaged were heated at 394 K (250" F) 

for at least 24 hours to remove any absorbed moisture. The strain-gaged room temperature 

specimens were baked out prior to application of the gages and then stored in a desiccator until 

test. 

7.1.2 TEST RESULTS AND DlSCLlSSlON 

7.1.2.1 BorodEpoxy. Because a large data base was already available for the B/E system, no 

baseline tensile testing was needed. However, for later comparisons with the residual strength 

data from the flight simulation exposum specimens it was desirable to have baseline unnotched 

and notched tensile data from the same batch of materiaL Therefore, room temperature 

unnotched and notched tensile tests were conducted on (0 O f 45 "Is crossply material using the 

same test conditions and specimen configurations as were used later in evaluating the BIE 

flight simulation specimen3 after 10,000 hours of exposure. The unnotched specimens were 

strain gaged so that Poisson's ratio values could also be determined. The results are listed in 

Table 7-3. A tensile had-strain diagram is shown in Figure 7-7 for one of the strain gaged 

specimens. The points plotted are the average of the two 1ongtitudina.I or transverse gages. 

Based on the few tests that were run, there was an approximately 30% decrease in the net tensile strength in (0 O f 45 "Is crossply BIE for the notched specimen configuration chosen. The 

Poisson's ratio values appear to be quite large (values for metals are generally about onethird) 

but are close to what is calculated hrn lamination theory for a crossplied Iayup. 

Also included in Table 7-3 are tensile data from the quality assurance evaluation of the Phase I1 

batch of BIE. The tensile strength of the Phase I material is approximately 15% higher than 

that of the Phase I1 material. The Phase I specimens were cut from an area of panel that was 

somewhat thin. A comparison of failure loads for the two sets of data shows less than a 5% difference. Additional data indicating that the Phase I baseline tensile strengths are not representative are the tensile data obtained from the stacking sequence fatigue study (Section 

11.3.4) and tensile data from short time thed aging exposures (Section 8.2.1). Both sets of 

results show good agreement with the Phase I1 quality assurance tensile data. 

7.1.2.2 GraphitdEpoxy. The initial plan for the Phase I GIE baseline tensile tests was to test at 

four temperatures. This was later changed to seven temperatures with three specimens each 

(total of 21 specimens). From the tensile strength versus temperature curve obtained, two of 

the temperatures would then be selected for an additional seven tests each (total of 14 specimens). It was planned to select these two temperatures in the region just below that where a 

distinct change in the slope of the strength versus temperature curve occurred and then to use 

these data to set the loads for the flight simulation exposures. For GIE, however, no fall-off in 

strength was observed out to the maximum test temperature of 450 K (350' F). Because of the 

, 

, 

7- 9 
125 

100 

75( 

N' 

E 

s 

2 

3 

a 

- 

w 

a 

50( 

2% 

0 

STRAIN (x 10-3) 

653217-21 

Figure 7-5 Typical Stress-Strain Curve for Tensile lest of [0°]6 G/E 

7-10 
ORPGINAL PAGE I5 

OE POOR QUALITY 

- 

- 93 

0 

0 

d 

a 

2 

Figure 7-6 Automatic Strain Recording Instrumentation for Strain Gaged Tensile Specimen 

SPEC. NO: AC01-3 

AREA: 9.42 X loe6 rn2 (0.0146 in2) 

MODULUS: 93.1 GNh2 (13.5 Msi) 

POISSON'S RATIO: 0.79 

I 1 1 

1.2 

1.0 

0.8 

6 

D 

0 

E 

'0, 

0.6 

0.4 

a2 

0 

-6 -4 -2 0 2 4 6 0 10 

STRAIN (x 10-31 6532 1 7 -23A 

Figure 7-7 Tensile load-Strain Diagram for [Oo f 4501, B/E at 297 K (75O F) 

7-1 1 
Tablo 74. Boullno Tonsilo Propod- of [Oo f 45O], B/E at Room tompomtun 

Spdnlen T.milo Strongih tonsilo Modulus Poisson's 

Numbu Configumtion MN/~ (kri) GWd (psi X I@) Raflo. 

Phase I Material 

ACO1-1 unnotched 596 86.4 96.5 a14.0 aO.76 

=13.5 * .79 

Average 619 89.7 94.2 13.7 .78 

-2 617 89.5 a93.1 a13.5 a .ao 

-3 - 643 -- 93.2 93.1 - 

AC02-1 Notched 

-2 

-3 

Average 

Phase I1 Material 

A- 1 unnotched 

-2 

-3 

Average 

a Strain gage measurement 

399 57.9 - 

476 69.0 - 

454 65.8 - - 

443 

- 

64.2 

539 78.2 79 11.4 - 

529 76.7 52 7.6 - 

524 76.0 81 - 

53 1 

-- 

77.0 71 

11.7 - 

10.2 

uncertainty in the maximum use temperature of GIE for very long times, the two additional 

tensile test temperatures were selected such as to cover a wide range of temperature. They 

were 394 K (250 O F) and 450 K (350 O F). As it later was found during the long-term flight simulation tests, both temperatures were above the msnimum use temperature of the epoxy 

systems for 10,000 hour senrice. 

Baseline tensile tests, both unnotched and notched. were performed on unidirectional and [O O 

2 45 OjS crossplied specimens in the longitudinal direction. Some unnotched tensile tests were 

also performed in the transverse direction on unidirectional material, and a limited amount of 

Poisson's ratio testing was conducted, Results are listed in Tables 7-4 through 7-9. Figure 7-8 

shows a typical example of a load-strain diagram obtained from one of the strain gaged 

Poisson's ratio specimens. The longitudinal unnotched and notched data are plotted as a function of temperature in Figures 7-9 and 7-10. A summary of all the G/E tensile properties is 

presented in Table 7-10. 

Modulus values are given for the majority of unnotched specimens. For some specimens, however, modulus values were not determined because of the change in the test plan discussed 

above. Specimens that had originally been scheduled for testing at room temperature had been 

prepared with doublers bonded on with low temperature adhesive. To test these samples at an 

elevated temperature, a small ring furnace that heated only the center section of the specimen 

was required. With this type of heater, the clamp-on extensometer could not be accomme 

dated. In addition to these specimens, a few others are shown without modulus values because 

of extensometer malfunctions. Several of the specimens were strain gaged in order to measure 

7-12 
Tablo 74. Bosolino Tmsik Propmr)ios of [0°]1 G/E 

Spoeimon Tomporaturr Tonsil. Strongih Tonsil. Modulus 

Numbor K (OF) MN/~ W) GN/W? (Msi) .I 

BUO1-1 

-2 

-3 

BU01-4 

-5 

-6 

BU01-7 

-8 

-9 

-22 

-23 

-24 

-25 

-26 

-27 

-28 

BUOl-10 

-11 

-12 

BU01-13 

-14 

-15 

BU01-16 

-17 

-18 

BUO1-19 

-20 

-21 

-29 

-30 

-31 

-32 

-33 

' -34 

-35 

218 

297 

394 

408 

422 

436 

450 

-67 917 

1050 

683 

avg 883 

75 1460 

1660 

- 

1540 

avg 1550 

- 

250 1650 

1440 

1490 

1460 

1560 

1520 

1350 

1630 

1450 

1570 

avg 1510 

- 

275 . $1660 

1740 

1610 

avg 1670 

- 

300 . 1540 

1520 

1450 

avg 1500 

- 

325 1600 

1660 

1600 

avg 1620 

- 

350 1630 

1630 

1550 

1580 

1630 

1630 

1590 

1310 

1630 

1680 

avg 1590 

- 

133 

153 

99 

211 

240 

223 

128 

225 

239 

209 

216 

212 

227 

221 

196 

236 

210 

228 

241 

252 

233 ' 

219 

242 

224 

221 

210 

218 

232 

240 

232 

235 

237 

237 

225 

229 

237 

236 

23 1 

190 

236 

243 

230 

- 

- 

- 

a143 

al 34 

al 45 

141 

a141 

a136 

a136 

161 

152 

141 

150 

144 

152 

146 

- 

- 

- 

- 

165 

166 

166 

a141 

a140 

a136 

139 

143 

141 

131 

138 

150 

150 

131 

154 

148 

154 

160 

144 

156 

155 

150 

- 

- 

- 

- 

- 

- 

- 

a20.8 

a19.4 

a21.0 

20.4 

a20.4 

a19.a 

a19.a 

23.3 

22.1 

20.4 

21.8 

20.9 

22.1 

21.2 

23.9 

24.1 

24.0 

a20.5 

a20.3 

a19.9 

20.2 

20.7 

20.5 

19.0 

20.1 

21.7 

21.8 

19.0 

22.3 

21.5 

22.4 

23.2 

20.9 

22.6 

22.5 

21.8 

- 

- 

- 

- 

- 

- 

- 

- 

a Strain gage measurement 

7-13 
Tablo 7-5. Basdino Tonsilo Pmportios of [OO f 45OIs G/E 

BCO1-1 

-2 

-3 

BC01-6 

-7 

-8 

BC01-4 

-5 

-16 

-22 

-23 

-24 

-25 

-26 

-27 

-28 

BCOl-9 

-10 

-17 

BCO1-11 

-18 

-19 

BCO1-12 

-13 

-20 

BC01-14 

-15 

-21 

-29 

-30 

-3 1 

-32 

-33 

-34 

-35 

218 

297 

394 

408 

422 

436 

450 

-67 

75. 

250 

275 

300 

325 

350 

416 

439 

493 

avg 449 

578 

447 

463 

avg 496 

478 

505 

443 

524 

518 

623 

596 

54 1 

521 

509 

avg 526 

494 

407 

552 

avg 484 

569 

460 

497 

avg 509 

508 

5 20 

477 

avg 502 

457 

522 

499 

514 

499 

430 

470 

558 

506 

544 

avg 500 

- 

- 

- 

- 

- 

- 

- 

60.4 

63.6 

71.5 

65.2 

83.8 

64.8 

67.2 

71.9 

69.4 

73.2 

64.2 

76.0 

75.2 

90.3 

86.4 

78.4 

75.5 

73.8 

76.2 

71.6 

59.0 

80.1 

70.2 

82.5 

66.7 

72.1 

73.8 

73.7 

75.4 

69.2 

72.8 

66.3 

75.7 

72.4 

74.6 

72.4 

62.4 

68.1 

81.0 

73.4 

78.9 

72.5 

- 

- 

- 

- 

- 

- 

- 

49 

43 

50 

47 

a61 

a54 

a56 

57 

- 

- 

- 

55 

46 

43 

48 

65 

74 

62 

56 

- 

- 

- 

54 

54 

48 

54 

51 

- 

- 

- 

- 

- 

53 - 

53 

- 

- 

46 

78 

70 

52 

47 

63 

53 

57 - 

58 

7.1 

6.2 

7.3 

6.9 

a 8.8 

a 7.9 

a 8.1 

8.3 

- 

- 

- 

8.0 

6.6 

6.2 

7.0 

9.4 

10.8 

9.0 

8.1 

- 

- 

- 

- 

- 7.8 

7.8 

- 

6.9 

7.8 - 

7.4 

- 

- 

7.7 - 

7.7 

- 

- 

6.6 

11.3 

10.1 

7.5 

6.8 

9.2 

7.7 

8.2 - 

8.4 

a Strain gage measurement 

7-14 
Tablo 7-6. Basolino Notehod Tonsilo ProporHor of [VI6 WE 

BU02-1 

-2 - 

-3 

BU02-4 

-5 

-6 

BU02-7 

-8 

-9 

-22 

-23 

-24 

-25 

-26 

-27 

-28 

BU02-10 

-11 

-12 

BU02-13 

-14 

-15 

BU02-16 

-17 

-18 

BU02-19 

-20 

-21 

-29 

-30 

-31 

-32 

-33 

-34 

-35 

218 

297 

394 

-67 

75 

250 

408 275 

422 300 

436 325 

450 350 

1100 

938 

1060 

avg 1030 

1340 

1310 

1490 

avg 1380 

1390 

1210 

1200 

1380 

1210 

1560 

1250 

1570 

1630 

1510 

avg 1390 

1640 

1630 

1590 

avg 1620 

1610 

1480 

1270 

avg 1450 

1400 

1320 

1510 

avg 1410 

bl 170 

1240 

1260 

b1160 

1340 

b1180 

b1260 

1590 

1340 

b 910 

avg 1350 

159 

136 

154 

150 

194 

190 

- 

216 - 

200 

202 

176 

174 

200 

f76 

226 

181 

228 

236 

219 

202 

238 

237- 

230 

235 

234 

214 

184 

21 1 

203 

192 

- 

- 

- 

219 - 

205 

b170 

180 

192 

bi6a 

195 

b171 

b183 

230 

194 

b132 - 

196 

a Net section strength, Le., based on total width less hole diameter. 

b Doubier failure. values not included in average. 

7-15 
tablo 7.7. Basollm N&hd Tonsilo Pr0portii.r of [OO f: 45O], G/E 

BC02-1 

-2 

-3 

BC02-4 

-5 

-6 

BC02-7 

-8 

-9 

-22 

-23 

-24 

-25 

-26 

-27 

-28 

BC02-10 

-11 

-12 

BC02-13 

-14 

-15 

BC02-16 

-17 

-18 

BC02-19 

-20 

-21 

-29 

-30 

-31 

-32 

-33 

-34 

-35 

218 -67 317 

309 

311 

avg 312 

297 

394 

408 

422 

436 

450 

75 

250 

275 

300 

325 

350 

367 

477 

391 

avg 412 

46 1 

398 

536 

324 

360 

429 

443 

385 

319 

358 

avg 401 

448 

367 

377 

avg 397 

456 

372 

353 

avg 394 

429 

425 

436 

avg 430 

445 

396 

553 

419 

334 

403 

396 

445 

398 

369 

avg 416 

46.0 

44.8 

45.1 

45.3 

53.2 

69.2 

56.7 

59.7 

66.9 

57.7 

77.8 

47.0 

52.2 

62.2 

64.3 

55.8 

46.2 

51.9 

58.2 

65.0 

53.3 

54.7 

57.7 

66.2 

54.0 

51.2 

57.1 

62.2 

61.6 

63.2 

62.3 

64.6 

57.5 

80.2 

60.8 

48.4 

58.4 

57.4 

64.5 

57.7 

53.5 

60.3 

- 

- 

- 

- 

- 

- 

- 

a Net section strength, Le., based on total width less hole diameter. 

7-16 
tablo 7-8. Baulino Tonsilo Propodios of [OO],2 G/E T.rfod in tho tranrvono Direction 

Sp.cimmn T.mpUUtUm Edgo Spajmon Wfdth tonsilo Strongth 

Numbor K (OF) Condition m (in.) MN/~ (kri) 

TBuo1-1 

-2 

-3 

-4 

-5 

-6 

TBUO1-A 

-3 

-c 

-1 1 

-12 

TBUO1-7 

-8 

-9 

010 

297 75 Poor 0.0025 

.0025 

.0025 

,0025 

.0025 

.0025 

297 75 Good .0013 

.0019 

.0019 

.0020 

.0019 

450 350 Good .0016 

,0019 

.0019 

.0020 

1.00 29 

1.00 40 

.99 27 

1.00 38 

1-00 38 

32 

avg 

- 

34 

1.00 

.50 54 

.75 59 

.75 64 

.80 57 

62 

avg 

- 

59 

.75 

.62 39 

.75 37 

.75 32 

.80 41 

avg 

- 

37 

4.2 

5.8 

3.9 

5.5 

5.5 

4.6 - 

4.9 

7.8 

8.6 

9.3 

8.2 

9.0 - 

8.6 

5.6 

5.4 

4.7 

6.0 - 

5.4 

tablo 7-9. Baulino Poisson's Ratio Data in Tonrion of G/E 

Tort 

Spuimon Lamina). fost Tompemturo Poisson's 

Numbor Oriontation Dirdon K (OF) Ratio 

BU01-4 

-5 

-6 

BUO1-7 

-8 

-9 

BU01-13 

-14 

-15 

BC01-6 

-7 

-8 

[o 0i6 Longitudinal 297 75 0.35 

.33 3 

.32 

avg .33 

[w6 Longitudinal 394 250 .30 

.30 

.37 

avg .34 

KJ 01~ Longitudinal 422 300 .39 

.32 

.35 

avg .35 

[Oo 2 45"], Longitudinal 297 75 .71 

.70 

.70 

avg .70 

7-17 
1.0 

aa 

r 

0 

0 

0.6 > 

E 

0, 

I 

I I 

SPEC. NO: BCO1-6 

AREA: 874 x m2 (a01355 in2) 

. MODULUS: 6117 GN/m2 (88 Msi) 

POISSON'S RATIO: 0.71 

" 

-6 -2 0 2 4 6 8 10 

STRAIN (x 10-3) 

6S3217-24 

Figuro 7.8 Tonsilo Load-Strain Diagram for [Oo & 45O],, C/E ot 297 K (750 F) 

TEMPERATURE (OF) 

1750 ru- 

\ E 240 g 

3 

=! 

200 m 

- 2 1500 x 

I 

c 

0 

W 

+ 

m 

z 

m 

z 1250 

a 

k 160 

5 1000 

W 120 z" 

3 500 

v) 

-4 

z 

c) + 

z 

c 750 

c I 

5 

z 

W 

U 

80 - a 

250 40 

0 0 

TEMPERATURE (K) 

200 250 300 350 400 450 

65321 7-21 

Figuro 7-9 Barelino Tensilo Propertks of G/E 

7-18 
TEMPERATURE (OF) 

-100 0 too 200 300 400 

1750 

240 6iE 

\ 

- I 1500 

200 I 

la 

5 1250 

ta 

y 1000 

1 60 a 

z 

+ 150 

1 20 2 

W 

W 

t 

a 

80 500 3 

3 

4a 250 

0 200 '0 250 300 350 400 450 

TEMPERATURE (KI 

-I 

m ... 

2 

en - 

m 

v) 

4 

a 

m 

z 

66321 7-26 

Figun 7-10 Basdin. Notehod Tonsilo Propodes of C/E 

Poisson's ratio. The modulus values measured on these specimens are noted in the tables: In 

general, these values are more reliabie than those determined with the mechanical 

extensometer. 

The effect of temperature on the tensile strength of G/E is shown on Figures 7-9 and 7-10. With 

the exception of the 218 K (-67" I?) results, the tensile strength exhibited almost no change 

with increasing temperature out to the maximum test temperature of 450 K (350 F). At 218 K 

(-67" F) the strength showed a moderate decrease for the crossplied material and a rather 

large one for the unidirectional material (almost 50%) for the unnotched specimens. This 

strength decrease at low temperatures is not uncommon for resin matrix composites. The 

effect is believed to be related to residual stresses in the material, which increase as the difference between the cue temperature and the test temperature is increased. Also shown on 

Figure 7-9 and 7-10 are the scatter bands for the various tests. For both unnotched and 

notched specimens the data scatter was numerically greater for the unidirectional material, 

but, percentagewise, the crossplied material showed a significantly greater degree of scatter. 

Because of the larger number of specimens tested at 394 K (250" F) and 450 K (350' F), the 

scatter bands at these temperature tended to be somewhat wider. 

In like manner to the tensile strength data, the tensile modulus was also not significantly 

affected by increasing temperature out to the maximum test temperatures of 450 K (350" F). 

At 218 K (-67 O F) the average value for the crossplied material was slightly lower than those 

obtained at higher temperatures, but all of the 218K (-67 " F) data fell within the scatter band 

of the tests conducted at 394 K (250O F). No modulus values were obtained from the 218 K 

(-67" F) tests on the unidirectional specimens because of extensometer malfunctioning. 

I 

7-19 
e4 

E 

t 

00- 

mom 

arb 

m 

Amea 

ON 

-0 

(0113 

m 

sp! 

8C I 

.- e 

2 

I 

I 

rd 

2 

I 

S 

3 

mo 

m0 

I 

i 

M 

K 

7-20 
A comparison of the ratio of the notched to the unnotched tensile strength of GIE at each of 

the test temperatures has been made in Table 7-1 1. For the unidirectional material the effect of 

a notch on the net section strength is quite small The effect of the notch was frequently seen 

during testing, particularly fatigue testing, in that longitudinal cracks propagated tangent to 

the edges of the notch (center hole) along the entire length of the test section. Examples of the 

failure mode of the uniktional notched specimens can be seen in Figure 7-11. Figure 7-12 is 

a doseup view of the bottom specimen showing the longitudinal splitting that has occurred 

tang& to the 0.0064 m (0.25 in.) hole. The net effect is to divide the specimen into two smaller 

parallel specimens. In a tensile test the two strips act as one and very little reduction in the net 

tensile strength was observed. During fatigue testing, however, as will be shown in Section 11, 

the notch was observed to either increase or decrease the fatigue life compared to unnotched 

specimens depending on the test conditions. For crossplied material the notched specimens 

gave a tensile strength that was 10 to 30% lower than the unnotched specimens. Similar effects of a center hole notch were observed for the HT-SI710 GIPI system in both unidirectional 

and crossplied specimens. No unidirectional specimens of the Celion 60001LARC-160 G/PI 

system were tested, but the notched crossplied specimens were also significantly lower in tensile 

strength than the unnotched specimens. 

Tensile data for tbe unidirectional material tested in the transverse direction are listed in Table 

7-8. The plan called for six specimens each to be tested at mom and one elevated temperature. 

The specimens, which were taken from the first GE laminate that was fabricated, were intended 

to be used as a quality assurance check on the material as well as for baseline tensile property 

determinations. The mng blade was used on the table saw when cutting the specimens. Six of 

these specimens were tested at room temperature, giving the low values shown in Table 7-8. 

An examination of the saw cut edges of the fded specimens revealed a very poor edge condition and one at which premature failure could readily be initiated. The remaining six untested 

specimens and three spares that were available were salvaged by trimrmn * g the chipped edges 

with the correct blade. The two types of edge conditions are shown in Figure 7-13 and 7-14. The 

mom temperature tensile strength of the somewhat naz~~w specimens with good edges was 

nearly double that of those with the badly damaged edges. Unlike longitudinal tensile properties, the transverse strength showed a considerable decrease at 450 K (350 O F). The transverse 

strength is a matrix dominated property and would be expected to be more temperature dependent that the fiber dominated longitudinal strength. 

The remaining baseline tensile data for the Phase I GIE, Poisson’s ratio values, are listed in 

Table 7-9. For the unidirectional material values of approximately onethird (similar to that 

obtained for most metals) were obtained at all three temperatures. Room temperature 

measurements on crossplied specimens gave values that were similar to those determined for 

the BIE system. 

Tensile data at room and 450 K (35OOF) were obtained for the Phase I1 batch of GIE during 

quality assurance testing (Table 6-5). The data, however, did not show good agreement with 

the residual strength data (Phase I1 material) obtained after long term flight simulation exposure. For this reason and because extra material was available, ten additional tensile tests were 

conducted at each temperature. The results (including the QA data) are listed in Table 7-12. 

The Phase I1 batch of GIE shows significantly higher tensile strength values than the original 

material purchased in Phase I. However, as wiU be seen in Section 12, the room temperature 

residual strength of the Phase I1 material after flight simulation exposure was considerabiy 

higher than the baseline data of Table 7-12. 

7-21 
Tablo 7-11. Ratio of Notched to Unnotched Tensile Strength of G/E at Various Temperatures 

Temperature 

K (OF) 

Ratio of Notched to Unnotched Strength 

[0oi6 [OO f 4501, 

218 -67 

297 75 

394 250 

408 275 

422 300 

436 325 

450 350 

1.2 

.9 

.9 

1.0 

1.0 

.9 

.9 

0.7 

.8 

.8 

.8 

-8 

.9 

.8 

Figure 7-1 1 Unidirectional G/E Notched Tensile Specimens 

7.1.2.3 Graphite/Polyimide 

HT-S/710 G/PI. The test plan for the HT-S/710 GlPI baseline tensile specimens was similar to 

that used for the GIE specimens in that three tests were performed at each of seven temperatures. It differed slightly, however, in that all remaining specimens were then tested at one of 

the seven original temperatures rather than at two. This temperature was the one that had 

been selected for the flight simulation exposure tests, 505 K (450 O F). Testing at one temperature was done to increase the tensile strength data base that was required for selection of the 

short-term flight simulation load levels. At the time these tests were performed very little data 

was available for the HT-S/710 system. Selection of the temperatures for the flight simulation 

tests is discussed in Section 12. 

7-22 ORIGINAL PAGE IS 

OF POOR QUALITY 
Figure 7-12 Closeup View of Unidirekfional G/E Notchred Tensile Specimens Showing longitudinal 

Splitting Tangent to the Center Hole 

7-29 

..- 

Figure 7-13 Transverse Tensile Specimens of G/E Showing Poor Edge Condition of Improperly Machined 

Specimen (upper) and Good Edge Condition of Properly Machined Specimen (lower) 

7-23 
Figure 7-14 Transverse tensile Specimens of G/E Showing Poor Edge Condition of Improperly Machined 

Specimen (upper) and Good Edge Condition of Properly Machined Specimen (lower) 

Baseline tensile tests, both unnotched and notched, were performed on unidirectional and [0 O 

2 45 crossplied specimens in the longitudinal direction. Some unnotched tensile tests were 

also performed in the trahsverse direction on unidirectional material, and a limited amount of 

Poisson's ratio testing was conducted. Results are listed in Tables 7-13 through 7-18. A typical 

load-strain diagram from one of the strain gaged Poisson's ratio specimens is shown in Figure 

7-15. The longitudinal unnotched and notched data are plotted as a function of temperature in 

Figure 7-16 and 7-17. A summary of all the HT-SI710 GlPI tensile properties is presented in 

Table 7-19. 

Modulus values were determined at four temperatures for the specimens tested in the longitudinal direction. A mechanical extensometer was used for the 218 K (-67 O F) tests while strain 

gages were used at room temperature and the two elevated temperatures. Poisson's ratio 

values were also determined for some of these specimens. For the material tested in the transverse direction, modulus values were obtained at room temperature only. A mechanical extensometer was used for these tests. 

The temperature dependence of the HT-SI710 GlPI tensile strength is shown in Figures 7-16 

and 7-17. At the maximum test temperature, 616 K (650 O F), the unidirectional unnotched and 

notched specimens had lost approximately 10% of the room temperature strength. The loss for 

the crossplied specimens was somewhat higher, 37% for the unnotched and 30% for the notched 

specimens. Because of the limited amount of tests and the scatter in the data, the temperature 

at which the tensile strength begins to decrease is difficult to determine precisely. The curves 

indicate this decrease to occur between 561 K (550" F) and 616 K (650' F). Some decrease in 

strength was also observed at 218 K (-67 O F) but the effect was generally less than that found 

for the GIE system. The spread in the HT-SI710 GlPI tensile data as shown by the scatter 

bands in Figures 7-16 and 7-17 was similar to that observed for GIE. The tensile modulus for 

O2JGINXL PAGE IS 

,olE HIOR QUALITY 

7-24 
Tablo 7-12. Barolino Tmrilo Proportior of [OO f 4501, G/E (Pharo II Moforial) 

B-1 297 

-2 

-3 

-17 

-18 

-19 

-20 

-21 

-22 

-23 

-24 

-25 

-26 

B-4 450 

-5 

-6 

-27 

-28 

.-29 

-30 

-3 1 

-32 

-33 

-34 

-35 

-36 

75 456 

459 

556 

487 

590 

572 

509 

604 

570 

554 

570 

538 

681 

550 

350 557 

584 

623 

578 

626 

632 

664 

614 

629 

647 

666 

689 

621 

568 

66.1 

66.5 

80.6 

70.7 

85.5 

82.9 

73.8 

87.6 

82.7 

80.4 

82.6 

78.1 

98.7 

79.7 

- 

80.8 

84.7 

90.4 

83.8 

90.8 

91.6 

82.4 

96.3 

89.1 

91.3 

93.8 

96.6 

90.1 

- 100 

50 

51 

53 

48 

53 

55 

50 

56 

54 

48 

59 

57 

57 

53 

47 

47 

50 

56 

52 

54 

48 

56 

50 

53 

52 

56 

57 

52 

7.2 

7.4 

7.7 

6.9 

7.7 

8.0 

7.3 

8.1 

7.9 

7.0 

8.5 

8.2 

8.2 

7.7 

6.8 

6.8 

7.3 

8.1 

7.6 

7.8 

6.9 

8.1 

7.2 

7.7 

7.6 

8.1 

8.3 

7.5 

- 

7-25 
toblo 7-13. Basdino tonsilo ProporHos of [PI, HT-S/710 G/PI 

a Strain gage measurement 

DUO1-1 

-2 

-3 

DU01-4 

66 

-6 

DU01-7 

-8 

-9 

-22 

-23 

-24 

-25 

DUO 1- 10 

-11 

-12 

DU01-13 

-14 

-15 

DU01-16 

-17 

-18 

DU01-19 

-20 

-21 

218 

297 

505 

533 

561 

589 

616 

-67 

75 

450 

500 

550 

600 

650 

1300 

1180 

1100 

avg 1190 

1260 

1190 

1120 

avg 1190 

1090 

1260 

1270 

1360 

1260 

1130 

1160 

avg 1220 

1270 

1260 

1210 

avg 1250 

1370 

1400 

1190 

avg 1320 

1150 

1200 

1240 

avg 1200 

1080 

1120 

1060 

avg 1090 

- 

- 

- 

- 

- 

- 

- 

188 

171 

160 

173 

182 

172 

163 

172 

158 

182 

184 

197 

182 

164 

168 

176 

184 

182 

176 

- 

- 

- 

- iai 

199 

203 

173 

192 

167 

174 

180 

174 

156 

162 

154 

157 

- 

- 

- 

144 

97 

134 

125 

a127 

a126 

a126 

126 

a121 

a128 

a126 

- 

- 

- 

- 

- - 

125 

- 

- 

- 

a141 

al 33 

a128 - 

134 

- 

- 

- 

- 

- 

- 

20.9 

14.1 

- 19.4 

18.1 

a18.4 

a18.2 

a18.3 

a17.6 

a18.6 

a18.3 

- 

18.3 

- 

- 

- 

- - 

18.2 

- 

- 

- 

7-26 
Tablo 7-14. Basolfno Tmsilo Proportier of [OO f 4501, HT-S/7lO G/PI 

Spoeimon Tempomtun Tonsilo Strength Tonsilo Modulus 

Numbor K (OF) MN/& (W GN/d (Msi) 

DCOI-I 

-2 

-3 

DC01-4 

-5 

-6 

DC01-7 

-8 

-9 

-22 

-23 

-24 

-25 

-26 

-27 

-28 

-29 

-30 

-3 1 

-32 

-33 

-34 

-35 

DCO1-10 

-11 

-12 

DC01-13 

-14 

-15 

DC01-16 

-17 

-18 

DC01-19 

-20 

-21 

218 

297 

505 

533 

561 

589 

616 

-67 387 

417 

411 

avg 

- 

405 

510 

505 

75 

450 

543 

avg 

- 

519 

423 

299 

441 

478 

' 525 

478 

437 

463 

496 

496 

485 

478 

. 560 

432 

512 

543 

449 

avg 

- 

470 

500 398 

410 

385 

avg 

- 

398 

550 257 

30 1 

463 

avg 

- 

340 

600 352 

432 

409 

avg 398 

650 305 

330 

- 

345 

avIt 

- 

327 

a Strain gage measurement 

56.1 

60.5. 

59.6 

58.7 

74.0 

73.2 

78.7 

75.3 

61.3 

43.4 

63.9 

69.3 

76.2 

69.4 

63.4 

67.1 

71.9 

71.9 

70.3 

69.3 

81.2 

62.7 

74.3 

78.8 

65.1 

68.2 

57.7 

59.5 

55.9 

57.7 

37.5 

43.7 

67.2 

49.4 

51.1 

62.6 

59.3 

57.7 

44.3 

47.9 

50.0 

47.4 

- 

- 

- 

- 

- 

- 

- 

7-27 
Tabio 7-15. Bamlino Notehod tonsilo Proportior of [0OL HT-SnlO G/PI 

DU02-1 

-2 

-3 

DU02-4 

-5 

-6 

DU02-7 

-8 

-9 

-22 

-24 

-25 

-26 

-27 

-28 

-29 

-30 

-31 

-32 

-33 

-34 

-35 

DUO2-10 

-11 

-12 

DUO2-13 

-14 

-15 

e 43 

DU02-16 

-17 

-18 

DU02-19 

-20 

-21 

218 -67 1180 

820 

972 

avg 991 

297 

505 

75 

450 

1310 

1140 

1320 

avg 1260 

1230 

1160 

1340 

986 

1250 

1040 

1110 

1130 

1060 

1080 

1120 

1210 

1240 

1190 

1140 

1070 

1050 . 

avg 1140 

533 500 1240 

1240 

1140 

avg 1210 

561 550 

589 600 

616 650 

1040 

1120 

1280 

avg 1150 

979 

1060 

1030 

avg 1020 

1130 

1140 

1050 

avg 1110 

a Net section strength, Le., based on total width less hole diameter. 

171 

119 

141 

144 

190 

166 

192 

183 

178 

169 

195 

143 

181 

151 

161 

164 

154 

157 

163 

175 

180 

173 

166 

155 

152 

166 

180 

166 

175 

151 

162 

186 

166 

142 

154 

150 

149 

164 

166 

153 

161 

- 

- 

- 

1 ao 

- 

- 

- 

- 
Tablo 7-16. Bardin. Notehod Tonsil. Propodos of [OO f 45O], HT-SnlO G/PI 

218 -67 328 47.5 

288 41.8 DCO2-1 

-2 

-3 

DC02-4 

-5 

-6 

DC02-7 

-8 

-9 

-22 

-23 

-24 

-25 

-26 

-27 

-28 

-29 

-30 

-31 

-32 

-33 

-34 

-35 

DC02-10 

-11 

-12 

DC02-13 

-14 

-15 

DC02-16 

-17 

-18 

DC02-19 

-20 

-21 

297 75 

505 450 

533 500 

589 

616 

561 550 

600 

650 

339 

avg 318 

367 

332 

372 

avg 357 

343 

352 

330 

334 

350 

350 

301 

332 

386 

354 

305 

373 

333 

363 

303 

292 

314 

avg 336 

361 

386 

308 

avg 352 

250 

355 

339 

avg 315 

242 

232 

27 2 

avg 249 

22 1 

262 

270 

avg 251 

49.1 

46.1 

53.3 

48.2 

54.0 

51.8 

49.7 

51.0 

47.8 

48.4 

50.8 

50.7 

43.7 

48.2 

56.0 

51.4 

44.2 

54.1 

48.3 

52.7 

43.9 

42.4 

45.5 

48.8 

52.3 

56.0 

44.7 

51.0 

36.3 

51.5 

49.2 

45.7 

35.1 

33.6 

39.4 

36.1 

32.0 

38.0 

39.2 

36.4 

- 

- 

- 

- 

- 

- 

- 

a Net section strength, Le., based on total width less hole diameter. 

7-29 
Tablo 7-17. Basdino tonsilo Propodor of [PI12 Ht-S/7lO G/PI tatod in tho Tmnnono Diractfon 

sw- Tomp~ro T.milo Strongdl tonsilo Modulus 

Numbor K (OF) MN/d (ksi) GN/d (W 

TDuo1-1 

-2 

-3 

-4 

-5 

-6 

TDUOl-7 

-8 

-9 

-10 

-11 

-12 

297 75 

505 450 

a Failed in handling 

21 

31 

21 

23 

20 

25 

24 

- 

(a) 

22 

(a) 

21 

28 

19 

22 

- 

3.0 

4.5 

3.1 

3.4 

2.9 

3.6 

3.4 

(a) 

(a) 

3.2 

3.1 

4.0 

2.8 

3.3 

5.2 

6.3 

4.8 

5.2 

4.8 

4.8 

5.2 

- 

- 

- 

- 

- 

- 

- 

0.75 

.91 

.69 

.76 

.69 

.69 - 

.73 

- 

- 

- 

- 

- 

tablo 7.10. Basolim Poisson's Ratio Data in tonsion of HTW710 G/PI 

T& 

Spocimon Laminato tat tompomtun Poisson's 

Numbor Oriontatton Dlnction K (OF) Ratio 

DU01-7 

-9 

-a 

DU01-13 

-14 

-15 

DC01-4 

-5 

-6 

Longitudinal 505 450 

Longitudinal 561 

[0° c 457, Longitudinal 297 

550 

75 

0.38 

.35 

.45 

.39 

SO 

.45 

.44 

avg .46 

.71 

.65 

.64 

avg .67 

7-30 
U 

- 0.8 

MODULUS: 128 GNh2 (183 Msi) = (14 

POISSON'S RATIO: a45 

-2000 -1000 

653217-31 

MICRO STRAIN 

Figun 7-15 tensilo Load-Struin Diagmm for [00]6 Hl-S/710 G/Pl at 500 K (&O0 F) 

1500 

c 

(1 

a 

z 1000 

750 

z 

w 

c 

LU 

I- 2 500 - 

3 

a 

250 

0 

65321 7-32 

Figuro 7-16 Basolino Tensile Properties of HT-S/710 G/PI 

7-3 1 
TEMPERATURE (OF) 

1500 

Si- - E 

1250 - 

I + 

c1 

z 1000 

z 

a 

w 

A 

z 

w 

IW 

IZ 750 

s 500 

5 

a 

250 

0 

0 0 100 200 300 400 500 700 

20 0 

160 

120 

ao 

40 

0 

200 250 300 

TEMPERATURE (K) 

Rgun 7-17 Bordino Notch4 Tonsilo Pr0por)ios of HT-S/710 G/PI 

the unidirectional material was not lowered by increasing temperatures out to the maximum 

test temperature of 561 K (550" F). The crossplied material did show a decrease in modulus, 

however, of about 15% at 561 K (550" F) and nearly 30% at 218 K (-67" F). 

A comparison of the ratio of the notched to unnotched tensile strength of HT-SI710 G/PI at 

each of the test temperatures has been made in Table 7-20. The results are very similar to those 

found for the GIE system. For the unidirectional material the effect of a center hole notch was 

quite small except at 218 K (-67 " F). The crossplied material, on the other hand, experienced 

decreases in tensile strength of 10 to 40% when tested in the notched configuration. 

Transverse tensile data are presented in Table 7-17 for room and elevated temperatures. 

Strength values were considerably lower than those presented earlier for the G/E system but 

were about 50% higher than values reported by others (ref. 24). The modulus values, however, 

. were considerably lower than those reported in reference 24. No significant effect of temperature on the transverse strength was observed out to 505 K (450" F). Evidence of the low 

transverse strength of this material is shown by the two accidental specimen failures that 

occurred when installing the specimens in the tensile machine grips. Only slight flexing of the 

specimens will result in fracture. 

Poisson's ratio values, Table 7-18, for unidirectional HT-SI710 G/PI were somewhat larger 

than those obtained for GIE. An increase in value with increase in temperature was also 

observed. In like manner to the BIE and G/E data, the room temperature Poisson's ratio for 

crossplied material was also quite large. 

7-32 
- 

0 * 

O 

L 

E 

E 3 

Y) 

1 

r; 

Y) 

0 

Y) 

A 

YI 

h 

h 

01 

w 

Y 

CI 

n 

I 

9 

Y 

z 

(* 

i 

a 0 .- 

Y 

2 

01 

C 

vl .- 

0 

01 

I 

I 

81 

0 

- QYfl 

00 

I00 

7-33 
tablo 7.20. Rdo of Notchd to Unnotchod Tonrib Strongth of Hf-S/7lO G/Pl at Various tompomtunr 

tompomturo Rdo of Notehod to Unnotehd Strongth 

K (OF) [OO f 4501, 

218 -67 0.8 

297 75 1.1 

505 450 .9 

533 500 1.0 

561 550 .9 

0.8 

.7 

.7 

.9 

.9 

589 600 .9 .6 

616 650 1.0 .8 

Celion 6000/LARC=160 G/PI. The plan for baseline tensile testing of the second G/PI system was 

to test three specimens at each of seven temperatures with the remaining 14 specimens equally 

divided between two of these temperatures. The temperatures selected for additional testing 

were room temperature and that used later for the flight simulation exposures, 450 K (350 OF). 

Selection of the temperatures for the flight simulation tests is discussed in Section 12. These 

two temperatures were the ones used for residual strength testing after both thermal aging 

and flight simulation exposures. 

Only [Oo 2 45'1, cmssplied tensile data were obtained for the Celion 60001LARC-160 GlPI as . 

no unidirectional panels were fabricated. Testing was performed on both unnotched and notched 

specimens in the longitudinal direction. Results are listed in Tables 7-21 and 7-22. A typical 

load-strain diagram from one of the strain gaged specimens is shown in Figure 7-18. The unnotched and notched tensile data are plotted as a function of temperature in Figures 7-19 and 

7-20. A summary of the Celion 60001LARC-160 G/PI tensile properties is presented in Table 

7-23. 

Modulus values were determined on the unnotched specimens at four temperatures using a 

mechanical extensometer. In addition, three each of the room temperature and 450 K (350 OF) 

specimens were also strain gaged for modulus measurements. As shown in Table 7-21, considerable scatter was observed in the modulus data obtained with the mechanical extensometer. 

The data from the strain gaged specimens, however, were much more consistent. They showed 

no effect of temperature on the modulus between room temperature and 450 K (350°F). 

The effect of temperature on unnotched and notched tensile strength of [0 f 45 OJS Celion 

6000/LARC-160 G/PI is presented in Figures 7-19 and 7-20. For both unnotched and notched 

specimens very little change in strength was observed over the entire range of temperatures 

studied. This is in contrast to the other organic matrix composite systems evaluated where 

decreases in strength at 218 K (-67 OF) were enerally found. The data scatter, as shown by 

the scatter bands in Figures 7-19 and 7-20, was somewhat less for this GlPI than for the 

HT-SI710 system. A comparison of the notched and unnotched tensile strengths of the ttvo 

G/PI systems shows the Celion 6000/LARC-160 material to be significantly higher at all temperatures. 

7-34 
Tabio 7-21. Basdin. Tonsiio Proportios of [Oo +45O], Colion 6000/LARC.160 G/PI 

~~ ~ 

Spoccimon Tomprmtun Tonsilo Strong)), Tonsil. Modulus 

Numbor K (OF) MN/& (ksi) GN/d (Msi) 

FCO1-1 218 -67 

-2 

-3 

FC014 297 75 

-5 

-6 

-7 

-8 

-9 

-10 

-11 

-12 

-13 

FC01-14 394 250 

-15 

-16 

FC01-17 450 350 

-18 

-19 

-20 

-21 

-22 

-23 

-24 

-25 

-26 

FC01-27 478 400 

-28 

-29 

FC01-30 505 450 

-3 1 

-32 

FC01-33 533 500 

-34 

-35 

a Strain gage measurement 

745 

738 

752 

avg 745 

758 

752 

703 

745 

703 

710 

674 

676 

68 1 

652 

avg 705 

745 

668 

800 

avg 738 

738 

73 1 

665 

703 

627 

638 

724 

738 

703 

690 

avg 696 

690 

696 

683 

avg 690 

703 

663 

703 

avg 690 

683 

717 

68 1 

avg 694 

- 

- 

- 

- 

- 

- 

- 

108 

107 

109 

108 

110 

109 

102 

108 

102 

103 

- 

97.7 

98.1 

98.7 

- 94.5 

102 

108 

116 

107 

107 

106 

102 

96.9 

- 

96.5 

90.9 

92.5 

105 

107 

102 

100 

101 

100 

101 

100 

102 

102 

100 

104 

101 

- 

- 99.1 

96.1 

- 

99.0 

- 96.8 

59 

60 

53 

57 

61 

62 

62 

52 

52 

44 

71 

74 

54 

81 

61 

66 

64 

66 

8.5 

8.7 

7.7 

8.3 

- 

a 8.8 

a 9.0 

a 9.0 

7.6 

7.5 

6.4 

10.3 

10.7 

7.9 

11.7 

8.9 * 

9.5 

9.3 

9.6 

65 

62 

63 

59 

55 

46 

54 

63 

63 

51 

63 

57 

9.5 

a 9.0 

a 9.2 

a 8.6 

8.0 

6.7 

7.8 

9.2 

9.2 

7.4 

9.2 

8.4 

- 

7-35 
Tablo 742. Baroilno Notehod Tonsilo Propdo8 of [OO fao], Caiion 6000/LARC=160 G/Pl 

Spufmon 

Numbor 

FC02-1 

-2 

-3 

FC024 

-6 

-6 

-7 

-8 

-9 

-10 

-11 

-12 

-13 

FC02-14 

-15 

-16 

FC02- 17 

-18 

-19 

-20 

-21 

-22 

-23 

-24 

-25 

-26 

FC02-27 

-28 

-29 

FC02-30 

-3 1 

-32 

FC02-33 

-34 

-35 

218 -67 

297 75 

394 

450 

250 

350 

478 400 

505 450 

533 500 

(a) Net section strength, Le., based on total width less hoie diameter. 

442 

40 1 

404 

avg 416 

381 

345 

396 

452 

487 

432 

500 

390 

419 

474 

avg 428 

458 

469 

395 

avg 441 

44 1 

481 

445 

468 

430 

483 

439 

517 

492 

554 

avg 475 

534 

414 

472 

avg 473 

555 

544 

500 

avg 533 

514 

564 

456 

avg 511 

64.1 

58.1 

58.6 

60.3 

55.2 

50.1 

57.4 

65.6 

70.7 

62.6 

72.5 

56.5 

60.8 

68.7 

62.0 

66.4 

68.0 

57.3 

63.9 

63.9 

69.8 

64.6 

67.9 

62.4 

70.0 

63.7 

75.0 

71.4 

80.3 

68.9 

77.5 

60.0 

68.5 

68.7 

80.5 

78.9 

72.5 

77.3 

74.5 

81.8 

66.1 

74.1 

- 

- 

- 

- 

- 

- 

7-36 
1 .o 

0.8 

& 

0 

0.6 g 

5 

n 

U 

0.4 

0.2 

0 

0 0.5 1.5 2.0 2.5 3.0 3.5 4.0 4.5 5.0 

65321 7-31 6 STRAIN (x 10-31 

Flgun 7-18 Tonsil. Load-Shin Dlagrarn for [OO f 1501, Celion 6000/LARC-160 G/PI at 297 K (75OF) 

TEMPERATURE (OF) r 

-100 0 100 200 300 400 500 600 7 00 

i 1 I I 1 1 1 I i 

200 

160 0 

z 

w 

a 

i 

a 

0 

TEMPERATURE (KI 65321 7-31 7 

Figure 7-19 Borelino Tonrile Propertior of [Oo 2 45O], Celion 6000/LARC-160 G/PI 

7-37 . 
- 1500 n?. c 

z 

s - 1250 

S 

c 

CI 

z 

- 

g 1000 

z 

W 5 750 

2 500 

3 

W 

c 

w 

c 

250 

0, 

200 

200 

- 

y. 

.- 

160 

+ 

c1 

z 

W 

120 E 

i 

a0 5 

cn 

w 

VJ 

c 

W 

c 

4 

40 

!i 

a 

0 

I I 1 1 I 1 1 

~ 

. r I 

I I 

250 300 350 400 450 500 550 600 650 

65321 7.318 

Figun 7-20 Bardino Notehod tonsilo Pmportior of [OO f 4501, Collon 6000/LARC-160 G/PI 

The notched to unnotched tensile strength ratios of the [O f 45 OIs Celion 60OO/WC-160 

G/PI specimens at each of the test temperatures are listed in Table 7-24. The effect of the notch 

was a reduction in the net section strength of 20 to 40%. Similar results were observed for the 

other G/PI system and both the epoxy systems with the [O O 5 45 OIs layups. 

7.1.2.4 Boron/AJuminum. The baseline tensile test plan for the B/Al system as originally proposed included tests at 218 K (-67 O F) and 297 K (75 ' F) as used for the resin matrix systems 

and at 561 K (550 ' F) and 700 K (800 O F). The 700 K (800' F) temperature was considered to be 

the maximum short-time use temperature and 561 K (550' F) the maximum use temperature 

for very long time applications. However, fatigue tests conducted on this program (see Section 

11), revealed the existence of a severe surface degradation effect during fatigue cycling of B/Al 

at 561 K (550 F) in an air environment. The effect was not observed at 505 K (450 ' F). Because 

of this severe surface degradation the flight simulation temperature initially selected for B/AL 

561 K (550 O F), was lowered to 505 K (450 ' F). At the time of this decision all baseline testing 

was completed except for the unnotched [0 t 45 'Is crossplied mated A change was made to 

indude several tensile'tests at 505 K (450' F) to assist in selection of the short-term flight 

simulation load levels. Three specimens were also tested at 616 K (650" F) to fill in the tensile 

strength versus temperature curve for unnotched [0 " k 45 "Is crossplied B/AL 

Baseline tensile tests, both unnotched and notched, were performed on unidirectional and [0 O 

45 OIs crossplied specimens in the longitudinal direction. Some unnotched tensile tests were 

also performed in the transverse direction on unidirectional material, and a limited amount of 

Poisson's ratio testing was conducted. Results are listed in Table 7-25 through 7-30. A typical 

load-strain diagram is plotted in Figure 7-21. The longitudinal unnotched and notched data are 

7-38 
b 

aa 

CD 

0 

aa 

(0 

(0 

aa 

(0 

b) 

Cr) 

PI 

- 

u) 

b 

PI 

I 

I 

7-39 
Tablo 7-24. Ratio of Notehod to Unnotcttod Tonsilo Strongh of [OO f 4501, 

Colion 6000/LARC=160 G/PI at Various Tompomtuns 

Tomprm)un 

K (OF) Ratio of Notchod to Unnotcbod Strongth 

218 -67 

297 75 

394 250 

450 350 

478 400 

505 450 

533 500 

0.6 

0.6 

0.6 

0.7 

0.7 

0.8 

0.7 

Spdmon Tomporatun Tonsilo Stnngth tonsilo Modulus 

Numbor K (OF) MNM (ksi) GN/d (Msi) 

EUO1-1 218 -67 1640 238 178 25.8 

-2 1620 235 197 28.6 

-3 1580 229 172 25.0 

4 1540 223 166 24.1 

-5 22.8 

avg 1600 232 174 

- 

25.3 

1610 - 234 - 157 

EU01-6 561 550 

-7 

-8 

-9 

EUO1-10 700 800 

-11 

-12 

-13 

1200 

1360 

1200 

1430 

avg 1300 

1360 

1150 

1380 

1320 

avg 1300 

- 

- 

174 

197 

174 

207 

188 

198 

167 

200 

192 

189 

a199 

a212 

a207 

206 

a192 

a194 

a214 

200 

- - 

- - 

a28.9 

a30.7 

a30.0 

29.9 

a27.8 

a28.2 

a31.0 

- - 

- 

29.0 

a Strain gage measurement 
toblo 7-26. Baulino Tonsilo Proportior of [OO f 4501, B/Al 

ECO1-1 

-2 

-3 

-4 

-5 

EC01-6 

-7 

-8 

-9 

-10 

-11 

-12 

-13 

-14 

-15 

ECO 1-1 6 

-17 

-18 

-28 

-29 

-30 

-3 1 

EC01-19 

-20 

-21 

-32 

-33 

-34 

-35 

EC01-22 

-23 

-24 ' 

EC01-25 

-26 

-27 

218 -67 

297 

505 

75 

450 

561 550 

616 650 

700 800 

499 

550 

525 

554 

511 

avg 

- 

528 

487 

545 

482 

458 

552 

565 

549 

459 

545 

514 

avg 

- 

516 

459 

447 

468 

454 

457 

471 

410 

avg 

- 

452 

450 

365 

416 

450 

474 

455 

464 

avg 439 

423 

397 

- 

499 

avg 7 440 

467 

338 

335 

avg 

- 

380 

72.4 

79.8 

76.1 

80.4 

74.1 

76.6 

70.7 

79.1 

69.9 

66.4 

80.0 

82.0 

79.6 

66.5 

79.1 

74.5 

74.8 

66.6 

64.9 

67..9 

65.9 

66.3 

68.3 

59.5 

65.6 

65.2 

53.0 

60.3 

65.3 

68.8 

66.0 

67.3 

63.7 

61.3 

57.6 

72.4 

63.8 

67.7 

49.0 

48.6 

55.1 

- 

- 

- 

- 

- 

- 

- 

- 

- 

- 

- 

155 

158 

160 

140 

167 

163 

169 

154 

141 

139 

155 

a104 

142 

100 

136 

126 

116 

124 

121 

a103 

123 

102 

120 

128 

103 

108 

112 

a108 

- 

7 

- 

- 

- - 

108 

a93.1 

- 

- - 

93.1 

22.5 

22.9 

23.2 

20.3 

24.2 

23.7 

24.5 

22.4 

20.4 

20.1 

22.4 

a15.1 

20.6 

14.5 

19.8 

18.3 

16.9 

18.0 

17.6 

- 

- 

a15.0 

17.9 

14.8 

17.4 

18.5 

15.0 

15.6 

16.3 

a15.6 

- 

- 

- - 

15.6 

a13.5 

- 

- - 

13.5 

a Strain gage measurement 

7-4 1 
Tablo 7-27. Eaulino Notehod tonsilo Pr0portfi.r of [VI6 B/Al 

EU02-1 

-2 

-3 

-4 

-5 

EU02-6 

-7 

-8 

-9 

EU02-10 

-1 1 

-12 

-13 

218 -67 1200 

1210 

1240 

1230 

1210 

avg 1220 

561 

700 

550 

800 

1190 

1130 

1140 

1130 

avg 1150 

986 

1100 

1080 

1030 

avg 1050 

a Net section strength, La, based on total width less hole diameter. 

174 

175 

180 

178 

176 

177 

173 

164 

165 

- 

164 - 

166 

143 

159 

157 

150 - 

152 

plotted as a function of temperature in Figures 7-22 and 7-23. A summary of all the BlAl tensile 

properties is presented in Table 7-31. 

Modulus measurements were made using both a mechanical extensometer and strain gages at 

temperatures to 561 K (550 O F). Above 561 K (550" F) a small number of tests was made using 

only strain gaged specimens. Modulus values were obtained at 218 K (-67 O F) for the unidirectional specimens, but extensometer malfunctions during tests on the crossplied specimens 

gave poorly defined stress strain curves that did not allow determination of modulus. For the 

material tested in the transverse direction, modulus was determined only at room temperature 

using a mechanical extensometer. 

The effect of temperature on the tensfie strength of BlAl is shown in Figures 7-22 and 7-23. No 

tests were conducted on the unnotched unidirectional material at room temperature, but data 

were available from the QA panels, and these data are plotted at mom temperature in Figure 

7-22. The curves show a gradual decrease in strength with increasing temperature for both 

orientations and configurations. The B/Al data are similar to those for HT-S/710 G/PI in that 

the decrease is somewhat greater for the crossplied material, both unnotched and notched, 

than for the unidirectional material. The tensile strength was slightly higher at 218 K ( -67 O F) 

for the unnotched specimens but essentially unchanged for the notched specimens. In like 

manner to both the G/E and HT-S/710 GlPI systems, the data scatter for BlAl was numerically 

greater for the unidirectional material, but, percentagewise, was greater for the crossplied 

7-42 
tablo 7-26. Barolino Natchod Tonsilo Pmpodor of [OO f 45O], B/Al 

EC02-1 

-2 

-3 

-4 

-5 

EC02-6 

-7 

-8 

-9 

-10 

-11 

-12 

-13 

-14 

-15 

E CO2- 16 

-17 

-18 

-19 

-20 

-21 

-22 

-23 

-24 

-25 

EC02-26 

-27 

-28 

-29 

-30 

-31 

-32 

-33 

-34 

-35 

218 -67 486 

492 

460 

442 

495 

avg 475 

297 75 

561 550 

700 800 

470 

465 

470 

467 

455 

496 

485 

490 

495 

501 

avg 479 

434 

433 

423 

461 

454 

439 

421 

451 

455 

462 

avg 443 

361 

361 

363 

377 

341 

361 

378 

371 

363 

351 

avg 363 

70.5 

71.3 

66.7 

64.1 

71.8 - 

68.9 

68.1 

67.4 

68.1 

67.7 

66.0 

72.0 

70.4 

71.0 

71.8 

72.7 

69.5 

62.9 

62.8 

61.4 

66.9 

65.9 

63.7 

61.1 

65.4 

66.0 

67.0 

64.3 

52.3 

52.4 

52.6 

54.7 

49.5 

52.3 

54.8 

53.8 

52.7 

- 

- 

50.9 - 

52.6 

7-43 

a Net section strength, Le., based on total width less hole diameter. 
tablo 7-29. Bosdino tonsilo Proportios of [OOL B/Al Tostad in tho tmnrvono Dirodion 

TEuo1-1 

-2 

-3 

-4 

-5 

-6 

TEUO1-7 

-8 

-9 

-10 

-11 

-12 

297 75 94.5 

122 

51.0 

74.5 

57.9 

76.5 

avg 79.4 

- 

700 800 44.1 

31.7 

38.6 

29.0 

35.8 

33.1 

avg 35.4 

- 

13.7 

17.7 

7.4 

10.8 . 

8.4 

11.1 

11.5 

6.4 

4.6 

5.6 

4.2 

5.2 

4.8 - 

5.1 

48.3 

67.6 

133 

121 

105 

140 - 

102 

7.0 

9.8 

19.3 

17.6 

15.2 

20.3 - 

14.9 

- 

- 0 

Tobio 7-30. Badino Poisson's Ratio Duta in tomion of B/Al 

TUt 

Swimon . ~ominoto TeSt Tempi&n 

Numlmr Oriontution Dimtion K (OF) 

Poisson's 

0.3 1 

EU01-6 

-7 

-8 

EU01-10 

-11 

-12 

[O 76 Longitudinal 561 550 0.31 

.34 

.34 

avg .33 

Longitudinal 700 800 .32 

.30 

.39 

material. The scatter bands for the notched tests, Figure 7-23, were considerably smaller than 

those obtained for the unnotched specimen tests, Figure 7-22. This effect was not observed for 

the resin matrix systems. 

No decrease in tensile modulus was found for the unidirectional BlAl out to 700 K (800" F). 

There was a decrease of about 15% however, for tests at 218 K (-67 " F). The [0 " h 45 "Is crossplied material exhibited a continuous decrease in modulus out to 700 K (BOO" F). The total 

change between room temperature and 700 K (800" F) was approximately 40%. 

7-44 
TEMPERATURE (OF) 

-1 

-- 1750 

E 

z 

I 

e 

0 

\ 

3 1500 

1250 

G 

1 100c Ly 

v) 

2 

w L 7% 

e 

U 

a 

3 

- 5 SO( 

251 

TEMPERATURE (K) 

Figuro 7-22 Baseline Tensile Propertior of B/Al 

240 I- 

=! 

s 

m 

+ 

m 

160 E 

m 

200 5 

r 

v) 

120 

2 

G1 + 

I 80 - 

- E 

40 

0 

65321 7-35 

7-15 
65321 7-36 

Ffgun 7-23 Basdim Notehod tonrilr Proportior of WAl 

The ratios of the notched to the unnotched tensile strengt& of BlAl have been calculated for 

several temperatures and are shown in Table 7-32. The data differ from those presented eariier 

for the resin matrix systems in that the crossplied specimens were not greatly affected by the 

notch; whereas, the unidirectional specimens experienced decreases in tensile strength of 

approximately 20% when tested in the notched configuration. 

Tensile data for unidirectionai material tested in the transverse direction are given in Table 

7-29. The effect of temperature on the transverse tensile strength was much more pronounced 

than that for longitudinal strength with values at 700 K (800O F) approximately half of those 

measured at room temperature. 

Results of the Poisson’s ratio tests are listed in Table 7-30. Values at both test temperatures 

were approximately onethird, similar to those found for most unreinforced metals. 

7.2 SHEAR TESTING 

Two methods of shear testing were utilized for both the resin matrix and the metal matrix 

systems. For the resin matrix materials, a torsion tube specimen configuration that measures 

in-plane shear and a short beam configuration that measures interlaminar shear were tested. 

Short beam shear testing, which requires a specimen thickness between 0.0015 m (0.060 in.) 

and 0.064 m (0.25 in.), was performed on all 12 and 24 ply resin matrix laminates. These data 

7-16 
I 

m 

2 

I 

I 

II 

II 

7-47 
tabk 7-32. Rot10 of N&hd to Unnotchod tmsilo Strongth of B/Al at Various tompom)unr 

t.mpOrcrhrn Ratio of Notchod to Unnotchd Strongdl 

K (OF) [w6 [OO f 1501, 

218 -67 0.8 

297 75 a .8 

561 550 .9 

700 800 .8 

0.9 

.9 

1.0 

1.0 

~ ~ ~~ 

aEstimated from curves of tensile strength ve!rsus temperature 

have been reported in the Laminate Quality Assurance Testing section. Shear strength prop 

erties of the 6 ply [0 O f 45 "Is material also became of interest, however, because of the matrixdominated failure modes observed in the 6 ply Phase I flight simulation specimens. A specimen made by bonding three or four pieces to obtain a sufficiently thick laminate was devised, 

and short beam shear strength was included in the baseline property testing for the thermal 

aging/compression study materials. For B/Al the torsion tube specimen and a slotted rectangular interfiber shear specimen were used. Specimens, test procedures, and results are presented in the following sections for the methods that used the tube and rectangular interfiber 

shear specimens. Only the test results are given for the short beam shear tests of the bonded 

18.or 24 ply specimens as the specimen dimehqions and test procedure are discussed in Appendix B, Quality Assurance Test Procedures. 

7.2.1. SPEQMEN DESIGN AND TEST PROCEDURE. The B/Al torsion tube specimen was 0.15 m (6 

in.) long by 0.024 m (1 in.) outside diameter with a nominal 0.001 m (0.040 in.) wall thickness. 

The specimens were fabricated from 6 plies of B/Al tape oriented with the boron filaments in 

the long-axis direction of the tube. Steel end fittings were diffusion bonded to the tube in the 

manufacturing process. Loading was accomplished by cylindrical male fixtures that were pinattached to the steel reinforced tube ends with 0.0064 m (0.25 in.) steel bolts. A sketch of the 

BIAl tube is shown in Figure 7-24. 

The GIE torsion tube specimen, also shown in Figure 7-24, was similar to the BIN tube except 

for the length and end fittings. These specimens were 0.089 m (3.5 in.) in length and had steel 

fittings adhesively bonded to the ends after fabrication and cutting to the correct lengths. To 

obtain a wall thickness of 0.001 m (0.040 in.) (to prevent buckling during test) it was necessary 

to use an 8 ply tube for the GIE specimens rather than the 6 plies used for the BIAl specimens. 

To determine the shear modulus, all of the torsion tube specimens were strain gaged prior to 

test. A half bridge, shear-torque rosette strain gage was bonded to each tube with M-bond 610 

adhesive. The gages and adhesive cure cycles that were used are: 

a Gages: 

BIN - FAED-12C-12-S3 

GIE - FAED-12C-12-SO 

7-48 
P 

DflAIL: 0 a001 m (a040 in) THICKNESS (6 PLY) E/Al TUBE 

@ a0024 m (a094 in) THICKNESS STEEL REINFORCEMENT 

@ a00071 m (aoza in.) THICKNESS STEEL REINFORCEMENT 

I 0.165 m ( 61/2 in) I 

OETAIL: 0 

0 

0 

a001 m ((1040 in) THICKNESS (8 PLY) A43501 G/E TUBE 

a0032 m (1/8 in.) WALL STEEL TUBE 

a0016 m (1/16 in) WALL SPLIT STEEL COLLAR 

653217.37 

Figun 7-24 Torsion Tube Configurations (a) B/Al Specimen and (b) G/E Specimen 

7-49 
b. Cure Cycles: 

Test Temperature Cure Cycles 

297 K (75" F') 

394 K (250" F) 

450 K (350" F) 

505 K (450" F) 

2 hours at 394 K (25OOF') 

2 hours at 394 K (250 O F) 

1.5 hours at 450 K (350" F) 

1 hour at 505 K (450" F) 

Application techniques were the same as those described earlier for the Poisson's ratio and tensile modulus specimens. 

Testing of the strain gaged tube specimens was conducted in an MTS Closed-Loop TensionTorsion machine. The specimens were loaded (in torsion) incrementally to failure with strain 

readings recorded automatically at each load step with a B & F Instruments, Inc. Multichannel Digital Strain Indicator, 161 Mini-System. The output was both a typewritten recording of torque versus strain and a punched paper tape that could be processed by a computer, 

programmed to produce a shear stress verms strain plot. In addition, a torquetwist curve was 

also recorded for each test. Typical torquetwist and stress-strain curves are shown in Figures 

7-25 and 7-26. An overall view of the torsion testing machine and the Mini-System strain 

recorder is presented in Figure 7-27. For elevated temperature tests the specimens were heated 

in the mid-section with a small ring furnace and held 10 minutes prior to twisting. A closeup 

view of a B/Al tube installed in the test machine is shown in Figure 7-28 prior to testing at 

room temperature and in Figure 7-29 before testing at elevated temperature. The white 

material packed around the tube above and below the ring furnace in Figure 7-29 is Fiberfrax 

insulation batting for maintaining good temperature uniformity. Temperature was monitored 

with a chromel-alumel thermocouple attached to the center of the tube and controlled manually 

to f 3 K (f 5 "F) during the short time period required for the test. 

Shear strengths were calculated from the torsion test data and tube dimensions by using the 

following thin walled tube approximation of the torsion formula for circular shafts: 

T 

2TDo 

r t (DO - t)3 

T = ultimate shear strength, N/m2 (Iblin2) 

T = Maximum torque, Nm (in-lb) 

Do = Outside diameter of tube, m (in.) 

t = Wall thickness of tube, m (in.) 

The shear modulus, G, was determined from the stress-strain wes as follows: 

G =--- 7 7 - 

Y €1-9 

where y (equal to €1 - €2) is the angle of twist and €1 and €2 are strain readings taken at a value 

of shear stress, 7, in the linear portions of the stress-strain curves (see Figure 7-26). 

The second type of shear specimen, which was used only for the BlAl system, is shown in 

Figure 7-30. The slots were cut in the specimens by the EDM method. All specimens were pre 

pared from 6 ply unidirectional material with the boron filaments oriented in the long direction 

7-50 . 
TWIST 4 

Figure 7-25 Typicoi TorquoTwist Cuwo for Tonion Test of Unidirectionai G/E Tube Specimen BUO6-13 

65321 7-38 

7-51 
10 

9 

8 

7 

6 

5 

4 

3 

2 

1 

0 

20.0 

65321 7-39 

-20.0 -10.0 0.0 10.0 

STRAIN (x 10-3) 

Figure 7-26 typical Stress-Strain Curve for Torsion Tost of Unidirectional G/E tube, Specimen 8UO6-13 

7-52 
n 

Q) 

n 

a 

c 

n 

a c 

e 

n 

a 

0 

P 

7-53 
Figure 7-28 Strain Gaged B/AI Tube Specimen Installed in Torsion Test Machine 

of the specimen. No end doublers were required because of the low failure loads. These low 

loads ensured failure in the test sections rather than the grips. The specimens were tested in 

Instron or Tinius-Olsen testing machines using Instron capstan grips with ball or universal 

joints to ensure axial loading. A head rate of 0.0025 m (0.01 in.) per minute was used. No strain 

measurements were made during testing of these BlAl interfiber shear specimens. Tests at 218 

K (-67" F) and 561 K (550" F) were conducted in a Conrad-Missimer Chamber. The 700 K 

(800O F) tests were performed with a small ring furnace. As in the tensile tests, temperature 

was measured with a thermocouple clipped to the surface of the specimen. Temperature was 

maintained at the desired level for 10 minutes prior to loading. 

7.2.2 TEST RESULTS AND DISCUSSION 

7.2.2.1 Graphite/Epoxy System. The initial plan for the GIE system was to test six tubes each 

at room temperature and 450 K (350 O F). This was later changed such that at elevated temper 

ature three were tested at 394 K (250 O F) and three at 450 K (350 O F). The revision was made to 

7-54 

(.)RIGINAL PAGE 1s 

OF POOR QUALITY: 
ORIGINAL PAGE IS 

Q! POOR QUALITY, 

Figure 7-29 Test Setup for Elevated Temperatures Torsion Testing of B/AI Tube 

better define the shear strength versus temperature relationship. Shear strength and shear 

modulus data for GIE are presented in Table 7-33. Shear modulus values were obtained from 

torquestrain diagrams such as the one shown in Figure 7-31. The effect of temperature was to 

reduce the strength by approximately 30% at both 394 K (250 O F) abd 450 K (350 O F). While 

the strength leveled off at 394 K (250 O F) and 450 K (350 O F), the shear modulus exhibited a 

continuous decrease out to 450 K (350' F). The percentage change in modulus was, however, 

significantly less than that for the shear strength. Data from one of the tests (tube no. 

BU06-37) was not usable because of overheating of the tube. The overheated tube failed at a 

torque level that was less than onetenth that of the other tubes tested at 450 K (350" F). 

Short beam shear strength of the 6 ply material used in the thermal aging/compression study 

is given in Table 7-34. These G/E specimens were made by bonding together three pieces of the 

6 ply panels to obtain 18 ply laminates. 

7-55 
61321 7-43 

ngun 1-30 B/Al In-plom Shoat Spoeirnon 

7.2.2.2 Boron/Aluminum Systom. Data for the two types of shear specimen used in testing of 

B/AI are listed in Tables 7-35 and 7-36. Strain measurements for determination of shear 

modulus were made only on the tube tests. A typical torquestrain diagram is shown in Figure 

7-32. The initial test plan was to conduct laminate shear testa at the four temperatures shown 

in Table 7-36 and to conduct the tube testa at 297 K (75 " F) and 561 K (550 " F). As discusssed 

in the baseline tensile section, 561 K (550" F) was considered to be the maximum use 

temperature of BlAl for very long time applications. Because of a surface degradation of the 

B/Al observed during fatigue testing on this' program at 561 K (550" F), the maximum use 

temperature was lowered to 505 K (450" F). At the time of this decision alI shear testing was 

completed except for the elevated temperature tube tests. A change was made to conduct 

these tests at the lower temperature, 505 K (450" F). 

The shear strength of B/N as determined from the tube tests, Table 7-35, was not lowered by 

increasing the temperature from room temperature to 505 K (450 " F). There was a decrease in 

shear modulus, however, of approximately 15%. For the laminate specimen tests, Table 7-36, a 

decrease in shear strength of 30 to 40% was observed at elevated temperatures. At room 

temperature higher shear strengths were obtained with the laminate specimen configuration 

than with the tubular configuration although, presumably, the two measure the same prop 

erty. The two groups of specimens were from different batches of material, which could account 

for some of the difference. Another possible reason for the lower values obtained from the tube 

tests was the presence of some small longitudinal cracks on the outside surface of half of the 

tubes. These cracks, shown in Figure 7-33, were shallow and did not extend beyond the first 

ply. Whether or not other cracks were present in any of the inner plies of the 6 ply tubes could 

not be determined. There was no correlation between the individual values of shear strength 

and the presence or absence of the surface cracks, however. While the longitudinal cracks in 

the tubes may well have been responsible for low values of ultimate shear strength, it is not 

believed that the cracks had much effect during the early portions of the tests where the shear 

modulus was determined. In the past, agreement of shear strength values determined by different test methods has generally been poor for composite systems. 

7-56 
Tablo 743. Baulino Intorfibor Shoar Proportier of G/E Unidirdonal Tubos 

Spodmon Tomporotur~ Shoar Strongth Shoar Modulus 

Numbor K (OF) MN/d (kri) GN/d (Msi) 

BU06-11 

-12 

-13 

-24 

-25 

-26 

BU06-38 

-40 

42 

BUO6-37 

-39 

-4 1 

297 75 94 

94 

106 

89 

92 

86 

avg 94 

- 

394 

450 

250 

350 

52 

72 

70 

avg 65 

(a) 

67 

70 

avg 68 

- 

- 

13.7 

13.6 

15.4 

12.9 

13.4 

12.5 

13.6 

7.6 

10.5 

10.1 

9.4 

(a) 

9.7 

10.2 

9.9 

- 

- 

- 

5.1 0.74 

6.1 .a8 

5.4 .78 

5.5 .80 

5.5 .80 

5.9 .85 

4.7 .68 

4.8 .70 

5.6 .a1 

5.4 .78 

5.0 .72 - 

(a) (a) 

4.4 .64 

4.5 - .65 

4.4 .64 

- 

. aSpecimen overheated 

STRAIN (x 10-3) 

65321 7-44 

Figuro 7-31 Shear load-Strain Diagram for [0°]8 G/E at 394 K (250O F) 

7.57 
Tablo 7-34. Bardino Shoe Boom Show Strongth of [0° k 45O], G/E at 297 K (75OF) 

(Thorrnol Aging/Compmrion Study - 18 Ply Laminator) 

Spoeimon 

Numbor 

BC102-1 

-2 

-3 

72 10.4 

68 9.9 

75 10.9 

avg 72 10.4 

- 

Tablo 745. Basolino Intorfibor Shoor Proporti- of B/Al Unidlrodonal Tubor 

Spoeimon T.mpOfatUn Shmr Strongth Shoor Modulus 

Numbor K (Of) MN/d . (kri) GN/d (Msi) 

EUO6-1 297 75 

-2 

-3 

4 

-5 

-6 

EU06-7 

-8 

-9 

-10 

-11 

-12 

505 450 

61 

90 

66 

61 

63 

61 

avg 67 

82 

64 

70 

65 

77 

b?4 

avg 72 

- 

- 

8.8 

13.0 

9.5 

8.8 

9.1 

8.8 

9.7 

11.9 

9.3 

10.1 

9.4 

11.1 

b10.8 

10.4 

- 

- 

5.0 7.3 

(a) (a) 

4.3 6.2 

4.5 6.5 

4.6 6.7 

4.2 6.1 

4.3 s.s 

4.1 5.9 

3.2 4.6 

3.6 5.2 

2.8 4.1 

5.2 7.5 

4.5 6.4 

3.9 5.6 

aStrain gage malfunction 

horque still rising - machine at maximum twist 

One of the tube specimens, EU06-12, did not fail before the test machine had reached its limit 

of twist, 100 degrees. At this point the torque was still increasing slightly. Consequently, the 

value of ultimate shear strength reported in Table 7-35 is slightly low. The unfailed, but 

twisted, specimen in shown in Figure 7-34 with a similar specimen that has failed in shear in 

the longitudinal direction of the tube. 

Data from one of the laminate specimens, EU06-32, was not reported because of improper 

machining. The slot on one side of the specimen was not sufficiently deep and several continuous boron filaments remained across the test section. With this configuration a shear 

failure was not possible thus invalidating the test. 

7-58 
Tablo 7-36. Barolino Intodibor Shoar Stnngth of [O0l6 B/Al Laminato8 

EU06-13 218 

-14 

-15 

-16 

-17 

-67 

EU06-18 

-19 

-20 

-21 

-22 

EU06-23 

-24 

-25 

-26 

-27 

EU06-28 

-29 

-30 

-3 1 

-32 

aSpecimen improperly machined 

297 75 

561 550 

700 800 

90 

100 

91 

100 

93 

avg 95 

97 

100 

94 

. 94 

81 

avg 93 

61 

52 

54 

62 

50 

avg 56 

66 

64 

66 

63 

65 

- 

- 

- 

- (a) 

13.1 

14.5 

13.2 

14.5 

13.5 

13.8 

14.0 

14.5 

13.6 

13.7 

11.8 

13.5 

8.9 

7.5 

7.8 

9.0 

7.2 

8.1 

9.5 

9.5 

9.6 

9.1 

(a) 

9.4 

- 

- 

7 

- 

One inteesting point in comparing the BlAl data with that for the resin matrix materials is the 

striking difference in values of shear modulus. Values for the metal ma& material are almost 

an order of magnitude higher than those obtained for the resin matrix systems. Shear strength 

values (at room temperature), on the other hand, differ by no more than 30%. Similar results 

are obtained when comparing transverse tensile data (modulus and strength) of metal and 

resin matrix systems. 

7.2.2.3 Graphite/Polyimido 

HT-SnlO G/PI. Because sufficient HT-SI710 prepreg was unavailable, it was not possible to 

fabricate the tube specimens necessary to conduct the G/PI shear tests. However, some shear 

data, for the same HT-S/710 system, had been obtained in a study performed at General 

Dynamics Convair during 1973. These dates are listed in Table 7-37 (unpublished data from 

7-59 
400 

300 

-I 

0 

a 

P 

c 

m 

200 z - CT 

100 

3 

61321 7-46 

Flgun 7-32 Shoar Load-Strain Diagram for [00]6 B/Al at 505 K (4500 F) 

AXIAL 

01 RECTION 

OF TUBE 

Ftgure 7-33 B/Al Tube Specimen, EU06-12, Showing Longitudinal Surfoce Crocks 

7-60 
7-47 

Figure 7-34 B/Al Tube Speeimonr After Shear Testing: Upper, Specimen EU06-1, Twisted to Failure in 

Shear: lower, Specimen EUM-12, Twisted to Maximum Capability of Test Machine 

W.G. Scheck). The GIPI tubular specimens were processed using the same fabrication techniques described in Section 5 for GIE and B/PI tubes and the autoclave cure cycle given in Section 5 for the HT-SI710 GlPI laminates. The tubes were 0.15 m (6 in.) long by 0.025 m (1 in.) outside diameter with a nominal wall thickness of 0.001 m (0.040 in.). Test techniques were similar 

to those described above for the B/Al and G/E tubes (Section 7.2.1). 

The data of Table 7-37 show a decrease in shear strength of G/PI of approximately 30% of the 

room temperature value for specimens tested at both 77 K (-320OF) and 589 K (600°F). The 

modulus, on the other hand, decreases slightly as the test temperature is raised from 77 K 

(-320 OF) to 589 K (600 OF). 

Celion 6000/LARC-160 G/PI. As no unidirectional layups were evaluated for the Celion 6000/ 

LARC-160 system, the only shear tests performed were those using the short beam configuration. Test results for the crossplied 12 and 24 ply panels have been reported in the Laminate 

Quality Assurance Testing section. The tests on the crossplied 6 ply panels were made on specimens fabricated by bonding together three pieces to make an 18 ply laminate. These data, 

which were used as the baseline shear data for the thermal aginglcompression study, are given 

in Table 7-38. 

7.3 FRACTURE TESTING 

The effects of stress risers (cutouts, cracks, damage zones, material discontinuities, etc.) in 

composite materials has been the subject of a number of recent investigations. Methods are 

7-61 
Tablo 7-37. Intorfibor Shoar Propodos of HTW710 G/Pl UnidlMonal Tubos 

Spuimrcr Tomporatun Shoar Strength Shoar Modulus 

Numbor K (OW MN/d (kri) GN/d (Msi) 

Tu-1 

A-3 

A-4 

Tu-4 

Tu4 

Tu4 

Tu-7 

A-5 

77 

297 

589 

-320 52 7.6 6.8 0.98 

44 6.4 7.2 1.04 

7.3 7.2 - 1.05 

7.1 1.02 

- - 

avg 49 7.1 

75 87 12.6 5.8 .84 

68 9.9 6.6 .96 

8.2 

avg 71 

- 

10.2 

- 57 6.1 - .89 

6.2 .90 

600 49 7.1 7.2 1.04 

7.1 

avg 

- 

49 7.1 

49 - 4.8 .69 

6.0 - .86 

Tablo 7-38. Basolino Shoe &om Shoar Stnngth of [Oo 4 Go], Cdion 6000/LARC-160 G/PI at 

297 K (75Of) (Thormal Aging/Comprossion Study - 18 Ply Laminates) 

Speimon 

. Numbor 

Shoar Strength 

I"/& (ksi) 

FC 10 2- 1 50 7.2 

-2 51 7.4 

-3 52 - 7.6 

avg 51 7.4 

being sought to predict both the static and fatigue behavior of composite materiais that contain such stress risers. Waddoups, Eisenman. and Kaminski (ref. 29) have developed a fracture 

mechanics based model, the Wearout Model, which has been moderately successful in predicting the degradation of coupon strength as a function of loading. Whereas in metals, crack 

growth rate (Le., dah) is nody measured and used as an indication of strength degradation such is not the case for composites. Flaw growth is usually not observed in composites, 

even though such crack growth has been observed for both B/Al and B/E (ref. 30 and 31). 

The objective of this study was to attempt to observe damage accumulation in flawed composite tension coupons and to reduce the data to a dddn format. Since crack growth is not 

readily observed in composites, changes in material compliance (load, P, divided by crack opening displacement, COD) were to be related to changes in effective crack length Given da/dn 

data for a composite material, conventional fracture mechanics theories can then be applied to 

predict effective crack length as a function of loading history, and failure characteristics. 

7-62 
7.3.1 SPECIMEN DESIGN. The study included six specimens each of four different materials to 

determine fracture behavior. The four materiala investigated are described in Table 7-39. For 

each material, six (0" k 45'1, specimens measuring 0.064 m (2.5 in.) by 0.30 m (12 in.) were 

fabricated and identified by dash numbers as shown in Table 7-40. Specimen configuration is 

shown in Figure 7-35. 

Initial flaws were applied by drilling a No. 13 hole, 0.0047 m (0.185 in.) diameter, in the center 

of the specimen and then extending symmetrical narrow slits, 0.00066 m (0.026 in.), from each 

side of the hole by ultrasonic impact grinding. The end of a flaw made in this manner in B/E is 

shown in Figure 7-36. Dash three specimens reguired that the slots be progressively increased 

during testing without losing the clip gage instrumentation. This was done by hand with a 

0.00013 m (0.005 in.) diameter diamond coated wire. All flaw lengths at the beginning of each 

test sequence are shown in Table 7-41. 

7.3.2 TEST PROCEDURE, R€SULTS AND DISCUSSION., The test phn iS Summarized in Table 7-42. 

The procedure for -1 and -2 tests was to try to initiate crack pwth, either visual or effective 

delta length, by progressively increasing static loads and then applying cyclic fatigue loads to 

extend the effective crack length. Clip gages were used to monitor COD, and any change in 

compliance, P/COD, was considered to represent an increase in effective crack length. These 

specimens would then be static pulled to determine their residual strengths. 

Table 7-43 shows results of compliance testing of -1 and -2 specimens. The expected systematic 

damage growth and corresponding change in slope of the compliance curve did not occur. Only 

isolated damage growth was observed, and this was mostly in the form of cracks parallel to 

surface fibers. More recent experiments suggest that synergistic effects of moisture and cyclic 

rate play a large role in damage growth in composites. The failed specimens are shown in 

Figures 7-37 and 7-38. 

The purpose of the -3 specimens was to determine load versus COD for progressively increased 

crack lengths. Clip gages were used to measure COD and flaws (slits) were lengthened in small 

increments by hand sawing with a diamond coated wire. For each flaw length the specimen was 

slowly cyded, three times, up to a maximum load of 1000 pounds. Test results for the four 

specimens are shown in Figures 7-39 through 7-42. Load divided by COD versus flaw length 

was then plotted for each material as shown in Figure 7-43. All four specimens yielded an expo. 

nential decaying type of me with the data points of the G/E and G/PI specimens lying on the 

same curve. The consistent shape of these curves warrants further investigation of a Load 

COD = Asflaw size fatigue model. The failed specimens are shown in Figure 7-44. 

The -4 specimens were fatigue tested with the 0.0047 rn (0.185 in.) diameter hole only. No slits 

were applied. COD was measured after cycle numbers 1,2,19,100,1000,10,000 and 2,500.000. 

Specimens were then static pulled to determine residual strengths. 

Fatigue test results are shown in Table 7-44. There was no fatigue crack growth observed in 

any of the four specimens. When these specimens were static pulled for residual strength, only 

one, AC05-4, failed at the 0.0047 m (0.185 in.) diameter hole. The other three failed in or near 

the grips. The failed specimens are shown in Figure 7-45. 

The objective of -5 and -6 specimen tests was to determine static strengths at two different 

flaw lengths. The -5's had 0.0286 m (1.125 in.) flaws and the -6's had 0.00349 m (1.375 in.) flaws. 

All were simply pulled to static failure. 

7-63 
tablo 7-39. Comporito Matorials Studiod 

Pan4 thieicnosr 

Matorial LWJP m in. Matorial ID 

550514 B/E [O" * 4501, 0.00076 0.030 AC05 

A-S13501-5 GIE . [O" f 45"], .00076 .030 BC05 

HT-SI710 G/PI [O" f 4501, .00076 .030 DC05 

BlAl [0 -C 45 '1, .00107 .042 EC05 

tablo 7-40. Spoeimon Idoniifkation 

Spocimonr 

5506/4 B/E 

A-Sl3501-5 G/E 

HT-SI710 GlPI 

B/Al 

AC05-I, AC05-2, AC05-3, AC05-4, AC05-5, AC05-6 

BC05-1, BC05-2. BC05-3, BC05-4, BC05-5, BC05-6 

DC05-1, DC05-2, DC05-3. DC05-4, DC05-5, DC05-6 

EC05-1, EC05-2, EC05-3, EC05-4, ECO5-5, EC05-6 

~~ 

DIMENSIONS IN rn (IN.) 

I 

a064 (251 ,500) 

65321 7-08 

Agun 7-35 Tost Sprcimon Configuration 

Results of static strength testa are shown in Table 7-45, and the failed specimens are shown in 

Figures 7-46 and 7-47. Results of all residual strength tests are shown in Table 7-46. The effect 

of flaw length on gross area static failure stress is plotted in Figure 7-48. 

7-64 
ORIGINAL PAGE IS 

POOR QUALITY 

Figure 7-36 End of Slot in Specimen ACOS-1 Showing Crack 

Values for the apparent stress intensity factor, 

that was pulled to failure. The following 

were calculated for each notched specimen 

om reference 32 was used. 

% ?4 = c ./iia(,a 2b tan -) .Ira 

2b 

where KQ = apparent stress intensity factor, MN/m2 J-m (ksi &I.) 

C = gross stress, MNlm2 (ksi) 

2a = crack length, m (in.) 

2b = specimen width, m (in.) 

F,,rure 7-49 shows the variation of KQ with crack length for both fatigued and nonfatigu-rl 

specimens. 

Four composite laminates were studied to determine their flaw growth- and fracture behavior. 

All tests were run at room temperature and moisture was not controlled. All fatigue cycling 

was at R = 0.1 and 30 Hz. Based on the test program and data reduction presented above, the 

following conclusions are drawn: 

1. Under these test conditions, cracks (or effective cracks) in crossplied composite laminates 

grow only when applied cyclic loads are very near the breaking strength. 

7-65 
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7-68 
ORIGINAL PAGE IS 

OF POOR QUALUX 

- 

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7-69 
0/E G /E G /PI B/Al 

65321 7-52 

Figure 7-37 Compliance Specimens (-1's) After Failure in Residual Strength Tests 

ORIGINAL PAGE IS 

QE EOOR QUALITY 

7-70 
ORIGIIV2.L I’dCE IS 

.OF POOR QUALITY 

0 /E G/E G /PI B/Al 

65321 7-53 

Figure 7-38 Compliance Specimens (-2’s) After Failure in Residual Strength Tests 

7-71 
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65321 7-58 

Figure 7-43 LoadKrack Okning Displacement Versus Crack Length 

7-76 
ORIGINAL PAGE IS 

OF POOR QUAJJIX 

BIE GIPI BIAl 

65321 7-59 

Figure 7-44 Compliance Specimens (-3's) After Failure in Residual Strength Tests 

7-77 
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7-78 
ORIGINAL PAGE IS 

OF POOR QUALITY: 

B/E G /E GPI B/Al 

653217-61 

Figure 7-45 Fatigue Specimens (-4's) After Failure in Residual Strength Tests 

7-79 
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cos 

er. 

ro' -?w 

cue 

99 

os 

-Q) 

Q)cD 

00 

00 

lard 

earIC1 

3% eam 

0 

99 

'99'9 

roareaeaeu 

c 

aq 

Q 

E 

.- 

10 

rea 

ea 

ea 

8 

i 

II 

6 

ea 

C 

L 

Y 

3 

a 

m 

E 

5 

6 

.- 

3 

z - 0 

.- % 

2 

E 

Y 

P) 

4 

c 

a 

n 

C 

.- 

0 

n 

m 

2 

5 

m 

7-80 
Table 74. Ruulh of Rosidual Shngih T.rh 

INITIAL NO. OF FINAL 

SPECIMEN FLAW SIZE FATIGUE h FLAW SIZE FAILURE 

IOENTIF1- Zq COMPLIANCE tr LOA0 

MATERIAL CAllON m(id CYCLES m (in) IN (Kip) 

GROSS 1 NET I Kn 

AREA A REA 

2. 

3. 

4. 

The apparent stress intensity factor, Q, calculated from fracture tests does not lie on a 

straight horizontal line when plotted versus visual crack length The testing showed only 

very limited evidence of growth of effective crack length. This leads to the conclusion that 

the fracture mechanics approach is not appropriate for characterizing crack growth in composite materials. Delaminations were not investigated, hence, no conclusions can be made 

regarding that type of flaw. 

Fatigue cycling showed no trend effects on stress intensity factors calculated from residual 

strength testa when compared to results from uncycled static tests. 

Results from the "-3" specimens, Figure 7-43, indicate that the BIAl is more resistive to 

crack opening displacement than the BIE. The GIE and GIPI possess the same resistance 

to crack opening displacement and are less resistant than the B/E. As the two graphite 

reinforced systems were the same, this would indicate that the fibers dominate the fatigue 

mechanism, Le., whether the matrix is epoxy or polyimide is of less importance. The large 

difference between the BlAl and BIE materials, on the other hand, indicates that for boron 

reinforced systems the matrix, either a ductile metal or a brittle resin, plays a major role in 

the fatigue mechanism. 

7.4 COMPRESSIVE TESTING 

Results of the Phase I thermal aging and flight simulation tests indicated that matrix properties such as compressive and shear strength were affected to a greater degree than tensile 

strength, a fiber dominated property. For this reason it was decided to include a limited 

number of residual compressive tests in the 10,000 hour flight simulation test program. To 

check out the feasibility of determining compressive strength of a 6 ply laminate such as that 

7-81 
BIE G /E GPI 

Figure 7-46 Static Strength Specimens (-5's) After Failure 

ORIGINAL PAGE f.3 

OF POOR QUALITY 7-82 

B/Al 

66321 7-62 
BIE G G /PI B/Al /E 

65321 7-63 

Figure 7-47 Static Strength Specimens (-6's) After Failure 

7.83 
0 0 0 0 

Lp v c) H 

I 1 I 

4 

i 

I I I I I I 

a a a a 0 0 0 

(? N u-2 a u-2 a 

N 5 3 

1-81 
u? c 

c 

1 c 

h 

c 

c 

c 

c 

I 

w 

m 

\ 

0 

W 

(1 

\ 

0 

h 

\ 

(1 

0 

x 

m 

x 

N 

0, 

x 

0 

z 

N I 

SI 

x 

UY 

x 

7.85 
used in the flight simulation exposures, a series of baseline compressive tests was conducted 

on laminated BIE, GIE, and GIPI specimens. The results obtained from the bminated specimen 

design were satisfactory, and similar compressive tests were performed on 'the 6 ply thermal 

aginglcompression study materials, Celion 6OOOILARC-160 GlPI and A-SI3501-5A GIE. Compressive tests were also conducted on the 24 ply thermal aginglcompression study materials. 

7.4.1 SPECIMEN DESIGN AND TEST PROCEDURE. At the the that the compressive tests were performed, an ASTM standard test method, D3410-75, had been prepared for determination of. 

compressive properties of oriented fiier composites. The specimen configuration, test fixture, 

and procedure used in the baseline tests were those called out in this ASTM Specification. 

Figure 7-50 shows the test ,specimen and test fixture. To obtain the required specimen thickness of 0.0015 m (0.060 in.) to 0.003 m (0.120 in.) for the 6 ply material it was necessary to bond 

together several pieces. For the Phase I BIE and GIE systems, four layers were used, while for 

the somewhat thicker Phase I HT-9710 GIPI and the thermal agingkompression study 

SPECIMEN FIXTURE 

E 

.I kl t- 

-- EDGE FRONT 

OlMENStONS rn 

A aoom 

B a00635 

C 0.05715 

in. 

ai57 

a250 

2250 

c 

DIMENSIONS rn in 

0 aoi 270 0.500 

E a00318 0.125 

F a06350 2500 

65321 7-67 

Figuro 7-50 Compression Test Specimen and Fixturm 

7-86 
ORIGINAL PAGE 

OF POOR QUALITY 

materials, Celion 6000/LARC-160 GIPI and A-S/3501-5A GIE only three layers were required. 

The procedure for fabrication of the laminated specimens was: 

1. Cut three or four pieces of [O O f 45 OIs composite approximately 0.038 m (1.5 in.) by 0.140 m 

(5.5 in.) with 0" plies in long direction. 

2. Heat at 394 K (250°F) for 24 hours to remove any absorbed moisture. 

3. Bond pieces together with fiberglass doublers at each end using EA 934 adhesive. 

4. Cut 0.0064 m (0.25 in.) wide specimens from the 18 or 24 ply laminates. 

5. Apply two strain gages to each specimen using the procedures described in Section 7.1.1. 

Gages used were BIE, FAE-12-1243; GIE and GIPI, FAE-12-12-SO. 

6. Store in desiccator until tested. 

A similar procedure was used to prepare the compression specimens from the 24 ply thermal 

agingkompression study materials except that no laminating was required. 

The compressive tests were conducted in an Instron testing machine using a fixture of the 

type described in Figure 7-50. The test setup is shown in Figure 7-51. After the specimen is 

loaded into the fixture, a slight preload is applied to exercise the gages and ensure that the 

specimen is properly positioned. The specimen is then loaded incrementally to failure with 

Figure 7-51 Test Setup for Compressive Tests 

7-87 
strain measurement3 recorded automatidy with the same equipment that was used for the 

tensile tests. The output was both a typewritten recording of load versus strain and a punched 

paper tape that could be processed by a computer, programmed to produce a stressstrain plot. 

Compressive modulus values were determined from the stress-strain curves. 

7.4.2 TEST RESULTS AND DISCUSS1ON. Compressive strength and moduhs data for the three 

Phase I resin matrix systems are given in Table 7-47, and typical load-strain awes for each 

system are shown in Figures 7-62 to 7-54. In agreement with results of others, the strength and 

modulus values of the boron reinforced system are considerably larger than either of the. 

graphite reinforced systems. The higher compressive properties of the BIE over the G/E and 

G/PI would be anticipated from a comparison of the reinforcing fibers. The boron fibers in the 

0 plies are Bligned along their entire length in the loading direction, whereas the much smaller 

diameter graphite fibers that are twisted into multi-filament tows are, on a micro scale, 

generally aligned slightly off axis to the compressive load. Both the strength and modulus 

values of the graphite systems are less than half of the values for the boron system. 

Data for the two thermal aginglcompression study materials are given in Tables 7-48 and 7-49. 

like manner to the tensile properties, the compressive properties of the Celion 6000/LARC-160 

G/PI are significantly higher than those of the HT-S/710 system. The laminated GIE data are 

in relatively good agreement for the two batches of material (Tables 7-47 and 7-48). The compressive strength of the 24 ply material, however, was considerably lower than the values 

obtained for the lsminated 6 ply materials. The low values for the 24 ply material can be partially accounted for by a somewhat larger cured ply thickness, 0.00015 m (0.0060 in.) compared 

to 0.00013 m (0.0051 in.) and 0.00014 m (0.0055 in.) for the 6 ply materials. 

8 Examples of typical load-strain curves for each material are shown in Figures 7-55 and 7-56. In 0 

7-88 
fablo 7-47. Baulino Comprorrivo PmpocH.r of Laminahd [0° -C 4S0], Crorrpliod Comporito Systoms at 

Room tompomturo 

Matorial Spocimm Comprorrivo Strongih Compnrrivo Modulus 

Syttom Numbor RAN/& (W GN/d (Msi) 

(4 (4 

550514 BIE ACO4-1 

-2 

-3 

A-Sl3501-5 GIE BC04-1 

-2 

HT-SI710 GlPI DCO4-1 

-2 

-3 

1459 

1665 

1497 

avg 1540 

571 

570 

avg 570 

382 

363 

386 

avg 377 

- 

- 

- 

211.6 

241.5 

217.1 

82.8 

82.6 

82.7 

55.4 

52.6 

56.0 

54.7 

223.4 

- 

- 

aStrain gage measurements 

103 

101 

101 - 

102 

47 

43 

45 7 

46 

48 

44 - 

46 

14.9 

14.7 

14.7 - 

14.8 

6.8 

6.3 - 

6.6 

6.6 

7.0 

6.4 - 

6.7 

653217-69A 

Figuro 7-52 Compnrrivo Load-Strain Diagram for [Oo f fio],21 WE at 297 K (750 F) 

7-89 
3 

a 

- 

0 

0 

-1 

STRAIN (x 10-31 66321 7-70 

Agun 7-53 Compmrivo load-Strain Diagram for [OO f 45O],& G/E at 297 K (750 F) 

I I I I I lo 

1 2 3 4 5 

STRAIN (x 10-3) 

653217-71 

Figun 7-54 Comprerrivo Load-Strain Diagram for [Oo k 45OIr3 HT-S/710 G/PI at 297 K (750 F) 

7-90 
Tablo 7-48. Bardino Compnrsivo Prapdos of Laminated [OO f 4501, Cmrrpliod Composito Syrtomr at 

Room Tompomturo - Thormal AgingKompmrian Study 

Matorial 

svCompmsivo MaduIur 

Sp.U'lMl Comprossivo Strongth GN/& (Msi) 

Numbu MN/& (ksi) (0) (4 

A-Sl3501-5A 

G/E 

BC101-1 536 

-2 514 

-3 376 

avg 542 

Celion 60001LARC-160 FC101-1 622 

G/PI -2 532 

-3 598 

avg 584 

. 0 

aStrain gage measurements 

77.8 

74.6 

83.5 

78.6 

- 

90.2 

77.2 

86.8 

84.7 

- 

55 

53 

52 

53 

50 

52 

54 

52 

8.0 

7.7 

- 7.5 

7.7 

7.3 

7.5 

7.8 

7.3 

- 

Tablo 7-49. Barolino Compmrivo Proportfos of 24 Ply [OO -C 45O],, Cmrrpliod Comporito Systems at 

Room Tomporohrro - Thormal AgingKomprosrion Study 

Matdal 

System 

Compressivm Modulus 

SM- Compmrivo Strongth GN/& (Msi) 

Numbu MN/& (ksi) (4 (4 

A-Sl3501-5A 

G/E 

BC101-4 

-5 

-6 

Celion 60001LARC-160 FC101-4 

G/PI -5 

-6 

447 

467 

402 

439 

565 

556 

652 

591 

64.8 

67.8 

63.6 

- 58.3 

82.0 

80.6 

94.5 

39 

44 

- 

41 

57 

57 

39 

85.7 58 

5.7 

6.4 

- - 

6.0 

8.2 

8.3 

8.6 

8.4 

- 

astrain gage measurements 

7-91 
SPEC. NO.: BC101 

MOOULUS: 44 GN/m2 (6.4 Msi) 

AREA: 2.28 X 10 2 m 2(0.0354 in2) 

4 

3. 

- 1.u 

- 0.8 

- 0.6 2 

P 

0 

f = 

- 

- 0.4 

- 

SPEC NO.: FCOl-5 

AREA: 1.01 X 10-5m2 (0.0156 h2) 

MODULUS: 62 GNld (9.0 Msi) 

65321 7-319 

ngum 7-55 Compmrrivo Load-Stmin Diagmm'for [0°&450]d G/E at 297 K (75OF) 

6S3217-320 

STRAIN (X 

0.8 

0.6 

I0 

P 

0 

0.4 

0.2 

0 

Figuro 7-56 Compressive -load-Strain Diagram for [P -C 45O]d Celion 6000/LARC-160 G/PI 

at 297 K (75OF) 

7-92 
SECTION 8 

THERMAL AGING 

To determine the effects of elevated temperature exposure on the stability of advanced composites an extensive thermal aging study was conducted. It was thought that these data, together 

with the fatigue data also generated during the program, would be very valuable in understanding the results of the flight simulation exposures. 

Initially, the thermal aging program was structured so that tensile strength changes were the 

primary means of following the effects of thermal exposure on the Phase I composite systems. 

Tensile properties were selected because the intent of the contract was to investigate material 

requirements and capabilities for the lower wing surfaces of a supersonic aircraft where the 

major loading is in tension. Early results from the flight simulation tests, however, indicated 

that a more useful measure of the changes occurring during thermal aging of composites might 

be obtained from residual compressive or shear strength testing. These are properties that were 

thought to be more likely to signal the early changes in a material's strength as a result of matrix 

degradation. For this reason a limited study of thermal aginglcompression effects was added 

during Phase 11. 

The thermal qhg section has been divided into three parts. The first is a description of the test 

equipment and the procedures used. The second presents and discusses the thermal aging/ 

tension experimental results. The third part covers the thermal aginglcompression study. 

In the thermal agindtension study all composite materials were aged for 10,000 hours in Phase I. 

Additional exposures, out to 40,000 hours for BlAl and 50,000 hours for the BIE, GIE, and HT-SI 

710 G/PI systems were carried out during Phase 11. At various times during aging, specimens 

were removed for examination and determination of residual tensile strength and, for BIN, also 

residual interfiber shear strength. Test specimens were taken from both unidirectional and 

10" f45"], cro~spliedlamjnates except for Celion 6000ILARC-160 GIPI for which only [O" +45"], 

material waa evaluated. All were 6 ply in thickness. The aging temperatures were: epoxy 394K 

(250°F) and 450K (350°F): HT-SI710 GIPI 505K (450°F) and 561K (550°F): Celion 6000ILARC160 GlPI 450K (350°F): and BlAl450K (350°F). 505K (450"F), 561K (550°F), and 700K (SOOOF). 

The resin matrix composites were aged at ambient pressure and at a reduced pressure of 0.014 

MNlm2 (2 psi) to simulate high altitude flight conditions. The BIN system was aged at ambient 

pressure only. A second oxygen pressure was selected for the thermal aging of the organic composites because. in general, the organics &e oxidation prone. Previous work had shown a direct 

correlation of thermal aging and oxygen pressure on residual strength of resin matrix composites (ref. 33). In actual use, a SCR vehicle would be at very high altitude during much of the time 

at which the structure has reached its maximum temperature. Thus, if oxygen pressure is taken 

into account, the organic materials can likely be used at higher temperatures and for longer times 

than would be predicted from ambient pressure data. 

For the thermal aginglcompression program two composites, GE and Celion 60001LARC-160 

GIPI, were selected for evaluation. The other polyimide, HT-S/710, would have been included in 

the study but was no longer commercially available at that time. Aging was conducted at 373K 

(212°F) and 394K (250°F) for GIE and 450K (350°F) and 505K (450°F) for GIPI for time periods 

8- 1 
out to 10,000 hours. All exposures were done at 1 atmosphere pressure. For both systems 6 ply 

(0" f45"], and 24 ply [Oo f45"], laminates were tested. The post aging evaluation consisted of 

weight change measurements, loom temperature compressive, interlaminar shear, and fatigue 

tests, and T, tests. As a part of the thermal agingkompression study a small number of Celion 

6OOO/LARC-160 G/PI samples coated with a solution of a polyphenylquinoxaline (PPQ) were also 

exposed. Prehmaq work at NASA-Langely Research Center had shown a significant improve 

ment in the long term thermooxidative stability of LARC-16OIGIPI laminates that were coated 

with PPQ. Sets of three specimens each were fabricated, coated, and aged for 10,000 hours at 

450K (35OOF) and 505K (450°F) followed by residual interlaminar shear and compressive 

strength determinations. A set of coated weight change specimens was also prepared for each 

temperature and weighed after 1000: 5000; and 10,000 hours. The PPQ coated specimens were 

exposed in the same thermal aging containers as the uncoated specimens. 

.. 

The following sections describe the test procedures and equipment for the thermal aging studies 

and present and discuss the experimental results. 

8.1 TEST EQUIPMENT AND TEST PROCEDURE 

The 700K (800°F) thermal age of B/AI was carried out in a small, commercial, electrical resistance oven. The built-in "on/off" temperature controller was replaced by a more accurate controller. After modifying the control system, the entire oven was enclosed with ceramic bricks to 

improve the temperature stability over the lengthy aging period. With this arrangement the 

temperature was held to 700 +6K (800" flO°F). 

All other exposures were conducted in specially constructed aging furnaces similar to the sketch 

in Figure 8-1. The heater plates consist of insulated wire sandwiched between two thin aluminum 

plates. The reduced pressure specimens are enclosed in containers consisting of an aluminum 

picture frame with vacuum fittings and thin aluminum covers. The top cover is sealed with a 

weld after the specimens are in place. The top is cut off and rewelded periodically during the 

50,000 hours to remove specimens for testing. This is less expensive and more.reliable than a 

gasket sealed container. The reduced pressure is obtained by a vacuum pump simultaneously 

pumping on one side of each container. Bleed air is valved into the other side of each to maintain 

0.014 MN/m2 (2 psi). This assures a constant fresh supply of air to each container. All furnace 

temperatures were initially equilibrium controlled, Le., a constant amount of power was supplied 

to the heaters. The power equals that lost through the insulation at equilibrium temperature. 

This method does not require the use of expensive controllers, and temperature fluctuations are 

generally smaller than those achieved with active controls. There is also essentially zero probability of a malfunction allowing the specimens to overheat. However, there was a problem that 

developed on very warm days when the temperature could rise slightly above the desired limits. 

To prevent this from happening wide still retaining the assurance that a failure in a power supply, thermocouple, or some control element would not cause overheating, a small modification 

was made. This was to add a controller to the system on about 10% of the available power. The 

furnace would operate as before, except when the outside temperature rose too high the controller would override the system and maintain the desired temperature range. The temperatures 

were generally maintained to f3K (25°F) with infrequent excursions to a maximum of f6K (f 

10°F). 

The procedure followed for the thermal agingkension exposures was to cut the specimen blanks 

to final size including machining of the B/Al interfiber shear specimens (Figure 7-30), drying the 

8-2 
GQC-CXAQ-85-005 

2 PSI 

HEATER PLATE 

VACUUM CONTAINER 

REOUCEO PRESSURE SPECIMENS 

SPACERS 

DOQ I ATM SPECIMENS 

HEATER PUT€ 

HEATER POWER 

0 THERMOCOUPLE 

613217-72 

Aguro 8-1 Thormal Aging Furnace Configuration 

res& matrix composites at 394K (250°F) for 24 hours, and loading into the aging furnaces. The 

specimens were supported at their ends by narrow strips of stainless steel so that almost the 

entire surface was exposed to the air atmosphere within the furnaces. At the required time inter 

vals the ovens were shut down, opened, and the specimens were removed and stored in a desiccator until doubler bonding and tensile testing. Only the resin matrix specimens were placed in the 

desiccator, as moisture absorption was not a problem with the BlAL Residual strength testing 

was performed at 297K (75°F) for the BIAI, 450K (350K"F) for the epoxies, 505K (450°F) and 

561K (550°F) for HT-S/710 G/PI, and 450K (350°F) for Celion 6000/LARC-160 G/PI. The tensile 

and shear specimen configurations and test procedures were the same as those described in Section 7. 

The procedure for the thermal aging/compression exposures differed slightly from that for the 

thermal aging/tensile exposures. Because of the requirement of performing shear and compressive testing on 6 ply material it was necessary to use the technique described earlier of bonding 

three pieces together to obtain specimens of sufficient thickness. For this reason oversize sample 

blocks were aged rather than precut specimens. The dimensions of these aging samples were: 

compression, 0.038 m by 0.14 m (1.5 in. by 5.5 in.) and shear, 0.038 m by 0.038 m (1.5 in. by 1.5 in.). 

The same sizes were used for the 24 ply material with test specimens again being machined after 

aging. This method may have eliminated edge effects of oxidation that would have been present 

for the thermal agingkension specimens. The tension-compression (R = -1) fatigue specimens 

were also prepared after aging. These were cut from 0.10 m by 0.28 m (4 in. by 11 in.) aging blanks 

of 24 ply material. Fatigue strengths of the 6 ply materials were not evaluated. The specimen 

configurations and test procedures were the same as those described in Section 7 for compression 

and shear and Section 11 for fatigue. In addition to mechanical property determinations, 

8-3 
GDC-CRAD-85-00s 

periodic weight change measurements were performed during the course of aging using 0.038 m 

by 0.038 m (1.5 in by 1.5 in.) samples of both the 6 ply and 24 ply materials. Finally, 0.038 m by 

0.038 m (1.5 in. by 1.5 ~IL) pieces of 6 ply laminates were exposed for determining T, changes and 

to use for metallographic examination after aging. 

In addition to the Celion 6000/LARC-160 specimens described above, a number of 6 ply specimens were coated with PPQ to evaluate its ability to improve the long term thermooxidative 

stability of the GlPI laminates. The PPQ was applied at a solution concentration of 12% in a 

mixed solvent of meta-cresol, mixed xylenes, and chloroform The following procedure was used 

for applying the solution to the specimens and for subsequent solvent removal. 

1. 

2. 

3. 

4. 

Thoroughly wipe the surface of the laminate with acetone using a lint-free cloth and, thereafter, handle with lint-free gloves. 

Dry in a forced air oven at 473K (392°F) for 2 hours. 

Apply the PPQ solution to the cooled laminate within a few hours after drying by brush coating a light coat using natural bristle brush in a well-ventilated hood. It was necessary to coat 

one side, dry, and then coat the other side. 

Remove the solvent mixture by stagedrying in a forced air oven from room temperature to 

333K (140°F) in 1/2 hour, hold at 333K (140°F) 1 hour, increase the temperature to 363K 

(194°F) and hold for 1/2 hour, increase the temperature to 393K (248°F) and hold for 1/2 hour* 

increase the temperature to 433K (320°F) and hold for 1/2 hour, and increase the temperature 

to 473K (392°F) and hold for 2 hours. After drying at 473K (392"F), cool to room temperature 

in the forced air oven relatively fast. 

8.2 THERMAL AGING/TENSION - TEST RESULTS AND DISCUSSION 

8.2.1 BORON/EPOXY AND Gi?APHITE/EPOXY. Residual tensile strength data for the epoxy systems are presented in Tables 8-1 through 8-4 for B/E aged at ambient pressure, Table 8-5 for BIE 

aged at reduced pressure, Tables 8-6 through 8-9 for GIE aged at ambient pressure, and Table 8- 

10 for G/E aged at reduced pressure. Agingwas carried out at 394K (250°F) and 450K (350°F) for 

both unidirectional and [O" f45"], crossplied specimens. The tensile strength data have also 

been plotted as a function of aging time for the two epoxy systems. These curves are found in 

Figures 8-2 through 8-7. A summary of the effects of 10.000 hours and 50,000 hours of thermal 

aging on the tensile strength retention of BE and G/E is shown in Table 8-1 1. In Table 8-1 1 and 

in others throughout the report that list percentage strength retention, whenever residual 

strength is greater than baseline strength a value of 100% retention has been used. 

394K, 0.10 MN/m2 (25OoF, 14.7 psi) EXPOSURE. After 25,000 hours of aging the appearance of the 

BIE and G/E specimens had not changed appreciably. Compared with unexposed material, the 

only observable differences were slight color changes that were first noted after 10.000 hours. No 

surface damage was found. After 50,000 hours the specimen condition still appeared to be excellent. Some shallow cracking at the edges of the crossply specimens was visible but did not appear 

to be of much consequence. During the cure of the doubler adhesive at 450K (350°F). however, 

delamination of the 50,000 hour crossplied G/E specimens occurred. Similar effects were 

observed for the G/E specimens aged at 450K (350°F) in one atmosphere air for 10,000 hours (see 

Figure 8-12). In neither case was the damage sufficient to preclude tensile testing. 

8-4 
Tabk 8-1 Thormal Aging Data for [Ool6 B/E in 0.10 MN/m2 (14.7 psi) Air 

394 250 450 350 Baseline avg 1380 

100 1390 

1350 

1370 

avg 1370 

- 

500 

1.000 

5,000 

1460 

1360 

1340 

avg 1390 

1330 

1390 

1380 

avg 1370 

' 1480 

1450 

1520 

- 

- 

10,000 

avg 1480 

1410 

1430 

1450 

25,000 

avg 1430 

1240 

1400 

1390 

50.000 

avg 1340 

1280 

1320 

1300 

avg 1300 

200 

202 

196 

198 

199 

212 

197 

194 

201 

193 

201 

200 

198 

214 

210 

220 

215 

204 

208 

211 

208 

180 

203 

202 

195 

186 

192 

188 

189 

- 

- 

- 

- 

- 

- 

- 
GOC-CRAO.85405 

Tablo 8-2 Thormal Aging Data for [Oo +45°], Crossply WE in 0.10 MN/m2 (14.7 psi) Air 

Tonsilo Stnngth TUt Aging 

T-m T.mp.rotum nnn (OF) K (Of) (hr) MN/d (kri) K 

250 450 350 Baseline avg 550 79.8 394 

100 310 73.9 . 496 71.9 

493 71.5 

avg 500 72.4 

478 69.4 

428 62.1 

474 

avg 460 66.7 

512 74.2 

504 73.1 

539 78.2 

avg 518 75.2 

535 77.6 

527 76.4 

533 77.3 

avg 532 77.1 

538 78.1 

534 77.4 

522 75.7 

avg 531 77.1 

443 64.2 

462 67.0 

404 58.6 

avg 436 63.3 

428 62.1 

388 56.3 

65.0 

avg 421 61.1 

- - 

- - 68.7 

- - 

- - 

- - 

- - 

- 448 - 

500 

1,000 

5,000 

10,000 

25,000 

50,000 

a-6 
GOC-CRAD-85-005 

Tablo 8-3 Thoma1 Aging Data for [Oob WE in 0.10 MN/m2 (14.7 psi) Air 

Tonsilo Strongth 

MN/mZ (kri) 

450 350 450 350 Baseline avg 1380 

5.000 1230 

1240 

1250 

avg 1240 

- 

10,000 1010 

1010 

1080 

avg 1030 

25,000 1130 

1160 

1030 

avg 1110 

- 

- 

50,000 986 

1010 

1000 

avg 1000 

- 

200 

179 

180 

182 

180 

146 

147 

156 

- 

150 

164 

168 

150 

161 

- 

143 

147 

145 

145 

- 

8-7 
Tablo 8-4 Thonnal Aging Data for [Oo -C4S0], Crossply B/E in 0.10 MN/m2 (14.7 psi) Air 

Tonsil. Shngth 

MN/m2 (kri) 

450 350 450 350 Baseline avg 550 79.8 

500 544 78.9 

46 1 66.9 

1.000 

5,000 

10,000 

25,000 

50,000 

517 

avg 507 

- 

581 

508 

493 

avg 527 

310 

319 

309 

avg 313 

339 

342 

- 

- 

75.0 

73.6 

- 

84.2 

73.7 

71.5 

76.5 

44.9 

46.3 

44.8 

45.3 

49.2 

49.6 

- 

- 

45.6 

avg 332 48.1 

310 45.0 

293 42.5 

301 43.7 

- 314 - 

- - 

avg 301 43.7 

325 47.2 . 297 43.1 

268 38.8 

avg 296 43.0 

- - 

8-8 
GDC-CRAD-85-005 

Tablo 8-5 thormal Aging Data for B/E Aged in 0.OlA MN/m2 (2 psi) Air 

450 

450 

350 [0"16 450 350 Baseline avg 

5,000 

10,000 

25,000 

50,000 

350 [O f4501, 450 350 Baseline avg 

5,000 

1380 

1520 

1290 

1400 

200 

220 

187 

203 

avg 1400 

1390 

1340 

1400 

avg 1380 

979 

1010 

1090 

avg 1030 

958 

1090 

1210 

avg 1090 

550 

519 

529 

590' 

- 

- 

- 

203 

201 

195 

203 

200 

142 

146 

158 

149 

139 

158 

176 

158 

79.8 

75.3 

76.7 

85.6 

- 

- 

- 

10,000 

25,000 

50,000 

avg 546 

379 

367 

381 

avg 376 

290 

297 

355 

avg 314 

332 

327 

350 

avg 336 

- 

- 

- 

79.2 

54.9 

53.2 

55.3 

54.5 

42.0 

43.1 

51.5 

- 

45.5 

48.1 

47.4 

50.8 

48.8 

- 
Tablo 8-6 Thoma1 Aging Data for [O"], G/E in 0.10 MN/m2 (14.7 psi) Air 

Tonsilo Strangth Aging T& Aging 

T.mpOfOtUm T.m~f0 Tim 

K (OF) K (OF) (hr) MN/m2 (kti) 

394 250 450 ' 350 Baseline avg 1590 230 

100 1520 220 

1590 230 

1540 224 

avg 1580 225 

1410 205 

1270 184 

1470 213 

- - 

- - 

avg 1390 20 1 

1480 214 

1560 227 

1450 210 

avg 1500 217 

1630 236 

1600 232 

1440 209 

avg 1560 226 

1650 240 

16%0 239 

1850 268 

avg 1720 249 

1670 242 

1650 239 

1520 220 

avg 1610 234 

1560 226 

1500 217 

1510 219 

- - 

- - 

- - 

- - 

- - 

avg 1520 221 

500 

1,000 

5,000 

10,000 

25,000 

50,000 

8-10 
GDC-CXAD-85905 

Tablo 8-7 Thormal Aging Data for [Oo -CdS0], Crorrply G/E in 0.10 MN/m2(14.7 psi) Air 

Tonsil. SWenqih 

MN/m2 . (kri) 

394 250 450 350 Baseline avg 500 72.5 

100 . 496 72.0 

542 78.6 

483 70.0 

avg 507 73.5 

- - 

500 475 68.9 

492 71.4 

- 494 - 71.7 

1,000 

5,000 

10,000 

avg 487 

494 

552 

626 

avg 557 

589 

559 

536 

avg 561 

559 

593 

585 

- 

- 

25,000 

avg 579 

405 

492 

70.7 

71.6 

80.0 

90.8 

80.8 

85.4 

81.8 

77.7 

81.4 

81.1 

86.0 

84.9 

84.0 

58.8 

71.3 

- 

- 

- 

- 44 1 - 63.9 

avg 446 64.7 

50,000 245 35.5 

258 37.4 

287 41.6 

avg 263 38.2 

- - 
GDC-CRAD-85905 

Tabk 8-8 fhonnal Aging Data for [Ool6 G/E in 0.10 MN/& (14.7 psi) Air 

450 

~ 

350 450 350 Baseline avg 1590 

100 1500 

1650 

1460 

avg 1540 

500 1630 

1620 

1550 

avg 1600 

1,000 1260 

1500 0 1540 

avg 1430 

1210 

1370 

1300 

avg 1290 

1360 

1270 

1210 

avg 1280 

958 

1010 

1070 

avg 1010 

50,000 986 

952 

945 

avg 958 

- 

- 

- 

- 

- 

- 

- 

5,000 

10,000 

25,000 

230 

218 

239 

212 

223 

236 

235 

225 

232 

182 

218 

223 

208 

175 

199 

188 

188 

198 

184 

176 

- 

- 

- 

- 

186 

139 

147 

155 

147 

143 

138 

137 

139 

- 

- 

8-12 
GDC-CRAD-85-005 

Tablo 8-9 Thrrmal Aging Data for [Oo -CaSo], Crossply G/E in 0.10 MN/d (14.7 psi) Air 

Tomile Strength 

MN/m2 (kri) 

450 350 450 350 Baseline avg 500 72.5 

500 494 71.6 

608 88.2 

559 81.1 

avg 554 80.3 

- - 

1.000 

5.000 

10,000 

552 

501 

563 

avg 539 

485 

440 

474 

avg 466 

265 

60 

238 

- 

- 

25.000 

50,000 

avg 188 

253 

316 

321 

avg 296 

251 

194 

272 

avg 239 

- 

- 

80.0 

72.6 

81.6 

78.1 

70.3 

63.8 

68.7 

- 

67.6 

38.5 

8.7 

34.5 

27.2 

36.7 

45.9 

46.5 

43.0 

36.4 

28.2 

39.5 

34.7 

- 

- 

- 

8-13 
Tablo 8-10 Thorrnai Aging Data for G/E Agod in 0.014 MN/m2 (2 psi) Air 

450 350 toole 450 350 Baseline avg 

5,000 

10,000 

25.000 

50,000 

450 350 [O IC,45qs 450 350 Baseline avg 

5,000 

10,000 

25,000 

50,000 

1590 

1730 

1740 

1560 

avg 1680 

1770 

1870 

1760 

avg 1800 

1630 

1620 

1440 

avg 1560 

1400 

1020 

1170 

avg 1200 

500 

584 

596 

571 

avg 584 

535 

498 

548 

avg 527 

432 

476 

485 

avg 465 

358 

340 

309 

avg 336 

- 

- 

- 

- 

.- 

- 

- 

- 

230 

251 

252 

227 

243 

257 

271 

255 

261 

236 

235 

209 

227 

203 

170 

174 

72.5 

84.7 

86.4 

82.8 

84.6 

77.6 

72.3 

79.5 

76.5 

62.7 

69.1 

70.3 

67.4 

51.9 

49.3 

44.8 

48.7 

- 

- 

- 

148 

- 

- 

- 

- 

- 

8-14 
GDC-CRAD-85-005 

240 

200 

160 

120 

80 

40 

0 

0.1 1 10 100 1000 10,000 100,000 

AGING TIME (HR) 

c 

r 

i 

z 

m 

--I 

m 

2 

v) 

7 

m 

a 

m 

z 

0 

-I 

I 

H. 

5 

Y 

x 

- 

65321 7.342 

Flguro 8-2 Tonsil. S)t.rrgth of WE at &OK (SOOF) Aftor lhrrnal Aging at 394K (250OF) 

240 

200 

c 

r 

=! 

= 

160 2 

rn 

--I 

m 

z 

120 

v) + 

1J 

z 

I 

m 

80 2 

- z 

40 

0 

0.1 1 10 100 1000 10.000 100,000 

AGING TIME (HR) 65321 7-343 

Figure 8-3 Tensile Strength of WE at 45OK (350'F) Aftor Thormal Aging at the Same Temperature 

8-15 
GOC-CRAO-85905 

240 

200 

c 

I2 

z 

160 5 

m 

2 

4 

2 

120 F 

v) 

-I 

z 

m 

z 

c) 

80 2 

x 

- D. 

40 

0 

0.1 1 10 100 1000 10,000 100,000 AGING TIME (HRI 

6S3217-344 

Flguro 8-4 Tonsilo Stnngth of B/E at 450K (350°F) Aftor Thormal Aging at tho Sam. Tomporaturo 

1750 

240 

1500 

200 

1250 

160 

1000 

120 

750 

80 5 00 

250 40 

0 0 

0.1 1 10 100 1.000 10,000 100,000 

AGING TIME (HR) 

c 

r 

i 

E 

m 

4 

m 

z 

5", 

r 

m 

v) 

-I 

z 

z 

2 

m 

65321 7-345 

Ffguro 8-5 Tonsil. Strength of G/E at 45OK (350°F) Aftor Thermal Aging at 394K (250°F) 

8-16 
GOC-CRAD-85-005 

240 

200 

c 

I160 s 

m 

-I 

m 

z 

r 

m 

v) 

-I 

a 

80 2 

2 

120 2 

4 

I 

- E 

40 

0 

0.1 1 10 100 1000 10,000 100,000 

AGING TIME (HR) 65321 7-346 

Rguro 8-6 Tmsik Strongth of G/E at 45OK (350°F) Aftor Thormal Aging at tho Same Temperature 

1750 

240 

c 

I 2 

9 

m - 1250 X 

t 

c1 160 2 

z 

2 

z 

Y 

7 

a 1000 

k m 

1 

z 750 

- 1500 (Y 

\ E 200 2 

5 

120 2 Lu 

VI 

u 

c 

Y + 

m 

z 

D + 

I 80 - 2 - E 500 

a 

- 

5 

40 250 

0 0 

0.1 1 10 100 1000 10,000 100,000 

AGING TIME (HA) 653217.347 

Figuro 8-7 Tensile Strength of G/E at 450K (350'F) After Thermal Aging at the Same Temperature 

8-17 
GOC-CRAD-85905 

Tabio 8-1 1 Summary of Tonsilo Strmgth Rotmion Aftor Thormai Aging of B/E and G/E 

MohriOl Aging Tomp Aging Pmswo Tost Tmnp Rotontion of Tonsilo Smgth (%) 

Systom Oriontation K (OF) MN/mZ (psi) K (OF) 10,000 hn 50,000 hn 

B/E [00i6 

[00i6 

[0° +450], 

[Oo f45"], 

[oo16 

[0° 245'1, 

GIE [0oi6 

(0' -C45O], 

[oo16 

[0° f45"], 

[w6 

[0° f45'], 

394 

394 

450 

450 

450 

450 

394 

394 

450 

450 

450 

450 

250 

250 

350 

350 

350 

350 

250 

250 

350 

350 

350 

350 

0.10 

0.10 

0.10 

0.10 

0.014 

0.014 

0.10 

0.10 

0.10 

0.10 

0.014 

0.014 

14.7 

14.7 

14.7 

14.7 

2 

2 

14.7 

14.7 

14.7 

14.7 

2 

2 

450 

450 

450 

450 

450 

450 

450 

450 

450 

450 

450 

450 

350 

350 

350 

350 

350 

350 

350 

350 

350 

350 

350 

350 

100 

97 

75 

60 

100 

68 

100 

100 

81 

38 

100 

100 

94 

77 

72 

54 

79 

61 

96 

53 

60 

48 

76 

67 

8-18 
The residual tensile strengths after aging at 394K (250°F) are shown in Figures 8-2 and 8-5. For 

the first 10,000 hours of aging, no degradation of the 350°F tensile properties was observed for 

either system. After 25,000 hours, however, a decrease in tensile strength of the crossply mate 

rial, both BIE and GIE, had occnrred. The effect was not large but appeared to be a real one. Some 

slight drop in the unidirectional tensile strength of the B/E was also found. Additional decreases 

in strength were observed for the 50,000 hour specimens, particularly the crossplied GIE. 

While'no changes in appearance, other than slight color differences, were noticed following aging 

for 25,000 hours, there were signiscant changes after tensile testing. The failed tensile coupons, 

particularly the GIE, were markedly different than those from material that had been aged for 

times up through 10,000 hours. The epoxy matrix at the surface of the 25,000 hour G/E specimens had broken up during test and mumbled off, revealing the the graphite fibers in the outer 

plies. At failure, the graphite tows opened up, giving a very fuzzy appearance to the test section 

of the specimens. The specimens aged at 450K (350°F) in one atmosphere air showed this effect 

to some degree after 5000 hours. Specimens aged 10,000 hours at 450K (350'F) were almost identical to the ones described above that were aged 25,000 hours at 394K (250'F). The appearance of 

the failed specimens is supporting evidence that the slight fallsff in tensile properties after 

25,000 hours marks the beginning of significant degradation to the epoxy matrix specimens 

when aged at 394K (250°F) in one atmosphere air. After 50,000 hours of exposure the failed GIE 

tensile specimens had lost not only the resin at the surface but much of the matrix throughout 

the thickness. Some changes in appearance from those after 10,000 hours were also observed in 

the as-tested B/E specimens aged 25,000 hours at 394K (250°F). There was very little damage to 

the 10,000 hour specimens after tensile testing except near the fracture location. For the 25,000 

hour specimens, however, the matrix at the surface was beginning to show evidence of breaking 

up during testing, revealing some of the boron fibers. The effect was not as pronounced as that 

seen in the GIE specimens after 25,;OOO hours. The breaking up and loss of the epoxy matrix dur 

ing tensile testing increased considerably for the 50,000 hour B/E specimens. Numerous boron 

fibers were visible in these specimens. In comparison with other exposures, the specimen condition was intermediate between the 10,000 hour, 450K (350"F), 0.014 MNlm2 (2 psi) and the 5000 

hour, 450K (350"F), 0.10 MNIm2 (14.7 psi) BIE specimens. 

Portions of faded test specimens before and after aging at 394K (250°F) are shown in Figure 8-8 

for BIE and Figure 8-9 for G/E. Evidence of the excellent condition of the B/E after 25,000 hours 

and the G/E after 10,000 hours can be seen by comparing the failed tensile specimens before and 

after aging. No significant differences are visible. On the other hand, the change in appearance of 

the failed specimens between 25,000 hours and 50,000 hours of exposure for B/E and between 

10,000 hours and 25,000 hours of exposure for G/E is quite distinct. 

45OK, 0.10 MN/m2 (SOOF, 14.7 psi) EXPOSURE. Thermal aging of the epoxy systems at 450K 

(350°F) produced significant degradation when exposure was carried out at ambient pressure. 

Degradation in tensile properties combined with a siwcant change in appearance of the specimens, particularly after tensile testing, occurred after 5000 hours. For these specimens the 

epoxy resin matrix, in many areas, crumbled away during tensile testing - exposing the boron or 

graphite filaments. Portions of failed test specimens of each material before and after aging at 

450 K (350°F) are shown in Figures 8-10 and 8-11. The change in appearance of the failed specimens between 1000 hours and 5000 hours of exposure is quite distinct. While not apparent from 

Figure 8-10 and 8-11, the loss of resin after 5000 hours was somewhat greater for the B/E than for 

the GIE material. One other observation was made during bonding of doublers to the 5000 hour 

tensile specimens. Prior to applying the adhesive, the ends were lightly polished with Scotch8-19 
GDC-CRAD-85-005 

653217-340 

Figure 8-8 B/E Before and After Thermal Aging at 394K (25OOF) in One Atmosphere Air and Tensile Testing 

at 450K (SOOF) 

8-20 
GDC-CRAD-85-005 

c 

ORIGINAL PAGE IS .I . 

50,000 HOURS ~- UNEXPOSED 10.000 HOURS 25,000 HOURS 

653217-349 

Figure 8-9 G/E Before and After Thermal Aging at 394K (25OOF) in One Atmosphere Air and Tensile Testing 

at 450K (350OF) 

8-21 
GDC-CRAD-85-005 

UNEXPOSED 1000 HOURS 5,000 HOURS 10,000 HOURS 25,000 HOURS 50,000 HOURS 

6532 17-350 

Figure 8-10 B/E Before and After Thermal Aging at 450K (350'F) in One Atmosphere Air and Tensile Testing 

at the Same Temperature 

8-22 
GDC-CRAD-85-005 

1.1 

UNEXPOSED 1000 HOURS 5000 HOURS 10,000 HOURS 25,000 HOURS 50,000 HOURS 

653217-351 

Figure 8-1 1 G/E Before and After Thermal Aging at 45OK (350°F) in One Atmosphere Air and Tensile 

Testing at the Same Temperature 

8-23 
GOC-CRAD-85-005 

^" . , 

DELAMINATION OCCURRED AFTER AGING DURING CURE OF DOUBLER ADHESIVE. 

653217-81 

Figure 8-12 G/E Tensile Specimens After 10,000 Hours of Thermal Aging at 450K (35OOF) in One 

Atmosphere Air , 

Brite. For the BIE samples exposed at ambient pressure, just a slight amount of rubbing 

removed the outer powdery layer of resin and revealed the boron fibers of the outer plies. Such an 

effect was not observed for the GIE specimens or the reduced pressure B/E specimens. 

After 10,000 hours of thermal aging at 456K (350°F), further degradation of the epoxy materials 

was observed. The GIE specimens showed a slight color change (to a darker brown) but, otherwise, were visually unchanged by the 10,000 hour exposure. However, during doubler bonding 

the [Oo f45'], specimens aged at ambient pressure all delaminated. Figure 8-12 shows the poor 

condition of these specimens. The tensile strength of these crossplied specimens was very low, 

while theunidirectional specimens showed a decrease of about 25%. The appearance of the specimens after test was similar to those aged for 5000 hours at 450K (350"F), see Figure 8-11. 

The epoxy matrix crumbled away during tensile testing leaving the bare graphite tows looking 

like a bundle of yarn. The relatively high tensile strength of the Unidirectional specimens, 1680 

MN/m2 (186 ksi), after 10,000 hours exposure is somewhat misleading as to the actual quality of 

the material. A transverse tensile or shear test that measures the matrix strength would 

undoubtedly give extremely poor results. 

The BIE specimens aged in ambient air were in very poor condition after the 10,000 hour exposure at 450K (350°F). Two of the crossplied specimens were delaminated when removed from the 

aging furnace, and slight flexing of the third specimen during handling also resulted in delamination. Figure 8-13 shows the two delaminated crossplied specimens prior to doubler bonding. For 

the unidirectional specimens much of the resin was gone from the surface, exposing many of the 

outer boron filaments. Some of these outer filaments were lost during the doubler bonding opera8-24 

ORIGINAL PAGE IS 

$33 POOR QUALITY 
GDC-CRAD-85-005 

ORIGINAL PAGE IS 

'a mR QUALITY 

DELAMINATION OCCURRED DURING AGING 

Figure 8-13 Crossplied B/E After 10.000 Hours of Thermal Aging at &OK (SOOF) in One Atmosphere Air 

tion. The unidirectional specimens had very little transverse stiffness and could easily be bent by 

hand. Again, the crossplied specimens showed a greater decrease in tensile properties than the 

unidirectional specimens. In like manner to the GIE material, the tensile results don't indicate 

how badly the BIE had been degraded by the 10,000 hours in ambient air at 450K (350°F). The 

resin matrix has retained almost no strength after the aging exposure. Figure 8-10 shows failed 

tensile specimens after various exposure times at 450K (350°F) in ambient air. The 10,000 hour 

specimens consist primarily of broken boron fibers and scrim cloth with much of the epoxy 

matrix having crumbled and fallen away from the test section between the doublers. 

In addition to the slight color changes, first observed after 10,000 hours, the GIE showed a few 

small edge cracks after 25,000 hours at 450K (350°F). Otherwise, the appearance was visually 

unchanged by the exposure. After 50,000 hours the surface of the specimens had a burned 

appearance with some loose graphite fibers at the edges. There was some slight edge cracking 

and indications that delamination might be starting. The specimens gave off a dead sound when 

dropped on a hard surface rather than the typical ringing sound associated with unaged mate 

rial. The specimens offered little resistance to bending. 

In contrast to the GIE material the BIE specimens were in very poor condition after the 25,000 

hour exposure at 450K (350°F). The matrix, because of severe oxidation, wasirery brittle and 

crumbled away at the touch. All specimens showed edge cracks, exposed and loosened surface 

fibers, and color changes. One of the specimens delaminated into three separate layers as a result 

of the handling required to perform the visual examination. After 50,000 hours, edge cracking, 

8-25 
loss of resin from the surface, and loose fibers had all increased. Several of the specimens were 

delaminated, and all offered little resistance to bending. In like manner to the G/E system the 

specimens gave off a dead sound when dropped on a hard surface. 

Because of the poor condition of the specimens after 25,000 hours and 50,000 hours of exposure, 

a room temperature cuing adhesive was selected for doubler bonding, and the tensile tests at 

450K (350°F) were conducted using a short ring furnace. With this furnace, only the center por 

tion of the specimens was heated to 450K (350°F), while the grips remained very close to the 

ambient temperature. Even so, all of the B1E specimens dnlaminated in the central section 

between the doublers during the doubler bonding operation. Two of the 25,000 hour and all three 

of the 50,000 hour crossply GIE specimen3 also delaminated during doubler bonding. Crossply 

specimens have typically shown more tendency toward delamination after thermal aging. In 

spite of the extensive delamination ail of these epoxy specimens could be tensile tested. These 

tensile results, Figures 8-3 and 8-6, are similar to those after 10,000 hours of aging and give no 

indication of how severely the BIE and GIE composites had been degraded by the 25,000 and 

50,000 hours of exposure. The resin matrix in each system had retained almost no strength as a 

result of the aging. A test that measures matrix strength, e.g., shear, compressive, or transverse 

tensile would have produced results very dose to zero. After test, the failed BIE specimens, Figure 8-10. consisted primarily of broken boron fibers and scrim cloth with most of the epoxy 

matrix having crumbled and fallen away from the test section between the doublers. For the G/E 

specimens, Figure 8-11. the resin matrix also crumbled away during tensile testing, leaving the 

bare graphite tows looking like a bundle of yarn. 

*OK, 0.014 MN/m2 (SOOF, 2 psi) EXPOSURE. Thermd agingof the epoxy systems at 450K (350°F) 

for 5000 hours produced significant degradation when exposure was carried out at ambient pressure but almost no effects when aging was done at a reduced pressure of 0.014 MNIm2 (2 psi). 

Specimens of both BE and G/E exposed at the reduced pressure showed essentially no reduction 

in 450K (350°F) tensile strength and fded in a manner typical of unexposed materid 

After 10,000 hours of aging the GIE specimens still retained 100% of the 450K (350°F) tensile 

strength and had a failure appearance after tensile testing similar to that of unexposed material. 

In contrast to the G/E material, the BIE specimens aged for 10,000 hours at 450K (350°F) and 

0.014 MNIm2 (2 psi) showed both a reduction in strength and a change in failure appearance. The 

unidirectional specimens retained 100% of the 450K (350°F) tensile strength, but the crossplied 

specimens only approximately 70% (Figure 8-4). Many bare filaments were visible in the failed 

specimens of both layups where the resin matrix had broken up. The effect was not as pronounced as that observed for the 450K (350°F) and 0.10 MNIm2 (14.7 psi) tests, but, compared to 

unexposed specimens where almost no bare filaments occur, the indications are clear that the 

matrix is starting to degrade even at reduced pressure. The difference between the 5000 hour and 

10,000 hour failure appearance can be seen in Figure 8-14. 

The first change in appearance after aging at reduced pressure but before tensile testing was 

observed for the 25,000 hour specimens. Both the BIE and GIE showed slight color changes and a 

few small edge cracks. After 50,000 hours the BIE gpecimens showed considerable degradation 

in the form of loss of resin, loosened boron fibers, delamination, and a dead sound when dropped. 

The condition was similar to the specimens exposed at one atmosphere for the same time and 

temperature. The GIE specimens, on the other hand, were not much different from the 25,000 

hour group other than a somewhat burned looking surface and slightly less ring when dropped. 

8-26 
GDC-CRAD-85-005 

8-27 
Tensile results for the specimens aged at reduced pressure are plotted in Figures 8-4 and 8-7. The 

degradation in 450K (350°F) tensile strength and the appearance of the failed B/E specimens 

after both 25,000 and 50,000 hours were very nearly the same as those for the one atmosphere 

specimens. At a reduced pressure of 0.014 iMN/m2 (2 psi) air, B/E begins to experience some 

matrix degradation between 5000 and 10,000 hours. By 25,000 hours, the degradation, as shown 

in Figure 8-14, is very severe. For the G/E specimens aged in 0.014 MNlm2 (2 psi) air, a slight 

decrease in tensile strength after 25,000 hours accompanied by a change in appearance of the 

failed test specimens indicates the beginning of matrix degradation. After 50,000 hours the tensile strength had decreased considerably, and the appearance of the failed specimens was much 

worse than the 25,000 hour group. Fded tensile specimens before and after thermal aging are 

shown in Figure 8-15. The resin loss and amount of graphite fibers exposed after 50,000 hours of 

aging at reduced pressure were about the same as observed for the 5000 hour, 450K (350°F), 0.10 

MNlrn2 (14.7 psi) specimens after tensile testing (see Figure 8-11). 

In summary, at 394K (250°F) and 0.10 MN/m2 (14.7 psi) matrix embrittlement of the B/E and 

G/E systems began between 10,000 and 25,000 hours, and slight decreases in tensile strength 

occurred. The effects were worse for the GIE. After 50,000 hours of exposure matrix embrittle 

ment was severe for both systems. At 450K (350°F) and 0.10 MNlmz (14.7 psi) matrix embrittle 

ment of the BIE and G/E systems began between 1000 and 5000 hours, and at the same time 

tensile strength losses were obsenred. Severe matrix embrittlement was found for all exposures 

from 5000 to 50,000 hours. For thennal aging at 450K (350'F) and 0.014 MNlm2 (2 psi), matrix 

embrittlement of the B/E system began between 5000 and 10,000 hours and was severe by 25,000 

hours. The G/E system was somewhat better with the beginning of embrittlement starting 

between 10,000 and 25,000 hours. By 50,000 hours the effect was severe. 

8.2.1.1 Microscopic Examination of tho Thermal Aging Specimens. After tensile testing, many of 

the thermai aging specimens were sectioned and mounted for study using both an optical microscope (metdograph) and scanning electron microscope (SEMI. These studies were intended to 

detect changes that occurred in the composites during exposure in order to assist in identifying 

degradation mechanisms. Additional examinations were conducted of the fractured surfaces 

(primarily with the SEM) to determine failure modes and to further study the degradation processes. The most conclusive results were obtained from examination of the G/E specimens, and, 

for this reason, the majority of this section discusses the G/E system. 

Figure 8-16 shows photomicrographs of GIE aging specimens using a metallograph and a SEM 

at lOOX and 500X magnifications. These pictures are for 10,000 hours at 450K (350°F) and 0.014 

MNlm2 (2 psi) pressure. Close examination of these photomicrographs shows only surface 

changes in the resin and some cracking, which probabiy occurred during post exposure tensile 

testing. Some of the cracks are located at boundaries between lamina and others within a particular lamina, as shown in the upper two photomicrographs. The lighter cast at the specimen edges 

in the SEM pictures will be shown later to be oxidation. Here it has occurred only at the surfaces 

exposed to oxygen. 

The pictures in Figure 8-17 taken at lOOX with the metallograph show good examples of how 

oxidation of the epoxy resin can proceed inward as a function of time and temperature. The upper 

pictures show that there is little difference between the as-received material and that aged for 

10,000 hours 394K (250°F) in one atmosphere air. The lower pictures, on the other hand, show 

extensive degradation in the outside plies after 5000 hours at 450K (350"F), and considerable 

degradation in both the inner and outer layers after 10,000 hours. The SEM pictures in 

8-28 
GDC-CRAD-85-005 

UNEXPOSED 10,000 HOURS 25,000'HOURS 50,000 HOURS 

65321 7-353 

Figure 8-15 G/E Before and After Thermal Aging at 450K (35OOF) in 0.014 MN/d (2 psi) Air and Tensile 

Testing at the Same Temperature 

8-29 Q 
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ORIGINAL PAGE Ts 

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Figure 8-18 taken from the same specimens show the identicdeffects, but even more deariy than 

the metallographic pictures. For example, the slight amount of surface oxidation after 10,000 

hours at 3944 (250°F) is easily distinguished in the SEM photograph in Figure 8-18, while 

scarcely visible in Figure 8-17. 

The changes from light to dark for pictures taken using the metallograph and dark to light using 

the SEM are explained by the amount of relief polishing around the individual fibem that 

increases with the degree of oxidation. When the epoxy resin oxidizes, it is more prone to cnunble, and differences in the amount of oxidation in the polished mounts can easily be detected 

using the SEM at higher magnifications. 

Figure 8-19 shows typically how this occurs. The as-received specimen has little and uniform 

relief for all plies as shown at 2500X using the SEM to examine a typical polished surface. The 

innerpliesof aspecimenagedfor 10,000 hoursat394K(25O0F)lookmuchthesameexcept for the 

filaments being oriented at 45 degrees to the surface. If one moves to the outer ply, the relief has 

increased as shown at the lower right of Figure 8-19. This effect becomes more pronounced for 

specimens aged at 450K (350°F) and 10,000 hours as shown in the lower left picture. The magnification was lowered to a lOOOX for clarity. 

The initial examinations were conducted after 10,000 hours of exposure. Additional studies were 

also made later in the program after time periods out to 50,000 hours. Figures 8-20 through 8-22 

show the results of these examinations for crossplied specimens of GE exposed to the three different aging conditions for 10,000,25,000, and 50,000 hours. The lighter cast, beginning at the 

specimen edges and spreading inward with exposure time is again caused by the amount of relief 

polishing around the individual fibers and is related to the degree of oxidization of the epoxy 

resin. The degree of oxidation, as shown in Figures 8-20 and 8-21, is quite similar for the first two 

conditions, ie., 394K (250°F) and one atmosphere air, and 450K (350°F) and reduced pressure. 

Slight-oxidation at the outer edges is visible after 10,000 hours. At 25,000 hours, considerable 

oxidation has occurred in the outer plies, and is beginning to extend into the second plies. The 

effect is somewhat more pronounced for the 450K (350"F), 0.014 MNlm2 (2 psi) exposure. After 

50,000 hours of aging, evidence of oxidation is visible in all the plies, with the effect again more 

pronounced for the 0.014 MNIm2 (2 psi) specimen. The metallographic results show slight effects 

after 10,000 hours, with major changes after 25,000 and 50,000 hours. These results agree well 

with both the residual tensile data and the condition of the specimens after tensile testing discussed earlier. Photomicrographs of specimens exposed to the third and most severe aging condition, 450K (350°F) and one atmosphere air, are presented in Figure 8-22. At 10,000 hours, the 

effects of oxidation are visible throughout the thickness of the laminate. After 25,000 and 50,000 

hours, extensive matrix embrittlement has occurred. These effects are particularly pronounced 

in the two middle plies, where resin loss is quite severe. 

Samples of BIE were also examined after thermal aging using both a metallograph and a SEM. 

No evidence of time dependent oxidation similar to that for the GIE specimens was observed. 

Overall, very little information was gained from examination of the aged BIE samples. This can 

be seen in Figure 8-23, which shows the 4.0 mil boron fibers and the epoxy matrix of specimens 

aged for 10,000,25,000, and 50,000 hours. No significant changes in the structure were found as 

the agingproceeded to 50,000 hours. As discussed later in Section 12, some evidence of increased 

matrix cracking was found for flight simulation specimens that had been exposed for the longer 

' time periods (36,000 hours). These specimens, however, were subjected to fatigue cycling and 

thermal cycling as opposed to only constant thermal exposure for the thermal aging specimens. 

8-32 
GDC-CRAD-85-005 

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GOC-CRAD-85-005 

8.2.1.2 Fmdogtaphr. In addition to the metallographic cross sectional examinations, a study 

was also made of the fracture surfaces of several of the specimens. Examples are presented in 

Figures 8-24 and 8-25. Typical fractographs of GE unaged and aged at 394K (250°F) and ambient pressure for 10,000 hours are shown in the upper portion of Figure 8-24. As might be pre 

dicted from the tensile results, the appearance of the fracture surfaces is very similar for the two 

specimens. The matrix is relatively intact and there is little fiber pullout. For specimens aged at 

450K (350°F) and ambient pressure for 5000 hours or longer, it was not possible to get a good 

SEM picture of the fracture. The very brittle resin had crumbled away from the filaments in the 

area of the fracture so that only a few filament ends would have appeared in the field of view. 

The pictures in the lower portion of Figure 8-24 show some of the filaments near the fracture 

locations in specimem aged at 450K (350°F). The filaments are quite different in appearance 

than those in the upper two photographs in the amount of resin adhering to the surface. The 

presence of this much resin on the graphite filaments indicates a severe degradation of the epoxy 

matrix, which accounts for the failure within the matrix rather than at the fiberlresin interface. 

Aninteresting example of the gradient of oxidation into the specimens is illustrated in the fractographs of Figure 8-25. These pictures were taken from areas near the surface of specimens aged 

at 394K (250°F) and reduced pressure. As shown in Figure 8-16 and 8-18 only a slight amount of 

oxidation (only at the outer surfaces) had occurred in these specimens in 10,000 hours. In Figure 

8-25 the depth to which the oxidation had progressed is readily visible. The resin near the outer 

surface has been embrittled by oxidation and has broken away fmm the filaments during tensile 

testing. Deeper into the specimen the fracture more closely resembles that of an unaged 

specimen. 

8.2.1.3 Glass Transition Temperature Studies. Further studies of the changes in the epoxy resin 

were made by measuring the glass transition temperature (Tg) of small samples cut from the G/E 

thermal aging specimens. These tests were performed using a DuPont 942 Thermomechanical 

Analyzer module in conjunction with a DuPont 990 Thermal Analyzer. The results are tabulated 

in Table 8-12. For aging at 394K (250°F) no change was observed for aging times out to 10,000 

hours, in agreement with the tensile results. For the reduced pressure, 0.014 NM/m2 (2 psi), exposures at 450K (350"F), however, signiscant increases in T, were found. These changes in the 

resin were not detected by either the mechanical property tests or microscopic examinations. No 

measurements could be made on the 5000 hour and 10,000 hour specimens aged at ambient pressure and 450K (350°F) or the 25,000 hour and 50,000 hour specimens aged at any of the three 

conditions because of the damage that occurred during tensile testing. 

The full significance of the changes in the T, values and their relationship to the changes in the 

resin system during aging are not clearly understood at this time. Additional T, measurements 

made on thermal aginglcompression specimens (see Section 8.3.1) and flight simulation specimens (see Section 12) were also of little value in understanding the resin system changes that 

occur during the thermal exposures of this program. 

8.2.2 GRAPHITE/POLYIMIDE 

8.2.2.1 HT-S/7lO G/PI. Residual tensile strength data for the HT-SI710 GlPI system are presented in Tables 8-13 through 8-16 for aging at ambient pressure and Table 8-17 for aging at 

reduced pressure. Aging was carried out at 505K (450°F) and 561K (550°F) for both unidirectional and [O" 245'1, crossplied specimens. Residual tensile strength testing was performed at 

8-39 
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8-40 ORIGINAL PAGT IS 

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DRIGINA4L P'1G-E TS 

OF POOR QUALITY 

10.000 HOURS 394K (25OoF), 101 kN/m2 (14.7 psi) 

(250x1 

10,000 HOURS 450K (35OoF), 13.8 kN/rn2 (2 psi) 

(400x1 

Figure 8-25 SEM Fractographs of G/E (Surface of Specimen Toward Viewer) 

8-41 

GDC-CRAD-85-005 

65321 7-90 
GOC-CRAO-85405 

Tablo 8-12 Glass Transition, T , Data for [Oo flSo], G/E 

0 

As-€&c&v& 463 374 

Aged 394 250 0.10 14.7 5,000 463 374 

Aged 394 250 0.10 14.7 10,000 464 375 

Aged 450 350 0.014 2 5,000 490 422 

Aged 450 350 0.014 2 10,000 505 449 

the same temperatures that were used for the thermal aging exposures. The tensile data have 

been plotted as a function of aging time and are presented in Figures 8-26 through 8-28. A summary of the effects of 10,000 hours and 50,000 hours of thermal aging on the tensile strength 

retention of G/PI is shown in Table 8-18. 

SOSK, 0.10 MN/rn2 and 0.014 MN/rn2 (GOOF, 14.7 psi and 2 psi) EXPOSURE. Exposures of up to 

25,000 hours at both one atmosphere and reduced pressure had essentially no effect on the 505K 

(450'F) tensile strength (see Figure 8-26 and 8-28). Compared with unexposed material there was 

no apparent difference in appearance for the one atmosphere specimens. The reduced pressure 

specimens did show some small spots, but these did not appear to be either serious or significant. 

After tensile testing the specimens again resembled unexposed specimens with no bare graphite 

fibers visible as a result of breaking up of an embrittled matrix. 

After 50,000 hours significant decreases in tensile strength were found in specimens aged at both 

ambient and reduced pressure. Specimen condition, however, was still quite good. Before tensile 

testing, a small amount of edge cracking and a slightly burned appearance were observed. After 

tensile testing a very slight loss of resin with a few bare graphite fibers could be seen. perhaps an 

indication of the start of matrix degradation. As shown in Figures 8-29 and 8-30, however, the 

effects were minor. 

There was one group of three specimens that gave unu&ally poor results. These were the unidirectional specimens aged for 500 hours at 595K (450°F). Residual strength values were obtained 

that were lower than all others of this group inciuding those aged for 50,000 hours. Examination 

of the broker specimens and the test data, as well as careful review of the specimen preparation 

and test procedures failed to reveal the cause of this anomally. Because the 500 hour data 

appeared to be in error, they were not included on the aging plot in Figure 8-26. 

561K, 0.10 MN/rn2 and 0.014 MN/m2 (55OoF, 14.7 psi and 2 psi) EXPOSURE. In contrast to the lower 

temperature exposures, those at 561K (550°F) produced significant decreases in tensile strength 

in times as early as 200 hours for ambient pressure (Figure 8-27). Although the strength 

decreases were quite large, very little change in the appearance of the specimens was observed 

after the first 10,000 hours of aging. The color and surface texture remained almost unchanged. 

Only slight edge cracks and some small spots on the surface with a charred appearance were 

8-42 
Tobk 8-13 Thormal Aging Data for [O0I6 HT-S/7lO G/PI in 0.10 MN/m2 (14.7 psi) Air 

505 450 505 450 Baseline avg 

500 

1,000 

5,000 

. 10,000 

25,000 

50,000 

1210 

827 

752 

758 

avg 779* 

1210 

990 

1370 

avg 1190 

1190 

1420 

1240 

- 

- 

176 

120 

120 

110 

113* 

175 

144 

198 

172 

- 

- 

173 

206 a 

180 

avg 1280 

1280 

1390 

1350 

avg 1340 

1350 

1210 

avg 1260 

848 

882 

965 

avg 898 

- 

. 1210 ’ 

- 

- 

186 

185 

20 1 

196 

194 

176 

196 

176 

183 

123 

128 

140 

130 

- 

- 

- 

*Not included in Figure 8-26 

8-43 
GOC-CRAD-85905 

Tablo 8-14 thormal Aging Data for [Oo faSol, Crorrply HFS/710 G/PI in 0.10 MN/m2 (14.7 psi) Air 

505 450 505 450 Baseline avg 

1,000 

470 

521 

443 

328 

5,000 

10,000 

25,000 

. 50,000 

avg 431 

435 

501 

576 

avg 504 

335 

525 

454 

avg 438 

- 

502 

463 

434 

avg 467 

. 410 , 

338 

34 1 

avg 363 

- 

- 

68.2 

75.5 

64.3 

47.5 

62.4 

63.1 

72.6 

83.5 

73.1 

48.6 

76.2 

65.9 

63.6 

72.8 

67.2 

63.0 

,67.7 

59.5 

49.0 

49.4 

32.6 

- 

- 

- 

- 

- 

8-44 
Tab10 8-16 Thormal Aging Data for [Oo 2aS0], Crorsply HFSnlO G/PI in 0.10 MN/m2 (14.7 psi) Air 

561 550 561 550 Baseline avg 434 

1,000 443 

474 

394 

avg 437 

- 

5,000 

10,000 

25,000 

50,000 

276 

276 

258 

avg 270 

192 

211 

274 

avg 226 

34 

48 

43 

avg 42 

- 

- 

- 

* 

* 

* 

63.0 

64.2 

57.1 

63.3 

40.0 

40.0 

37.4 

39.1 

27.8 

30.6 

39.7 

32.7 

5.0 

6.9 

6.2 

6.0 

68.7 

- 

- 

- 

- 

* 

* 

* 

*Specimens not suitable for tensile testing . 

8-46 
GDC-CXAD-85-005 

Tab10 8-15 fhormal Aging Data for [OO], HTW710 G/PI in 0.10 MN/m2 (14.7 psi) Air 

561 550 561 550 Baseline avg 

200 

500 

1,000 

5,000 

10.000 

25,000 

1320 191 

1290 187 

1280 185 

- 1350 - 196 

avg 1310 189 

1170 170 

1300 189 

1200 174 

avg 1220 178 

- - 

1100 159 

1170 169 

1190 

avg 1150 

- 

1170 

1070 

1010 

avg 1080 

1117 

876 

807 

avg 933 

254 

260 

277 

- 

- 

50.000 

avg 264 

* 

* 

* 

173 

167 

- 

169 

155 

147 

157 

- 

162 .. 

127 

117 

135 

- 

36.9 

37.7 

40.2 

38.3 

- 

* 

* 

* 

*Specimens not suitable for tensile testing 

8-45 
GOC-CRAD-85-005 

Tablo 8-17 Thormal Aging Data for HTS/710 G/PI Agod in 0.014 MN/m2 (2 psi) Air 

505 450 [0oi6 505 450 Baseline avg 1210 176 

5,000 1370 198 

1210 175 

215 

avg 1350 196 

- 1480 - 

10,000 1280 185 

1140 166 

1180 171 

avg 1200 174 

- - 

25,000 1210 176 

1330 193 

172 1190 

avg 1240 180 

- - 

50,000 1082 157 

882 128 

130 

avg 953 138 

- 896 - 

561 550 1001~ 561 550 Baseline avg 3 1320 191 

5,000 1300 

980 

1300 

avg 1190 

- 

10.000 1060 

1010 

970 

avg 1010 

- 

25,000 1010 

1060 

979 

avg 1020 

50,000 696 

752 

avg 763 

- 

841 

- 

188 

142 

188 

173 

- 

154 

146 

140 

147 

- 

147 

154 

142 

148 

- 

101 

122 

109 

111 

- 

8-47 
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1250 

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500 

250 

0 

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200 

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120 

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0.1 1 10 100 1000 10.000 1 00 ,000 

AGING TIME (HRI 

Ffguro 8-26 Tonsilo Strongth of HFSnlO G/PI at 50% (A50.F) Aftor Thermal Aging ot tho Same 

lomporaturo r 

found. After tensile testing, the specimens aged for 10,000 hours or less were almost identical to 

unexposed tensile specimens (Figures 8-31 and 8-32) indicating no change in the failure mode as a 

result of the exposure. 

A major change in specimen condition occurred after 25,000 hours in one atmosphere air. These 

specimens experienced considerable weight loss and, in the case of the crossplied material, severe 

edge delamination as shown in Figure 8-33. Both orientations were warped and curved, badly 

discolored with a burned appearance, and were covered with orange and white spots. Compared 

to the one atmosphere specimens the reduced pressure specimens looked very good. They did, 

however, show some loss of resin with some exposed fibers and were somewhat faded as compared to unexposed material. For the specimens aged in one atmosphere air, 25,000 hours of 

exposure severely degraded the tensile strength with losses in both the crossplied and unidirectional materid of 80% or more (Figure 8-27). The specimens aged at reduced pressure showed 

approximately the same tensile strength (Figure 8-28) as those aged for 10,000 hours (75 to 80% 

of that for unaged material). After tensile testing, the one atmosphere specimens were markedly 

different from those from material that had been aged for times up through 10,000 hours. The 

polyimide matrix had broken up during test and crumbled off revealing the graphite fibers. At 

failure the graphite tows opened up and resembled pieces of yam. Figure 8-31 shows the change 

in appearance in failed tensile specimens after 25,000 hours of exposure at 561K (550"F). The 

25,000 hour reduced pressure specimens were discolored after tensile testing but showed very 

little resin loss or exposed fibers as shown in Figure 8-32. 

8-48 
GOC-CRAD-85-005 

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AGING TIME (HR) 65321 7-350 

Figuro 8-27 Tonsil. Strongth of Ht-S/7lO G/PI at 561K (SOOF) Aftor fhormal Aging at tho Sam. 

Tomporotvro. 

After 50,000 hours of aging in one atmosphere air almost no resin was left in the specimen, only 

graphite fibers. The specimens were badly warped and delaminated and were no longer suitable 

for tensile testing. Figure 8-34 shows these specimens and the reduced pressure specimens that 

were in considerably better condition. The reduced pressure specimens were discolored and 

showed some loss of resin and exposed surface fibers but could be tested. The residual tensile 

strength, as shown in Figure 8-28, was slightly less than 60% of the unexposed value. The failed 

specimens showed evidence of the beginning of matrix embrittlement with loss of the matrix 

material and numerous exposed graphite fibers (see Figure 8-32). 

In summary, the HT-SI710 GlPI system survived 25,000 hours of thermal aging at 505K (450°F) 

with no effects. Some decrease in tensile strength was measured after 50,000 hours, but matrix 

degradation was not observed. Raising the aging temperature to 561K (550°F) reduced the time 

at which tensile strength decreases were observed to 1000 hours. In like manner to the 505K 

(450°F) exposures, this initial fall-off in tensile strength was not accompanied by matrix embrittlement. After 25,000 hours the material was partially delaminated and showed high weight loss. 

The tensile strength was greatly reduced and severe matrix embrittlement had occurred. Degradation after 50,000 hours was such that the specimens could not be tested. Again, aging effects 

were less severe for exposures conducted in a reduced pressure environment. 

m 

2 

c) 

The results of the metallographic examinations of the thermal aging specimens of HT-Sl710 

GlPI are shown in Figures 8-35 through 8-37. The changes in appearance as a result of the thermal exposures were far less obvious for the GlPI system than for the G/E system. For specimens 

aged at 505K (450"F), the only indication of matrix embrittlement was observed in specimens 

8-49 
GDC-CRAD-85405 

500 

250 

o* 

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P 

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I I111111 I 1111111 I I111111 I I111111 I111111 I IIIuIl, 

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c 

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160 

120 

a0 

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a 

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0.1 1 10 lG0 1000 10,000 100,000 

AGING TlME (HR) 65321 7-360 

Agure 8-28 Tenrllo Strength of Unidirectional HT-S/7lO G/PI at tho Indicated Temperature After Thermal 

Aging in 0.014 MN/m2 (2 psi) Air at the Sam0 Temperature 

Table 8-18 Summary of Tensile Strength Retention After Thermal Aging of HT-SfllO G/PI 

~ ~~~ 

Aging Tmnprrotun Aging Pmrurm Tost Tompemtun Rotention of Tensilo Stnngth (%) 

0ri.ntotfon K (OF) MNh2 (psi) K (OF) 10,000 hrs 50,000 hrs 

[00i6 505 450 0.10 14.7 505 450 100 74 

[Oo f45'1, 505 450 0.10 14.7 505 450 93 77 

[0oi6 505 450 0.014 2 505 450 99 78 

[00i6 561 550 0.10 14.7 561 e 550 71 0 

10' 245'1, 561 550 0.10 14.7 561 550 52 0 

[0'16 561 550 0.014 2 561 550 77 58 

8-50 
ORIGINAI; PAGE IS 

,OF POOR QUALIm .. 

GDC-CRAD-85-005 

25,000 HOURS 50,000 HOURS ._ 

UNEXPOSED 10,000 HOURS 

65321 7-361 

Figure 8-29 HFS/710 G/PI Before and After Thermal Aging at 505K (45O'F) in One Atmosphere Air and 

Tensile Testing at the Same Temperoture 

8-51 
GDC-CRAD-85905 

UNEXPOSED 10,000 HOURS 25,000 HOURS 50,000 HOURS 

65321 7-362 

Figure 8-30 HT-SnlO G/PI Before and After Thermal Aging at SOSK (GOOF) in 0.014 MN/m2 (2 psi) Air and 

Tensile Testing at the Same Temperature 

8-52 ORIGINAL PAGE IS 

OE POOR QUALITY 
ORTGTNAI; PAGE IS 

.OF POOR QUALITY 

GDC-CRAD-85-005 

6532 17-363 

Figure 8-31 HT-S/710 G/PI Before and After Thermal Aging at 561K (SSOOF) in One Atmosphere Air and 

Tensile Testing at the Same Temperature 

8-53 
GDC-CRAD-85-005 

25,000 HOURS 50,000 HOURS - UNEXPOSED 10.000 HOURS 

65321 7-364 

Figure 8-32 HT-SnlO G/PI Before and After Thermal Aging at 561K (550OF) in 0.014 MN/m2 (2 psi) Air and 

Tensile Testing at the Same Temperature 
GDC-CRAD-85-005 

653217-365 

Figure 8-33 Crossplied HT-SnlO G/PI After 25,000 Hours of Thermal Aging at 561K (55OOF) in One 

Atmosphere Air Showing Severe Edge Cracking and Delamination, 5X' 

exposed for 50,000 hours in one atmosphere air. This specimen, Figure 8-35, shows more relief 

polishing around the graphite fibers, more porosity, and more fiber-matrix separation than the 

specimens aged for shorter times or at reduced pressure. The effects, particularly the relief polishing, are more pronounced in the center than at the edges. Visual effects starting at the edges 

and moving inward, as seen in the G/E system, were not observed for the G/PI system. The effect 

of oxygen pressure can readily be seen, however, as evidenced by the two 50,000 hour specimens 

in Figure 8-35. The start of matrix degradation, as determined by the condition of the 50,000 

hour, one atmosphere specimens after tensile testing, appears to be confirmed by the metallographic results. 

Raising the aging temperature to 561K (550°F) greatly increased the degree of matrix degradation of the HT-S/710 G/PI system. Comparison of Figures 8-35 and 8-36 shows that in one atmosphere air, 10,000 hours at 561K (5550°F) has a slightly greater effect on the microstructure than 

50,000 hours at 505K (450°F). The amount of relief polishing around the graphite fibers is more 

pronounced, while the porosity and fiber-matrix separation are about the same. After 25,000 

hours at 561K (550°F) in air, the matrix is severely embrittled with considerable loss of resin 

from the center plies. The final photomicrograph in Figure 8-36, 50,000 hours at reduced pressure, shows little effect to the matrix. At higher magnification, however, (Figure 8-37) considerable relief polishing has occurred. This metallographic evidence of the onset of matrix 

embrittlement agrees well with the tensile data and the appearance of the failed tensile specimens. The remaining photomicrographs in Figure 8-37 clearly show the effect of matrix condition on the degree of relief polishing around the graphite fibers. 

8-55 
GDC-CRAD-85-005 

Figure 8-34 HT-SDlO G/PI After 50,000 Hours of Thermal Aging at 561K (55OOF) - left and Center 

Specimens, One Atmosphere Air: Right Specimens, Reduced Pressure 

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8-59 
8.2.2.2 Celion 6000/LARC-160 WPI. The thermal aging study of Celion 6000/LARC-160 G/PI was 

performed using only [0" f45"], crossplied specimens. Exposures were conducted at 450K 

(350°F) and 5058 (450°F) for time perid out to 10,000 hours. Both one atmosphere and 0.014 

MNlm2 (2 psi) environments were included. Post aging evaluation consisted of weight change 

measurements (one atmosphere specimens only), 450K (350°F) tensile tests, and metallographic 

and SEM examinations. Results of the weight chat3 measurements are presented in Tables 8- 

19 and 8-20. Residual tensile strength data are tabulated in Tables 8-21 and 8-22 for aging at 

ambient pressure and Tables 8-23 and Table 8-24 for aging at reduced pressure. The tensile data 

are also plotted as a function of aging time in Figures 8-38 and 8-39. A summary of the effects of 

10,000 hours of thermal aging on the tensile strength retention of this G/PI is shown in Table 

8-25. Six extra specimens were available after 10,000 hours of exposure for each of the aging 

conditions. These were tensile tested and are included in the data of Tables 8-21 through 8-24. 

The appearance of the specimens was relatively unchanged by the 10,000 hours of exposure. 

Some slight color changes and slight changes in the ring when the specimens were dropped were 

observed. After 5000 hours some edge cracking had occurred in the specimens aged at 505K 

(450°F). Similar edge cracking was found in the specimens aged at 450K (350°F) after 10,000 

hours. The degree of edge cracking in these 6 ply laminates was not severe. 

Duringcure of the adhesive for bonding the doublers seven of the nine 10,000 hour, 450K (350"F), 

one atmosphere specimens delaminated. None of the other specimens experience this problem. 

The delamination did not prevent tensile testing. 

After tensile testing, the specimens that had been aged at reduced pressure yere similar in 

appearance to the unexposed baseline tensile specimens. Specimens aged at one atmosphere 

were somewhat more broken up following tensile testing after 5000 hours of exposure. There was 

no visible difference between specimens aged at the two temperatures, however. After 10,000 

hours of agingin ambient air, the specimens (for both temperatures) showed an increased amount 

of breaking up and splintering. Some exposed fibers were visible where the matrix had crumbled 

away. The amount of bare fibers was not large. 

At 450K (350°F) weight changes, as shown in Table 8-19, were small with the maximum being a 

loss of 1.5% after 10,000 hours. At 505K (450"F), however, significant weight loss occurred after 

5000 hours (6%) and increased to about 9% after 10,000 hours. No weight change measurements 

were made on the reduced pressure specimens. 

As shown in Figure 8-38 thermal aging at 450K (350°F) had little effect for specimens exposed at 

reduced pressure with a decrease in tensile strength of only 8%. Ambient exposure at this temperature, on the other hand, reduced the tensile strength by 35% after 10,000 hours. By way of 

comparison, the crossplied G/E material lost 40% of the strength for the same exposure. 

Reduced pressure exposure at 505K (450'F) had about the same effect on residual tensile 

strength as did the 450K (350°F) exposure. Aging at 505K (450'F) in one atmosphere air was not 

detrimental to tensile strength retention as that at 450K (350°F). A decrease of approximately 

20% was measured after 10,000 hours. This is in contrast to thermal aging effects measured on 

several other composite systems where higher aging temperatures have resulted in increased 

loss in strength. Compared to the HT-S/710 GlPI system where crossplied specimens survived 

25,000 hours at 505K (450'F) and one atmosphere with almost no loss in tensile strength, the 

Celion 600O/LARC-160 system possessed considerably less thermal stability. 

8-60 
GDC-CRAD-85-005 

Tablo 8-19 thonnal Aging Data (Weight Changos) for Celion 600O/LARC-160 G/PI 

in 0.10 MNh2 (14.7 psi) Air 

450 350 500 

1.000 

f0.006 

+0.008 

+0.006 

+0.004 

+0.005 

+0.006 

5,000 

10,000 

0.000 

-0.001 

-0.001 

-0.056 

-0.045 

-0.063 

+0.2 

+0.2 

+0.2 

avg +0.2 

+0.1 

+0.1 

+0.2 

avg CO.1 

- 

- 

0.0, 

0.0 

0.0 

avg 0.0 

- 

-1.6 

-1.3 

- -1.7 

avg -1.5 

8-6 1 
' GOC-CRAD-85905 

Tabio 8-20 Thormal Aging Data (Woight Changor) for Ceiion 6000/LARC-l60 G/PI 

in 0.10 MN/d (14.7 psi) Air 

505 450 500 

1.000 

5,000 

10,000 

-0.02 

-0.02 

-0.02 

-0.03 

-0.03 

-0.03 

-0.21. 

-0.21 

-0.19 

-0.31 

-0.31 

-0.32 . 

-0.6 

-0.6 

- -0.5 

avg -0.6 

-0.7 

-0.8 

-0.8 

avg -0.8 

-5.9 

-5.8 

-5.5 

avg -5.7 

4.7 

-8.9 

-8.7 

avg -8.8 

- 

- 

- 
Tabk 8-21 fhormol Aging Data for [Oo kUo], Celion 6000/LARC-160 G/PI in 0.10 MN/d (14.7 psi) Air 

450 350 450 350 Baseline avg 

500 

1.000 

5,000 

10,000 

. 

696 

597 

596 

628 

avg 607 

518 

564 

578 

avg 553 

506 

550 

avg 528 

482 

423 

496 

432 

479 

433 

465 

419 

428 

- 

- 

* 

- 

avg 451 

*Doubler bond failure 

101 

86.6 

86.5 

91.1 

88.1 

- 

75.2 

81.8 

83.8 

80.3 

73.4 

79.8 

76.6 

69.9 

61.4 

72.0 

62.7 

69.5 

62.8 

67.5 

60.7 

62.1 

65.4 

* 

- 

- 

8-63 
Tablo 6-22 Thormal Aging Data for [Oo &Go], Cllion 600O/LARC-160 C/Pl in 0.10 MN/m2(14.7 psi) Air 

~~ ~~ ~ ~ ~~ 

505 450 450 350 Baseline avg 

500 

696 

609 

663 

685 

1,000 

5,000 

10,000 

avg 652 

667 

579 

607 

avg 618 

571 

562 

566 

avg 566 

514 

516 

510 

549 

58 1 

521' 

* 536 

585 

618 

avg 547 

- 

- 

- 

101 

88.3 

96.2 

99.4 

94.6 

96.8 

84.0 

88.1 

89.6 

82.8 

81.5 

82.1 

82.1 

74.6 

74.8 

74.0 

79.6 

84.2 

75.6 

77.7 

84.8 

89.7 

79.4 

- 

- 

- 

- 
GOC-CRAO-85405 

' 

Tablo 8-23 fhormal Aging Data for [Oo f45°], Celion 600O/lARC-160 G/PI in 0.014 MN/m2 (2 psi) Air 

450 350 450 350 Baseline avg 696 

5,000 565 

613 

614 

avg 597 

10,000 687 

685 

685 

623 

596 

629 

630 

632 

584 

avg 639 

- 

- 

101 

81.9 

88.9 

89.0 

86.6 

99.7 

99.4 

99.3 

90.4 

86.4 

91.3 

91.4 

91.7 

84.7 

92.7 

- 

- 

Tablo 8-24 Thormal Aging Data for [Oo 245'1, Colion 600O/LARC-160 G/PI in 0.014 MN/m2(2 psi) Air 

505 450 450 350 Baseline avg 696 

10.000 

5,000 562 

555 

570 

avg 563 

- 

621 

588 

593 

635 

624 

391 

658 

573 

607 

avg 610 

- 

101 

8 1. 5 

80.5 

82.6 

81.5 

90.1 

85.3 

86.0 

92.1 

90.5 

85.7 

95.4 

83.1 

88.0 

88.5 

- 

- 

8-65 
GDC-CRAD-85905 

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o 0,lO MN/m2 (14,7 PSI) 

- 20 

I IIllllI I I111111 I111111l I lllllll I1111111 I I111 

AGING TIME (HR) 

m 

a 

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65321 7-37 1 

Rgun 8-38 Tonsilo Strmgth of Collon 600O/LARC-160 G/Pl at 45OK (35O'F) After fhermol Aging a) &OK 

(350'F) and tho Indicotod Prosruro 

In summary, the effects of 10,000 hours of thermal aging in one atmosphere on Celion 60001 

LARC-160 G/PI were significant losses in tensile strength, beginning as early as 500 hours, and 

some matrix degradation. Tensile strength losses were higher for the 450K (350°F) exposures, 

35%, than for those at 505K (450"F'), about 20%, and were accompanied by greater matrix degradation. Weight loss during aging, however, was higher at the higher temperature. For the reduced 

pressure exposures no matrix degradation was observed, and strength losses were only of the 

order of 10% after 10.000 hours. 

Samples of Celion 6000LARC-160 G/PI in the unexposed condition and after 5000 and 10,000 

hours of exposure were mounted, polished, and examined using optical and scanning electron 

microscopes. At lower magnifications the degree of microcracking and relief polishing at the 

graphite fibers was found to be related to both the temperature and length of time of aging 

(increasing with both increasing temperature and time). At higher magnification, evidence of oxidation, similar in nature to that observed in the G/E system, was found. Figure 8-40 shows the 

results of the higher magnification examinations. When the polyimide resin matrix oxidized, it 

was more prone to crumble and resulted in an increased amount of relief polishing around the 

individual fibers. After 10,000 hours of aging at both temperatures, considerable oxidation had 

occurred in the outer plies but almost none in the center plies. This would be expected for an 

oxidation mechanism where attack begins at the outer surfaces and proceeds inward. Careful 

examination of Figure 8-40 also reveals an effect of temperature on oxidation where the degree of 

relief polishing around the graphite fibers is slightly greater at 505K (450°F) than at 450K 

(350°F). This oxidation effect was not observed for the HT-S/710 G/PI system. 

8-66 
GDC-CRAO-85405 

700 100 

A 

"E 600 \ 

z 80 ZE 

Io 

- 

500 

60 z 

a t; 400 

& 

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w 

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Iw 

I40 

a 200 z 

5 a 20 

100 

0 0 

0.1 1 10 100 I000 10,000 100,000 

1 

- E 

AGING TIME (HR) 

65321 7-372 

Figuro 8-39 Tonsilo Strongth of Celion 6000/LARC-160 G/PI at 450K (350OF) After Thormal Aging at 505K 

(45OOF) and tho Indicoted Prossuro 

8.2.3 BORON/ALUMINUM. In Phase I of the thermal aging test program exposures were conducted on B/E, GIE, HT-SI710 G/PI, and BIAl for time periods up to and including 10,000 hours. 

For the B/E, GIE, and G/PI systems additional specimens had been prepared and placed in the 

thermal aging furnaces at the beginning of Phase I for removal during Phase I1 at 25,000 and 

50,000 hours. No such provision was made for the B/Al system. However, at the beginning of 

Phase I1 some excess material purchased for fabrication of the 25,000 hour flight simulation 

specimens was available, and BIAI specimens were prepared and placed in the aging furnaces for 

Table 8-25 Summary of Tonsilo Strongth Retention After 10,000 Hours of Thormal Aging for 

[Oo f45'1, Crossplied Celion 6000/LARC-160 G/PI 

Aging Tomporoturo Aging Pmrura Tost Tomporoturo Rotontion of Tonsilo 

K (OF) MNh2 (Psi) K (OF) Strength (%) 

450 350 350 65 

450 350 0.014 2 450 350 92 

505 450 0.10 . 14.7 450 350 79 

505 450 0.014 2 450 350 88 

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exposures of 25,000 and 40,000 hours. The 40,000 hour figure was chosen so that all the material 

systems would complete the thd aging exposures at about the same time. As will be shown 

the two batches of BIAL, obtained 3 years apart, exhibited considerably Merent behavior during 

thermal exposure. For this reason the Phase I and Phase I1 results are discussed separately. 

8.2.3.1 Pharc. 1. The Phase I BlAl thermal aging study was conducted on unidirectional and 

[O" f45"], cn>ssplied laminates for time period3 of 5000 and 10,000 hours. Exposures were car 

riedout in one atmosphere air at temperatures of 450K (350"F), 561K (550°F). and 700K (800°F). 

FoUowhg aging, residual tensile and interfiber shear tests were performed at room temperature. 

Longit~~dinal tensile strengths of theunidirectional and crossplied materials were determined to 

evaluate fiber degradation while transverse tensile and longitudinal shear properties of the unidirectional material were obtained to study matrix effects. The data from the various exposures 

are presented in Tables 8-26 to 8-29. A summary of the effects of 10,000 hours of thermal aging 

on the strength retention of BlAl is given in Table 8-30. The tensile and shear strength data have 

also been plotted as a function of agins time. These curves are shown in Figures 8-41 to 8-43. 

At 561K (550°F) and 700K (800'F) extensive degradation was observed in the tensile strength of 

material tested in the 0 de- direction of both layups. These decreases were as high as 80% at 

700K (800°F) and 60% at 561K (550°F). Decreases were also observed (particularly in the unidirectional material) at 450K (350°F). although not as large. The physical appearance of the 

specimens was unchanged after 10,000 hours of aging at the two lowest temperatures. At the 

highest temperature, 700K (800'F), some surface discoloration was observed. Of more importance, however, was oxidation and loss of the exposed ends of the boron fibers in the transverse 

tensile and interfiber shear specimens. Figure 8-44 is a SEM photomicrograph of. the edge of a 

transverse tensile specimen that was aged for 5000 hours at 700K (800°F). The boron fibers are 

oxidized to a depth of appro&mately 0.00064m (0.025 in.). After 10,000 hours of aging, loss of the ' 

boron fibers extends to a depth of approximately 0.0015m (0.06 in.). No oxidation of the boron 

was found in specimens aged at the other temperatures. 

Because of the notches h the interfiber shear specimens were fabricated prior to aging, oxidation 

of the boron filaments had a significant effect on the residual shear strength. At 450K (350°F) 

and 561K (550°F). where no oxidation was found, no decrease in shear strength was observed. 

However, at 700K (800°F) decreases of 30 and 60% were measured after 5000 and 10.000 hours 

respectively. The machined edges at each end of the shearing regions were exposed to the air 

environment during aging and extensive oxidation of the boron occurred. This can be seen in the 

fracture surfaces of shear specimens shown in Figures 8-45 and 8-46. These figures show samples 

of B/Al after aging periods of 5000 'and 10,000 hours and, for comparison, one that had not been 

exposed prior to shear testing. The length of the shear region is approximately 0.0025m (0.10 in.) 

as shown in Figure 7-30. For the unexposed specimen, Figure 8-45, where no oxidation of the 

boron has occurred, the fracture surface is uniform along its length, The primary failure mode 

was matrix failure with some failure at the filament-matrix interface. In the specimen aged for 

5000 hours at 700K (800°F). Figure 8-45, loss of boron by oxidation is readily visible at either end 

of the sheared region. This attack extends for approximately 0.00064m (0.025 in.) from each end 

and has consumed about half of the fiber length, With the loss of the filaments, the area over 

which fracture occurred was considerably reduced and the specimen had lower strength. The 

residual shear stress was calculated using the original shearing area. It is this effect that 

accounts for the apparent 30% decrease in shear strength after 5000 hours at 700K (800°F). Had 

the notches .been machined in the specimens after aging, the residual shear strengths would 

probably have shown little change from the unexposed value. The failure mode in the center por8-69 
GOC-CRAD-85905 

Tablo 8-26 Thormal Aging Data for [0"]6 B/Al Agod in Ono Atmorphoro Air and Tostod at 

297K (75OF) - Longitudinal Tod Dfrodlon 

Tonsil. Strongtk 

MN/ml (kri) 

450 350 Baseline avg 

5,000 

561 550 

700 800 

10,000 

5,000 

10,000 

5,000 

1450 

1320 

1230 

1070 

avg 1210 

1010 

986 

869 

avg 955 

841 

855 

889 

avg 862 

855 

703 

786 

avg 781 

- 

- 

- 

- 

327 

63 1 

315 

10,000 

avg 424 

263 

320 

318 

avg 300 

- 

210 

191 

179 

155 

175 

146 

143 

126 

138 

122 

124 

129 

125 

124 

102 

114 

113 

47.4 

91.5 

45.7 

61.5 

38.1 

46.4 

46.1 

- 

- 

- 

- 

- 

43.5 

8-70 
Tablo 8-27 Thormal Aging Dot0 for [0"b B/Al Agd in Ono Atmorphoro Air and Tortd at 

297K (75°F) - Tronrvooo Tort Dirodion 

Aging Tompuotura Aging Tllm Tonsilo Strongth 

K (OF) (hr) MN/d (W 

450 

561 

700 

350 Baseline avg 

5.000 

550 

800 

10,000 

5,000 

10,000 

5.000 

79.3 

159 

143 

116 

avg 139 

104 

121 

avg 106 

102 

104 

avg 102 

108 

- 

93.8 

- 

98.6 

- 

74.5 

75.2 

avg 85.9 

104 

101 

94.5 

10,000 

avg 100 

80.0 

76.5 

avg 85.5 

100 

11.5 

23.1 

20.7 

16.8 

20.2 

15.1 

13.6 

17.5 

- 

~~ 

15.4 

14.3 

14.8 

15.1 

14.7 

15.7 

10.8 

10.9 

12.5 

15.1 

14.7 

13.7 

14.5 

11.6 

11.1 

14.5 

12.4 

- 

- 

7 

8-71 
Tablo 8-28 normal Aging Data for [O0I6 B/Al Agod in Ono Atmorphon Air and Tostod at 

297K (75OF) - Intorfibor Skoar in longitudinal Dirodion 

Tonsil. Strongth 

MN/d (kri) 

450 Baseline avg 

5,000 

561 

700 

550 

800 

10,000 

5,000 

10,000 

5,000 

- 10.000 

93.1 

82.0 

93.1 

avg 94.7 

- 109 

109 

100 

108 

avg 

- 

106 

85.5 

93.8 

91.7 

avg 90.3 

97.2 

91.7 

91.0 

avg 93.3 

42.1 

77.2 

79.3 

avg 66.2 

71.7 

18.6 

24.8 

avg 38.4 

- 

- 

- 

13.5 

11.9 

13.5 

15.8 

13.7 

15.8 

14.5 

15.7 

15.3 

12.4 

13.6 

13.3 

13.1 

14.1 

13.3 

13.2 

13.5 

6.1 

' 11.2 

11.5 

9.6 

10.4 

2.7 

3.6 

5.6 

- 

- 

- 

- 

- 

- 

8-72 
Tablo 8-29 Thormal Aging Data far [Oo f45'1, Crosspiid B/Al Agod in Ono Atmosphuo Air and 

Tdod at 297K (75OF) - Longitudinal Test Dirmion 

450 

561 

700 

350 

550 

800 

Baseline avg 

5,000 

10,000 

5,000 

10,000 

5,000 

10,000 

516 

430 

467 

535 

avg 477 

459 

444 

428 

- 

avg 444 

276 

251 

257 

avg 261 

184 

214 

210 

- 

avg 203 

185 

198 

avg 184 

142 

143 

154 

avg 146 

168 

- 

- 

74.8 

62.3 

67.7 

77.6 

69.2 

66.6 

64.4 

62.1 

- 

64.4 

40.4 

36.4 

37.3 

37.9 

26.7 

31.1 

30.4 

29.4 

24.4 

26.8 

28.7 

26.6 

20.6 

20.7 

22.3 

21.2 

- 

- 

8-73 
Tabk 8-30 Summary of Room Tompomtun Strongth Rotontion After 10,000 Houn of Thormal Aging 

in On. Atmosphoro Air for B/Al 

Laminah TUt Shengtfr Rltmtion (76) Ah? 10,OOO Houn 

Orimntution DI&oo P=+P*r)y UOK (350°F) 561K (SOOF) 7OOK (MW)OF) 

54 

100 

100 

39 

21 

100 

41 

28 

tion of the aged specimen is similar to that of the unexposed specimen in Figure 8-45 with the 

addition of a small amount of filament splitting. In the two end portions, failure has occurred in 

the narrow aluminum regions between the hollow areas previousiy occupied by the oxidized 

boron filaments. Figure 8-46 shows the two fracture surfaces (specimens are tested in double 

shear - Figure 7-30) of a specimen aged for 10,000 hours at 700K (800'F). For this specimen 

oxidation of the boron was more extensive on one side than on the other. As can be seen in Figure 

8-46, oxidation has consumed almost all the boron on the surface on the left. The opposite 

sheared surface, on the right in Figure 8-46, shows somewhat less filament oxidation. The failure 

w 

c 

w 

c 

I 

x 

a 

5 

1500 

1250 

1000 

750 

500 

250 

0 

ai 1 10 100 1000 10,000 

AGING TIME (hr) 

66321 7-99 

Figure 8-41 Tensilo Strength of B/AI at 297K (75°F) After Thormal Aging at Indicated Temperature 

8-74 
GDC-CRAD-85405 

I I I I1111 

w 

- 4 

so w 

c 

w 

c 

Z 25 

a 

Id 

0 

m 

v) 

-8 3: *m 

z 

0 

-4 

I 

I?. 

-4 i;: - 

I 1 I I1111 I I 1 I I Ill I I 1 11111 1 1 I 1 I Ill 

ai 1 10 100 1000 10.000 

AGING TIME (hr) 

653217-100 

Figun 8-42 Tranrvooo Tonsilo Strength of B/Al at 297K (75OF) Aftor Thormal Aging at Indicated 

Tompomture 

& - ;* 5 - %. 

3 

0 I I I11111 I 1 I I I Ill I I I I I Ill 1 I I I1 Ill I I I I I Ill 

0.1 1 10 100 1000 10,000 

AGING TIME (hr) 653217-101 

Figuro 8-43 Shoar Stength of Unidirectional B/Al at 297K (75OF) After Thermal Aging at Indicated 

Temperature 

8-75 
GDC-CRAD-85-005 

Figure 8-44 Edge View of B/AI After 5000 Hours of Thermal Aging at 700K (8OOOF) in One Atmosphere Air 

Showing Loss of Boron Through Oxidation 

mode of the specimens aged for 10,000 hours was again matrix dominated but showed an 

increase in the amount of filament splitting over that observed after 5000 hours.. 7 

Table 8-30 shows 100% retention of the transverse tensile strength for all the thermal aging 

exposures. These results, however, are somewhat misleading because of the low values obtained 

from the baseline tests. The baseline specimens were taken from a different panel than was used 

for the thermal aging tests and showed considerable scatter when tensile tested. The transverse 

tensile data from aged material, Table 8-27, indicate some decrease in strength has occurred for 

the 10,000-hour exposures, at 56K (550°F) and 700K (800°F). These decreases are thought to be 

the result of changes in the boron and boron aluminum interface rather than degradation of the 

aluminum matrix. Figures 8-47 through 8-49 show the fracture surfaces of transverse tensile 

specimens that were tested after the following thermal treatments: 

a. Figure 8-47, unexposed. 

b. Figure 8-48,5000 hours at 450K (350°F). 

c. Figure 8-49, 10,000 hours at 561K (550°F). 

In the unexposed specimen, Figure 8-47, the primary failure mode is matrix failure with some 

failure near, but not quite at, the filament-matrix interface. As thermal aging temperatures and 

times increase the failure mode changes from primarily matrix dominated to one in which, first, 

increased fracture at the filament-matrix interface occurs and, second, for more severe exposures, considerable filament splitting is observed. Figure 8-48 shows the increased interface fracture. A good example of filament splitting is illustrated by Figure 8-49. Two boron fibers have 

fractured lengthwise through their centers revealing the fine tungsten cores used in fabricating 

8-76 

ORIGINAL PASF '19 

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GDC-CRAD-85-005 

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Figure 8-47 Fracture Surface of Transverse Tensile Specimen of Unexposed B/Al Showing Predominantly 

Matrix Failure 

the filaments. An element scan taken with the SEM is also shown in Figure 8-49. With this technique the higher the atomic number of the constituent of the specimen the lighter the trace that is 

shown on the screen of the SEM. The two very light lines indicate the tungsten cores, the darkest 

areas are the low atomic number boron fibers, and the remaining grey area is the aluminum 

matrix. The change in the failure mode and the accompanying decrease in transverse tensile 

strength apparently occurs because the boron filaments become weakened as a result of interdiffusion between the boron and the 6061 aluminum alloy at the fibermatrix interface during thermal exposure. Electron microprobe studies (reference 34) have shown that degradation of boron 

fibers can occur at elevated temperatures as aresult of the diffusion of various elements from the 

aluminum alloy to the filaments. An excellent review of interfacial effects on the strength of BlAl 

after thermal exposure can be found in reference 35. 

The effect of boron fiber oxidation during aging at 700K (800'F) was shown to have contributed 

considerably to the decrease in shear strength. For the transverse tensile specimens aged at 

700K (800'F) the contribution was considerably less because of the lower percentage of test section affected, e.g., after 5000 hours at 700K (800'F) oxidation has consumed approximately 

0.00064m (0.025 in.) from each end of the test section or 50% of the 0.0025m (0.10 in.) long shear 

region but only 10% of the 0.0127m (0.50 in.) wide tensile specimen. Figure 8-50 shows the extent 

of oxidation inward from the surface of transverse tensile specimens aged for 5000 and 10,000 

hours. The appearance of split tubes is similar to that of Figure 8-45 and 8-46 for the shear specimens. The fracture surface away from the ends again reveals filament splitting and interfacial 

fracture in addition to the matrix failure. 

8-79 
GDC-CRAD-85-005 

MATRIX INTERFACE 

Figure 8-48 Fracture Surface of Transverse Tensile Specimen of B/Al After 5000 Hours of Thermal Aging at 

45OK (350°F) in One Atmosphere Air Showing Both Matrix and Interface Failure Modes 

Further evidence that the decrease in tensile and shear strength of BlAl during thermal aging 

was the result of degradation of the boron filaments rather than the aluminum matrix was found 

during examination of the failed [0" f45"], crossplied specimens aged for 5000 hours at 561K 

(550°F). The outer 0 degree ply (on both sides) showed a series of closely spaced, 0.0013m 

(0.05 in.) to 0.0025m (0.10 in.), cracks across the width of the specimens. An overall view of these 

specimens showing the surface cracks as well as the necking that occurred during tensile testing 

is presented in Figure 8-51. Magnified views of the surface cracks are shown in Figure 8-52. 

Examination of the cracks with a stereographic microscope revealed that the cracks extended 

only through the outer 0 degree plies and that the boron filaments in these outside plies were 

broken at each one of the surface cracks. This can be seen on the right in Figure 8-52. It is believed 

that the boron filaments became weakened as a result of diffusion to the filament-matrix interface during thermal exposure. Following the multiple failures along the brittle filaments in the 

outside 0 degree layers, the load was picked up by the 45" plies. With the 0 degree plies no longer 

active, the specimens necked down considerably prior to final failure. This necking, which is not 

typical of [0" +45"], material that has not been exposed, is clearly visible in Figure 8-51. In the 

specimens aged for 10,000 hours at 561K (550"F), surface cracks were not as prevalent as those 

observed in the 5000 hour specimens but were similar in appearance. A very small number of 

surface cracks were also found after tensile testing the specimens aged at 450K (350°F) while 

none were observed in the specimens aged at 700K (800°F). Apparently the fracturing of the 

0 degree outer boron filaments with the attendant surface cracks is highly dependent on the 

degree of filament embrittlement and is optimum after 5000 hours of exposure at 561K (550°F). 

8-80 ORlGINAL PAGE TS 

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Figure 8-51 B/Al After 5000 Hours of Thermal Aging at 561K (SOOF) in One Atmosphere Air and Tensile 

Testing at Room Temperature Showing Transverse Surface Cracks and Necking in.the Test 

Section 

8.2.3.2 Phase II. The Phase I1 BlAl thermal aging evaluation was similar to that conducted in 

Phase I in that both unidirectional and [0" f45"], crossplied laminates were thermally exposed 

in one atmosphere air. Also, post-exposure tensile and shear tests and metallographic and SEM 

examinations were performed. The 561K (550°F) aging temperature was retained, but the 700K 

(800°F) age was discontinued because of the severe degradation after 10,000 hours in Phase I, 

and the 450K (350OF) age was changed to 505K (450°F) to match the flight simulation exposure 

temperature. The exposure times were 25,000 hours and 40,000 hours. Data from the various 

exposures are presented in Tables 8-31 to 8-34. A summary of the effects of 25,000 and 40,000 

hours of thermal aging on the strength retention of BlAl is given in Table 8-35. 

Visual examination of the BlAl specimens after 25,000 hours and 40,000 hours of aging revealed 

no changes in the aluminum surface for either exposure temperatures. There was, however, oxidation and loss of the exposed ends of the boron fibers. At 505K (450°F) oxidation of the fine 

tungsten core in the center of the boron fibers was visible after 25,000 hours. After 40,000 hours 

some oxidation of the boron had occurred. The effect was considerably more pronounced for the 

specimens aged at 561K (550°F). Figure 8-53 shows the holes formerly occupied by the boron 

fibers in the edges of specimens aged at 561K (550°F) for 25,000 hours and 40,000 hours. 

After 25,000 hours the fine tungsten core has been oxidized, leaving a yellow-green residue. This 

can be seen in some of the holes in Figure 8-53. Oxidation of the boron has also occurred, although 

the amount is not very great for 25,000 hours. After 40,000 hours no traces of the tungsten oxida8-83 
GDC-CRAD-85-005 

8-84 
GOC-CRAD-85-005 

Table 8-31 Thermal Aging Data for [Ool6 B/Al Aged in One Atmosphere Air and Tostd 

at 297K (75°F) - Longitudinal Test Diroction 

Tonsil. S)nngth 

MN/m2 (kri) 

505 

561 

450 

550 

Baseline avg 

25,000 

40.000 

25.000 

40,000 

1500 

862 

896 

1430 

avg 1060 

1450 

1340 

1230 

avg 1340 

1360 

1510 

1180 

avg 1350 

917 

1230 

820 

avg 990 

- 

- 

- 

- 

218 

125 

130 

208 

154 

210 

195 

* 179 

195 

198 

219 

171 

196 

133 

179 

119 

144 

- 

- 

- 

- 

8-85 
GDC-CRAO-85905 

Table 832 Thermal Aging Data for [O"], B/Al Aged in One Atmosphere Air and Todd 

at 297K (75OF) - Transverse Test Direction 

Tonsilo Strmgth 

MN/rn2 (kri) 

505 450 Baseline avg 

25.000 

79.3 

109 

109 

103 

561 550 

40,000 

~ 

avg 110 

105 

103 

101 

25,000 

40,000 

avg 103 

91.7 

93.1 

91.0 

avg 91.7 

91.0 

86.2 

99.3 

- 

avg 92.2 

11.5 

15.8 

15.8 

15.0 

16.0 

15.2 

14.9 

14.7 

14.9 

- 

- 

13.3 

13.5 

13.2 

13.3 

- 

13.2 

12.5 

14.4 

13.4 

- 

. 

8-86 
GDC-CUD-85405 

Tablo 8-33 Thormal Aging Data for [Ool6 WAl Agod in Ono Atmorphoro Air and Testod 

at 297K (75'F) - Intorfibor Shoar in Longitudinal Direction 

Tonsilo Strongtth 

MN/m2 (kri) 

505 450 Baseline avg 

25,000 

40,000 

93.1 

88.2 

87.6 

89.6 

avg 88.2 

91.0 

86.9 

78.6 

- 

13.5 

12.8 

12.7 

13.0 

12.8 

13.2 

12.6 

11.4 

7 

561 550 25,000 

avg 85.5 

51 

50 

51 

12.4 

7.4 

7.2 

7.4 

40,000 

avg 50 

47 

52 

46 

7.3 

6.8 

7.6 

6.6 

avg 48 ~~ 

7.0 

8-87 
GOC-CRAO-85405 

Table 8-34 thormal Aging Doto for [Oo f45"], Crossplied B/Al Aged in Ono Atmosphero Air and Test4 

at 297K (75°F) - Longitudinal Test Dlrodion 

505 450 Baseline avg 625 90.6 

561 560 . * 

25,000 

40,000 

25,000 

40,000 

573 83.1 

585 84.9 

584 84.7 

avg 581 84.2 

584 84.7 

593 86.0 

574 83.3 

avg 584 84.7 

576 83.6 

574 83.2 

552 80.0 

avg 567 82.3 

525 76.1 

518 75.1 

558 80.9 

avg 534 * 77.4 

- - 

- - 

- - 

- - 

8-88 
GDC-CRAD-85-005 

Tablo 8-35 Summary of Room Temperature Strength Retention After fhormol Aging in 

One Atmorphoro Air for B/Al 

TOSt Aging T7mo Strength Rotontion (%) Laminato 

Oriontation Dlmtion PW=V (hr) 505K (GOOF) 561K (550'F) 

40,000 89 

[00i6 [goo1 Tensile 25,000 100 

40,000 100 

[o ' 16 DO1 Shear 25,000 95 

40,000 92 

[0" f45"], Pol Tensile 25,000 93 

40,000 93 

*See text for method of calculation of this value. 

90 

66 

100 

100 

54 

52 

91 

85 

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ORIGTNAL PAGE IS 

OF POOR QUALITY 
tion products are visible, but oxidation of the boron is quite pronounced. Similar oxidation of the 

boron was observed during Phase I (Figure 8-44) but only in specimens aged at 700K (800°F). 

The longitudinal residual tensile hits, Tables 8-31 and 8-34, indicate that the B/Al material 

used in the Phase I1 work behaved quite differently than did the Phase I material. For example, 

at 561K (550°F) the loss in strength for both the unidirectional and crossplied specimens was 

considerably less for the 40,000 hour specimens (Phase I1 material) than for the 5000 hour or 

10,000 hour specimens (Phase I material). The 505K (450°F) data also show the Phase I1 material 

to be superior when compared to the Phase I results for 450K (350°F) aging. 

There was an unexplained inconsistency in the 505K (450°F) Phase I1 data for the 25,000 hour 

exposure. Two of the unidirectional specimens gave residual strength values considerably lower 

than the third and also lower than the three 40,000 hour specimens. Metallographic examination 

failed to resolve the differences. It is presumed that the two low values are not representative of 

this material and, consequently, the percent retention value for 25,000 hours at 505K (450°F) in 

Table 8-35 has been calculated using only the one tensile value. 

A comparison of the quality assurance data obtained for the Phase I material and the Phase I1 

material shows that the greatest difference between the two batches of material is the thickness. 

The Phase I1 B/Al sheets were approximately 12% thicker. The photomicrographs obtained during QA testing show that one consequence of the thickness difference is that the outside layers of 

boron fibers in the Phase I material are much closer to the surface than those of the Phase I1 

material. Whether this could be responsible for the loss of strength found during Phase I (perhaps as a result of oxygen penetration, for example) could not bedetermined. Additional studies 

withmetallographic sectioning and optical and SEM examinations failed to reveal the reason for 

the significant difference in thermal aging behavior of the Phase I and Phase I1 batches of B/Al. 

While the longitudinal tensile results varied considerably for the two batches of material, the 

transverse tensile data and interfiber shear data appeared to fit reasonably well with the Phase I 

results. This would indicate that the problem lies with the boron fiber andlor fiber-matrix interface rather than with the aluminum alloy matrix. In like manner to the Phase I results, the 

decrease in strength at 561K (550°F) was caused by oxidation of the boron fibers beginning at the 

exposed ends. For the transverse specimens, the percentage of test section affected was small, 

and the decrease in strength with time was also small. For the interfiber shear specimens aged at 

561K (550"F), the oxidation effect occurred over the entire length of the shearing area and the 

decrease in shear strength was almost 50% for both 25,000 hours and 40,000 hours. 

Figure 8-54 shows the extent of oxidation on the fkacture surface of a transverse tensile specimen 

after 40,000 hours at 561K (550°F). The tungsten core has oxidized to amuch greater extent than 

the boron. The effect of the tungsten oxidation appears to be to embrittle the boron fiber as 

shown in Figure 8-54. The depth to which the tungsten core has oxidized appears to affect the 

nature of the failure. Splitting of the boron fibers occurred at the specimen edges where the tungsten has oxidized; otherwise failure was in the aluminum matrix and at the fiber-matrix interface. 

For specimens aged at 505K (450°F) for either time period the depth of oxidation was no more 

than 0.00025m (0.01 in.). At 561K (550"F), oxidation was considerably increased with depths of 

0.0013m (0.05 in.) after 25,000 hours and 0.0020m (0.08 in.) after 40,000 hours. Figure 8-56 shows 

the fracture surface of the same specimen shown in Figure 8-54 but at a location away from the 

edges where oxidation of the fibers had occurred. Of interest in this photomicrograph is the 

8-91 
GDC-CRAD-85-005 

SPECIME 

EDGE , 

* 

_- g, 

38X 6532 17-375 

Figure 84% Portion of Fracture Surface of Transverse Tensile Specimen of Phase II B/AI After 40,000 Hours 

of Thermal Aging at 561K (550°F) in One Atmosphere Air Showing Extent of Embrittlement of 

Boron Through Oxidation of the Tungsten Core 

absence of filament splitting. For the Phase I material, splitting was observed after only 10,000 

hours at 561K (550°F). This observation would appear to indicate that the boron fibers in the 

Phase I1 B/Al had a considerably improved thermal stability over that of the Phase I fibers. 

The results of the metallographic examination of the interfiber shear specimens showed a considerable difference in the fracture appearance of specimens aged at the two temperatures. At the 

lower temperature almost no oxidation of the fibers was found, and failure was in shear, primarily 

in the aluminum matrix. For the 40,000 hour specimens, some interface failure was observed as 

shown in Figure 8-56, but, as with the transverse specimens, no filament splitting was present. 

At 561K (550"F), oxidation of the filaments (primarily the tungsten cores) had occurred over the 

entire shearing length of the specimens. As shown in Figure 8-56, the boron filaments were broken into many small fragments with the oxidized remnants of the tungsten cores readily visible. 

Failure did not take place in one shear plane as was seen for the 505K (450°F) specimens but 

occurred in more or less a tearing manner across several planes in a direction 45 degrees to the 

direction of loading. The weakened boron filaments appear to have been responsible for the lowered shear strength values of these specimens. 

Transverse surface cracks of the type found on Phase I [O" +45"], crossplied specimens after 

tensile testing (see Figure 8-52) were observed on several of the Phase I1 specimens. They were 

few in number, however, as shown on a typical specimen in Figure 8-57. This may be further evidence of the improved thermal stability of the boron filaments in the Phase I1 B/Al material. 
ORIG'N:%L PAGE IS 

Q? POOR QUALITX 

GDC-CRAD-85-005 

MATRIX INTER FACE 

653217-376 

Figure 8-55 Fracture Surface of Transverse Tensile Specimen of Phase II B/Al After 40,000 Hours of Thermal 

' Aging at 561K (SOOF) in One Atmosphere Air Showing Both Matrix and Interface Failure 

Modes (Note Absence of Filament Splitting) 

In summary of these data, the B/6061 borodaluminum used in the Phase I1 portion of the program was much superior to that used during Phase I. The Phase I data indicated a maximum use 

temperature of only 450K (350°F) for exposures longer than 10,000 hours. The Phase I1 data, on 

the other hand, indicate that a maximum use temperature of 505K (450°F) for 40,000 hours of 

exposure is possible. Some precautions may be required, however, to protect the material edges 

to prevent oxidation of the boron fibers. 

8.3 THERMAL AGING/COMPRESSION - TEST RESULTS AND DISCUSSION 

8.3.1 GRAPHITE/EPOXY. The results of the weight change measurements, the compressive 

strength and modulus tests, and the shear strength tests are given in Tables 8-36 and 8-37. The 

data have also been plotted as a function of aging time for the two laminate thicknesses and the 

two aging temperatures in Figures 8-58 through 8-60. The effect of aging on the fully reversed 

tension-compression fatigue strength of the 24 ply laminate was also determined. These data are 

listed in Table 8-38, and S-N curves are plotted in Figure 8-61. 

All specimens were examined visually with a stereo microscope after aging (1000,5000, or 10,000 

hours). For the GIE specimens aged at 373K (212"F), no changes in appearance were observed for 

any of the time periods. The only difference noted was a slight change in the ring of the 24 ply 

specimens when dropped after the 10,000 hour exposure. At 394K (250"F), slight edge cracking 

was seen after 5000 hours in the 24 ply specimens and after 10,000 hours also in the 6 ply speci8-93 
GDC-CRAD-85-005 

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POOR QUALITY 

6X 6532 17-370 

Figure 8-57 Phase It B/Al After 40,000 Hours of Thermal Aging at 505K (45OOF) In One Atmosphere.Air and 

. Tensile Testing at Room Temperature Showing 'Slight Transverse Surface Cracking Near 

Fracture 

mens. Both layups showed a slight loss .in ring after 10,000 hours. In general, all the specimens 

were in excellent condition except for the slight edge cracking. Examples of these cracks are 

shown in Figure 8-62 for 6 ply and 24 ply specimens aged 10,000 hours at 394K (250°F). 

Weight change for the GIE system was small during thermal aging as shown in Figure 8-58. The 

respective weight losses were as expected for the four combinations of aging temperature and 

specimen thickness. The greatest effect, which was only 1%, was for the 6 ply material at 394K 

(2 50 OF). 

Figure 8-59 shows that the 10,000 hours of aging at both temperatures did not lower the compressive strength. In fact, values were increased after aging, particularly for the 24 ply specimens. Similar results were obtained for the compressive modulus measurements. These 

apparent increases in compressive properties may be the result of faulty baseline property tests. 

However, this is not believed to be the explanation, as the specimen preparation, test procedures, 

and appearance of the failed specimens were the same for all the tests. The superior properties 

after exposure may be the result of a postcure effect at the aging temperatures. 

For G/E, the only detrimental effects of thermal aging were the shear strength losses of the 6 ply 

material. As shown in Figure 8-60, these decreases were found as early as 1000 hours and became 

quite large by 10,000 hours. Since the effect was observed only in the 6 ply material, the possibility exists that the method of specimen preparation (bonding of three 6 ply lamhates to get an 18 

ply specimen) may have influenced the data. However, based on results from both the Phase I 

8-95 
GDC-CRAD-85-005 

Tablo 8-36 Thormal Aging Data for 6 Ply [Oo kaSO], AW3501 G/E Aged in 

0.10 MN/m2 (14.7 psi) Air and Tested at Room Tomporaturo 

Aging WM 

Tompomiurr Aging Timr Changr bmpmsivr Strongtll Compmsivr Modulus Show Strongtk 

K (OF) (hd (W MN/mZ (kd) GN/m2 (fw MN/m2 (ksi) 

373 212 Baselineavg 542 78.6 53 7.7 72 10.4 

1.000 -0.1 616 89.3 52 7.5 67 9.7 

avg -0.1 620 89.9 53 7.7 64 9.3 

4.1 623 90.3 52 7.6 55 8.0 

- 4.1 - 620 - 90.0 - 56 - 8.1 - 70 - 10.2 

5,000 0.0 510 74.0 57 8.2 63 9.9 

0.0 490 71.1 56 8.1 65 9.4 

8.8 

avg 0.0 504 73.1 57 8.2 63 9.1 

- -0.1 - 512 - 74.2 - 57 - 8.3 - 61 - 

10,000 -0.1 546 79.2 53 7.7 23 3.4 

-0.1 580 84.1 57 8.3 23 3.3 

avg -0.1 571 82.8 55 8.0 23 3.3 

e -0.1 586 85.0 55 8.0 23 3.3 P - - - - - - - 

394 250 1.000 -0.3 

-0.3 

- -0.3 

avg -0.3 

5.000 -0.5 

-0.5 

- -0.5 

avg -0.5 

565 

643 

572 

594 

558 

597 

561 

572 

- 

- 

82.0 

93.3 

83.0 

86.1 

81.0 

86.6 

81.3. 

83.0 

- 

- 

52 

69 

56 

59 

51 

53 

52 

52 

- 

- 

7.5 

10.0 

8.1 

8.5 

7.4 

7.7 

7.6 

7.6 

- 

- 

63 

60 

64 

63 

52 

48 

50 

30 

- 

- 

9.2 

8.7 

9.3 

9.1 

7.6 

7.0 

7.2 

7.3 

- 

- 

10,000 -1.0 562 81.5 54 7.8 41 6.0 

-1.1 549 79.6 54 7.9 45 6.5 

-1.1 6.2 

avg -1.1 561 81.4 34 7.9 43 6.2 

- - 573 - 83.1 ' - 55 - 8.0 - 43 - 

8-96 
GOC-CRAO-85-005 

Table 8-37 Thermal Aging Data for 24 Ply [Oo f45"], A-S/3501 G/E Aged in 

0.10 MN/m2 (14.7 psi) Air and Tested at Room Temperature 

__ 

Wnq WgM 

T.Wlpmhm Aging Timo Change bmpmsive Strongtth Comprmsriva Modulus Shoar Str.ngth 

373 212 Baseiineavg 439 63.6 41 6.0 56 8.1 

1.000 0.0 623 90.3 58 8.4 76 11.0 

0.0 

0.0 

avg 0.0 

5.000 0.0 

0.0 

0.0 

avg 0.0 

10,000 0.0 

0.0 

0.0 

avg 0.0 

394 250 1.000 -0.1 

-0.1 

-0.1 

avg -0.1 

5.000 -0.2 

-0.2 

- 

- 

- 

- 

-0.2 , 

661 

676 

653 

644 

636 

583 

620 

684 

649 

662 

665 

578 

583 

618 

- 

- 

- 

- 

593 

607 

598 

631 

avg -0.2 

10.000 -0.5 

-0.5 

612 

606 

603 

95.8 

98.1 

94.7 

93.4 

92.2 

84.5 

90.0 

99.1 

94.1 

96.0 

- 

- 

96.4 

83.8 

84.5 

89.6 

86.0 

88.1 

86.8 

91.5 

- 

88.8 

87.9 

87.5 

59 8.3 

57 8.3 

58 8.4 

63 9.1 

62 9.0 

58 8.4 

61 8.8 

61 8.9 

65 9.4 

59 8.5 

61 8.9 

53 7.7 

48 7.0 

51 7.4 

51 7.4 

52 7.5 

53 - 7.7 

56 8.0 

53 7.7 

54 7.8 

56 8.1 

- - 

- - 

- - 

- - 

- - 

76 

79 

77 

68 

68 

53 

63 

70 

73 

77 

74 

83 

75 

70 

76 

61 

54 

54 

57 

60 

59 

- 

- 

- 

- 

- 

11.0 

11.5 

11.2 

9.8 

9.9 

8.0 

9.2 

10.2 

10.6 

11.2 

10.7 

12.1 

10.9 

10.1 

- 

- 

- 

11.0 

8.8 

7.9 

7.9 

8.2 

8.7 

8.3 

7.2 

avg -0.5 595 86.3 54 7.9 56 8.1 

- -0.5 - 576 - 83.6 - 54 - 7.8 - 50 - 

8-97 
GOC-CRAO-85405 

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Oi 

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-2 s 

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c-4 

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-6 

-8 

I 

0 6PLY 373K (212°F) 

[I 24 PLY 373K (212°F) 

h 6 PLY 394K (250°F) 

0 24 PLY 394K (250°F) 

I I111111 I 1111111 I I II1111 I I I I lllL 

0.1 1 10 100 1000 10,000 

AGING TIME (HR) 65321 7-379 

Flguro 8-58 Woight Chango of G/E Aftor Thermal Aging ai tho Indicatod Temporature 

and Phase I1 flight simulation tests of this specimen configuration the methodis believed to be a 

satisfactory one. Careful visual examination of the failed specimens, both baseline and aged, 

with a stereo microscope failed to reveal any information on the degradation process involved in 

the loss in shear strength. 

In agreement with the residual compressive and shear strength values for 24 ply material, the 

residual fatigue strength showed no decrease as a result of the 10,000 hour thermal age at either 

373K (212°F) or 394K (250°F). The data, plotted in Figure 8-61, show that the scatter bands for 

each temperature overlap those of the baseline data. Individual cwes for the three data sets 

were not drawn because of their close proximity to each other. The net effect is that the thermal 

aging has not changed the fatigue properties of the 24 ply GIE material. 

Metallographic aamples were prepared from 6 ply specimens before aging and after 5000 hours 

and 10,000 hours of exposure at both 373K (212°F) and 394K (250°F). The results were the same 

as found in the Phase I thermal aginghension study of GIE. Slight oxidation effects could be seen 

to be just starting in the outside plies after 10,000 hours of aging. The specimens resembled 

those in Figures 8-17 and 8-18. No explanations were uncovered to explain the high losses in 

. shear strength of the 6 ply material after aging. It was thought that the slight edge cracking 

might have contributed, but the polished samples of material aged for 10,000 hours at each temperature showed no evidence of internal matrix cracking. Figure 8-63 shows samples after 10,000 

hours of aging at each of the temperatures. The condition of both specimens with regard to 

matrix cracking is the same as unexposed material (see Figure 8-17). 

8-98 
GOC-CRAO-85-005 

x 

2 

I2 

w 

t 

w 

X 

Icl 

2 

W 

- 

= 

a 

a 

a 

100 

80 

60 

40 

20 

0.1 1 10 100 1000 10,000 0 

AGING TIME (HR) 

65321 7-380 

Figure 8-59 Retention of Compressive Strength of G/E at 297K (75°F) Aftor Thormal Aging at the Indicated 

Temperature 

Glass transition temperatures (T,) were also measured on samples cut from the same specimens 

that provided the metallographic samples. The results are given in Table 8-39. The T, values 

appear to offer no assistance to understanding the changes that occur in the epoxy resin during 

long term thermal exposure in air. 

. 

In summary, 10,000 hours of thermal aging in one atmosphere air at 373K (212°F) and 394K 

(250°F) produced no significant reduction in compressive strength or compressive fatigue 

strengthof GIE. A slight weight loss, about 1%, was found after 10,000 hours at 394K (250°F) for 

6 ply materid The major effect of the thermal exposure was a considerable deciease in shear 

strength of 6 ply material. These reductions were observed as early as 1000 hours and increased 

to 40910 or more by 10,000 hours. A mechanism to account for the poor shear strength retention of 

GIE during thermal aging was not found. 

By way of comparison with the thermal agingkompression data, the thermal agingkension 

results showed no decrease in tensile strength after 10,000 hours of exposure in air at 394K 

( 250 " F). 

8.3.2 CELJON 6000/LARC=160 GRAPHITE/POLYIMIDE. The results of the weight change measurements, the compressive strength and modulus tests, and the shear strength tests are given in 

Tables 8-40 through 8-42. These data have also been plotted as a function of aging time for the 

two laminate thicknesses and the two aging temperatures in Figures 8-64 through 8-66. Table 

8-42 and Figure 8-64 give the data for the PPQ coated material. The results of aging on the 

8-99 
GDC-CRAO-85905 

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c 

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0 

. 0 6PLY 373K 

0.1 1 10 100 1000 10,000 

AGING TIME (HR) 

65321 7-381 

Rgure 8-60 Retention of Short barn Shear Strength of G/E at 297K (75'F) After thermal Aging at tho 

Indicated Ternperaturo 

R = -1 fatigue strength of the 24 ply laminate are presented in Table 8-43 and Figure 8-67. 

Table 844 compares the results of the uncoated and PPQ coated samples of 6 ply material. 

Vial examinations were performed on the specimens at intervals of 1000 hours, 5000 hours, 

and 10,000 hours during aging. For the exposures at 450K (350°F) the major changes noted were 

for the 24 ply samples. After 5000 hours, slight edge cracking had occurred and by 10,000 hours 

the cracking was quite extensive. These cracks are readably visible in Figure 8-68. For the 6 ply 

specimens aged at 450K (350°F) very slight edge cracking appeared to be starting in the specimens that had been exposed for 10,000 hours. Overall, the appearance of the GlPI material after 

10,000 hours was excellent with no surface degradation, color changes, or decreases in the ring of 

the specimens when dropped on a hard surface. 

Aging at 505K (450°F) also hadlittle visual effect on the specimens except for edge cracking. For 

the 24 ply samples extensive cracking was observed after 1000 hours. The number of cracks was 

about the same after 5000 hours, but they were considerably wider. The 10,000 hour samples 

were similar in appearance to the 5000 hours samples. Figure 8-69 shows the degree of edge 

crackingin the 24 ply lamhates after 1000 hours and 10,000 hours. For the 6 ply material no edge 

cracks were observed until after 10,000 hours of esposure. Figure 8-70 shows an esample of edge 

cracks and what appears to be delamination. The degree of cracking is less than that found in the 

24 ply material. 

After the observation of these edge cracks the question arose as to the depth to which they 

extended. Some spare specimens of the 24 ply material had been included in the 10,000 hour 

8- 100 
GOC-CRAD-85-005 

Tablo 838 Axial Fdiguo Data for [Oo f45"], G/E Aftor 10,OOO Hours of Thormal Aging, 

297K (75OF), R =: -1, Unnotchd 

373 212 B- 1 0-3-1 0 

-11 

-12 

-13 

-14 

-15 

-16 

-17 

-18 

394 250 B-10-3-28 

-29 . 

-30 

-32 

-3 1 

-33 

-34 

-35 

-36 

241 

276 

224 

258 

241 

224 

276 

207 

190 

224 

207 

224 

241 

224 

241 

258 

258 

280 

35 

40 

32.5 

37.5 

35 

32.5 

40 

30 

27.5 

32.5 

30 

32.5 

35 

32.5 

35 

37.4 

37.5 

40.6 

7,000 

9,000 

894,000 

14,000 

11.000 

48,000 

6,000 

3,460,000 

11,450,000 Did not fail 

25,000 Fixture misaligned 

12,210,000 

4,341,000 

705,000 

1,190,000 

1,415,000 

13,000 

14,000 

10,000 

exposures and were available fof examination. Two samples, one with 10,000 hours at 450K 

(350°F) and one with 10,000 hours at 505K (450"F), were sectioned in the plane of the laminate 

near the mid-plane. Using a stereo microscope one could follow the extent of the edge cracks into 

the specimens. As shown in Figure 8-71 these edge cracks extended all the way through the plane 

of the 0.038m by 0.038m (1.5 in. by 1.5 in.) samples. The pieces shown in Figure 8-71 are onequarter of the original samples so that the lower right comer of each is approximately the center 

of the specimen. 

The PPQ coated 6 ply samples were also examined after 1000 hours, 5000 hours, and 10,000 

hours. For the 450K (350°F) exposures no changes in appearance were detected. Two of these 

specimens after 10,000 hours are shown in Figure 8-72. At 505K (450°F) peeling of the coating 

was visible after 1000 hours and increased slightly after 5000 hours. For the three specimens 

examined one showed extensive peeling after 1000 hours while the other two showed only slight 

peeling after 5000 hours. There was very little change in appearance from 5000 hours to 10,000 

hours. Figure 8-73 shows two of these specimens after 10,000 hours of exposure, one with extensive peeling of the coating and one with slight peeling. Because the peeling and cracking occurred 

early in the aging exposure and did not change much with time, it would appear to be an application problem with the coating not adhering properly. 

Using chloroform as a solvent, the PPQ coating was removed after 10,000 hours of aging prior to 

preparation of the compression and shear specimens. Examination of the samples after coating 

removal showed them to be in excellent condition and, unlike the uncoated 6 ply material, to be 

free of edge cracking. Figure 8-74 shows a typical edge from one of these samples after removal of 

the PPQ coating. 

8-101 
GOC-CRAD-85905 

I I OUNEXPOSED 

0 AGED AT 373K (212°F) 

AAGED AT 394K (250°F) 

I 

r65321 7-382 

Figure 8-61 Axial Fotigue Properties of [Oo *45"] G/E Before and After 10,000 Hours of Thermal Aging, rl. . 297K (75OK). R = -1, Unnotchd 

Weight change for the Celion 6000/LARC-160 G/PI during thermal aging is plotted in Figure 

8-64. Significant weight losses occurred at both aging temperatures during the 10,000 hours of 

exposure. As expected, greater losses were found at the higher aging temperature, and the 6 ply 

laminates lost more weight, percentagewise, than the 24 ply laminates. The effect of the PPQ 

coating on the 6 ply material was a slightly lower weight loss at all of the timetemperature conditions. Comparison of the weight loss of the 6 ply compressive specimens with the 6 ply tensile 

specimens (Table 8-20) after aging at 505K (450°F) gives very good agreement. At 450K (350"F), 

however, the tensile specimens (Table 8-19) showed weight losses somewhat lower than those of 

the compressive specimens. The aging conditions and weighing procedures were identical for the 

two groups of specimens. The weight change differences would appear to be related to the two 

different panels from which the specimens were obtained. 

In contrast to the weight change data, but in agreement with the thermal aging tensile data, the 

compressive strength in most cases showed a greater decrease with aging time at 450K (350°F) 

than at 505K (450°F). As shown in Figure 8-65 this is particularly evident for the 24 ply material 

and the 6 ply specimens after 5000 hours. For two of the data sets (6 ply at 450K and 24 ply at 

505K) the results showed greater decreases in compressive strength after 5000 hours than after 

10,000 hours. Similar inconsistencies were found in the compressive modulus values for these 

two data sets. No explanation was found for these results, but data scatter cannot be ruled out. 

For 6 ply Celion 6000/LARC-160 G/PI the loss of compressive strength and tensile strength (Section8.2.2.2) after 10,000 hours of thermal agingin air at both450K (350°F) and 505K (450°F) was 

significant with decreases in excess of 20%. 

8- 102 
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GDC-CRAD-85-005 

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8- 104 
GOC-CRAO-85-005 

Tablo 8-39 Glass Transition Temporaturo, T , Data for [Oo -CdS0], G/E 

9 ~ 

Aging TImo Tg Aging Tempomtun 

condition K (OF) (hr) K (OF) 

As-Received 441 333 

Aged 

Aged 

Aged 

373 212 

373 212 

373 212 

1.000 434 322 

5,000 419 295 

10,000 440 332 

Aged 394 250 1,000 428 311 

Aged 

Aged 

394 250 5,000 449 348 

394 250 10,000 420 296 

For both the G/E and the G/PI systems the most damaging effects of the thermal aging were 

found in the shear strength measurements. The 6 ply G/PI material aged for 10,000 hours at 

450K (350°F) retained only 35% of its room temperature shear strength. Aging at 505K (450°F) 

was only slightly better with a retention of about 50%. Again, the unexpected result of greater 

property loss in the 6 ply material when aged at the lower temperature was observed. For the 24 

ply material, shear strength losses after 10,000 hours at 505K (450°F) were also quite large. 

However, the 24 ply material showed good resistance to aging at the lower temperature with over 

90% retention of shear strength after 10,000 hours. The effects of thermal aging on shear 

strength are plotted in Figure 8-66. 

The thermal aging tension-compression fatigue data for the 24 ply Celion 6000/LARC-160 G/PI 

are plotted as S-N curves in Figure 8-67. Unlike the GIE material, which showed no effects of 

aging on fatigue strength, the G/PI fatigue strength was lowered significantly after 10,000 hours 

of aging at both temperatures. Room temperature fatigue strength was about 15% lower after 

aging at 450K (350°F) and 30% lower for the 505K (450°F) aging exposure. 

As discussed earlier the PPQ coating was applied to a batch of specimens in an attempt to 

improve the long term themooxidative stability during aging. Table 8-44 summarizes the 

results of this study. The coated samples showed slightly better results in almost every case. 

These included less edge cracking, lower weight loss, and better shear strength retention. For 

compressive strength and modulus, the coated samples gave better results than uncoated when 

aged at 505K (450°F) but poorer results for the 450K (350°F) age. These preliminary results indicate that the PPQ coating has potential for providing increased themooxidative stability to 

GiPI laminates. 

Metallographic examinations were made on 6 ply samples aged for 5000 and 10,000 hours at each 

temperature. The same matrix oxidation effects described earlier for the thermal aging/tension 

samples of Celion 60001LARC-160 (Figure 8-40) were also found for the thermal 

agingkompression samples. Extensive internal matrix cracking was also observed. Figure 8-75 

shows the extent of cracking after 5000 hours at the lower aging temperature and after 5000 

8- 105 
GOC-CRAD-85-005 

Tabk 8-40 Thormal Aging Data for 6 Ply [Oo -C45'], Colion 6000/LARC-160 G/Pl Agad in 0.10 MN/d (14.7 psi) 

Air and Tostod at Room Tempomturo 

A9m weigh( 

Tomporoturo Aging nmo Chongo bmpmsivo Strongth bmpnuivo Modulus Show Strongth 

K (Ot) (hr) (XI MNh2 (k4 CN/d (AIW) MN/m2 (kai) 

450 350 Baselineavg 584 84.7 52 7.5 51 7.4 

1,000 +0.1 592 85.8 54 7.8 50 7.3 

+0.1 634 92.0 52 7.6 44 6.4 

+0.1 88.3 52 6.6 

avg +0.1 612 88.7 52 7.6 47 6.8 

5,000 -0.4 397 57.6 50 7.2 47 6.8 

-0.4 390 56.5 50 7.2 44 6.4 

-0.4 58.6 50 6.1 

avg -0.4 397 57.6 50 7.2 44 6.4 

- - 609 - - - 7.5 - 46 - 

- - 404 - - - 7.2 - 42 - 

10.000 -2.3 

-2.3 

- -2.3 

avg -2.3 

505 450 1,000 -1.2 

-1.0 

-0.9 

avg -1.0 

- 

5,000 -5.6 

-5.5 

-5.6 

avg -6.6 

- 

10.000 -8.2 

-8.1 

- -8.2 

avg -8.2 

461 

408 

398 

422 

588 

587 

629 

602 

477 

407 

472 

452 

423 

436 

40 1 

421 

- 

- 

- 

- 

66.8 

59.2 

57.7 

61.2 

85.2 

91.3 

87.3 

69.2 

59.1 

68.5 

- 

85.3 

- 

65.6 

61.4 

63.3 

58.2 

61.0 

54 

54 

56 

54 

51 

51 

52 

51 

52 

54 

52 

47 

45 

46 

46 

- 

- 

- 

- 

- 

7.9 

7.8 

8.1 

7.9 

7.4 

7.4 

7.5 

7.4 

7.5 

7.8 

7.6 

6.8 

6.5 

6.7 

6.7 

- 

- 

19 2.8 

18 2.6 

16 2.3 

18 2.6 

52 7.5 

52 7.6 

43 6.2 

49 7.1 

43 6.3 

46 . 6.6 

45 6.5 

45 6.5 

27 3.9 

24 3.5 

30 4.4 

27 3.9 

- - 

- - 

- - 

- - 

. 

8- 106 
GDC-CRAO-85905 

Tabio 8-41 Thormal Aging Data for 24 Ply [Oo 245'1, Celion 

psi) Air and Tostod at Room Tomporaturo 

ooo/LARc-1 0 G/PI Agod in 0.10 MN/d (14.7 

450 350 Baselineavg 

1.000 

5.000 

10,000 

0.0 

0.0 

0.0 

avg 0.0 

0.0 

0.0 

0.0 

avg 0.0 

- 

- 

-0.6 

-0.5 

-0.5 

505 450 1.000 

5,000 

10.000 

avg -0.5 

-0.2 

-0.2 

-0.2 

avg -0.2 

- 

-2.3 

-2.6 

- -2.7 

avg -2.5 

-3.7 

-4.8 

-5.4 

avg -4.6 

- 

591 85.7 

490 71.0 

481 69.7 

521 75.5 

497 72.1 

449 65.1 

470 68.2 

492 71.4 

470 68.2 

373 54.1 

45 1 65.4 

428 62.0 

- - 

- - 

417 60.5 

531 77.0 

519 75.3 

536 77.8 

401 58.1 

483 70.0 

415 60.2 

433 62.8 

467 67.7 

454 65.8 

520 75.4 

480 69.6 

559 81.1 

- - 

- - 

- - 

58 

49 

48 

48 

48 

45 

48 

49 

48 

47 

46 

47 

47 

46 

46 

43 

45 

47 

47 

47 

51 

50 

51 

50 

- 

- 

- 

- 

- 

- 

- 

8.4 

7.1 

7.0 

7.0 

7.0 

6.5 

7.0 

7.1 

6.9 

6.8 

6.7 

6.8 

6.8 

6.7 

6.6 

6.3 

6.5 

6.8 

6.8 

6.8 

7.4 

7.2 

7.4 

7.3 

- 

- 

- 

- 

- 

- 

- 

50 7.3 

54 

57 

63 

39 

53 

50 

57 

53 

48 

40 

50 

- 

- 

46 

63 

67 

58 

63 

44 

55 

41 

47 

29 

32 

33 

31 

- 

- 

- 

7.9 

9.2 

8.5 

7.7 

7.2 

8.2 

8.3 

- 

7.7 

7.0 

5.8 

7.2 

6.7 

9.2 

9.7 

8.4 

9.1 

6.4 

8.0 

5.9 

6.8 

4.2 

4.6 

4.8 

4.5 

- 

- 

- 

- 

a- io7 
GDC-CRAD-85-005 

Tablo 8-42 Thonnal Aging Data for PPQ Coatod 6 Ply [Oo r45°]s Ceiion 60OO/LARC-l60 G/Pl Agod in 0.10 MN/ 

m2 (14.7 psi) Air and Todd at Room Temporaturo 

Aging Wgh 

T.mp.fOtUm Aging Timo Change brnp~ivo Strongth bmpmdvo Modulus Shoar Strongth 

K (OF) (hrf (75) MNh2 (k4 CN/m2 (M.i) MN/m2 (kd) 

Baseline avg 584 84.7 52 7.5 51 7.4 

450 350 1,000 0.1 

0.0 

0.0 

avg 0.0 

5,000 -1.0 

0.0 

0.0 

avg 0.0 

- 

- 

10,000 -1.6 397 57.6 51 7.4 22 3.2 

- 1.9 259 37.6 44 6.4 21 3.0 

-1.7 249 36.1 50 . 7.2 21 3.1 

19 2.8 

15 2.2 

- - - - - - 19 - 2.7 

avg -1.7 302 43.8 43 7.0 19 2.8 

505 450 1.000 -1.0 

-0.8 

-1.0 

avg -0.9 

- 

5.000 -5.3 ' 

-4.9 

-5.1 

avg -5.1 

- 

10,000 -8.0 422 61.2 52 

-7.6 446 64.7 52 

-8.0 434 63.0 49 

- - - - 

avg -7.9 434 63.0 51 

7.6 34 4.9 

7.5 30 4.3 

7.1 24 3.5 

26 3.8 

30 4.3 

30 4.4 

7.4 29 4.2 

- - - 
Figure 844 Weight Chango of Celion 6000/LARC-160 G/PI After Thermal Aging ai the Indicated 

Temperature 

hours at the higher temperature. Neither of these specimens had been tested prior to polishing so 

that the cracks are the result of thermal aging and not mechanical testing. It was thought that 

these internal cracks might be responsible for the decrease in mechanical properties, Le., compressive, shear, and fatigue strengths after aging. However, cracks of this type were not seen in 

the 6 ply GIE material for any of the aging conditions up to 10,000 hours of exposure, although 

large decreases in shear strength were found. It is possible that the compressive and fatigue 

strength losses in the G/PI are related to the internal cracking, but a mechanism to explain the 

large shear strength decreases was not found. Certainiy, fatigue strength would be affected by 

the presence of matrix cracks. . 

Glass transition temperatures (T,) were determined from samples cut from the same specimens 

that were used for the metallographic examinations. These results are listed in Table 8-45 for the 

two aging temperatures. Unlike the data from the GIE material, the T, values show a trend to 

lower values as aging proceeds. The effect is somewhat greater for the higher aging temperature. 

In summary, thermal aging of Celion 60001LARC-160 G/PI at 450K (350°F) and 505K (450°F) 

produced significant reductions in compressive, shear, and fatigue strengths after 10,000 hours 

of exposure. In like manner to the thermal aging/tension results for this material, the compressive and 6 ply shear strength losses were higher for the 450K (350°F) exposures. The other properties, however, showed greater decreases at the higher aging temperature. Weight loss was also 

greater at the higher temperature. For the 6 ply specimens these losses were quite large, greater 

than 8%. Again, as with the G/E system, the property most affected was the shear strength of 

8- 109 
GOC-CRAO-85.005 

100 

I 

6 PLY .450K (%? I 24 PLY 450K (350' 

65321 7-386 

0.1 1 10 100 1000 10,000 

AGING TIME (HR) 

Figuro 8-65 Retontion of Comprosrivo Strength of Colion 6000/LARC-160 G/PI at 297K (75'F) After Thormal 

Aging at tho Indicatod Temporaturo 

the 6 ply material. Losses of 65% were found after 10,000 hours at 450K (350°F). Extensive edge 

cracking, particularly in the 24 ply specimens, was found as early as 1000 hours into the esposure. After 10,000 hours these cracks extended throughout the matrix. One set of 6 ply specimens was coated with PPQ in an effort to improve the thermooxidative stability of the G/PI 

material. Preliminary results were promising, with slight improvements over the uncoated specimens in host every case. 

8-1 10 
GOC-CRAO-85.005 

100 

80 

60 

40 

I 

0 6 PLY ' 450K (350°F) 

0 24 PLY 450K (350'F) A 6PLY 505K (450°F) 0 24 PLY 505K (450'F) I f 

2o 00.1 1 1 10 100 1000 10,000 

AGING TIME (HR) 

653217.387 . 

Figure 8-66 Retontion of Short Beam Shear Strength of Celion 6000/LARC-160 G/PI at 297K (75OF) After 

Thermal Aging at the lndicoted Temperature 

8-111 
GDC-CRAD-85-005 

Tabio 84 Axial Fotiguo Data for [Oo b45'], Colion 6000/LARC-160 G/PI Aftor 10,000 Hours of 

Thonnal Aging, 297K (75"F), R = -1, Unnotched 

Aging Temporaturo Spocimon Maximum Strosr 

K (OF) Numbu MNh2 (kri) Cycla Commonfs 

450 350 F-10-3-10 

-1 1 

-12 

-13 

-14 

-15 

-16 

-17 

-18 

505 450 F-10-3-28 

-29 

-30 

-3 1 

-32 

-33 

-34 

-35 

-36 

258 

293 

207 

224 

241 

241 

276 

190 

172 

207 

172 

138 

138 

138 

155 

190 

190 

172 

37.5 

42.5 

30 

32.5 

35 

35 

40 

27.5 

25 

30 

25 

20 

20 

20 

22.5 

27.5 

27.5 

25 

1,840 

400 

350,000 

122,000 

- 

94,000 

12,810 

38,320 

- 

1.280 

35,760 

22,760 

21,810 

6,531,000 

8,277,000 

440 

592,000 

7,569,000 

Machine malfunction 

Machine malfunction 

Bond failure 

Failed under doubler 

Bond failure 

Bond failure. 

Failed under doubler 

Failed under doubler 

8-1 12 
GDC-CRAD-85-005 

A 

40 

30 

20 

102 1 o4 105 106 

NUMBER OF CYCLES 65321 7-388 

Figuro 8-67 Axial Futfquo Pr0porti.r of [Oo f45"], Celion 6000/LARC-l60 G/PI Before and After 10,000 

Hours of Thormal Aging, 297K (75OF), R = -1, Unnotchod 

P 

Tablo 8-44 Effot, of PPQ Coating on Weight Chango and Comprosrivo and Show propertior of 6 Ply 

[Oo f45"], Celion 6000/LARC-l60 G/PI Aftor 10,000 Houn of Thermal Aging in Air e 

Aging Tampomtun 

Itom K (OF) Uncoatd PQQ Coated 

Weight Change (%I 450 

505 

Compressive Strength 450 

(% Retention) 

505 

Compressive Modulus 450 

(90 Retention) 

505 

Shear Strength 

(% Retention) 

450 

505 

350 -2.3 

450 -8.2 

350 72 

450 72 

350 100 

450 89 

350 35 

450 53 

-1.7 

-7.9 

52 

74 

93 

99 

38 

57 
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s 

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C N 

z 

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U 

0) 

C 

5 

e 

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0) 

m 

w 

w 

. 0 

8-1 15 
GDC-CRAD-85-005 

LUA 66321 7-391 

Figure 8-70 Edge Cracking in 6 Ply Celion 6000AARC-160 G/PI After 10,000 Hours of Aging at 505K (450OF) 

, in Air 

ORIGINAL PAGE B 

OF POOR QU8-1 16 
ORIGINAL PAGE TS 

OF POOR QUALITY 

- 

LL 

0 

VI 

d 

0 

- 

GDC-CRAD-85-005 

Y, 

0 

VI 

u. 

0 

VI 

m 

0 

1 

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0 In 

e 

a 

w 

Iz 

W 

W 

Id 

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aN 

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rl 

rN 

rl In 

(0 

c 

* 

H 

0 

L 

2 

0 

I 

0 

0 

8-117 
GDC-CRAD-85-005 

65321 7-393 

Figure 8-72 PPQ Coating on 6 Ply Celion 6000/LARC-160 G/PI Showing Good Adherence After 10,000 Hours 

of Thermal Aging'at 450K (350°F) in Air 

ORIGINAL PA&; -7 

OF POOR QUALITY 

8-118 
GDC-CRAD-85-005 

6632 17-394 

Figure 8-73 PPQ Coating on 6 Ply Celion 6000/LARC-160 G/PI Showing Varying Degrees of Peeling After 

10,000 Hours of Thermal Aging at 505K (450OF) in Air 

PPQ COATING REMOVED. N O T E A B S E N C E OF EDGE CRACKS 

65321 7 -395 

Figure 8-74 Edge View of 6 Ply Celion 6000/LARC-160 G/PI After 10,000 Hours of Thermal 

Aging at 505K (45OOF) in Air 

8-1 19 
0 

w 

w 

0 

w 

I 

a 

Y 

X 

0 

GDC-CRAD-85-005 

0 

s 

3 

pc 

\ 

0 

* 

5 

8 

8- 120 
GDC-CRAD-85-005 

Tablo 8-45 Class Transition Tornperatvn, Tg, Data for EO0 kaSO], Celion 6000/LARC-160 G/PI 

As- Received 

Aged 450 350 

Aged 450 350 

Aged 450 350 

Aged 505 450 

Aged 505 450 

Aged 505 450 

643 698 

1,000 646 702 

5,000 622 660 

10,000 603 626 

1,000 646 702 

5,000 598 616 

10,000 592 606 

. 

8-121 
ENVIRONMENTAL AGING 

At the time that this study was initiated, most of the data generated on advanced composite 

materials had been initial strength data without regard to environmental conditions. The small 

amount that was available was generally for only relatively short periods of exposure compared to the lifetime of a commercial airliner. This portion of the program was intended to 

evaluate the composite systems as a function of exposure to moisture, ambient aging, and 

atmospheric con taminants over relatively long periods of time. 

In Phase I, 24-hour water boil, &week humidity, and 20-week and 52-week ambient aging tests 

were conducted on the resin matrix composites as accelerated means of simulating long-term 

ambient exposures. Residual strength testing (flexure) of unidirectional and [Oo f 45"Is2 

crossplied specimens at room and elevated temperatures was performed after exposure and 

compared to tests on control (unexposed) specimens. For the B/AI system the Phase I environTensile and shear specimens of both coated and bare B/A were placed in an outdoor industrial 

seacoast atmosphere corrosion test facility maintained by General Dynamics Convair Division 

in San Diego, California. Residual strength was determined on specimens after 10,000 hours of 

exposure. 

For the Celion 6000/LARC-160 GlPI that was added after completion of Phase I, 24-hour water 

boa 6-week humidity, and 20-week and 52-week ambient aging exposures were carried out on 

[O O 2 45 "JS2 crossplied material No unidirectional Celion 6000/LARC-160 laminates were evaluated in the environmental aging study. After exposure, residual flexural strength tests were 

conducted at room and three eievated temperatures. 

mental study was a determination of the effects of corrosion and atmospheric con taminants. 

The Phase I1 task in the Environmental Aging portion of the program was a 50,000-hour outdoor exposure to determine the effects of weather and atmospheric con taminants on flexural 

properties of uncoated B/E, G/E, and HT-S/710 G/PI and tensile and shear properties of uncoated and coated B/AL These exposures were initiated eariy in Phase I. One batch of B/M 

specimens was removed during Phase I after 10,000 hours for examination and residual property testing. No interim evaluations were made on the resin matrix materials. 

The following sections describe the test procedures for the environmental aging studies and 

present and discuss the experimental results. 

9.1 RESiN MATRIX COMPOSiTE SYSTEMS 

9.1.1 SPECIMEN DESIGN AND TEST PROCEDURES. The longitudinal flexure specimen was used for 

all the environmental aging studies of the resin matrix composites. The flexure specimen does 

not provide data that can be used directly for design but is an effective and relatively inexpensive method for making comparisons between materials or evaluating effects of various expoposures. The specimen configuration is shown in Figure 9-1. As no universally accepted 

specification for flexural testing of composites was available, the test methods that were used 

were generally in accordance with ANWASTM D 790-71 and Federal Test Method Standard 

No. 406, Method 1031. Details of the test procedure are given in Appendix B. 

9- 1 
SPEC1 MEN 0 IM ENS10 NS 

LENGTH (1) 

WIOTH (b) = a01270 rn (0.500 in.) . THICKNESS (d) 

= (2076 rn (30 in.) 

8 

= APPROXIMATELY 0.0015 rn (0.060 in.) 

SPANnHICKNESS RATIO (i) a 32 TO 1 

LOAOING HEAO AN0 REACTION SUPPORTS ARE (200635 rn (0.250 in.) 

DIAMETER STEEL ROO 

OVERHANG MUST BE THE SAME OVER EACH EN0 

t 

d 

65321 7-1 11 

,- 

Flgun 9-1 longitudinal Fiexural Spocirnen 

Prior to conducting the moisture environmental exposures, the flexure specimens were heated 

at 394 K (250 OF) for 24 hours to ensure that all were in a dry condition at the start of exposure. 

The control specimens were also dried for 24 hours at 394 K (250 OF) before testing. All testing 

after ejrposure was accomplished within four hours after the exposure cycle had been 

completed. 

The 24hour water boil exposures were carried out in a reflux condenser using deionized water. 

An American Instrument Co., Inc. ClimateLab, Complete Humidity-Temperature Apparatus 

was used to conduct the &week humidity exposures. The chamber was maintained at 322 K 

(120 OF) with a controlled environment of 95 to 100% relative humidity. The ambient aging test 

specimens were kept in a desiccator at 294 f 3 K (70 f 5'F) and 52% relative humidity. The 

humidity level was maintained by a saturated aqueous solution of Na2Cr207*7H20. Specimens 

were removed from the desiccator and tested after 20 weeks and a second batch after 52 weeks. 

Painted redwood racks were made to hold the 50,000-hour outdoor exposure specimens. The 

racks were placed in an outdoor industrial-seacoast atmosphere, corrosion test facility (Figure 

9-2) maintained by General Dynamics Convair on the rooftop of a building in San Diego, 

California approximately 800 meters (0.5 mile) from San Diego Bay. 

The Phase I material environmental aging exposures were conducted with specimens cut from 

both unidirectional and [0 O f 45 'Is2 crossplied panels. Residual strength testing was done at 

297 K (75'F) and 450 K (350°F) on the unidirectional specimens for a€l exposure conditions. 

9-2 
Figuro 9-2 Outdoor Corrosion Test Facility 

For the crossplied specimens, only 450 K (350 OF) tests were perfoqned for the humidity, ambient aging, and weather exposures. The 24-hour water boil-crossplied specimens were tested at 

both 297 K (75°F) and 450 K (350°F). For the new material addition, Celion 60001LARC-160 

G/PI, only [0 " f 45 "Is2 crossplied specimens were used with residual strength testing conducted at 297 K (75 OF), 394 K (250"F), 450 K (350"F), and 505 K (450°F). Testing at elevated 

temperatures was accomplished by heating the specimens to the test temperature in a ConradMissimer chamber (held in a standard test machine-Instron, Tinius-Olsen or Baldwin), 

holding at temperature for 10 minutes before test, and then loading to failure. It has been 

found that the 10-minute hold is a good compromise for getting good temperature stabilization 

without driving off all the moisture. 

9.1.2 TEST RESULTS AND DISCUSSION. Residual flexural strength data after 24-hour water boil, 

6-week humidity, and 20-week and 52-week ambient aging exposures are presented in Table 9-1 

for BE, Table 9-2 for GIE, Table 9-3 for HT-S/710 GIPI, and Table 9-4 for Celion 6000/LARC-160 

for GIPI. Listed in the tables are the control and exposed values and the percent retention of ' 

flexural strength after the different environmental ages. In order of increasing effect on mechanical properties, the various exposures are generally ranked as one might predict, i.e., 20-week 

ambient, 52-week ambient, and either 6-week humidity or 24-hour water boil. The results are 

typical of those obtained by others (ref. 24 and 36). 

For epoxy systems, crossplied laminates traditionally have shown better 450 K (350°F) 

strength retention after exposure than unidirectional laminates. No definite theory has been 

- 

9-3 
Tablo 9-1. Hods of Moirtuun and Ambknt Aging on WE laminates 

F1oxurol Strongth/Tompo~ro 

297 K (750 F) 45OK (SOOF) 

Rotontion Rotention 01 Exporuro Orientation MN/d (kri) % MN/m2 (kri) /a 

Control 10 "112 2100 

2060 

304 

298 

1680 244 

1740 252 

1770 257 

1730 25 1 

889 129 

820 119 

299 - 

300 

2060 

2070 

[o" f 45"],2 924 

958 

134 

139 

135 - 

136 

931 

938 

882 128 

864 125 

1900 

1990 

276 

288 

415 60.2 

341 49.4 

428 62.1 

94 395 37.2 

43 1 62.3 

437 63.4 

758 110 

94 542 78.6 

724 105 

738 107 . 

586 * 85.0 

100 683 99.0 

690 100 

648 94.0 . 

586 85.0 

641 93.0 

1060 154 

1000 145 

1030 149 

100 1030 * 149 

869 126 

834 121 

284 - 

283 

1960 

1950 23 

63 

39 

74 

59 

100 

53 

94 

134 

130 

119 

128 

298 

318 

309 

308 

- 

- 

[O" 2 45"],, 924 

896 

820 

880 

6- Week [O "112 2060 

Humidity 2190 

2130 

2130 

[O" f 45"],, 

20-Week [O "I12 2190 

Ambient Age - 2120 

317 

308 

322 - 

316 

2220 

2180 

[O" 2 45"],, 

952 138 

885 128 

320 

334 

965 140 

862 125 

52-Week [O "112 2210 

Ambient Age 2300 

2320 

2280 

[Oo C 45"],2 

337 - 

330 

896 130 

100 908 132 

882 128 

772 112 

772 112 

809 117 
tablo 9-2. Effods of Moirtuun and Ambiont Aging on G/E Laminato8 

F1oxurol Strongth/Tomporutun 

297 K (75O F) 450K (SOOF) 

9 Retention Retontion 

Exporuro Oriontation MN/d (ksi) Y. MN/rn2 (ksi) % 

Control [O "112 

[O " f 45 "I,, 

10 "112 

[O" f 45"],, 

20-Week 

Ambient Age 

[O" t 45"],, 

[O "112 

[O" rt 45"],2 

52-Week 10 "112 

Ambient Age 

[O" rt 45"],, 

1720 

1710 

1620 

1680 

882 

779 

910 

857 

1850 

1620 

1580 

1680 

855 

834 

820 

836 

1720 

1640 

1680 

1680 

1960 

1690 

1900 

1850 

1670 

1660 

1540 

1620 

294 

248 

235 

244 

128 

113 

132 

124 

268 

235 

229 

244 

124 

121 

119 

121 

250 

238 . * 

244 

244 

7 

- 

- 

- 

- 

284 

245 

275 - 

268 

100 

98 

100 

100 

96 

1220 

1460 

1210 

1230 

658 

710 

676 

68 1 

424 

405 

464 

43 1 

345 

415 

334 

365 

405 

372 

388 

388 

268 

289 

27 2 

276 

634 

690 

793 

706 

609 

479 

535 

541 

605 

564 

598 

589 

394 

392 

376 

387 

177 

211 

176 

188 

95.4 

98.1 

98.8 

61.5 

58.7 

67.3 

62.5 

50.1 

60.2 

48.4 

52.9 

58.7 

54.0 

' 56.3 ' 

56.3 

38.9 

41.9 

39.4 

40.1 

91.9 

103 

100 

115 

102 

88.3 

69.5 

77.6 

78.5 

87.8 

81.8 

86.7 

85.4 

57.1 

56.8 

34.5 

56.1 

33 

54 

30 

41 

54 

79 

45 

67 

9-5 
Tablo 9-3. Efiods of Moirturo and Ambiont Aging on Ht-S/710 G/Pl Laminatos 

Ffoxuml Strongth/Tomporotun 

297 K (75O F) 450K (SOOF) 

Rotomtion Rotontion 

Exporuro Orionfation MN/d (kri) % MN/d (kri) % 

Controi 

&Week 

Humidity 

20-week 

Ambient Age 

52-W~k 

Ambient Age 

lo "I12 1330 

993 

1220 

1180 

[O" f 45"],, 665 

567 

701 

644 

[O "112 1140 

1020 

841 

1000 

[O" h 45"],, 578 

567 

351 

499 

[O O112 938 

. 889 

1190 

1010 

[O" -c 45"],2 

[O "I 12 1150 

1210 

1280 

1210 

[O" f 45"],, 

[O "112 1290 

1180 

1100 

1190 

[O " f 45 "Is* 

193 

144 

177 

171 

96.4 

82.2 

101.7 

93.4 

166 

148 

122 

145 

83.9 

82.3 

50.9 

72.4 

136 

129 

172 

146 

- 

- 

- 

167 

176 

186 

176 

- 

- 

- 

187 

171 

159 

172 

- 

- 

- 

85 

78 

85 

100 

100 

869 

952 

807 

876 

538 

66 1 

439 

546 

848 

986 

938 

92 4 

476 

374 

385 

41 2 

1080 

848 

848 

925 

348 

50 1 

486 

445 

703 

690 

827 

740 

523 

703 

590 

605 

869 

93 1 

1020 

940 

537 

323 

458 

306 

126 

138 

117 

127 

78.0 

95.8 

63.7 

79.2 

123 

143 

136 

134 

69.0 

54.3 

55.9 

59.7 

156 

123 

123 

134 

50.6 

72.6 

70.5 

64.5 

102 

100 

120 

107 

102 

75.9 

85.5 

87.8 

126 

135 

148 

136 

77.9 

75.9 

66.4 

73.4 

100 . 

75 

100 

81 

84 

100 

ZOO 

93 

9-6 
0 "E ss 

c 

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c 

0 

m 

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0 

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53 

0 

53 

m 

Q, 

0 

2 

8 

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c 

0 

0, 

0 

2 

9-7 
advanced to explain this, but contributing factors might indude specimen configuration, builtin thed stresses from curing, compressive mode of failure after exposure, etc. An exami~- 

tion of the epoxy flexural strength data (Tables 9-1 and 9-2) shows that after long term aging, 

the failure strengths for both unidirectional and crossplied laminates are nearly the same. This 

would be anticipated for a matrix dominated failure mode. The mom temperature flexural 

strengths were nearly unaffected by any of these exposures, while those at 450 K (350 OF) were, 

in general, severely degraded; the only exceptions were the crossplied BIE material after the 

20- and 52-week ambient age. The unidirectional specimens were particularly affected with 

losses in the 450 K (350'F) flexural strength of from 77 to 41% depending on the type of 

exposure. 

The effects of the moisture exposures on the polyimide specimens were generally less damaging than for the epoxy systems. For the HT-SI710 some decreases in flexural strength of the 

crossplied material were observed at 450 K (350°F) after the 24hour water boil and 6-week 

humidity tests, but no significant effects were obsenred as a result of the 20- and 52-week 

ambient ages. The Celion 6000/LARC-160 behaved even better with almost no loss in flexural 

strength at elevated temperatures and only minor losses at room temperature. 

The Phase I1 task in the environmental aging portion of the program was a 50,000-hour outdoor exposure to determine the effects of weather and atmospheric con taminants on flexural 

properties of uncoated BIE, GIE, and HT-SI710 GIPI. At the completion of 50,000 hours of 

exposure it was discovered that six of the nine GIE specimens and four of the nine G/PI specimens were missing. Whether the missing specimens were the result of vandalism or unusual 

weather conditions could not be determined. After exposure the recovered specimens were 

photographed and examined by eye and with a low power stereo microscope. Before and after 

photographs of the three systems are shown in Figures 9-3 through 9-6. The foilowing is a 

description of the condition of the different batches of material: . 

BIE - The outer glass scrim cloth layer on the upper surface had been loosened and was no 

longer present on seven of the specimens and was quite loose on the other two. There were no 

loose boron fibers, but some shallow grooves were visible between the fibers, perhaps an indication of the loss of resin from the surface. The lower sides of the specimens did not show much 

change. The specimens were held in the racks only at the ends so that both upper and lower 

surfaces were exposed to the atmosphere. The lower surfaces were, however, shielded from 

direct sunlight, rain, and any falling particulate matter. 

GIE - The three GIE specimens had a considerable number of graphite fibers exposed on the 

upper surfaces as a result of the loss of resin. The undersides showed only a very few exposed 

fibers. 

GIPI - On the upper surfaces of the specimens the outer portion of the .polyimide resin had 

also been lost so that considerable graphite fibers were exposed. On the surfaces facing the 

rack no loose fibers were visible. 

Residual flexural strength data after the 50,000 hours of exposure are presented in Table 9-6 

for the three materials. Also listed are the percent retention values based on the control data 

shown in Tables 9-1 through 9-3. For BIE the effect of the 50,000 hours of exposure was more 

severe than that for the 20-and 52-week ambient exposures but less than those observed after 

the 24-hour water boil or the 6-week humidity tests. The room temperature strength was nearly 

9-8 
653217-321 

Figure 9-3 Corrosion Racks Containing BIE and G/E Specimens - Before Exposure 

unchanged while the 450K (350 OF) properties were degraded approximately 50 and 25 percent 

respectively for the unidirectional and the [0 O A 45 "1 crossplied layups. The loss in elevated 

temperature strength is believed to have been primarily the result of absorbed moisture in like 

manner to the results obtained during the Phase I studies. The limited results obtained from 

the three unidirectional G/E specimens indicate that the 50,000 hours of exposure produced 

effects at 450K (350 OF) that were no worse than those obtained after any of the exposures conducted during Phase I with the possible exception of the 20-week ambient age of the crossplied 

layup. While the HT-S/710 GlPI data are also limited, it appears that the 50,000 hour weather 

exposure has been more damaging to the flexural properties of this system than any of the 

Phase I exposure conditions. 

Fiber reinforced resin matrix composites (particularly epoxies) absorb moisture from the 

atmosphere. This absorption causes two basic problems, Le., 1) losses in matrix dominated 

mechanical properties, and 2) changes in physical dimensions. The effects are mostly reversible 

in that heat or vacuum will drive off the moisture and will return the composites to their dry 

control properties and dimensions. The absorption of moisture primarily occurs in the resin 

matrix, and the resulting changes in composite mechanical and physical properties are a result 

of this absorption by the resin. Therefore, the type of reinforcement does not greatly influence 

the moisture effects. This was seen in the results presented earlier in that both the boron and 

graphite reinforced epoxies exhibited considerable property degradation from the absorbed 

moisture. 
Figure 9-4 Corrosion Racks Containing WE and G/E Specimens - Exposed for 50,000 hours in 

an Industrial-Seacoast Atmosphere 

The general moisture problem was discovered by General Dynamics Convair Division in 1970 

(ref. 8 and 24). The problem manifested itself initially by unusually low longitudinal flexure 

strength at 450 K (350°F) in GIE after short terms of aging at ambient temperature and 

humidity in air conditioned laboratories. Extensive work on this problem (ref. 36) led to the 

conclusion that the cause of the problem was plasticization of the epoxy resin by moisture, 

which subsequently lowers the glass transition temperature of the resin. The high temperature 

composite properties generally affected by moisture absorption are shear strength and 

modulus, edge compressive strength, bolt bearing and flat-wise tensile strengths. Since the initial work at General Dynamics Convair significant efforts have been made and are still in progress at several laboratories in this country to determine moisture effects on mechanical 

9-10 

ORIGINAL PAGE IS 

OF POOR QUALITY 
6532 17-323 

Figuro 9-5 Corrosion Rock Containing HT-S/710 G/PI Specimens - Before Exposure 

properties of polymer matrix composites. Many of these early results were presented in 1975 

at a U.S. Air Force Materials Laboratory and Society of Aerospace Materials and Process 

Engineers Workshop, which examined the environmental durability of resin matrix composites. A summary of the data presented along with the conclusions and recommendations of 

the participants can be found in reference 37. 

9.2 METAL MATRIX COMPOSITE SYSTEM 

9.2.1 SPECIMEN DESIGN AND TEST PROCEDURE. The tensile and interfiber shear specimen con- 

. figurations used for the BlAl environmental study were identical to those described previously 

in Section 7, Baseline Testing. All specimens were machined from 6 ply unidirectional BlAl. 

Tensile specimens were cut from both the longitudinal and transverse directions while the 

interfiber shear specimens were cut from only the longitudinal direction. 

In addition to determining the susceptibility of B/A1 composite material to corrosion and 

atmospheric contaminants, an objective of the study was to evaluate an elevated temperature 

corrosion protection system for BlAL In previous work at General Dynamics Convair (ref. 38) 

it was found that BlAl composite materials are somewhat more susceptible to corrosion than 

aluminum structures, especially at the edges, where both boron and aluminum are exposed to 

the corrosive environment. In these same studies, effective corrosion resistance was achieved 

by a paint finish system applied over a chemical film treatment. These systems, however, were 

effective to only approximately 422 K (300°F), somewhat below the range desired for the SCR 

application. Later work on a corrosion protection system for BlAl composites serviceable up to 

589K (600 OF) was also accomplished at Convair (ref. 39). While not completely effective in 

preventing corrosion of the aluminum, the polyimide, silicone, and fluorocarbon coating 

systems studied showed promise. Using these results as a starting point, a limited test program was conducted to develop a coating for use on these environmental exposures. The finish 

system that was selected consisted of a chemical conversion coating followed by an epoxy 

primer and a polyimide topcoat. Before the chemical conversion coating, the specimens were 

degreased, alkaline cleaned, and deoxidized. The epoxy primer was applied over the chemical 

9-1 1 
0 

0 

0 

v) 

9 

9-12 

8 

2 

I 

c.. 

W 

0 

P 

x 

W 
tablo 9-5. Effufs of 50,000-hour Wdor Exposure on Floxural Strongth of WE, 

WE, and HTW710 G/P! 1aminat.r 

Floxural Strongth/Temporutun 

297 K (75O F) A5OK (SOOF) 

Retontion Retention 

Motorial Orionfation MN/d (kri) % MN/d (kri) % 

BIE 

GiE 

[O "112 2020 

1860 

[O" f 45"],, 

. , 

1830 

1900 

[O" f 45"],, 

HT-SI710 G/PI [0 "112 972 

945 

958 

[O" f 45"],2 

293 

270 

266 - 

276 

141 

137 - 

139 

869 

800 

779 

92 816 

5 68 

550 

820 

646 

786 

685 

736 

341 

632 

724 

81 678 

425 

126 

116 

113 

118 47 

82.4 

79.7 

93.7 75 

119 

114 

99.3 

107 37 

49.4 30 

91.6 

105 

98.3 77 

61.7 78 

conversion coating and allowed to dry. Aluminum powder pigment (5910 by weight) was added 

to the polyimide coating material before application over the primer. After application of the 

aluminized polyimide coating, the specimens were cured at 436 K (325 OF) for 14 hours. This 

coating has a maximum use temperature of about 389 K (600°F). Half of the BlAl specimens 

used for the environmental exposure tests were coated and the other were bare. 

The same type of painted redwood racks that were used for the resin matrix specimens were 

made to hold the B/Al specimens during exposure. The racks were placed in the outdoor industrial-seacoast atmosphere, corrosion test facility shown in Figure 9-2. Half of the coated and 

half of the bare specimens were removed after 10,000 hours for examination and residual 

strength testing. The remaining half of the specimens were removed after 50.000 hours of 

exposure. Tensile and interfiber shear tests were conducted at room temperature using the pre 

cedures described in Section 7. 

9.2.2 TEST RESULTS AND DISCUSSION. After 10,000 hours of outdoor exposure no change in 

appearance was observed in the coated B/Al specimens, but the bare specimens all showed 

evidence of extensive surface corrosion with some concentrated attack at localized areas (pitting). This attack was much heavier on the side of the specimen facing away from the rack. The 

samples were held in the racks only on the ends so that both front and back surfaces were 

exposed to the environment. Photographs of the specimens mounted in the wooden racks 

before exposure are shown in Figure 9-7, and after 10,000 hours in Figure 9-8. The corrosion of 

the uncoated specimens is clearly visible in Figure 9-8. 

9-13 
0- 

0 

- Q) 

0- 

0 

- Q) 

0- 

0 

L 

2' 

2 

a 

0 

X 

W 

2 

m 

0 - 

Q) 

I 

I) 

c 

Q) 

- E 

U 

Q) 

a 

v) 

e 

0 

0 

0 

u 

- 

u 

L 

JNGINAL PACE rs 

9-14 al? PQOR QUALITY 
9-15 
The results of post esposure tensile and shear testing are presented in Tables 9-6 and 9-7. Only 

typical values of the unesposed strengths are given because the environmental aging 

specimens were cut from different panels than those used for the baseline tests. No effects to 

either tensile or shear properties were observed for the coated samples. For the uncoated 

specimens, significant decreases were found in the transverse tensile strength and the shear 

strength, while the longitudinal tensile strength was unaffected by the 10,000-hour exposure. 

A surface effect such as corrosion of the bare aluminum would be expected to decrease the 

matrix controlled properties, ie., transverse tensile and shear, but have little effect on 

longitudinal tensile, a fiber controlled property. 

To check the extent and nature of the surface corrosion, a metallographic examination of bare 

and coated specimens was performed. No change in appearance was found for the coated 

specimen after the 10,000-hour exposure as the coating had provided complete protection. 

Examination of the uncoated specimen, on the other hand, revealed extensive intergranular 

attack at the surface, which extended to a depth of approximately 0.00007 m (0.003 in.). An 

example of these corrosion crack regions is shown in Figure 9-9. Some surface disruption has 

also occurred near one of the boron filaments. 

The uncoated interfiber shear specimens, see Figure 7-30, had been prepared prior to placing in 

the test facility so that the machined edges as well as the surfaces were exposed to the atmosphere. These specimens showed a decrease in shear strength of approximately onethird. Later, 

a second set of shear specimens was prepared by remachining half of each of the three failed 

specimens. For these specimens, only the surface of the test area had been exposed and subjected to corrosive attack. The shear strength of these specimens, Table 9-7, was only slightly 

lower than the coated materiaL The configuration of the shear specimen used for these tests is 

one in which stress concentrations build up at the machined surfaces at either end of the shear- 

' ing regions. If these surfaces are exposed to a corrosive environment, the subsequent corrosion' 

cracks can initiate failures at loads substantially lower than those obtained from undamaged 

specimens. It is this phenomenon that is believed to have been responsible for the significant 

effec't on shear strength of machining the notch before or after exposure. 

Photographs of the 50,000 hour B/Al specimens before and after exposure are shown in Figured 

9-10 and 9-11. Again, the coating provided excellent protection with no visible failures. The 

upper surfaces of the specimens contained a small amount of debris but no evidence of corrosion. The bottom surfaces were quite clean. The uncoated specimens, on the other hand, showed 

considerable evidence of surface corrosion with extensive pitting. The attack was much more 

severe on the side of the specimen facing away from the rack. This upper surface als~ contained a great deal of debris, much of which had the crystalline appearance of salt. These 

uncoated 50,000 hour specimens showed much greater surface degradation than did those 

(Figure 9-8) that had been removed after 10,000 hours. At the completion of the 50,000 hours of 

exposure it was found that two of the notched shear specimens had been broken in the shearing regions. These are the first and last of the set of shear specimens shown in Figure 9-11. No 

reason could be found for these unexpected failures. 

The results of post-exposure tensile and shear testing are presented in Tables 9-6 and 9-7 along 

with the data from the 10,000 hour specimens. The tensile properties of the coated specimens 

in both the longitudinal and transverse directions were still relatively unaffected even after 

50,000 hours of exposure. For the bare specimens, however, both orientations showed lower 

strengths after exposure. The longitudinal specimens that showed no effects after 10.000 

hours were down in tensile strength from 4 to 12% after 50,000 hours. 

9-16 
tablo 9-6. tonsiio Data for Bar. and Coatod [Ool6 B/AJ Aftor Atmorphoric Contaminants 

Exporurm in an Industrial-Soacoasi Environmont 

typical Strmgth Strmngth Aftor Stnngth Aftor 

Unoxpod 10,000 Hours 50,000 Hours 

Condition Oriontation PAN/& (W PAN/& (kri) MN/& (kri) 

[90 "1 17.6 

18.4 

139 

90.3 

20.2 

13.1 

121 

127 

134 

22 127 

19.3 - 

18.5 

13.1 - 

15.5 

90.3 

152 106.5 

Coated [90 "1 176 

156 

25.6 

22.7 

152 

190 

22.1 

27.3 

24.2 - 

24.2 

23.5 - 

24.4 

167 

22 166 

162 

132 168 

[O "1 1410 

' 1400 

205 

203 

226 

211 

185 

194 

202 - 

194 

1280 

1340 

1390 

1340 

1560 

1450 210 1460 

Coated [O "I 1500 

1500 

217 

218 

202 

212 

192 

227 

195 

205 

-. 

1320 

1560 

1340 

1410 

1390 

1450 210 1460 

The average tensile strength of the bare transverse specimens had decreased nearly 30% after 

50,000 hours of exposure. The 10.000 hour decrease had been about onehalf this value. Again, 

corrosion of the aluminum matrix would be expected to affect the matrix controlled transverse 

strength to a greater degree than the fiber controlled longitudinal strength, and such was the 

case. The other matrix controlled property, interfiber shear strength, was also significantly 

decreased by the 50,000 hours of exposure (approximately 30%). However, in this case, both 

the bare and the coated specimens were affected. Remachining the shear configuration in one 

half of each of the failed specimens and retesting did not appreciably affect the shear strength 

values as was found after the 10,000 hour exposures. The decrease in shear strength of the 

uncoated material after both the 10,000 hour and 50,000 hour exposures is thought to have 

been caused by the surface corrosion of the aluminum at either end of the shearing regions. No 

explanation has been found, however, for the shear strength drop in the coated specimens that 

did not experience any surface corrosion. To verify the absence of corrosion on the coated 

specimens and to determine the extent and nature of the corrosion on the bare specimens, additional metallograhpic examinations were conducted. As shown in Figure 9-12, no change in 

appearance was found in a coated specimen after the 50,000 hour exposure as the coating had 

provided complete protection. Examination of an uncoated specimen on the other hand, 

revealed extensive intergranular attack near the surface, which extended to a depth of approximately 0.00018 m (0.007 in.). The attack, shown in Figure 9-13, was considerably deeper and 

more widespread than that observed for a specimen that had been exposed for 10,000 hours 

(Figure 9-9). 

9-17 
= 

H ?Yo9 * 

c 

0 

0 

C 

I 

L 

c 

m 

2 

2 r( 

2 

2 4 

NI 

13 

I 

0 

2 

2 d 

2 

2 

d 

II 

I 

0 

9-18 
Figure 9-9 Intergranular Corrosion Cracks in Unidirectional B/Al After 10,000 hr of Atmospheric 

Contaminants Exposure in an Industrial-Seacoast Environment 

9-19 
P 

N 

0 

0 

w 

tU 

0 

0 

?! 

a 

: 

X 

Ly 

n 

0 

1 

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9-20 

GPJGINAE PAGE rs 

DB KXlR QUALITY 

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0- 

0 - Q) 

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0 

a 

0- 

0 

- 03 

0- 

1 

0 
Figure 9- 12 Coated, Unidirectional B/Al After 50,000 hr of Atmospheric Contaminants Exposure 

in an Industrial-Seacoast Environment. No Corrosion has Occurred 

Figure 9-13 Intergranular Corrosion Cracks in Unidirectional B/Al After 50,000 hr of Atmospheric 

Contaminants Exposure in an lndustrial-Seacoast Environment 

9-22 
SECTION 10 

CREEP TESTlNG 

The objective of this task was to determine experimentally the creep and creep rupture behavior 

of the four advanced composite materials selected for evaluation in Phase I of the SCR program. Axial creep tests were performed at two temperatures for exposure times of 100 and 1000 

hours for each system. The two test temperatures were the proposed maximum use temperature for each material and one selected lower temperature in the potential application range. 

Results obtained during Phase I of this contract study have since proven these proposed maximum use temperatures to have been overly optimistic for all four composites. The lower temperatures were, however, well chosen, and agree quite closely with those used for the flight 

simulation tests. Creep properties were determined for two layups, unidirectional and [0 O k 45 'Is 

crossply, of each materid All specimens were taken from 6 ply laminates. The creepicreep rupture test program is summarized in Table 10-1. 

For the Celion 6000/LARC-160 G/PI that was added after completion of the Phase I creep testing, a decision was made to eliminate the creep testing portion from the baseline material characterization. This was done because of the results obtained in this section on the four Phase I 

composites. Composite materials with 0 O plies aligned in the direction of loading were observed 

to undergo very little creep deformation prior to failure. These small amounts of strain and the 

scatter in the measurements combined to make creep testing a questionable parameter for 

material characterization of filamentary reinforced composites. 

10.1 SPECIMEN DESIGN 

No standard specimen for creep testing of composites had been established at the time the 

creep testing task was initiated. Previous work at General Dynamics Convair (ref. 24 and 38) 

with a straight sided specimen similar to that used for tensile testing had given good results 

for material with no fibers aligned in the loading direction. For specimens with fibers in the 

loading direction the high loads required resulted in frequent bond failures in the end doublers. 

A slightly modified specimen was used for this program. To reduce misalignment and ensure 

axial loading the grips were changed from wedge to pin loaded. Tapered titanium doublers 

were bonded to the grip areas to minimize stress concentrations. The length of the specimens 

was such that it would help prevent premature grip failures by permitting the grip ends to 

extend outside the furnace. A length of 0.56 m (22 in.) was selected, and the BIE and BIAl specimens were prepared. No problems were encountered with bond failures in the B/E specimens, 

but four of the B/A1 specimens failed prematurely in the doubler bonds. As heating of the 

adhesive during testing was thought to have contributed to the problem, the length of the GIE 

and GIPI specimens was increased to 0.71 m (28 in.) to further extend the @p ends outside the 

furnace. While che longer specimen undoubtedly helped, the combination of high loads and 

temperatures also resulted in several bond failures in both the GI'E and GiPI systems. The configuration of the creep specimen is shown in Figure 10-1. 

10.1 
Tablr 10- 1. Summary of CreopKroop Rupturo Test Program 

Lominatr Tempuature Tlmo Numbor of 

Motorial Oriontation K (OF) (hr) Spdmons 

- 

550514 BIE [0" k 45 "Is 

[o oi6 

A-Sl3501-5 GIE [O" f 457, 

HT-2l710 GIPI [0" -C 45'1, 

[o 0i6 

BlAl [O" 4 45 "1, 

394 250 

450 350 

450 350 

394 250 

450 350 

394 , 250 

450 350 

505 450 

561 550 

505 450 

561 550 

505 450 

561 ' 550 

505 450 

561 550 

100 

1000 

100 

1000 

100 

1000 

100 

1000 

100 

1000 

100 

1000 

100 

1000 

1000 

100 

100 

100 

100 

1000 

100 

1000 

100 

100 

4 

2 

5 

2 

4 

1 

5 

1 

4 

2 

6 

6 

6 

6 

3 

3 

3 

3 

6 

6 

10.2 TEST EQUIPMENT AND TEST PROCEDURE 

All creep tests were performed in air using Arcweld Creep Testing Machines, Model JE, 5440 

kg (12,000 lb) or Model EE, 9070 kg (20,000 lb), and Arcweld Power Positioning Furnaces, 

Model F8. The furnaces are 0.43 m (17 in.) in length with 0.064 m (2.5 in.) diameter ceramic muffle tubes in which the specimens are radiant heated. The resistance heated furnaces have three 

zones that can be manually adjusted to minimize temperature gradients along the central test 

section of the specimens. A centrally located viewing port is available for optical measure 

ments. Temperature control for the furnaces is provided by Wheelco controllers using chromelalumel thermocouples inserted through the wire heating elements and in contact with the outside of the muffle tube. The test temperature was measured with three chromel-alumel thermocouples attached at 0.0444 m (1.75 in.) intervals to the gage section of the test specimen with 

high temperature fiberglass tape. Temperature variations at any one location were generally 

less than k 1.75 K (3" F) while temperature uniformity along the 0.0889 m (3.50 ir~) section 

was generally maintained to f 2.8 K (5O F). Strain measurements were made with an optical 

10-2 
LOADING HOLES ON 

CENTER LINE OF 

SPECIMEN TO WITHIN 

?: 0.000064 rn (0.0025 in.) 

BONDED OOUELERS4 PLACES 

DIM ENSIONS 

IN rn (in) 

65321 7- 1 18 

Figure 10.1 Croop lest Spocirnen 

10-3 
method employing a platinum sliding-strip extensometer mechanically attached to the 

specimen. The optical equipment consisted of a Gaertner filar micrometer 50X magnification 

microscope mounted on the furnace and sighted through the central viewing port. Illumination 

of the platinum extensometer was provided by a lamp system within the microscope. 

A creep machine and furnace with a specimen in test is shown in Figure 10-2. A closeup view of 

the Gaertner microscope in place on the furnace can be seen in Figure 10-3. Figuse 10-4 shows 

the platinum sliding-strip extensometer attached to a composite specimen prior to loading into 

the creep testing apparatus. Also visible are the chromel-alumel thermocouples. 

Prior to testing, the resin matrix composite specimens were dried at 394 K (250" F) for 24 

hours. The specimen dimensions were measured, and the creep machine loads were calculated 

for the particular stress level desired, Fiducial marks were lightly penciled at 0.0508 m (2 in.) 

intervals in the region where the platinum extensometer was to be attached in order to determine the total strain by before and after measurements with a cathotometer. These measure 

ments were later discontinued when the scatter in the readings was found to be of the same 

magnitude as the total strains obtained during creep testing. The platinum extensometer was 

attached, and the three thermocouples were taped to the specimen, above, below, and opposite 

the extensometer. The specimen, which was connected to the upper and lower pullrods, was 

then carefully loaded through the preheated'furnace and attached to the loading arms of the 

creep machine. A small load, 0.45 kg (1 lb), was applied while the temperature was stabilized, 

and the microscope was focused and positioned on the platinum extensometer. The full load 

was then gradually applied, and strain measurements were begun immediately. The elapsed 

time between the initial small load application and the full load application was approximately 

0.25 hour. The difference in measurements between zero and full load is due to elastic strain. 

These values are shown on the individual creep curves but are not reported in the total strain 

values presented in the tables. Optical measurements were continued for either 100 hours or 

for 1000 hours or until specimen failure. At the conclusion of the tests for those specimens that 

did not fail, the load, except for 0.45 kg (1 lb), was removed, and a final strain measurement was 

recorded. The final measurement was generally not made at exactly 100 or 1000 hours because 

of a one shift operation in the testing laboratory. Consequently, the creep curves are plotted 

out to the times of the last measurement. The values listed in the data tables for total creep 

strain, however, have been, when possible, taken from the creep versus time wes at either 

100 or 1000 hours. 

10.3 TEST RESULTS AND DISCUSSION 

The creep testingprogram, which is summarized in Table 10-1, consisted of 100- and 1000-hour 

tests at two temperatures and two layups for each of the four material systems. The data are 

presented in tabular form and as creep strain versus time curves. The elastic strain values 

obtained during loading are shown on the individual creep wes but are not reported in the 

total creep strain values presented in the tables. Also shown on the creep curves are the strain 

values after unloading at the conclusion of a test in which failure did not occur. The difference 

between the initial full load strain and the final unloaded strain should be equal to the permanent plastic creep strain. However, as a result of the very sd creep strains obtained in these 

tests the agreement was generally poor. Only for the [0 O f 45 'Is crossplied BlAl and B/E was 

the agreement reasonably close. Again, these values were not reported in the tabular data but 

can be obtained from the creep curves. 

10-4 

Figure 10-3 Furnoce Mounted Filar Micrometer Microscope for Creep Measurements 

10-6 

ORIGINAL PAGE 13 

OF POOR QUALIm 
. 

Figure 10-4 Creep Specimen with Platinum Sliding-Strip Extensometer and Thermocouples Attached 

The specimens that failed during load application and for which no curves could be drawn are 

identified in the tables. Also noted in both the tables and on the curves are the specimens that 

failed during testing or that experienced doubler bond failures during testing. 

10.3.1 BORON/EPOXY. Results of creep testing of BIE at 394 K (250" F) and 450 K (350" F) are 

listed in Table 10-2. Data are presented for 18 specimens. 

Creep strain versus time curves are presented in Figures 10-5 and 10-6 for crossplied material 

at 394 K (250" F), in Figures 10-7 through 10-9 for unidirectional material at 450 K (350" F), 

and in Figures 10-10 and 10-11 for crossplied material at 450 K (350" F). 
cn 

L 

1 

0 

c 

Q, 

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5: 

m 

0 

m 

v 

0 

ua 

m 

10-8 
1.0 

a9 

0.8 

a7 

a6 

st 

$03 

a 

0 

0.4 

0.3 

QT 

a1 

1 

1 

1 

1 

1 

1 

1 

ELASTIC STRAIN 

j 1 1 

1 1 

0 I Ill 1 I I 1 ' Ill 1 I I1 Ii I/ 

0.1 1 10 loo 1000 

i I 'I11 

1 1 I I Ill1 1 11 I I Ill 

SPEC NO. STRESS. MNld (Ut) 

0 ACS-11 310f4S) 

0 AC57 379(55) 

A ACS4 345(50l 

1 I I I1111 1 I 1 I ! II!, 

1.0 

0.8 

a7 

0.6 

st 3 as 

0.4 

a 

u 

03 

a2 

a1 

0 

I 

SPEC NO. STRESS. MN/d (GI1 

0 ACS4 414(60l 

A A610 448(651 

Figure 10-5 Croep Strain for [Oo f 45O], B/E at 394 K (2500 F) 

633217- I21 

Figure 10-6 Creep Strain for [Oo f 45O], B/E at 394 K (2500 F) 

10-9 
0.1 1 10 too low 

TIME (hr) 

SPEC NO. STRESS. MN/d (KSl1 

0 AU51 891(130) 

A AU5-7 lOJO(150l a53217 122 

Figuro 10-7 Creep Stroin for [OO], B/E e 450 K (3500 F) 

0.1 1 10 100 

TIME (krl 

SPEC NO. STRESS. MN/d (KSII 

0 AU54 931 (1351 

A AU5.2 985 11401 

653217 123 

Figure 10.8 Creep Stroin for [Ool6 B/E a) 450 K (3500 F) 

10-10 
0.2. 

ai 

0 

Figun 10-9 Croep Strain for [Ool6 WE at 450 K (3500 F) 

I 

I I 

A I 

I 

ELASTIC STRAIN 

0 

I1 8 I I I Ill1 I I I '"11 I I I I1111 I I ! llI!l 

1 1 10 100 tow 

TIME lhr) 

SPEC NO. STRESS. MNd (MI) 

0 ACS.13 YS(5Ol 

A ACS-1 414(601 

653117.125 

I 

I 

I 

I 

Ill11 I I 1 I II I11111 I 111' 1 I * /!I1 

Figuro 10-10 Creep Strain for [Oo -C as0], B/E at 450 K (350O F) 

10.11 
55321 7.128 

Figuro 10-11 Ctoop Strain for [go f 45O], B/E at 450 K (3JOO F) 

The data show that the amount of creep obtained in B/E material with boron filaments aligned 

in the loading direction is very small. No specimen gave values of more than a few tenths of a 

percent of permanent creep strain. For example, specimen AC5-8 (crossplied) exhibited only 

0.13% creep in 1000 hours at 394 K (250 O F) when loaded at 79% of the average ultimate tensile 

strength, and specimen AU5-5 (unidirectional) exhibited only 0.27% creep in 1000 hours at 450 

K (350 O F) when loaded at 88% of the average ultimate tensile strength. Specimens that failed 

prior to completing 100 hours of exposure also underwent very little creep as indicated by 

measurements just before fracture. See, for example, Figures 10-6 and 10-9. One specimen, 

AU5-3, Figure 10-7, did show a sharp rise in the creep curve before fdure, but the total creep 

strain.was still less than 0.2%. Before-after measurements were not successful in determining 

the total creep strain of failed specimens. For specimens that fractured in the gage section the 

combination of low creep strain and the difficulty of accurate fit-back of the failed specimen 

halves made the measurements unreliable. 

In analyzing the creep data an effort was made to prepare composite creep curves for 0.05, 

0.10,0.15,0.20, and 0.25% creep by determining the time periods required for varying amounts 

of creep strain as a function of stress level. The limited amount of creep data available, the 

small amount of creep strain measured, and the scatter in the data combined to give results 

that were unsuitable for the plotting of this type of creep curve. This was found to apply to ail 

the other composite systems as well as the BIE. 

10.3.2 GRAPHITE/EPOXY. Results of creep testing of GIE at 394 K (250 O F) and 450 K (350 O F) 

are listed in Table 10-3. Data are presented for 24 specimens. 

10-12 
50 

c 

20 

2s 

CE 

00 

BZ -cI 

z2 

.-. .-. 

I 

B 

2 

.-. 

ddead 

0 

9999 ma 

99 

ua 

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0 00- 8oFq loo 00 

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tea 

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arooomea~oooo coQJ-4d4 4 

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10-13 
Creep strain versus time curves are presented in Figures 10-12 through 10-14 for unidirectional 

material at 394 K (250" F), in Figures 10-15 through 10-18 for crossplied material at 394 K 

(250 " F), in Figures 10-19 and 10-20 for unidirectional material at 450 K (350 O F), and in Figure 

10-21 for crossplied material at 450 K (350" F). 

The G/E data are very similar to those for BIE in that the amount of permanent creep strain 

was very sd for all test conditions. For only two specimens was the creep ,mater than 0.1% 

and for these specimens the values were below 0.2%. An example of the small amount of creep 

observed in the GI'E material can be seen for specimen BC5-10 (crossplied). This specimen 

showed only 0.17% creep in 570 hours when loaded to 500 MN/m2 (72.5 ksi) or 95% of the 

average ultimate tensile strength. A similar specimen, when tested at 517 MN/m2 (75 hi), 

failed during application of the load in the test machine. Two of the crossplied specimens failed 

in this manner (during loading) but at stresses that were within the scatter band found during 

baseline testing at the same temperatures. Three of the unidirectional specimens also failed 

during loading. These failures, however, occurred at stresses well below the scatter band 

obtained in the baseline test program. Differences in material strengths were not the cause of 

the premature failures because both the baseline and the creep specimens were all cut from the 

same panel, LRC-1OA. Of the three specimens that failed on loading, one failed at the center, 

the second failed in two locations - at the center and approximately 0.04 m (1.5 in.) from the 

top doubler, and the third, also failing at two locations, at the edge of each doubler. It may 

have been the rapid application of load that led to the early failure of these specimens. The 

loads were added in incremental steps of 181 kg (400 lb) to reach a value of slightly more than 

907 kg (2000 Ib). 

10.3.3 GRAPHITE/POLYlMIDE. Results of creep testing HT-Sl710 GlPI at 505 K (450" F) and 561 K 

(550" F) are listed in Table 10-4. Data are presented for 24 specimens. 

1.0 

a9 

aa 

a7 

0.6 

3 2 as 

a4 

a 

U 

0.3 

a2 

ai 

0 

1 10 1 DO 1 ow ai 

TIME (hrl WEC NO. STRESS. MNld (KSII 

3 8UU 11% 1130) 

6U2 17. t 27 

Figure 10-12 Creep Strain for [OO], G/E at 394 K (2500 F) 

10-14 
1.0 

0.9 

a8 

0.1 

as 

j as I 

I I 

I a 

U I 

I 0.4 j ' 

a3; ! ! 

!I 

I I 

ELASTlC STRAIN 

TIME (hd SrCC NO. STRESS, MWd fK.Ul 

0 BUS-14 931 (135) 

A BUS4 IS (14001 

Figure 10-13 Cmp Strain for [Ool6 WE at 394 K (2500 F) 

1.0 

0.9 

a8 

a7 

as 

z 

yI as 

a4 

L 

YI 

U 0 

0.3 

a2 

01 

0 

0.1 1 10 too 1000 

TIME Ihr) 

WEC NO. STRESS. MN/& (ai) 

0 BUS4 1000 (145) 

653217 129 

Figure 10.14 Creep Strain for [00l6 G/E at 394 K (250O F) 

10.15 
12 

1.1 

1.0 

0.9 

an 

a7 

0.6 

as 

0.4 

ai 

a2 

a1 

0 

ai 1 10 1 00 1000 

0 BCS.7 414 (60) 653217.130 

ABCS.13 448(6S) 

SQ€C NO. STRESS. MWrd (KUt TIME (krl 

Figure 10-15 Creop Shin for [OO f Go], G/E at 394 K (250O F) 

8 

0 Y 

U 

a 

1.1 

1.0 

0.9 

aa 

07 

a1 

as 

ao 

B.3 

02 

ai 

0 

0.1 1 10 100 tow 

SPEC NO. STTRESS. MHIJ (KSI) TlPE (hr) 

65321 '31 0 8CS.a 446(65) 

A 8CS.11 436 (67.51 

Figure 10-16 Creep Strain for [OO -C 4501, G/E at 394 K (250O F) 

10-16 
12 

1.1 

1 .o 

0.3 

ai 

&I 

a8 

;z 

5 as 

a1 

y1 

Y 

a 

01 

a2 

ai 

0 

0.1 1 10 100 tow 

TIME lhr) 

653217.132 

Figun 10-17 Creop Strain for [OO f aO], G/E at 394 K (250O F) 

0.1 1 10 1000 

TIME (hr) 

SPEC NO. STRESS. MWI~ (KS) 

0 BCS.10 500 (72.51 553217 133 

Figure 10-18 Creop Strain for [Oo f 45O], G/E at 394 K (250O F) 

10-17 
1.2 

1.1 

1.0 

0.9 

as 

a7 

ac 

8 5 a5 

U X 

a4 

Q.3 

0.2 

01 

0 

IO 

TlME (hr) 

1000 

653217.134 

Figure 10-19 Crwp Stroin for [Ool6 G/E at 450 K (SOo F) 

1.2 

1.1 

1.0 

0.8 

aa 

a7 

0.6 

8 

0 

Y 

U 

w 0.5 

a 

0.4 

ai 

0 

ai 1 10 101 

SEC NO. STRESS. MN/~ (KSI) TlME (hr) 

0 BUS4 1034(150) 

A 8U5-13 1069(155) 

Figure 10-20 Creep Strain for [OO], G/E at 450 K (SOo F) 

10.18 

653211.135 
1.2 

1.1 

1.0 

Qo 

ai 

-. 

b.3 

0.. 

ai 

I I 

ELASTIC STRAIN 

- 

0.' ' I 1 I!) !I I I I IIIII ! 1 Ill I!! ! 1 IIIII 

TIME (hi) 

653217.136 

Flguro 10-21 Croop Strain for [Oo f 45O], G/E at 450 K (SOo F) 

Creep strain versus time cufves are presented in Figure 10-22 for unidirectional material at 550 K . 

(450 O F), in Figures 10-23 through 10-27 for crossplied material at 550 K (450 O F), in Figures 

10-28 and 10-29 for unidirectional material at 561 K (550 O F), and in Figure 10-30 for crossplied 

material at 561 K (530" F). 

Creep strain values obtained for the GlPI system at both test temperatures were also very low. 

For only one specimen, DCS-1 (crossplied), was a value greater than 0.1% recorded. Examples 

of the small creep extensions that were measured are the 0.0370 creep for specimen DU5-1 

(unidirectional) in 100 hours at 550 K (450 O F) when loaded at 85% of the average ultimate tensile strength and the 0.07% creep for specimen DC5-8 (crossplied) in 100 hours at 361 K (550 ' 

F) when loaded at 91% of the average ultimate tensile strength. 

The data obtained for the GiPI system were somewhat limited because of several doubler bond 

failures in the unidirectional specimens. These were caused by the combination of high shear 

stresses in the adhesive because of the high applied test loads and heating from the furnaces. 

Several specimens failures also occufred during application of the loads at the beginning of 

testing. In like manner to the G/E loading failures, the GlPI crossplied specimens failed at 

stresses within the baseline tensile data scatter band while the unidirectional specimens failed 

at stresses slightly below the baseline scatter band. 

10.3.4 BORON/ALUMINUM. Results of creep testing of BlAl at 305 K (450" F) and 361 K 

(330" F) are listed in Table 10-5. Data are presented for 24 specimens. 

10-19 
I 

I 

z 

c( 

In 

CD 

aa 

In 

0 

In 

- a 

d 

(0 

In 

v m 

CD In 

0 

0 

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ea 

5 

9 

ea 

ea 

5 

0 

v 

r( 

In 

CD 

aa 

3000 

10000 c(4dcl I I 

d 

CD 

In 

m m 

In 

v 

tl 

0 

I 

0 

10-20 
a1 

I 

I 

I! 

I 

S?EC NO. STRESS, MNld (KS) 

0 oum lOWf14S) 

AOUS~ iou(isa 

ELAVC STRAIN 

Flguro 10.22 Crmop Strain for [Ool6 G/Pl at 505 K (4500 F) 

eb 

! 

! 

G53217.137 

I I I I 

I I 

I I 

ELASTIC STRAIN 

Flguro 10-23 Creop Strain for [Oo f 45O], G/PI ai 505 K (450° F) 

10-21 

(JKIGINAL PAGE IS 

8E: BOOR QUALITY 
1.0 

i I 

I 

i I j 

aa 

aa 

a7 

0.1 

a 

3 as 

a4 

- 

v E 

0.3 

42 

01 

0 

I 

a3 

a2 

ai 

0. 

I 

I (> 

I 

I 

1 

' I !I1 I I I I! I 111 I I1 I !Ill I !, I( 

SPEC NO. STRESS. MMld (KSI) 

0 OC53 278(40) 653217-139 

Figun 10-24 Creep Strain for [OO f: 4501, G/PI at 505 K (4500 F) 

65321 7.140 

Figure 10-25 Creep Strain for [Oo -C 45O], G/PI at 505 K (450O F) 

10-22 
If 

1.1 

1.0 

a9 

a8 

a8 

e 3 a5 

a4 

a 

U 

e3 

0.2 

ai 

0 

0.1 1 10 100 i om 

TlME (hr) S?€C NO. STRESS. MWId (KSI 

0 OC52 310(45l 653217 141 

31 -I 3 a5 

z 

u 

Flguro 10-26 Croop Strain for [OO f 4501, G/Pl at 505 K (4500 F) 

1 ELPSTlCSfRAlN 

0.6 . 

u.4. 

0.3 

a2 

ai 

0 1' I;(, I 11 I Ill1 

10-23 

1 

1 0 

1 

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1 

II I I Ill/ I , I 41 
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ELASTIC STRAIN 

I 

I 

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I . 

' I I I I I Ill I II I Ill1 I ! Ill1 Ill , , ,,/I I!. 

0.6 

3 5 as 

z 

Y 

0.4 

Ll 

65321 7.143 

! 9 1 11 i 

1; I 

I ,O I 

I I 

I 

I I I 

j 

I I 

ELASTIC STRAIN 

Flguro 10-28 Crmop Strain for [Ool6 G/P! at 561 K (SOo F) 

0.6 

3 5 as 

z 

Y 

0.4 

el 

I2 

1.1 

1.0 

I I]; 

1; I 

j I ,O I 

_____ ______ 

9 ! 1 

I I 

I 

I I I 

ELASTIC STRAIN 

1" 

I I I. 

0.1 1 10 100 1000 

SPEC NO. STRESS. MNlmz (KSI) 

0 OUII 896(130) 

A OUI8 931 (135) 

0 OUC7 9651140) 

TIME Ihr) 

653217.144 

Figure 10-29 Creep Strain for [Ool6 G/PI at 561 K (5500 F) 

10-24 
Creep S'train versus time curves are presented in Figures 10-31 and 10-32 for unidirectional 

material at 505 K (450 F), in Figures 10-33 and 10-34 for crossplied material at 505 K (450 " F), 

in Figures 10-35 through 10-37 for unidirectional material at 561 K (550" F), and in Figures 

10-38 through 1040 for crossplied material at 561 K (550" F). 

Creep strains measured for BlAl specimens were somewhat higher than for other systems but 

were still quite small. The unidirectional material gave values less than 0.2% with the crossplied 

material reaching a maximum value of only 0.31%. Several failures were recorded during test 

prior to completing the required 100- or 1000-hour time periods. There were also several 

doubler bond failures in the more highly loaded unidirectional specimens. These failures, both 

in the BlAl composite material at each test temperature and in the doubler bonds at 561 K 

(550" F), occurred at stress levels above approximately 85% of the average ultimate tensile 

strength for the 505 K (450" F) tests and above approximately 75% for the 561 K (550" F) 

tests. One specimen, EC5-12, failed during loading at a stress level, 421 MN/m2 (61 hi), slightly 

below the average ultimate tensile strength, 439 MNlrn2 (63.7 hi), of crossplied BlAl at 561 K 

(560" F). The specimen had been mistakenly identified as an unidirectional specimen and was 

intended to have been tested at a much higher load. 

The conclusion drawn from creep testing of the four composite materials is that boron or 

graphite filament reinforced materials with 0 " plies aligned in the direction of loading undergo 

very little creep even at stress levels that result in fracture in short exposure times. Because of 

the low creep strains observed in these materials, it would appear that stress rupture tests 

10-25 
0 

ua 

v 5: 

ua 

4 

CD In 

w J 

w w 

c 

CD 

ua 

- rl) 

ua 

b 

tl 

0 

I 

0 

10-26 
1.0 1 1 

03 

0.2 

ai 

0 

aa 

a7 

0.6 

t 

a 

- 

as 

a4 

v 

! I I 

b A 

. 

"" I I IIIII I I1 I I Ill I I I I I Ill I I I ;,!,! 

2 as a 

v 

0.4 

0.3 

0.2 

ai 

0. 

Figurm 10-31 Creop Strain for [0°]6 B/Al at 505 K (450° F) 

I ;I I - I io i 

! 

ELASTIC STRAIN 

1 

I 

~ 

I 

I !I1 1 I , , l,,##l ' ' I, I!, ! : I ,$I,' i 

SPEC NO. STRESS. MNld (KSII 

0 EU5-9 1103(1601 

A EU5-10 1172 (170) 653217 147 

Figure 10-32 Creep Strain for pol6 B/AI at 505 K (450O F) 

10.27 
1.0 I i 

1 10 100 1 ow ... 

TIME (hrl 

ai 

SPEC NO. 

0 ECS-5 

h ECS-11 

a ECI~ 

Figurm 10-33 Creop Strain for [go C aso], 8/AI at 505 K (4500 F) 

. 1.0 

W 

0.7 

8 

; 0.5 

a 

U 

a4 

0.3 

ai 

0 

SPCC no. STRESS. MN~ (KSI) TIME lhrl 

0 ECS-I 379(55) 

A ECS-10 414(60) 

0 ECSd W(65) 

Figure 10-34 Creep Strain for [Oo 2 4S0], B/Al at 505 K (4500 F) 

10-28 
1.0 

an 

a: 

a7( 

Figuro 10-35 Creop Strain for [Ool6 B/Al at 561 K (5500 F) 

I 

I 

0 U U 1 I I[ 

0 0 

L 

1.0 

ao 

a: 

ai 

aa 

3 - 

," 0.5 

z 

ar 

a3 

az 

ai 

U 

0 

0.1 1 10 100 1000 

TIME (hr) 

SPEC NO. STRESS. MN/d (KSII 

0 EU51 862 (1251 

A EU55 8% (1301 653217 151 

~ 

z I 

0.8 

Figure 10-36 Creep Strain for [O0l6 B/Al at 561 K (550O F) 

I 

1 

10-29 

505 

a 

U 

a4 I I1 I 

0.3 

nz 

ai 

0. 

ELASnC STRAIN I I; I 

I I 

. 6 

I I I I l11l1 I I I IIII! I I I IO1 I ! ([I I [I ' 
1.0 

as 

aa 

a7 

as 

3 

5 a5 YI 

Y E 

0.4 

I . , I 

0.3. 

az 

ai 

01 

Flguro 10-37 Creop Strain for [(lo]* 8/AI at 561 K (5500 F) 

EhklC STRAIN I I1 

I I I 

- 

' ' I I I Ill11 I I I )!Ill I I Ill I I1 I I I I I I'I 

a4 

a3 

02 

ai 

6532 17.153 

r 

I I ELASTtC STRAIN 

I ! c * 

I 

I I Ill I I' ! !I I I Ill I I IIIII II 1 I (Ill! 0' 

Figure 10-38 Creep Strain for [Oo f aO], B/Al at 561 K (550O F) 

10-30 
1.0 

I 

a7 

I7 I I 0 I IlII I I I I Ill1 

a3 

I I I I1111 I Ill I II! 

ai 

0 

I 

Figurm.lO-39 Creep Stmin for [Oo f Go], B/Al at 561 K (5500 F) 

1.0 

a9 

0.1 

ai 

a8 

5 es a 

v 

0.4 

43 

0.2 

ai 

0 

ai 1 10 100 1000 

TIME lhr) 

SPEC NO. STRESS, MN/d (KSI) 

0 ECS-3 328 (47.5) 

A ECS.2 - 345 (501 

853217 155 

figure 10-40 Creep Strain for [Oo -C 45O], B/Al at 561 K (5500 F) 

10-31 
could well be substituted for creep measurements. A plot of creep stress versus time for 

various creep strains from 0.05% to rupture would likely show a family of curves that were 

very closely spaced. Since the expense in obtaining creep data is far greater than that for 

stress rupture testing, a considerable savings could be realized. 

10-32 
SECTION 11 

BASELINE FATlGUE TESTING 

Fatigue testing made up a significant portion of the SCR study with four distinct parts: 1) conventional constant amplitude tests at room and elevated temperatures with flawed and 

unflawed test specimens: 2) constant amplitude tests on material that survived 10,000 to 

36,000 hours of flight simulation exposure: 3) short-term flight simulation tests (random 

fatigue loads, constant elevated temperature) of 100 and 200 hours: and 4) long-term flight 

simulation (random fatigue loads, cyclic temperature) for up to 40,000 hours. The constant 

amplitude tests on unexposed material characterize the composites. Tests after flight simulation exposure provide data to compare the residual fatigue properties to original properties. 

The short-term flight simulation tests generate wearout analysis model data necessary to set 

the stress levels for the long term flight simulation tests. Finally, the long-term tests provide 

data on the effect of simulated supersonic flight service environments for each composite 

system. The bulk of these tests, those that senre as baseline for later comparisons, were performed on unexposed composite materials. It is these tests that are described in this section. 

In some cases, the data were available in the literature, and no additional tests were required. 

As a result of considerable early testing and extensive use, the BIE material fitted this 

category and only limited testing was necessary (ref. 40 and 41). 

. 

Test parameters that were considered in the baseline constant amplitude fatigue tests were 1) 

specimen layup,'2) specimen configuration, 3) test temperature. and 4) stress ratio. Specimens 

of both unidirectional and [0" & 45'1 crossply layups were evaluated. All test specimens were 

loaded in the direction of the 0" fibers. Flawed and unflawed specimens were included. The 

flawed specimens were fabricated with a center hole having a diameter equal to onefourth of 

the specimen width. For isotropic material a specimen with this configuration would have a 

theoretical stress concentration factor, K, of 2.43 (ref. 28). Testing was conducted at room temperature for all systems and, in addition, at 450 K (350' F) for G/E and Celion 6000ILARC-160 

GlPI and at 505 K (450 " F) for HT-S/710 G/PI and B/AL Fatigue tests were performed at stress 

ratios (R) of 0.1 and -1 for all systems except BIE and, in addition, at R = 10 for the Celion 

6000/LARC-160 GIPI. R is defined as the algebraic ratio of the minimum stress to the maximum 

stress $one cycle (thus loading for the R = 0.1 tests is tension-tension, for the R = -1 test is 

fully reversed tension-compression, and for the R = 10 tests is compression-compression). 

In addition to the baseline tests described above, three other fatigue studies are included in 

this section. The first was a limited program to investigate the effects of stacking sequence on 

fatigue strength of the [O" f 45'1 family of laminates. The second was a determination of 

Poisson's ratio of specimens after 10 million fatigue cycles (runout). The third was a determination of baseline fatigue strength of the thermal agingkompression study materials. These 

tests were done on unnotched 24 ply [0 ' -I- 45 'Is4 specimens at room temperature and a stress 

ratio, R, of -1. 

11.1 SPECIMEN DESIGN 

Specimens tested at a stress ratio, R, of 0.1 were fabricated from 6 ply [0 'I6 and [0 & 45 'Is 

laminates. For R = -1 tests, where compressive as well as tensile loads are applied, a thicker 

11-1 
specimen was used. Information obtained hm IITRI (private communication from KE. Hofer, 

1973) indicated that fatigue testing 6 ply composite material at R = -1 gave unreliable results. 

The problem was related to compressive buckling and could be overcome by the use of thicker 

material. For this reason the R = -1 specimens were cut from 12-ply [0 O]12 and [0 O h 45 "Is2 

laminntes and the R= 10 specimens were cut from 24 ply [0 O h 45 laminates. Cutting, hole 

machining, and doubler bonding techniques were those described in Section 6. Details of the 

specimen geometry for the various configurations and test conditions are listed in Table 11-1. 

The length of the R = 0.1 specimens that were tested at elevated temperature was twice that 

of the room temperature specimens to allow the ends to extend well away from the furnace during testing. With this technique doubler bond failures were eliminated. Because of the compressive loads present during the R = -1 and R= 10 testing it was not feasible to extend the 

length of these specimens for the elevated temperature conditions. Doubler bond heating of 

these specimens was minimized by reducing the size of the furnace as described in Section 

11.2. The widths of the various specimens were the same as those chosen for the baseline unnotched and notched tensile tests (Section 7), except that the R = -1, and R= 10 unnotched 

specimens were all 0.025 m (1.0 in.) in the fatigue tests rather than 0.013 m (0.5 in.). 

11.2 TEST EQUIPMENT AND TEST PROCEDURE 

Testing was performed on Baldwin-Sonntag SFlU and SFlOU fatigue machines at a constant 

frequency of 30 Hz. Both grip ends of the specimens were placed between two Eshaped steel 

fixtures that were bolted to the machine loading platens. The specimens were secured to the 

steel plates by two -013 m (0.5 in.) diameter bolts that straddle the edges of the grip ends of the 

specimen. To further increase the gripping power of the fixtures, a special abrasive cloth (Bay 

State Gritcloth, size 320 grit) was placed at each specimen-to-fixture interface. This cloth had 

two beneficial features for a secure grip: 1) the open weave design, which increases the frictional holding force, and 2) the abrasive coating on both 'surfaces of the cloth, which becomes 

embedded in the contact surfaces of both the fixtures and tabs. The two bolts were torqued to 

I level that just secured the specimen, the alignment of the specimen was carefully checked to 

ensure axial loading, and final tightening was completed. Figures 11-1 through 11-3 show one 

of the fatigue testing machines and close-up views of the gripping fixtures with both R = 0.1 

and R= -1 specimens in place. 

For elevated temperature tests of the R = 0.1 specimens a 0.15 m (6 in.) long clamshell furnace 

was used. The length of the specimens was such as to dow the ends to extend out of the furnace far enough to prevent excessive heating of the doubler adhesive. This factor and the 

excellent gripping power of the end fixtures combined to eliminate doubler bond failures at 

elevated temperature. Figure 11-4 shows the furnace arrangement with a specimen loaded in 

the fatigue machine. Temperature was controlled to 2 3 K (2 5" F) with a Minneapolis 

Honeywell Brown Electronik controller using a chromel-alumel thermocouple atached to the 

specimen at the center of the gage section. 

For the R = -1 and R = 10 specimens a much smaller furnace was required for the elevated 

temperature tests because of the limited amount of space available between the gripping fixtures. Most tests were conducted using a small ring furnace of the type shown in Figure 11-6. 

This could be used on all tests that were run in the smaller SFlU fatigue machine but not in 

the larger SFlOU because of space limitations. The unidirectional B/AI specimen loads were 

such as to require the larger SFlOU fatigue machine. For these specimens a successful technique for elevated temperature testing was developed by using a flexible heating tape, which 

11-2 
Tablo 11.1. Dotails of Fatiguo Spocimon GoomHolo Ter) 

Speeimm R Test thicknosr Lmgtl, Width Dfamotor S.ction 

Configuration Valuo Twnporutun (Plios) rn (in.) rn (in.) m (in.) rn (in.) L) 

Unnotched 0.1 

.1 

-1 

-1 

10 

10 

Notched .1 

.1 

-1 

-1 

10 

10 

R.T. 

Elevated 

R.T. 

Elevated 

R.T. 

Elevated 

R.T. 

Elevated 

R.T. 

Elevated 

R.T. 

Elevated 

6 

6 

12 

12 

24 

24 

6 

6 

12 

12 

24 

24 

0.23 9 

.46 18 

so 4 

so 4 

.lo 4 

.lo 4 

.23 9 

.46 18 

.lo 4 

.lo 4 

.lo 4 

so 4 

0.013 0.5 

.013 0.5 

-025 1 

.025 1 

.025 1 

.025 1 

a.025 1 

a.025 1 

.025 1 

,025 1 

.025 1 

.025 1 

- 

- 

- 

- 

- 

- 

b.0064 0.25 

b.0064 0.25 

.0064 0.25 

.0064 0.25 

.0064 0.25 

.0064 0.25 

0.10 4 

.33 13 

.013 0.5 

.013 0.5 

.013 0.5 

.013 0.5 

so 4 

.33 13 

.013 0.5 

.013 0.5 

.013 0.5 

.013 0.5 

afar BIAL 0.013 (0.5) 

bfor BIAL 0.00318 (0.125) 

~~~ 

was wrapped around the short test secti011 of the specimens. The heater was 0.61 m (24 in.) in 

length, 0.013 m (0.5 in.) wide, rated at 200 watts and had a maximum use temperature of 873 K 

(1112 " F). Figure 11-6 shows a specimen with the heating tape in place ready for test. 

Prior to testing, the resin matrix composite specimens were dried at 394 K (250" F) for 24 

hours. No baking of the B/Al specimens was required. The specimen dimensions were 

measured, and the static and dynamic loads calculated for the particular stress level and R 

value desired. The notched stress values were based on net section, Le., width minus hole 

diameter. After loading the specimen in the machine and carefully checking the alignment, the 

load levels were set, and the test initiated. The tests were continued to failure or until a maximum of 10 million cycles was obtained. The number of load cycles was measured with a 

counter driven by the motor rotating the dynamic load. This counter had a least count of 1000 

cycles so that any failures below this value were not registered. Nine specimens were tested for 

each combination of: 1) material system, 2) laminate orientation, 3) specimen type (unflawed or 

flawed), 4) R value, and 5) test temperature. The test data were plotted on semi-logarithmic 

paper to obtain the characteristic S-N curves. 

Initially the S-N curves were drawn to show a straight line of best fit of the data points as 

determined by linear regression analysis. The high degree of scatter in some of the data sets 

and the uncertainty of how to input the runout specimens often gave S-N curves that did not 

appear to be consistent with those from data sets with little scatter. For this reason the 

method was changed, and all curves have been drawn by hand. 

11-3 
Figure 11-1 Fatigue Testing Apparatus 

11-4 
Figure 11-2 Fatigue Specimen Mounted in Gripping Fixture for R = 0.1 Test at Room Temperature 

11.3 TEST RESULTS AND DISCUSSION 

The overall baseline fatigue testing program is summarized in Table 11-2. This table shows the 

number of specimens and stress ratios that were employed at each test temperature for the 

five material systems. Of the total number of test specimens shown in Table 11-2, all are included in the data presented in this section although a few are not valid because of equipment 

malfunctions, doubler bond failures, or incorrect test loads or temperatures. These specimens 

are identified in the tables and figures. The data are presented in both tabular form and as S-N 

curves, one for each stress ratio and temperature. 

11.3.1 BORON/EPOXY AND GRAPHITE/EPOXY. Because of the large amount of testing previously conducted by others on the B/E system, no baseline fatigue testing was required. 

11-5 
Figure 11-3 Fatigue Specimen Mounted in Gripping Fixture for R = - 1 Test at Room Temperature 

However, to determine more precisely the effects of the flight simulation exposure it was 

desirable to have baseline fatigue data from this same batch of material using the same 

specimen configuration and test technique as used during the post exposure evaluation. For 

this reason baseline fatigue data were determined for B/E for the same test conditions and 

specimen configurations as were later used in evaluating the BIE flight simulation specimens 

after 10,000 hours of exposure. These were room temperature tests on unnotched and notched 

[0 O It 45 'I9 crossply material tested at an R value of 0.1. The results are listed in Table 11-3 and 

plotted in Figure 11-7. The slopes of the S-N curves are quite different for the two specimen 

configurations. The notched material has a relatively flat curve lying near 414 MN/m2 (60 ksi). 

The S-N curve for the unnotched material is much steeper, with fatigue strengths below 105 

cycles greater than the notched specimens but with lower strength for lifetimes greater than 

105 cycles. 

For GIE a complete set of fatigue tests was performed. The results are listed in Tables 11-4 

through 11-11 and individual SN curves are plotted in Figures 11-8 through 11-15. On these 

S-N curves the individual data points have been included. In the discussion of the fatigue data, 

several of the SN curves have been combined to more clearly show the effects of the test 

parameters, i.e., stress ratio, temperature, or specimen configuration. On these curves the data 

points have been omitted for clarity. 

Several points are apparent from an examination of the G/E baseline fatigue data. The test 

parameter that exhibited the greatest influence on the fatigue strength was the stress ratio, R. 

In every case the tests conducted at R = -1 gave considerably lower fatigue properties than 
Figure 11-4 Furnace Arrangement for R = 0.1 Fatigue Test .at Elevated Temperature 

ORKTI’J QL PAGE rs 

OF POOR QUALITY 

11-7 
2 

a 

a 

c 

Q) 

U 

c 

LL 

a 

a 

L 

m 

C 

a 

? 

c 

c 

11-8 
Figure 11-6 Elevated Temperature, R = - 1 Fatigue Test Setup Showing Use of Flexible Heating Tape 

11-9 
Tablo 11-2. Summary of Futlguo Test Program 

550514 BIE [OO c 45aj 

.-\-S13501-5 GiE [O O C 45 "1 

[O "I 

HT-SI710 GlPI [Oo f 45'1 

(0 "1 

Celion 60001 [OO c 457 

BlrU [' f 457 

WRC-160 GiPI 

Unnotchd 

Notched 

unnotched 

Notched 

Notched 

unnotched 

Notched 

unnotched 

Notched 

Unnotched 

Notched 

Unnotchd 

Notched 

unnotched 

Notched 

UMOtched 

5 -- 

5 

9 -9 

9 9 

9 -9 

9 9 

9 9 

9 - 9 

9 9 

9 - 9 

9 99 

9 99 

9 9 

9 9 

9 9 

9 9 

-- 

- 

- 

- 

- 

aTest Temperature was 561 K (560') 

9- 9 

9- 9 

9- 9 

9- 9 

-e9-9 

9- 9 

9-9 

9- 9 

as 9 

as 9 

as 9 

as 9 

~~ 

Tablo 11-3. Axial Futlguo Data for [Oo k 4501, B/E, R = 0.1 

Spocimon Tomporatun Spocimm Maximum 

Numbor K (OF) TYPO MN/d (ksi) Cycl.8 Commonis 

AC72-1 

-2 

-3 

-4 

-5 

AC71-4 

-5 

-6 

-7 

-8 

297 75 Unnotched 414 

448 

345 

483 

310 

297 75 Notched 345 

414 

483 

43 1 

448 

60 

65 

50 

70 

45 

50 

60 

70 

62.5 

65 

227,000 

3 1.000 

3,534,000 

8,000 

10,116,000 Did not fail 

10,090,000 Did not fail 

1,386,000 

< 1,000 

462.000 

< 1,000 

those run at R values of 0.1. Generally, the effect of a notch (center hole) was to reduce the 

fatigue strength, and an increase in temperature from 297 K (75" F) to 450 K (350" F) also 

caused a drop in strength. Figures 11-16 and 11-17 compare the results of axial fatigue testing 

at room temperature and stress ratios of R = 0.1 and R = -1 for both unnotched and notched 

specimens. These data show the fatigue life for a stress ratio of R = -1 to be much lower than 

for R = 0.1. For both stress ratios, the fatigue life for notched specimens is less than for 

11-10 
500 

Ui 

E 

z 

zz 

W 

d 

\ 

400 u 

w 

a. 

VJ 

VJ 

w 

a 

a 

a 

z 

300 

X 

a 

a 

200 

0 UNNOTCHED 

0 NOTCHED 

- 

W 

-40 s: 

c) 

< 

c) 

Im 

30 2 

- E. 

1 I I I I Id I I I I*IIIIIA 1 I I I 11111 I I I I I Ill1 I I I I I IId*O 

1 o2 1 o3 1 o4 lo5 lo6 1 o7 

NUMBER OF CYCLES 66321 7-161 

Figure 11.7 Axial Fatigue PropodOS of [OO f fi0], WE at 297 K (75O F) and a Stress Ratio, R, of 0.1 

unnotched specimens except the unidirectional material tested at R = -1, where no notch 

effect was observed. Similar data are presented in Figures 11-18 and 11-19 for testing at 450 K 

(350 O F). Again the fatigue strength is greatly reduced for the R = -1 tests compared to those 

conducted at a stress ratio of R = 0.1. The presence of a notch is shown to reduce the fatigue 

strength for the R = 0.1 tests, but not for the R = -1 tests. Further, the unidirectional 

notched specimens performed better than those that were unnotched. The reason for this 

behavior is not clearly understood at this time, but may be related to the phenomenon of crack 

blunting brought on by the combination of high compressive loading and a matrix softened 

because of elevated temperature exposure. 

Two problems were experienced early in the GIE fatigue test program. The first was doubler 

bond failures in the room temperature, R = - 1, unnotched, crossply specimens. After the first 

two specimens failed, the remaining doublers were removed and rebonded using a modification 

of the bonding procedure and substituting a higher shear strength adhesive. These specimens 

gave satisfactory results. The second problem occurred with the room temperature, R = 0.1, 

unnotched, unidirectional specimens. Six of these nine specimens failed in the G/E but inside 

the doublers. These failures are thought to have been caused, in part, by nonuniform clamping 

pressure applied by the test machine grips because of the warped condition of the doublers. 

Phanoou in the nrihesivp hnnding nrncess were made to eliminate this Droblem the remaining 

One measure of the fatigue resistance of a material is the percent of the ultimate strength at 

which the fatigue limit or endurance limit (run-out) occurs. Fatigue limit is the limiting value of 

the stress below which a material can presumably endure an infinite number of stress cycles, 

Le., the stress at which the S-N curve becomes horizontal. For certain materials and test conditions no fatigue limit occurs and some value of N (number of cycles) is chosen. Many of the S-N 

11-11 
Tablo 11-4. Axial Fatiguo Data for [Oo k 45O], G/E. Unnotdtd, Rs0.1 

T.rt Maximum 

Sp4m.n Tompomtun Strosr 

Numbu K (OF) MN/d (kri) Cycl.8 Commonts 

BC72-20 

-21 

-22 

-23 

-24 

-25 

-26 

-27 

-28 

BC72-47 

48 

-49 

-50 

-5 1 

-52 

-53 

-54 

-55 

297 

450 

75 414 

483 

448 

43 1 

448 

43 1 

43 1 

448 

465 

350 445 

276 

414 

379 

310 

396 

362 

3 28 

379 

60 

70 

65 

62.5 

65 

62.3 

62.5 

65 

67.5 

50 

40 

60 

55 

4s 

57.5 

52.5 

47.5 

55 

11,830,000 

c 1,000 

c 1,000 

c 1,000 

340,000 

c1,ooo 

1,121,000 

1.274,OOO 

60,000 

6,181,000 

1,000 

3,000 

1,000 

12,165,000 Did not fail 

10,140,000 Did not fail 

10,165,000 Did not fail 

10,000,000 Did not fail 

10,000,000 Did not fail 

CuNes obtained during testing of the composites in this study were of this type. For these 

curves a vdue of 107 cycles was used. 

The percentages of the ultimate strengths at lo7 cycles have been calculated for each of the 

groups of G/E specimens and are listed in Table 11-12. For [0 O k 45 "1 crossply material tested 

in tension-tension fatigue, R = 0.1, the values average 86% for both room and elevated temperature and unnotched and notched configurations. Similar tests on unidirectional material 

averaged about 60% of the ultimate strength at lo7 cycles. Values for the same types of tests 

except for tension-compression fatigue, R = -1, were considerably lower, 34% for crossply 

and 28?0 for unidirectional. To show how these values compare to typical aircraft aluminum 

sheet alloys, Table 11-13 was prepared. Data were obtained (ref. 42) for 0.0023 m (0.090 in.) 

thick 2024-T3 and 7075-T6 aluminum alloy sheet at room temperature. For tension-tension 

fatigue the G/E composite is far superior to the aluminum alloys particularly for notched configurations. For tension-compression fatigue the advantage is not as great although the 

notched G/E is still considerably higher than either of the notched aluminum alloys. For either 

type of fatigue loading the GiE material is less sensitive to a notch than the aluminum alloys. 

11.3.2 GRAPHITE/POLYlMlDE 

11 3.2.1 HTW710 G/P!. A complete set of fatigue data was also obtained for the HT-S/710 G/PI 

system. For this material the elevated temperature selected was 505 K (450 O F). The results of 

the numerous tests are listed in Tables 11-14 through 11-21 and individual S-N curves are plotted in Figures 11-20 through 11-27. Again, all the data points have been included in these S-N 

plots. These curves have been combined in Figures 11-28 through 11-31 to show the effects of 

the different test parameters. 

11-12 
Tabio 11.5. Axiai Fdguo Dafa for [Oo f 45OIs2 WE, Unnotchod, R = -1 

~~ 

Tort Maximum 

Spocimm Tompoaturo Strosr 

Numbor K (Of) MN/d (kri) Cyd.. Commonts 

BC72-11 

-12 

-13 

-14 

-15 

-16 

-17 

-18 

-19 

BC72-38 

-39 

-40 

-4 1 

43 

-44 

-45 

-46 

-47 

297 75 414 

276 

345 

310 

276 

241 

241 

207 

172 

450 350 207 

138 

172 

241 

190 

207 

190 

172 

224 

60 

40 

50 

45 

40 

35 

35 

30 

25 

30 

20 

25 

35 

27.5 

30 

27.5 

25 

32.5 

e 1,000 

16,000 

1,000 

3,000 

3,000 

5,000 

10,400,000 

5.000 

10,159,000 Did not fail 

15,560,000 Did not fail 

28,000 Bond failure 

87,000 Bond failure 

3,000 

2,131,000 

38,000 

38,000 

9.000 

10,007,000 Did not fail 

In like manner to the results obtained with the G/E material, the factor having the most pronounced effect on fatigue strength was the stress ratio, R. In all cases, with the other parameters identical, specimens tested in an R value of -1 gave fatigue strengths at lo3 cycles 

(determined from S-N curve) that were less than those obtained at 10 cycles for specimens 

tested at an R value of 0.1. 'he effect of a notch was generally to reduce the fatigue strength 

but, again, there were exceptions. Notched specimen awes were higher than unnotched for 

both [0 "1 material at 297 K (75 O F) and R = 0.1, Figure 11-29, and [0 O f 45 "1 material at 505 K 

(450" F) and R = 0.1, Figure 11-30. For [O"] material at 297 K (75O F) and R = -1, Figure 

11-29, and at 505 K (450" F) and R = -1, Figure 11-31, the unnotched and notched curves 

crossed each other. The HT-S/710 GlPI fatigue data, primarily the [O"] results, showed a good 

deal more scatter than was found in the G/E data, and this may have contributed to the inconsistencies in the effects of a notch on HT-S/710 G/PI fatigue strength. The effect of increasing 

the temperature for the R = -1 tests was a slight reduction in fatigue strength. For the R = 0.1 

tests both decreases and increases in fatigue properties were observed depending on the laminate orientation and the specimen configuration. Additional fatigue testing would be useful in 

determining the effects that a notch or an increase in test temperature contribute to the fatigue 

strength of HT-S/710 GIPI. 

In testing one group of specimens (12 ply, unidirectional, notched, R = -1, room temperature, 

Table 11-21, Figure 11-27) many of the failures occurred in the composite under the bonded end 

doublers rather than in the test section through the center hole. This type of failure is not uncommon for unidirectional composites tested in fatigue. During fatigue cycling vertical cracks 

initiate at the edges of, and tangent to, the center hole, dividing the specimen into three 

pdel strips. Failure can then occur anywhere along the two outside load carrying strips and 

11-13 
Tablo 11-6. Axial Fdiguo Data for [Oo k 4!S0], G/E, Nofcbod. R = 0.1 

TOd Maximum 

Spocimon Tomporatun Stmr 

Numbor K (OF) MN/m2 (ksi) cyci~ Commortts 

BC71-10 

-11 

-12 

-13 

-14 

-15 

-16 

-17 

-18 

BC71-28 

-29 

-30 

-3 1 

-32 

-33 

-34 

-35 

-36 

297 75 276 

345 

310 

328 

345 

379 

448 

414 

396 

450 350 345 

362 

379 

310 

328 

345 

345 

345 

336 

40 

50 

45 

47.5 

50 

55 

66 

60 

57.5 

50 

52.5 

50 

45 

47.5 

50 

50 

50 

48.75 

10,468,000 

c 1.000 

12,723.000 

10,313,000 

10,116,000 

15,284.000 

c 1,000 

c 1.000 

< 1.000 

10,006,000 

c 1.000 

c 1,000 

13,368,000 

10,125,000 

C 1.800 

26.000 

6 1.000 

e 1,000 

Did not fail 

Did not fail 

Did not fail 

Did not fail 

Did not fail 

Did not fail 

Did not fail 

Did not fail 

Failed under doubler 

often is initiated at the higbly stressed regions at the end doublers. The same failure mechanism was prevalent in the unidirectional BlAI specimens tested under the same conditions. 

A table similar to the one presented earlier for the GIE data has been prepared to show the 

percentage of the HT-Si710 GlPI ultimate strengths at which the fatigue limit (or 107 cycles) 

occurred. The results, Table 11-22, were quite similar to the GIE system except for being 

slightly lower in value. Tests at an R value of 0.1 give approximately 75% for the crossply 

material and 60% for the unidirectional material. Again, for stress ratios of -1 the values were 

considerably lower, 2'7% and 24% respectively for the crossply and unidirectional layups. In a 

comparison with similar data for typical aircraft aluminum sheet alloys (Table 11-13) the 

HT-S/710 G/Pl is superior, particularly for the notched materid While the advantage is not as 

great as that shown by the GIE system, GlPI retains these excellent fatigue properties up to 

505 K (450" F), well above the maximum use temperatures of either the epoxy or aluminum 

systems. To show the same comparison with a structural material having a higher maximum 

use temperature than the aluminum alloys, Table 11-23 was prepared. Data were obtained (ref. 

42) for 0.0016 m (0.063 in.) thick Ti-6A1-4V titanium alloy sheet at 297 K (75" F) and 589 K 

(600 O F). The GlPI composite is as good or better than the titanium alloy up to 505 K (450 F) 

for all comparisons except the unnotched unidirectional material tested in tension-compression 

fatigue. Again, data scatter in the [0"] results may be a factor. In like manner to the G;E - 

aluminum alloy comparison the advantage held by the HT-Sl'ilO G/PI over the titanium alloy 

is particularly evident in the notched data. 

11.3.2.2 Celion 6000/LARC-160 G/PI. As discussed previously only [0 " k 45 "1 crossplied specimens of the Celion 60001LARC-160 system were evaluated in the fatigue study. However, in 

addition to the R = 0.1 and -1 tests a series of R = 10 tests was included. The results of the 

11-14 
tablo 11.7. Axial Fotiguo Data for [OO f 45.1, G/E, Notehod. R = - 1 

TOSt Maximum 

Spoeimon tomporaturo Stmr 

Numbor K (OF) MN/d (kri) Cycla Commontr 

BC71-1 

-2 

-3 

-4 

-5 

-6 

-7 

-8 

-9 

BC71-19 

-20 

-21 

-22 

-23 

-24 

-25 

-26 

-27 

297 75 276 

207 

138 

69 

172 

241 

207 

103 

121 

450 350 207 

172 

172 

190 

155 

172 

190 

207 

155 

40 

30 

20 

10 

25 

35 

30 

15 

17.5 

30 

25 

25 

27.5 

22.5 

25 

27.5 

30 

22.5 

17,000 

164,000 

282.000 

127;OOO 

24.000 

4,000 

15,418,000 Did not fail 

17,308,000 Did not fail 

10,500,000 Did not fail 

17.000 

665,000 

434,000 

1,689,000 

670,000 

89,000 

1,000 Machine malfunction 

12,160,000 Did not fail 

12,470,000 Did not fail 

tests are listed in Tables 11-24 through 11-29 and S-N awes are plotted in Figures 11-32 

through 11-37. Additional S-N curves have been prepared to show the effects of the various test 

parameters. These are shown in Figures 11-38 and 11-39. The elevated temperature selected 

was 450 K (350°F), the same as that used for the flight simulation exposure testing. 

The effect of stress ratio, R, is clearly shown in FiMs 11-38 and 11-39. The application of 

compression loading during the fatigue cycle significantly lowered the fatigue properties of the 

Celion 60001LARC-160 G/PI. Completely reversed, tension-compression, R = - 1, loading gave 

the lowest fatigue lives for each combination of snmhnen cnnfieuration and test temperature. 

Fatigue strengths determined from the compression-compression tests, R = 10, were some 

what higher than those for R = -1 but well below those for the R = 0.1 loading. The effect of a 

notch was, without exception, to lower the fatigue strength of this system. The effect was 

moderate, ranging from 4 to 11% at room temperature and 1 to 19% at elevated temperature. 

increasing the test temperature to 450 K (350 OF) generally lowered the fatigue strength. These 

decreases were also moderate ranging from 1 to 18% at 107 cycles. Some crossover of the S-N 

cuwes for the two temperature was observed. This occurred for the R = 0.1 and -1 tests in 

the lower cycle regions of the S-N curves. 

Some testing difficulties were experienced with two series of fatigue tests. The first was for the 

6 ply, unnotched, R = 0.1, room temperature group of specimens. All failures except one 

occurred at the composite in the region under the bonded end doublers. Failures during tensile 

testing are frequently found at this location because of high gripping stresses. It is thought 

that the end failures did not sigmficantly alter the S-N curve for this group of specimens. The 

11-15 
Tablo 11-8. Axial Futiguo Data for [OO], G/E, Unnotchod, R = 0.1 

Test Maximum 

Sp-imon Tompomtun Stross 

Numbor K (OF) MN/d (kri) , Cyclor Commomts 

B U7 2- 20 

-21 

-22 

-23 

-24 

-25 

-26 

-27 

-28 

BU72-47 

-48 

-49 

-50 

-5 1 

-52 

-53 

-54 

-55 

297 75 1.030 

965 

965 

896 

93 1 

931 

965 

1,000 

1,070 

450 350 965 

896 

1,030 

1,000 

979 

948 

983 

1,000 

965 

150 

140 

140 

130 

135 

135 

140 

145 

155 

140 

130 

150 

145 

142' 

137.5 

142.5 

145 

140 

42,000 

17,000 

8,000 

14,620,000 

70,000 

10.184.000 

17.000 

47,000 

19,000 

1,073,000 

10,156,000 

2,000 

5,000 

316,000 

9,482.000 

103,000 

1.000 

8,548,000 

Failed under doubler 

Failed under doubler 

Did not fail 

Failed under doubler 

Did not fail 

Failed under doubler 

Failed under doubler 

Failed under doubler 

Did not fail 

Did not fail 

second problem area was for the 24 ply, unnotched, R = 10 tests at room temperature. Four of 

these specimens failed at the doubler bonds rather than in the composite. These bond failures 

were attributed to the much higher loading required because of the greater cross sectional area 

of the 24 ply laminates. Again, the S-N curve, Figure 11-34, is thought to be representative of 

this group of specimens based on the samples that failed in a conventional manner. 

The fatigue data for Celion 6000/LARC-160 G/PI have been summarized in Table 11-30 by 

comparing the runout strength (107 cycles) with the ultimate tensile strength for the various 

test series. The results were very similar to those for HT-S/710 G/PI for the R = 0.1 tests and 

slightly better for the R = 1 tests. The same advantage over aluminum and titanium alloys 

held by the HT-SI710 material is also shown by Celion 60001LARC-160 G/PI except for the 

somewhat lower maximum use temperature, 450 K (350 OF) versus 505 K (450 OF). 

11 3.3 BORON/ALUMINUM. Results of the BlAl fatigue testing program are listed in Tables 

11-31 through 11-38. These data have been plotted in S-N diagrams in Figures 11-40 through 

11-47. The fatigue curves have also been combined in Figures 11-48 through 11-51 to show the 

effects of the different test parameters. The elevated temperature initially selected for testing 

was 561 K (560 F), but this was later reduced to 505 K (450 ' F) because of severe surface degradation of the aluminum. Consequently, half of the elevated temperature tests (those 

at R = 0.1) were conducted at 561 K (550" F) and the other half (R = -1) at 505 K (450' F). 

Figures 11-48 through 11-51 again show the pronounced effect that the stress ratio, R, has on 

the fatigue strength. For all combinations of orientation, temperature, and specimen configuration that were evaluated, the R = -1 tests gave values well below those of R = 0.1. The 

' 

11-16 
Tablo 11-9. Axial fdguo Data for [0°],2 G/E, Unnotehod, R = - 1 

Tost Maximum 

Spocimon Tompomtun Strur 

Numbor K (OF) MN/d (ksi) Cycl.. Commonh 

BU72-11 

-12 

-13 

-14 

-15 

-16 

-17 

-18 

-19 

BU72-38 

-39 

-40 

43 

-44 

-45 

46 

-1 

-2 

-3 

297 75 620 

483 

517 

448 

414 

448 

500 

500 

534 

450 350 552 

379 

414 

483 

448 

345 

310 

276 

241 

207 

90 

70 

75 

65 

60 

65 

72.5 

72.5 

77.5 

80 

55 

60 

70 

65 

50 

45 

40 

35 

30 

c 1,000 

12,000 

3,517,000 

10,003,000 Did not fail 

12,592,000 Did not fail 

10,202,000 Did not fail 

3,000 

19,000 

5,000 

< 1.000 

75,000 

37,000 

3,000 

31,000 

307,000 

56,000 

3,215,000 

5,269,000 

10,904,000 Did not fail 

effect of the center hole in the notched specimens, generally, was a decrease in fatigue 

strength. For some groups of specimens the effect was slight and for one, the unidirectional 

material tested at 297 K (75 O F) and R = -1, the notched fatigue strength was well above the 

unnotched strength at the lower end of the SN me. The effect of raising the test 

temperature from 297 K (75O F) to 561 K (550' F) was to significantly reduce the fatigue 

strength at lo7 cycles. For fatigue lives of lo5 or less and at an R value of 0.1, the effect of 

temperature on fatigue life was not very great. The specimens. however, exhibited extensive 

surface attack in the outer aluminum layer. Because of this poor performance the R = - 1 test 

temperature was lowered to 505 K (450 O F). As shown in Figures 11-50 and 11-51 the slopes of 

the S-N mes for these tests were similar to those obtained at room temperature, and the 

fatigue strengths were only slightly lowered. 

Examination of the B/Al specimens that were fatigue tested at 561 K (550 O F) revealed severe 

degradation of the al& matrix material, as shown in Figures 11-52 and 11-53. The surface of the specimens showed multiple cracks and roughness caused by localized deformation. 

The surface appearance suggests that the degradation may have been caused by grain growth 

and embrittlement at the grain boundaries. A temperature gradient analysis of the hot zone of 

the furnace together with an examination of the degraded specimens indicated a minimum 

threshold temperature of 505 K (450" F) to 533 K (500" F) to cause matrix degradation. To 

better substantiate this, two of the specimens originally scheduled for 561 K (550' F) were 

fatigue tested at 505 K (450" F) under the same conditions (stress level and environment) as 

used on the 561 K (560 " F) tests. Test data is reported in Table 11-36. Visual and metallographic 

11-17 
Tablo 11-10. Axial Fdguo Data for [VI, G/E, Notehod, R = 0.1 

Test Maximum 

Sprcimon Tomporahrro Strosr 

Numbmr K (OF) MN/d (kri) cychs Commonts 

BU71-10 

-11 

-12 

-13 

-14 

-15 

-16 

-17 

-18 

BU71-28 

-29 

-30 

-3 1 

-32 

-33 

-34 

-35 

-36 

297 73 896 

862 

93 1 

965 

879 

914 

862 

845 

8 27 

450 350 965 

896 

827 

8 27 

931 

862 

862 

879 

914 

130 

125 

135 

140 

127.5 

132.5 

125 

122.5 

120 

140 

130 

120 

120 

135 

125 

125 

127.5 

132.5 

307,000 

1.688.000 

36,000 

183,000 

401,000 

26,000 

625,000 Failed under doubler 

1.116,OOO 

12,630,000 Did not fail 

1.000 

425,000 

55,000 

1.000 

1.000 

1,071,000 

15,415,000 Did not fail 

12.1 43,000 Did not fail 

11,050,000 Did not fail 

examinations of the specimens revealed little or no matrix degradation at the lower temperature. The effect of environment was also briefly examined. Two fatigue tests were performed at 

561 K (550" F) in an inert atmosphere (nitrogen.gas for the first test and argon gas for the 

second one) to determine the effect, if any, on matrix degradation. The tests were mn at the 

same temperature, stress levels, and for approximately the same number of cycles as previous 

tests performed in air (see Table 11-34). Visual and metallographic examinations indicated a 

significant decrease in matrix degradation. This implies that oxygen and/or moisture play an 

important role in the degradation process. One other test variable was stress leveL For tests in 

air at 561 K (550" F) the effect was observed at all cyclic stress levels evaluated, although the 

number of cycles for degradation was substantially greater at the low stress levels than for the 

higher stress levels. A minimum threshold stress level was not determined. However, one is 

believed to be required since notched fatigue specimens (center hole notch) did not show matrix 

degradation in the unstressed areas (see Figure 11-52). Additional studies of the surface 

* degradation problem, particulariy effects of environment, should be conducted to identify the 

degradation mechanism and to develop solutions,' e.g., protective coatings or change of matrix 

alloys. 

A table similar to those presented earlier for both G/E and GlPI data has been prepared to 

show the percentage of the B/AI ultimate strengths at which the fatigue limit (or lo7 cycles) 

occurred. The results for the eight BlAl fatigue curves are presented in Table 11-39. In comparing these data with those for the organic matrix materials, several differences are apparent. 

For tension-compression loading the BlAl system exhibits relatively higher fatigue limits than 

the organic matrix systems at both room and elevated temperatures. For tension-tension 

11-18 
Tabio 11-11. Axial Fdguo Doto for [0°],2 G/E, Notehod. R = - 1 

Tost Maximum 

Spoecimm T~rnpomtun Str.rr 

Numbor K (OF) MN/d (ksi) Cyc1.r Commonts 

BU71-1 

-2 

-3 

-4 

-5 

-6 

-7 

-8 

-9 

BU71-19 

-20 

-21 

-22 

-23 

-24 

-25 

-26 

-27 

297 75 8 27 

552 

483 

517 

500 

500 

517 

483 

483 

450 350 552 

483 

414 

448 

448 

517 

465 

43 1 

534 

120 

80 

70 

75 

72.5 

72.5 

75 

70 

70 

80 

70 

60 

65 

65 

75 

67.5 

62.5 

77.5 

< 1,000 

3,000 

25,000 

17,000 

35,000 

14,000 

92,000 

2,382.000 

2,000 

58,000 

9,770,000 

2,000 

5,020,000 

17,000 

448.000 

12.000 

10,000,000 Did not fail 

10,155,000 Did not fail 

loading at room temperature the unidirectional BlAl data are comparable with those for GIE 

and GIPI while the crossply data are similar to the GlPI but lower than the G/E data. At 561 K 

(550 F) the BlAl was inferior to either the GIE at 450 K (350 O F) or the GlPI at 505 K (450 O F). 

In a comparison with similar data for the aluminum alloys listed in Table 11-13 the BlAl is 

superior. This is particularly apparent for the notched specimens. In like manner to the G/PI 

system, BIA1 retains the excellent fatigue properties to 505 K (450' F), well above that for 

aluminum alloys. To show the same comparison with the Ti-6Al-4V alloy as was presented 

earlier for GIPI, Table 11-40 was prepared. The titanium alloy sheet data are the same as those 

listed in Table 11-23. With the exception of the unnotched, R = 0.1 results at 561 K (550 O F) 

(surface degradation temperature) the BlAl is superior to the titanium alloy. This is particularly 

demonstrated by the notched fatigue data. 

1 1.3.4 STACKING SEQUENCE EVALUATION. A limited program was conducted to investigate the 

effects of stacking sequence on the [0 O f 45 family of laminates. Two systems, BE and GIE, 

were evaluated. Room temperature tensile and fatigue tests (R = 0.1 were made on specimens 

of the following three laminate stacking sequences: 

[OO, 45 O, -45 O, -45 O, 45 O, 0 "1 

[45 O, 0 O, -45 O, -45 O, 0 O, 45 "1 

[45 3 -45 O, 0 O, 0 O, -45 O, 45 "1 

11-19 
. 

500 

2 TESTS 

mi - 

g 400 

c 

-E 

- 450K (350°F) w 

4 4 u C 

e 

w 

300 v) 

w a 

5 

t; 

a 

a 

0 297K (75'F) 

450K (350'F) 

200 

i I 1 I11111 I I I I I Ill 8 

0 a 

---- 

40 r : 

m 

r 

- 'A. 

- 30 

I I 1 I I111 I 1 I I I Ill I Ill ' I11,20 

65321 7-1 62A 

500 

ru- E 

\ 

- z 400 w 

-1 

u 

-3 

P= 

w 

Figuro 11-8 Axial Fatiguo Proporties of Unnotched [Oo k 45O], G/E at a Strosr Ratio, R, of 0.1 

- 70 

0 FAILED IN TEST SECTION 

0 DOUBLER BOND FAILURE 

9 

0 - 60 

L 

3 

0 x, 

-50 

E 

NUMBER OF CYCLES .- 

65321 7- 163 

Figuro 11-9 Axial Fatiguo Propertior of Unnotched [Oo f 45O],2 G/E at a Strort Ratio: R, of - 1 

11-20 
v) 

z 

: 200 

z 

z 

- 

Figuro 11-10 Axial Fotiguo Proportier of Notched [Oo k Go], G/E at a Stress Ratio, R, of 0.1 

F 

m 

R 

1 

5. 

- 30 

. I I I I1 Ill 1 I I I1 Ill I 1 I 1 I111 I I I 1 Illl I I I 111~20 

400 

z 

a 

w 

0 

v) 

VJ 

u 5 200 

VJ 

zz 

a 

5 

x 

U 

z 100 

1 

297 K (75' F) 

450 K (350' F) 

1 I I I I Ill I I I I1 Ill 

! 1 o3 

"V 

0 297 K (75' FI 

0 450 K (350' F) E 

- 50 ' E 

0 l 

I I 

I 1 1 1 I Illl I 1 1 1 Illll I 1 I llll~o 

104 1 o5 lo6 10' 

NUMEER OF CYCLES 653217-165 

Figure 11-11 Axial Fatigue Properties of Notched [Oo -C aS0],2 G/E at a Stress Ratio, R, of -1 

11-21 
1100 

FiE 

z = 

w 

d 

\ 

- 

:loo0 

u 

w 

P 

v) 

v) 

w 

a 

a 

a 

a 

X 

a 

z 

900 

- 

800 

1 

0 450 K (350" F) 

w 

d 

u 

>. 

u 

d 

SF 

- 2. 

- 120 

a 

w 

p. 

I I I I I Ill I I I I I Ill 

v) 

v) 

w 

a 

I I I I1111 I I 1-1 I Ill I I I I I Ill 

a 

5 

600 

- 

-\ I 

80 

I 2 PO I NTS -1 

500 

400 

90 

&a: 

0 g z 

rm 

65321 7- 166 

300 

200 

Figurm 11-12 Axial Fatiguo Propmrtior of Unnotchod [0°]6 G/E at a Stress Ratio, R, of 0.1 

G> 

v) & W 

m 

I- 

' 0 m 

0 297 K (75' F) 

0 450 K (350'F) 

I I I I1111 I I I11111 I I I I IIII I I I11111 

Figure 11-13 Axial Fatigue Properties of Unnotched [0°]12 G/E at a Stress Ratio, R, of - 1 

11-22 
- 1000 K - 

< 

z 

s 

w 

4 

u 

- 

z 900 e 

w 

0 

v) 

v) 

w 

a 

k 2 800 

3 

a 

x 

z 

700 

lo2 lo4 io7 

NUMBER 3F CYCtES 

65321 7. I 68 

Figure 11-14 Axial Fatigue Proporties of Notehod [0°]6 G/E at a Stress Ratio, R, of 0.1 

700 

ri5 

2 

s 

e 600 

e 

- 

- 

w 

> 

u 

w 

0 

v) 

v) 

w 

a 

g 500 

H 

a 

2 

X 

U 

z 

Jon 

1 ! 1 o4 lo5 

NUMBER OF CYCLES I 

65321 7-169 

Figure 11-15 Axial Fatigue Properties of Notched [0°]12 G/E at a Stress Ratio, R, of -1 

1 1-23 
500 

ry= 400' -E 

z 

- H 

w 

i 

u 

> 

u 

," 300 

k 

a 200 

Q 

v) 

m 

w 

a 

5 

- 

X 

U 

a 

100 

Figun 11-16 Axial Fatiguo Proporties of [Oo f 45O] G/E at 297 K (75O F) 

r > 

- 70 

# 

R=ai - 60 

L 

z 

> 

I 

-- - - -- - - - -- - - - --- - - - - x 

- 50 r v) + 

v) 

v) 

Rs-1 7) 

-40 w \ m \ 3EI 

c) 

-c 

c) 

.\ 

\ 

\ - 20 

I I I I I Ill I I I I1111 1 1 11 IIII I I111111 I 1 1 1 I Ill 10 

-\ 

U N N OTCH ED 

- -- NOTCHED 

II II I 

NUMBER OF CYCLES 

65321 7-1 71 

Figure 11-17 Axial Fatiguo Properties of [OO] G/E at 297 K (75O F) 

11-24 
400 I I 

fI I I 

w 

u 

> 

u 

w 

0 

v) 

v) 

w 

A 300 

a 

E 200 v) 

A=-1 - 

I 

UNNOTCHED 

U. 

z 

1 I 1 I 1 I Ill1 I 1 I 111111 I I I1 Ill 

1 o2 lo3 1 o4 

m 

v) 

v) 

1' 20 g 

j10 

1 o6 lo7 

NUMBER OF CYCLES 653217-172 

Figuro 11-18 Axial Fatiguo Propor)ios of [Oo C 4501 G/E at 450 K (3500 F) 

1000 

140 

900 

ru- E 

\ 120 5 aoo - 

w 

A 

0 

U 

w 

v) 

VJ 

w 

100 700 

a 600 

a A=-1 - ao 

500 - 

{ 60 

-- -- 

5 ---- -- 

a 400 

z 

z UNNOTCHED 

- - -- NOTCHED 300 

i 1 I I I I111 I I I I1111 I I I I I Ill 1 I 1 I1 Ill 

1 o2 1 o3 1 o4 lo5 lo6 

NUMBER OF CYCLES 

Figure 11-19 Axial Fatiguo Properties of [OO] G/E at 450 K (350O F) 

1 1-25 

65321 7- 173 
Tablo 11-12. Strongth at Fdguo Limit or lo7 Cycles Expmsod as Porrmtago of Ultfmaia Tonsilo 

Stnngth for G/E 

Sp.cirn.cl Tornpuatum R Mguo strongill at lor Cydes 

Orientdon Configuration K (OF) Valuo (%) 

[O" 2 4501 

[O "I 

unnotched 

Notched 

Unnotched 

Notched 

unnotched 

Notched 

Unnotched 

Notched 

Unnotched 

Notched 

Unnotched 

Notched 

unnotched 

Notched 

Unnotched 

Notched 

297 

297 

450 

450 

297 

297 

450 

450 

297 

297 

450 

450 

297 

297 

450 

450 

75 

75 

350 

350 

75 

75 

350 

350 

75 

75 

350 

350 

75 

75 

350 

350 

0.1 

.1 

.1 

.1 

- 1.0 

-1.0 

-1.0 

- 1.0 

.I 

.1 

.1 

.1 

-1.0 

- 1.0 

- 1.0 

-1.0 

90 

92 

77 

83 

35 

29 

34 

37 

60 

60 

60 

64 

31 

35 

13 

32 

Tablo 11-13. Comparison of Strongth at Fdguo Limit or lo7 Cydos Exprosrod as Percontag. of Ultimate 

Tonsilo Strength for G/E and Two Aluminum Alloys Tostod at Room Temperoturo 

Fatiguo Strongth at lo7 Cyclor (% of F,) 

Spocimon R Graphito/Epoxy Aluminum Alloy 

Configuration Valuo [OO f as.] [OO] 2024-M 70750T6 . 

unnotched 0.1 90 60 49 47 

aNotched .1 92 60 27 24 

Unnotched -1.0 35 31 29 24 

aNotched -1.0 29 35 16 13 

aGiE, center notch theoretical Kt = 2.43 

Al Alloys, edge notch K, = 2 
tablo 11.14. Axial Futiguo Data for [Oo f 45O], Hl-S/710 G/PI, Unnotchod, R = 0.1 

TOSt 

Spocimon Tempomturo Maximum Stress 

Numbor K (OF) MN/d (ksi) Cycl.r Commonts 

DC72-20 

-21 

-22 

-23 

-24 

-25 

-26 

-27 

-28 

DC72-47 

-48 

-49 

-50 

-5 1 

-5 2 

-53 

-54 

-55 

297 75 414 

345 

276 

310 

293 

310 

328 

328 

328 

505 450 345 

276 

207 

310 

241 

259 

276 

276 

27 6 

60 

50 

40 

45 

42.5 

45 

47.5 

47.5 

47.5 

50 

40 

30 

45 

35 

37.5 

40 

40 

40 

< 1,000 

1,000 

5,622.000 

1,000 

12,045,000 Did not fail 

10,184,000 Did not fail 

3,000 

14,000 

3.000 

1,000 

18,000 

10,008,000 Did not fail 

4,000 

14,974,000 

10,230,000 Did not fail 

6,764,000 

30,000 

6,000 Failed under doubler 

tablo 11-15. Axial Fatiguo Data for (00 * 45O],2 HT-S/710 G/PI. Unnotchod. R = - 1 

' tost 

Spocimm lwnporutun Maximum Stnrr 

Numbor K (OF) MN/d (ksi) Cycbr Comments 

DC72-11 

-12 

-13 

-14 

-15 

-16 

-17 

-18 

-19A 

DC72-38 

-39 

-40 

-4 1 

-42 

-43 

44 

-45 

-4 6 

297 75 172 

190 

121 

155 

190 

172 

156 

207 

138 

505 450 138 

103 

69 

172 

121 

155 

207 

241 

190 

25 

27.5 

17.5 

22.5 

27.5 

25 

22.5 

30 

20 

20 

15 

10 

25 

17.5 

22.5 

30 

35 

27.5 

169,000 

77,000 

10,150,000 Did not fail 

15,394.000 Did not fail 

20,000 

56,000 

4,647,000 

1,000 

1,036,000 

2,908,000 

245,000 

58.000 

1,419.000 

5,000 

< 1.000 

3,000 

10.004,OOO Did not fail 

10,200,000 Did not fail 

11-27 
tablo 11-16. Axial Fatiguo Data for [OO f 45O], HT-S/7lO G/PI, Notehod. R = 0.1 

TOSi 

Spocimom Tompwatun Maximum Strorr 

Numbor K (OF) MN/d (ksi) Cycior Comments 

DC71.10 297 

-11 

-12 

-13 

-14 

-15 

-16 

-17 

-18 

DC71-28 505 

-29 

-30 

-3 1 

-32 

-33 

-34 

-35 

-36 

75 310 

276 

293 

310 

302 

297 

293 

276 

297 

450 276 

310 

345 

379 

362 

345 

345 

328 

310 

45 

40 

42.5 

45 

43.75 

43.125 

42.5 

40 

43.125 

40 

45 

50 

66 

52.5 

50 

50 

47.5 

45 

<l,OOO 

10,200,000 

10,141,000 

< 1,000 

3.000 

2.000 

10,175,000 

10,288.000 

15,370,000 

15,375,000 

3,068,000 

14,604,000 

< 1,000 

< 1,000 

1,000 

117,000 

6,000 

< 1.000 

Did not fail 

Did not fail 

Did not fail 

Did not fail 

Did not fail 

Did not fail 

Failed near doubler 

Did not fail 

Tablo 11-17. Axial Faffguo Data for EOo f 45O],2 HTW710 G/Pl. Notchod, R = -1 

Tost 

Spocimon tomporatun Maximum Sfrorr 

Numbor K (OF) MN/~ (ksi) Cyclor Commonts 

DC71-1 297 75 

-2 

-3 

-4 

-5 

-6 

-7 

-8 

-9 

-20 

-21 

-22 

-23 

-24 

-25 

-26 

-27 

DC71-19 505 450 

41 6 

207 30 

138 20 

121 17.5 

103 15 

138 20 

172 25 

155 22.3 

86 12.5 

103 15 

121 17.5 

86 12.5 

138 20 

155 22.5 

172 25 

129 18.75 

112 16.25 

95 13.75 

10,256,000 Did not fad 

< 1,000 

462,000 

330,000 

3,421,000 

114,000 

1.000 

47,000 

2,591,000 

1,137,000 

27.000 

11,000 

1.000 

60,000 

2,018,000 

2,569,000 

12,532,000 Did not fail 

10,220,000 Did not fail 

1 1-28 
Tablo 11-18. Axial Fdguo Data for [00]6 HT-S/710 G/PI, Unnotchod, R = 0.1 

T& 

Spocimon Tempomtun Maximum Stross 

Number K (of) MN/& (ksi) Cycles Cornmonts 

DU72-20 

-21 

-22 

-23 

-24 

-25 

-26 

-27 

-28 

DU72-47 

-48 

49 

-50 

-5 1 

-52 

-63 

-54 

-55 

297 75 758 

690 

620 

656 

655 

672 

690 

672 

'655 

505 450 690 

758 

724 

724 

741 

724 

776 

810 

845 

110 

100 

90 

95 

95 

97.5 

100 

97.5 

95 

100 

110 

105 

105 

107.5 

105 

112.5 

117.5 

122.5 

3,000 

3,000 

17,648,000 Did not fail 

9,000 Bond failure 

10,290,000 Did not fail 

7,000 

2,000 Bond failure 

2,000 

52,000 

12,000 

4,000 

17,256,000 Did not fail 

100,000 

12.0 15,000 

3,480,000 

70,000 

10,180,000 Did not fail 

< 1,000 

Tablo 11-19. Axial Fafiguo Data for [00]12 HT-S/710 G/PI, Unnotchod, R = -1 

Td 

Spocimon Tomporaturm Maximum Stress 

Numbor K (OF) MN/& (ksi) Cydes Comments 

DU72-11 

-12 

-13 

-14 

-15 

-16 

-17 

-18 

-19A 

-46 

DU72-38 

-39 

-40 

-4 1 

-42 

-43 

-44 

-45 

297 75 345 

241 

276 

259 

207 

483 

448 

379 

310 

138 

505 450 276 

310 

345 

414 

379 

276 

207 

293 

50 

35 

40 

37.5 

30 

70 

65 

55 

45 

20 

40 

45 

50 

60 

56 

40 

30 

42.5 

19.000 

8,404,000 Failed under doubler 

39.000 Failed under doubler 

70,000 

1,000 

22.000 

39,000 Failed under doubler 

10,584,000 Did not fail 

< 1,000 

10,257,000 Did not fail 

1.1 54,000 

6,452.000 

155,000 

7,000 

3.000 

6.000 

< 1.000 

12,232,000 Did not fail 

11-29 
Tablo 11-20. Axial Fatlguo Data for [00]1 HTW710 G/PI, Notehod, R = 0.1 

Tost 

Spoeimm Tempomturm Maximum Strorr 

Numbor K (OF) MN/& (kti) Cyclor Commonts 

DU71-10 297 75 

-11 

-12 

-13 

-14 

-15 

-16 

-17 

-18 

DU71-28 505 450 

-29 

-30 

-3 1 

-32 

-33 

-34 

-35 

-36 

653 

724 

793 

862 

896 

879 

827 

810 

845 

758 

620 

552 

690 

586 

586 

569 

655 

586 

95 

105 

115 

125 

127.5 

120 

117.3 

122.5 

110 

90 

80 

100 

85 

85 

82.5 

95 

85 

130 

10.1 60,000 

20,000,000 

10,175,000 

685,000 

4,000 

1,000 

3,000 

634,000 

421,000 

2,000 

58,000 

10,111.000 

6,000 

9.000 

4,000 

3,640,000 

5,885,000 

10,,000,000 

Did not fail 

Did not fail 

Did not fail 

Did not fail 

Did not fail 

Fabrication of the test panels and preparation of the tensile and fatigue specimens were accomplished using the same procedures described in Section 5. Tensile and fatigue tests were conducted using the techniques also described earlier. 

Results of the tensile tests of the three stacking sequences for each material are given in Table 

11-41. The data show that the position of the 0" layers in the layup appears to influence the 

tensile strength of both systems. The strength is lowest with the 0 O layers on the outside and 

increases as these layers move toward the center. The effect is small for BIE and my not be 

signrficant because of the very limited test data. Comparison of these data with the baseline 

data in Section 7 shows some discrepancies. The baseline BE' strengths of 619 MNlm2 (89.7 

ksi) for the Phase I material and 531 MN/m2 (77.0 hi) for the Phase I1 material fall above and 

below the corresponding value of 550 MNlm2 (79.8 hi) obtained for the [0 O, 45 O, -45 OIs stacking sequence paneL While these differences are quite large for data scatter no other explanation could be found. For the G/E IOo, 45 O, -45 'Is stacking sequence panel the two tensile 

strength.values fell within the wide scatter band of the baseline tensile test data for the Phase 

I material The tensile modules values obtained from the stacking sequence panels showed considerable variation among the three layups of the G/E material and for both materials even 

more variation from the baseline values of Table 7-3 and 7-5. The baseline values were obtained 

from strain gage measurements, which have generally been more reliable than the extensometer method used on the stacking sequence specimens. The large variation of modulus with 

stacking sequence for G/E would not be expected and is thought to be the result of extensome ter errors. 

, 

11-30 
Tablo 11-21. Axial Fotfguo Data for [00]12 HT-S/7lO G/PI, Notehod. R = -1 

Tost 

Spocimon Tempomturo Maximum Strorr 

Numbor K (OF) MN/& (ksi) Cydor Commmnts 

DU71-1 297 75 

-2 

-3 . 

4 

-5 

-6 

-7 

-8 

-9 

-20 

-21 

-22 

-23 

-24 

-25 

-26 

-27 

DU71-19 305 450 

83 

110 

414 

345 

310 

328 

379 

345 

414 

483 

345 

414 

276 

310 

241 

207 

310 

345 

a12 

al 6 

60 

50 

45 

47.5 

55 

50 

60 

70 

50 

60 

40 

45 

35 

30 

45 

50 

10,268,000 

15,377,000 

2,000 

12,000 

1,443,000 

3,000 

25,000 

12,000 

1.068,OOO 

< 1,000 

7,000 

2,000 

1,893,000 

6,000 

3,867,000 

10,149000 

10,021,000 

10,173,000 

Did not fail 

Did not fail 

Failed under doubler 

Failed under doubler 

Failed under doubler 

Failed under doubler 

Failed near doubler 

Failed under doubler 

Did not fail 

Did not fail 

Did not fail 

aLow stress levels were the result of an error in calculating the fatigue machine loads. 

4Si 

mE 

z 

w 300 

- 

z 

- 

-1 

u 

z 

E 200 

3 

a 

w 

v) 

v) 

c 

v) 

I = 

X 

2 100 

lo2 1 o4 1 o5 1 o6 10' 

653217-1 74A 

NUMBER OF CYCLES 

Figure 11-20 Axial Fatiguo Properties of Unnotched [Oo f 45O], HTW71O G/PI at a Stress Ratio, R. of 0.1 

11-31 
3 

3i 

X 

U 

a 

100 

- 

NUMBER OF CYCtES 

0 a FAILED NEAR DOUBLER -20 5 505 (450'F) 

Figure 11-21 

6532 17-330 

Axial Fatigue Pt0poC)ios of Unnotchod [Oo 2 aS0],2 HT-W710 G/PI at a Stress 

Ratio, R. of -1 

Figure 11-22 Axial Fatigue Propottior of Notched [OO f 4501, HT-S/710 G/PI at a Stress 

Ratio, R, of 0.1 

11-32 
300 

mE 

2 

a 

w 

d 

0 

e 

w 

0 

v) 

en 

w 

\ 

- 

g 200 

a 

s 

a 

I X 0 297K (75OF) 

a 0 505K (450'F) s 

Ea 100 

0 I I I11111 

lo2 1 o3 1 o4 1 o5 1 o6 lo7 

65321 7-1 76 NUMBER OF CYCLES 

wE 

2 

s 

\ 

- 

y 800 

c 

u 

a 

w 

v) 

v) 

100 

k 

2 

5 X 

a 

600 

Figure 11-23 Axial Fatigue Properties of Notched [Oo f HT-S/710 G/PI at a Stress 

Ratio, R, of -1 

o FAILED IN'TEST SECTION 

0 DOUBLER BONO FAILURE 

- 20g 

x 

0 3 

L 

c 

0 

0 a 

0 

.O ' 

n 

0 297K (75'F) 0 0 

0 505K (450°F) 

I 1 I 11 Ill I I I11111 I I 1 I Ill1 I I I11111 I I I1111 80 

900 I 

Figuro 11-24 Axial Fotiguo Propertior of Unnotchod [00l6 HTW710 G/PI at a Stress Ratio, R, of 0.1 

1 1-33. 
z 200 

U 

s 

- 

0 297 K (75' F) 

0 505 K (450' F) 

930 2 

-$r 

- z. 

0 297K (75°F) 

0 505K (450' F) 

- 20 

. I I I I I Ill 1 I I I I Ill I 111l1l1 I I I I I Ill I I l11l1l 

I I111111l I I I I1 Ill1 1 I I1 1 11 111 I I I IIIII I I1ll111 

1 o7 

65321 7- 170 

lo3 lo4 1 o5 lo6 

NUMBER OF CYCLES 

Figure 11-26 Axial Fatigue Properties of Notched [0°]6 HT-S/710 G/PI at a Stress Ratio, R, of 0.1 

11-34 
- tl - 70 

0 3 - 60 3 

B 

x 

- 297 K (75' F) 

x, 

505 K (450' F) 0 2 SPECIMENS 

0 

0 

0 

2 300 

- =i 

c) 

< 

c) 

r 

0 I 

0 297 K (75' F) 

30 n 50s K (4500 FI 

-E - 

0 FAILED AT NOTCH 

FAILED UNDER DOUBLER - 20 

I I111111 I IIIIIIl I 1 I 1 lltl I I I I1111 

1 o3 lo4 1 o5 1 o6 lo7 ! 

NUMBER OF CYCLES 65321 7- 180 

I 

t I 

Flguro 11.27 Axial Faiiguo Proporifi.r of Notehod [00]12 HT-S/710 G/PI at a Strorr Ratio, R, of - 1 

0 

-; I R - 0.1 

---- NOTCHED 

I I1 I1111 I I I I IIII I I IlllU 

w 

A 

u 

> 

u 

cy 

v) 

w 

U 

ZE 

3 

a 

200 

G 

3 

x 100 

a 

z - UNNOTCHED I I 

I--"i--1.6 c z 

I lo 

1 o2 lo3 lo4 lo5 1 o6 10' 

NUMBER OF CYCLES 

Figuro 11-28 Axial Fatiguo PtoporHor of [Oo -I- 4S0] Hf-S/710 G/PI at 297 K (75O F) 

653 2 1 7- 1 8 1 

11-35 
900 

64- c 800 

700 

- 

z 

- 5 

d 

0 

> 

w 

’ 140 

- - ----- 

-- ----.I- - 120 

--- ----- -2- - .) L 

Rs0.1 4 

k 

\ - 100s 

z= 

d 

u 

> 

u 

0 

a 

w 

L 

- --- - 2. 

UNNOTCHED 

---- NOTCHED - 10 

I I1111 1 I1 1111. I I I I I Ill I I Illll 

rn 

CA 

Y 

111.0 

- 

a 

Figure 11-30 Axial Fotigue Properties of [OO -C 4S0] HT.S/710 G/PI at 505 K (450O F) 

11-36 
cn 

w 

a 

G400, 

s 

z 

U = 

300 

200 

The fatigue tests were all carried out at mom temperature at an R value of 0.1 using unnotched 

specimens. The test results are tabluated in Tables 11-42 and 11-43 and S-N cwes are plotted 

in Figures 11-54 and 11-55. These limited data indicate that the position of the 0 layers in the 

composite has an effect on both tensile and fatigue strength. In like manner to the tensile 

results the fatigue properties are lowest with the 0 O layers on the outside and highest with the 

0" layers at the center. The S-N curves for the BE system show the relationship between 

fatigue strength and stacking sequence quite clearly. The same effect, while apparent in the 

GIE curves, is not as distinct because of the increased scatter of the data. 

. 

V 

rn 

< 

f - 60 0" 

R =-I CI 

m 

L 7r 

c 

---..-- 

UNNOTCHED -402 \ 

---- NOTCHED 

I I 111111 1 1 I I1 Ill I 4 I 11111 I 1 I I1111 1 I 1 IIl11.20 

The data presented in this section indicate that the [45 O, -45 O, 0 "1, sequence of the [0 O f 45 "1 

family of laminates would have the best room temperature tensile and fatigue properties. For 

all testing of crossplied laminates on the balance of the test program, however, the stacking 

sequence that was selected was the [0 O, 45 O, -45 'Is sequence. The reason for this choice was 

that the joint strength to attachments, Le., doublers, is much higher when made to an unidirectional ply than to an off-axis ply. Consequently, to minimize doubler bond failures in the 

various types of test specimens the [0 O, 45 O, -45 OIs layup was used for all crossplied panels. 

11 3.5 POISSON'S RATIO EVALUATION. One of the tasks included in the constant amplitude 

baseline fatigue study was a series of residual Poisson's ratio tests on specimens that survived 

10 million fatigue cycles at room temperature at a stress ratio of -1. Determinations were 

made on unidirectional specimens of GIE, HT-SI710 GIPI, and BIAL Strain gage installation 

and test procedures were the same as those described in Section 7. The condition of the specimens (stress level and number of fatigue cycles) and the results of the Poisson's ratio tests are 

tabulated in Table 11-44. Also included are the tensile strength and modulus values obtained 

during the same tests. For the organic matrix materials, GIE and GIPI, the 10 million fatigue 

cycles appeared to have little effect on the value of Poisson's ratio. For the metal matrix 

system, BIAL however, one specimen exhibited a drastic decrease in Poisson's ratio with the 

11-37 
Tabio 11-22. Strongfh at Fatiguo Limit or lo7 Cydor Expressed os Poreontogo of Uhimato Tonsilo Strongth 

for HT-S/710 G/PI 

Spcimon Tempomturo R fotfgum ~trongih at 107 Cydes 

Oriontation Configumtion K (OF) Voluo (76) 

[O" & 4501 

[O "I 

Unnotched 

Notched 

unnotched 

Notched 

Unnotched 

Notched 

Unnotched 

Notched 

Unnotched 

Notched 

Unnotched 

Notched 

unnotched 

Notched 

unnotched 

Notched 

297 

297 

505 

505 

297 

297 

505 

505 

297 

297 

505 

305 

297 

297 

505 

505 

75 

75 

450 

450 

7 5' 

75 

450 

450 

75 

75 

450 

450 

75 

75 

450 

450 

0.1 

.1 

.1 

.1 

-1.0 

-1.0 

-1.0 

- 1.0 

.1 

.1 

.1 

.1 

-1.0 

- 1.0 

- 1.0 

-1.0 

60 

85 

57 

100 

30 

24 

26 

28 

55 

63 

60 

50 

19 

25 

23 

27 

second showing a moderate loss. Modulus values were slightly lower after fatigue cycling for 

the GIE and GlPI systems. All of the specimens showed significant decreases in tensile 

strength. It should be pointed out, however, that the specimen configuration, 2.5 cm (1 in) by 

10.2 cm (4 in.) with a 1.3 cm (0.5 in.) gage section was not a good one for tensile testing and was 

not the one used for the baseline as received tests. Also, all but one of the specimens failed 

under the end doublers. 

The purpose of these tests was to attempt to determine the amount of damage accumulated 

after 10 million stress cycles at different values of maximum stress. Based on the residual 

Poisson's ratio results alone, the resin matrix composite systems, GIE and GIPI, appear to 

have experienced only slight damage while the BlAl metal matrix system was significantly 

affected by the fatigue cycling. The more highly stressed BIAl specimen showed the greater 

decrease in Poisson's ratio. Additional testing would be required to verify these findings. 

11 3.6 BASELINE FATIGUE DATA FOR THE THERMAL AGINGKOMPRESSION STUDY. The find portion of the baseline fatigue study was determination of S-N curves for the 24 ply A-SI3501 G/E 

and Celion 60001LARC-160 GIPI to be evaluated in the Thermal AgingKompression study. All 

of these tests were conducted at room temperature at a stress ratio of -1. The specimen configuration was the same as the unnotched, 24 piy, R = 10 specimen described in Table 11-1. 

For GIE, fatigue testing was performed in the same machine (SFlOU) used for the Phase I 

baseline test program at a constant frequency of 30 Hz. Loads for the somewhat thicker lami11-38 
11-39 
Tablo 11-24. Axial Fdguo Data for [Oo f 45O], Colion 6000/LARC-160 G/PI, Unnofchod, R = 0.1 

TOd Maximum 

Spoc'i mon Ternporoturo Strmrr 

Nurnbor K (OF) MN/d (kri) Cyclor Comrnmnts 

FC72-19 

-20 

-21 

-22 

-23 

-24 

-25 

-26 

-27 

FC72-28 

-29 

-30 

-3 1 

-32 

-33 

-34 

-35 

-36 

297 75 345 

414 

483 

517 

500 

483 

448 

465 

483 

450 350 483 

517 

448 

414 

379 

483 

500 

517 

552 

50 

60 

70 

75 

72.5 

70 

65 

67.5 

70 

70 

75 

65 

60 

56 

70 

72.5 

75 

80 

10,056,000 

10,281,000 

563,000 

1,000 

1,000 

28,000 . 

4,453,000 

12,849,000 

3,740,000 

147,000 

82,000 

2,356,000 

7,503,000 

81.000 

56,000 

9.000 

< 1.000 

< 1,000 

Did not fail 

Did not fail 

Failed under doubler 

Failed under doubler 

Failed under doubler 

Failed under doubler 

Did not fail 

Failed under doubler 

Failed under doubler 

nates of Celion 60001LARC-160 G/PI were higher than those for the G/E and could not be held 

by the conventional L-shaped steel fixtures used for the other composite materials. The slipping 

problem was solved by using large hydraulic grips. These large gripping fixtures, however, 

required a change to an MTS Closed-Loop Tension-Torsion machine. The cycling capabilities of 

the MTS machine with the large hydraulic grips installed limited the test frequency to 5 Hz 

rather than the 30 Hz used for all other constant amplitude fatigue testing. Figure 11-56 shows 

the test machine and grips with a specimen ready for testing. 

Results of the fatigue tests are listed in Table 11-45 for the two materials. S-N curves are 

plotted in Figures 11-57 and 11-58. A comparison of these curves with those for the 12 ply 

materials presented previously (Figure 11-9 for GIE and Figure 11-33 for G/PI) shows good 

agreement, particularly for the G/PI. The effect of fatigue cycling the G/PI at 5 Hz rather than 

30 Hz appeared to be slight. As a check, one of the specimens that was tested at the lowest 

load was run successfully using the 30 Hz test setup. This data point, (Specimen F3-8) fell on 

the best fit line for the 5 Hz data. 

11-40 
Table 11-25. Axial Fdguo Data for [OO 4 450ld Cdion 6000/LARC=160 G/Pl, Unnokhd, R = - 1 

Tut Maximum 

fpuimon Temperature Stress 

Number K (OF) MN/d (ksi) Cydes Comments 

FC72-1 

-2 

-3 

4 

-5 

-6 

-7 

-8 

-9 

FC72-10 

-11 

-12 

-13 

-14 

-15 

-16 

-17 

-18 

297 75 212 

345 

276 

241 

259 

276 

224 

259 

241 

450 350 207 

172 

190 

224 

259 

293 

310 

328 

207 

30.8 

50 

40 

35 

37.5 

40 

32.5 

37.5 

35 

30 

25 

27.5 

32.5 

37.5 

42.5 

45 

47.5 

30 

12.9 16,000 Did not fail 

< 1,000 

1,000 

728,000 

74.000 

2,000 

5 19,000 

39,000 

34,000 

176,000 

10,192,000 Did not fail 

10.030,OOO Did not fail 

838,000 

110,000 

19,000 

4,000 

< 1,000 

4,409,000 

Tablo 11-26. Axial Fdgue Oota for [OO f AP], Cdion 6000/LARC-160 G/PI, Unnotehed, R = 10 

TOSt Maximum 

Spocimen Temporatun S tross 

Number K (OF) MN/d (kti) Cydes Comments 

FC72-37 

-38 

-39 

-40 

-4 1 

-42 

-43 

-44 

-45 

FC72-46 

47 

-48 

-49 

-50 

-51 

-5 2 

-53 

-54 

297 75 310 

345 

379 

276 

379 

362 

293 

328 

310 

450 350 310 

345 

276 

293 

310 

328 

328 

336 

336 

45 

50 

55 

40 

55 

52.5 

42.5 

47.5 

45 

45 

50 

40 

42.5 

45 

47.5 

47.5 

48.75 

48.75 

298,000 

49,000 

4,000 Bond failure 

30,673,000 Did not fail 

3,000 Bond failure 

2.000 Bond failure 

455.000 Bond failure 

1,199,000 

10.1 12,000 Did not fail 

17,000 

1.000 

9,008,000 

394,000 

2 15.000 

275,000 

566.000 

118,000 

10,167,000 Did not fail 

11-41 
Tablo 11-27. Axial Fatiguo Data for [OO f 45O], Celion 6000/LARC-160 G/Pl, Notehod. R = 0.1 

~ 

Test Maximum 

Spoeimon Temporaturo Strosr 

Numbor K (OF) MN/d (kri) Cydos Comrnonts .) 

FC71-19 

-20 

-2 1 

-22 

-23 

-24 

-25 

-26 

-27 

FC71-28 

-29 

-30 

-3 1 

-32 

-33 

-34 

-35 

-36 

297 75 483 

500 

517 

465 

465 

465 

414 

448 

483 

450 350 379 

483 

414 

448 

465 

483 

517 

500 

345 

70 

72.5 

75 

67.5 

67.5 

67.5 

60 

65 

70 

55 

70 

60 

67.5 

70 

75 

72.5 

50 

65 

1,080.000 

18,000 

< 1.000 

< 1.000 

< 1,000 

<l,OOO 

< 1,000 

< 1,000 

10,000,000 Did not fail 

10,000,000 Did not fail 

8,058,000 

7,859,000 

4,250,000 

1,567,000 

1,101,000 

< 1.000 

1,502,000 

10,030,000 Did not fail 

Tablo 11-28. Axial Fatiguo Data for [Oo f Cmlion 6000/LARC-160 G/PI, Notehod, R = - 1 

Test Maximum 

Spocimon Tornpomturo Strosr 

Numbor K (OF) MN/d (kri) Cyclos Cemmonts 

FC71-1 

-2 

-3 

-4 

-5 

-6 

-7 

-8 

-9 

FC71-10 

-1 1 

-12 

-13 

-14 

-15 

-16 

-17 

-18 

297 75 207 

224 

241 

259 

276 

293 

190 

241 

207 

450 350 241 

276 

259 

224 

224 

190 

172 

155 

207 

30 

32.5 

35 

37.5 

40 

42.5 

27.5 

35 

30 

35 

40 

37.5 

32.5 

32.5 

27.5 

25 

22.5 

30. 

149,000 

419,000 

24,000 

23,000 

2,000 

2,000 

10,031,000 

64,000 

292,000 

Did not fail 

312,000 

3,000 

6,000 

14.000 

76,000 

1,386,000 

3,468,000 

668.000 

10.5 18,000 Did not fail 

11-42 
Tablo 11-29. Axial Futiguo Data for [OO f 4501, Celfon 6000/LARC-160 G/PI, Notchod, R = 10 

Test Maximum 

Spoeimon Tompomturo Strosr 

Nurnbor K (OF) MN/d (ksi) Cych Commonts 

FC71-37 

-38 

-39 

-40 

41 

42 

-43 

-44 

-45 

FC7146 

-47 

48 

49 

-50 

-5 1 

-52 

-53 

-54 

297 75 310 

241 

276 

293 

302 

284 

284 

276 

276 

450 350 293 

276 

24 1 

259 

276 

310 

267 

293 

45 

35 

40 

42.5 

43.75 

41.25 

41.25 

40 

40 

42.5 

40 

35 

37.5 

40 

45 

38.75 

42.5 

- 

<l,OOO 

21,455,000 

10,372,000 Did not fail 

6,000 

1,000 

6,000 

42,000 

75,000 

4,004,000 

4,000 

6,000 

7,390,000 Temp. over-shoot 

10,060,000 Did not fail 

10,000 

< 1.000 

- Machine malfunction 

1,704,000 

3,000 

11-43 
600 

500 

400 

300 

200 

100 I I I I I Ill 

- ir 0 450K (35OoF) - u. - 30 I 

I I 1 IIIII 1 I I I I Ill I I I I I Ill I I I Ill 

o a FAILED IN TESTSECTION 

0 FAILED UNOER DOUBLER I 

600 

- 

500 - 

- 

40 0 

.-50 i 

300 KOF) I I 

\ 

I - 

I 

0 I 

80 

70 E 

E 

> 

= 

60 v) 

-I 

J 

VJ 

m 

m 

n 

40 g 

c) < 

m 

F 

c. 

102 103 1 04 105 106 107 

65321 7-331 NUMBER OF CYCLES 

200 

100 

Figure 11-32 Axial Fatlguo Pmportfes of Unnotched [0° f 4501, Colion 6000/LARC-l60 G/PI at a 

Stress Ratio, R, of 0.1 

---,-g 297K (75'F) / ---TL 430 - 

0 297K (75OF) I I 3 

- 20 I 

I 450K (350'F) 

I I ! 1111t 1 1 I ((((I I I I llIl11 I I I I(l(I I I I IIIII 4 

Figuro 11-33 Axial Fatigue Properties of Unnotched [OO f 450Is2 Coiion 6000/LARC-160 GP! at a 

Stress Ratio, R, of -1 

11-44 
600 

< 500 - 

< 

z 

- s 

w 

A 

u 400 

a 

w 

0 

v) 

v) 

w 

a 

s 

a 

z 

a 

300 

X 

200 a 

100 

0 FAILED IN TEST SECTION 

102 104 105 106 107 

NUMBER OF CYCLES 

65321 7-333 

Figuro 11-34 Axial Fotiguo ProporHor of Unnotchd [OO f 45'5 Celion 6000/LARC-160 G/PI at a 

600 

500 

400 

300 

200 

100 

102 1 o4 105 106 107 

653217.334 

NUMBER OF CYCLES 

-- 

Stress Ratio, R,. of 10 

Figure 11.35 Axial Fatigue Properties d Notched [Oo k Go], Celion 6000/LARC-160 G/PI at a 

Stress Ratio, R, of 0.1 

11-45 
600 

500 

400 

Figure 11-36 Axial Fatigue Properties of Notched [Oo f 45O],2 Celion 6000/LARC-160 G/PI 

at a Stress Ratio, R, of -1 

* 

- 80 

3 

> 

x 

- 60 

60C 

5 00 

400 

300 

200 

100 

- 

Figure 11-37 Axial Futigue Properties of Notched [OO f 45O],, Celion 6000/LARC-160 G/PI 

at astress Rotio, R, of 10 

11-46 

cn + 

J 

m 

50 g 

TI 

m 

300 

200 

. 

100 

rJ 

m 

' 

0 297K f75'F) 

0 450K f35O'F) 

P I I I I Ill I I I I I Ill I I 1 I I Ill I t 1 I I Ill I I l!'lll 
600 

500 

400 

300 

200 

too 

- UNNOTCHED 

- -- NOTCHED - 20 

102 103 1 04 105 106 107 

NUMBER OF CYCLES 

65321 7-337 

I I I I! Ill I I I I I Ill I I I I I Ill I I I I I Ill I I I I lllL 

figure 11-38 Axial Fatigue Properties of [Oo f Go] Colion 6000/LARC-160 G/PI at 297 K (750F) 

600 

500 

400 

300 

200 

100 

- 80 

z R a 0.1 (A 

- 70 5 

x = 

c 

m 

J 

- - - -= R 

- z. 

- 

- 20 

UNNOTCHED 

-0- NOTCHED 

I I I1l111 I I I I I Ill 1 I I IIl[l I I I IIIII 1 I I I :Ill 

Figure 11-39 Axial Fatigue Properties of [OO f Go] Celion 6000/LARC-160 G/PI at 450 K (350OF) 

11-47 
Tablo 11-30. Strongth at Futigue Limit or lo7 Cycles Exprossod as Percentago of Ultimo). Tensilo Strongtfr 

for Celion 6000/LARC-160 G/PI 

Spdmon Temporaturo R Fatiguo Strength ot lo7 Cyclor 

Oriontation Configuration K (OF) Value ("/.I 

[O 4 45 "1 Unnotched 

Notched 

Unnotched 

Notched 

Unnotched 

Notched 

Unnotched 

Notched 

unnotcfied 

Notched 

Unnotched 

Notched 

297 

297 

450 

450 

297 

297 

450 

450 

297 

297 

450 

450 

75 

75 

350 

350 

75 

75 

350 

350 

75 

75 

350 

350 

0.1 

.1 

.1 

.1 

10 

10 

10 

10 

-1.0 

- 1.0 

-1.0 

-1.0 

66 

100 

36 

81 

42 

64 

42 

55 

30 

45 

27 

33 

Tablo 11-31. Axial Fatigue Data for [0° -C Go], WAl, Unnotched, R = 0.1 

Test Maximum 

Specimon Temperature Strosr 

Number K (OF) MN/d (kti) Cycler Comments 

EC72-20 

-21 

-22 

-23 

-24 

-25 

-26 

-27 

-28 

EC72-47 

-48 

-49 

-50 

-5 1 

-52 

-53 

-54 

-56 

297 75 262 

379 

414 

345 

448 

483 

517 

465 

414 

561 560 414 

379 

345 

310 

276 

241 

224 

207 

448 

38 

55 

60 

50 

65 

70 

75 

67.5 

60 

60 

56 

50 

45 

40 

35 

32.5 

30 

65 

17.45 1,000 Did not fail 

1,876,000 

72,000 

716,000 

1.686.000 

17.000 

122,000 

2,627,000 

53,000 

201,000 

571,000 

1,780,000 

2,691,000 

9,385,000 

9,341,000 

10,551,000 

66,000 

33,360,000 Did not fail 

11-48 
tablo 11-32. Axial Fatiguo Data for [Oo f 4501, B/Al, Unnotchd, R = - 1 

Ted Maximum 

Spdmon Tempomtun . Strorr 

Numbor K (OF) MN/d (ksi) Cyctor Commonts 

EC72-11 

-12 

-13 

-14 

-15 

-16 

-17 

-18 

-19 

EC72-38 

-39 

-40 

-4 1 

-42 

-43 

-44 

-45 

-46 

297 75 345 

379 

310 

276 

207 

241 

190 

414 

362 

505 450 276 

207 

345 

379 

310 

241 

138 

172 

259 

50 

55 

45 

40 

30 

35 

27.5 

60 

52.5 

40 

30 

50 

55 

45 

35 

20 

25 

37.5 

38,000 

8.000 

72,000 

448,000 

6,610,000 

788,000 

1,000 

2,000 

58,000 

861,000 

4,000 

2,000 

7,000 

52,000 

15,437.000 Did not fail 

10,006,000 Did not fail 

10,111,000 Did not fail 

38,000 

tablo 11-33. Axial Faiiguo Data for [Oo k 4501, B/Al, Notched, R = 0.1 

TOSt Maximum 

Spocimon temporatun Strorr 

Numbor K (OF) MN/d (kri) Cyclor Comments 

EC71-10 

-11 

-12 

-13 

-14 

-15 

-16 

-17 

-18 

E C7 1-28 

-29 

-30 

-3 1 

-32 

-33 

-34 

-35 

-36 

297 75 345 

276 

414 

379 

310 

448 

396 

362 

328 

561 550 345 

276 

241 

207 

310 

345 

414 

448 

224 

50 

40 

60 

55 

45 

65 

57.5 

52.5 

47.5 

50 

40 

35 

30 

45 

50 

60 

65 

32.5 

492.000 

24.000 

1,192.000 

5,000 

901.000 

2.913.000 

10,150.000 Did not fail 

10,130,000 Did not fail 

10,085,000 Did not fail 

No data Temp. over-shoot 

2,634,000 

5,490,000 

790.000 

477,000 

91.000 

79,000 

8,247,000 

10.1 00,000 Did not fail 

11-49 
Tablo 11-34. Axial Fatiguo Data for [OO f 450],2 WAl, Notchod, R = - 1 

~ ~~ ___ 

Test Maximum 

Spoeimom Tempomturo Stress 

Numbor K (OF) MN/d (ksi) Cycles Commonts 

EC71-1 

-2 

-3 

-4 

-5 

-6 

-7 

-8 

-9 

-20 

-21 

-22 

-23 

-24 

-25 

-26 

-27 

E C71- 19A 

297 73 276 

310 

241 

345 

207 

379 

172 

328 

224 

505 450 207 

241 

276 

172 . 

138 

310 

345 

224 

259 

40 

45 

35 

50 

30 

55 

25 

47.5 

32.5 

30 

35 

40 

25 

20 

45 

50 

32.5 

37.5 

247,000 

60,000 

547,000 

8,000 

6,480,000 

< 1,000 Failed under doubler 

10,195,000 Did not fail 

7,000 

1,227,000 

739,000 

42,000 

10,000 

2,309,000 

4,000 

2,000 

26,000 

12.000 

12,710,000 Did not fail 

Tablo 11-35. Axial Fatiguo Data for [Ool6 B/Al, Unnotchod, R = 0.1 

Test Maximum 

Spoeimon Temperatun Stress 

Numbor K (OF) MN/d (kri) Cyclos Commonts 

EU72-20 

-21 

-22 

-23 

-24 

-25 

-26 

-27 

-28 

EU72-47 

-48 

-49 

-50 

-5 1 

-52 

-53 

-54 

-55 

297 75 758 

1,000 

93 1 

862 

827 

827 

1,030 

896 

1,100 

561 560 862 

827 

793 

758 

724 

8 27 

965 

965 

1,030 

110 

145 

135 

125 

120 

120 

150 

130 

160 

125 

120 

115 

110 

105 

120 

140 

140 

150 

12,140,000 Did not fail 

302,000 

54,000 

76,000 

2,000 

798,000 

48,000 

18P.000 

2,185,000 

3,209,000 

3,104.000 

3,761,000 

4,080,000 

2,980,000 Nitrogen Atmosphere 

1,364,000 

1,856,000 Argon Atmosphere 

610,000 

30,559,000 Did not fail 

1 1-50 
Tablo 11-36. Axial Fotiguo Doto for [0°]12 WAI, Unnotchod. R = -1 

Test . Maximum 

Spoeimon Tomporaturo Str.8 

Numbor K (OF) MN/d (ksi) Cyc1.r Commonts 

EU72-11 297 75 

-12 

-13 

-14 

-15 

-16 

-17 

-18 

-19 

-46 

-39 

-40 

-4 1 

-42 

-43 

-44 

-45 

EU72-38 505 450 

690 

655 

638 

638 

620 

552 

586 

603 

569 

414 

425 

345 

379 

414 

620 

483 

552 

586 

100 

95 

92.5 

92.5 

90 

80 

85 

87.5 

82.5 

60 

61.7 

50 

55 

60 

90 

70 

80 

85 

3,000 

229,000 

< 1,000 

8,000 

9,000 

646,000 

21,000 

6,000 

12,000 

10,000,000 Did not fail 

10,087,000 Did not fail 

10,120,000 Did not fail 

10,153,000 Did not fail 

10,005,000 Did not fail 

< 1.000 

1,332,000 

51,000 

5.000 

Tablo 11-37. Axial Fatlguo Data for [OO], WAl, Notehod, R = 0. I 

Tost Maximum 

Spocimon Tornpomtun Str.8. 

Numbor K (OF) MN/d (kri) Cyc1.r Comments 

EU71-10 

-11 

-12 

-13 

-14 

-15 

-16 

-17 

-18 

EU71-28 

-29 

-30 

-3 1 

-32 

-33 

-34 

-35 

-36 

297 75 

561 5 50 

561 550 

561 560 

505 450 

505 450 

561 550 

561 580 

561 550 

561 550 

1,000 

896 

827 

758 

1,030 

965 

1,100 

93 1 

793 

862 

690 

6 20 

620 

862 

827 

965 

896 

758 

145 

130 

120 

110 

150 

140 

160 

135 

115 

125 

100 

90 

90 

125 

120 

140 

130 

110 

76,000 

4,634,000 

4,546,000 

39,000 

120,000 

< 1.000 

1,469,000 

1,761,000 

930,000 

3,732,000 

5,482,000 

996,000 

1,057,000 

60,000 

330.000 

2,974,000 

10,002,000 Did not fail 

10,179,000 Did not fail 

11-51 
tablo 11-38. Axial Fatiguo Data for [O0Il2 WAl, Notched, R = -1 

Test Maximum 

Spdmon Temporaiun Strorr 

Numbor K (OF) MN/d (ksg Cycler 

~~ 

Comments 

EU71-1 

-2 

-3 

-4 

-5 

-6 

-7 

-8 

-9 

EU71-19A 

-20 

-21 

-22 

-23 

-24 

-25 

-26 

-27 

297 75 586 

620 

655 

724 

758 

793 

690 

552 

586 

505 450 552 

483 

414 

448 

566 

379 

517 

465 

43 1 

85 

90 

95 

105 

110 

115 

100 

80 

85 

80 

70 

60 

65 

82.1 

55 

75 

67.5 

62.5 

310,000 

560,000 

887,000 

34,000 

9.000 

3,000 

62,000 

12,486,000 

2,977,000 

6,000 

18,000 

4,253,000 

3,931,000 

4,000 

10,829,000 

9,000 

2,398,000 

554,000 

Failed under doubler 

Failed under doubler 

Failed under doubler 

Failed under doubler 

Failed under doubler 

Failed under doubler 

Did not fail 

Failed under doubler 

500 

mE 

z 

H 

w 

> 

u 

w 

0 

v) 

v2 

w 

\ 

u’ 400 

a 

a 

t; 300 

z 

5 

X 

U 

z 

200 

1 

0 297K (75’F) 

0 561 K (550°F) 

104 1 o5 1 Q6 10’ 

653217-185 

NUMBER OF CYCLES 

Figure 11-40 Axial Fatigue Properties of Unnotched [Oo f 45O], WAI at a Stress Ratio, R, of 0.1 

11-52 
-; 400 

< 

2 

- a 

w 

-1 

u 

>. 

300 W 

Q 

v) 

v) 

W 

a lz 

s 200 

z 

x 

U 

=E 

- 

100 

6S3217- 186 

Flgure 11-41 Axial Fatigue Propedies of Unn0teh.d [OO f 45OIrZ B/Al at a Stress Ratio, R, of -1 

- 

r 

0 297K (75°F) 

0 505K (450OF) - 

I 1 I11111 I I I Ill I I111111 I I I1 IIII I I I I Ill 
." 

NUMBER OF CYCLES 

653217.188 

Figuro 11-43 Axial Fatiguo Proportii.. of Notchod [OO f 450],2 B/Al at a Stress Ratio, R. of -1 

1100 

mi 

E 

\ 

- z 1000 

w 

-I 

u 

900 

a 

w 

p. 

g aoo 

3 

w 

c 

a 

v, 700 

5 

2 600 

5 

500 

1 

NUMBER OF CYCLES 653217-189 

Figuro 11-44 Axial Fatiguo Properties of Unnotched [00]6 B/Al at a Stress Ratio, R, of 0.1 

1 1-54 
700 

600 

500 

W 

m 

0 297K 561K (75"50°F) FI 505 K (450' F) 5: z 

300 

1 

NUMBER OF CYCLES 65321 7- 190 

Ftguro 11-15 Axial Fuiiguo Propor)lor of Unnotchd [00]12 B/Al at a $)toss Ratio, R, of - 1 

1100 

ru2 

=E 

3 1000 

_- 

900 

aoo 

700 

Figuro 11-46 Axial Fatiguo Proportior of Notched [0°]6 at a Strors Ratio, R, of 0.1 

11-55 
Figurw 11-48 Axial Fotigue Proportier of [Oo f 45O] B/Al at 297 K (750 F) 

11-56 
1100 I I rt I I t 160 

400 

w 

u 

>- 

0 

w 

0 

a 

- 60 

1 I I I I1111 I I 1 I1 Ill I 1 I 11l11 I I I 1 I Ill I I I111L 

v) 

v) 

w 

a 

G 

3 

a 

I 

- 

X 

R=(Ll< 

1000 1 - 140 

""" I fI 

v) I: 

lo2 

500 

Fin 

E 

z 

w 

A 

0 

> 

u 

w 

a. 

v) 

\ 

z 400 

a 

300 a 

k 

s 

a 

z 

x s 20(1 

lo3 lo4 to5 1 o6 10' 

NUMBER OF CYCLES 65321 7- 194 

Figuro 11.49 Axial Fatiguo Propodor of [OO] B/Al at 297 K (75O F) 

1 o2 1 o3 lo4 1 o5 to6 lo7 

653217-195 

. 

Figurr 11-50 Axial Fatigur Proportier of [OO f 4501 B/Al at Elevated Temporaturer 

11-57 
1 o6 lo7 

65321 61 96 NUMBER OF CYCLES 

Figun 11-51 Axial Fdguo ProporHos of [OO] B/Al at Elovatod Tempora)uror 

1 OOC 

Nw ’ 800 

a 

z 

3 

w 

700 

w 

a 

400 

Figuro 11-52 Surfaco Roughoning of B/Al Spocimon Fatiguod at 561 K (550O F) in Air. 

(Roughening is confinod to rtresrod amas on eithor lido of tho hole.) 

11-58 
Figure 11-53 Area of Degraded B/Al Surface Showing Reticulation and Exposed 

(Aluminum-Covered) Boron Fiber 

1 1-59 
Tablo 1 1-39. Strongth at Fdiguo Limit or lo7 Cycler Exprersod as Poreontogo of Ulti~wato Tonsilo Strongth 

for WAl 

Spoeimon Tomporafure R Fatiguo Strongth at lo7 Cyclos 

Orientdon Configuration K (OF) Valuo (7%) 

[O" 4 45'1 

[O "I 

Unnotched 

Notched 

unnotched 

Notched 

unnotched 

Notched 

unnotched 

Notched 

unnotched 

Notched 

Unnotched 

Notched 

unnotched 

Notched 

Unnotched 

Notched 

297 

297 

561 

561 

297 

297 

505 

505 

297 

297 

561 

561 

297 

297 

505 

505 

75 

75 

550 

550 

75 

75 

450 

450 

75 

75 

550 

560 

75 

75 

450 

450 

0.1 

.l 

.1 

.1 

-1.0 

-1.0 

-1.0 

-1.0 

.1 

.1 

.1 

.1 

-1.0 

-1.0 

-1.0 

- 1.0 

74 

71 

50 

47 

39 

41 

38 

32 

57 

64 

a38 

a43 

38 

45 

33 

33 

aDetermined by extrapoiation. 

11-60 
n 

a 

4 

is 

CI 

0 

Y 

n 

a 

U 

- 

c 

4 

0 

0 

m 

5 

a 

L 

c 

= 

5 

a 

L 

a 

m 

0 

Y 

- 

c 

n 

Y 

0 

0 

v) 

v 

x 

v) 

0 nn 

0 

0 

Y 

LI 

Y 

0 

v) 

h 

X 

h 

01 

cy 

v 

A 

Y 

0 

0 

v) 

v 

Y n ov) 

-0 vv) 

Y 

+I 

on 

OY 

v) 

h 

Y 

h 

01 

cy 

Y 

Y 

Q) 

a =: 

- €; 

e: 

aCS 

y 

$Z 0 

U 

ea 

Ita 

0 In 

a0 

m 

a 

e 

ro 

s 

P 

v 

p1 

T! 

0 

B 

e 

Y 

0 

E 

E 

3 

2 

11-61 
. , 

tablo 11-41. Room tomporatun Tonsilo Data for B/E and G/E Stacking Soquonco tort Panols 

Matorial Stadcing fenrilo Strongth tonsilo Modulus 

Syrtom Soquonco MN/d (ksi) GN/d (Msi) 

BIE [OO, 45 O, -45 5 25 

576 

550 

556 

584 

570 

576 

- 

[45 O, 0 O, -45 'IS 

- 

[45 9 -45 O, 0 "1, 

GIE 

" 

580 - 

578 

[O", 45 O, -45 506 

539 - 

5 23 

[45 O, oo, -4501, 664 

652 

658 

699 

- 

[45 O, -45 O, 0 "1, - 675 - 

687 

76.2 

83.5 - 

79.8 

80.6 

84.7 

82.6 

83.5 

- 

84.1 - 

83.8 

73.4 

78.2 - 

75.8 

96.3 

94.5 

95.4 

101.4 

- 

97.9 - 

99.6 

66 9.6 

61 8.8 - 

63 

- 

9.2 

63 9.2 

63 9.1 - 

63 

- 

9.2 

61 8.9 

61 8.9 - 

61 

- 

8.9 

48 6.9 

47 6.8 - 

47 

- 

6.8 

52 7.5 

52 7.5 

32 7.3 

55 8.0 

- - 

56 8.1 - 

55 

- 

8.0 

11-62 
Tablo 11.42. Room Tomporatun Axial Fatiguo Data for WE Stacking Soquoncs Test Panolr, Unnotchod. 

R = 0.1 

Spoeimom Stacking Maximum Strorr 

Numbor Soquonco MN/d (ksi) Cydor Comments 

A62-1 

-2 

-3 

-4 

-5 

-6 

-7 

-8 

-9 

A63-1 

-2 

-3 

-4 

-5 

-6 

-7 

-8 

-9 

A61-1 

-2 

-3 

-4 

-5 

-6 

-7 

-8 

-9 

[O", 45 ] -W], 448 

414 

379 

345 

379 

345 

310 

310 

379 

[45O, 09 -45'1, 448 

414 

379 

448 

414 

448 

414 

379 

396 

[45 -45 O, OOJ5 448 

414 

43 1 

448 

414 

483 

448 

43 1 

414 

65 

60 

55 

50 

55 

50 

45 

45 

55 

65 

60 

55 

65 

60 

65 

60 

55 

57.5 

65 

60 

62.5 

65 

60 

70 

65 

62.5 

60 

1,000 

1.000 

3,000 

294,000 

178,000 

132,000 

3,015,000 

10,145,000 

8,000 

2,000 

81.000 

2,372,000 

1,000 

34,000 

9,000 

21,000 

5,733,000 

111.000 

58,000 

407,000 

16,000 

4,000 

4,950,000 

1,000 

2,000 

7,000 

1,825.000 

Did not fail 

Did not fail 

Failed under doubler 

Failed under doubler 

Failed under doubler 

Did not fail 

11-63 
Tabk 11-43. Room Tompomtun Axial Fatiguo Data for G/E Stacking Soquonco tat Punolr, Unnotchd. 

R = 0.1 

Spoeimon Stacking Maximum. Strors 

Numbor Soquonco MN/d (kri) Cydor Commonts 

B62-1 

-2 

-3 

-4 

-8 

-6 

-7 

-8 

-9 

B63-1 

-2 

-3 

-4 

-5 

-6 

-7 

-8 

-9 

B61-1 

-2 

-3 

-4 

-5 

-6 

-7 

-8 

-9 

[O", 45 ', -45'1, 379 

414 

448 

483 

448 

414 

43 1 

431 

431 

[45O, O', -4501, 483 

448 

483 

5 17 

448 

414 

414 

43 1 

431 

379 

552 

517 

483 

448 

483 

465 

465 

448 

[45 O, -45 ', 0 '1, 

55 

60 

65 

70 

65 

60 

62.5 

62.5 

62.5 

70 

65 

70 

75 

65 

60 

60 

62.5 

62.5 

55 

80 

75 

70 

65 

70 

67.5 

67.5 

65 

13,000 

39,000 

36,000 

1,000 

10,000 

1.000 

2,000 

34,000 

2,000 

34.000 

1,000 

1,000 

22,000 

7,220,000 Did not fail 

114,000 Failed under doubler 

2,306,000 ' Did not fail 

33,000 

21,000 

1,000 

1.000 

3.000 

1.000 

3,000 

3,000 

2,001.000 Did not fail 

2,157.000 Did not fail 

2,295,000 Did not fail 

11-64 
500 

450 

400 

350 

300 

1 o2 lo4 1 o5 lo6 lo7 

NUMBER OF CYCLES 65321 7- 199 

550 

N- 

\ E 

- H 500 

z 

w 

d 

u 

> 

0 

a 

w 

CL 

v) 450 v) 

w 

a 

Iv) 

I 

3 

z 

x 

a 400 

z 

- 

Figurr 11-51 Axial Fatigue of B/E Laminabs Showing Eft& of Stocking Sequence, 297 K (750 F), R = 0.1 

I I UTS - 80 

MNh2 (ksd 

0 [45', .45',0°1 687 99.6 

-- 

e 0 A [45', O', -45'1 658 95.4 - 75 

523 75.8 - 

- 70 

. 

- 55 

11 I Ill11 I I I I I Ill I Ill I Ill I I I I Ill I 

e 

I I I IIIL 

0 

65321 7-200 

Figure 11-55 Axial Fatigue of G/E Laminates Showing Effect of Stacking Sequence, 297 K (750 F), R =0.1 

11-65 
tablo 11-44. Results of Roridual Poisson's Rotio testing of Run-Out Futiguo Spmcimom at Room Tompomturo 

tonsilo Tonsilo 

Matoriai Spoccimon Numbor of Futiguo Strors Poisson's Strongth Modulus 

System Numbor Cyclor MN/d (kri) Ratio MN/d (kri) (GN/mZ (Mri) 

Condition 

BU72-15 

BU72-16 

HT-S/710 - 

G/PI. [0 "112 DU72-15 

DU72-46 

B/A1, [0"112 - 

EU72-16 

EU72-46 

Baseline 

10.0 x 106 

12.6 

10.2 

Baseline 

10.6 

10.3 

Baseline 

10.0 

10.1 

0.33 

483 70 .28 

414 60 .33 

448 65 .3 1 

b.39 

207 30 .41 

138 20 .33 

.28 

552 80 .10 

414 60 .23 

1540 

a1.100 

a1.000 

a1,280 

1.190 

a807 

a945 

1,450 

a470 

1,010 

224 141 

a159 130 

a145 123 

a186 129 

172 126 

a117 118 

a137 117 

210 194 

147 201 

%a 191 

20.4 

18.8 

17.9 

18.7 

18.3 

17.1 

17.0 

28.2 

27.7 

29.2 

aFailed under doubler 

b505 K (450" F) value 

11-66 
65321 7-339 

Figure 11-56 Fatigue Testing Machine with Hydraulic Grips Installed 

11-67 ORIGINAL PAGE IS 

OF POOR QUW 
Tablo 11-15. Basolino Axial Fatiguo Data for [OO & Go],, AW3501 G/E and Celion 6000/LARC.160 G/PI 

Thorrnal Aging/Comprosrion Study Matorials, 297K (75OF), R = - 1, Unnotchod 

500 I - 0 FAILED IN TEST SECTION 

FAILED UNDER DOUBLER 

-- - 

= 400 1 

z 

w - u 

> 

P I - u 300 a 

Maximum 

Matorial Spocimon Stress 

Numbor Numbor MN/d (kri) Cycles 

70 

60 5 

X 

E 

50 5j 

a 

- 

v) + 

40 

Comments 

w 

p. 

v) 

v) 

w 

c 

v) 

= 200 

s 

a 

9 5 100 

a 

Ok 

GIE B3-1 

-2 

-3 

-4 

-5 

-6 

-7 

-8 

-9 

mu r) v) 

7 

m 

)30 

-< c) 

m 

e. 

0 

0 - 7 

- 20 I 

- 

- 10 

I I I11111 I I11l11l I II IIll1 I1 I I I!O! I I I I Ill1 0 

i 

G/PI F3-1 

-2 

-3 

-4 

-5 

-6 

-7 

-9 

-a 

207 

241 

276 

258 

224 

207 

241 

241 

224 

238 

276 

310 

310 

328 

258 

238 

207 

293 

30 

35 

40 

37.5 

32.5 

30 

35 

35 

32.5 

34.5 

40 

45 

45 

47.5 

37.5 

34.5 

30 

42.5 

978,000 Failed under doubler 

453,000 

10,000 

13,000 

8,452,000 

27,000 

28,000 

215,000 Bond failure 

61,470 

500 Bond failure 

8,980 

340 

59,800 

883,860 

13,820 

10,287,000 Did not fail 

9,000 Failed under doubler 

10,040,000 Did not fail 

Figuro 11-57 Axial Fotigue Properties of [OO f 4S0Ica G/E, 297K (75OF), R = -1, Unnotched 

11-68 
5 00 

- -- 

z 

I 400 

- 

w 

4 

u 

> 

c 

w 

P 

v) 

v) 

w 

300 

E 200 

2 

9 

2 100 

v) 

a 

102 104 105 106 107 , 

65321 7-341 NUMBER OF CYCLES 

Figure 11-58 Axial Fatigue Propedier of [Oo f 45°)d Celion 6000/LARC-160 G/PI, 297 K (75OF), 

R = -1, Unnotched 

11-69 
SECTlON 12 

FLIGHT SlMULATlON TESTlNG 

12.1 INTRODUCTION 

The objective of this portion of the program was to subject in real time each of the composite 

systems to the load-temperature history that would be experienced by a supersonic transport 

during its 50,000-hour life. To do this a flight simulation apparatus was constructed that was 

capable of loading a number of large specimens of each composite system using a random 

loading pattern while the specimens were heated to the temperatures corresponding to those 

expected to be encountered during flight. As the composites had different strengths and maximum use temperatures, it was necessary that the flight simulator be capable of providing the 

load and temperature suited to each composite. The apparatus had to be large enough to allow 

the testing of a number of replicates of each system to provide adequate statistical data for 

proper evaluation of the endurance of the materiaL In addition, the simulator had to be capable 

of automatically testing a large number of specimens reliably for an extended period of time. 

A load spectrum was generated using the exceedance curves for a supersonic cruise aircraft, 

and this purely random spectrum was recorded on tape in a form suitable for load generation in 

real time so that 25,000 two-hour flights (50,000 hours) could be simulated. Each flight contained climb, cruise, descent, and landing loads with the temperatures corresponding to each 

portion of the flight. A typical load and temperature profile is shown in Figure 12-1. 

The origLnal intent of the program was to use the wearout model developed by Halpin (ref. 6) to 

predict the life and set the peak load levels for the flight simulation tests of all the composite 

systems. Once a random load spectrum had been generated, the test conditions for each composite system could be set by increasing or decreasing the magnitudes of the entire spectrum. 

This was done by setting the magnitude of the peak load in the spectrum. To set these peak 

loads, accelerated tests were NI~ by increasing the frequency of the spectrum. A quarter of a 

lifetime and half a lifetime of loads were applied in 100 hours and 200 hours respectively. This 

was done with the specimens held at a constant temperature corresponding to what was considered to be the maximum use temperature for each individual composite system. These accelerated tests, referred to as short-term tests, were conducted so that the loads normally applied 

during a twehour flight were applied in about one minute. Figure 12-2 shows a flight picked at 

random and plotted to show the nature of load variation as the specimen is held at its maximum use temperature. These data in conjunction with the baseline properties were to be used 

to calculate, by means of the wearout theory, the peak load to be applied during flight simulation testing of each composite. In Phase I short term tests were conducted on the five 

materials originally selected for evaluation: BIE, G/E, BIPI, HT-SI710 GIPI, and BIAl. The 

BIPI system however, was removed early from the short term testing because of numerous 

specimen failures (see Appendix A). Of the remaining systems only the BIE and G/E proved to 

be amenable to the wearout model for selecting long term loads. The difficulty encountered for 

the G/PI and BIAl materials was that for the loads applied during accelerated testing the 

residual strength was observed to increase between onequarter lifetime and one-half lifetime. 

This was contrary to the model, which is based on strength decreasing with time. The manner 

in which the flight simulation loads were determined for the G/PI and BIAl specimens as well 

12-1 
tua c 

z= = 

I, 

,=mz I 

Y 

c 

UY 

d 9 c 

t 

c 

7 

U 

c 

u 

U 

C 

I 

I 

0 

N 

I 

0 Y 

4 

1 %s 

12-2 
2TlMES 

100 HOUK (1/4 Lite) 

v 200 Houn (1/2 Life) 

T = CONSTANT 

1 FLIGHT PER MINUTE TIME 

653217-202 

Ffguro 12-2 Typical Accdomtod or Shor+-Torm Flight Simulation Cydo Showing loads 

and Tomporntun Pmfilo 

as the details of the wearout analysis of the BIE and GIE specimens are discussed later in the 

section. 

Following the short-term tests the four composite systems were subjected to the Phase I 

10,000 hours of flight simulation exposure. The GIE specimens experienced a high failure rate 

and the BIE specimens suffered from extensive visual degradation during these 10,000 hours. 

Because of this greater than expected damage to the epoxy systems new specimens were pre 

pared, and the long term tests were restarted for BIE and GIE. The maximum test temperature 

was reduced using the& aging data obtained earlier in the program. For the HT-SI710 GIPI 

and BIAl systems a portion of the specimens with 10,000 hours of exposure were removed for 

residual property evaluations. Flight simulation testing of the remaining specimens continued 

for the 50,000 hour Phase I1 evaluation. Shortly after the start of the Phase I1 flight simulation tests the new material, Celion 6000ILARC-160 GIPI, was added to the program. Long 

term specimens were prepared and introduced into test for 10,000 hours of exposure. No short 

term testing of the Celion 6000ILARC-160 system was conducted because of the poor results 

obtained with the HT-SI710 GIPI and the delays that would have occurred in flight simulation 

testing of the other systems. Determination of the long term flight simulation loads for the 

Celion 6000ILARC-160 specimens is discussed later in the section. 

12.2 DESCRIPTION OF SIMULATOR 

The flight simulation apparatus was built to test five composite materials and has the capability of cycling loads and temperatures on 100 test specimens simultaneously. The apparatus is 

12-3 
divided into 10 independent test setups each of which contains 10 identical specimens. The 

flight simulation testing equipment is shown in Figure 12-3. There are 10 load frames, five of 

which are facing forward, and five more directly behind the front five. The small cabinet at the 

right of the simulator is the control console, which is shown in greater detail in Figure 12-4. 

12.2.1 1st FIXTURES. The test fixtures consist of 10 basic load frames, 10 whiffletree assemblies, attachment hardware for 100 specimens, 100 removable specimen stiffener assemblies, 

20 heat lamp rack assemblies, and lateral support hardware. 

The lateral support between a whiffletree assembly and its specimens is a beam containing five 

linear motion bearings. They allow motion in the load direction only. The travel of the clevis fitting, which passes through a bearing, is limited and adjustable in each direction. Specimen or 

doubler failures are detected by means of microswitches activated when the clevis motion 

exceeds the preset values in either loading direction. Activation of a microswitch immediately 

dumps the hydraulic pressure and reduces the load to zero. 

12.2.2 LOAD PROGRAMMER AND CONTROLLERS. The MTS digital load programmer and associated load controllers are shown in Figure 12-5. Ten of the 60 available controllers are used for 

the 10 basic load frames. The load programmer and controllers program and control the fatigue 

loading spectrum and synchronize the heating and cooling cycles. 

12.2.3 PROGRAMMER SAFETY FEATURES. The load' programmer provides the following selfchecking features that automatically stop all test operations if activated. 

*, . .. ~ _". 

- , " ,. ... ^., ,,. 

Figure 12-3 Flight Simulation Equipment 

12-4 
Figure 12-4 SCR Flight Simulation Apparatus - Control Console 

12-5 
Figure 12-5 Programmer and load Controllers 

a. 

b. 

C. 

The load controller ERROR DETECTOR is activated if the error between the command 

and feedback system exceeds a preset value. 

Program HOLD is activated if a malfunction occurs in timing, parity, or the power supply. 

END OF PROGRAM code at the end of a punched tape places the program in “hold.” If 

the test program is not manually restarted within a few minutes, all operations are 

deactivated. 

12.2.4 HYDRAULIC SYSTEM. The hydraulic system consists of a hydraulic power source with oil 

reservoir level and oil temperature detectors, a manifold pressure dump valve, distribution 

manifolds and lines, and hydraulic actuating cylinders with pressure dump valves, filters, and 

servo valves. 

Two hydraulic power sources were used during the test program. One 50 gpm pump of the 200 

gpm Pump Station was used for the short-term tests. A 7 gpm portable hydraulic pump was 

used for the long-term tests. 

12.2.5 HEATING SYSTEM. The test specimens are heated by 800-watt infrared quartz heat 

lamps. The heat is’directed to the specimens by parabolic reflectors as shown in Figure 12-6. 

The lamps and reflectors are positioned so that they heat a 0.05 m (2 in.) strip on each half of 

the specimens. As discussed more fully in Section 12.3, this arrangement doubles the number of 

residual strength specimens that can be obtained from each long-term specimen. For each test 

setup, the five upper lamps and the five lower lamps are assembled in an upper rack and a lower 

12-6 

ORIGiNAL PAGE IS 

QF POOR QUALI‘IX 
i 

bREFLECTOR =.- 

- 

-SPECIMEN / 

I 

\ "EAT/ 

LAMP 

-0- 0- 

allm (45 IN.) 

I 

65321 7-206 

Figuro 12-6 Hating System Showing Armngomont of Specimens, Quartz Heatlarnpr, and Parabolic 

Rofl.eton 

3 

12-7 

I 
rack Each lamp rack is an independent heating system with its own temperature controller, 

control thermocouple, monitor thermocouple, and temperature error detection thermocouple. 

The temperature cycling for the long-term test is synchronized with the loads by using the 

STEP code in the load programmer. STEP codes are programmed at the start and end of the 

heating period. The first STEP code activates the low temperature checking circuit for 10 milliseconds. If the specimen temperature is not less than 311 K (100" F), all test operations for 

that specific setup are deactivated. If the specimen temperature is less than 311 K (100" F), 

then power is applied to the heat lamps. The power is not controlled until the desired specimen 

temperature is obtained. The temperature controller and its control thermocouple will maintain the desired temperature within +. 6 K (lo" F). The temperature error detector thermocouple makes an independent and continuous check against over-temperature occurrences. 

(This thermocoupie is also used by the low temperature checking circuit.) In case of an overtemperature occurrence, all test operations for that specific setup are deactivated. The second 

STEP code removes the power to the heat lamps, and the specimens are ailowed to COOL 

12.2.6 COOLJNG SYSTEM. A large centrifugal blower driven by a five horsepower motor provides the cooling air for the test specimens. The cooling air is distributed to 50 locations by a 

duct system. At each location a nozzle directs the cooling air to a pair of test specimens. Air 

flow at each location can be adjusted by physically changing the nozzle outlet area. These 

nodes are shown near the top in Figure 12-3. Figure 12-7 is a closeup view of the nozzles in 

operation during the long-term tests. 

12.2.7 CONTROL CONSOLE. The control console, shown in Figures 12-4 and 12-8, contains the 

temperature control and checking circuits, malfunctiodshutdown logic circuits, and system 

control and indicator circuits. 

The control console front panel contains the following controls and indicators. 

a. 

b. 

C. 

d. 

e. 

f. 

g* 

h. 

Ten setup counters to maintain the load programmer block count for each set up. 

Ten dump switches/indicators to manually dump or indicate an automatic dump of any 

setup. 

Ten JOG-RESET toggle switches to energize the hydraulics of each setup. 

PROGRAMMER counter to indicate load programmer block count. 

RUN and HOLD switches/indicators to provide remote control of load programmer. 

LOAD and UNLOAD switches and indicators to provide remote control of the hydraulic 

pump pressure to all setups. 

DUMP switchlindicator to completely shut down the hydraulic pump. 

OVER TEMPERATURE, LOW TEMPERATURE, SPECIMEN FAILURE, LOAD 

ERROR, RAM OVERTRAVEL, MANUAL LOCK indicators to indicate which malfunction caused a test to stop. The numbered setup indicator will also indicate which setup is 

related to the malfunction. 

12-8 
12-9 

0FJGINA.L PAGE IS 

OE POOR QUALITY 
Figuro 12-8 Control Consolo 

12.2.8 DATA RECORDING SYSTEM. All data are recorded on a high speed digital recorder. The 10 

loads and 20 monitor thermocouples can be recorded on magnetic tape at a maximum sampling 

rate of six times per second. Data can then be printed out in engineering units. 

Data are recorded periodically to confirm that loads appear to be reasonable and that the temperature cycle profile is correct. Loads cannot be confirmed exactly because the programmed 

loads are not known due to the random nature of the spectrum. 

12.2.9 MONITOR/SAFETY SYSTEM. A Metrascope is used as a monitor and a limit detect safety 

system. Channels 1 through 10 display the tension loads up to 100% with the limits set to 

105%. Channels 11 through 20 display the compression loads (1 through 10) up to 25% with 

12-10 
the limits set at 30%. Channels 21 through 40 display the temperatures of each monitor thermocouple for each of the 20 lamp racks. The temperature limit is set 5% above the maximum 

temperature for a given setup. 

If any limit is exceeded, the main electrical circuit breaker is thrown to its OFF position. All 

heat lamps go off, the hydraulic pressure at the manifold is dumped, the load programmer goes 

into HOLD, and ail 10 setups are individually dumped. The channel that exceeded the limit is 

indicated by a flashing line on that channel. 

12.2.10 CHECKOUT. Tests were performed to check out the loading mechanism and the distribution of loads by the whiffletree assemblies. A set of 10 strain gaged aluminum specimens 

was prepared and calibrated. The specimens were installed in the flight simulator and loaded 

to three load levels in the region of interest. Both the total applied load and the 10 individual 

loads were found to be within desired limits. The individual loads on each specimen were within 

23% of the desired load. 

During checkout of the loading system, a dummy BIE short-term exposure specimen was 

intentionally loaded to failure in the simulator to determine the effects of individual specimen 

failures on the load levels of similar specimens within the same whiffletree arrangement. Each 

whiffletree con& five pairs of specimens. The specimens are loaded by a command from the 

programmer. This load program command signal is continuously compared to the signal from 

the load cell, and if the difference is more than a fixed amount, the loading mechanism shuts 

down. To check the response times of this arrangement in the event of specimen failure, the 

loads were monitored on two adjacent specimens, the specimen paired to the failed specimen, 

and one specimen in the adjacent pair. The strain gage and recorder arrangement used had a 

combined response time of well under 0.01 second. At failure, the load on the specimen paired 

to the failed specimen dropped to zero in 0.01 second. A compressive load of 10% of the tensile 

load was recorded for a duration of 0.01 second. The adjacent specimen lost about 20% of its 

tensile load for 0.02 second and recovered to 95% of the original load, which it maintained until 

the load was dumped at 0.18 second; the load then decayed linearly to zero in about 1 second. 

12.2.1 1 AUTOMATIC OPERATION. To do long-term testing at the lowest cost and to be able to 

complete 50,000 hours testing in the shortest total time, the simulator must be automatic and 

be capable of operating unattended a high percentage of the time. During the Phase I 10,000 

hour tests the simulator operated on the average of 125 hours per week or about 75% of the 

.time available. The monitoring and safety features were largely responsible for the off time. 

They were, however, very important in order to avoid destroying or damaging specimens during flight simulation testing. There were a number of other factors that have restricted the 

average run time to 125 hours per week. When equipment of this type is operated year after 

year, periodic maintenance must be performed. Recorders, heat lamps, hydraulic cylinders, 

bearings and electronic components must be serviced or replaced. The equipment must be shut 

down to remove broken specimens. When shutdown occurs in early evening or during the 

weekend, additional time is lost because the resin matrix specimens must be dried before operation is resumed. 

During the Phase I1 tests, which consisted of almost 40,000 hours of exposure, simulation 

operation dropped to an average of about 100 hours per week. The Phase I1 flight simulation 

required almost seven years to complete. Reasons for the decrease in operating efficiency were 

numerous but were generally related to the many years of almost continuous operation of the 

apparatus. 

12-1 1 
One of the major problems was frequent parity errors in the programming system with subsequent shutdown of the test apparatus. The problem was finaUy resolved with a parity error 

resetting meter after many months of intermittent operation. Another lengthy delay was 

caused by a faulty tape reader. In-house repair was not successful so that the device had to be 

returned to the manufacturer for refurbishing. Similar problems were experienced with the 

Metrascope, the primary monitor and limit detect safety system, and estensive repair and 

replacement of several components were required. During the latter portion of the test program the flight simulation apparatus required considerable maintenance, primarily because of 

numerous hydrauiic leaks. Other problems that affected the operation of the flight simulation 

apparatus and contributed to the reduction in the weekly average run times are listed below: 

Programmer problems with the tape drive, tape reader, bearings, lamps, etc. 

Temperature control and faulty overtemperature indications 

Load control problems 

Hydraulic pump or line failures and oil leaks 

System oil spills 

Hydraulic pumps overhaul 

Servoblock repeated failures 

Power supply failure (28 Vdc) 

Controller circuit failure problems 

Power failures 

Servo amplifier problems 

Various electronic and mechanical component failures 

Doubler bond failures (primarily BIAI) 

Specimen failures 

Sys tem calibration and specimen inspections 

High priority mainline contract tests 

Flood damage in programmer room during heavy rains 

Cooling system motor failure 

12.3 SPECIMEN CONFIGURATION 

The unnotched specimen configuration for the long-term tests is shown in Figure 12-9. The 

composite portion of the specimen was 0.076 m (3 in.) by 0.46 m (18 in.). The tapered doublers 

bonded at each end of the specimen extended the overall length to 0.56 m (22 in.). Aluminum 

spacers of the same thickness as the composite were used to fill the gaps between the titanium 

doublers in the regions where the pin holes were located. The specimens were pin loaded to ensure uniform loading for the whiffletree arrangement used. Half of the resin matrix long-term 

specimens were prepared with notches consisting of 0.0064 m (0.25 in.) diameter holes. As 

shown in Figure 12-10, these holes were located so that they were centered in each of the six 

0.025 m (1 in.) by 0.23 m (9 in.) post-test specimens to be cut from the notched long-term 

12-12 
9) 

--?- 

a00318 (am MIN. 

LOADING HOLES ON CENTER 

LINE OF SPECIMEN TO WITHIN 

~a000064m (0.0025 IN.) 

- BONDED DOUBLERS4 PLACES 

DIMENSIONS IN 

rn (IN.) 

65321 7-209 

Figure 12-9 long-Term Flight Simulation Specimen 

12-13 
+ 

I 

0 

0 

a 

en 

u 

E 

s 

3 

I 5 I M 

u 

C 

Y 

0 

Y) 

E 

t 

c 

ta 

C 

0 

L 

L 

0 

C 

0 

& 

- 

12-14 
specimens after flight simulation exposure. These specimens were used to determine the 

residual notched tensile.or fatigue strength of the composite after flight simulation. The unnotched long-term specimens were also cut into smaller pieces for residual strength testing. 

The on@ plan was to cut each long-term specimen into twelve 0.013 m (0.5 in.) by 0.23 m (9 

in.) post-test specimens. As discussed later, this plan was slightly modified as a result of the 

condition of the specimens at the completion of 10,000 hours of esposure. All of the post-test 

tensile and fatigue specimens used bonded doublers as shown at the right in Figure 12-10. Two 

sets of quartz heat lamps 0.23 m (9 in.) apart were positioned so that they heated a 0.05 m (2 in.) 

strip on each half of the long-term specimen (see Figure 12-6). This arrangement placed the 

0.05 m (2 in.) heated zones at the centers of the post-test specimens. The edges of the heated 

zones were 0.025 m (1 in.) from the doublers so that the adhesive was not subjected to the 

operating temperature of the composite during simulation. Temperature measurements taken 

during testing showed the titanium end doublers to be approximately 69 K (125" F) cooler than 

the 408 K (275X F) heated zones of the epoxy specimens and approximately 117 K (210' F) 

cooler than the 505 K (450 O F) heated zones of the G/PI and BlAl specimens. One modification 

was made to the BlAl specimens to increase the efficiency of heating during flight simulation 

exposure. The heated zones on the sides facing the quartz heat lamps were coated with flat 

black, heat resistant paint to reduce reflection from the aluminum surface. 

The requirements for the short-term flight simulation specimens were less restrictive than for 

the long-term specimens because only residual tensile strength properties were to be obtained 

after the 100 and 200 hours of esposure. With this in mind a short-term specimen was designed 

that could be tensile tested after flight simulation with no further modifications, Le., machining or doubler bonding. The specimen that was used is shown in Figure 12-11. A further advantage in using the 0.025 m (1 in.) wide specimen rather than the 0.076 m (3 in.) wide long-term 

configuration was a considerable savings in material. To satisfy the pin loading requirement of 

the flight simulation apparatus it was necessary to use the 0.05 m (2 in.) wide end doublers 

shown in Figure 12-11. The space between each set of doublers not occupied by the end of the 

composite material was filled with aluminum sheet of the same thickness as the composite. 

Also shown in Figure 12-11 is the notched specimen used for half of the resin matrix short-term 

tests. The 0.0064 m (0.25 in.) diameter holes, one in each half, are centered in the 0.05 m (2 in.) 

heated zones in like manner to the long-term specimens. 

12.4 RANDOM LOAD SPECTRUM 

The objective of the SCR spectrum generation task was to provide a load simulation that 

would be representative of flight conditions during the life of a supersonic cruise aircraft and 

would be acceptable and remain acceptable to all concerned parties during this long real-time 

flight simulation test program. Developments (ref. 43 and 44) in random flight-by-flight load 

simulation have produced a simple method for generating a random load sequence that 

matches the desired load-exceedance relationships on a mission segment basis. These studies 

have shown the importance of including the random character of the individual load, adhering 

to flight-by-flight profiles, and including the ground-air-ground load for each flight. 

Several assumptions were made to define the loading spectrum. The mission profile and loads 

data were obtained from reference 4. These data were also used for the study described in 

reference 3. The life of a supersonic cruise aircraft was assumed to consist of 25,000 flights of 

two hours duration each. For this work, each flight was assumed to be an "operational flight" 

as defined by reference 4. Each flight consisted of a 10-minute climb, a 90-minute cruise. a 

12-15 
PO5' '"? 

r- 61 + -1 

1 I 

I I 

1 I 

I 

1 

I 

I 

I 

-* 

I 

I 

I 

I I 

I I I 

+ 

4.5) 

0.038 

0.025 

I I 

I 

I I 

I I 

I I 

I I 

I 

I I 

I 

r-- 1 A 

a 114 (4.5) 

(1)- IC. 

W 

- 

LOAOING HOLES ON CENTER 

u 

LINE OF SPECIMEN TO WITHIN 

5 0.000064m (0.0025 IN.) 

M 

(a) UNNOTCHED DIMENSIONS IN m (IN.) (b.1 NOTCHED 65321 7-21 1 

Figuro 12-1 1 Short-Tom Flight Simulation Specimonr 

12-16 
20-minute descent, and a single landing load. The number of loads per flight and the number of 

loads in each of the segments were modeled after the reference 4 load sequence. However, the 

loads were randomized such that all the loads expected to occur in 10,000 flights were included. 

Since the 90-minute cruise was scheduled to include only about eight loads per flight, the mean 

load was held for 44 minutes before and after the imposition of a two-minute cruise loading 

period. The climb and descent loads were distributed over the entire climb and descent time. 

To simulate the mission thermal profile for a Mach 3 cruise, an elevated temperature was applied to the specimens. From room temperature at the beginning of climb, the temperature was 

increased uniformly to the maximum use temperature for each material Once the random load 

spectrum was generated, the magnitude of the loads and the maximum use temperature or 

cruise temperature was adjusted to suit each of the composite systems being studied. Both the 

load magnitudes and cruise temperatures were set using data from the baseline measurements 

and test data obtained by using the load spectrum at an accelerated rate. The loading rate was 

increased so that the loads nody occurring during a two-hour flight would occur in one 

minute. These accelerated tests were run at the cruise temperature and are referred to as shortterm tests. 

12.4.1 SPECTRUM GENERATlON FOR THE LONG-TERM TESTS. The segment esceedance curves 

from reference 4 are shown in Figures 12-12, 12-13, and 12-14. The Gaussian process line 

needed for the random simulation is shown in each figure. The mapping function data are 

shown in Table 12-1. 

The No values for each segment were arrived at by the following argument. The long-term test 

simulation of reference 4 considered 10,000 flights. In this simulation it was desired to maintain the number of load values at 80 per flight. Since 20 dwell signals were required per flight 

to maintain the mean mise load during heating and one landing load was included in each 

flight, there were 101 loads per flight for a total of 1.01 million loads for the 10,000 flights. For 

this work these 1.01 million loads are applied in 25,000 flights, thus requiring about 40 loads 

per flight. Since there are actually only 32 minutes of loading per two-hour flight, the overall 

cyclic frequency, N, of load values is calculated as: 

N, = 40 cycles132 minutes = 1.25 cpm 

= 75,000 cyclesll000 hours 

The average number of mean crossings per unit time, N, is defined as the product of the irregularity factor (R = 0.9) and N,, giving: 

No = (0.9)NP = 67,500 cycles11000 hours 

The 10-minute climb, 90-minute cruise, and 20-minute descent se,ments represent 31%, 

6.25%, and 62.5% of the total loading time, respectively. For this program spectrum, the 

climb, cruise, and descent segments contain 70%, 12%, and 18% of the total number of loads. 

The No values for each segment were calculated as: 

No (Climb) = No (0.7M0.31) = 152,000 cydkseg hr 

No (Cruise) = No (0.12)/(0.0625) = 129,600 cydkseg hr 

No (Descent) = No (0.18)/(0.625) = 19,440 cydkseg hr 

12-17 
vj 

a 

X 

. , 

w 

0 

0 

0.2 0.4 0.6 0.8 1.0 1.2 1.4 

653217-212 e2 (LOAOP 

Aguto 12-12 Climb Exceodanco Curvo 

These values along with the other mission parameters were input to the General Dynamics 

Simulated History of Aircraft Missions (SHAM) computer procedure, which generated the 

load history (ref. 44). A typical mission is shown in Figure 12-15. 

12.4.2 SPECTRUM FOR THE SHORT-TERM TESTS. To set the magnitudes of the long-term loads for 

each composite system some accelerated tests were performed. This was accomplished by increasing the frequency of loading to the point where a half of a lifetime of loads could be applied to specimens of each composite system in 200 hours. These groups of specimens were 

held at constant temperatures. The temperatures chosen were their maximum use 

temperatures. These tests, which were run for 100 hours and 200 hours. are referred to as the 

short-term tests. The objective of the short-term tests was to apply as many mission loads as 

12-18 
100,000.0 

iomao 

1,000.0 

100.0 

N, = 129,600 CYC/k SEG. HRS. 

0.1 0.2 0.3 0.4 0.5 0.6 

02 (LOAD 12 . Figun 12-13 Cruiro Excoedance Curvo 

653217-213 

possible to the specimens in a short time. The random character of the loads, the fight-byflight sequence, and the landing load were all maintained in the accelerated spectrum. 

It can be shown that the minimum periods of the long-term test climb, cruise, and descent segments are 8.5, 10, and 66.25 sedload, respectively. To obtain a maximum cyclic rate of 50 cpm, 

the minimum period for each segment should be 0.6 sec/load. The acceleration factors were 

thus calculated as: 

(AF) climb = 8.510.6 = 14.167 

(AF) cruise = 1010.6 = 16.67 

(AF) descent = 66.2510.6 = 110.417 

12-19 
100.000.0 

1o.ooao 

1,ooao 

ioao 

10.0 

1.0 

ai a2 0.3 a4 a5 0.6 

e2 (LOAD I 2 

Flguro 12-14 Descent Exceedance Cunre 

653217-214 

These acceleration factors and the long-term test mission data were input to the SHAM proce 

dure to produce the short-term test history. A flight lasts about 60 seconds and is divided into 

approximately 25 climb loads in 42 seconds, six cruise loads in seven seconds, and seven descent loads and one landing load in 11 seconds. A typical load sequence is shown in Figure 

12-16. The accelerated spectrum results in an overall rate of one minute per flight, so that in 

100 hours of testing, about one fourth of a cruise aircraft lifetime (6000/25,000) can be imposed. 

12.5 WEAROUT MODEL 

At the time these tests were started a wearout model had been developed and used on epoxy 

matrix composite systems. The fatigue and residual strength data from the short-term service 

12-20 
Tablo 12-1. Mapping Function Data 

Gaussian Real N 

Load load 0 

Crossing/ 

(k rogmont hr) 

152,000 

100,000 

10,000 

1,000 

100 

10 

5 

2 

1 

130,000 

100,000 

10,000 

1,000 

100 

10 

1 

19,000 

8,500 

1,000 

100 

10 

3 

1 

0.5 

0 

0.04 

.27 

.50 

.73 

.96 

1.03 

1.13 

1.2 

0 

0.01 

.09 

.165 

.245 

.32 

.4 

0 

.035 

.12 

.215 

.305 

.355 

.4 

.43 

Climb Segment 

0 0 

0.04 0.2 

.095 .520 

.205 .707 

.375 354 

.655 .980 

.79 1.015 

1.05 1.063 

1.43 1.095 

Cruise Segment 

0 0 

0.01 0.1 

.045 .3 

.130 .406 

.225 .495 

.325 .576 

.425 .632 

Descent Segment 

0 0 

.035 .187 

.06 .346 

.12 .464 

.18 .562 

.24 396 

.34 .632 

.44 .656 

0 

0.2 

.308 

.453 

.612 

309 

389 

1.025 

1.196 

0 

0.1 

.212 

.361 

,474 

.570 

.652 

0 

.187 

.245 

.346 

.424 

.490 

.583 

.663 

0.224 

.130 

.142 

2.53 

2.16 

0.324 

- 

history tests on two of the four composite systems, Le., GIE and B/E specimens, were analyzed 

using the wearout model developed in references 6 and 45 and previously used in references 46 

and 47. 

The wearout or fatigue model is reviewed here and examples of its use will be given for the 

epoxy system in the section on short-term test results. This model is based on four assumptions (three physical and one mathematical form). 

a. Materials fail because of the presence of preexisting flaws. 

b. Flaws grow in a deterministic manner. The manner is determined by material properties, 

state, and magnitude of the stresses at the flaw perimeter; history of applied stress: and 

the thermal and mechanical history. 

c. The critical load for a structure is a function of instantaneous f&~w state, and the distribution of residual strengths defines the flaw field statistics. 

12-21 
v) 

2 

>. 

k 

a 

a 

a 

a 

k 

m 

a 

a 

v)- 

a 

0 

0 

-1 

3 

a 

c 

U 

w 

v) 

-70 

-a0 

-90 

-100 

100 

90 

a0 

70 

60 

50 

40 

30 

20 

10 

I 

. 

J 

6532 1 7-2 15 

Figuro 12-15 Typical Mission for Long-Term Testing 

12-22 
cn 

k 

z 

a 

> 

a 

< 

a 

k 

a 

a 

VJeo 

0 

0 

-L 

a 

3 

a 

Io 

w 

0 

VJ 

90 

80 

70 

60 

50 

.) 

, 

. 

65321 7-216 

10 

0. 

-1 0 

-20 

-30 

Figure 12-16 Typical Mission for Short-Term Testing 

. 4 

, 

12-23 
d. Damage rate accumulation is of the functional fom 

=M C', ac 

at 

- 

where C is the flaw length and r is the flaw growth rate exponent. 

The residual strength function FR(Q can be shown to have the form 

1 

where: 

A, = FdF,, with F, being the truncation load a measure of rms stress and F, being 

a reference rms stress level. 

The term A, is related to the historydependent constant (M) in Equation (1). 

The term r is the exponent of the flaw state (C) in Equation (1). 

The term -Us is the slope of the In F--h t curve and is related to r. 

The term t refers to time and to is a reference time condition. 

If the initial assumptions are correct, the results of varying stress history, environmental conditions, and geometry are contained in A,. 

Note that Equation (2) is a strictly monotonic decreasing function of time. Therefore, time t 

must be measured after the initial stress relaxation occurs. If this point in time is considered to 

be the time origin, t = 0, Equation (2) may be considered applicable for all positive values of t. 

F (t = 0) is the initial static strength of the component possessing a Weibull (extreme value) 

distribution: 

where (yo and 0, are the shape and scale parameters, respectively, for the Weibull distribution 

Introducing the Weibull flaw distribution, Equation (3). into Equation (2) gives the timevarying residual strength distribution of the fom 

This equation is represented functionally as 

f (P, F,)= A,t 

Therefore, the constant A, is the slope of the linear regression equation fitted to the empirical 

data. The iteration sequence is begun with an assumed value of "r" and repeated until the error 

term, i.e., the probabilistic squared difference between the model and the observed data, 

becomes a minimum. 

12-24 
for FR 

and 

P (FR(t) = 0) = 1 - P(F,(t) > F-) 

The probabilistic Equation (4) does not involve the reference time to, which appears in Equation (2), and agrees with Equation (3) when t = 0. For fixed time t, it follows from Equation (4) 

that the transformed variable FR(t)B(r-l) has a truncated threeparameter Weibull distribution. 

The distribution depicted in Equation (4) has a limiting form for (F&B$(r-l)<< 1, which gives 

the lifetime distribution as: 

af 

P(T > t) = P [ FR(t) > F-] = exp - (t/&) 

where: 

If specimen behavior is found to be consistent with the model, the following responses would 

be expected. 

a. The lifetime shape parameter would be independent of rms stress, temperature, and 

humidity. 

b. Both the truncation load-endurance data and initial shape parameter to fatigue shape 

parameter relationship will yield consistent values of r. 

c. Characteristic strengths and endurances will shift log-linearly. 

The program used to fit the wearout model to the experimental data sample is based on 

minimizing the “error” between the prediction (model) and the data. The model parameters a0, 

8, and to and the experimental data are available as “known” values for input. The program 

uses an iteration procedure to minimize the error (variance). The iterative soiution may be 

accomplished by determining either the value of “r” or of “A,” that gives a minimum error. 

12-25 
When the value of r is to be used in iteration, a value of A, must first be determined. The probabilistic model may be written in the form: 

where: 

and 

A? = 1 

The application of this model to the short-term test data is shown in Section 12.6. That section 

will clarify the determination of parameters A,, s, and AI. 

a, P - Weibull shape and scale parameters for the distribution of the tensile 

data. 

010 ' Po - Weibull parameters at time zero. 

ae Bf 2(r-1) ( r - l)A 4A f 

010 ' p2,k-11 

- -' 

CYTv PT - Weibull parameters at a temperature, T. 

PRT - Room temperature value. 

- Calculated using pooled CY by formula 

1 PMLE 

n 

P= [ El x'] 

n 

where n is number of specimens. 

PLT 

U 

e 

aT 

- 

- 

13 for long term test. 

Relates to slope of Gaussian process line. Slope = - - 1 

See Figure 12-13. 2 02 

- 

- 

Load for exceedance curves expressed in number of g's. 

-9 PT ratio of 0 at temperature to P at room temperature. 

PRT 

12-26 
A1 

A4 

C 

f c 

fr 

f T 

S 

t 

T 

where F,, is the truncation load, a measure of rms stress, 

and F, is a reference mu stress level, 

This term is related to the history dependent constant (M) in Equation (1). 

Flaw length. 

1 

(3)- Q!o is the penalty for least of three as a function of the shape 

parameter, aq 

1 

(100)- Q! is the risk factor obtained from failure statistics by assuming a 1% static failure rate is acceptable. 

- 

- 

Strength retention factor at elevated temperature obtained from the 

In 9 versus T curve. 

Static strength 

Residual strength function. 

Acceleration due to gravity. 

Proportionality constant in Equation (1). 

R N, is the average number of mean crossings per unit time. 

Overall cyclic frequency of loads. 

Probability function. 

Flaw growth rate exponent. 

Irregularity factor. 

The term - s is the slope of the In 

related to r. 

Time. to is a reference time condition. 

Temperature. 

12.6 SHORT-TERM FLIGHT SlMULATlON 

F,, versus In t' curve and is 

The short-term tests were intended to provide the data necessary to set reasonable load levels 

for the long-term tests. These load levels were extremely important to the success of the longterm tests. A selection that was too high could result in failure of all the specimens prior to 

completion of 50,000 hours of exposure. Conversely, a very low load level might result in 

neither specimen failures nor measureable changes in residual strength after 50,000 hours. The 

12-27 
short-term tests were accomplished by accelerating the loading rate such that half a lifetime of 

loads was applied in 200 hours. During loading the specimens were exposed to the maximum 

use temperature for each composite materid system. The loads for the short-term tests were 

set to induce some fatigue failures. Specimens that did not fail were tested for residual 

strength. These fatigue life and residual strength data were then used with the static baseline 

data to fit the wearout model and set the loads for the long-term tests. 

The short-term tests were additionally intended to check out the flight simulation apparatus 

and to provide the modes and statistics of failure. 

12.6.1 TEST PLAN AND PROCEDURES. For each of the four Phase I material systems a total of 40 

specimens, in groups of 10, were subjected to the short-term flight simulation exposure tests. For 

B/Al, both unidirectional and [0 O k 45 OIs crossplied materials were tested while only [0 O & 45 OIs 

crossplied material was tested for the resin matrix systems. The unidirectional B/Al was included because its significantly higher transverse strength and transverse and shear modulii, 

than those of the resin systems, offer a much greater potential for structural applications. The 

origmally scheduled unidirectional resin matrix specimens were changed to a notched [Oo & 45 OIs 

crossply configuration. Tests were conducted for time periods of 100 and 200 hours. At the 

accelerated rate of loading that was used, this represented onefourth and onehalf lifetimes 

respectively. All tests on any one system were conducted at the same temperature and load 

IeveL 

Following the 100- and 200-hour exposures, all unfailed specimens were tested for residual tensile strength at the same temperature as used for the short-term tests. 

12.6.1.1 Tempwaiure Soldion. Selection of the short-term test temperatures was to be based 

on the results of the baseline tensile tests. It was planned to select a temperature in the region 

just below that where a distinct change in the slope of the strength versus temperature curve 

occurred. Generally, however, no distinct fall-off in strength was observed out to the maximum 

tensile test temperatures used for any of the materials. At the time that the short-term tests 

were ready to begin, the thermal aging exposures had not been running long enough to provide 

data for selection of the flight simulation temperatures. Data from the literature for use in 

selecting the maximum use temperatures for 50,000-hour service was ais0 lacking. However, at 

the time that the long-term tests were expected to start it was thought that sufficient thermal 

aging data would then be available and that a rational selection of exposure temperatures 

could be made at that time. Therefore, the temperatures selected for the short-term tests were 

in the region covered by the thermal aging tests. For the epoxy systems 422 K (300 O F) and for 

BlAl505 K (450" F) were chosen. These temperatures were midway between those at which 

thermal aging exposures were currently in progress. To obtain a better comparison between 

the BlAl and the HT-S/710 GlPI systems, 505 K (450 O F) was also chosen for GlPI even though 

the midway temperature of the thermal aging tests was 533 K (500' F). 

12.6.1.2 Selection of Short Term Loads. The loads for the short-term tests were set as high as 

possible to produce early fatigue failure and provide a basis for subsequent stress variations. 

The room temperature static ultimate strength is reduced by two independent factors to arrive 

at the desired load. One of these factors is strength retention at elevated temperature, fT. It 

was obtained from a retention cume as described below. The other factor is a risk factor, f, and 

was obtained from failure statistics by assuming that a 1% static failure rate is acceptable. 

12-28 
Because no baseline tests were performed on B/E, its reduction factors were obtained from 

existing data. The procedure for the reduction of the GIE, HT-SI710 GIPI, and B/Al data is 

described here. 

The baseline data provides a starting point for setting the peak loads to be used during the 

short-term tests. The raw failure load data for a given materid, notch condition, and temperature were fit with a Weibull distribution by linear regression to obtain the scale parameter, fl, 

and the shape parameter, a The temperature retention me was plotted in the form of In aT 

versus temperature, where aT is the ratio of /3 at a given temperature to fl at room temperature. 

The portion of this curve that exhibits linearity permits statistical pooling of the data over 

that temperature range. The pooling was accomplished by scaling each data point (load value) 

to its raw fl estimate and using this ratio as the random variable. The resulting CY was the pooled 

CY for that material and notch condition over the temperature range of interest. Using maximum 

likelihood estimation techniques, a better estimate of the fl for each temperature was obtained 

from the original load data and the pooled a These new 0 estimates were used to refine the In aT 

versus T me so that the temperature factor, fT, could be obtained directly from the plot. The 

risk factor, fr, was calculated from the pooled 01 as the first failure of 100 samples, or 

-1 - 

CY 

f, = (100) 

The permissible peak load was then the product: 

The procedure for load selection is demonstrated below using the baseline tensile data for 

unnotched [Oo f 45 'Is G/E in a typical calculation: 

1. Fit a Weibull distribution to each set of raw load data (a distribution for each temperature, 

material, and notch condition). 

2. Plot temperature retention in the form of a In aT versus T curve, where aT is PTIPRF 

3. From Step 2, determine which sets of data may be pooled (based on linear part of In aT versus T curve). 

4. Pool several temperatures by using Xl& as the random variable. Obtain a pooled CY for that 

material and notch condition. 

5. Recompute fl~ based on pooled 01 and original load data (use maximum likelihood 

estimator). Refine aT versus T with flwLE. 

6. Compute risk factor based on 1% failure rate fr = (100) 

-1 - 

CY 

7. Find temperature retention factor (fT) from refined aT versus T. 

Using the 35 tensile strength values in Table 12-2, the procedure is followed to calculate the 

peak short-term load for GIE. 

12-29 
Tabk 12-2. Baulino Tonsilo Data - Unnotchod [Oo f 45O], G/E 

BCO1-1 

-2 

-3 

BC01-6 

-7 

-a 

BCO1-1 

-16 

-22 

-23 

-24 

-25 

-26 

-27 

-28 

BC01-9 

-10 

-17 

BCO1-11 

-19 

BC01-12 

-13 

-20 

BC01-14 

-15 

-21 

-29 

-30 

-31 

-32 

-33 

-34 

-35 

-18 

218 -67 3870 

3879 

4359 

297 75 5031 

3905 

4377 

394 250 4257 

4101 

3914 

4871 

44448 

5782 

4804 

4848 

4670 

4724 

408 275 4492 

4097 

5026 

422 300 4919 

4270 

4715 

436 325 4657 

5249 

4404 

150 350 4008 

4884 

4670 

4715 

4635 

3941 

4270 

5293 

4670 

5338 

a70 

872 

980 

1131 

878 

984 

957 

922 

1095 

1000 

1300 

1090 

1050 

1062 

1010 

921 

1130 

1106 

960 

1060 

1047 

990 

901 

1098 

1050 

1060 

1042 

886 

960 

1190 

1050 

1200 

a80 

ioao 

iiao 

4034 907 4190 942 11.79 0.89 0.96 

.93 

1.04 

4444 999 4702 1057 7.40 1.0 1.07 

.93 

.a3 

4644 1044 4875 1096 9.61 1.04 .a7 

.a4 

.ao 

1.00 

.91 

1.19 

.99 

.99 

.96 

.97 

4537 1020 4750 1068 9.34 1.01 .95 

.a6 

1.06 

4635 1042 4790 1077 13.16 1.02 1.03 

.98 

.a9 

4662 104s 4786 1076 17.09 1.02 .97 

1.03 

.92 

4644 1044 4853 1091 10.49 1.03 .83 

1.01 

.96 . 

.97 

.96 

.a 1 

.aa 

1.09 

.96 

1-10 

4710 1059 1.01 

4715 1060 1.01 

4715 1060 1.01 

4862 1093 1.04 

1. The fl and CY columns of the table list the. value calculated from the data at each 

temperature. 

2. Figure 12-17 shows a plot of In aT versus temperature using the values listed in the aT = 

3. From Figure 12-17, it was determined to pool all the data. 

4. All temperatures were used and the data shown in the Loadlo column was ordered and used 

to find a pooled CY. The computer makes this calculation. However, for clarity the nature of 

this calculation is shown in Table 12-2 and the Weibull plot in Figure 12-18. The pooled CY 

for this case is 12.71. 

fl/pR~ COh~ 

12-30 
5. Recompute fl~ based on CY = 12.71 and original load data for each temperature. The maximum likelihood estimator formula is 

P 

P 

\ 

c a4 

J- a3 

a2 

I1 

0.1 

Results of this calculation are shown in Table 12-2 in the flMu and fl/flRT columns. Now use 

this data to refine In a~ versus T plot using PM~. These points leave Figure 12-17 essentially 

unchanged. 

I 

I I I I I I 

6. 

7. 

8. 

-- 1 

Risk factor based on 1% failure rate was then calculated as fr = (100) 12.71 = 0.70. 

The temperature retention factor was considered to be 1.0 for this case. 

The peak load is equal to: 

0.70 X 1.0 X [4684 N (1053 lb) ] = 3278 N (737 lb) 

Because the tensile specimens were 0.013 m (0.5 in.) wide, the peak load per inch is 6557 N 

(1474 lb). 

In a similar manner the baseline test data for the other composite systems were analyzed for 

both the notched and unnotched cases. The BlAl data were calculated for unidirectional and 

crossplied materials. The reduction factors and calculated peak loads are given in Table 12-3. 

12-31 
m, 

PI 

cy 0 

(L+ N/l) Nl Nl 

12-32 
Tablo 12-3. Summay of Rdudon Factors, loads, and Strerros for Short-Term Tests 

.rook ~aod %anding LW~ oak ~tms Landing ~tms -no g ~tms 

Material Condition f, ft N (Ib) N (Ib) MN/d (krl) MN/d (kri) MN/d (kri) .) - 

BIE Unnotched 0.79 0.80 6218 1398 

Notched .79 .80 3483 783 

GiE Unnotched .70 1.0 6557 1471 

Notched .68 1.0 3616 813 

BlAl [O0& .77 .91 25.3105690 

[Oo 2 45 '1, .71 32 8451 1900 

GtPI Unnotched 68 1.0 4946 1112 

Notched .68 1.0 2593 583 

Note: Stresses based on the following thicknesses: 

BIE. Unnotched: 0.00076 m 10.030 in.) 

BIE. Notched: .00074 m LO29 in.) 

GiE. Unnotched: .00076 m LO30 in.) 

GiE. Notched: .00076 m LO30 in.) 

GIPI. Notched: .00089 rn LO35 in.) 

BlAL [0 2 45 7,: .00097 m LO38 in.) 

GiPI, UMOtChed: 

BlAL [OO],: 

.00094 m LO36 h.) 

.00104 m LO41 hJ 

-1557 -350 

- 872 -196 

-1637 -368 

- 903 -203 

-6325 -1422 

-2117 - 476 

-1237 - 278 

- 649 - 146 

321 

248 

338 

249 

956 

345 

207 

141 

46.6 - 80 

36.0 - 62 

49.1 - 85 

36.1 - 62 

138.8 -239 

50.0 - 86 

30.1 - 52 

20.5 - 35 

- 11.7 

- 9.0 

- 12.3 

- 9.0 

-34.7 

- 12.5 

- 7.5 

- 5.1 

161 

124 

169 

125 

478 

173 

104 

71 

23.3 

18.0 

24.6 

18.1 

69.4 

25.0 

15.1 

10.3 

a Peak load - f,f.+3RT 

Landing load is defied as negative 25% of peak load 

One g stress is defined as 50% of peak stress 

12.6.1.3 Test Procedure. Before installing the short-term specimens in the flight simulation 

apparatus, width and thickness measurements were determined, and the average cross-sectional 

areas were calculated for each set of 10. These areas were used to obtain the flight simulation 

stresses and, for specimens that did not fail, to calculate the residual tensile strengths after 

post-exposure testing. The specimens were then installed in the test machine, thermocouples 

were attached, the compression stiffening grids were bolted in place, and all resin-matrix specimens were heated to 422 K (300" F) for 24 hours to remove any absorbed moisture. Before 

starting the tests, each set of 10 spechaens was proof loaded to the maximum load value that it 

would experience during the 100- or 200-hour. short-term exposure. 

Once started, the short-term testing of any one set continued uninterrupted until completion 

unless specimen failure or equipment malfunction occurred. If the test was halted for 24 hours 

or more, an additional moisture bakeout of four hours was performed on all except the B/Al 

specimens. 

At the completion of the 100- or 200-hour flight simulation exposures, the specimens were 

removed from the simulator and visually examined with a stereo microscope to check for 

changes in appearance or the presence of surface flaws such as cracks or delaminations. After 

the visual examination, the short-term specimens were tensile tested at the same temperatures 

at which they had been exposed. Testing was performed in a Tinius-Olson Universal testing 

machine at a head rate of 0.0013 m (0.05 in.) per minute. Heating of the specimens was accomplished with two ring furnaces that encircled the specimen at the two areas that had been 

12-33 
heated during the simulation exposures. This technique was used rather than an environmental chamber to prevent the possibility of premature failures in the doubler adhesive. The two 

ring furnaces were independently controlled by manually adjusting two power supplies. Temperature was monitored with thermocouples attached to the two test sections. The specimens 

were heated to the desired test temperature in approximately 10 minutes and then held at 

temperature for an additional 10 minutes before testing. Temperature was held to 2 3 K (5 O F) 

during both the hold period and while testing. Figure 12-19 shows a specimen in the testing 

machine with the two ring furnaces in place. If the resin matrix specimens were not tensile 

tested within 48 hours of completion of the flight simulation exposure, they were given a bakeout treatment at 394 K (250" F) for at least 24 hours to remove any absorbed moisture. 

12A1.4 Problem Anas. In addition to providing the data necessary to set the load levels for 

the long-term tests, the short-term test program was also intended to check out the loading 

facility and the suitability of specimen design and fabricatioa. Numerous deficiencies in the 

flight simulation apparatus were detected and corrective measures were taken. Some of the 

equipment problems included: malfunctioning of solid-state power control switches, voltage 

transients that interfered with the temperature control circuits, lack of proper safety valves in 

the hydraulic system, improperly positioned lamps and thermocouples, and intermittent shutdowns of the entire system caused by both electronic and mechanical difficulties &th the pre 

gramming systems. Because of the complexity of the flight simulation apparatus and the 

extremely long run times involved in the test program, it was not unexpected that equipment 

malfunctions and component wearouts would occur from time to time. 

One problem involving support equipment rather than the simulator was caused by movement 

of the compression stiffeners (steel grids bolted to the specimens to prevent buckling), see 

Figure 12-7, during fatigue cycling. These fixtures tended to rotate slightly and/or ride too low 

on the specimens, leaving an unsupported column of composite approximately 0.0025 m (0.1 

in.) long at the top. During the compressive loading portion of the cycle some buckling or bending could occur. It was most noticeable for the unidirectionai BlAl where the compressive loads 

were considerably greater than those for the other materials. Spacers were prepared that, 

when inserted between the lower grips and the stiffeners, supported the stiffener, assured that 

they were centered about the specimen, and greatly reduced the unsupported column of composite material. The spacers also reduced rotation of the stiffeners during fatigue loading. One 

other problem caused by the compression fixtures was abrasion of the soft BlAl specimens 

during fatigue cycling (a-form of fretting corrosion). The addition of the spacers reduced the 

relative movement between the stiffener and specimen and, thereby, eliminated much of the 

wearing. Further reduction in the abrasion of the aluminum surface of the BlAl specimens was 

accomplished by application of a stratified Teflon coating to the compression stiffeners. By the 

time these modifications were completed, however, considerable surface abrasion had occurred 

on the B/Al short-term specimens. 

A problem encountered early in the testing of the BlAl short-term specimens was premature 

failure at the edge of the end doublers caused by undercutting of the BlAl during removal of 

excess adhesive when fitting the compression stiffeners. The problem was eliminated in the 

long-term tests by modifying the fabrication procedures. Another problem was a bond failure 

at the titanium doublers of the crossply specimens. A solution to this problem was found by 

installing a Ti-6A1-4V rivet on each end of the specimens through the titanium doublers and 

the B/A1 material A more difficult problem was the attachment of end doublers to the 

12-34 
ORIGTNAL PAGE IS 

.OF POOR QUALITY 

Figure 12-19 Test Setup for Resldual Strongth Determinations After 1 OO-hr and 200-hr Short-Term 

Simulation Exposures 

12-35 
unidirectional specimens. For these specimens, the test loads were much higher than used for 

the crossply specimens with correspondingly higher shear stresses at the doublers. For 

unidirectional specimens, rivets did not solve the problem. Failures occurred in the doublers by 

shearout of the B/A1 behind the rivets. Other techniques that were considered were brazing, 

diffusion bonding, diffusion braze bonding, spot welding, and mechanical clamping. The 

method selected was a clamping device on each end of the BlAl unidirectional short-term 

specimens. The clamp was placed near the tip of the doublers in the tapered portion. To assure 

uniform contact between the clamps and the doublers, flats were milled on each side at the contact areas. The purpose of the clamps was to increase the normal force reacting on the 

specimen and to decrease the tendency of the doublers to peel during flight cycle loading. An 

overall view of a BlAl specimen is shown in Figure 12-20 with a closeup view of the tapered section of the doubler shown in Figure 12-21. This technique proved to be successful for the shortterm tests and was also used for the long-term testing of both the crossplied and unidirectional 

B/AL Figure 12-7 shows the clamps in use during the long-term tests. 

12.6.2 SHORT-TERM TEST RESULTS. Following the short-term flight simulation exposures, all 

specimens were removed from the test apparatus, examined visually, and then tensile tested 

for residual strength at the same temperature as was used for the flight simulation tests. The 

following sections describe the results of these examinations and tensile tests. 

12.6.2.1 Post Simulation Examinations 

WE, [OO -C 450],, unnotched, 100 hours. After 100 hours the major change in the specimens was 

a distinct discoloring in the two heated zones of each specimen. These color changes were very 

similar to those observed during the thermal aging tests at 450 K (350" F) and, hence, were 

considered normal for these exposure conditions. A slight amount of surface damage in the 

L 

- 

Figure 12-20 Overall View of B/AI Shor)-term Specimen with Flats Milled in Titanium Doublers for 

Clamps 

12-36 
4 

I 

613217.221 i 

Figure 12-21 Cloroup-View of Tapered Section of Titanium Doubler Showing Flat Area for Attaching 

Clamp 

areas where the compression stiffeners supported the specimens was visible. No fibers were 

damaged, however, and the effect on the residual tensile properties was considered to be 

negligible. 

WE, [OO f 45O],, unnotched, 200 hours. The added 100 hours of flight simulation exposure did 

not change the appearance of these specimens from that described above. 

B/E [OO -C &PI,, notched 100 hours and 200 hours. The overall appearance of the notched 

specimens was identical to the unnotched samples previously discussed. In the vicinity of the 

notches, 0.0064 m (0.25 in.) diameter holes in the center of the heated areas, no apparent 

damage was visible. In machining these notches, by drilling with a diamond tool., a very small 

amount of chipping occurred on the back side of each hole. Small portions of some of the boron 

filaments were broken off, and several of the filaments were exposed where the epoxy matrix 

had been removed. No degradation of either the exposed filaments or the matrix was visible 

after the 100- or 200-hour exposures. 

G/E, [OO f 450],, unnotched, 100 hours. Seven of the 10 specimens in this group survived the 

100 hours of exposure. Discoloration in the heated areas was not as noticeable as for the BIE 

specimens and, again, was quite similar to that found during thermal aging of GIE at 450 K 

(350' F). Only one of the seven specimens (BC84-7) experienced any visible damage. In this 

specimen, a small crack was found in the outer 0 O ply in one of the heated areas. This crack and 

delamination of the outer ply is shown in Figure 12-22, an edge view of the specimen. Also visible in the photo is a crack that passes through the second, third, and fourth plies and extends 

for several mm between the fourth and fifth plies. The effect of this flaw on the residual tensile 

strength, however, was minor (see Section 12.6.2.2, Table 12-6). 

12-37 
Three of the specimem in this set failed during the flight simulation exposure, one in 10.6 

hours and the other two both at 54.2 hours. The first, BC84-6, is shown in Figure 12-23. While 

three breaks occurred along the length of the specimen, it is thought that the failure initiated 

in the heated zone. Multiple fractures of this type are common to the!e long specimens. This is 

especially true when failed in a "soft" test machine with the sudden release of a large amount 

of stored energy. 

The second specimen (BC84-2.54.2 hours) also failed in one of the heated areas and was also a 

fatigue failure. This specimen was of interest in that a crack had also formed at the other 

heated area but did not propagate sufficiently to fail the specimen at that location. An edge 

view of this crack is shown in Figure 12-24. The crack, while not deep, has extended through all 

six plies. Some delamination of the outer layers is also evident. 

The third specimen (BC84-1,54.2 hours), which was in the same set of two as BC84-2, failed in 

the doubler. Although it has been established that no transfer of load from a failed specimen to 

an adjacent one occurs, the failure of BCS4-2 is thought to have been responsible in some 

unknown way for the apparently simultaneous failure of BC84-1. The time of failure and location of fracture make it improbable that this was a true random fatigue failure. Hence, this 

specimen was discounted in the data analysis. 

G/E [OO -C 450],, unnotched, 200 hours. These 10 specimens completed 200 hours with no 

failures. The specimens experienced color changes similar to the others of this system and 

minor abrasion from the compression stiffeners. No edge cracks were found except for some 

slight splintering of the outer plies in several of the specimens. These effects were small and 

were observed only at the edges. 

Figure 12-22 Edge Crack and Delamination in Specimen BC84-7, G/E, Unnotched, Aftor 100-hr Simulation 

Exposure at 422 K (300O F) 

' 

12-38 

ORIGINAL PAGE rs 

OF POOR QUALI'I% 
ORIGINAL PAbE 

OF POOR QUALITY, 

INITIAL FAILURE BELIEVED TO HAVE INITIATED 

IN THE HEATED ZONE (ARROW). 

65321 7-223 

Figure 12-23 Specimen BC84-6, G/E, Unnotched 

1 EDGE VlMl OF A CRACKED REGION. SPECIMEN 

EVENTUALLY FAILED IN ANOTHER AREA AFTER 

54.2 HOURS OF SIMULATION EXPOSURE AT 422 K (3OO0F). 

65321 7-224 SX 

-- 

Figure 12-24 Specimen BC84-2, G/E, Unnotched 

12-39 
G/E [OO f 450],, notched, 100 hours. These 10 specimens completed 100 hours with no failures. 

The specimens experienced color changes similar to the unnotched specimens and minor abrasion from the compression stiffeners. Minor cracking of the inner surface of several of the holes 

was visible. With the exception of one specimen, these cracks were very shallow and did not 

appear to be very significant. The exception was specimen BF84-8, which contained a crack in 

the outer one or two plies extending from one of the holes for approximately 0.0025 m (0.1 in.). 

Figure 12-25 shows the location and extent of the crack on the surface of the specimen while 

Figure 12-26 shows the crack extending into the inner surface of the hole. Smaller branch 

cracks are also visible in the two middle plies. An overall view of this specimen is shown in 

Figure 12-27. 

G/E, [OO f 450],, notched, 200 hours. These specimens, all of which completed 200 hours, were 

very similar in appearance to the previous group. The only difference was at the inner surface 

of the holes. After 200 hours, the number of fine cracks was increased over that observed for 

the 100 hour specimens. These cracks were still small and quite shallow, however. These 

cracks, which were found in about half of the specimens, did not seem to influence the residual 

tensile properties. 

B/Al, [OO f 450],, unnotched, 100 hours. Examination of these specimens with a stereo micre 

scope revealed the presence of numerous transverse cracks in the BlAI in the regions adjacent 

to the end doublers. These cracks were found in seven of the 10 specimens of this group. Figure 

12-28 is an enlargement of the flawed area in one of these specimens. All of the transverse 

cracks were located in regions that had been undercut during the process of removing excess 

adhesive that had flowed from under the titanium end doublers during bonding. It was 

necessary to remove this adhesive so that the compression fixtures would fit securely to the 

- 

Figuro 12-25 Specimen BF84-8, G/E, Notchod, After 100-hr Simulation Exposure at 422 K (3000 F) 

12-40 
Figun 12-26 VIow of Main Crack and Sovoral Branch Cracks at Edgo of Hole In Spocimon BFU-8, G/E, 

Notchod, Aftor 100-hr Simulation Exposure at 422 K (300O F) 

I 

CRACK SHOWN IN FIGURES 12- 25 AND 12- 26 

LOCATED AT ARROW. 

65321 7-227 .a 

Figuro 12-27 Overall View of Specimen BF84-8, G/E, Notched After 100-hr Simulation Exposure at 422 K 

(300O F) 

12-41 
test section between the doublers. The combination of a very hard adhesive and soft annealed 

aluminum led to the problem. Also, only about 0.00001 m (0.0004 in.) of aluminum covers the 

first layer of boron filaments, so that very little abrasion was required before exposing the 

boron filaments. The consequences of the undercutting of the specimens with the attendant 

transverse cracking were twofold. Examination of the specimen (EC84-9) that broke after 65.4 

hours of exposure suggests that failure was premature and was caused by the growth of these 

transverse cracks in an abraded region. This is readily apparent from a comparison of the failed 

edge of specimen EC84-9 (Figure 12-29) with the transverse cracks shown in Figure 12-28. The 

second consequence of these transverse cracks was that residual strength testing of the exposed 

specimens in their present condition would probably be of little value because of the likelihood of 

premature failure of at least seven of the samples (those that were cracked). Therefore, rather 

than test the full size specimens, subsize coupons 0.013 m (0.5 in.) by 0.15 m (6 in.) were cut 

from an area adjacent to the end doublers, one from each short-term specimen. Figure 12-30 is 

a sketch showing the details of how the subsize specimens were prepared. Note that half of the 

heated test section is covered by one of the doublers on the subsize specimen. Cutting of the 

residual strength coupons in this manner had the advantage that they could be taken from 

undamaged areas of the short-term specimens. Several specimens showed surface damage 

where the compression stiffeners had been clamped. Again it was the low abrasion resistance 

of the soft aluminum that contributed to this surface damage. 

B/Al, [OO =t 450],, unnoiched, 200 hours. During the first 100 hours of exposure, doubler bond 

failures were experienced on five of the specimens. At the end of 100 hours, all specimens were 

removed from the apparatus, new doublers were attached where bond failures had occurred, 

and rivets were installed in the doublers of all 10 specimens. No more bond failures were 

observed during the remainder of the test. Four specimens did fail in the BIAl, however, before 

completing 200 hours of exposure. All four failed at the same time (one with 105 hours of test 

TRANSVERSE * 

Figure 12-28 Specimen ECW-4, B/Al, Crossply, After 100-hr Simulation Exposure at 505 K (GOOF) 

12-42 
Figure 12-29 Specimen EC84-9, B/Al, Crossply, Which Failed After 65.4 hr of Simulation Exposure at 505 K 

(450O F) 

time and three with 136 hours). Three failed in the heated zones, and the fourth specimen failed 

about 0.008 m (0.3 in) from a doubler. Failure undoubtedly occurred during a very high load 

portion of the random load spectrum, but the probability of four true fatigue failures occurring 

simultaneously is quite small. All four specimens had abraded regions at the site of the fractures. Of these, three appeared to have been severe enough to possibly have initiated the fracture. Figure 12-31 shows one of these specimens. Fracture occurred 45" to the length of the 

specimen except in the region of the abrasion. Test coupons were cut from the heated areas of 

these specimens at the ends away from the fractures. 

The remaining six specimens completed 200 hours without failures. Examination of these 

specimens with a stereo microscope revealed the wear marks caused by the relative motion of 

the specimens and compression stiffeners. This is a type of fretting corrosion that in aluminum 

is generally accompanied by a black powdery residue of aluminum oxide. The effect is clearly 

visible in Figure 12-31. Also noticeable on the specimens were numerous surface cracks located 

in the heated zones. An example of these cracks is shown in Figure 12-32. In all cases, these 

longitudinal cracks extend from regions where one or more of the outer layer boron fibers have 

fractured. Although not observed during the original examination, a recheck of the 100-hour, 

short-term specimens also revealed cracks of this type. For the 10 specimens, however, only 

two very small cracks were found. Apparently, the approximate time period required to form 

these cracks was between 100 and 200 hours. Similar surface cracking was observed during 

constant amplitude fatigue testing of B/Al at 561 K (550" F). The surface roughness seen on 

the fatigue specimens was not present on the short term specimens, however. For the few constant amplitude fatigue tests performed at 505 K (450" F) (same temperature as used for shortterm tests), surface roughness was also not observed. 

' 

12-43 
I 

I 

I 

I 

I 

.I 

I4 

w :I r 

2 

2 

z 

w 

a 

r 

0 

v) 

2 

ul 

u 

w 

0. 

v) 

w 

II 

0 

c .- 

VI 

e 

0 

P 

a 

0 

U 

f 

E 

5 

a8 

c 

- 

0 

a 

3 

H 

Y 

w 

Iw 

v) 

3 

v) 

N 

m 

12-44 
3x 653217-231 

Flgure 12-31 Enlarged VIew of Failed End of Specimen EC82-7, B/Al, Crossply, Which Failed After 135.8 

hr of Simulation Exposure at 505 K (4500 F) 

8X 653217-232 

Figuro 12-32 Enlarged View of Surface Cracks Found in Heated Zones of B/Al Crossply Specimens After 

200 hr of Simulation Exposure at 505 K (4500 F), Specimen EC82-3 

12-45 
WAl, [0016, unnotched, 100 and 200 hours. Considerable difficulty was experienced during shortterm testing of the unidirectional BlAl specimens. The major problem involved shear failures 

in the doubler bonds. These failures generally occurred at the upper bonds where test temperatures reached slightly higher values (406 K (270 ' F) maximum versus 372 K (210 F) maximum 

at the lower bond). The doubler bond failures were the result of the considerably higher loads 

required for the unidirectional material than were required for the crossplied BlAl specimens. 

These high loads, in combination with several other factors, also led to numerous failures in the 

BlAL These included: 

.) 

a. Failures in the BlAl very close to the edge of the doublers. At the bottom end, these failures 

were initiated by the stiffening grids gouging into the soft aluminum during the compression load. Additional failures occurred at regions that had been undercut during the process of removing excess adhesive that had flowed from under the titanium end doublers during bonding. These failures occurred during the high tensile loads. 

b. The compression stiffening grids could ride too low on the specimens, leaving an unsupported column of BlAl approximately 0.0025 m (0.1 in) long. During the compressive load, 

some buckling could occur. In severe cases, longitudinal splitting and/or failure occurred. 

c. Specimens that were riveted to increase doubler reliability all failed in the doublers by 

shearout of the B/Al behind the rivets. 

d. Abrasion of the soft aluminum surface of the specimens with eventual breakage of the 

underlying boron fibers was caused by movement of the stiffening grids during fatigue 

cycling. These damaged regions could ultimately initiate failure. 

One of the objectives of the short-term tests was to uncover these testing problems and devise 

solutions prior to the start of the 50,000-hour, long-term test program. The mechanical clamping technique discussed earlier provided a very noticeable improvement in doubler bond 

reliability. Spacers inserted between the lower grips and compression stiffeners greatly 

decreased the unsupported column of BlAl and in turn reduced the incidence of compressive 

buckling failures. The combination of these spacers and the Teflon coating on the stiffeners 

significantly reduced the fretting fatigue abrasion of the soft aluminum surface of the B/AI 

specimens. 

Because of these testing difficulties, the number of specimens that survived the 100- and 

200-hour short-term exposures was less than originally planned. Nine specimens with 100 

hours and six with 200 hours were available for residual strength testing. Several of the 

specimens were close to failure at one end as a result of small buckles. Many specimens also 

exhibited damage in the form of wear marks from stiffener abrasion and surface cracks and 

edge delaminations. It was possible, however, to cut subsize coupons, 0.013 m (0.5 in.) by 0.15 

m (6 in.), from good areas of each of the short-term specimens as shown previously in Figure 

12-30. The surface cracks were much less pronounced in the unidirectional specimens than 

observed in the 200-hour crossply specimens. Again no surface roughening similar to that 

found after fatigue testing at 516 K (550" F) was seen on the unidirectional short-term 

specimens. 

12-46 
HT-SnlO G/PI, [OO f 450],, unnotched, 100 hours. In like manner to the unidirectional BIAl, the 

number of unnotched G/PI short-term specimens was somewhat less than originaUy planned. 

The reasons for this were twofold: 

a 

b. 

During proof loading of the G/PI system, nine of the unnotched specimens failed. These 

specimens, however, failed in a manner similar in appearance to the baseline tensile 

specimens. The reason for the premature failures was thought to be either: 1) an error in the 

applied load, or 2) lower than expected material strength The first reason seemed unlikely 

as three independent strain measuring systems were used to monitor the loads. However, 

as a further check, 10 strain gaged dummy specimens of aluminum were installed in the 

flight simulation apparatus. Load checks in the region of interest indicated that the total 

applied load and the 10 individual loads were in agreement for all the measuring systems. 

To check the tensile properties of the panel (LRC-59) from which the short-term specimens 

were cut, two specimens were obtained from the small piece remaining from the paneL 

Several additional tensiles were also cut from the failed short-term specimens. These were 

then tensile tested at 505 K (450" F) using both a nod loading rate, 0.0013 m (0.05 in.) 

per minute, and a rate similar to that used during the proof test. The results, which were 

independent of load rate, showed the short-term panel to have an average tensile strength 

about 69 MN/m2 (10 ksi) lower than the panel from which the baseline data were obtained. 

This same 69 MN/m2 (10 ksi) lower value was found for the panel (LRCBO) that supplied 

the notched short-term specimens. Using the data from panels LRC-59 and 60 and that 

from additional specimens cut from the short-term specimens that had failed during the 

proof test, revised short-term loads were determined and testing was begun. However, 

because of the loss of the specimens during the first proof test, only seven specimens each 

were available for the 100-hour and 200-hour unnotched short-term exposures. No notched 

short-term specimens had failed so that full sets of 10 were available for the 100-hour and 

200-hour time periods. 

During the short-term testing of the G/PI specimens, almost all the 200-hour unnotched 

specimens were failed during a compressive overload. The accident occurred as the result of 

an operator error when an incorrect hydraulic valve was opened. Testing was interrupted 

while seven additional specimens were fabricated from spare material. Again a full set of 

specimens could not be prepared because of a lack of extra material Just before completing 

the short-term tests, another compressive overload occurred. This incident resulted in four 

additional specimen failures, three unnotched and one notched. The cause of the overload 

was traced to the load programmer. The malfunction, which could have occurred later during long-term testing, was corrected. Again, one objective of the short-term test program 

was to uncover the problems and equipment deficiencies that, if not found, could be very 

costly during the long-term test program. 

The five unnotched specimens of this group completed the 100 hours of exposure with almost 

no change in color or surface texture. There was a slight amount of surface damage on four of 

the specimens from the stiffening grids, but this was not believed to have had any effect on 

residual tensile properties. All of the specimens showed slight edge cracking, generally in the 

heated regions or near the doublers. A typical example can be seen in Figure 12-33. Again, 

these edge cracks probably did not affect the residual tensile strength, although in a compression or shear test they likely would have been important. 

12-47 
HTW710 G/PI, [OO -C 4501,. unnotched, 200 hours. The added 100 hours of flight simulation 

expoposure did not change the appearance of these specimens from that described above. 

HT-S/7lO G/PI, [OO =t 450],, notched, 100 hours and 200 hours. The overall appearance of the 

notched specimens was identical to the unnotched specimens previously discussed. Slight surface damage from the stiffening grids and shallow edge cracks were visible on most of the 

specimens. Figure 12-34 shows an edge view of one of the edge cracked specimens. During 

residual strength testing, failure occurred at the hole in this notched specimen approximately 

0.004 m (1.6 in.) from the edge crack. Two of the 200-hour specimens also showed slight cracks 

at the holes. The cracks were small and were not thought to have been a factor during residual 

strength testing. 

12.6.2.2 Residual Tensile Strength Data. Following the examination of the 100-hour and 

200-hour short-term flight simulation specimens, residual tensile strength tests were conducted. These tensile tests were made at the same temperatures as the simulation exposures, 

422 K (300 O F) for the B/E and GIE systems and 505 K (450 O F) for both the HT-SI710 G/PI and 

BIAl materials. Full sets of residual strength data were obtained for the epoxies except for the 

100-hour unnotched GIE where three simulation failures occurred. Only partial sets of data 

were available from the GlPI and BlAl systems because of the numerous testing and material 

problems previously discussed in Section 12.6.2.1. The test results, both tensile strengths and 

failure loads, are given in Tables 12-4 and 12-5 for BE, Tables 12-6 and 12-7 for GIE, Tables 

12-8 and 12-9 for G/PI, and Tables 12-10 and 12-11 for BlAL The widths of the unnotched resin 

matrix specimens were 0.025 m (1.0 in). The notched specimens were also 0.025.m (1.0 in.) wide 

but with 0.0064 m (0.25 in.) diameter holes. All the B/Al residual strength specimens were 

0.013 m (0.50 in.) in width. 

9x - 4 653217-233 

Figure 12-33 Edge Cracking (arrow) in Specimen DC82-5, G/PI, Unnotched, After 100-hr Simulation 

Exposure at 505 K (4500 F) 

12-48 
. , 66321 7-234 t 

Figure 12-34 Edge Cracking (arrow) in Specimen DF81-5, G/PI, Notched, After 100-hr Simulation 

Exposure at 505 K (450° F) 

12-49 
Tablo 12-4. Short-Term Rosidual Stnngfh Data, [Oo f 454],, B/E. Unnotchod, Exporod at A22 K (3000 F) and 

Tonsilo Ted at 422 K (3000 F) 

AC82-1 

-2 

-3 

4 

-5 

-6 

-7 

-8 

-9 

-10 

AC83-1 

-2 

-3 

4 

-5 

-6 

-7 

-8 

-9 

-10 

100 

100 

100 

100 

100 

100 

100 

100 

100 

100 

200 

200 

200 

200 

200 

200 

200 

200 

200 

200 

10.7 

10.7 

10.7 

8.94 

9.83 

9.65 

9.92 

9.90 

10.2 

10.0 

ave 10.1 

9.54 

8.01 

8.72 

9.83 

9.65 

9.90 

9.30 

8.45 

9.59 

8.67 

ave 9.17 

2.415 

2.405 

2.410 

2.010 

2.210 

2.170 

2.300 

2.230 

2.225 

2.250 

2.260 

2.145 

1.800 

1.960 

2.210 

2.170 

2.225 

2.090 

1.900 

2.155 

1.950 

2.060 

541 78.4 

553 80.2 

552 80.1 

474 68.8 

5 23 75.9 

502 72.8 

539 78.2 

516 74.8 

518 75.2 

533 77.3 

525 76.2 

483 70.1 

405 58.8 

456 66.2 

516 74.9 

502 72.8 

512 74.2 

487 70.6 

443 64.2 

501 72.6 

462 67.0 

477 69.1 

12-50 
.) 

Tabie 12-6. Short-Term Residual Strength Data, [Oo -C 45O],, WE, Notched, Exposed at 422 K (3000 F) and 

Tensile Ted at 422 K (300O F) 

Notched 

load Tensile Strength 

Number (hr) kN (kip) MN/d (kri) 

Specimen Exporufa nrrn 

AF84-1 

-2 

-3 

-4 

-5 

-6 

-7 

-8 

-9 

-10 

AF82-1 

-2 

-3 

-4 

-5 

-6 

-7 

-8 

-9 

-10 

100 

100 

100 

100 

100 

100 

100 

100 

100 

100 

200 

200 

200 

200 

200 

200 

200 

200 

200 

200 

6.12 

6.41 

6.32 

6.23 

6.85 

6.04 

6.05 

5.38 

6.36 

6.61 

ave 6.24 

5.96 

6.49 

5.98 

4.98 

6.32 

6.41 

5.78 

5.78 

5.69 

5.56 

ave 5.90 

1.377 

1.440 

1.420 

1.400 

1.540 

1.357 

1.360 

1.210 

1.430 

1.485 

1.400 

1.340 

1.458 

1.345 

1.120 

1.420 

1.440 

1.300 

1.300 

1.280 

1.250 

1.330 

'416 (a) 60.4 

452 65.5 

445 64.6 

433 62.8 

485 70.3 

427 62.0 

423 61.3 

379 55.0 

446 64.7 

467 67.8 

437 63.4 

420 60.9 

45 1 65.4 

430 62.3 

354 51.4 

47 1 68.3 

475 68.9 

419 60.7 

419 60.7 

403 58.4 

412 59.8 

425 61.7 

a Net section strength, Le., based on total width less hole diameter. 

12-51 
lablo 12.6. Shod-Tum Residual Stnngfh Data, [0° -C 45O],, G/E, Unnotchod, Exporod at 422 K (3000 F) and 

tensilo Todd at 422 K (300O F) 

Load Tensile Strmgih Spocimm Expun nrno 

Numbor (hr) kN (kip) MN/d (ksi) 

BC84-1 

-2 

-3 

4 

-5 

-6 

-7 

-8 

-9 

-10 

BC81-1 

-2 

BC82-3 

-4 

-5 

-6 

-7 

-8 

-9 

-10 

a54.2 

a54.2 

100 

100 

100 

a10.6 

100 

100 

100 

100 

200 

200 

200 

200 

200 

200 

200 

200 

200 

200 

- 

- 

7.09 

8.09 

9.34 

8.94 

8.45 

9.96 

ave 9.00 

- 

11.1 

10.3 

10.0 

8.27 

9.74 

7.34 

6.56 

9.67 

8.10 

9.56 

ave 8.96 

10.1 

- 

1.595 

1.818 

2.100 

2.010 

2.499 

1.900 

2.240 

2.020 

2.325 

1.860 

2.250 

2.190 

1.650 

1.475 

2.175 

1.820 

2.260 

2.150 

2.020 

- 

- 

426 

445 

492 

b465 

565 

460 

501 

479 

547 

405 

552 

5 23 

38 1 

365 

484 

420 

5 16 

501 

469 

- 

- 

- 

61.8 

64.5 

71.4 

b67.4 

81.9 

66.7 

72.7 

69.5 

79.4 

58.7 

80.1 

75.8 

55.2 

52.9 

70.2 

60.9 

74.8 

72.6 

68.1 

- 

a These specimens failed during flight simulation exposure testing. 

b This specimen had a small crack in the test section after the 100-hr simulation exposure. Failure during 

tensile testing occurred at this location. 

12-52 
Tablo 12-7. Short-Tom Raidual Stnngth Data, [Oo f dS0],, WE, Notehod, Exposd at 422 K (3000 F) and 

Tonriio Tod at 422 K (3004 F) 

BF84l 

-2 

-3 

-4 

-5 

-6 

-7 

-9 

-10 

BF82-1 

-2 

-3 

-4 

-5 

-6 

-7 

-8 

-9 

-10 

100 

100 

100 

100 

100 

100 

100 

100 

100 

200 

200 

200 

200 

200 

200 

200 

200 

200 

200 

5.69 

5.36 

6.05 

5.60 

6.29 

6.05 

6.29 

5.78 

6.00 

ave 5.90 

5.69 

4.98 

5.20 

5.87 

6.16 

5.65 

5.43 

5.69 

. 5.38 

4.17 

ave 5.42 

1.280 

1.205 

1.360 

1.260 

1.415 

1.360 

1.415 

1.300 

1.350 

1.330 

1.280 

1.120 

1.170 

1.320 

1.385 

1.270 

1.220 

1.280 

1.210 

.937 

1.220 

403 

405 

439 

412 

434 

424 

436 

425 

450 

( 4 

58.4 

58.8 

63.6 

59.7 

62.9 . 61.5 

63.2 

61.6 

66.2 

425 61.7 

392 56.9 

367 53.3 

414 60.0 

444 64.4 

452 65.6 

438 63.5 

429 62.2 

465 67.4 

394 57.1 

343 49.8 

414 60.0 

a Net section strength, Le., based on total width less hole diameter. 

12-53 
Tablo 12-8. Short-Term Residual Strongth Data, [Oo -C 450],, HTW710 G/PI, Unnotchod, Exporod at 505 K 

(4500 F) and Tensilo Tested at 505 K (4500 F) 

Load Tonsilo Stnngth Spodmon Exposun nnn 

Numbor (hr) kN (kip) MN/d (kri) 

DC81-12 

DC82-5 

-9 

-10 

-12 

DC85-2 

-4 

-5 

-7 

-9 

-10 

100 

100 

100 

100 

100 

200 

200 

200 

200 

200 

200 

8.12 

7.83 

7.41 

7.34 

8.32 

ave 7.80 

7.03 

7.34 

7.58 

7.03 

7.90 

7.72 

ave 7.43 

1.825 

1.760 

1.665 

1.650 

1.870 

1.750 

1.580 

1.650 

1.705 

1.580 

1.775 

1.735 

1.670 

402 58.3 

395 57.3 

425 61.7 

388 56.3 

393 57.0 

401 58.1 

316 45.9 

339 49.1 

304 55.7 

348 50.5 

426 61.8 

368 53.4 

364 52.7 

12-54 
Tablo 12.9. Short-Term Rosidual Strength Data, [Oo f 45O],, HT-S/7lO G/PI, Notched, Exposed at 505 K 

(4500 F) and Tensilo Tested at 505 K (GOo F) 

load Tonrib Strength Spuimm Exposun Tlmo 

Numbu (hr) kN (kip) MN/& (kri) 

DF81-1 

-2 

-3 

-4 

-5 

-6 

-7 

-8 

-9 

-10 

DF82-1 

-2 

-3 

-4 

-5 

-7 

-8 

-9 

-10 

100 

100 

100 

100 

100 

100 

100 

100 

100 

100 

200 

200 

200 

200 

200 

200 

200 

200 

200 

4.11 

4.83 

4.89 

4.07 

5.38 

4.71 

4.29 

4.56 

4.67 

3.83 

ave 4.53 

5.11 

3.91 

5.60 

4.20 

4.91 

4.50 

4.92 

4.78 

5.27 

ave 4.80 

0.925 

1.085 

1.100 

.915 

1.210 

1.060 

.965 

1.025 

1.050 

.860 

1.020 

1.148 

380 

1.260 

.945 

1.104 

1.012 

1.105 

1.075 

1.185 

1.080 

(a) 

241 34.9 

301 43.7 

295 42.8 

240 34.8 

325 47.1 

294 42.6 

266 38.6 

285 41.3 

286 41.5 

234 34.0 

277 40.1 

319 46.3 

236 34.2 

328 47.6 

247 35.8 

305 44.2 

265 38.5 

308 44.6 

295 42.8 

311 45.1 

290 42.1 

a Net section strength, Le., based on total width less hole diameter. 

12-55 
Tablo 12-10. Sho*Tonn Rosidual Stnngt)l Data, [Oo -C 45O],, B/Al, Unnotchod, Exporod at 505 K (4500 F) 

and Tensilo Ted at 505 K (4500 F) 

EC84-1 

-2 

-3 

4 

-5 

-6 

-7 

-8 

-9 

-10 

EC82-1 

-2 

-3 

4 

-5 

-6 

-7 

-8 

-9 

-10 

100 

100 

100 

100 

100 

100 

100 

100 

q5.4 

100 

a105 

a136 

200 

200 

200 

200 

a136 

200 

200 

a136 

7.56 

7.12 

7.34 

7.21 

5.78 

7.29 

6.56 

6.14 

7.07 

ave 6.90 

- 

6.20 

5.27 

5.12 

5.43 

5.72 

6.52 

4.94 

4.25 

4.49 

6.83 

ave 5.48 

1.700 

1.600 

1.650 

1.620 

1.300 

1.640 

1.475 

1.380 

- 

1.590 

1.550 

1.395 

1.185 

1.150 

1.220 

1.285 

1.465 

1.110 

.955 

1.010 

1.535 

1.230 

604 

566 

574 

561 

457 

580 

519 

471 

- 

556 

543 

509 

415 

40 1 

423 

434 

507 

383 

323 

348 

521 

426 

87.6 

82.1 

83.3 

81.4 

66.3 

84.1 

75.3 

68.3 

80.7 

78.8 

- 

73.8 

60.2 

58.1 

61.3 

63.0 

73.6 

55.5 

46.8 

50.5 

75.6 

61.8 

a These specimens failed during flight simulation exposure testing. 

12-56 
Tablo 12-1 1. Short-Tom Rosidual Strongfh Data, [PI6 WAl, Unnotehod, Expod at 505 K (4500 F) and 

Tonsilo T4 at 505 K (4500 F) 

Load Trmilo Stnngth 

kN (kip) MN/d (kri) 

EU85-1 

-3 

-6 

-9 

EU86-3 

-6 

EU87-11 

-12 

-14 

EU87-1 

-2 

-5 

-6 

-7 

-8 

100 

100 

100 

100 

100 

100 

100 

100 

100 

200 

200 

200 

200 

200 

200 

17.6 

19.0 

17.8 

17.3 

17.3 

18.0 

21.8 

21.7 

12.5 

ave 18.1 

21.1 

21.5 

20.5 

21.3 

23.0 

20.9 

ave 21.4 

3.950 

4.280 

4.010 

3.900 

3.900 

4.050 

4.910 

4.880 

2.820 

4.080 

4.750 

4.850 

4.600 

4.780 

5.180 

4.700 

4.810 

1460 211 

1610 233 

1490 216 

1410 205 

1390 201 

1550 225 

1560 226 

1560 227 

882 128 

1430 208 

1460 211 

1520 221 

1450 210 

1520 221 

1630 236 

1480 214 

1510 219 

12.6.3 SHORT-TERM ANALYSIS. Using the residual strength measurements from Section 12.6.2 

and the baseline data from Section 7.1.2, the peak loads for the long-term, real-time flight 

simulation tests were calculated. The wearout model equations discussed previously in Section 

12.5 were programmed into a desk top computer (HP9830) to calculate the parameters r, A,, af, 

and Pf starting with q,, Po, and the short-term residual strength data. The results for GIE and 

BIE are listed in Table 12-12. For HT-SI710 GIPI and BlAl the wearout model was not successful (to be discussed later). Since the long-term specimens were three times as wide as the 

short-term specimens they were expected to exhibit a reduced static strength. This expectation was based on the increase in complexity (least of three identical specimens). The strength 

reduction can also be thought of as a scaling effect since the long-term specimens have three 

times the stressed volume of the short-term specimens. The expected strength of the long-term 

specimen is PLT = fc Pm where the scaling factor, fc = (3) is the penalty for least of three 

as a function of the shape parameter, ao. The values of f, are listed in Table 12-12. 

-1 - 

12-57 
ea......- 

rouaea... 

0 

a?? 40' 

12-58 
A few other assumptions are necessary to predict the loads appiied in the long-term tests to 

achieve the desired results. The problem is to apply loads high enough to produce measureable 

effects without failing all specimens in fatigue. A reliability of 0.80 at one lifetime was the 

desired goal. In other words, one out of five specimens used for the long-term tests would be 

expected to fail in fatigue prior to the completion of the 50,000 hours of exposure. 

One remaining parameter of the wearout model, s, must also be determined. The parameter, s, 

is the exponent on the stress or load ratio that defines the power law time response to changing 

loads. This parameter is obtained by subjecting specimens to various magnifications of the 

random spectrum to obtain a random S-N me. The slope of the random S-N curve is - 1 - 

9 

Since no such tests were performed for the current program, s has been estimated from data in 

reference 46 and 47 to be 11.0. The assumption of s is not felt to be too serious in this case 

because the loads required for the desired long-term response are not very different from the 

loads applied in the short-term tests. 

To calculate the appropriate Al for the desired results, Equations (5) and (7) (Section 12.5) are 

combined by eliminating ?f to arrive at: 

p 2(r-1) (-m) i/af 

A1=[ (r-1) O A, t 

The results of this calculation are also listed in Table 12-12. Using the model parameters of 

Table 12-12 and the wearout model equations of Section 12.5 the peak long-term loads were 

obtained by an iteration process to give 80% survival at one lifetime (50,000 hours). These 

values are given in Table 12-12. 

The computer was used to plot the wearout model curves shown in Figures 12-35 through 

12-38 for the unnotched and notched GIE and BIE systems. The horizontal peak load lines have 

been set for 80% survival at one lifetime. 

Wearout model curves could not be plotted for the HT-SI710 G/PI and BlAl systems because 

the wearout model did not work for these materials. In some cases the short-term residual 

strength value, 0, and the shape parameter, a, increased with time. The 200-hour or half life 

time residual strength values were actually higher than the 100-hour or quarter lifetime values, 

and the spread in the data was decreased. One method initially considered for determining the 

long-term peak loads for GlPI and BlAl was to take two-thirds of the "A" basis allowable tensile strength value for the particular composite considered. However, it was thought that this 

method did not place sufficient importance on the shape parameter, a, which for BlAl was 

higher than for GIPI. The method that was used was to essentially reapply the risk factor, 

-1 

a f, = (100) , to the short-term peak load using the worst case a as determined from the 

baseline, 100-hour, or 200-hour strength data. This amounted to setting the long-term loads for 

these two systems by applying the square of the risk factor, fp times the complexity factor, f, 

to the baseline 0 for each system at the temperature at which the long-term simulation was to 

be performed. The results of these calculations for HT-Sl710 GlPI and BlAl are given in Table 

12-13. 

12-59 
12 

10 

a, - 

2 

-* 

X 

Io 

z 

Iv) 

d 

-2 

3 

- 

LY 6' a 

= 

z4 

a 

2. 

0 

10-2 

Ffguro 12-35 Woarout Modo4 for Unnatchod [Oo f 4501, G/E 

0 - 25 

0 

n 

V 

rn 

z 

a OPR ES10 U A -I L STRENGTH 

O=FATIGUE FAILURE I 

- 1.0 E 

3, 

010 = 1271 Po = 2.12 to = 0 r = 7 A1 = 1 A4 = 395 af = 1.06 s = 11 . 8 

Pf = 3.47 

- 0.5 

I I I I I Ill I 1 I I IIII I I I I IIII 0 

10'1 1 10 

tablo 12-13. Long-Torm Peak loads for HT-S/710 G/PI and B/Al 

bPeak load 

Matorial Condition kN/m (kip/in) aT f, f ,z fc kN/m (kip/in) 

HT-SI710 GlPI Umotched 366 2.09 10.7 0.65 0.42 0.90 140 0.800 

Notched 233 1.33 10.0 .63 .40 .90 83.4 0.476 

BlAl [O']s 1566 8.94 14.7 .73 .53 .93 778 4.44 

[O" k 45'1, 427 2.44 13.6 .71 .51 .92 200 1.144 

BT = fTpRT where T = 505 K (450' F) 

b Peak Load = fr2f& 

12-60 
ul 

$3 

3 

0 

w ul 

- 

a 

2 

1 

0 

2.5 

2.0 

n 

m 

2 

c) + 

I 

0 RESIDUAL STRENGTH 

- 

- 0.50 

= 1203 0, = 1.29 1, = 0 r = 7 A1 = 1 AQ = 0.564cuf= l.OOs= 11.0 

pf 6.16 - 0.25 

I I I I I Ill I I I I1111 I I I I I Ill 0 

a 

a 

3 ORESIDUAL STRENGTH 

z4 

0 

ul 

- - 

2 

1 I 

- 0.5 ol,=19.5p,=2.30t,=Or=5Al 11 A4= 131ru~=2.44~~=1.50~=11 

12-61 

0 I I I11111 I I I I I Ill I 1 I I 1111 0 
z4 

a 

4 

Iv). 

2 - 3. v) 

W 

a 

2 

1 

0 

65321 7-238 

cn 

\ I I I 0 

c PEAKLOAD - .8 B 

Y m 

m 

z 

0 --E -I RESIDUAL STRENGTH I 

x 

z 

0.950 0.90 0.5 0.1 

- .4 

Cyo = 19.5 0, = 1.45 to = 0 r = 5 A1 * 1 A4 = 2.77 Cyf = 2.44 = 1.76 S = 11 

- .2 

I I I I I Ill I 1 I I1 Ill 1 I 1 IIIllo 

Figuro 12-38 Woarout Modd for Notchad [OO -C aSo], B/E 

The wearout model also could not be used to calculate the long term loads for Celion 

60001LARC-160 as no short term testing was performed on this system. The procedure for 

long term load selection, for both unnotched and notched specimens, was similar to that de 

scribed earlier for the short term loads (Section 12.6.1.2). The baseline unnotched and notched 

tensile strength data were used to find the values of the scale parameter 0 and a pooled value 

for CY, the Weibull shape parameter. For Celion 60001LARC-160 the data for all temperatures 

(Tables 12-14 and 12-15) were suitable for pooling. The risk factor based on 1% failure was 

calculated as fn = (lOO)-l/a. Since the pooled CY for the unnotched samples was 28 and the CY for 

the ten specimens at 450K (350°F) was 22 (Table 12-14) the more conservative value of 22 was 

used. This approach was also used in earlier calculations for other systems. A temperature 

retention factor (fT) of 1.00 was chosen for the unnotched case for 450K (350°F) and 1.08 for 

the notched case for use at 450K (350 OF). The factor fc was (6)-l/a for the unnotched specimens 

since the baseline specimens were 0.013 m (0.50 in.) wide and the long term specimens were 

0.076 m (3.00 in.) wide. For the unnotched specimen it was (3) since the baseline specimens were 

0.025 m (1.00 in.) wide. Both factors turned out to be equal because of the difference in the 

pooled CY values, f, = 0.92. ~ 

12-62 
Tablo 12-14. Barelino Tonsilo Data of [OO f 45O], Celion 6000/LARC-l60 G/PI 

POOLED a 28 

Iding Load for Iding Load 

Tompuaturo 0.013 m (0.50 in.) 3p.c por Inch P 

Spufmon No. K (OF) kN (kip) kN/m (kip/ln) kN/m (kfp/in) a 3/P, Load/P kN/m (kip/in) J/P, 

FCO1-I 

-2 

-3 

-4 

-5 

-6 

-7 

-8 

-9 

-10 

-11 

-12 

-13 

-14 

-15 

-16 

-17 

-18 

-19 

-20 

-21 

-22 

-23 

-24 

-25 

-26 

-27 

-28 

-29 

-30 

-31 

-32 

-33 

-34 

-35 

218 

297 

394 

450 

4723 

505 

533 

-67 7.219 

7.495 

7.740 

75 7.473 

7.540 

6.984 

7.340 

7.073 

7.059 

6.695 

6.939 

6.672 

6.157 . 250 7.762 

6.850 

7.829 

7.206 

6.695 

6.628 

6.308 

6.628 

7.397 

7.473 

7.028 

400 7.139 

7.028 

7.006 

450 7.095 

6.539 

7.028 

500 6.872 

7.006 

7.028 

350 7.428 

7.286 

1.623 568.4 

1.685 590.2 

1.740 609.4 

1.680 588.4 

1.695 593.7 

1.570 549.9 

1.650 577.9 

1.590 556.9 

1.505 527.1 

1.560 546.4 

1.500 525.4 

1.465 513.1 

1.745 611.2 

1.540 539.4 

1.760 616.4 

1.670 584.7 

1.620 567.4 

1.505 527.1 

1.490 521.9 

1.418 496.6 

1.490 521.9 

1.638 573.7 

1.663 564.9 

1.680 588.4 

1.580 553.4 

1.605 562.1 

1.580 553.4 

1.575 551.6 

1.595 558.6 

1.470 514.8 

1.580 553.4 

1.545 541.4 

1.575 551.6 

1.580 553.4 

1.587 555.8 

3.246 597.2 

3.370 

3.480 

3.360 566.0 

3.390 

3.140 

3.300 

3.180 

3.174 

3.010 

3.120 

3.000 

2.930 

3.490 604.2 

3.080 

3.520 

3.240 

3.010 

2.980 

2.836 

2980 

3.276 

3.226 

3.360 

3.160 

3.210 558.3 

3.160 

3.150 

3.190 550.9 

2.940 

3.160 

3.090 551.3 

3.150 

3.160 

3.340 568.4 

3.410 41 1.06 0.95 595.1 3.398 1.05 

.99 

1.02 

1.05 

.97 

1.02 

.98 

.98 

.93 

.97 

.93 

.91 

.89 

1.02 

1.0 

3.232 24 1.00 1.04 567.6 3.241 1.00 

3.450 24 1.07 1.01 605.2 3.456 107 

3.246 22 1.00 1.03 568.1 3.244 1.00 

.93 

.92 

.87 

.92 

1.01 

1.02 

1.04 

.97 

.99 

.99 

. .93 

1.00 

1.00 

1.00 

3.188 120 0.99 1.01 556.2 3.176 0.98 

3.146 40 .97 1.01 549.2 3.138 .97 

3.148 145 .97 .98 549.5 3.138 .97 

.ORIGINAL PAGE IS 

Q! ROOR QUALITY 

12-63 
tablo 12-15. Barolino Notehod Tonsilo Data of [OO f 4501, Colion 6000/LARC-l60 G/Pl 

POOLED a =: 14 . .) 

Bmklng Load for Broaklng Lood 

u 4au Tompomluro 0.013 rn (0.50 in.) Sp.c pu Inch 

Spulmm No. K (OF) kN (kip) kN/m (kip/in) kN/m (kiplin) u P/P, Load/B kN/m (kip/in) B/& 

FCO1-1 

-2 

-3 

-4 

-5 

-6 

-7 

-8 

-9 

-10 

-11 

-12 

-13 

-14 

-15 

-16 

-17 

-18 

-19 

-20 

-21 

-22 

-23 

-24 

-25 

-26 

*27 

- 28 

-29 

-30 

-3 1 

-32 

.-33 

-34 

-35 

218 

297 

394 

450 

478 

505 

533 

-67 7.073 

6.717 

6.850 

75 6.361 

5.961 

6.539 

7.473 

8.140 

7.295 

8.096 

6.405 

6.494 

7.028 

250 7.206 

7.295 

5.916 

350 6.761 

7.918 

7.562 

7.918 

7.073 

8.155 

7.651 

8.007 

7.651 

8.318 

400 8.585 

6.405 

7.562 

450 9.341 

8.674 

7.962 

500 8.007 

8.985 

7.117 

1.590 371.3 

1.510 352.5 

1.540 359.5 

1.430 334.0 

1.340 330.5 

1.470 343.2 

1.680 3923 

1.830 427.3 

1.640 383.0 

1.820 425.0 

1.440 336.2 

1.460 341.0 

1.580 369.0 

1.620 378.3 

1.640 383.0 

1.330 310.5 

1.520 355.0 

1.780 415.6 

1.700 397.0 

1.780 415.6 

1.590 371.3 

1.840 429.6 

1.720 401.6 

1.800 420.3 

1.720 401.6 

1.870 436.6 

1.930 450.6 

1.440 336.2 

1.700 397.0 

2100 490.3 

1.950 455.3 

1.790 418.0 

1.800 420.3 

2.020 471.6 

1.600 373.5 

2120 365.0 

2013 

2.053 

1.907 383.9 

1.887 

1.960 

2240 

2.440 

2187 

2.427 

1.920 

1.947 

2107 

2160 371.4 

2187 

1.773 

2.027 415.0 

2.373 

2267 

2373 

2.120 

2.453 

2.293 

2.400 

2293 

2493 

2573 415.4 

1.920 

2.267 

2.800 468.4 

2.600 

2.387 

3.400 440.3 

2.693 

2.133 

2.084 50 0.95 1.02 362.1 2.068 0.93 

0.97 

.99 

.86 

.a9 

1.02 

1.11 

1.00 

1.11 

.88 

.a9 

.96 

1.03 

.84 

1.00 

.96 

1.00 

.89 

1.04 

.97 

1.01 

.97 

1.05 

.8 1 

.96 

.97 

.89 

1.07 

.85 

2192 10 1.0 .87 389.6 2.225 1.0 

2.121 16 .97 1.02 370.6 2.116 .95 

2.370 22 1.08 .86 411.9 2.352 1.06 

2.372 10 1.08 1.08 421.7 2.408 1.08 

2.675 18 1.22 1.05 465.8 2.660 1.2 

2.514 12 1.15 .95 442.7 2.528 1.14 

12-64 
Table 12-16 shows the reduction factors and the calculated peak loads for the Celion 6000/ 

LARC-160 G/PI long term flight simulation tests. The loads used on the HT-S/710 G/PI are 

also listed for comparison. The peak loads for Celion 6000/LARC-160 were considerably higher 

than those calculated for HT-S/710. The scale parameter 0 was found to be consistently higher 

and the shape parameter CY was also much higher. - 

Tablo 12-16. Long-Tom Peak Loads for [(P f 4501, G/PI 

“& bQeak Load 

Motorial Condition kN/m (kip/in) aT f r f ? f, kN/m (kip/in) 

HT-SI710 Unnotched 366 2.09 10.7 0.65 0.42 0.90 140 0.800 

Notched 233 1.33 10.0 .63 .40 .90 83.4 .476 

Celion 60001 unnotched 567 3.24 22 .81 .53 .92 338 1.93 

LARC-160 Notched 443 2.53 14 .72 .51 .92 198 1.13 

~ ~~~ 

a PT = f@RT where T = 505 K (450O F) for HT-SI710 and 450 K (350°F) for Celion 60001LARC-160 

b Peak Load = fr2f,& 

Using the load levels specified in Table 12-16 one of the unnotched Celion 60001LARC-160 

G/PI specimens failed after only 1.3 hours of exposure. Because of this early failure another 

look was taken at the process of load selection. The original unnotched loads were calculated 

using the baseline unnotched tensile data obtained at seven different temperatures. All 35 

specimens were taken from panel 111-5. Because the flight simulation specimens were taken 

from panels 111-8,111-9, and 111-10, revised loads were calculated using the quality assurance 

(QA) tensile data for these three panels. The QA data consisted of three tests each at 297 K 

(75 OF) and 450 K (350 OF) for each panel. The Weibull scale and shape parameters calculated 

from these data are given in Table 12-17 along with the values used in the on@ load de 

terminations. Using the QA data from panels 111-8, 111-9, and 111-10 resulted in a reduction in 

the peak long term load from 338 WIm (1.93 kipIin) to 277 kN/m (1.582 kipIin). 

This load level also proved to be excessive with a second unnotched specimen failure occurring 

after 138 hours of exposure. As these specimens were intended to have an 80% expectation of 

completing 50,000 hours of exposure, a further reduction in the peak load level was necessary. 

To increase the data base for the Celion 6000ILARC-160 GIPI system, QA tensile data from all 

the 6 ply panels were used in calculating a revised peak load. The Weibull scale and shape 

parameters calculated from QA tensile data from seven test panels (42 specimens) are listed in 

Table 12-17. The recalculated peak load was 183 kN/m (1.043 kip/in). This value was now lower 

than the peak load that had been chosen for the notched specimens (based on notched baseline 

tensile data). Since the notched baseline tensile strengths were consistently lower than the unnotched baseline tensile strengths it was felt that the same relationship should be held for the 

flight simulation peak loads. As no notched tensile data were obtained during QA testing, the 

value of ‘;YT (a measure of data scatter) obtained from the QA unnotched tensile data was used 

to recalculate fr and f, These values were then used with the 0T obtained from the notched 

baseline tensile data to calculate a revised peak load of 143 kNlm (0.819 kiplin) for the notched 

12-65 
Tablo 12-17. Long-Tom Pea& Loads for [VJ f 4501, G/P! 

a& ha& Load 

Matorial Condition kN/m (kip/in) crT f , f ,z f, kN/m (kip/in) 

~ 

ORIGINAL LOADS 

HT-SI710 Unnotched 366 2.09 10.7 0.65 0.42 0.90 140 0.800 

Notched 233 1.33 10.0 .63 .40 .90 83.4 0.476 

Celion 6000/ Unnotched 567 3.24 22 .81 .53 .92 338 1.93 

LARC- 160 Notched 443 2.53 14 .72 .51 .92 198 1.13 

REVISED LOADS 

QADatafrom Unnotched 595 3.397 14.4 .73 .53 .88 277 1.582 

Panels 8.9 & 10 

QADatafrom Unnotched 562 3.212 9.8 .625 .39 .83 183 1.043 

Panels 4.5.7, 

8,9,10 & 11 

UT from Notched 443 2.53 9.8 .625 .39 .83 143 .819 

QA Data 

a PT = fT&T where T = 505 K (450" F) for HT-S/710 and 450 K (350°F) for Celion 60001LARC-160 

b Peak Load = fr2fCpT 

flight simulation tests (see Table 12-17). With the recalculated peak loads both the unnotched 

and notched specimens completed 10,000 hours of flight simulation exposure without additional failures. 

The peak long-term loads for the five materials, Tables 12-12, 12-13, and 12-17, were used to 

calculate the peak, landing, and one g stresses presented in Table 12-18. 

12-66 
Tablo 12-18. Long-Term Peak Loads and Stressos 

Td Peak Load 

Long-term for 0.076 rn 

Pd Load (3.0 in) Sp.cimm Peak Stress landing Strns On, g Sh.r, 

Mddd Condition kN/m (kip/in) kN (kip) MN/d (kri) MN/d (kri) MN/d (kri) 

BIE 

GIE 

unnotched 

Notched 

Unnotchd 

Notched 

HT-Sl710 GlPI Unnotched 

Notched 

Celion 60001 Unnotched 

LARC-160 GlPI Notched 

BlAi IO "Is 

[O" 2 45'1, 

230 1.314 17.5 3.94 302 43.8 

132 .752 10.1 2.26 239 34.6 

232 1.325 17.7 3.98 305 44.2 

132 .756 10.1 2.27 232 33.6 

140 300 10.7 2.40 153 22.2 

83.4 .476 6.36 1.43 125 18.2 

183 1.043 13.9 3.13 243 34.8 

143 .819 10.9 2.46 191 27.3 

778 4.44 59.3 13.32 747 108.3 

200 1.144 15.3 3.43 208 30.1 

-75.5 -10.9 

-60.0 -8.7 

-76.5 -11.1 

-57.9 -8.4 

-38.6 -5.6 

-31.7 -4.6 

-60.0 -8.7 

-47.7 -6.8 

-187 -27.1 

-51.7 -7.5 

151 21.9 

120 17.3 

153 22.1 

116 16.8 

76.5 11.1 

63 9.1 

122 17.4 

95 13.7 

374 54.2 

104 15.1 

12.7 LONG-TERM FLIGHT SIMULATlON 

The main emphasis of the entire program was on the long-term tests and that is why they were 

approached so carefully. The objective of the long-term test program was to determine the 

behavior of advanced composites in the operating environment of supersonic cruise aircraft. 

Specimens were to be exposed for real-time periods of 10,000, 25,000, and 50,000 hours. The 

tests would be unsuccessful if the environmental conditions (loads or temperatures) were set 

either too high or too low. If the conditions could be set so that a given number failed by the 

end of the required 50,000 hours of exposure, not only would the behavior of the composites be 

determined but also a predictive capability would be demonstrated. 

12.7.1 TEST PLAN AND PROCEDURES. For each material system of Phase I, BIE, GIE, HT-SI710 

GIPI, and BlAL a total of 20 long-term specimens was installed in the flight simulation apparatus. The resin matrix specimens were all fabricated from 6 ply panels having a [0 O -C 45 OIs 

crossplied orientation. Ten were unnotched and 10 were notched (see Figures 12-9 and 12-10 for 

specimen configurations). The BIAl specimens, all of which were unnotched, were fabricated, 

half from 6 ply unidirectional and half from 6 ply [0 O f 45 'Is crossply panels. Five specimens 

from each materiaUorientatiodconfiguration set were to be exposed for 10,000 hours 

(equivalent to 5000 of the two-hour flights) for the Phase I evaluation. The other five 

specimens of each set were to remain in the simulation apparatus until 50,000 hours (25,000 

flights) of exposure had been completed (Phase 11) or until excessive specimen failures made it 

uneconomical to continue operation of the simulator. 

Several modifications to the Phase I1 test plan were made at the completion of 10,000 hours of 

exposure. The first was the addition of three unnotched long-term specimens to each materid 

orientation set. These specimens were installed in the flight simulation apparatus after the 

10,000-hour specimens were removed. The added specimens were to be exposed for 25,000 

hours to provide additional residual strength data between the on@y scheduled 10,000- 

and 50,000-hour time periods. The second modification was caused by a high failure rate of the 

GIE specimens and extensive visual degradation of the BIE specimens during the first 10,000 

12-67 
hours. All of the original B/E and G/E flight simulation specimens were removed, and new 

specimens were fabricated to replace them. These specimens, which were all unnotched, were 

installed in the simulator as foilows: 

Bay 5 5 BE, 10,000-hr specimens 

5 G/E, 10,000-hr specimens 

Bay 6 3 BiE, 25,000-hr specimens 

3 GIE, 25,000 hr specimens 

Bay 7 5 B/E, 50,000-hr specimens 

5 GIE., 50,000-hr specimens 

The random load spectrum remained unchanged except that the load levels were now the 

average of those applied previously for the two epoxy materials. The change (shown below for 

the maximum tensile loads) was quite small, less than 1%. 

Previous BIE load: 17.53 MN (3942 lb) 

Previous GIE load: 17.68 MN (3975 lb) 

Revised load: 17.60 MN (3958 Ib) 

An increase in testing efficiency was possible by intermixing the two systems as the flight 

simulation exposures could be performed in three bays rather than the four required if the systems were run separately. The maximum test temperature during flight simulation testing of 

BIE and GIE was reduced to 373 K (212°F). This temperature was determined from a log time 

versus 1/T plot of G/E thermal aging data obtained earlier in the program. The we 

represents the timeItemperature combinations at which significant degradation of the epoxy 

matrix begins. Extrapolation of the curve to 50,000 hours gave the new test temperature. 

The final modification to the Phase I1 test pian was the addition of the new material, Celion 

6000LARC-160 GIPI. Ten unnotched and 10 notched specimens were prepared and installed 

in the flight simulation apparatus for 10,000 hours of exposure testing. 

The flight simulation cycle, shown earlier in Figure 12-1, is two hours in length and consists of 

three parts: climb, cruise, and descent. The timetemperatureload history of a typical flight is 

shown in Table 12-19. Each flight is concluded with one compressive load cycle. The load spectrum is applied in a random manner. 

Table 12-19. Time=Tempemturo-Load History for Typical long-Term Flight Simulation Cyde 

Climb Cruise Descent 

Item (Heating) (Held at Constant Temperature) (Cooling) 

Time 10 minutes 44 minutes 2 minutes 44 minutes 20 minutes 

aLoad 25 cycles constant 6 cycles constant 7 cycles 

Wne additional compressive load per twehour flight. 
The cruise portion of the simulation cycle was conducted at a constant temperature for each 

material system. The intent of the program was to select the temperature that was considered 

to be the maximum use temperature for each composite for longtime (50,000-hour) service. For 

the HT-S/710 GlPI and BlAl materials, the short-term results and data from the thermal aging 

studies then in progress indicated that 505 K (450" F) was probably a good upper limit, and 

this temperature was selected. For the two epoxy systems, the short-term results obtained at 

422 K (300" F) gave no reason to change the exposure temperature, and 1000-hour thermal 

aging data at 394 K (250 F) and 450 K (350 " F), which was then available, also indicated no 

problem with the 422 K (300" F) temperature. Therefore, the long-term flight simulation 

exposures were begun using these temperatures for the cruise portions of the cycle. Shortly 

after starting the tests, the 5000-hour thermal aging results for the epoxy systems became 

available. The data indicated that severe degradation of the epoxy matrix with sipficant 

decreases in tensile strength had occurred at 450 K (350 " F). No effects were observed at 394 K 

(250" F). Primarily because of the thermal aging data, but also as a result of discussions with 

various individuals who were well versed in composite usage and research and development, 

the maximum exposure temperature for the epoxy long-term tests was lowered to 408 K 

(275" F). At the time that the temperature was changed, the epoxy specimens had received 

only approximately 1000 hours of exposure. As discussed earlier the 408 K (275" F) cruise 

temperature was also found to be excessive, and, following the 10,000 hours exposure of 

PhaseI, new epoxy specimens were prepared for Phase 11 and the maximum exposure 

temperature was again lowered, this time to 373 K (212 " F). The selection of 450 K (350 " F) as 

the maximum use temperature for the Celion 60001LARC-160 system was based on minimal 

data from the then current literature. Sufficient thermal aging results from this study were not 

yet available. 

Selection of the maximum load levels for the long-term flight simulation exposure tests has 

been discussed earlier in the short-term test results analysis section. A summary of the types 

and numbers of long-term specimens with the test conditions for each is given in Table 12-20. 

Before installing the long-term specimens in the flight simulation apparatus, width and 

thickness measurements were made at several locations in the test sections of all specimens, 

and average cross-sectional areas were calculated for each set of specimens. These measure 

ments were used to obtain the flight simulation stresses shown in Table 12-20 and could be used 

for the residual strength determinations for specimens that experienced surface degradation 

during the long-term exposure tests. The specimens were then installed in the test machine, thermocouples were attached, the compression stiffening grids were bolted in place, and all resinmatrix specimens were heated to their respective cruise temperatures and held for 24 hours to 

remove any absorbed moisture. Before starting the tests, each set of 10 specimens was proof 

loaded to the maximum value that would be experienced during the 50,000 hours of exposure to 

the random load spectrum. 

At the completion of 10,000 hours of flight simulation testing, all unfailed Phase I specimens 

were removed from the simulator and visually examined with a stereo microscope to check for 

changes in appearance or the presence of surface flaws such as loose or bare fibers, cracks, or 

delaminations. Photographs were made of typical specimens and of those with significant 

changes in appearance as a result of the 10,000 hours of exposure. The large, 0.076 m (3 in.) by 

0.46 m (18 in.), long-term specimens were then cut into smaller pieces, which were fabricated 

into various types of specimens and tested for determination of residual strength values. 

I 

12-63 
tablo 12-20. Summary of Long-tom Test Progrom 

cruise 

Numbor of Spuimons Tomprature PwkStrmss Matuial 

syrtmm Oriontailan bnfigumtion 10,ooO hr 2S.ooO hr 50,000 hr K (Of) MNM (kri) 

BiE 

550514 

BIE 

5505i4 

GiE 

GiE 

GIPI 

GIPI 

A-Sl3501-5 

AS1350 1-5 

HT-Si71 0 

Celion 60001 

' IARC-160 

BlAl 

[0° 2 457, unnotched 

[0° f 457, unnotched 

Notched 

Notched 

Notched 

Notched 

Notched 

Notched 

[oo f 45'1, UMotched 

[0° 2 457, unnotched 

[0° f 457, unnotched 

. [Oo f 45'1, Unnitched 

[0 O f 45 "1, Unnotched 

Pol, Unnotched 

10 

10 

5 

0 

10 

10 

5 

0 

5 

5 

10 

10 

5 

5 

0 

0 

3 

0 

0 

0 

3 

0 

3 

0 

0 

0 

3 

3 

0 

0 

5 

0 

0 

0 

5 

0 

5 

5 

0 

0 

5 

5 

408 275 302 43.8 

408 275 239 34.6 

373 212 303 44 

408 275 305 44.2 

408 275 232 33.6 

- - - 

373 212 303 44 

- - - 

505 450 153 22.2 

505 450 125 18.2 

450 350 236 34.2 

450 350 184 26.7 

505 450 208 30.1 

505 450 747 108.3 

Similsr procedures were carried out for Phase I1 specimens that completed 25,000 hours of exposure. These included BIE. GIE, and a portion of the BIAl specimens. The three 25,000 hour 

HT-SI170 GIPI specimens failed to complete the required 25,000 hours of exposure. 

The objective of the Phase I1 flight simulation test program was to subject the BIAL HT-SI710 

GIPI, BIE, and GIE specimens to 50,000 hours of exposure followed by residual strength 

testing. None of the BIAl or GIPI specimens survived the simulation testing to 50,000 hours so 

that the originally scheduled post-exposure test plan could not be accomplished. However, one 

half of each failed simulation specimen was salvaged for residual strength testing, and a re 

vised test plan with considerably fewer tests was conducted. For the BIE and GIE systems the 

decision was made to halt testing after about 36,000 hours in order to have sufficient material 

for residual strength testing. Two of the B/E specimens had already failed (near 33,000 hours) 

and the likelihood of additional failures was high. 

12.7.2 PHASE I LONG-TERM TEST RESULTS. Once started, the long-term tests were intended to 

continue uninterrupted 24 hours a day, seven days a week except-for brief shutdowns to 

change the program tapes, retighten the BIN doubler clamps, remove failed specimens, and 

provide equipment maintenance. Some unexpected problems did occur from time to time, however, and these often required extended amounts of downtime to correct. The monitoring and 

safety devices were also responsible for numerous shutdowns when test conditions tended to 

stray slightly from the allowable ranges. Any shutdown during nonworking hours could not be 

corrected until the next working day and could result in losses of one to 16 hours of exposure 

(up to 62 hours on a weekend) depending on the time of shutdown. If the tests were halted for 

24 hours or more an additional four hours were lost because of the moisture bakeout of the 

12-70 
resin matrix specimens. At the completion of the short-term tests and at regular intervals during the long-term tests, the flight simulation apparatus was given an overhaul, which, as a 

minimum consisted of changing the piston rod seais in the hydraulic cylinders and replacing 

the quartz heat lamps. Any worn bearings or bushings in mechanical systems and any questionable electronic components were replaced as necessary. 

The short-term test program was very successful in checking out the flight simulation apparatus, the suitability of the test specimen design, and the test procedures required to ensure 

long-time reliable operation during the long-term exposures. Many problems were uncovered 

and corrected. Some additional problems were encountered, however, after the long-term tests 

were begun. One problem that occurred with the compression stiffening grids affected only the 

long-term specimens. If the grids rotated slightly during testing, the attachment screws, see 

Figure 12-39, could bear against the specimen edges and in time could wear away a portion of 

the material, forming an edge notch This could not occur with the short-term specimens 

because of the much narrower widths involved. Several long-term specimens were found where 

the grid screws had begun to bear against the edges, but, with the exception of one GIE 

specimen, the effects were very slight. The GIE specimen failed prior to 10,000 hours with the 

failure location near one of the end doublers away from the heated zone. There was some edge 

damage in the vicinity of the fracture site, and this damage was believed to have been the 

cause of failure. The problem of the screws bearing against the specimen edges was eliminated 

by adding aluminum strips between the edges of the specimens and the attachment screws, 

Figure 12-39. During the short-term testing a problem of abrasion of the soft BIAl specimens 

during fatigue cycling was identified. This was caused by the relative motion of the specimens 

and compression stiffeners and led to a form of fretting corrosion. In an attempt to reduce this 

abrasion, the BlAl compression stiffeners were covered with a stratified Teflon coating, a 

single coat, self-lubricating finish designed for relatively low temperature heat curing. During 

baking, a fluorocarbon-rich surface stratum is formed that provides the low friction properties. 

The technique helped to reduce the wearing of the surface, but at the completion of short-term 

testing the Teflon coating had worn away in several locations. It was also noted that the 

uncoated stiffener grids used on the short-term resin matrix specimens had also caused some 

abrasion of these specimens. Consequently, all the grids were modified for the long-term tests. 

However. because of the high cost of stratified Teflon, the grids were covered with an adhesive 

backed fiberglassITeflon tape. This reduced, but did not eliminate, the surface wearing problem. One other problem occurred with the unidirectional BIAl long-term specimens. From 

time to time specimens exhibited failures in the adhesive bonds between the BlAl and the 

titanium doublers. The adhesive, HT-424, which is quite porous, slowly oxidized and eventually failed under the much larger loads used for the unidirectional BIAL No damage was done to 

the specimens, but testing time was lost while the specimens were removed and new end 

doublers were attached. This problem continued throughout the test program. The remainder 

of the problems were with the equipment and involved periodic failure or erratic behavior of 

mechanical, electronic, or hydraulic components as a result of the rather severe service re 

quirements of the flight simulation apparatus. 

Periodically during the 10,000 hours of exposure in Phase I, the test specimens were examined 

in the test apparatus for changes in appearance or damage as a result of the simulation cycles. 

During two of these examinations all of the compression stiffening grids were removed so that 

the specimens could be studied more closely and, where desired, photographs could be made. It 

was during the first of these examinations that the problem of the grid screws bearing against 

the specimen edges was found. Figure 12-40 shows the most severe example of this edge 

damage. This specimen, BC92-4, later failed at 9669 hours but not at the edge notches. At the 

12-71 
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Figuro 12-40 G/E long-Tom Spoeimon 602-4 Showing Edge Damago (arrows) from Compression 

Stiffenor Scrows 

time of the first examination, the two epoxy systems had accumulated approximately 4500 

hours of exposure and the HT-Sl170 GlPI and BlAl systems approximately 1700 hours. The results of the examination of the four composite systems were as follows: 

a. 

b. 

C. 

d. 

B/E-The two heated zones on each specimen were discolored and on several of the 

unnotched specimens exposed and broken boron fibers were visible. Several specimens also 

had shallow edge notches from the stiffener grid screws. 

GIE-Some slight discoloring of the heated zones was visible and several specimens were 

edge notched from the stiffener screws. At the time of this examination several of the 

unnotched specimens had already failed. The unfailed ones, however, appeared to be in 

good condition. None of the notched specimens had failed, but three of the 10 showed considerable damage in the form of cracks and delaminations at the edges and at the holes. 

Figure 12-41 shows the specimen with the most extensive damage (BF91-4). This specimen 

was removed because failure appeared imminent. 

GIPI-After 1700 hours of exposure the only visible changes were very slight wear marks 

on the edges of one of the unnotched specimens. 

BIAl-These specimen had also been exposed for only 1700 hours and were in good condition. Three specimens had slight edge notching from the stiffener screws and there was 

some surface abrasion from the stiffeners. 

12-73 
Figure 12-41 G/E Long-Term Specimen BF91-4 Showing Cracks and Dolaminations (arrows) 

at Edges and Holes 

The next thorough examination was made just before the completion of 10,000 hours on the 

epoxy specimens (approximately 9000 hours for the epoxies and 6000 for the HT-SI710 GIPI 

and BIA1 specimens). The results of the examination of the four composite systems was as 

follows: 

a. BIE-These specimens showed the most damage as a result of the flight simulation exposures even though, in contrast to the GIE system, none had failed. The two heated zones on 

each specimen were discolored and in several instances showed exposed and broken fibers 

where portions of the epoxy resin had fallen away. In the notched BIE specimens some 

longitudinal surface cracks tangent to the holes were observed. 

b. G/E-At the time of this examination 11 of the 20 GIE specimens had failed or had been 

removed from the exposure apparatus because of showing significant damage. Of the 

remaining specimens the only visible effects were color changes in the heated zones at the 

top and bottom of each specimen and some small longitudinal cracks tangent to the holes 

in the notched specimens. These cracks were believed to extend only through the outer 0 O 

ply in both the B/E and GIE specimens. 

c. GIPI-These specimens appeared to be in excellent condition with only slight color changes 

in the heated zones. 

d. BIAl-The only visible effects were longitudinal stripes on the surface where some of the ' Teflon tape from the stiffening grids had adhered. 

Figures 12-42 through 12-45 are photographs of typical specimens of each composite material 

that were made during this examination with the specimens installed in the flight simulation 

apparatus. Only the BIE specimens, Figure 12-42, show any visible damage. 

12-74 
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Figure 12-42 B/E long-Term Specimens After Approximately 9000-hr of Exposure Showing Broken and 

Exposed Boron Fibers 

12.75 
Figure 12-43 G/E long-Term Specimens After Approximately 9000-hr of Exposure 

12-76 
Figure 12-44 G/PI long-Term Specimens After Approximately 6000-hr of Exposure 

12-77 
Figuro 12-45 B/Al Long-term Spwimonr Aftor Approximatoly 6000-hr of Exporuro 

12-78 
In setting the load levels for the long-term flight simulation tests, the expectation was for 80% 

of the specimens of each system to complete 50,000 hours of exposure. The 20% failure rate 

would amount to, on average, four specimens for each composite material system. In general, 

these failures would be expected to occur during Phase I1 of the test program. For the G/E 

system a much higher failure rate was observed. Twelve of the 20 G/E specimens either failed 

outright or were removed from test because of severe damage. Some of the badly damaged 

specimens were removed for further inspection and because of the likelihood of imminent 

failure were not returned. The others were in such poor condition that when a specimen in the 

same setup failed the damaged one was also removed to decrease the amount of equipment 

downtime. Table 12-21 lists the G/E specimens that did not survive the 10,000 hours of flight 

simulation and also shows the amount of exposure time log,@ when removed because of 

failure or significant degradation. 

Tabio 12-21. G/E Long-Tom Spocimonr Romovd Bocauso of Failuro or Significant Degradation 

Spocimon Exporuro 

No. TYPO (hr) 

Failuro or Damago 

Location 

BC91-1 

BC91-2 

BC91-3 

BC91-4 

BC91-5 

BC92-3 

BC92-4 

BC92-5 

BF91-2 

BF91-3 

BF91-4 

BF91-5 

Unnotched 

Unnotched 

UMOtched 

Unnotched 

unnotched 

unnotched 

UMOtched 

unnotched 

Notched 

Notched 

Notched 

Notched 

4515 

6683 

6683 

4524 

2696 

4053 

9669 

4524 

7520 

7520 

4549 

4549 

Heated zone: failed 

Heated zone: not failed 

Heated zone: failed 

Heated zone: not failed 

Heated zone: failed 

Heated zone: failed 

Heated zone: failed 

Near doubler: failed 

Heated zone: failed 

Heated zone: not failed 

Heated zone; not failed 

Heated zone: not failed 

All of the specimens except one failed in either the upper or lower heated zones. The specimens, 

which are 0.46 m (18 in.) long, are heated in two zones that are 0.11 m (4.5 in.) on either side of 

the centerline. Figure 12-46 shows three of the specimens, all of which failed in one of the 

heated sections. The two specimens with over 4000 hours of exposure also showed considerable 

damage in the heated zones opposite the failures in the form of edge delaminations. This can be 

seen in Figure 12-47. The 2669-hour specimen, BC91-5, did not show delamination in the 

unfailed heated zone. Another difference between the 2669-hour specimen and the two 

4000-hour specimens of Figure 12-46 is the appearance of the fracture. The 2669-hour specimen 

has failed in a manner similar to the unexposed baseline specimens with little loss of the epoxy 

resin from around the graphite tows. After 4000 hours of exposure, however, the epoxy resin 

12-79 
Figure 12-46 G/E long-Term Specimens That Failed During Flight Simulation Testing Before Completion 

of 10,000.hr of Exposure 

Figure 12-47 G/E long-Term Specimen 802-3 Showing Edgo Dolaminations in Heated Zone. (Specimen 

had failed at opposite heated zone after 4053-hr of exposure) 

12-80 
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matrix has begun to crumble during failure leaving many bare graphite tows. Similar effects 

were observed during tensile testing of the 450 K (350 O F) thermal age specimens after 5000 or 

more hours of exposure. Closeup views of the two different fracture appearances are shown in 

Figure 12-48 and Figure 12-49. The failure mechanism that was responsible for these early failures of the GIE specimens appeared to be a combination of several factors. Based on earlier 

thermal aging data the time at temperature, 4000 hours at 408 K (275" F), itself, is not sufficient to explain the brittle nature of the epoxy matrix. Added to this, however, are the effects 

of the thermal expansion stress during the 2000 thermal cycles and the tensile and compressive loads. As the matrix becomes embrittled as a result of oxidation the thermal and mechanical stresses act to produce microcracking and eventual loss of the outer epoxy layer. This 

can lead to accelerated oxidation as the filaments in the outer layers fail, thereby exposing the 

underlying resin to the oxidizing environment. The compressive load which occurs once per 

flight cycle is also important because of the lowered compressive resistance of the graphite 

tows as the matrix is degrading. It is these compressive loads that are responsible for the 

numerous delaminations observed in the heated zones of many of the specimens. A good example of a specimen that was about to fail by this mechanism is shown on the right in Figure 

12-50. The edge view of the same specimen, Figure 12-51, shows clearly the loss of matrix on 

both surfaces with the exposed and broken graphite filaments and the extensive delamination 

caused by the compressive loading. This specimen would undoubtedly have failed very soon if 

it had not been removed from test. 

The specimen on the left in Figure 12-50 was the only GIE specimen that did not fail in the 

heated zone. A closeup view of the fracture is shown in Figure 12-52. The arrows in the figure 

Figure 12-48 G/E long-Term Specimen BC91-5 Showing Fracture Appearance After Failure in 2669 hr. (No 

significant loss of resin from around the graphite tows has occurred) 

12-81 
Figure 12-49 G/E long-Term Specimen BO?-? Showing Fracture Appearance After Failure in 4515 hr. 

(loss of resin from around graphite tows is readily apparent) 

Figure 12-50 G/E long-Term Specimens That were Removed from Test Before Completion of 10,000 hr of 

Exposure 

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Figure 12-51 Edge View of G/E Long-term Specimen BC91-4 Showing Exposed and Broken Graphite 

Filaments and Extensive Delamination. (Spocimen was removed from test after 4524-hr) 

Figure 12-52 G/E Long-Term Specimen BC92-5 Showing Fracture Appearance After Failure in 4524-hr 

(Failure was probably the result of edge notches (at locations shown) caused by rotation 

of the compression stiffener grids) 

1283 
point to locations where stiffener grid screws could bear against the specimen edges if the 

grids rotated slightly. It appears likely that it was edge notches of this type that caused the 

early failure of specimen BC92-5. The fracture location and appearance were also quite dif- . ferent from the other failures. 

One other specimen failure occured during the long-term testing prior to completion of 10,000 

hours. This was unidirectional B/Al specimen EU91-1 which failed out of the heated zone after 

2454 hours. The specimen, shown in Figure 12-53, apparently failed because of a misaligned 

bearing in one branch of the whiffletree used to apply the fatigue loads. This caused the compressive loads to be misaligned with the axis of the specimen and resulted in the stiffening grid 

being pushed into the specimen at one end. Figure 12-54, a closeup of the fracture, shows where 

the four bearing surfaces of the stiffener were forced into the BlAl at the edge of the end 

doubler and initiated the failure. 

None of the B/E or G/PI long-term specimens failed during the 10,000 hours of flight simulation exposure in Phase I. 

At the completion of 10,000 hours of flight simulation exposure the Phase I long-term 

specimens were removed from the test apparatus. In addition, because of the high failure rate 

of the G/E and the extensive surface damage visible on the B/E, all remaining Phase I1 epoxy 

specimens were also removed. The revised flight simulation test plan for the BIE and GIE 

systems was described in Section 12.7.1. The specimens were examined visually with the aid of 

a stereo microscope and photographed. Based on the number and condition of the specimens 

Figure 12-53 B/Al Long-Term Specimen That Failed During Flight Simulation Testing Before Completion of 

10,OOO hr of Exposure 

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Figure 12-54 B/Al long-Term Specimen EU91-1 Showing Fracture Appearance After Failure in 2454 hr. 

(Failure was initiated at locations (arrows) where stiffener grid was forced into the B/Al) - ~ __ 

and information gained during the overall test program, a residual strength testing plan was 

prepared. As will be discussed later, the test plan was somewhat different from the one 

originally scheduled. Cutting plans were prepared and the long-term specimens were machined 

into residual strength specimens, which were then tested. The results of the specimen examinations and residual strength tests are presented in the following sections. 

12.7.2.1 Post Simulation Examinations - Phase I 

B/E, [OO f 4501,. unnotched and notched. All of the B/E specimens survived 10,000 hours of 

flight simulation exposure without fracture at temperatures to 408K (275 OF). The specimens 

did, however, experience considerable surface damage in the heated areas. The outer portions 

of the epoxy matrix had crumbled away in numerous places revealing the boron fibers, many of 

which had failed. In contrast to the G/E specimens no edge delaminations were observed at 

either the outside edges or at the holes in the notched specimens. Photographs showing both 

sides of three unnotched and three notched specimens are presented in Figures 12-55 to 12-58. 

The B/E laminates have a glass scrim cloth as the outer layer on one side. This side shows 

distinct color changes (to a light brown) in the heated zones. On the other side the epoxy resin 

has darkened somewhat, particularly in the heated zones. The effect of shielding of the surface 

by the compression stiffening grids is very evident. A few loosened and broken boron fibers are 

visible in Figure 12-55 and 12-57, the non scrim-cloth sides. A better view of these exposed 

boron fibers is shown in Figure 12-59 for an unnotched specimen and in Figure 12-60 for a notched specimen. These regions where the matrix has broken away at the surface and exposed 

the fibers were observed only in the heated zones in the areas not covered by the stiffening 

grids. Exposed fibers were also found in the heated areas on the scrim-cloth side of the B/E 

specimens, but these were generally hidden under the glass cloth. Figure 12-61 shows a closeup 

1285 
65321 7-255 

Figure 12-55 B/E long-Term Specimens After 10,000 hr of Flight Simulation Exposure 

Figure 12-56 B/E long-Torm Specimens After 10,000-hr of Flight Sim 

(Glass scrim cloth side) 

iulation Exposure 

12-86 
Figure 12-57 WE Na 

i 

kchod Long-term Spdmons Aftor 10,000-hr of Flight Simulation Exposure 

65321 7-258 

Figure 12-58 B/E Notched Long-Term Specimens After 10,000-hr of Flight Simulation Exposure 

(Glass scrim cloth side) 

12-87 
Figuro 12-59 B/E long-Torm Spocimen AC91-5 After 10,000-hr of Flight Simulation Exposure Showing 

loosened and Broken Boron Fibers 

Figure 12-60 B/E Notched long-Term Specimen AF91-5 After 10.000 hr of Flight Simulation Exposure 

Showing loosened and Broken Boron Fibers 

12-88 
view of the heated zone on the scrim-cloth side. The disturbed areas of the glass cloth are caused by loosened and broken boron fibers beneath the cloth. 

G/E, [OO f 4!P], unnotched and notched. only two unnotched and six notched GIE specimens 

from the original 20 Phase I and Phase I1 specimens survived the 10,000 hours of exposure at 

maximum temperatures of 408K (275 OF). The two unnotched specimens are shown in Figure 

12-62 and 12-63. Unlike the failed specimens these appeared to be in reasonably good condition. No surface or edge cracks or delaminations were found. The heated zones had undergone' 

some color changes, particularly between the stiffening grids. These areas had a worn or eroded looking appearance and, in several places, the graphite tows were very close to being exposed as a result of the loss of the epoxy matrix on the outer surface. The notched specimens that 

survived the 10,000 hours of flight simulation exposure were similar in appearance to the unnotched specimens except for some damage at the holes. Four of the specimens had 

longitudinal cracks near the holes and some delaminations at the edges of the holes. Two of the 

notched specimens that completed 10,000 hours are shown in Figure 12-64 and 12-65. 

Specimen BF92-3 is one with cracks at the holes but they are not readily visible in this 

photograph. The material on the surface of the specimens at one end is residue of the adhesive 

used to attach the thermocouples. 

One of the specimens that appeared to be in good condition during the second inspection with 

the grids removed (9404 hours of exposure) is shown in Figure 12-66. Compressive failure in 

one of the heated zones occurred soon after the inspection, 9669 hours. No evidence of the 

impending failure was seen during the examination after 9404 hours. A higher magnification 

view of this specimen and two others that were removed from test prior to 10,000 hours 

Figure 12-61 B/E long-Term Specimen AC91-4 After 10,000 hr of Flight Simulation Exposure. (Glass scrim 

cloth side. The disturbed areas of the glass cloth are caused by loosened and broken boron 

fibon below the cloth) 

12-89 
Figuro 12-62 G/E long-Term Spoeimonr Ahr 10,OOO hr of Flight Simulation Exposun 

Figure 12-63 

12-90 
Figuro 12-61 C/E Notehod long-Tom Spodmons After 10,OOO hr of Flight Simulation Exposuro 

Figure 12-65 G/E Notched long-Term Specimens After 10,000 hr of Flight Simulation Exposure (Opposite 

sides of specimens in Figure 12-64) 

12-91 
Figure 12-66 G/E Long-Term Spocirnon 802-4 After 9669 hr of Flight Simulation Exposure. (Spdmen 

failed (arrows) in compression at hmtd zone) 

because of severe damage is shown in Figure 12-67. The damage level in these specimens 

increases from left to right in the photograph. Inspection of the specimens suggests a failure 

mechanism slightly different from that proposed early in the test program and described in 

Section 12.7.2. In these specimens and also those in Figure 12-68, initial failure is thought to 

occur by delamination of the outer [OO] plies in the heated zones between the stiffening grids. 

The combination of thermal cycling stresses and the mechanical fatigue stresses acting at the 

ply interface overcomes the relatively low interlaminar strength, and delamination occurs. The 

outer plies retain their strength during the tensile loads, but during the one compressive load 

per cycle all the load is carried by the low strength +45 plies. At this point two effects occur. 

The 245 O plies begin to fail from the tensile and compressive fatigue loads, and oxidation with 

accompanying breaking up of the matrix of the outer plies is accelerated. The first specimen in 

Figure 12-67, BC91-2, was removed from test shortly after delamination of the outer plies 

occurred in the regions between the grids. Some damage to the inner +45 O plies was visible but 

the amount was small. The specimen still maintained its rigidity. In the regions that were 

beneath the 0.00318 m (0.125 in.) wide grids, transverse cracks can be seen. The second 

specimen, BC92-4, had sustained an increased amount of damage before being removed from 

test. Delamination and transverse cracking of the outer plies was much more extensive and the 

inner f45 O plies were badly broken. The specimen was no longer rigid and could easily be bent 

at the damaged heated zone. Some bare graphite fibers were also visible at the edges of the 

cracks and delaminations. These fibers were typical of those found in badly oxidized thermal 

aging test specimens, i.e., rather fuzzy appearing with no resin adhering to them. The fii 

specimen in Figure 12-67, BC91-4, shows the greatest degree of degradation of the three. In 

addition to delamination-of the outer plies and considerable inner ply damage, the surface has 

12-92 

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Flguro 12-67 G/E Long-Torm Speeimmr that wore Removed from T.) Won Comploilon of 10,000 hr of 

Exporun. (Increasing Iovdr of damage in the rp.cimonr aro visible in moving from loft to 

right) 

also experienced a great deal of tensile failure of the graphite fibers. Much of the epoxy resin 

matrix from the outer plies has oxidized and crumbled away, leaving bare graphite fibers 

exposed. This specimen had the least rigidity of the three. For these three specimens total time 

of flight simulation exposure was apparently not a factor in the onset of initial failure. 

A similar set of the notched specimens showing varying degrees of damage is pictured in 

Figure 12-68. The first, BF91-5, shows the onset of delamination of the outer 0" plies, some 

longitudinal cracking near the holes, and some slight damage to the inner [+45"] plies. The 

degree of delamination varies from one side of the specimen to the other. This can be seen from 

the closeup photographs of Figure 12-69. Severe delamination has occurred at one edge, a 

lesser amount in the center as shown at the three holes, and none at all at the opposite edge. 

The second specimen, BF91-4, had delaminated throughout the heated zone and shows 

increased longitudinal cracking and the onset of transverse cracking in the regions under the 

stiffening grids. The ridigity or resistance to bending at the heated zone has decreased 

somewhat. The third specimen, BF91-3, is very close to fracture with severe cracking of all the 

plies. The epoxy surface is heavily oxidized with many graphite fibers exposed. The specimen 

has almost no bending resistance across the notched region. The final specimen of the set, 

BF91-2, has fractured with considerable damage throughout the heated zone. In like manner to 

the 5000-and 10,000-hour 450 K (350" F) thermal aging specimens, many of the graphite fibers 

are visible as a result of the loss of the brittle oxidized epoxy resin. 

, 

12-93 
Figure 12-68 G/E Notched Long-Term Specimens that wore Removed from Test Bdoro Completion of 

10,OOO hr of Exposure. (Increasing Iovols of damage in tho specimons are visible in 

moving from left to right) 

Figure 12-70 shows the effect of exposure time on the fracture appearance with respect to loss 

of resin from the graphite fibers. In 2669 hours, oxidation has not embrittled the matrix sufficiently to cause the resin to break away from the fibers during failure. The second specimen, 

BC91-1, after exposure of 4515 hours begins to show the bare graphite fibers caused by the 

crumbling away of the oxidized matrix. The effect increases with exposure time as can be seen 

in the other two specimens in Figure 12-70. 

G/PI, [Oo -C GO],, unnotched and notched. All of the Phase I HT-S/710 G/PI long-term specimens completed 10,000 hours of flight simulation exposure at maximum temperatures of 505K 

(450°F). Photographs showing both sides of typical unnotched and notched specimens are 

presented in Figures 12-71 to 12-74. All of the specimens were discolored in the regions between the stiffening grids in the heated zones. All of the specimens also had exposed graphite 

fibers in these same areas. Generally the exposed fibers were found on the side of the specimen 

that faced the quartz heating lamps. However, for three specimens loose surface fibers were 

observed on both surfaces. Typically, these regions of exposed graphite fibers were approximately 0.025 m (1 in.) in length in the center of the heated zones. Figure 12-75 shows a closeup 

view of these damaged areas on an unnotched specimen and Figure 12-76 the same for a notched specimen. The amount of exposed fibers on the notched specimens was less than that found 

on the unnotched specimens. Some slight edge cracks were found in the heated zones of a few 

of the specimens. An example is shown in Figure 12-77. At 10,000 hours of exposure the edge 

cracks did not appear to be a serious problem. The appearance of the holes in the notched 

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Figuro 12-70 G/E long-Torm Spoeimons that Failod During Flight Simulation Tosting Woro Complotion 

of 10,000 hr of Exposun. (Spoeimons show progrossivo loss of rosin from tho graphito 

fibon with oxposum timo) 

.ORIGINAL PAGE 18 

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65321 7-271 

Figure 12-71 G/PI long-Term Specimens After 10,OOO hr of Flight Simulation Exposure 

12-96 
Figure 12-72 G/PI long-term Spocimens Aftor 10,000 hr of Flight Simulation Exposuro (Opposite sides of 

specimens in Figure 12-71) 

Figure 12-73 G/PI Notched long-term Specimens After 10,000 hr of Flight Simulation Exposure 

12-97 
ORIGINAL PAGE E3 

OE POOR QUAIJTX 

65321 7-274 

Figure 12-74 G/PI Notched long-Term Specimens After 10,OOO hr of Flight Simulation Exposure (Opposite 

sides of specimens in Figure 12-73) 

Figure 12-75 G/PI Long-Term Specimen DC91-3 After 10,000 hr of Flight Simulation Exposun Showing 

Exposed Graphite Fibers in Heated Zone 

12-98 
ORIGIN AI 

R r, 

QU 

Figure 12-76 C/PI Notchod Long-tom Spoeimon DF91.2 After 10,OOO hr of Flight Simulation Exporuro 

Showing Exporod Crophito Fibers in Hoatd Zono 

I 

r 

9x 663212277 

Figure 12-77 G/PI Long-Term Specimon DC91-1 Showing Slight Edge Cracks in Heated Zone After 10,000 

hr of Flight Simulation Exposure 

12-99 
specimens was generally quite good. Slight horizontal cracks tangent to the holes were observed in three of the specimens. These were very short and appeared to extend only through the 

outer ply. Some slight delaminations or cracks were present in some of the holes, but, again. 

these also appeared to be very shallow and not a serious problem at this time. 

B/Al, [OO f 4501, and [OO],, unnotched. All of the [0 O f 45 'Is crossplied specimens and four of 

the five unidirectional BlAl long-term specimens completed 10,000 hours of exposure at maximum temperatures of 505K (450OF). Front and back views of typical specimens are shown in 

Figures 12-78 and 12-79 for crossply material and in Figures 12-80 and 12-81 for unidirectional 

materiaL The heated zones on the sides facing the quartz heat lamps have been coated with flat 

black, heat resistant paint to reduce reflection from the aluminum surface and thereby increase the efficiency of the heaters. The only noticeable effects on these photographs are the 

vertical stripes where the stiffener grids were attached. These markings are residue from the 

Teflon coating on the grids, and in some areas, regions of wear where the soft aluminum surface has been eroded by the rubbing action of the stiffener during fatigue cycling. Careful examination of the surface in the heated zones reveals a roughened surface and some small 

shallow cracks. The effect is sipiiar to, but not as extensive as, that observed in the specimens 

but was more difficult to see on the painted side. A magnified view of the heated zone on the 

unpainted side of a crossplied specimen is shown in Figure 12-82. A photograph taken from the 

unheated center section of the same specimen is also shown in Figure 12-82 for comparison. 

The roughened surface was found on both the crossplied and the unidirectional specimens. One 

other observation was made on the unidirectional specimens. The specimens were fIat except 

in the heated zones where the surface was very irregular (wavy) in the transverse direction. 

Edge cracks and delaminations were found in two of the crossplied specimens, generally in the 

heated zones. Figure 12-83 is an enlarged view of the worst area found on these two specimens. 

An ultrasonic C-scan of the same specimen is shown in Figure 12-84. The regions marked on 

the C-scan are the areas of the specimen where edge cracks were observed with a stereo 

microscope. The correlation was quite good. The C-scan indicates the cracks to extend approximately 0.006 m (0.25 in.) into the specimen. 

I that were fatigue tested in air at 561 K (550 O F). The surface roughening occurred on both sides 

12.7.2.2 Residual Strength Data - Phase 1. The original test plan for residual strength determinations assumed that all five specimens of each set would complete 10,000 hours of exposure 

and that all of the specimen width would be available for cutting into subsize specimens. For 

two of the materials, G/E and unidirectional BlAl, all of the Phase I specimens did not survive 

10,000 hours. Also, earlier tests on material from the unfded end of a G/E specimen, BC92-3, 

that had failed in 4053 hours showed considerable variation in tensile properties across the 

width of the specimen. Specimens that were shielded by the stiffening grids gave tensile values 

of 551 MNlm2 (79.9 ksi) and 545 MNlrn2 (79.0 ksi) while a specimen that was cut from between 

the grids gave 163 MNlm2 (23.5 ksi). Because of these results the decision was made to cut the 

specimens from the unnotched long-term specimens from the material between the grids. This 

reduced the number of 0.013 m (0.5 in.) wide residual strength specimens that could be 

obtained from one long-term specimen from 10 to six. For the notched long-term specimens the 

entire width was used (=.was originally planned) so that the number of residual strength 

specimens remained the same (six per specimen). During the course of the test program, particularly for the thermal aging and the flight simulation tests, it became apparent that compressive and shear properties of the composites were of considerable importance in addition to 

12-100 
Figure 12-78 Crossply B/Al Long-Term Specimens After 10,OOO hr of Flight Simulation Exposure 

Figure 12179 Crossply B/Al long-Term Specimens After 10,000 hr of Flight Simulation Exposure (Opposite 

sides of specimens in Figure 12-78) 

12-101 
Simulation Exporuro 

LIGINAL PAGE E 

POOR QUALTTY 

Figuro 12-81 Unidiroctional B/Al long-Term Specimonr After 10,000 ht of Flight Simulation Exporuro 

(Opposite ridor of rpocimenr in Figuro 12-80) 

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W 

z 

0 

N 

0 

W 

W 

I 

12-103 
Figuro 12-83 B/Al long-Term Spwimon EC91-1 After 10,OOO hr of Flight Simulation Exposure Showing 

Edge Cracking in Heated Zone 

the tensile characteristics toward which the study was primarily directed. The combination of 

these three factors 1) less than 100% survival, 2) loss of material from under grids, and 3) 

increased importance of residual compressive and shear properties resulted in a change in the 

Phase I residual strength test plan. Table 12-22 shows both the original and the revised plans 

for the G/E and B/E systems. Similar information for the HT-S/710 G/PI and B/A1 systems is 

listed in Table 12-23. The number of tests from the GE unnotched long-term specimens was 

limited by the amount of material available (two and onehalf specimens). Two specimens survived 10,000 hours and the unfailed end of BC92-4, which failed in 9669 hours, was undamaged 

and considered usable. The number of tests from the G/E notched, B/E unnotched, and B/E 

notched long-term specimens was intentionally kept small because the poor condition of the 

specimens did not appear to warrant a full scale residual strength test program. The 10,000 

hour flight simulation exposure of the epoxy systems was repeated using less severe test conditions as described in Section 12.7.1. The revised test plans for the HT-S/710 G/PI and B/A1 

systems were developed to derive as much useful data as possible from a limited amount of 

test material, all of which was used. 

All of the 0.013 m (0.5 in.) wide unnotched tensile and fatigue and B/i notchedtensile and 

shear specimens were cut from the 0.01 m (0.7 in.) wide portions of the flight simulation 

specimens between the stiffening grids. The approximate locations of these specimens (six per 

long-term specimen) can be seen in Figure 12-81 where the residue of the Teflon tape has outlined the location occupied by the grid during flight simulation testing. Bonding of doublers 

was required on only one end of each of the residual tensile and fatigue specimens as the 

titanium doubler from the flight simulation specimen was still attached to the other end. 

12-104 OjtiGiNAL PAGE IS 

OR POOR QUAUTX 
ORIGINAL PAGE IS 

OF POOR QUALITY 

- 

W 

z 

0 

N 

0 

W 

IW 

I 

a 

- 

- 

W 

2 

0 

N 

0 

W 

k 

W 

X 

7 

- 

I 

11 

I 

.d m J 

c 

N 

0 

In 

In 

0 

0 

9 

z 

t 

a 

* e 

c 

L. 

t 

op 

2 

2 

iz 

3 

.o 

12-105 
Tablo 12-22. Raidual Sirmgth Tost Plan for Pharo I B/E and G/E Aftor 10,000 hr of Flight Simulation Tdng 

Numbor of Spoeimmr 

Tomporaturm Matorial Flight Simulation Tost Original Rovirod 

Syrtom Spoeimon Typo TYPO K (OF) Plan Plan 

BIE [0" f 45O], Unnotched 

[O" f 45 '1, Notched 

GIE [Oo C 45O], Unnotched 

[0" f 45 OIs Notched 

Tensile 

Notched Tensile 

Fatigue 

Compressive 

Tensile 

Notched Tensile 

Notched Tensile 

Notched Fatigue 

Notched Tensile 

Tensile ' 

Notched Tensile 

Fatigue 

Compressive 

Tensile 

Notched Tensile 

Notched Tensile 

Notched Fatigue 

Notched Tensile 

297 

297 

297 

297 

450 

450 

297 

297 

450 

297 

297 

297 

297 

450 

450 

297 

297 

450 

75 

75 

75 

75 

350 

350 

75 

75 

350 

75 

75 

75 

75 

350 

350 

75 

75 

350 

10 

10 

10 

0 

5 

5 

10 

10 

5 

10 

10 

10 

0 

5 

5 

10 

10 

5 

6 

0 

9 

7 

0 

0 

6 

9 

0 

4 

0 

7 

2 

0 

0 

6 

9 

0 

The notched resin matrix flight simulation specimens were 0.073 m.43 in.) wide with three 

equally spaced 0.0064 m (0.25 in) diameter holes across the width. Shearing of the specimens 

would produce three identical 0.025 m (1 in) wide residual strength specimens with a centrally 

located hole. For composite material, however, the specimens must be cut with a saw having a 

finite blade width, approximately 0.0015 m (0.06 in.). Consequently, the notched tensile and 

fatigue specimens could be cut from the flight simulation specimen in at least three different 

ways: equal widths with offset holes: unequal widths with centered holes: and equal widths 

with centered holes. A check of reference 28 showed that the various configurations had very 

little difference in K, factors. The one selected was specimens of equal widths with offset holes 

(J121&x in the two outer specimens. The difference in (J values for the specimen with a centered 

hole and one with an offset hole was negligible (3.28 versus 3.30). Again, bonding of doublers 

was required on only one end. 

The residual compressive tests were performed using the specimen and test procedures 

described in Section 7. To get a sufficiently thick specimen for compressive testing from the 6 

ply flight simulation specimens it was necessary to bond three or four layers together. From 

these 18 or 24 ply panels the 0.0064 m (0.25 in.) wide compressive specimens could be cut. Two 

12-106 
tabio 12-23. Rnidual Stnngth tost PIon for Phaso I HTW710 G/PI and B/AI Aftor 10,OOO hr of Flight 

Simulation tosting 

Numlaor of Spocimons 

tomporaturo Motorial Flight Simulation ter) Original Revised 

Sythm Spocimon Typo TYPO K (OF) Plan Plan 

HT-SI170 GlPI [0° 2 45 "Is Unnotched Tensile 

Notched Tensile 

Fatigue 

Compressive 

Tensile 

Notched Tensile 

[Oo 2 45 '1: Notched Notched Tensile 

Notched Fatigue 

Shear 

Notched Tensile 

BIN [0" f Go], Unnotched Tensile 

Notched Tensile 

Fatigue 

Compressive 

Tensile 

Notched Tensile 

Tensile 

Notched Tensile 

Fatigue 

Shear 

Compressive 

Tensile 

Notched Tensile 

Shear 

[Oo]6 Umotched 

297 

297 

297 

297 

505 

505 

297 

297 

297 

505 

297 

297 

297 

297 

505 

505 

297 

297 

297 

297 

297 

505 

505 

505 

75 

75 

75 

75 

450 

450 

75 

75 

75 

450 

75 

75 

75 

75 

450 

450 

75 

75 

75 

75 

75 

450 

450 

450 

10 5 

10 5 

10 9 

0 4 

5 5 

5 0 

10 10 

10 10 

0 4 

5 5 

10 4 

10 4 

10 8 

0 4 

5 4 

5 4 

10 4 

10 4 

10 7 

5 3 

0 4 

5 0 

5 0 

5 0 

specimens, made by lamihating three or four 0.01 m (0.75 in.) wide pieces together, used the 

entire width of one end of a flight simulation specimen. Four compressive specimens, using one 

entire flight simulation specimen, were made for the crossplied B/E, GIPI, B/AL and the unidirectional BlAl materials. The limited amount of GIE available allowed only two compressive 

specimens of this material Since there were no failures of the B/E flight simulation specimens 

and the number of residual strength specimens was already reduced, sufficient material was 

available for another type of laminRted compressive specimen. These were made by bonding 

four pieces together that were cut from the areas between the grids. Two flight simulation 

specimens provided enough material for three of the 0.0064 m (0.25 in.) wide compressive 

specimens. These two types of B/E compressive specimens were intended to show the effect of 

the stiffening grids on shielding the composite material during flight simulation exposure. 

Residual interfiber shear strength tests were performed on the unidirectional B/Al material ' 

No unidirectional resin matrix specimens were included in the flight simulation test program, 

12-107 
however, and, therefore, no residual shear strength tests were originaUy scheduled. Enamination of the failed G/E flight simulation specimens indicated that interlaminar shear strength 

played an important role in the premature failures and obtnining residuai shear properties 

would be very useful. There was one half of a 10,000-hour notched HT-S/170 G/PI simulation 

specimen that was not used for residual strength testing, and a method was devised to use this 

material to fabricate four short beam shear specimens. As shown in Figure 12-85 three strips 

were cut from between the grid marks in the heated zone and bonded together to make an 18 

ply laminate. The four short beam shear specimens, 0.015 m (0.60 in.) by 0.0064 m (0.25 in.), 

were cut from material just above and below the holes. As the heated zone in the simulation 

specimen was 0.05 m (2 in.) long and the holes 0.0064 m (0.25 in.) in diameter, there was sufficient room for cutting the specimens above and below the holes. The layout of the specimens 

was as shown in Figure 12-85. 

B/E, [OO f 450],, unnotchd and notched. Residual tensile, notched tensile, and compressive 

data for B/E after 10,000 hours of flight simulation exposure are listed in Table 12-24. Elastic 

modulus values are included for all unnotched tensile and compressive tests. Compressive data 

are shown for the two methods of test specimen fabrication described earlier. Baseline tensile, 

notched tensile, and compressive values are also included in Table 12-24. Residual axial fatigue 

data at room temperature and a stress ratio, R, of 0.1 for unnotched and notched specimens are 

presented in Table 12-25, and S-N curves are plotted in Figures 12-86 and 12-87. Also included 

on the mphs are S-N mes (no data points) for unexposed material tested under the same 

conditions (presented eariier in Section 11, Baseline Fatigue Testing). 

The unnotched tensile strength was considerably lowered as a result of the 10,000 hours of flight 

simulation exposure while the notched strength was lowered only 10%. The notched properties 

of the resin matrix systems have generally shown smaller decreases than the unnotched after 

exposure. This may, in part, result from the method of specimen cutting where portions of the 

notched specimens had been shielded by the stiffening grids, whereas the unnotched specimens 

were cut from material between the grids. Compressive strength was degraded slightly more 

than the unnotched tensile strength The residual compressive data also show the effect of the 

stiffening grid in protecting the underlying material. Specimens cut from the entire width had a 

compressive strength approximately onethird greater than specimens cut from between the 

grids. As shown in Figures 12-86 and 12-87, the fatigue properties were considerably lowered by 

the long-term exposure cycling. 

WE, [OO f SO],, unnotched and notched. Residual tensile, notched tensile, compressive, and 

elastic modulus values for G/E after 10,000 hours of flight simulation exposure are listed in 

Table 12-26. Data for unexposed material are also included in the table. Residual axial fatigue 

data at room temperature and a stress ratio, R, of 0.1 for unnotched and notched specimens are 

presented in Table 12-27, and S-N curves are plotted in Figures 12-88 and 12-89. S-N curves for 

unexposed material are shown on the fatigue plots for comparison. 

Eight unnotched and four notched long-term specimens of GIE did not complete 10,000 hours 

of exposure, either because of complete failure or significant damage in the heated zones. In 

spite of this the residual tensile properties, particularly for the unnotched specimens, were not 

severely degraded. The compressive strength was, however, greatly affected. After 10,000 

hours of flight simulation the compressive strength was reduced by approximately onehalf. 

This decrease would be anticipated if the failure mechanism proposed for the specimens that 

12-108 
~~ 

I 

I 0- . 

1 

I 

1 

r------- ---- 1 

\ 

2 

s 

L 

J 

LL 

12-1 09 
Tablo 12-24. Strongth and Modulus Proportfos of [OO f 45O], B/E at 297 K (750 F) Woro and Ahr 10,000 hr 

of Flight Simulation Exposun 

Strongth Modulus Spocimon Flight Simulation 

Numbor Spocimon Numbor MN/d (ksi) GN/d (Msi) 

Unnotched Tensile: 10,000 hr 

A912-1 

-2 

A913-1 

-2 

A9141 

-2 

AC91-2 392 

-2 373 

-3 498 

-3 400 

-4 362 

4 337 

av 394 

Baseline av 618 

Notched Tensile: 10.000 hr 

AF913-1 AF91-3 

-2 -3 

-3 -3 

AF915-1 -5 

-2 -5 

-3 -5 

452 

411 

412 

315 

366 

424 

av 397 

Baseline av 461 

Compressive: 10.000 hr (entire specimen width used) 

A9C-1 

-2 

-3 

-4 

AC91-1 834 

-1 934 

-1 834 

-1 1086 

av 918 

Compressive: 10,000 hr (cut from between grids) 

A9C-5 AC91-5 741 

-6 -5 & AC92-5 709 

-7 AC92-5 754 

. av 735 

Baseline av 1540 

56.8 

54.1 

72.2 

58.0 

52.5 

48.9 

57.1 

89.7 

65.6 

59.6 

59.7 

45.7 

53.1 

61.5 

57.5 

64.2 

120.9 

135.4 

120.9 

154.9 

133.0 

107.5 

102.9 

109.3 

106.6 

223.4 

68.9 10.0 

73.8 10.7 

86.2 12.5 

76.5 11.1 

71.7 10.4 

60.0 8.7 

72.8 10.6 

94.5 13.7 

84.1 12.2 

82.0 11.9 

86.2 12.5 

82.7 12.0 

83.8 12.2 

79.3 11.5 

72.4 10.5 

82.0 11.9 

77.9 11.3 

102 14.8 

12-110 
tablo 12-2S. Rosidual Axial Fatigue Data for [OO * 45O], WE at 297 K (750 F) After 10.000 hr of Flight 

Simulation Exposure, R = 0.1 

Maximum Stmr Spoeimon Flight Simulation Spocimen 

Numbor Spocimon Numbor TYPO MN/m2 (kri) Cyc1.S Comments 

A912-3 

A913-3 

-4 

-5 

-6 

A914-3 

4 

-5 

-6 

AF9 13-4 

-5 

-6 

AF9 14- 1 

-2 

-3 

AF915-4 

-5 

-6 

AC91-2 

-3 

-3 

-3 

-3 

-4 

4 

-4 

4 

AF91-3 

-3 

-3 

4 

-4 

-4 

-5 

-5 

-5 

unnotched 310 

276 

293 

276 

310 

293 

241 

259 

293 

Notched 345 

414 

379 

362 

345 

328 

310 

293 

276 

45 

40 

42.5 

40 

45 

42.5 

35 

37.5 

42.5 

50 

60 

55 

52.5 

50 

47.5 

45 

42.5 

40 

<1,000 

7,975,000 

47,000 

5,872,000 

<1,000 

600,000 

15,234,000 Did not fail 

10.245.000 Did not fail 

10,000,000 Did not fail 

1.000 

1.000 

<1.000 

< 1,000 

< 1,000 

< 1,000 

149,000 

1,213,000 

12,729,000 Did not fail 

did not survive the flight simulation exposures was correct. The residual fatigue properties 

were significantly lowered with the unnotched material again experiencing a greater decrease 

than the notched. 

G/PI, [W -C 450],, unnotched and notched. Results of the 297 K (75 O F) and 505 K (450 O F) tensile tests on uMotched and notched HT-S/710 G/PI after 10,000 hours of flight simulation exposure are presented in Tables 12-28 and 12-29. Also included are the baseline data for comparison purposes. Two types of notched tensile specimens were tested. The first, Table 12-28, 

was cut from unnotched flight simulation specimens and was notched after exposure. The second, Table 12-29, was cut from the notched flight simulation specimens after being exposed in 

the notched condition. Table 12-30 gives the compressive and short beam shear data for G/PI 

before and after flight simulation exposure of 10,000 hours. Results of the residual fatigue 

tests on the unnotched and notched GlPI material at 297 K (75" F) and a stress ratio, R, of 0.1 

are listed in Table 12-31 and plotted as S-N curves in Figures 12-90 and 12-91. Data for unexposed GlPI tested under the same conditions are shown on the fatigue plots for comparison. 

12-111 
500 

cuE 

z 

- 

400 W 

2 

u 

>. 

u 

W 

p. 

a 

rA 

300 a 

- 70 

E 

c 

rn 

= 

0 u - 0 m 

EXPOSED 

7r 

- 30 3 

I1111111 I I I I1111 I I 111111 1 1 I11111 I I I I IIII 20 

103 1 04 105 106 107 

NUMBER OF CYCLES 

65321 7-286 

Axial Fatiguo Proportios of [V -C 45O], B/E Boforo and Aftor 10,000 hr Flight Simulation 

Exposuro. 297 K (75OF), R=0.1, Unnokhod 

500 

cuE 

z 

I 

\ 

- 

400 u 

>- 

u 

CT 

w 

0 

rA 

rA 

w 

a 300 

s 

a 

z 

z 

x 

< 

z 200 

I I I11111 I 111111 I I I11111 I 11111 

102 103 1 04 105 1 

NUMBER OF CYCLES 

170 

-40 

30 % 

-uLlulI 20 

i 107 

653217-287 

Figuro 12-87 Axial Fatiguo Proportios of [OO f 4!P]s B/E Boforo and After 10,000 hr of Flight Simulation 

Exposure, 297 K (75OF), R=0.1, Notched 

12-112 
Toblo 12-26. Shngth and Modulus Proportla of [OO f 45O], G/E ot 297 K (750 F) Bdon and Aftor 10,000 hr 

of Flight Simulation Exporuro 

Shngth &aduiur Spocimon Flight Simulation 

Numbor Spoccimon Numbor MN/d (kri) GN/d (Msi) 

Unnotched Tensile: 10,000 hr 

B921-1 

-2 

B922-1 

-2 

BC92-1 429 

-1 427 

-2 410 

-2 411 

av 419 

Baseline av 496 

Notched Tensile: 10.000 hr 

BF911-1 

-2 

-3 

BF923-1 

-2 

-3 

BF91-1 

-1 

-1 

BF92-3 

-3 

-3 

413 

445 

377 

388 

470 

362 

av 409 

Baseline av 412 

Compressive: 10.000 hr (entire specimen width used) 

B9C-1 

-2 

BC92-1 

-1 

316 

293 

av 304 

Baseline av 570 

62.2 

62.0 

59.5 

59.6 

60.8 

71.9 

59.9 

64.6 

54.7 

56.3 

68.2 

52.5 

59.4 

59.7 

45.8 

42.5 

44.2 

82.7 

39 5.7 

46 6.7 

46 6.6 

50 7.2 

45 6.6 

57 8.3 

46 6.6 

47 6.8 

46 6.7 

46 6.6 

12-113 
tablo 12-27. Rosidual Axial Fatfguo Data for [go f 45O], G/E at 297 K (750 F) Aftor 10,OOO hr of Flight 

Simulation Exporun, R = 0.1 

~ 

Maximum ttrosr Spocimon Flight Simulation Spocimon 

Numbor Spocimon Numbor TYPO MN/d (ksi) Cycles Commonts 

3 

B922-3 

-4 

-5 

-6 

B924-1 

-2 

-3 

BF925-1 

-2 

-3 

-4 

-5 

-6 

BF923-4 

-5 

-6 

BC92-2 

-2 

-2 

-2 

BC924 

-4 

-4 

BF92-5 

-5 

-5 

-5 

-5 

-5 

-3 

-3 

-3 

UMOtChed 379 

345 

362 

328 

310 

276 

224 

Notched 276 

310 

345 

328 

362 

379 

328 

362 

241 

55 <1,000 

50 6,968,000 

52.5 3,000 

47.5 1,000 

45 4,000 

40 18,000 

32.5 10,466,000 Did not fail 

40 

45 

50 

47.5 

52.5 

55 

47.5 

52.5 

35 

1,829,000 

1,143,000 

2,000 

2,571.000 Bond failure 

112.000 Bond failure 

1,000 

630,000 

1,662,000 

5,778,000 

12-114 
653217-288 NUMBER OF CYCLES 

500 I 

- 

0 FA1 LEO TEST SECT1 ON 

0 DOUBLE BONO FAILURE h( 

\ E 

z - 

0 UNEXPOSED 5 400 w 

--I 

u 

> 

0 0 0 

- a 

2 300 --P-~- 1 

- a 

z 

7 

X 

0 Lu 0 0 0 

El 

2 

cn 

0 

0 

200 - 

I I I11111 I 1 I I1111 I 1 I I1111 I 1 1 I I Ill 1 I I IIILL20 

Figuro 12-88 Axial Futiguo Propod08 of [OO f A5O], G/E Wore and Aftor 10,000 hr of Flight Simulation 

Exporun, 297 K (75O F), R = 0.1, Unnotehod 

70 

z 

60g 

+z 

s 

- 

c 

v) 

50: m 

cn 

v) 

W 

m n40 c) 

< 

c) 

Im 

30 2 

x 

12-115 
Tablo 12-28. Strength and Modulus Propodos of [OO f 45O], Ht=S/7lO G/PI Bofon and Aftor 10,OOO hr of 

Flight Simulation Exposun. (Unnotchod flight simulation rpocimon) 

* Tomporatun Strongth Modulus 

Spocimon Flight Simulation 

Numbor Spdmon Numbor K (OF) MN/m2 (kri) GN/d (Msi) 

Unnotched Tensile: 10,000 hr 

D912-1 DC91-2 

-2 -2 

-3 -2 

-4 -2 

-5 -2 

D914-6 4 

D915-1 -5 

Baseline 

D912-6 DC91-2 

-1 -3 

-2 -3 

-3 

-4 -3 

Baseline 

Notched Tensile: 10,000 hr 

D915-2 DC91-5 

-3 -5 

-4 -5 

-5 -5 

-6 -5 

Baseline 

297 75 

av 

av 

505 450 

av 

av 

297 75 

av 

av 

405 

33 1 

225 

3 23 

510 

423 

372 

370 

519 

334 

411 

303 

225 

296 

314 

470 

385 

198 

336 

346 

199 

293 

357 

58.8 

48.0 

32.7 

46.8 

74.0 

61.4 

54.0 

53.7 

75.3 

48.5 

59.6 

43.9 

32.7 

43.0 

45.5 

68.2 

55.8 

28.7 

48.8 

50.2 

28.8 

42.5 

51.8 

47 

46 

41 

46 

48 

46 

46 

53 

36 

54 

42 

37 

41 

42 

47 

- 

6.8 

6.6 

5.9 

6.7 

6.9 

6.7 

6.6 

7.7 . 

5.2 

7.8 

6.1 

5.4 

6.0 

6.1 

6.8 

- 

4 

12-116 
Tablo 12-29. Notchod tensilo Sfrmgth of [P f Go], HT-S/710 G/PI Bofon and Ah 10,OOO hr of Flight 

Simulation Exporun. (Notehod flight simulation spocimon) 

.) Spocimon Flight Simulaiion tomporntun Strmngth 

Numbor Spocimon Numbor K (OF) MN/d (ksi) 

Notched Tensik 10.000 hr 

DF911-1 

-2 

-3 

-4 

-5 

-6 

DF9 12- 1 

-2 

-3 

DF914- 1 

DF91-1 297 75 

-1 

-1 

-1 

-1 

-1 

-2 

-2 

-2 

-4 

av 

Baseline av 

DF9143 

-4 

-5 

-6 

DF915-1 

DF91-4 

4 

-4 

4 

-5 

505 450 

av 

Baseline av 

348 

427 

342 

315 

355 

247 

261 

323 

376 

271 

326 

357 

256 

294 

377 

397 

336 

332 

336 

50.5 

61.9 

49.6 

45.7 

51.5 

35.8 

37.8 

46.9 

54.5 

39.3 

47.4 

51.8 

37.2 

42.7 

54.7 

57.6 

48.7 

48.2 

48.8 

While no HT-SI710 G/PI specimens failed during the 10,000 hours of flight simulation exposure, the unnotched tensile strength at both room and elevated' temperatures decreased by 

about onethird. The residual notched tensile data appear to indicate a difference in the 

average amount of degradation depending on whether the notch was machined in the specimen 

before or after exposure. Those specimens that were notched after exposure showed a greater 

average decrease in strength. Further analysis of the data, however, indicates that the location 

from which the residual notched tensile specimen was cut from the flight simulation specimen 

may be more important. Six notched tensile specimens are cut from a flight simulation 

specimen, two from the center and four from the outer edges (see Figure 12-10 for cutting plan). 

The center specimens are shielded during esposure by two metal strips of the compressive stiffening grids while the outer specimens are each shielded by only one. The locations of these 

strips with respect to where the six residual strength specimens are cut can be seen in Figure 

12-65. In addition to having two, presumably, higher strength portions, the center specimens 

are symmetric with respect to these shielded portions. The outer specimens have a shielded 

portion on only one side of the center hole, and consequently, may experience uneven loading, 

12-117 
Tablo 12-30. Compnirivo and Short ham Shwr Strongth Proportioa of [OO f Go], HT-S/710 G/Pl at 297 K 

(750 F) Bdon and Aftor 10,OOO hr of Flight Simulation Exporuro 

Stnngth Modulus Spoeimon Flight Simulation 

Numbor Spocimon Numbor MN/d (kri) GN/d (MA) 

Compressive: 10,000 hr 

D9C-1 DC91-1 

-2 -1 

-3 -1 

-4 -1 

Baseline 

Short Beam Shear: 10.000 hr 

D9S1 DF91-5 

-2 -5 

-3 -5 

-4 -5 

Baseline 

326 47.3 41 5.9 

299 43.3 41 6.0 

264 38.3 39 5.6 

274 39.7 41 5.9 

av 291 42.2 40 5.8 

av 377 54.7 46 6.7 

18 2.6 

18 2.6 

16 2.3 

17 2.4 

av 17 2.5 

av 37 5.4 

Tablo 12-31. Rosidual Axial Fatiguo Data for [OO = 45O], HT-S/7lO G/PI at 297 K (75O F) Aftor 10.000 hr of 

Flight Simulation Exporuro, R = 0.1 

Maximum Strosr Spocimon Flight Simulation Spoeimon 

Numbor Spocirnon Numbor TYPO MN/d (ksi) Cycler Comments 

D913-5 

-6 

D914-1 

-2 

-3 

-4 

-5 

DF9 12-4 

-5 

-6 

DF9 13- 1 

-2 

-3 

-4 

-5 

-6 

D F9 14-2 

DC91-3 

-3 

4 

-4 

-4 

4 

-4 

DF91-2 

-2 

-2 

-3 

-3 

-3 

-3 

-3 

-3 

-4 

Unnotched 293 

241 

207 

172 

190 

198 

190 

Notched 276 

270 

283 

345 

310 

310 

293 

293 

293 

293 

42.5 

35 

30 

25 

27.5 

28.75 

27.5 

40 

39.1 

41.1 

50 

45 

45 

42.5 

42.5 

42.5 

42.5 

< 1,000 

< 1,000 

< 1,000 

12.262,OOO Did not fail 

10,256,000 Did not fail 

10,269,000 Did not fail 

15,405,000 Did not fail 

10,219,000 Did not fail 

< 1.000 

1,099,000 

< 1.000 

4,000 

1.000 

53,000 

2,424.000 

< 1.000 

6,000 

12-118 
PiE 

1300 

=E 

W 

-J 

u 

z 

a 

w 

cn 

W 

200 

a 

k 

2 9 100 

a 

- 

UNEXPOSED 

J 

50 

E 

D 

x, 

-40: 

cn 

--I 

a 

Figuro 12-90 Axial Fatiguo Proporti08 of [OO .C 4501, Hf-S/710 G/PI Before and Aftor 10,000 hr of Flight 

Simulation Exporuro, 297 K (75OF), R=0.1, Unnotchod 

1 1 1 I I Ill 1 I I11111 

400 

(Y 

E 

z 

a 

2 300 

- 

- 

w 

u 

?i 

a 

w 

a 

v) 

v) 

200 

t; 

r 

a 

5 

- 

X 20 

a 

100. 

. I I111111 I 1 I11111 11111111 I I 1 I1111 I I 1 11I~lO 

102 103 1 04 105 106 107 

EXPOSED 

c) 

- 20 

R 

- 2. 

- 10 

I I I11111 I I I11111 I I I I1111 0 

x 

P 

x 

E’ 

= c 

rn 

VJ 

VJ 

NUMBER OF CYCLES 653217-291 

Figure 12-91 Axial Fatigue Properties of [Oo f 4501, HT-S/710 G/PI Before and After 10,000 hr of Flight 

Simulation Exposure, 297 K (75OF), R=0.1, Notched 

12-119 
which could initiate premature failure. To check out these effects the outer specimen (dash 

numbers 1.3.4. and 6) and center specimen (dash numbers 2 and 5) notched tensile strengths 

are compared in Table 12-32. The data are listed by temperature, without regard to whether 

the notches were introduced before or after flight simulation. For the room temperature tests, 

where sufficient data are available there is a distinct difference in the two types of specimens. 

The more heavily shielded center specimens were approximately 25% higher in notched tensile 

strength. In like manner to the G/E material the GlPI experienced a significant reduction in 

compressive strength after 10,000 hours of exposure. The decrease was not as great as that for 

G/E, however, approximately 25% as compared to 50%. It is interesting to note that the loss in ' 

tensile strength was greater than that for compressive strength. Again, however, specimen 

location may be playing a role as the compressive specimens were prepared from the entire 

width of the specimens while the tensile specimens were cut from the unshielded areas. The 

residual short beam shear data indicate an apparent major reduction in interlaminar shear 

strength has occurred as a result of the 10,000 hours of flight simulation esposure. This 

degradation, if real, correlates well with the failure mechanism proposed earlier for the GIE 

specimens. The data, however, are limited, and the effect of bonding 6 ply laminates together 

to form 18 ply short beam shear specimens requires additional study to determine the usefulness of the results. The unnotched fatigue properties were greatly affected with a lowering of 

the S-N curve of approximately 40% over the entire range from lo3 to 107 cycles. The notched 

fatigue strength, on the other hand, was just slightly lowered by the 10,000 hours of esposure. 

B/A, [OO f 4501, and [00l6, unnotched. Results of the room and elevated temperature tensile 

tests on unnotched and notched BlAl before and after 10,000 hours of flight simulation espe 

Tablo 12-32. Effod of Spocimon Location on Rosidual Notchod fenrilo Strongth of [P f 45O], HT-S/7lO G/PI 

Aftor 10,000 hr of Flight Simulation Exporun 

Outor Spoccimon Centor Spocimen 

Spocimon Numbor K (OF) MN/rn2 (ksi) MN/m2 (ksi) 

Tomporaturo Strongth Strongth Flight Simulation 

DC91-5 

-5 

-5 

DF91-1 

-1 

-1 

-1 

-2 

-2 

-4 

DF91-4 

-4 

-4 

'-5 

297 75 198 

336 

199 

348 

342 

315 

247 

261 

376 

271 

av 289 

505 450 256 

294 

397 

336 

av 321 

28.7 

48.8 

28.8 

50.5 

49.6 

45.7 

35.8 

37.8 

54.5 

39.3 

41.9 

37.2 

42.7 

57.6 

48.7 

46.6 

385 

346 

427 

355 

323 

367 

377 

377 

55.8 

50.2 

61.9 

51.5 

46.9 

53.2 

54.7 

54.7 

12-120 
sure are presented in Tables 12-33 and 12-34. Residual compressive data for both crossply and 

unidirectional BlAl are presented in Table 12-35. No baseline compressive tests were conducted on either of the BlAl layups. Also included in Table 12-35 are interfiber shear strength 

data from unidirectional BlAl before and after flight simulation esposure. Results of the 

residual fatigue tests on the unnotched crossply and unidirectional B/Al material at 297 K 

(75 O F) and a stress ratio, R, of 0.1 are listed in Table 12-36 and plotted as S-N curves in Figures 

12-92 and 12-93. Fatigue curves for unexposed BlAl tested under the same conditions are also 

included in Figures 12-92 and 12-93 for comparison. 

tablo 12-33. Unnotchod and Notehod Tonsilo Propod- of [go k 4501, B/Al Bdoro and Aftor 10,000 hr of 

Flight Simulation Exposun 

tompomturo Strongth Modulus Spoccimon Flight Simulation 

Numbor Spocimon Numbor K (OF) MN/m2 (ksi) GN/m2 (Msi) 

Unnotched Tensile: 10.000 hr 

E C911-1 EC91-1 297 

-2 -1 

-3 -1 

-4 -1 

Baseline 

EC911-5 EC91-1 505 

-6 -1 

EC912-1 -2 

-2 -2 

Baseline 

Notched Tensile: 10,000 hr 

EC912-3 EC91-2 297 

-4 -2 

-5 -2 

-6 -2 

Baseline 

EC913-1 EC91-3 505 

-2 -3 

-3 -3 

-4 -3 

Baseline 

75 

450 

75 

450 

409 

399 

406 

30 1 

av 379 

av 516 

264 

415 

404 

359 

av 360 

av 452 

303 

478 

478 

364 

av 406 

av 479 

283 

303 

38 1 

354 

av 330 

av 443 

. 59.3 103 15.0 

57.9 134 19.4 

58.9 160 23.2 

43.7 159 23.1 

55.0 139 20.2 

74.8 154 22.4 

38.3 - - 

60.2 143 20.7 

58.6 - - 

52.1 - - 

52.3 

65.6 121 17.6 

44.0 

69.4 

69.3 

52.8 

58.9 

69.5 

41.4 

44.0 

55.3 

51.3 

47.9 

64.3 

12-121 
Tablo 12-34. Unnotchod and Notehod tonsilo Propodia of [PI6 B/AI at 297 K (750 F) Beforo and Aftor 10,000 

hr of Flight Simulatlon Exposun 

Strongth Modulus Spocimon Flight Simulation 

Numbor Spocirnon Numbor MN/d (ksi) GN/d (Msi) 

Unnotched Tensile: 10.000 hr 

EU913-1 EU91-3 

-2 -3 

-3 -3 

-4 -3 

Baseline 

Notched Tensile: 10,000 hr 

EU913-5 EU91-3 

-6 -3 

EU9 14-1 -4 

-2 -4 

aBaseline 

1070 155 231 33.9 

972 141 240 34.8 

1000 145 170 24.6 

1290 187 172 25.0 

av 1080 157 203 29.5 

av 1450 210 198 28.7 

869 126 

1080 157 

896 130 

924 134 

av 942 137 

av 1220 177 

aEstimated from notched tensile strength versus temperature curve for [O"], BIAl. 

Figure 7-23. 

12-122 
tablo 12-35. Comprossivo and Intorfibor Shoar Strongth Propodia of B/Al at 297 K (750 F) Bdoro and Aftor 

10,000 hr of Flight Simulation Exporuro 

Strongth Modulus Spoeimrcl Flight Simulation 

Numbor Spoeimon Numbor Oriontation MN/m2 (kri) GN/d (Msi) 

Compressive: 10.000 hr 

E C9C- 1 EC91-5 

-2 -5 

-3 -5 

-4 -5 

EU9C-1 EU92-4 

-2 -4 

-3 -4 

-4 -4 

Interfiber shear: 10.000 hr 

EU915-4 EU91-5 

-5 -5 

-6 -5 

Baseline 

[Oo 2 45 "1, a >910 

a>1250 

a>1420 

a>1020 

av 

10 ai6 

1550 

1870 

av 1570 

(a) 

[o ai6 69.6 

72.4 

70.3 

av 70.8 

av 93.1 

225 

27 1 

(a) 

227 

10.1 

10.5 

10.2 

10.3 

13.5 

143 

147 

117 

112 

130 

212 

197 

207 

207 

206 

20.8 

21.3 

17.0 

16.3 

18.8 

30.7 

28.5 

30.0 

30.0 

29.8 

aDoubler bond shear failure. 

12-123 
Tablo 12-36. Rosidual Unnotchod Axial Fatiguo Data for B/Al at 297 K (750 F) Afhr 10,000 hr of Flight 

Simulation Exposuro, R = 0.1 

Maximum Strosr Sp.cimor~ Flight Simulation 

Numbor Spocimon Numbor Oriontation MN/d (kri) Cycior Commonts 

E C913-5 

-6 

EC914-1 

-2 

-3 

-4 

-5 

-6 

EU914-3 

-4 

-5 

-6 

EU915-1 

-2 

-3 

EC91-3 

-3 

-4 

-4 

-4 

-4 

-4 

-4 

EU91-4 

-4 

-4 

-4 

-5 

-5 

-5 

[0" -C 45'1, 345 

379 

310 

396 

276 

362 

414 

293 

827 

689 

827 

862 

965 

1000 

1070 

50 

53 

45 

57.5 

40 

52.5 

60 

42.5 

120 

100 

120 

125 

140 

145 

1% 

824,000 

130.000 

1,789,000 

7,000 

190,000 

38,000 

2,986,000 

- Equipment 

Malfunction 

16,022,000 Did not fail 

13,704,000 Did not fail 

2,953,000 

2,377.000 

27,000 

133.000 

28,000 

The BlAl system exhibited a loss in mechanical properties of approximately onefourth after 

10,000 hours of flight simulation exposure. This loss was generally observed for both unnotched 

and notched tensile strength of both crossplied and unidirectional layups and for interfiber shear 

strength of the unidirectional mated Tests conducted at both room and elevated temperatures gave similar results. Doubler bond shear failures were a problem with the BlAl compressive specimens. The crossply specimens were particularly bad with all failing prematureiy in 

the doubler bonds. The failure loads were quite high, however, equal to a compressive stress in 

one of the crossply specimens of 1420 MNlm2 (206 ksi). The residual fatigue results for the 

crossplied material showed a decrease of, again, approximately onefourth in the runout stress 

(107 cycles) after 10,000 hours of exposure. The unidirectional BIA1, on the other hand, showed 

very little change from the baseline fatigue data. 

PHASE I SUMMARY 

Table 12-37 summarizes the 10,000-hour, long-term flight simulation residual strength data 

obtained from the four composite systems. The percentage decrease in the various mechanical 

properties is listed for each materiaL The fatigue data are summarized in Table 12-38. In this 

table the fatigue strength at lo7 cycles, expressed as a percentage of the unexposed tensile 

strength, is compared for unexposed material and for that which has survived 10.000 hours of 

flight simulation exposure. 

12-124 
W 

>- 

u 

w 

d 4011 

a 

1- I Illlll 

cn 

cn 

w 

a 

s.3011 

- 

v 

O\ - 

- 

I I111111 I I I11111 1 I 1 I1111 1 I 11 Ill 

H 

a 

5 

X 

u 

a 

u 2 700 

5 600 

z s 

U 

a 

cn 

w 

a 

500 

U 

a 

200 

V rn 

WJ 

WJ 

D 

m 

n 

--c c) 

r 

m 

-80 & 

<,)loo c) 

k 

1 I I I I Ill I I I 1 I Ill 1 1 I1 1111 I 1 I I I Ill I I I 111 11 60 

1 

70 

L 

B 

60 

c 

z 

cn 

V 

rn 

n 

!O 

NUMBER OF CYCLES 65321 7-292 

Figure 12-92 Axial Fatigue Propodes of [W f 4501, B/Al Before and After 10,000 hr of Flight Simulation 

Exposure. 297 K (75OF). R= 0.1, Unnotcched 

Figure 12-93 Axial Fatigue Properties of [Ool6 B/Al Before and After 10,000 hr of Flight Simulation 

Exposure, 297 K (75OF), R=0.1, Unnotched 

12-125 
Tablo 12-37. Summary of 10,000-hr Flight Simulation Residual Strongth Duta 

T.mporaturo Material Propody Loss 

Systom Prop.* - K (OF) 

550514 B/E 

[O" 2 45'1, 

A-Si3501-5 GIE 

[0" -C 45 "1, 

. I 

HT-S/710 G/PI 

(0' -C 45'1, 

BlAl 

[O" & 45 OIs 

BIN 

[O "16 

Ftu 

Et 

Fa 

E, 

Ftu 

Et 

Fa 

EC 

FtU 

Et 

Ftu 

Et 

aNotched F, 

aNotched Ftu 

CNotched Ftu 

aNotched F,, 

aNotched Ftu 

Fa 

EC 

Fsu 

Ftu 

Et 

Ftu 

CNotched F,, 

CNotched Ftu 

Ftu 

Et 

Fsu 

CNotched F,, 

297 

297 

297 

297 

297 

297 

297 

297 

297 

297 

297 

297 

505 

505 

297 

297 

505 

297 

297 

297 

297 

297 

505 

297 

505 

297 

297 

297 

297 

75 

75 

75 

75 

75 

75 

75 

75 

75 

75 

75 

75 

450 

450 

75 

75 

450 

75 

75 

75 

75 

75 

450 

75 

450 

-75 

75 

75 

75 

36 

23 

10 

b40-52 

"18-24 

15 

20 

1 

47 

0 

29 

14 

33 

10 

18 

8 

'1 

23 

13 

54 

26 

10 

20 

15 

26 

25 

0 

23 

24 

aNotched flight simulation specimens 

bLoss greater for material cut from between grids 

CUnnotched flight simulation specimens 

12-126 
tablo 12-37, Summary of 10,000-hr Flight Simulation Rosidual Stnngth Data - Concludod 

Matorial Tompomturo Prop.* Lost 

Systom Prop.* 3 K (OF) (%I 

CNotched F,, 297 75 

Fsu 297 75 

aNotched flight simulation specimens 

 LOSS greater for material cut from between gritis 

Wnnotched flight simulation specimens 

23 

24 

tablo 12-38. Strongth at Fatiguo Limif or lo7 Cyclor Exprossod as Percontogo of Ultimato fensilo Strength 

Befon and After 10,000 hr of Flight Simulation Exposuro, R = 0.1, 297 K (750 F) 

Matorial Sprcimon Fotiguo Strongth at 107 Cycler 

Systom Configuration Condition (% of Unexposed F,,,) 

5505/4 BIE 

[O" f 457, 

A-S13501-5 G/E 

[Oo f 45"], 

HT-S/710 G/PI 

[Oo f 45O], 

B/Al 

[O" -C 45'1, 

B/Al 

[O "16 

Unnotched 

Notched 

.Unno tched 

Notched 

Unnotched 

Notched 

Unnotched 

Unno tched 

Unexposed 

Exposed 

Unexposed 

Exposed 

Unexposed 

Exposed 

Unexposed 

Exposed 

Unexposed 

Exposed 

Unexposed 

Exposed 

Unexposed 

Exposed 

Unexposed 

Exposed 

52 

45 

92 

62 

90 

49 

92 

69 L 

60 

37 

84 

78 

74 

54 

57 

53 

12-127 
12.7.3 PHASE II LONG TERM TEST RESULTS. The Phase I1 flight simulation test program consisted 

of the following: 

B/E and G/E. New spechem were prepared and flight simulation testing of the epoxy systems 

was restarted with the maximum test temperature reduced to 373K (212°F). All specimens 

were unnotched. Specimens successfully completed 10,000 hours and 25,000 hours of exposure 

without failure and were tested for residual strength properties. Two of the B/E specimens 

failed after about 33,000 hours. The three remaining B/E specimens and the five G/E specimens were removed after about 36,000 hours for residual strength testing. 

HT-S/710 G/PI. Exposure testing was continued on the unnotched and notched specimens in 

the flight simulation apparatus that had completed 10,000 hours of exposure. In addition, 

three unnotched specimens were prepared from the second batch of prepreg (same material as 

used on the G/PI new material screening evaluation) for 25,000 hours of exposure. Two of these 

25,000 hour specimens were damaged while being installed in the simulator. The third failed 

prematurely after only 5760 hours. The screening evaluation had shown this batch of material 

to have very poor thermo-oxidative resistance so that the early failure was not unexpected. 

However, at the same time as this failure two of the 50,000 hour unnotched specimens in 

the same bay of the simulator also failed (15,896 hours). The reason for these failures could 

not be determined but did not appear to be normal wearout from the thermal and fatigue 

environment. 

One other premature failure occurred to a notched specimen as a result of a tensile overload at 

20,398 hours. This failure was traced to operator error. The remainder of the HT-S/710 GlPI 

specimens failed at times ranging from 15,021 hoursto 39,856 hours for the unnotched and 

17,852 hours to 27,890 hours for the notched. These failures were believed to have been the 

result of wearout from the thermal and fatigue environment. 

WAI. Flight simulation testing was continued on the unidirectional and crossplied specimens 

that had completed 10,000 hours in Phase I. In addition, three unidirectional and three crossplied specimens were prepared for evaluation after 25,000 hours of exposure cycling. One 

major incident occurred where six crossplied specimens failed during an accidental compressive overload at 20,957 hours. The accident was the result of the combination of equipment 

complexity and operator error. The specimen failures were located at one end about 0.013m 

(0.5 in.) from the titanium doubler. Three tensile specimens, 0.013m (0:5 in.) by 0.15m (6 in.), 

were prepared from the failed end of one of the specimens (see Figure 12-94) and tested at room 

temperature. The results along with the baseline and 10,000 hour residual strength data are 

given in Table 12-39. As the long-term specimens still retained over 50% of their tensile 

strength it was decided to continue flight simulation testing with the revised specimen configuration show in Figure 12-94. This configuration proved to be satisfactory with the three 

25,000 hour specimens completing testing and the other three lasting about 30,000 hours 

before failure. 

With the exception of the 25,000 hour BlAl specimens, four of which survived the 25,000 

hours, all of the remaining specimens failed to complete 50,000 hours. Failure times varied 

from 17,709 hours to 31,447 hours for the crossply material and 12,605 hours to 31,584 hours 

for the unidirectional material. With the exception of the 12,605 hour specimen, all of the other 

failures were believed to have been the result of wearout from the thermdfatigue cycling. 

12 -128 
B/AL LONG TERM SPECIMENS 

k0.051 (2) -I k0.051 (2) 

LOCATION OF / COMPRESSIVE FAILURE 

ABOUT 0.013 (0.5) --cf t-- FROM TITANIUM OOUBLER 

MACHINE 

3 TENS1 LES E-3 CUT SPECIMEN AT MIDPOINT 

-- III BOND LONG TITANIUM DOUBLER 

I L0.15 (6)--4 

TO THE GOOD HALF 

I 1 \ UIMCIY~IUIY~ IIY m tin.) 

653.271.397 

Flgun 12-94 Plan for Salvaging Croupliod B/Al Simulation Spocimrns Damagod by Compressive Ovrrload 

Ceiion 6000/LARC- 160 G/PI. Ten specimens each of unnotched and notched configuration were 

prepared from crossplied material. 

Two problems occurred at the start of exposure testing of this system. The 

failure of two specimens, as discussed earlier in Section 12.6.3 was corrected 

the peak loads. The second was a series of doubler bond failures after 0.5 hour, 

hours. 

first, premature 

by recalculating 

13 hours. and 24 

12-129 
tablo 12-39. Unnotchod Tonsilo Strongth of [P 2 Go], B/Al Bdoro 

and Ahr Flight Simulation Exposuro 

Strong* 

MN/IT? (kri) 

Baseline 

10,000 

20.957 

516 74.8 

379 55.0 

268 38.8 

In an attempt to solve the second problem all of the specimens were removed from the flight 

simulation apparatus and new sets of titanium doublers were bonded to their ends. Removal of 

the original doublers was quite difficult, and some surface damage occurred on 80% of the 

specimens. The location, type, and extent of damage was noted and recorded for each 

specimen. Surface preparation techniques were the same as those used successfully during 

Phases I, but a new batch of the HT-424 adhesive was used. One change was the addition of a 

clamping device near the tip of the doublers in the tapered portion. This technique was used 

successfully on the BlAl specimens where the high flight simulation loads had caused numerous doubler failures during the short-tern test program. After installation of the clamps no 

further doubler bond failures occurred. 

With the exception of the two early failures, all of the Ceiion 60001LARC-160 specimens completed 10,000 hours of exposure and were tested for residual strength properties. 

As the Phase I1 specimens of the five composite systems were removed from simulation testing (as a result of completing 10,000 hours or 25,000 hours, because of failure or extensive 

damage, or, as for the epoxy specimens, after 36,374 hours to ensure sufficient material for 

residual strength testing), they were exarnined visually with the aid of a stereo microscope and 

photographed. Residual strength testing plans were formulated including the use of the unfailed 

ends of failed specimens. Cutting plans were prepared and the long-term specimens were 

machined into residual strength specimens, which were then tested. The results of the specimen examinations and residual strength tests are presented in the following sections. 

12.7.3.1 Post-Simulation Examinations - Phase ti. 

WE, [OO f 450],, unnotchd. The replacement specimens of BIE completed 10,000 hours of 

flight simulation exposure at 373K (212OF). All of the specimens survived the 10,000 hours 

without fracture. The condition of the specimens was excellent. There were no loose surface 

fibers and no edge cracks or delaminations as had been observed on the Phase I specimens exposed at 408K (275°F). Distinct color changes were visible on the B/E specimens at the heated 

zones on both sides. Photographs showing both sides of three of the 10,000 hour specimens are 

presented in Figures 12-95 and 12-96. Examination of the surface of the specimens at 20 to 

lOOX magnification revealed a network of fine cracks in tkc heated zones but none in the areas 

that were not heated. Similar mim&g was found on the heated zones of the Phase I longterm specimens of BE and G/E after 10,000 hours but not on the HT-S/710 G/PI specimens exposed at 505K (450°F) for 10,000 hours. Neither the 10,000 hour the4 aging specimens or 

the 100 and 200 hour short-term specimens of either B/E or G/E experienced microcracking. 

12-130 
ORIGINS PAGE IS 

OF POOR QUFiguro 

. I 

Figure 

12-95 

12-96 

WE long-Term Spa dmom After 10,OOO hr 

5- :I? 3K4 

of Flight Sin dation 

WE long-term Spwimons After 10,OOO hr of Flight Simulation 

(Opposite sides of specimens in Figure 12-95) 

Exposuro - 

Exposure - 

373K 

* 373K 

(212OF) 

1 (212OF) 

12-131 
These observations indicate that some combination of extended thermal exposure, thermal 

cycling and fatigue loading are necessary for the initiation of surface microcracks in the epoxy 

maThe replacement specimens of B/E also compieted 25,000 hours of flight simulation exposure 

at 373K (212°F) without fracture. The condition of the specimens, as shown in Figures 12-97 

and 12-98, was excellent. There were no loose surface fibers and no edge cracks or delaminations as had been observed on the Phase I specimens exposed at 408K (275°F) for much 

shorter periods of time. The only readily noticeable effects were distinct color changes at the 

heated zones on both sides. Again, examination of the surface of the specimens at 20 X magnification revealed a network of fine cracks in the heated zones similar to those seen after 10,000 

hours. In gene4 the 25,000 hour specimens were very similar in appearance to the specimens 

that had been removed after 10,000 hours of exposure. 

Two of the B/E specimens failed during flight simulation exposure after passing the 33,000 

hour mark. As shown in Figure 12-99, failure occurred in the heated zone for both specimens 

and appeared to have been the result of wearout from the thermal and fatigue environment. 

Figure 12-100 shows a closeup view of one of the fractures with numerous bare longitudinal 

boron fibers visible. 

The failure mechanism for the two BIE long-term specimens, in like manner to that previously 

described for the Phase I G/E long-term failures, was a combination of several effects. Matrix 

degradation, in the form of microcracking in the heatediones, occurred through a combination 

of oxidation and fatigue loading. Both were necessary, ie., microcracking was not observed on 

thermal aging specimens (no applied stresses) nor was it found to any extent in the center portion of the flight simulation specimens (no heating). As the microcracking progressed with exposure time, small portions of resin became loosened and the outer boron fibers (longitudinal) 

were no longer supported. During the one compressive load in each 2 hour cycle the outer [0 ") 

fibers that were free could no longer carry the compressive load, which was dumped into the 

weaker [45 "1 fibers. The microcracking and loss of resin also opened additional paths for oxidation into the inner plies, further destroying the load-bearing capabilities of the composite. 

Delamination of the outer [0 "1 plies as a result of thermal cycling stresses and fatigue loading 

may also have contributed to failure. This effect was not as evident in the BIE failures as it was 

in the Phase I GIE failures, however. Examination of the lower heated zones (unfailed ends) of 

these two specimens revealed some loosened and broken boron fibers similar to those observed 

on the Phase I BIE specimens after 10,000 hours. Lowering the maximum flight simulation 

temperature from 408K (275 OF) to 373K (212 OF) greatly delayed the onset of surface microcracking sufficient to e-xpose the outer layer of boron fibers. 

After 36,374 hours of exposure some loose surface fibers were observed on the remaining three 

specimens. To ensure sufficient material for residual strength testing, the flight simulation 

tests were halted for both epoxy systems. The likelihood of completing 50,000 hours appeared 

to be small with the two earlier failures after 33,000 hours. The three specimens are shown in 

Figures 12-101 and 12-102. In general, the specimens were in good condition. Examination at 

25 x showed some microcracking in the heated zones with occasional loose boron fibers. There 

were distinct color changes but no edge cracking. 

Metallographic and SEM examinations were conducted on BIE long-term specimens exposed 

for 10,000, 25,000, and 36,374 hours. In like manner to the thermal aging specimens, no time 

dependent oxidation similar to that for G/E specimens was observed. This can be seen in the 

12 -132 
Figuro 12-97 B/E Long-Torm Spimons Aftor 25,000 hr of Flight Simulation Exposure - 373K (212OF) 

- -~ 

Figure 12-98 WE Long-Torm Spocimons Aftor 25,000 hr of Flight Simulation Exposure - 373K (212OF) 

(Opposite sides of sp.cimons in Figure 12-97) 

12-133 
41tlMIlS 

- - __ 

Figure 12-99 WE Long-Torm Spoeimons that Failod During Flight Simulation Testing 

Elofor. Compldon of 50,000 hr of Exposuro 

. . . - 

Figuro 12-100 WE Long-Torm Spimon AC95-3 Showing Fracturo Appoaranco After Failure in 33,063 hr 

(Loss of rosin from around boron fibers is readily apparent) 

12-134 
CLI.-I 44 

Figure 12-101 B/E Long-term Specimens After 36,374 hr of Flight-Simulation Exposure - 373K (212OF) 

Figure 12-102 B/E Long-Term Specimens After 36,374 hr of Flight Simulation Exposure - 373K (212OF) 

(Opposite sides of specimens in Figure 12-101) 

12 -135 
SEM photomicrographs in Figure 12-103. The effect of relief polishing around the individual 

fibers that increases with the degree of oxidation of the epoxy resin, as found for the GIE 

system, is not present in this series of photos. What can be seen, however, is an increase in 

matrix cracking with flight simulation exposure. The effect is first visible in the 25,000 hour 

specimen with the loss of small pieces of resin at many of the fiber-matrix interfaces. After 

36,374 hours resin loss has increased, and considerable matrix cracking is present between the 

fibers. Based on the thermal aging results for BIE thermal exposure alone is not sufficient to 

cause either surface microcracking or matrix microcracking. The flight simulation results 

show that the combination of fatigue loading and thermal exposure are responsible for the 

microcracking observed in the B/E system. Thermal cycling may also contribute to the effect. 

t 10,000 HOURS t t 25,000 HOURS t 

t 36,374 HOURS 

A R ROWS IN DTCAXSPEC I M EN ED G ES 

t 

653217-406 

Figure 12-103 Photomicrographs (SEM) of B/E After Flight Simulation Exposure - 373K (212OF), lOOX 

12-136 
G/E, [OO k 450],, unnotched. The replacement specimens of GIE performed even better than the 

B/E during flight simulation exposure at 373 K (212°F). All 13 specimens completed exposure 

cycling (36,374 hours) without failure. Comparing this with the Phase I G/E results (eight of 10 

failures at 408K (275°F) in 10,000 hours) shows the great influence that the maKimurn test 

temperature had on exposure life. Both the 10,000 hour and the 25,000 hour specimens were in 

excellent condition at the completion of exposure. There were no loose surface fibers, edge 

cracks, or delaminations. Inspection with a stereo microscope revealed a network of fine cracks 

in the heated zones of both the 10,000 hour and 25,000 hour sets of specimens. After 36,374 

hours of flight simulation the specimens showed increased effects of the exposure. In the 

heated zones numerous graphite fibers were visible where the outer layer of epoxy resin ap 

peared to have eroded away. Figure 12-104 shows a magnified view of one of these regions of 

exposed fibers. Increased microcracking and color changes were also present in these 

specimens. Even so, the condition of the specimens as shown in Figures 12-105 and 12-106 was 

quite good for 36,374 hours of testing. 

As discussed in the thermal aging section, the extent of oxidation of the G/E could be ready 

detected by metallographic techniques, especially with the SEM. To study the degree of oxidation, as well as the amount of microcracking of the matrix as a result of flight simulation exposure, specimens were examined after 10,000, 25,000, and 36,374 hours. Figures 12-107 and 

12-108 are metallographic and SEM results showing the entire cross section for the three exposure periods. The metallographic photomicrographs, Figure 12-107, while not showing the 

effects of oxidation as clearly as the SEM, do show the increase in microcracking with exposure 

time. These cracks appear to initiate at the surface and have penetrated the outside ply after 

25,000 hours of exposure. 

After 36,374 hours, cracks are also observed in the second plies although not always lined up 

with outside ply cracks. Very little microcracking of this type was observed in the G/E thermal 

Figure 12-104 Magnifid View of Surface of G/E Specimen BC95-3 After 36,374 hr of Flight Simulation 

Showing Loose Graphite Fibers 

12-137 
*ut: la) 

Figure 12-105 G/E Long-Term Specimonr Athr 36,374 hr of Flight Simulation Exporuro - 373K (212OF) 

ORIGINAL PAGE E 

OF POOR QUALITY 

Figure 12.106 G/E Long-Term Spoeimonr After 36,374 hr of Flight Simulation Exposure - 373K (212°F) 

(Opposite sides of specimonr in Figure 12-105) 

12 -138 
t 10,000 HOURS t t 25,000 HOURS t 

ARROWS INDICATE SPECIMEN EDGES 653217-410 

Figure 12-107 Photomicrographs (Metallograph) of G/E After Flight Simulation 

Exposure - 373K (212OF). lOOX 

aging specimens. The SEM photomicrographs, Figure 12-108, in addition to showing the 

extent of microcracking, also clearly show the increase in oxidation with exposure time. The 

lighter cast, beginning at the specimen edges and spreading inward with exposure time, was 

shown in the thermal aging study to be related to oxidation of the epoxy resin. 

The changes from dark to light are caused by the amount of relief polishing around the individual fibers that increases with the degree of oxidation. When the epoxy resin oxidizes, it is more 

12-139 
t 10,000 HOURS t t 10,000 HOURS I 

t 36,374 HOURS t 

ARROWS INDICATE SPECIMEN EDGES 653217-411 

Figuro 12.108 Photomicrographs (SEM) of G/E Aftor Flight Simulation Exposure - 373K (212OF), lOOX 

prone to crumble, and differences in the amount of oxidation in the polished mounts can easily 

be detected using the SEM. As would be expected, oxidation proceeds inwardly along the 

cracks at a faster rate than through the matrix. This effect is clearly shown in Figure 12-108. 

Slight oxidation at the outer edges is visible after 10,000 hours, increases somewhat after 

25,000 hours, and is considerable after 36,374 hours. Higher magnification views of the outer 

plies of specimens aged for these three time periods are shown in the SEM photomicrographs 

of Figure 12-109. The effect of oxidation on the amount of relief polishing around the graphite 

12 -140 
fibers shows the slight increase in oxidation from 10,000 hours to 25,000 hours and the considerable increase after 36,374 hours. 

In comparison with thermal aging results, the flight simulation exposures at 373K (212 OF') pm 

duced less oxidation than either the 394K (250°F) ambient or 450K (350OF) reduced pressure 

exposures after comparable the periods. The effect of the fatigue loading and the thermal 

cycling during fight simulation testing produced matrix microcracking but did not appear to 

increase the rate of oxidation nody experienced during thermal aging. 

HT-S/710 G/PI [OO f 4501,. unnotched and notched. In contrast to the excellent performance of 

the G/E mated, all of the HT-S/710 G/PI specimens failed during flight simulation exposure 

at maximum temperatures of 505K (450°F) or were removed from test because of such poor 

condition that failure was imminent. Table 12-40 lists the GlPI long term flight simulation 

specimens with their respective failure times. Three of the specimens in Table 12-40 failed at 

the same time, one after less than 6000 hours and the other two just prior to 16,000 hours. An 

investigation of the specimens and the circumstances of the incident indicated that the initial 

lablo 12-40. Summaq of HT-S/7lO G/Pl Flight Simulation Exposuro Spocimon Failum 

Spdmon Spdmon Expour- Hours lost S.ri.8 

Numbu TYP at Failun (H-4 Commonts 

DC92-1 

DC92-2 

DC92-3 . DC92-4 

DC92-5 

DC93-1 

DC93-2 

DC93-3 

DC91-S 

DF92-1 

DF92-2 

DF92-3 

DF92-4 

DF92-5 

DF92-S 

unnotched 

unnotched 

Unnotched 

Unnotched 

unnotched 

UMOkhed 

unnotched 

Unnotched 

unnotcfied 

Notched 

Notched 

Notched 

Notched 

Notched 

Notched 

15,896 

15,896 

39,856 

39,856 

19,400 

0 

0 

5,760 

15,021 

18,302 

18,302 

27,890 

20,398 

27,890 

17,852 

50,000 

50,000 

50,000 

50,000 

50,000 

25,000 

25,000 

25,000 

SP50,000 

50,000 

50,000 

50,000 

50,000 

spare 

(a) 

(a) 

Wearout failure in heated zone 

Wearout failure in heated zone 

Weamut failure in heated zone 

Damaged during installation 

Damaged during installation 

(a) 

Wearout failure in heated zone 

(b) 

Wearout failure in heated zone 

(b) 

Failure in tensile overload 

Wearout failure in heated zone 

Wearout failure in heated zone 

(a) a failed at the same time, see text. 

(b) Did not fail but removed from test because of poor condition 

12 -141 
3HGINAL PAGE IS 

OF POOR QUALITY 

failure was to the 6000 hour specimen. This specimen had been fabricated from the screening 

test material that had previously shown poor thermu-oxidative properties. Failure of the other 

two specimens could not be explained but did not appear to have been the result of normal 

wearout from the thermal and fatigue cycling. One other specimen, DF92-4, also failed 

prematurely after 2C,398 hours. Failure was traced to a tensile overload caused by operator error. The bulk of the HT-S/710 G/PI flight simulation failures appeared to be wearout failures 

from the thermal and fatigue exposure cycling. 

Figure 12-110 shows three of the failed specimens. The failures were generally of two types. 

The earlier failures, e.g., DC92-5 in Figure 12-110, showed more shear transfer between plies 

than the later failures, DC92-3 and DC92-4 in the same figure. The failure mode for the longer 

time specimens was more of a flat fracture that was straight across the width, transverse to 

the loading direction. Somewhat similar fracture behavior was also shown by the notched 

specimens as shorn in Figure 12-111. Here, however, the fracture paths were considerably 

influenced by the three holes in the center of the heated zones. 

Examination of the HT-SI710 GIPI specimens with a stereo microscope showed that the surface microcracking found in the heated zones of the epoxy specimens was not prebent, Some 

edge cracking was vfsible, increasing with exposure time. The major effect of the flight simulation was shown by the loss of polyimide resin at the surface of the heated zones with exposure 

of the underlying graphite fibers. The amount of loose fibers increased with exposure time as 

shown in Figures 12-112 and 12-113 for specimens cycled 19,400 hours and 39,856 hours 

respectively. The sp&nen failures, usually at the upper heated zones, were characterized by 

considerable loose fibers and a yellow-orange color at the fracture. In like manner to the thermal aging specimens, these fracture areas also did not show the fuzzy graphite fiber appearance 

caused by embrittlerllent of the resin matrix. 

10,000 HOURS 25,000 HOURS 36,374 HOURS 

65321 7-412 

Figure 12-1 19 Photomicrographs (SEM) of G/E After Flight Simulation Exposure - 

Outside Plies - 373K (212OF), 2500X 

12 -142 
Figuro 12-1 10 HT-S/710 G/PI Long=Tenn Spocimons That Failed During Flight Simulation Testing 

Beforo Completion of 50,OO hr of Exposure 

I -- 

Figure 12-1 11 HT-S/7lO G/PI Notched Long-Term Specimens That Failed During Flight Simulation' 

Testing Before Completion of 50,000 hr of Exposure 

12-143 
Figure 12.112 HT-SnlO G/PI Specimen DC92-5 After 19,400 hr of Flight Simulation Exposure Showing 

Exposed Graphite Fibers in Heated Zones of Unfailed and Failed Ends 

Figure 12-1 13 HT-S/710 G/PI Specimen DC92-3 After 39,856 hr of Flight Simulation Exposure Showing 

Exposed Graphite Fibers in Heated Zones of Unfailed and Failed Ends 

12-144 
The notched specimens were similar in appearance to the unnotched as to the degree of loose 

surface fibers, edge cracking, and color changes. Close-up views of the fracture regions of two 

specimens are shown in Figure 12-114. The nature of the failure for the specimen on the right of 

the figure, DF92-5, is characteristic of the flat fractures seen for the longer exposure times. 

The other specimen, DF92-4, failed accidently from a tensile overload. Two of the notched 

specimens did not fail but were removed from test because of being in very poor condition. 

Closeup views of the heated zones where considerable damage had occurred are shown in 

Figure 12-115. An almost continuous series of cracks across the width of each specimen can be 

seen in the photographs. These specimens offered almost no resistance to bending in these 

regions. 

In addition to edge cracking near the heated zones, the notched specimens also contained 

numerous cracks or delaminations at the edges of the 0.0064m (0.25 in.) holes. Again, the 

degree of cracking at the holes increased with exposure time. Figure 12-116 shows closeup 

views of the three holes at the unfailed end of one of the last notched specimens to be removed 

from test. This specimen exhibited the most cracks in the regions of the holes of any of the notched specimens. 

Metallographic and SEM examinations were conducted on HT-S/710 G/PI specimens after 

10,000,19,400,27,890, and 39,856 hours of flight simulation exposure. The results were similar 

to those found for the same material after thermal aging at 505K (450OF) where no metallographic changes were observed until after 50,000 hours. The flight simulation specimens 

showed neither the matrix cracking or surface oxidation effects typical of those previously 

described for the G/E flight simulation specimens. 

Figure 12-117 contains SEM photomicrographs of cross sections of HT-S/710 G/PI flight 

simulation specimens after various exposure times. The unexposed as-received material is 

shown in Figure 8-35. Because of the length of the program, the photomicrograph were pre 

pared at different times by different SEM operators. The as-received and 10,000 hour photomicrographs were shot in the secondary electron mode, the others in the backscattered mode. 

No changes in the microstructure could be found in the course of the 39,856 hours of exposure. 

Ceiion 6000/LARC-160 G/PI, [Oo .C 45O],, unnotched and notched. All of the Celion 

60001LARC-160 G/PI specimens completed 10,000 hours of flight simulation exposure at maximum temperatures of 450K (350 OF) without failure. At the completion of 10,000 hours, all the 

specimens were removed from the simulation apparatus for examination prior to residual 

strength testing. Visually, the specimens were in very good condition as shown in Figures 

12-118 and 12-119 for the unnotched configuration and Figures 12-120 and 12-121 for the notched. There were some color changes but generally less than those observed for the epoxy 

samples after similar exposure times. No loose surface fibers were present in the heated zones 

of any of the specimens. There were a few edge cracks and some cracking or delamination at 

some of the holes in the notched specimens. Generally these cracks were not severe. Also found 

were some surface cracks in the heated zones of both the unnotched and notched specimens. 

None of these cracks were present in the unexposed material or in the center section of the 

specimens that were not directly exposed to the quartz lamp heaters. Figure 12-122 shows 25 X 

magnification views of unexposed material, unheated center section material, and material in 

the heated zones. The longitudinal surface cracks are readily visible in the heated zone. Similar 

surface cracks were also found on thermal aging samples after 10.000 hours of exposure at 

both 450K (350 OF) and 505K (450 OF). Examples of these are shown in Figure 12-123. 

12 -145 
__. 

0~924 20.398 HOURS 0~925 27,890 HOURS653217-417 

Figun 12-1 14 HT-S/710 G/PI Notchod Long-Torm Specimens That Failed During Flight Simulation 

~~92-1 18,302 HOURS DF92-3 27,890 HOURS 

6532 17-41 8 

Figuro 12-1 15 HT-S/710 G/PI Notehod Long-Torm Specimens That Won Romovod From Flight Simulation 

Testing Before Completion of 50,000 hr of Exposuro B~auso of Imminent Failure 

12 -146 
LEFT HOLE CENTER HOLE RIGHT H,OLE 

653217-419 

Figure 12-1 16 HT-SnlO G/PI Notchod Long-Torm Spodmon DF92-3 Aftor 27,890 hr of Flight Simulation 

Exposure Showing Variation in Degreo of Cracking in Holes Across Width of Specimen 

12 -147 
I 10,000 HOURS t t 19,400 HOURS t 

t 27,890 HOURS t 39,856 HOURS t 

ARROWS INDICATE SPECIMEN EDGES 653217-420 

Flgure 12-1 17 Photomicrographs (SEM) of HT-W710 G/PI After Flight Simulation Exposure - 

505K (450OF), lOOX 

It is interesting to note the circumstances under which surface microcracking occurs for the 

different resin matrix composites. For the epoxies, both thermal exposure and fatigue cycling 

were required. No surface microcracking was observed for HT-SI710 GIPI on either thermal 

aging or flight simulation specimens. Finally, for Celion 60001LARC-160 GIPI, thermal expe 

sure alone was sufficient to cause microcracking while fatigue cycling alone was not. 

.ORIGINAL PAGE TS 

12-148 OE POOR QUALITY 
-. - - __ . 

Figure 12-118 Celion 6000AARC~160 G/PI long-Term Specimens After 10,000 hr of 

Flight Simulation Exporuro 

Figure 12-119 Celion 6000/LARC-160 G/PI long-term Specimens After 10,000 hr of Flight 

Simulation Exporuro (Opposite sides of specimens in Figure 12-1 18) 

12-149 
Figure 12-121 

Celion 6000/LARC-160 G/PI Notched long-Term Specimens After 10,000 hr of 

Flight Simulation Exposure 

Celion 6000/lARC-160 G/PI Notched long-Term Specimens After 10,000 hr of 

Flight Simulation Exposure (Opposite sides of specmens in Figure 12-120) 

12 -150 
UNEXPOSED 

EXPOSED 

UNHEATED CENTER SECTION 

EXPOSED 

HEATEDZONE 653217-425 

Figure 12-122 Photomicrographs (SEM) of Celion 6000/LARC-160 G/PI Before (upper) and 

After (lower) 10,000 hr of Flight Simulation Exposure, 25X 

12-151 I 
. . 450K (35OOF) 

505K (45OOF) 

65321 7-426 

Figure 12-123 Photomicrographs (SEM) of Celion 6000/LARC-160 G/PI After10,OOO hr of Thermal Aging 

in One Atmosphere Air at Indicated Temperature Showing Surface Microcracks, 25X 

Metallographic and SEM examhations showed the effects of 10,000 hours of flight simulation 

to be quite similar to 10,000 hours of thermal aging except for an increase in matrix micro 

cracking or void formation. This was not unexpected with the fatigue loading present in the 

flight simulation exposure cycles. Figure 12-124 shows these differences in the microstructure 

for the two test conditions after 10,000 hours at 450K (350OF). At higher magnifications the 

SEM revealed evidence of oxidation similar to that shown in Figure 8-40 for thermal aging at 

450K (350 OF); i.e., considerable oxidation in the outer plies but almost none in the center plies. 

12-152 OIUGINAL PAGE TS 

03 EOOR QUALITY 
ARROWS INDICATE 

SPECIMEN EDGES 

t UNEXPOSED I 

1 (A) FLIGHTSIMULATION 1 t (B) THERMAL AGING t 

65321 7-427 

Figure 12- 124 Photomicrographs (Metallograph) of Celion 6000/WRC- 160 G/PI 

Before (upper) and After (lower) 10,000 hr of (a) Flight Simulation Exposure 

and (b)Thermal Aging in Air at 45OK (SOOF), lOOX 

B/AI, [OO -C 4501, and [00l6, unnotched. Table 12-41 lists the 15 Phase I1 B/Al fight simulation 

specimens and the number of hours of exposure that each completed. For the 25,000 hour 

group, four of the six survived. None of the 50,000 hour specimens reached their goals. Failure 

times ranged from just under 8000 hours to almost 32,000 hours. The following is a description 

of the foul; groups of specimens: 25,000 hours, [0 O k 45 "1,; 50,000 hours, [0 O f 45 'Ig; 25,000 

hours, [0 OI6; 50,000 hours, [0 'I6. 

12-153 
Tabb 12-41. Summary of B/Al Pharo II Flight Simulation Exposum 

Spocimom Expourn Hours Tost Soda 

Numbor Oriontation Comploted (Hours) Commontr 

EC92-5 

E C93- 1 

EC93-2 

EC93-3 

EU91-2 

EU92-1 

EU92-2 

EU92-3 

. EU93-1 

EU93-2 

EU93-3 

EC92-1 [0" 2 45"], 29,327 50,000 

EC92-2 [O" 2 45 "I, 31,447 50,000 

EC92-3 [0" f 45"], 17,707 50,000 

EC92-4 [0" 2 45'1, 19,519 50,000 

[O " f 45 "1, 30,447 50,000 

27,402 

27,402 

27,402 

17,519 

17,702 

31,584 

12,605 

7,692 

23,588 

25,000 

25,000 

25,000 

25,000 

50,000 

50,000 

50,000 

50,000 

25,000 

25,000 

25,000 

(a), Wearout failure in heated zone 

(a), Wearout failure in heated zone 

Wearout failure in heated zone 

Wearout failure in heated zone 

(a), Wearout failure in heated zone 

(a), (b), Did not fail 

(a), (b), Did not fail 

(a), (b), Did not fail 

Failed in unheated center section 

Failed near doubler at edge of heated zone 

Wearout failure in heated zone 

Failed at edge of heated zone 

Failed near doubler 

Wearout failure in heated zone 

Did not fail 

(a) Specimen failed in the simulation apparatus during compressive overload at 20,957 hours. 

(b) Because of operator error the 25,000 hour specimens were not removed from test until 27,402 

Specimen was modified and returned to test. See Figure 12-94. 

hours. 

The three 25,000 hour crossplied specimens were among those that were modified to the configuration shown in Figure 12-94 after failing during the compressive overload at 20,957 hours. 

These revised specimens successfully completed 27,402 hours of exposure, after which they 

were removed from the simulator for examination and residual strength testing. A bookkeeping error was responsible for not halting the exposure cycling after 25,000 hours as 

originally intended. 

Figures 12-125 and 12-126 show the two sides of the three crossplied specimens after removal 

at 27,402 hours. While the three specimens completed cycling without failure, there was considerable surface damage and edge cracking. Some surface wrinkling was also present. 

Magnified views of the surface of one of the specimens are shown in Figure 12-127. The surface 

roughening, first noted on the 10,000 hour specimens (Figure 12-82), is quite severe after 

27,494 hours. Extensive longitudinal cracking of the aluminum and fracture of the outer ply 

boron filaments can also be seen in the photographs. These surface cracks and broken 

filaments were found on both sides of two of the three specimens. The third, EC93-3, had only 

the roughened surface and edge cracking. 

12 -154 
Figure 12-125 Crorsply B/Al Long=Term Specimens After 27,402 hr of Flight Simulation Exposure 

(Specimonr were redoublered after compression ovorload failure) 

- -- 

Figure 12-126 Crorrply B/Al long-Term Specimens After 27,402 hr of Flight Simulation Exposure 

(Opposite sides of specimens in Figure 12-125) 

12-155 
a 

9 

x 

E! 

m 

C 

3 

e 

V 

Q 

2 v) 

a, 

U 

: v) 

-a 

a, 

C 

a, 

s 

al 

0 

a 

a 

al 

C 

0 

c 

3 

x 

c) 

a, 

3 

L 

8 

a 

X 

YI 
- .- 

Figure 12.128 Crossply B/Al long-Term Specimens That Failed During Flight Simulation Testing 

Before Completion of 50,000 hr of Exposure 

1 

- - 

Figure 12-129 Crossply B/Al long-term Specimens That Failed During Flight Simulation Testing Before 

Completion of 50,000 hr of Exposure (Opposite sides of specimens in Figure 12-128) 

12 -157 
- 

5X EC92-3 5X EC924 663217-433 

Figure 12-130 B/Al long-Term Sp.cimonr EC92-3 (17,707 hr) and EC92-4 (19,519 hr) Showing Roughened 

Surfaces, Surface Cracking in the Aluminum, and Fractured Outer-Ply Boron Filaments 

For the 50,000 hour, [Oo f 45'1, group of specimens, two failed prior to completing 20,000 

hours. The remaining three failed during the compressive overload at 20,957 hours, were 

modified to the configuration shown in Figure 12-94, and were reintroduced to test. The three 

modified specimens completed approximately 30,000 hours of exposure before final failure. 

Figures 12-128 and 12-129 are front and back views of the first two specimens. The fracture 

areas showed some necking down and numerous surface wrinkles. The wrinkles or bulges were 

quite severe for specimen EC92-4. Surface cracking, both in the Al and the boron filaments in 

the outer ply was extensive in the region of the failures. This can be seen in the magnified 

views of Figure 12-130 taken from near the failures. As was observed for the 10,000 hour specimens the degree of surface roughening varied considerably between the heated areas and the 

unheated center sections of these specimens. Figure 12-131 shows the difference after 19,519 

hours of exposure. Also shown in Figure 12-131 are a series of connected openings or holes in 

the surface of the AL These openings are aligned in the longitudinal direction of the specimens 

parallel to the boron filaments in the outer ply. 

The remaining three specimens of this group are shown in Figures 12-132 and 12-133. All three 

failed near the 30,000 hour mark. The failure appearances were quite different among these 

specimens. Fracture occurred along a direction 45 degrees to the load in one of the specimens, 

-- 

12 -158 
10 X HEATED ZONE 10 X CENTER SECTION 65321 7-434 

Figure 12-131 B/Al Long-tern Specimen EC92-4 After 19,519 hr of Flight Simulation Exposure 

Showing Roughened Surface Obsewed in Heated Zones (Photograph on right 

shows considerably smoother surface of unheated center section for comparison) 

EC92-1. Damage to the surface was also considerably less than for the other two specimens. 

Specimen EC92-5 showed extensive necking in contrast to the other two. Failure for this 

specimen was along two intersecting 45 degree directions. For the third specimen, EC92-2, 

fracture occurred almost perpendicular to the loading direction for two thirds of the width and 

at 45 degrees for the remainder. Higher magnification views of these failures and the surface 

away from the failures are shown in Figures 12-134, 12-135, and 12-136. 

The three 25,000 hour unidirectional specimens are shown in Figures 12-137 and 12-138. Only 

one specimen survived the 25,000 hours of exposure. The first failure, specimen EU93-1, 

occurred in less than 8000 hours. The failure was similar to that described earlier for specimen 

EU91-1 in Section 12.7.2 where- the stiffening grid was forced into the B/Al at the edge of an 

end doubler during a compressive load. Figure 12-139 shows where the four bearing surfaces of 

the stiffening grid were forced into the B/AL Figure 12-140 is a SEM photomicrograph of one 

of these areas. The effect of the grid bearing on the soft A1 matrix was a smearing over of the 

M, which is readily visible in the photomicrograph. The second specimen failure, EU93-2, 

occurred in the heated zone at 23,588 hours and appeared to be a wearout failure. The overall 

condition of these specimens was somewhat better than the crossply specimens after similar 

exposure periods. Fewer surface and edge cracks and a lesser degree of surface roughening 

were observed for the unidirectional specimens. Figure 12-141 shows the specimen surface 

after 25,000 hours of exposure for areas in the heated zone and the unheated center section. 

12 -159 

o! mR QUALITY 
Figure 12. ,132 Crorrpl 

Before 

compra 

Y 

1s C 

,in 

B/Al long-Term Spdmenr That Failed Dur 

:omplotion of 50,000 hr of Exposure (Spocirn 

don overload failure 

sing 

ens 

Flight Simulation Testing 

were redoublered after 

Figure 12-1 133 

LtiiLJU PAGE E3 

i ]eOOR QUALITY 

.. 

Crorrply B/Al long-Term Specimens That Failed During Flight Simulation Testing Bet01 

Completion of 50,000 hr of Exposure (Opposite sides of specimens in Figure 12-132) 

12-160 

re 
1x 1 ox 653217-437 

Figure 12-134 B/Al long-Term Specimen EC92-1 After 29,327 hr of Flight Simulation Exposure 

Showing Fracture and Surface Away 

1x 

__. 

from Failure 

1 ox 65321 7-438 

Figure 12-135 B/Al long-Term Specimen EC92-2 After 31,447 hr of Flight Simulation Exposure 

Showing Fracture and Surface Away from Failure 

12 -161 
1x 1 ox 65321 7-439 

___ 

Figuro 12-136 B/Al long-Term Specimen EC92-5 Aftor 30,477 hr of Flight Simulation Exposure 

Showing Fracture and Surfaco Away from Failure 

653217-440 

Figure 12-137 25,000 hr Unidirectional B/Al long-Term Spdmenr After Flight Simulation Exposuro. 

Two Specimens Failed Prior to 25,000 hr 

C>KiGIN& p~~s 1s 12 -162 

POOR QUA LIT^ 
Figure 12-138 25,000 hr Unidirectional B/Al long-Term Speximens After Flight Simulation Exposure 

(Opposite sides of specimons in Figuro 12-137) 

Figure 12-139 B/Al long-Term Specimen EU93-1 Showing Fracture Appearance After Failure in 7692 hr 

(Failure was initiated at locations (arrows) where stiffener grid was forced into the B/Al) 
. 13X 653217-443 

Figuro 12-140 SEM Photomicrograph of Rogion Whore Stiffonor Grid was Forced into Aluminum Matrix 

Near End Doubler of Spocimen EU93-1 Ahr 7692 hr of Flight Simulation Exposure (Contact 

aroa locatod betwoen arrows) 

1OXHEATEDZONE __ 10X CENTER SECTION 653217-444 

Figure 12-141 B/Al long-Term Specimen EU93.3 After 25,000 hr of Flight Simulation Exposure Showing 

Roughoned Surface Observed in Heated Zones (Photograph on right shows considerably 

smoother surface of unheated center section of comparison) 

12-164 
ORIGINAL PAGE 1s 

POOR QUA LIT^ 

The degree of surface roughening is less than that shown in Figure 12-131 for a crossplied 

specimen after 19,519 hours of exposure. 

The 50,000 hour unidirectional BlAl specimens did not fare any better than the crossply specimens with all failing to reach the 50,000 hour god Failure times ranged from 12,605 hours to 

31,584 hours. Front and back views of these specimens are shown in Figures 12-142 and 

12-143. The first failure, EU92-3, occurred after 12,605 hours at the edge of the heated zone. 

Other than the fracture the specimen condition was very good. Whether this early failure was 

the result of wearout could not be determined. 

The next two failures, both after slightly less than 18,000 hours, occurred away from the 

heated zones. One was near the specimen center and the other near one of the end doublers. 

Neither specimen appeared to have failed as a result of wearout. The condition of these 

specimens was better than crossply specimens after similar exposure times. Some surface rip 

pling and roughening were visible in the heated zones, but very few surface cracks were present. For specimen EU92-1 there was some evidence that the compression stiffeners may have 

lead to the failure in a manlrer similar to that described earlier for specimen EU93-1. 

The last specimen of this group, EU92-2, failed after 31,584 hours in the heated zone in what is 

thought to have been the result of wearout. Again, the degree of surface cracking in the region 

of the failure was considerably less than observed previously for the crossply specimens. There 

was considerable surface roughening in the heated zones, however, as shown in Figure 12-144. 

The degree of surface roughening has increased substantially over that present after 25,000 

hours as can be seen by comparison with the unidirectional specimen shown in Figure 12-141. 

. - __ -- ___ -- 

Figure 12-142 Unidirectional B/Al long-Term Specimens That Failed During Flight Simulation Testing 

Before Completion of 50,000 hr of Exposure 

12-165 
65321 7-446 - . - 

Figure 12-143 Unidirectional B/Al long-Term Specimens That Failed During Flight Simulation Testing 

Before Completion of 50,000 hr of Exposure - (Opposite sides of specimens in Figure 12-142) 

1OX CENTER SECTION 65321 7-447 ___~. 1OX HEATED ZONE 

Figure 12-144 B/Al long-Term Specimen EU92-2 After31,584 hr of Flight Simulation ExposureShowing 

Roughened Surface Observed in Heated Zones (Photograph on right shows considerably 

smoother surface of unheated center section for comparison) 

12-166 
12.7.3.2 Residual Strongtb Data - Phaso II. Residual strength test plans were formulated for 

the Phase I1 materials prior to the start of exposure. These plans stressed tensile and fatigue 

tests of both unnotched and notched configurations at ambient and elevated temperatures. It 

was necessary to modify these plans, however, for several reasons. Of prime importance was 

the failure of any of the specimens to survive 50,000 hours of exposure. Specimens either 

failed, or, in the case of the epoxies, were removed from test prior to reaching 50,000 hours. For 

the specimens that failed to complete the required number of exposure hours, their unfailed 

ends could generally be salvaged for residual strength testing. Other factors affecting the 

change in test plans were the increased importance of residual compressive and shear properties and loss of test material from under the compressive grids. 

In keeping with the rationale of Phase I, unnotched tensile and fatigue and short beam shear 

specimens were machined from material that was not shielded by the compression stiffening 

grids. For the notched tensile, notched fatigue, and compressive specimens the entire width of 

the flight simulation specimens was used. To get sufficiently thick specimens for compressive 

and short beam shear testing it was necessary to bond several layers together. All required 

three layers (18 ply) except the epoxy compressive specimens, which required four layers (24 

ply). No compressive specimens were made from the BIAl material because of the testing difficulties encountered with the 10,000 hour compressive tests of Phase I. Specimen preparation 

techniques, configurations, and test procedures were generally identical to those used during 

Phase I. In Phase I, bonding of doublers was required on only one end of each of the residual 

tensile and fatigue specimens as the titanium doubler from the flight simulation specimen was 

still attached to the other end. In Phase 11, it was often necessary to bond doublers to both 

ends of the specimens because of the poor condition of the adhesive holding the original 

titanium doublers on the long term specimens. This was particularly true for the longer 

exposure times. 

WE, [Oo f 45.1, and G/E; [Oo f 45O],, unnotchd. Because of the high failure rate of the GIE 

specimens and the extensive visual degradation of the BIE specimens during Phase I, all of the 

original BIE and GIE flight simulation specimens were removed at the end of 10,000 hours. 

New specimens were fabricated to replace them, and flight simulation testing was restarted 

with.the maximum test temperature reduced from 408K (275°F) to 373K (212°F). All of the 

replacement specimens were unnotched. 

The specimens of both BIE and GIE completed 10.000 hours and 25,000 hours without failure. 

Two BIE specimens failed near 33,000 hours. Because of the likelihood of additional failures 

and in order to obtain the maximum amount of residual strength data, flight simulation 

testing of the remaining specimens was halted after 36,374 hours. 

The test plan for residual strength testing of the epoxy specimens is given in Table 12-42. 

Material was available for test with accumulative exposure periods of 10,000 hours, 25,000 

hours, and 36,374 hours. The number of tensile and fatigue tests of the 36,374 hour B/E 

material had to be reduced because of the two flight simulation specimen failures near 33,000 

hours. 

12 -167 
Tablo 12.42. Rosidual Strength Tost Plan for Phaso I1 B/E and G/E Aftor Flight Simulation Tosting 

T.mp.&Um Numbor of Spocimom Matorial Flight Simulation 

Syrtom Specimon Typo Tost Typo K - ("F) 10,000 hr 25,000 hr 36,374 hr 

BIE [0 2 45'1, Unnotched Tensile 297 75 8 3 6 

NotchedTensile 297 75 4 0 3 

Fatigue 297 76 9 6 6 

Shear 297 75 4 4 4 

Compressive 297 75 4 4 4 

GIE [0 f 45'1, Unnotched Tensile 297 75 8 3 8 

NotchedTensile 297 75 4 0 4 

Fatigue 297 75 9 6 9 

Compressive 297 75 4 4 4 

Shear 297 75 4 4 4 

. Residual tensile and notched tensile data before and after 10,000 hours, 25,000 hours, and 

36,374 hours of flight simulation exposure are listed in Tables 12-43 and 12-44 for B/E and 

Tables 12-45 and 12-46 for GIE. Elastic modulw values are included for the unnotched tensile 

tests. The tensile strength values have been plotted as a function of exposure time in Figures 

12-145 and 12-146. Residual compressive and short beam shear data before and after the same 

exposure periods are listed in Tables 12-47 and 12-48 for BIE and Tables 12-49 and 12-50 for 

G/E. Elastic modulus values are given for the compressive tests. The compressive and shear 

strength values have also been plotted as a function of exposure time and are shown in Figures 

12-147 and 12-148. Residual axial fatigue data at room temperature and a stress ratio, R, of 0.1 

for unnotched specimens are presented in Tables 12-51 and 12-52, and S-N cwes are plotted in 

Figures 12-149 and 12-150. Also included on the graphs are SN curves (no data points) for 

unexposed material tested under the same conditons (presented earlier in Section 11. Baseline 

Fatigue Testing). 

12-168 
tablo 12-43. Strongth and Modulus Proportior of [Oo f 45.1, WE at 297K (75OF) Bofon 

and Aftor Flight Sirnulation Exporuro 

StrWlgth Modulur Spuimon Flight Simulation 

Numbor Spocimon Numbor MNIIT? (ksi) GN/m2 (Msi) 

Unnotched Tensile: 10,000 hr 

A931-1 

-2 

-3 

-4 

-5 

-6 

A935-4 

-6 

AC93-1 

-1 

-1 

-1 

-1 

-1 

-5 

-5 

Unnotched Tensile: 25,000 hr 

A941-1 AC94-1 

-2 -1 

-3 -1 

Unnotched Tensile: 36.374 hr 

A951-1 AC95-1 

-2 -1 

-3 -1 

-4 -1 

-5 -1 

-6 -1 

Baseline (Phase I1 Material) 

492 

457 

551 

468 

487 

527 

533 

av 498 

474 

447 

482 

av 471 

483 

450 

336 

419 

450 

414 

av 413 

av 531 

71.3 

66.3 

79.9 

67.9 

70.6 

76.4 

77.3 

72.3 

68.7 

64.9 

69.9 

70.1 

68.3 

- 

65.3 

48.7 

60.8 

65.2 

60.1 

59.4 

59.9 

77.0 

- 

79.3 

66.9 

69.6 

70.3 

68.3 

71.7 

71.7 

70.3 

63.4 

61 

59 

61 

60 

- 

69 

69 

70 

90 

85 

77 

70.3 

- 

- 

11.5 

9.7 

10.1 

10.2 

9.9 

10.4 

10.4 

9.2 

10.2 

- 

8.8 

8.6 

8.9 

8.8 

- 

10.0 

10.0 

10.1 

13.1 

11.1 

10.2 

- 

12.4 

12-169 I 
tablo 12-44. Notehod tonsilo Strongth of [Oo f 45'1, WE at 297K (75'F) Bdoro and Aftor 

Flight Simulation Exporuro (Unnotchod Flight Simulation Spoeimon) 

Spocimon Fllght Simulation 

Numbor Spoeeimon Numbor 

Strongth 

MN/& (kri) 

Notched Tensile 10,000 hr 

A9341 AC93-4 

-2 -4 

-3 -4 

A935-5 -5 

Notched Tensile: 36,374 hr 

A9544 

-5 

-6 

Baseline (Phase I Material) 

AC95-4 

-4 

-4 

396 

441 

380 

av 410 

423 

385 

403 

av 383 

361 

av 443 ' 

57.4 

63.9 

55.1 

61.3 

59.4 

55.8 

58.4 

55.5 

64.2 

52.4 

12 -170 
tablo 12-45. Strongth and Modulus Propodos of [Oo f 45'1, G/E at 297K (75OF) Bofon 

and Aftor Flight Simulation Exporuro 

Strongth Modulus Spcrcimon Flight Simulation 

Numbor Speccimon Numbor MN/~ (ksi) GN/d (Md) 

Unnotched Tensile: 10.000 hr 

B931-1 

-2 

-3 

-4 

-5 

-6 

B935-4 

-6 

BC93-1 

-1 

-1 

-1 

-1 

-1 

-5 

-5 

Unnotched Tensile: 25,000 hr 

B941-1 BC94-1 

-2 -1 

-3 -1 

Unnotched Tensile: 36,374 hr 

B953-1 

-2 

-3 

-4 

-5 

-6 

B955-4 

-6 

BC95-3 

-3 

-3 

-3 

-3 

-3 

-5 

-5 

Baseline (Phase I1 Material) 

747 

66 1 

686 

587 

705 

554 

492 

609 

av 630 

632 

647 

583 

av 621 

623 

574 

652 

609 

487 

598 

520 

487 

av 540 

av 550 

108.3 

95.9 

99.5 

85.1 

102.3 

80.4 

71.4 

88.3 

91.4 

91.7 

93.8 

84.6 

90.0 

90.3 

83.3 

.94.5 

88.3 

70.7 

86.8 

75.4 

70.6 

82.5 

79.7 

61 

61 

59 

61 

59 

56 

52 

52 

58 

46 

48 

48 

47 

59 

53 

54 

64 

62 

74 

79 

57 

67 

53 

8.8 

8.9 

8.6 

8.9 

8.5 

8.1 

7.5 

7.6 

8.4 

6.6 

7.0 

7.0 

6.9 

8.5 

7.7 

7.9 

9.3 

9.0 

10.7 

11.4 

8.2 

9.1 

7.7 

12-171 
Tab10 12-46. Natchod Tensilo Strength of [Oo f 45'1, G/E at 297K (75OF) Won and Aftor 

Flight Simulation Exposuro (Unnotchd Flight Simulation Spdmon) 

Spocirnon Flight Simulation 

Number Speccirnon Number 

Notched Tensile: 10,000 hr 

B934-1 

-2 

-3 

B935-5 

Notched Tensile: 36,374 hr 

B952-4 

-5 

-6 1 

Baseline (Phase I Material) 

BC93-4 478 

-4 385 

-4 459 

-5 456 

av 445 

BC95-2 476 

-2 488 

-2 426 

av 463 

av 412 

69.4 

55.9 

66.5 

66.1 

64.5 

69.0 

70.9 

61.8 

67.2 

59.7 

12 -172 1 r-/ - I 
700 

600 

500 

L 400 

- 90 

- 80 

-70 

OUNNOTCHED TENSILE STRENGTH 

NOTCHED TENSILE STRENGTH 

- 

0 c 

2 

0 r9 - 4 

Figuro 12-146 Roridual tonsilo Strength of [OO f 4501, G/E at 297K (75OF) Aftor Flight Simulation 

Exporuro at a Maximum Temporaturo of 373K (212OF) 

12 -173 

L. a 400 - 

300 - 

200 - 

- 

100 - . - , 

I I I 

0 10,000 20.000 30,000 40,000 

0 

m 

-60 ;;t 

z 

c" 

-50 v) 

--I 

a -40 

c1 

4 

I -30 - 

20 

10 

0 

E 

u 

- 

N - - E 

0 

t; 

w 

g300 - 4 

Iw 

t4 

5 200 - 3 0 UNNOTCHED TENSILE STRENGTH 

0 NOTCHED TENSILE STRENGTH - 3 

100 - 

- 

0 I I I 

0 10,000 ' 20,000 30,000 40,000 

400 

90 

80 

c 

I- 

-70 $ 

5 

E 

-60 m 

-50 

v1 

-I 

z 

c1 

-40 g 

-30 2 

20 

10 

.a 

x 

L"-. 

u 
Tablo 12-47. C~mpmssivo Stnngth of [Oo f 45'1, WE at 297K (75OF) Won 

and Aftor Flight Simulation Exporuro 

Strongth Modulus Spuimon Flight Simulation 

Numbor Spoccimon Numbor MN/~ (ksi) GN/d (Msi) 

Compressive: 10.000 hr 

A9C-8 AC93-3 

-9 -3 

-10 -3 

-11 -3 

Compressive: 25,000 hr 

A9C- 12 AC94-3 

-13 -3 

-14 -3 

-15 -3 

Compressive: 36.374 hr 

A9C- 16 AC95-5 

-17 -5 

-18 -5 

-19 -5 

Baseline (Phase I1 Material) 

1,160 169 69.0 10.0 

1,230 179 70.3 10.2 

1,170 170 71.0 . 10.3 

1.170 170 69.6 10.1 

av 1.190 172 70.3 10.2 

1,040 

924 

1.090 

1,090 

av 1.030 

151 

134 

158 

158 

150 

492 71.4 

430 62.4 

826 119.9 

687 99.7 

av 609 88.4 

68 

66 

67 

68 

67 

9.9 

9.5 

9.7 

9.9 

9.8 

74 10.8 

58 8.4 

65 9.4 

60 8.7 

64 9.3 

av 1,540 223 102 14.8 

12 -174 
Tablo 12-48. ShoH Eoam Shoar Strongth of [Oo f 45'1, B/E at 297K (75OF) Bdoro and Aftor 

Flight Simulation Exporuro 

Spdmon Flight Simulation 

Numbor Spuimon Numbor 

Short Beam Shear: 10.000 hr 

A9S1 

-2 

-3 

-4 

AC93-5 

-5 

-5 

-5 

av 

Short Beam Shear 25,000 hr 

A9S5 

-6 

-7 

-8 

Short Beam Shear: 36,374 hr 

A9S13 

-14 

-15 

-16 

Baseline (Phase I Material) 

AC94-2 

-2 

-2 

-2 

AC95-4 

-4 

-4 

-4 

av 

av 

av 

66 

65 

65 

68 

66 

45 

48 

45 

41 

45 

30 

30 

25 

29 

29 

66 

9.5 

9.4 

9.4 

9.9 

9.6 

6.5 

7.0 

6.5 

5.9 

6.5 

4.3 

4.4 

3.6 

4.2 

4.1 

9.5 

12 -175 
tablo 12-49. Compnrrivo Strength of [Oo 2 aO], G/E ot 297K (75OF) Before 

and Ahr Flight Simulation Exporuro 

Strongth Modulus 

Spoeirnm Flight Sirnulotion 

Numbor 3pocimon Numbor MN/~ (W GN/d (Msi) 

Compressive: 10.000 hr 

B9C-3 BC93-3 

-4 -3 

-5 -3 

-6 -3 

Compressive: 25,000 hr 

B9C-7 BC94-3 

-8 -3 

-9 -3 

-10 -3 

Compressive: 36.374 hr 

B9C-11 BC95-1 

-12 -1 

-13 -1 

-14 -1 

Baseline (Phase I Material) 

551 

5 27 

576 

576 

av 558 

620 

622 

574 

618 

av 608 

470 

497 

564 

592 

av 531 

av 570 

79.9 

76.5 

83.6 

83.5 

80.9 

89.9 

90.2 

83.2 

89.6 

88.2 

68.2 

72.2 

81.9 

85.8 

77.0 

82.7 

46 

43 

46 

43 

45 

51 

50 

50 

50 

50 

48 

50 

50 

54 

51 

46 

6.7 

6.3 

6.6 

6.3 

6.5 

7.4 

7.3 

7.2 

7.2 

7.3 

7.0 

7.2 

7.3 

7.8 

7.3 

6.6 

12 -176 
Tablo 12-50. Short Boom Shoar Stnngth of [Oo f 45'1, G/E at 297K (75OF) Edon and Aftor 

Flight Simulation Exporuro 

Spcimon Flight Simulation 

Numbor Speimon Numbor MN/~ (ksi) 

Short Beam Shear 10,000 hr 

B9S1 

-2 

-3 

-4 

Short Beam Shear: 25.000 hr 

B9S5 

-6 

-7 

-8 

Short Beam Shear 36,374 hr 

B9S-9 

-10 

-11 

-12 

Baseline (Phase I Material) 

BC93-5 

-5 

-5 

-5 

av 

BC94-2 

-2 

-2 

-2 

av 

BC95-2 

-2 

-2 

-2 

av 

av 

66 

61 

63 

62 

63 

68 

68 

68 

70 

68 

53 

59 

45 

55 

53 

74 

9.5 

8.9 

9.2 

9.0 

9.2 

9.8 

9.8 

9.9 

10.2 

9.9 

7.7 

8.6 

6.5 

8.0 

7.7 

10.7 

12 -177 
u 

2 75 

H 

5 50- a 

25 

- 

c) 

- 500 - OCOMPRESSIVE STRENGTH -20-3 2 

-2 

-1 

OSHORT BEAM SHEAR STRENGTH 

250 - - 10 

7 8 

I 1 I -0 -0 

U 

(kd 

100 9

s80- 

I0 

70- 

IVY 

6o 

$ 50- 

40- 

VY 

VY 

0 = 

E 30 

W + 

U 

-I 

3 

20 

10 

OCOMPRESSIVE STRENGTH 

OsHORT BEAM SHEAR STRENGTH 

(MN/m2) (MNlm2) (ksi) 

u 

- 600 7- 0 4 60 

500 - 

- 400 - 

300 - 

- 200 - 

- 

100 - 

- 

0- 0- I I I lo Jo 

0 10,000 20,080 30,000 40,000 

-50 , 

2 rn 

- 30 

- io 

- 10 

12 -178 
Tablo 12-51. Rosidual Axial Fotiguo Data for [Oo * 45OIs B/E at 297K (75O) Aftor 

Flight Simulation Exposuro, R = 0.1, Unnotchod 

~~~ 

Maximum Stross 9 Spoeimm Flight Simulation 

Number Spoccimon 

Numbor MN/& (kri) Cyckr Commonts 

10,000 hr 

A932- 1 

-2 

-3 

-4 

-5 

-6 

A934-4 

-5 

-6 

25,000 hr 

A941-4 

-5 

-6 

A942-4 

-5 

-6 

36,374 hr 

A952-1 

-2 

-3 

A953-1 

-2 

-3 

AC93-2 

-2 

-2 

-2 

-2 

-2 

-4 

-4 

-4 

AC94-1 

-1 

-1 

-2 

-2 

-2 

AC95-2 

-2 

-2 

-3 

-3 

-3 

345 

414 

276 

310 

379 

396 

328 

293 

362 

362 

310 

310 

345 

379 

328 

310 

276 

293 

328 

284 

302 

50 

60 

40 

45 

55 

57.5 

47.5 

42.5 

52.5 

52.5 

45 

45 

50 

55 

47.5 

45 

40 

42.5 

47.5 

41.25 

43.75 

639,000 

1.000 

10,223,000 

3,162,000 

441,000 

304,000 

2,493.000 

15,440,000 

484,000 

200,000 

199.000 

12,000.000 

145,000 

6,000 

349.000 

6,000 

9,462,000 

347,000 

4,500 

1,771,000 

10,000 

Did not fail 

Did not fail 

Failed under doubler 

Did not fail 

Failed under doubler 

12 -179 I 
tablo 12-52. Rosidual Axial Fatiguo Data for [Oo -C 45.1, G/E at 297K (75') Aftor 

Flight Simulation Exporuro, R = 0.1, Unnotchod 

Flight Simulation Maximum Sbr 

Spoeimon Spdmrn 

Numbrr Numbrr MN/& Cydrr Commrnts (W 

10.000 hr 

B932-1 

-2 

-3 

-4 

-5 

-6 

B934-4 

-5 

-6 

25,000 hr 

B941-4 

-5 

-6 

B942-4 

-5 

-6 

36,374 hr 

B954-1 

-2 

-3 

-4 

-5 

-6 

B955-1 

-2 

-3 

BC93-2 

-2 

-2 

-2 

-2 

-2 

-4 

-4 

-4 

BC94-1 

-1 

-1 

-2 

-2 

-2 

BC95-4 

4 

-4 

-4 

4 

-4 

-5 

-5 

-5 

414 

483 

552 

586 

517 

448 

500 

534 

552 

552 

483 

414 

379 

517 

492 

586 

448 

483 

517 

552 

414 

603 

379 

414 

60 

70 

80 

85 

75 

65 

72.5 

77.5 

80 

80 

70 

60 

55 

75 

71.3 

85 

65 

70 

75 

80 

60 

87.5 

55 

60 

10,286,000 

2,034.000 

.1.238,000 

1,000 

11.000 

10,332.000 

60,000 

< 1,000 

355,000 

5,000 

594,000 

1,069,000 

10,000.000 

740,000 

118,000 

11,000 

2,850,000 

645,000 

117,000 

29,000 

2.080.000 

2,000 

8,039,000 

895,000 

Did not fail 

Did not fail 

Failed under doubler 

Failed under doubler 

Did not fad 

Failed under doubler 

Lowering the maKimurn test temperature for flight simulation exposure of the epoxies from 

408K (275°F) to 373K (212°F) resulted in a significant improvement in the performance of 

these materials. The GIE system was particularly impressive with only the shear strength exhibiting a major decrease in value after 36,374 hours of exposure. For the BIE system the 

losses in tensile and compressive strengths after 36,374 hours were comparable to those in 

Phase I in only 10,000 hours of exposure. 

The unnotched and notched tensile strengths of the BIE material were slightly lowered (less 

than 10%) by 10,000 hours of flight simulation exposure. The modulus was unchanged. For the 

GIE system the unnotched and notched tensile strengths and the modulus were higher for the 

12 -180 
600 I rn 

0 10,000 HR EXPOSURE 

U 25,000 HR EXPOSURE- 80 

A 36,000 HR EXPOSURE I .- 500 

450 

400 

350 

300 

250 

200 

102 104 105 106 

NUMBER OF CYCLES 653.217452 

Flgun 12-149 Axial Fdguo Proportios of [OO 2 4501, B/E Bdon and Aftor Flight Simulation 

Exporun, 297K (7SoF), R = 0.1, Unnotchod 

25,000 HR EXPOSURE 

I I 1 I1 Ill I I I I1111 I I I I1 Ill I I111111 

102 1 04 105 106 107 

NUMBER OF CYCLES 

653.21 7-453 

Flgure 12-150 Axial Fdguo Propertior of [OO * 4501, G/E Before and After Flight Simulation 

Exposure, 297K (75OF), R = 0.1, Unnotchd 

12 -181 
exposed material than for the baseline specimens. The effect of 10,000 hours of flight simulation exposure on the matrix properties, compressive and shear strengths, was not very large 

except for the BIE compressive strength, which showed a decrease of approximately 25%. The 

results were much improved over those of Phase I for both epoxy systems. As shown in 

Figures 12-149 and 12-150 the fatigue properties of BIE were somewhat lowered by the 10,000 

hours of long-term exposure cycling while the GIE fatigue strengths after cycling were higher 

than those obtained during the baseline fatigue tests. 

As shown in Table 12-43, the tensile strength of the BIE material was lowered approximately 

10% by 25,000 hours of flight simulation exposure. The GIE material behaved even better with 

essentially no change in the tensile strength as a result of the exposure. Modulus values were 

decreased for both systems by about 10%. The effect of 25,000 hours of flight simulation exposure on the matrix properties, compressive and shear, was slight for the G/E but substantial 

for the BIE system. Compressive and shear strength of the BIE decreased by about onethird 

after 25,000 hours of cycling. The GIE showed no change in compressive strength and a 7% 

loss in shear strength. A further decrease in the fatigue strength of the BIE was observed, but, 

again the amount was smaL For GIE the 25,000 hours of exposure produced a decrease in 

fatigue strength only for fatigue lifetimes of 106 or greater. 

The 36,374 hours of exposure had a significant effect on the BIE system. Residual strength 

losses ranged from about 20% for tensile to near 60% for both compressive and shear. The two 

flight simulation failures after approximately 33,000 hours of exposure were undoubtedly the 

result of these very low residual compressive and shear strength properties. The GIE system, 

on the other hand, survived the 36,374 hours of flight simulation exposure with little loss in 

tensile or compressive strength. The major effect was a decrease in shear strength of 28%. The 

BIE showed a further decrease in fatigue strength while the G/E S-N curve was very similar to 

that obtained after 25,000 hours of exposure. The data are plotted in Figures 12-149 and 

12-150. 

In summary of these data, the BIE and GIE systems survived 25,000 hours of flight simulation 

exposure at a cruise temperature of 373K (212°F) with little visual damage. Some microcracking of the surface was observed in the heated zones but was not thought to be serious. 

Mechanical property degradation was minimal for the GIE material. For the B/E system, 

however, a 30% decrease in compressive and shear strengths and a 10% decrease in tensile 

strength were observed. After 36,374 hours of exposure the condition of the specimens was 

also quite good, although a few loose fibers were now visible in the heated zones of both the 

BIE and GIE specimens. Some increase in surface microcracking was also present. Mechanical 

property degradation of the GIE system was limited to a slight change in compressive 

strength but a significant loss in shear strength. The BIE system was severely degraded by the 

36,374 hours of exposure, particularly for the matrix properties, compressive and shear 

strengths. 

HT-S/710 G/PI, [Oo f 45OIs, unnotched and notched. The amount of residual strength data 

obtained for the HT-SI710 GIPI system was less than that originally planned because of two 

factors. First, the three 25,000 hour specimens were fabricated from the second batch of 

HT-SI710 prepreg ordered later in the program. This batch of material was found to have very 

poor thermooxidative properties during the new material screening tests. While two of the 

three 25,000 hour specimens were damaged during installation and were unusable, the third 

12 -182 
failed to complete 6,000 hours of flight simulation exposure. It is believed that had the two 

other specimens not been damaged they also would not have survived much more than 6,000 

hours of test. The second factor that reduced the number of residual strength specimens was 

the failure of all the unnotched and notched 50,000 hour specimens prior to completion of 

50,000 hours. The test parameters of temperature and fatigue loading were selected such that 

80% of the specimens were expected to survive 50,000 hours. This did not occur, and in fact, all 

of the specimens failed prior to reaching 40,000 hours. 

It was possible, however, to salvage a half of each failed flight simulation specimen for fabrication of residual strength specimens. The residual strength test plan was selected to emphasize 

tensile strength as well as the matrix dominated properties, compressive strength and shear 

strength. The limited material available for specimens did not allow any residual fatigue 

testing or testing at other than room temperature. Table 12-53 shows the number of specimens 

of each type and the approximate number of hours of exposure each received. The exact 

exposure times can be found in the data tables that follow. 

Table 12-53. Residual Strength Test Plan for HTW710 G/PI After Flight Simulation Testing 

Number of Spocimens Per Exposure Time 

Test Type 16,000 hr 18,000 hr 20,000 hr 28,000 hr 40,000 hr 

aUnnotched Tensile 6 3 

bNotched Tensile 6 3 3 

BCompressive 2 2 2 

@hear 4 4 

%ut from unfailed ends of unnotched flight simulation specimens 

bCut from unfailed ends of notched flight simulation specimens 

Residual tensile and notched tensile data before and after flight simulation exposure are listed 

in Tables 12-54 and 12-55. Elastic modulus values are included for the unnotched tensile tests. 

The tensile strength values have been plotted as a function of exposure time in Figure 12-151. 

Also included in Figure 12-151 are data obtained in Phase I for 10,000 hours of exposure. 

Residual compressive and short beam shear data before and after flight simulation exposure 

are listed in Tables 12-56 and 12-57 and plotted as a function of time in Figure 12-152. The 

residual strength data for 10,000 hours obtained in Phase I are also shown on the curves. In 

Figure 12-153 the residual tensile and compressive modulii are plotted for the various exposure 

times. 

12 -183 
Table 12-54. Strength and Modulus Properties of [Oo f aSo], Hf-S/710 G/PI at 297K (75OF) Before 

and After Flight Simulation Exposure 

Strength Modulus Specimen Flight Simulation 

Number Specimen Number MN/& (kti) GN/m2 (Msi) 

Unnotched Tensile: 15.458 hr 

D91S1 DC91-S 

-2 -S 

-3 -S 

D921-1 DC92-1 

-2 -1 

-3 -1 

Unnotched Tensile: 39,856 hr 

D924-1 DC92-4 

-2 -4 

-3 -4 

Baseline 

188 

189 

287 

310 

430 

375 

av 296 

46 

56 

179 

av 94 

av 519 

27.3 

27.4 

41.6 

45.0 

62.3 

54.4 

43.0 

6.7 

8.1 

25.9 

13.6 

75.3 

39 

54 

41 

44 

56 

44 

46 

32 

54 

31 

39 

53 

5.7 

7.9 

5.9 

6.4 

8.1 

6.4 

6.7 

4.7 

7.9 

4.5 

5.7 

7.7 

At the completion of the Phase I 10,000 hours of flight simuation exposure the HT-SI710 GIPI 

material had experienced significant losses in shear strength (54%). tensile strenth (29%). and 

compressive strength (23%). With the additional exposure during Phase I1 out to 40,000 

hours, the residual tensile and compressive strengths continued to decrease. The tensile 

strength, in particular, was severely affected with a loss of over 80% of the unexposed baseline 

value. Both the compressive and shear strengths decreased almost 60%, but the major portion 

of these changes occurred in the first 10,000 hours for shear and 20,000 hours for compressive 

strength. In like manner to the epoxy systems, the notched tensile strength was least affected 

by flight simulation exposure with a loss in value of approximately 25% in 28,000 hours. No 

notched tensile strength data were obtained for 40,000 hours as none of the notched flight 

simulation specimens cdmpleted more than 28,000 hours of exposure. 

It should be kept in mind that comparisons of the various properties may not be entirely valid 

because of the role of specimen location on strength levels. The tensile and shear specimens 

were cut from the unshielded areas of the flight simulation specimens while the compressive 

and notched tensile specimens were prepared from the entire width. Also, the effect of bonding 

6 ply laminates together to form 18 ply compressive and short beam shear specimens requires 

additional study to determine the validity of the results. 

In summary of these data none of the HT-SI710 GIPI specimens survived 50,000 hours of 

flight simulation exposure. Specimens experienced wearout failures from the thermal and 

I fatigue exposure cycling at time periods ranging from 15,000 hours to 40,000 hours. Unlike the 

. 

I 12 -184 
epoxy systems, no surface microcracking was observed. However, loss of polyimide resin from 

the surface in the heated zones occurred, thereby exposing the underlying graphite fibers. 

Cracking and delamination at the edges and at the holes was also found. Both surface resin 

loss and edge cracking increased with exposure time. Residual strength properties obtained 

from the undamaged ends of failed specimens were generally poor. Tensile, compressive, and 

shear strengths had all decreased by 50% or more in the first 20,000 hours of exposure. The effect of the cyclic fatigue loading and perhaps the thermal cycling can be seen by comparing 

these results with those from the thermal aging tests at the same temperature. After 25,000 

hours of thermal aging at 505K (450°F) in one atmosphere air, almost no loss in tensile 

strength was measured. 

table 12-55. Notchod tensile Strength of [Oo f dS0], HT-S/710 G/PI at 297K (75OF) Before and 

After Flight Simulation Exposure. (Notched flight simulation specimen) 

Specimen Flight Simulation 

Number Specimen Number 

Strength 

MN/m2 (kri) 

Notched Tensile: 18,077 hr 

DF92S-1 

-2 

-3 

DF922-1 

-2 

-3 

Notched Tensile: 20,398 hr 

DF924-1 

-2 

-3 

Notched Tensile: 27,890 hr 

DF923-1 

-2 

-3 

DF92-S 

-s 

-S 

-2 

-2 

-2 

DF92-4 

-4 

-4 

DF92-3 

-3 

-3 

257 

3 04 

134 

324 

326 

277 

av 270 

37.3 

44.1 

19.4 

47.0 

47.3 

40.2 

39.2 

3 19 46.2 

221 32.1 

244 35.4 

av 261 37.9 

245 35.5 

316 45.8 

242 35.1 

av 268 38.8 

Baseline av 357 51.8 

12 -185 
'-454 

Figure 12-151 Residual Tensile Strength of [Oo f 45O], HT-S/710 G/PI at 297K (75OF) After Flight 

Simulation Exposure at a Maximum Temperature of SOSK (450OF) 

Table 12-56. Compressive Strength Properties of [O" f 45.1, HT-S/710 G/PI at 297K (75OF) Before 

and After Flight Simulation Exposure 

Strength Specimen Flight Simulation 

Number Specimen Number MN/& (ksi) 

Modulus 

GN/m2 (Msi) 

Compressive: 15,896 hr 

D9C-5 DC92-2 

-6 -2 

Compressive: 19,400 hr 

D9C-7 DC92-5 

-8 -5 

Compressive: 39,856 hr 

D9C-9 DC92-3 

-10 -3 

Baseline 

248 

232 

av 240 

193 

188 

av 191 

135 

180 

36.0 

33.7 

34.8 

28.0 

27.3 

27.6 

19.6 

26.1 

37 

37 

37 

29 

30 

30 

30 

29 

5.4 

5.3 

5.4 

4.2 

4.4 

4.3 

4.3 

4.2 

av 158 22.8 29 4.2 

av 377 54.7 46 6.7 

12 -186 
table 12-57. Short Beam Shear Strength Properties of [Oo f 45'1, HT-S/710 G/PI at 297K (75OF) 

Before and After Flight Simulation Exposure 

70 

I + 

(I 

560 

a 

z50- 

c 

- > 

v) 

40 a. 

0 

!A 

c 

z 

a30 

g20- 

5 

10 1 

Specimen Flight Simulation 

Number Specimen Number 

- 

- 

- 

- 

- 

OL 

Strength 

MN/& (ksi) 

0 COMPRESSIVE STRENGTH 

OSHORT BEAM SHEAR STRENGTH 

- 

- 

0 

100 - - 

0 I I I 

Short Beam Shear 18,302 hr 

-50-7 

-6 

-5 

-4 

40 

30 

-20-3 

-2 

-1 

10 

0 &O 

D9S5 

-6 

-7 

-8 

Short Beam Shear: 27,890 hr 

D9S9 

-10 

-11 

-12 

Baseline 

DF92-1 

-1 

-1 

-1 

DF92-5 

-5 

-5 

-5 

21 

20 

17 

20 

av 20 

14 

19 

15 

18 

av 17 

av 37 

3.0 

2.9 

2.4 

2.8 

2.7 

2.0 

2.8 

2.2 

2.6 

2.4 

5.4 

. d 

(ksi) 

c 

I- + 

b 

rn 

- 

5 

v) 

X 

rn 

D 

z 

rn 

z 

n 

-455 

Figure 12-152 Residual Compressive and Short Beam Shear Strength of [OO .C 4S0], Ht-S/710 G/PI at 

297K (75OF) After Flight Simulation Exposure at a Maximum Temperature of 505K (45OOF) 

12 -187 
t 

' I OCOMPRESSIVE MODULUS 

0 I I I 0 

0 10,000 20,000 30,000 40,000 

Figure 12-153 Residual Tenrilo and Compnrrive Modulus of [OO -C Go], HT-S/710 G/PI at 297K (75OF) 

FLIGHTSIMULATION EXPOSURE TIME (hr) 653.217-456 

Aftor Flight Simulation Exposuro at a Maximum Tomprraturo of 505K (GOOF) 

Celion 6000/LARC-160 G/PI, [Oo 2 45"],, unnotched and notched. The original plan for flight 

simulation testing of this system included exposure times of 10,000 hours and 25,000 hours at 

a maximum temperature of 450K (350°F). Because of time and financial restrictions, the 

25,000 hour exposures were not funded. Consequently, at the completion of 10,000 hours there 

was an extra set of both unnotched and notched specimens available for residual strength 

testing. This excess material was used for compressive and shear tests and for some additional 

tensile, notched tensile, and notched fatigue tests. The original and revised test plans for 

Celion 60001LARC-160 GlPI are shown in Table 12-58. Results of the 297K (75°F) and 450K 

(350°K) tensile tests on unnotched and notched GlPI after 10,000 hours of flight simulation exposure are presented in Tables 12-59 through 12-61. Also included are the baseline data for 

comparison purposes. Two types of notched tensile specimens were tested. The first, Table 

12-60, was cut from unnotched flight simulation specimens and was notched after exposure. 

The second, Table 12-61, was cut from the notched flight simulation specimens after being exposed in the notched condition. Table 12-62 gives the compressive and short beam shear data 

for GlPI before and after flight simulation exposure of 10,000 hours. Results of the residual 

fatigue tests are listed in Tables 12-63 and 12-64. The data in Table 12-63 were obtained from 

specimens cut from unnotched flight simulation specimens and tested at 297K (75 :F) and a 

stress ratio, R, of 0.1 in both unnotched and notched configurations. The data in Table 12-64 

were obtained from notched specimens cut from notched flight simulation specimens and 

tested at 297K (75°F) and 450K (350°F) and a stress ratio, R, of 0.1. The fatigue results are 

plotted as S-N curves in Figures 12-154 through 12-157. Curves for unexposed Celion 

6000lLARC-160 GlPI tested under the same conditions are shown on the fatigue plots for comparison. 

12 -188 
table 12-58. Residual Strength test Plan for Celion 6000/LARC-160 G/PI After 

10,000 hr of Flight Simulation Testing 

Number of Specimens 

9 Revised Flight Simulation Temperohre Original 

Specimen Type Test Type K (OF) Plan Plan 

IO' f 45'1, Unnotched Tensile 

Notched Tensile 

Fatigue 

Compressive 

Notched Fatigue 

Tensile 

Notched Tensile 

[Oo -C 45O], Notched Notched Tensile 

Notched Fatigue 

Shear 

Notched Tensile 

Notched Fatigue 

297 75 

297 75 

297 75 

297 75 

297 75 

450 350 

450 350 

297 75 

297 75 

297 75 

450 350 

450 350 

10 10 

10 10 

10 10 

0 4 

0 9 

5 10 

5 5 

10 10 

10 10 

0 8 

5 10 

0 12 

12-189 
tablo 12-59. Strongth and Modules Proporties of [Oo 2 45'1, Celion 6000/LARC-160 G/PI Boforo and 

Aftor 10.000 hr of Flight Simulation Exposure. (Unnotchod flight simulation specimen) 

Tempomtu re Strongth Modulus Spoeimon Flight Simulation 

Numbor Spoeimon Numbor K (OF) MN/d (ksi) GN/d (Msi) 

Unnotched Tensile: 10,000 hr 

F911-1 

-2 

-3 

-4 

-5 

-6 

F913-1 

-2 

-3 

F914-1 

Baseline 

F912-1 

-2 

-3 

-4 

-5 

-6 

F9 13-4 

-5 

-6 

F914-4 

Baseline 

FC91-1 

-1 

-1 

-1 

-1 

-1 

-3 

-3 

-3 

-4 

FC91-2 

-2 

-2 

-2 

-2 

-2 

-3 

-3 

-3 

-4 

297 75 510 

482 

647 

674 

645 

703 

581 

63 1 

598 

678 

av 614 

av 703 

450 350 540 

510 

542 

649 

574 

587 

630 

5 96 

584 

73 1 

av 594 

av 696 

73.9 

69.9 

93.9 

97.7 

93.5 

84.2 

91.5 

86.8 

98.3 

89.1 

102 

102 

78.3 

73.9 

78.6 

94.2 

83.2 

85.1 

91.4 

86.5 

84.7 

86.2 

106 

101 

57 

63 

63 

74 

63 

61 

104 

66 

57 

83 

69 

59 

57 

74 

71 

70 

57 

53 

51 

51 

45 

74 

60 

58 

a 8.3 

a 9.1 

a 9.2 

10.7 

9.2 

8.9 

15.1 

9.6 

8.3 

12.1 

10.0 

8.6 

a 8.3 

=10.7 

a10.3 

10.1 

8.3 

7.7 

7.4 

7.4 

6.5 

10.7 

8.7 

8.4 

gage measurement 

12 -190 
fable 12-60. Notched Tensile Strength of [Oo f Go], Celion 6000/LARC-160 G/PI Before and After 

10,000 hr of Flight Simulation Exposure. (Unnotched flight simulation specimen) 

fompemture Strength Spoeimm Flight Simulation 

Number Specimen Number K (OF) MN/m2 (ksi) 

Notched Tensile: 10.000 hr 

F923-1 

-2 

-3 

-4 

-5 

-6 

F924-1 

-2 

-3 

' F922-1 

Baseline 

F922-2 

-3 

-4 

-5 

-6 

Baseline 

FC92-3 

-3 

-3 

-3 

-3 

-3 

-4 

-4 

-4 

-2 

FC92-2 

-2 

-2 

-2 

-2 

297 75 618 

548 

589 

49 1 

592 

552 

a 

603 

a 

547 

av 567 

450 350 

av 427 

568 

587 

591 

607 

583 

av 587 

av 475 

89.7 

79.5 

85.4 

71.2 

85.8 

80.1 

a 

87.4 

a 

79.4 

82.3 

62.0 

82.4 

85.1 

85.7 

88.0 

84.6 

85.2 

68.9 

aFailed under doubler 

12 -191 
Table 12-61. Notchod Tensilo Strongth of [ao f 45.1, Celion 600O/lARC-160 G/PlBoforo and Aftor 

10,000 hr of Flight Simulation Exposure. (Notched flight simulation specimon) 

Temporature Stnngth Specimon Flight Simulation 

Numbor Specimon Numbor K (OF) MN/m2 (kN 

Notched Tensile: 10.000 hr 

FF911-1 

-2 

-3 

-4 

-5 

-6 

FF913-1 

-2 

-3 

FF914-1 

Baseline 

FF912-1 

-2 

-3 

-4 

-5 

-6 

FF913-4 

-5 

-6 

FF9 14-4 

Baseline 

FF91-1 

-1 

-1 

-1 

-1 

-1 

-3 

-3 

-3 

-4 

FF91-2 

-2 

-2 

-2 

-2 

-2 

-3 

-3 

-3 

-4 

297 75 443 

643 

a 

419 

510 

389 

a 

623 

663 

528 

av 527 

450 350 

av 427 

a 

a 

592 

36 1 

44 1 

43 1 

456 

432 

532 

332 

64.3 

93.4 

a 

60.8 

74.0 

56.4 

a 

90.4 

96.2 

76.7 

76.5 

62.0 

a 

a 

85.9 

52.4 

64.0 

62.5 

66.1 

62.7 

77.2 

48.2 

av 447 64.9 

av 475 68.9 

,Failed under doubler 

12 -192 
Table 12-62. Compressive and Short Beam Shear Strength Properties of [Oo 2 45"],Celion 

6000/LARC-160 G/PI at 297K (75OF) Before and After 10,000 hr of Flight 

Simulation Exposure 

Specimen Flight Simulation 

Number Specimen Number 

Strength Modulus 

MN/m2 (ksi) GN/m2 (Msi) 

Compressive: 10.000 hr 

F9C-1 FC91-5 

-2 -5 

-3 -5 

-4 -5 

Baseline 

Short Beam Shear: 10,000 hr 

F9S 1 

-2 

-3 

-4 

-5 

-6 

-7 

-8 

Baseline 

FF92-3 

-3 

-3 

-3 

-3 

-3 

-3 

-3 

565 

544 

596 

492 

av 549 

av 584 

21 

24 

21 

21 

35 

35 

30 

35 

av 28 

av 51 

81.9 41 5.9 

78.9 47 6.9 

86.5 48 7.0 

71.5 56 8.2 

79.7 48 7.0 

84.7 52 7.5 

3.1 

3.5 

3.0 

3.1 

5.1 

5.1 

4.4 

5.1 

4.0 

7.4 

12 -193 
Table 12.63. Residual Axial Fatigue Data for [Oo 2 45.1, Celion 6000/LARC-160 G/PI at 297K (75OF) After 

10,000 hr of Flight Simulation Exposure, R = 0.1. (Unnotched flight simulation specimen) 

Maximum Stress Spocimen Flight Simulation Specimen 

Number Specimen Number TYPO MN/d (kri) Cycles Comments 

F914-2 

-3 

-5 

-6 

F921-1 

-2 

-3 

-4 

-5 

-6 

F9244 

-5 

-6 

F925-1 

-2 

-3 

-4 

-5 

-6 

FC91-4 

-4 

-4 

-4 

FC92-1 

-1 

-1 

-1 

-1 

-1 

FC92-4 

-4 

-4 

-5 

-5 

-5 

-5 

-5 

-5 

Unnotched 448 

483 

465 

500 

414 

517 

43 1 

379 

500 

396 

Notched 607 

559 

538 

512 

488 

467 

443 

419 

373 

65 

70 

67.5 

72.5 

60 

75 

62.5 

55 

72.5 

57.5 

88.1 

81.1 

78.0 

74.3 

70.8 

67.7 

64.2 

60.7 

54.1 

191,000 

2,000 

97,000 

459,000 

2,503,000 

c 1.000 

1,005,000 

3,000 

10,077,000 Did not fail 

12,423,000 Did not fail 

c 1.000 

2,000 

51,000 

1,480.000 

451.000 

9,000 

82.000 

194,000 

13,067,000 Did not fail 

12 -194 
Table 12-64. Residual Axial Notched Fatigue Data for [Oo f 4P], Celion 6000/LARC-160 G/PI 

After 10,000 hr of Flight Simulation Exposure, R = 0.1. (Notched flight 

simulation specimen) 

Temperature Maximum Stress 

Number Specimen Number K (OF) MN/m2 (ksi) Cycles Comments 

Specimon Flight Simulation 

FF9142 

-3 

-5 

-6 

FF915-1 

-2 

-3 

-4 

-5 

-6 

FF921-1 

-2 

-3 

-4 

-5 

-6 

FF922-1 

-2 

-3 

-4 

-5 

-6 

FF91-4 

-4 

-4 

-4 

-5 

-5 

-5 

-5 

-5 

-5 

FF92-1 

-1 

-1 

-1 

-1 

-1 

-2 

-2 

-2 

-2 

-2 

-2 

297 75 516 

493 

423 

374 

353 

562 

400 

445 

377 

423 

450 350 483 

448 

500 

448 

414 

396 

379 

552 

276 

345 

310 

483 

74.8 

71.5 

61.3 

54.2 

51.2 

81.5 

58.0 

64.5 

54.7 

61.4 

70 

65 

72.5 

65 

60 

57.5 

55 

80 

40 

50 

45 

70 

5,000 

8,000 

2,000 

175,000 

2,000 

258,000 

19,000 

53,000 

7,000 

10,057,000 Did not fail 

123,000 

- Machine malfunction 

11,000 

72,000 

359,000 

297,000 

1,156,000 

2,000 

11,395,000 

170,000 

1,541,000 

3.000 

12 -195 
300 

200 

Figure 12-154 Axial Fatigue Propodes of [OO f 45O], Celion 6000/LARC-160 G/PI Before and After 

10,000 hr of Flight Simulation Exposure, 297K (75OF). R = 0.1, Unnotched. (Unnotched 

flight simulation specimen) 

I 

L 

- 40 

I I I I1111 I 1 I I I Ill I I I I I Ill I I I I1111 I I I I 11y30 

N 

E 

2 

s 

\ 

- 

W 

u 

z 

v) 

v) 

W 

a 

c 

v) 

H 

3 

x 

H 

r 

a 

- 700 

- 

600 

0 

0 

UNEXPOSED 

10 500 

400 

/o 

0 \-- 

100 

90 

- - * 

Y -80 2; 

-70 2 

a 

2 

u 

>. 

VJ 

W 

60 

5 

Figure 12-155 Axial Fatigue Properties of [Oo f fiO], Celion 6000/LARC-160 G/PI Before and After 

10,000 hr of Flight Simulation Exposure, 297K (75OF), R = 0.1, Notched. (Unnotched 

flight simulation specimen) 

300 

200 

12 -196 

0 2 

-50 2 

z 

- 40 

I I I I I Ill I I I I I Ill 1 I I 1 I Ill I I I I1111 I I I I I 11330 
700r 

- 

600 

0 

Axial Fatiguo Proporties of [Oo * 45O], Celion 6000/LARC-160 G/PI Before and After 

10.006 hr of Flight Simulation Exposure, 297K (75OF), R = 0.1, Notched. (Notched flight 

simulation spocimon) 

-'loo 

90 

?= 

(II 

-80 

500 

400 

300 

200 

Figure 12-157 Axial Fatiguo Properties of [Oo f 45O3, Celion 6000/LARC-l60 G/PI Before and After 

10,000 hr of Flight Simulation Exposure, 45OK (35OoF), R = 0.1, Notched. (Notched flight 

simulation specimen) 

- 

0 -60 5 . 

I 

- 40 

+ I I I I I Ill I I I Ill11 I I I I I Ill I I I I I Ill I I I I 11~30 

12 -197 

- 

- 

800 

0 

4 

100 

90 

1 

VI 

- 

2t 

-80 2; 

700 

500 

400 

300 

200 

t 

2 

u 

>- 

u 

v3 

UJ 

-70 ;;j 

a 

-60 5 

;I 

2 

0 -50 2 

0 

3 

H 

\ 

<)40 

I I 1 I1111 I I I I1111 I I I I1111 1 I I 11111 I I I I 111i30 
After 10,000 hours of flight Simulation exposure at a maximum test temperature of 450K 

(350°F) the Celion 6000/LARC-160 material exhibited slight losses in unnotched tensile 

strength and compressive strength, almost no loss in notched tensile strength, but a severe 

reduction in shear strength. The results for the Celion 60001LARC-160 were considerably better than those for the HT-S/710 G/PI after 10,000 hours. In both cases, however, shear strength 

was drastically reduced by the flight simulation exposure. Surprisingly, the flight simulation 

residual properties of Celion 6000/LARC-160 were higher than those obtained after 10,000 

hours of thermal aging at the same temperature. Here again the property most degraded was 

the shear strength. 

Two groups of notched tensile specimens were tested. One group was notched after flight 

simulation exposure, and the other was notched before exposure and exposed in the notched 

condition. While both groups of specimens retained their strength after flight simulation, the 

specimens that were exposed in the unnotched condition possessed significantly higher values 

after exposure, particularly at the elevated test temperature. 

With the extra flight simulation specimens of Celion 60001LARC-160 G/PI that were available 

it was possible to increase the amount of residual fatigue testing over that originally scheduled. Both unnotched and notched residual fatigue curves, Figures 12-154 and 12-155, of 

specimens cut from unnotched flight simulation specimens showed slight decreases from the 

baseline curves. These decreases were more pronounced at the high cycle portions of the 

curves. For flight simulation specimens that were in the notched condition during exposure the 

residual fatigue properties (only notched residual strength specimens were tested) showed a 

greater decrease from the baseline. Again the decrease was greater for higher cycles. The S-N 

curves for specimens cut from notched flight simulation specimens are shown in Figures 

12-156 and 12-157. Also apparent from these curves is that the effect of test temperature on 

both baseline and residual fatigue strength was small. In like manner to residual notched tensile strength, the major effect on residual fatigue strength is the configuration of the flight 

simulation specimen. That is, material exposed in the notched condition shows a greater 

decrease in residual notched properties than does material exposed in the unnotched condition. 

B/Al, [0" ,+45"], and [Ool6, unnotched. The amount of residual strength data for B/Al, like that 

for the HT-S/710 G/PI system, was considerably less than originally planned because of 

numerous specimen failures during flight simulation exposure at maximum tempera tures of 

505K (450°F). Some of the 25,000 hour specimens survived, but all of the 50,000 hour 

specimens failed before reaching that goal. Table 12-65 shows the test plan for the material 

that could be salvaged from the few specimens that survived and the unfailed ends of those 

that did not. 

12 -198 
Table 12-65. Residual Strength Test Plan for B/AI After Flight Simulation Testing - 297K (75'K) 

Unnotched Tensile Strength 

Exposure 

Time 

(W 

Number of Residual Strength Specimens 

[Oo 2 45.1, (001~ 19001~ L) 

13,000 

18,000 

20.000 

21,000 

24.000 

b 

25.000 

27,000 

32,000 

3 3 

11 

6 

6 

Because of the small amount of material available for specimens and the widely ranging failure 

times, only room temperature tensile testing was performed. In addition to the [0" 2 45'1, and 

[0"]6 specimens, a sd number of 0.076111 (3 in.) by 0.013 m (0.5 in.) transverse tensile 

specimens were cut from the heated zones of specimens that had failed in such a manner so to 

preserve enough heated zone material for testing. These [go"] specimens were cut from the 

failed ends of the flight simulation specimens. 

Results of the unnotched tensile tests on BlAl before and after flight simulation exposure are 

presented in Tables 12-66 through 12-68. Elastic modulus values are included for the 

longitudinal [O" f 45'1, and [O"], tests. The longitudinal tensile strength data have also been 

plotted as a function of exposure time in Figures 12-158 and 12-159. Data obtained in Phase I 

for 10,000 hours of exposure are also plotted on the graphs. 

In Phase I the BlAl system showed losses in mechanical properties of approximately one 

fourth after 10,000 hours of flight simulation exposure. For the [0" k 45'1, crossplied material 

the tensile strength continued to decrease with exposure time although not at as great a rate. 

After 27,000 hours the tensile strength was slightly less than onehalf of the unexposed value. 

Modulus values for the exposed [0" k 45'1, BlAl showed considerable scatter. The data appear 

to indicate little change in modulus for the first 20,000 hours of simulation testing but a severe 

decrease after 27,000 hours. 

12 -199 
Table 12-66. Strength and Modulus Properties of [Oo f 45'1, B/Al ot 297K (75OF) Before and After 

Flight Simulation Exposure 

~ ~~~ 

Strength Mod u I us Specimen Flight Sirnulation 

Number Specimen Number MN/~ (ksi) GN/d (Msi) 

Unnotched Tensile: 17,707 hr 

EC923-1 EC92-3 

-2 -3 

-3 -3 

av 

Unnotched Tensile: 19,519 hr 

EC924-1 EC92-4 

-2 -4 

-3 -a 

av 

Unnotched Tensile: 20.957 hr 

EC925-1 EC92-5 

-2 -5 

-3 -5 

av 

Unnotched Tensile: 27,402 hr 

EC931- 1 

-2 

-3 

EC932- 1 

-2 

-3 

EC933-1 

-2 

-3 

EC93-1 

-1 

-1 

-2 

-2 

-2 

-3 

-3 

-3 

av 

Baseline av 

257 

276 

338 

290 

303 

301 

255 

286 

278 

268 

257 

268 

299 

225 

49 

207 

228 

217 

271 

288 

217 

222 

516 

37.3 

40.0 

49.0 

42.1 

43.9 

43.7 

37.0 

41.5 

40.3 

38.8 

37.3 

38.8 

43.4 

32.6 

7.1 

30.0 

33.0 

31.4 

39.3 

41.7 

31.5 

32.2 

74.8 

79 

193 

97 

123 

179 

125 

150 

151 

120 

80 

44 

42 

74 

105 

115 

84 

83 

154 

- 

11.5 

28.0 

14.1 

17.9 

25.9 

18.2 

21.8 

22.0 

17.4 

11.6 

6.4 

6.1 

10.8 

15.2 

16.7 

12.2 

12.1 

22.4 

- 

For the unidirectional B/Al tested in the fiber direction the tensile strength decreased by 25% 

after the first 10,000 hours of exposure but very little thereafter, out to almost 32,000 hours. 

An exception to this was the data from the specimen that failed at 12,605 hours. However, two 

of these three specimens failed at locations that were not in the heated zone of' the flight 

simulation specimen from which they were cut. For the [0°]6 material no decrease in elastic 

modulus was observed for the 32,000 hours of flight simulation. 

12 -200 
Table 12-67. Strength and Modulus Properties of [OO& B/Al at 297K (75OF) Before and After 

Flight Simulation Exposure 

Strength Modulus Specimen Flight Simulation 

Number Specimen Number MN/& (W GN/m2 (Msi) 

Unnotched Tensile: 12,605 hr 

aEU923-4 EU92-3 

a -5 -3 

-6 -3 

Unnotched Tensile: 17,519 hr 

EU9 12-1 EU91-2 

-2 -2 

-3 -2 

-4 -2 

-5 -2 

Unnotched Tensile: 17,702 hr 

EU921-1 EU92-1 

-2 -1 

-3 -1 

4 -1 

-5 -1 

-6 -1 

Unnotched Tensile: 25.000 hr 

EU933-1 . EU93-3 

-2 -3 

-3 -3 

-4 -3 

-5 -3 

-6 -3 

Unnotched Tensile: 31,584 hr 

EU922-1 EU92-2 

-2 -2 

-3 -2 

Baseline 

av 

av 

av 

av 

av 

av 

660 

448 

1240 

786 

1250 

1150 

1000 

1050 

979 

1080 

1360 

1370 

1310 

952 

1050 

1160 

1200 

1030 

1450 

1250 

848 

1030 

1180 

1130 

1050 

1100 

896 

1020 

1450 

96.5 

65.0 

180.0 

114 

181 

167 

145 

152 

142 

157 

197 

199 

190 

138 

152 

169 

174 

150 

210 

182 

123 

150 

171 

164 

152 

159 

130 

147 

210 

190 

213 

207 

203 

200 

183 

158 

191 

218 

190 

209 

221 

205 

299 

165 

197 

216 

206 

223 

197 

222 

23 1 

218 

216 

223 

210 

213 

215 

198 

27.6 

30.9 

30.0 

29.5 

29.0 

26.5 

22.9 

27.7 

31.6 

27.5 

30.3 

32.0 

29.7 

43.3 

23.9 

28.6 

31.3 

29.9 

32.3 

28.5 

32.2 

33.5 

31.6 

31.3 

32.3 

30.5 

30.9 

31.2 

28.7 

aSpecimen failed outside of heated zone 

12-201 
table 12-68. transvooe tensile Properties of [Ool6 B/Al at 297K (75OF) Before and After 

Flight Simulation Exposuro 

Strength Spoccimen Flight Simulation 

Number Specimen Number MN/& (W 

r 

600 - 

- 

- 

- L? 

z 

Iw 

a 

z300 

e200 - 

- 

- - 

v) 

z 

w 

I- - 

a 

z 

3 

7100 - 

- 

I I I 

0 10,000 20,000 30,000 40.000 

Transverse Tensile: 12,605 hr 

80 

70 

c 

? 60 

D 

-I 

-50: m 

2 

C" 

40 

v) 

--I 

z 

e, 

30 Z 

-202 

E 

10 

,O 

EU923-1 

-2 

-3 

Transverse Tensile: 23,588 hr 

EU932-1 

-2 

-3 

-7 

-8 

-9 

Baseline 

EU92-3 

-3 

-3 

EU93-2 

-2 

-2 

-2 

-2 

-2 

59 8.6 

66 9.5 

62 9.0 

av 62 9.0 

72 

65 

65 

64 

61 

61 

av 65 

av 79 

10.4 

9.4 

9.5 

9.3 

8.8 

9.4 

8.8 

11.5 

Figun 12-158 Roridual tenrilo Strength of [Oo f GO], B/Al at 297K (75OF) After Flight Simulation 

Exposure at a Maximum temporature of 505K (GOOF) 

12-202 
0 I I I 

0 10,000 20,000 30,000 40,000 

FLIGHT SIMULATION EXPOSURE TIME (hr) 553.217462 

Figuro 12-159 Roridual Tenrilo Strength of [V -C Go], B/Al at 297K (75OF) After Flight Simulation 

Exporuro at a Maximum Temperature of 5O5K (45OOF) 

The limited data for transverse tensile strength show a loss of about 20% after 24,000 hours of 

exposure. 

Examination of the flight simulation specimens after exposure both with a low power, stereo 

microscope and using metallographic techniques revealed damage only at the surface and 

specimen edges. The damage consisted of surface roughening, longitudinal cracking of the 

aluminum, and fracture of the outer ply boron filaments. This surface damage is believed to 

have been responsible for the decrease in tensile properties observed during residual strength 

testing. The loss in tensile strength for the crossplied material was greater than that for the 

unidirectional material for two reasons. First, the surface damage was more severe, as noted in 

Section 12.7.3.1 for the crossplied specimens. Second, the [0" k 45'1, crossplied material obtains a large percentage of its strength from the two [OO] plies. As these are located on the outsides, any surface damage will be concentrated in the high strength portion of the specimens. 

PHASE ll SUMMARY 

Table 12-69 summarizes the Phase 11, long term flight simulation residual strength data 

obtained from the five composite systems. The percentage change in the various mechanical 

properties is listed for each material. For several entries in the table data was not available at 

exactly 25,000 hours or 36,000 hours. These values were obtained from the property versus exposure time curves. The fatigue data from the two epoxy systems and Celion 60001LARC-160 

GlPI are summarized in Table 12-70. In this table the fatigue strength at 107 cycles, expressed 

as a percentage of the unexposed tensile strength is compared for unexposed material and for 

that which has survived various periods of flight simulation exposure. 

12-203 
Table 12-69. Summary of Phase II Flight Simulation Residual Strength Data 

Material Temperature % Property Loss In 

System Property K (OF) 10,000 hr 25,000 hr 36,000 hr 

5505I4 BIE F tu 297 

[O' 2 45'1, 

Et 297 

aNotched Ftu 297 

A-S13501-5 GIE 

[O' f 45'1, 

HT-SI710 GIPI 

[O' & 45'1, 

Fcu 297 

EC 297 

FSU 297 

F tu 297 

Et 297 ' 

aNotched F,, 297 

Fcu 297 

EC 297 

FSU 297 

F tu 297 

Et 297 

bNotched Ftu 297 

Fcu 297 

EC 297 

FSU 297 

75 

75 

75 

75 

75 

75 

75 

75 

75 

75 

75 

75 

75 

75 

75 

75 

75 

75 

6 

0 

7 

23 

31 

0 

0 

0 

0 

2 

2 

14 

29 

10 

8 

23 

13 

54 

11 

14 

- 

33 

34 

32 

0 

10 

- 

0 

0 

7 

59 

21 

23 

54 

30 

55 

22 

0 

14 

60 

37 

57 

0 

0 

0 

7 

0 

28 

76 

25 

- 

58 

35 

12-204 
Table 12-69. Summary of Phase II Flight Simulation Residual Strength Data - Concluded 

Temperature % Property Loss In Material 

System PmPeW K (OF) 10,000 hr 25,000 hr 36,000 hr 

Celion 60001 

LARC-160 G/PI 

[Oo ,C 45'1, 

BlAl 

[Oo k 45'1, 

BlAl 

[OOI, 

B/AI 

[90oi6 

Ftu 297 75 

Et 297 75 

Ftu 450 350 

Et 450 350 

aNotched Ftu 297 75 

aNotched Ftu 450 350 

bNotched Ftu 297 75 

bNotched Ftu 450 350 

Fcu 297 75 

E, 297 75 

FSU 297 75 

F tu 297 75 

Et 297 75 

FtU 297 75 

Et 297 75 

Ftu 297 75 

13 

0 

15 

0 

0 

0 

0 

6 

6 

7 

46 

26 

10 

25 

0 

17 

- 

55 

37 

29 

0 

21 

aUnnotched flight simulation specimens 

bNotched flight simulation specimens 

c31,584 hr 

12-205 
Table 12-70. Strength at Fatigue limit or lo7 Cycles Expressed as Percentage of Ultimate tensile 

Strength Before and After Flight Simulation Exposure, R = 0.1 

Fatigue Test 

temperature Fatigue Strength at lo7 

Material Specimen Cycles 

System Configuration Condition K (OF) (% of Unexposed FtV) 

550514 BIE 

[0" k 45'1, 

A-Sl3501-5 GIE 

[O" f 45'1, 

Celion 60001 

LARC-160 GlPI 

[O" If: 45'1, 

Unnotched Unexposed 

Exposed 10.000 hr 

Exposed 25,000 hr 

Exposed 36,374 hr 

Unnotched Unexposed 

Exposed 10,000 hr 

Exposed 25.000 hr 

Exposed 36,374 hr 

Unnotched Unexposed 

Exposed 10.000 hr 

Notched Unexposed 

Exposed 10,000 hr 

Exposed 10.000 hr 

Unexposed 

Exposed 10,000 hr 

297 

297 

297 

297 

297 

297 

297 

297 

297 

297 

297 

297 

297 

450 

450 

75 

75 

75 

75 

75 

75 

75 

75 

75 

75 

75 

75 

75 

350 

350 

60 

57 

57 

52 

81 

85 

73 

70 

67 

58 

100 

93 

77 

81 

61 

aUnnotched flight simulation specimens 

bNotched flight simulation specimens 

12-206 
SECTION 13 

THERMAL EXPOSURE TEST PROGRAM SUMMARY 

This section presents an overview of the thermal exposure test programs with emphasis on 

trends observed for each of the materials. Specific tests included in the summary are the base 

line tensile, compressive, and shear tests; the thermal aging tests: the short-term tests: the 

long-term tests; and the baseline and post-flight simulation constant amplitude fatigue tests. 

The thermal aging tests are included in the summary even though they generally showed that 

significant aging effects occur at temperatures higher than the selected maximum use temperature. In the combined load/environment tests, some of the thermal aging effects were triggered at lower temperatures by the combination of load cycles, thermal cycles, and time at 

temperature. 

Test data are presented in two ways. To aid the observation of trends, strength (stress) has 

been plotted as a function of the various tests performed, where the tests are ordered either in 

increasing test time or test sequence. Selected individual data have been plotted to observe 

data scatter. For the constant amplitude fatigue tests, the cyclic stress at lo7 cycles (aI07) has 

been plotted to indicate the relative severity of the effect of fatigue cycling. Ten million cycles 

were chosen in keeping with summary charts presented previously in both Sections 11 and 12. 

When fatigue effects are indicated by this simplified plotting method, the actual S-N curves 

should be referred to for more insight into the relative effects. In addition, to show the effects 

of the thermal exposures on specific properties, residual strengths have been compared to 

baseline strengths by means of bar charts. The bar charts have been prepared for crossply B/E, 

G/E, G/PI, and BlAl and unidirectional B/Al as these were the layups and materials for which 

flight simulation data were available. For the resin matrix composites, the residual notched 

tensile and notched fatigue values were obtained from notched flight simulation specimens. 

The residual notched tensile data for BlAl were obtained from material that was notched after 

flight simulation. The fatigue strengths shown on the bar charts were also for 107 cycles in 

keeping with sumxiary charts presented previously in both Sections 11 and 12. 

For the B/E and G/E systems, two sets of flight simulation exposure and residual strength 

results are presented. The first, obtained in Phase I, is for 10,000 hours of exposure at a maximum temperature of 408 K (275 OF). The second set of results, obtained during Phase 11, is for 

exposure out to 36,374 hours at a maximum temperature of 373 K (212°F). In Phase I both 

unnotched and notched flight simulation specimens of BIE and G/E were exposed. Only 

unnotched specimens were evaluated in Phase 11. The materials from which the specimens 

were fabricated were obtained at different times, and, particularly for the GIE system, comparisons between the Phase I results and those of Phase I1 were not always possible. 

The new material addition in Phase 11, Celion 6000ILARC-160, was flight simulation tested in 

both the unnotched and notched condition but only for 10,000 hours. 

Also added to the test program during Phase I1 was an evaluation of thermal agingkompression 

effects for two of the resin matrix systems, A433501 GIE and Celion 60001LARC-160 GIPI. 

Some of these data are included in the bar charts for the Celion 60001LARC-160 G/PI. The thermal agingkompression specimens of G/E were from a different batch of prepreg than either the 

I 

13-1 
thermal agingltensile specimens or the flight simulation specimens, and, for that reason, those 

data are not included. 

B/E - PHASE I 

B/E, [OO f 450],, unnotched. Figures 13-1 and 13-2 summarize the available tests for the unnotched BIE. As no baseline tensile testing was performed (except for three tests at room 

temperature), the mean strength at 422 K (300 OF) was estimated to be 545 MNIm2 (79 ksi) and 

the Weibull shape parameter, a, was obtained from reference 40 as 19.5. This value of CY 

translates to a coefficient of variation of about 5%, and indicates that BI5505 BIE is a wellcontrolled repeatable composite system. 

The short-term tests with a peak stress of 321 MNlm2 (46.6 ksi), at a constant 422 K (300°F). 

showed the typical reduction in mean strength and increase in scatter with time. No fatigue 

failures occurred, and the residual strength failure mode was the same as that for the room 

temperature baseline tests. 

Wearout analysis was performed to select the peak stress of 302 MNlm2 (43.8 ksi) for the longterm tests. . 8 

After the first 1000 hours, the cruise temperature was lowered from 422 K (300°F) to 408 K 

(275 OF). No fatigue failures occurred in the 10,000 hours of Phase I. Residual tensile strength 

tests showed more scatter than expected, with five specimens between 338 MNIm2 (49 ksi) and 

400 MNlm2 (58 ksi) and one at 496 MNlm2 (72 ksi). The average decrease at room temperature 

was approximately 35%. The residual compressive tests showed even greater reductions in 

properties with a decrease of slightly more than 50%. Although the residual strength failure 

mode was the same as the baseline, the specimens were severely oxidized in the heated zones. 

Recall that the wearout analysis had assumed that no thermal aging effects would be experienced below the apparent maximum use temperature. This assumption was violated by the 

long-term chemical and physical changes experienced by the epoxy resin. In the Phase I1 tests 

the cruise temperature was changed to 373 K (212°F) for the epoxy systems. The thermal 

aging results (Figure 13-2) have shown that with no applied loads or temperature cycling no 

degradation of tensile properties, and probably none for compressive properties as well, occurs 

during 10,000 hours exposure at 394 K (250 OF). However, at 450 K (350 OF) significant tensile 

and severe compressive property degradation was observed. No thermal aging data were 

obtained at the flight simulation temperature of 422 K (300 OF) or 408 K (275 OF). The post longterm fatigue tests at R = 0.1 and 297 K (75°F) indicated a 0107 value of 276 MNlm2 (40 ksi) 

compared to the baseline value of 324 MNlm2 (47 ksi). 

The combination of load, temperature, and thermal cycling has severly degraded the unnotched BIE specimens at a temperature at which thermal aging alone would probably not 

cause extensive damage. 

B/E, [Oo -C 45O],, notched. Figure 13-3 summarizes the trend data for the notched BIE. The 

residual properties after flight simulation exposure are summarized in Figure 13-2. 

The 422 K (300°F) baseline mean strength was calculated to be about 462 MNIm2 (67 ksi) net 

stress and the same CY value of 19.5 was assumed. The short-term specimens were cycled to a 

peak stress of 248 MNlm2 (36 ksi) at 422 K (300 OF) and exhibited the typical wearout effects of 

reduced mean and increased scatter. Failure modes were net tension. 

13 -2 
100 

600 

500 

400 

300 

200 

loo 

- 

- 

- 

- 

- 

t STATIC 

Y 

0 

(? 0 

Y 

- 

N N t 

c 

4 

3 

Y 

+ E v) 

Y 

w P 

*/ 

100 HR ZOO HR 

SHORT TERM 

Y 

Y u.) c (P - z Y 

a 

a 

t;; 

c 

N 

m Iv) Y v) 

\ 

10,000 HA 

LONG TERM 

R=.1 

BASELINE FATIGUE 

STRENGTH. 107 CYCLES 

65321 7.21 

100 

90 

80 

70 

60 

50 

40 

30 

20 

10 

0 

Figure 13-1 Phase I Unnotched [OO f 4501, B/E 

13 -3 
Ill 

N 

m 

In 

Nr 

N 

0 

0- 

N 

v) -- 

e 

zc 

3 g- -- 

I 

Ia 

z 

w 0- a= 

!FJ 

v), 

h 

0, 

v) 

v) 

NOb 

0 

z0, 

m 

0, 

m 

re 0, 2- 

I 

Ia 

a 

5 

5 

0, * 

m 

0 

N0, 

%- 

0, 

2- 

0- 

U m N 

- N 

0 

0 

co 

0 

0 

e 

0 

z 

c 

0 

0 I 0 

c 

r 

L W 

2 

n 

m E 

3 

n 

0 1" X 

YI 

% 

NE = z - o= 

- 

r- \ 

r--0 

m mx - I- 

- 

L 

L L 

a, 

U 

YI 

\ 

m 

CI N 

0 

v) 

v 

+I 

0 

0 

Y 

0 

u 

a! 

a! 

- 

L L 

n 

2 

L 

- 

U 

0 

v) 

- Ir 

Y 

rQ) 

N 

LL 

0 

0 

v) 

m 

Y 

0 

v) 

b 

- 0 

- 2 

Y 

rm 

hl 

W - a 

IW '2 a 

a 52 

a 0 U 

U IW 

I 

u 

IO 

z 

z 

0 

W 

I 

0 

I0 

a z 

I0 

w 

X 

u 

W 

v) 

v) 

w 

Q 

2 

a 

5 

ou 

+ 

0 

z 

z 

a 

13 -4 
700 

600- 

500 

400 

300 

200 

100 

0 

Y 

a 

E 

X 

N 

N 

- 

-w Y-4 

2; -I- 

:& 

- zz 

- 

- 

- 

- 

STATIC 

- . 

.. 

.. .* Ia 

a 

a 

. v1 W VI + I- 

. v) Y 

. W 0 

i 

1 I 

100HR ' 200HR 

SHORT TERM 

w A 

a 

z 

w - IY Y 

k * 

c N 

Y Y 

(D m 0 0 N 

- - 

c 

I4 

v) 

v) 

0 

W . a H 

i7, 

X 

W a. 

a . 

\ . 

10,000 HR 

LONG TERM 

- 70 

- 60 

= -50 f 

2 

-40 E 

I 

IL 

v) 

30 

20 

IO 

----

BASELINE FATIGUE 

STRENGTH. IO7 CYCLES I 

Figure 13-3 Phase I Notched [Oo & as0], B/E 

13 -5 
Wearout analysis was used to select the long-term peak stress of 239 MNlm2 (34.6 ksi). The 

temperature of the cruise leg was reduced from 422 K (300 OF) to 408 K (275 OF) after about 

1000 hours. No fatigue failures occurred in 10.000 hours. The residual strength specimens also 

failed in net tension with the values scattered between 317 MNlm2 (46 ksi) and 455 MNlm2 (66 

ksi) and having an average value of 400 MNlm2 (58 ksi). Again the epoxy resin sustained 

severe chemical and physical damage in the heated zones from the combination of load, 

temperature, and oxidation. 

The R = 0.1 fatigue strengths for 107 cycles at 297 K (75°F) before and after the long-term 

tests were 407 MNlm2 (59 ksi) and 276 MNlm2 (40 ksi), substantiating the degrading effects of 

flight simulation exposure. 

B/E - PHASE II 

B/E, [Oo -C GO],, unnotched. Because of the severe degradation that occurred during Phase I, 

new flight simulation specimens (all were unnotched) were fabricated for Phase 11, and testing 

was restarted using a maximum exposure temperature of 373 K (212°F). Three batches of 

specimens were included for evaluation: after 10,000 hours, 25,000 hours, and 50,000 hours of 

exposure. Two of the 50,000 hour specimens failed after just over 33,000 hours. Because.additional failures appeared imminent, and, to preserve sufficient material for residual strength 

testing, exposure was halted near 36.000 hours. 

Figures 13-4 and 13-5 summarize the results of the Phase I1 testing of the unnotched BIE. Included in the bar charts of Figure 13-5 are results of the 50,000 hour thermal aging studies for 

comparison with the 10,000 hour and 36,000 hour flight simulation residual strength data. 

No fatigue failures occurred during the first 10,000 hours or 25,000 hours of Phase 11. Residual 

tensile strength of the B/E had decreased slightly but was still 90% of the baseline value after 

25,000 hours. Almost no loss in R = 0.1,297 K (75 OF) fatigue strength was found after 10,000 

hours or 25,000 hours. For the specimens removed from test after 36,000 hours, a tensile 

strength loss of 22% was measured. The post long-term fatigue test gave a 0107 value after 

36,000 hours at 373 K (212°F) the same as that after 10,000 hours at 408 K (275°F). While the 

fiber dominated properties, tensile strength and tensiletensile fatigue strength, were not 

greatly affected by the 36,000 hours of exposure, the same was not true of the matrix 

dominated properties. Both the compressive and shear strengths exhibited losses of about 

60%. Unlike the Phase I specimens cycled to a maximum temperature of 408 K (275"F), the 

Phase I1 specimens showed almost no surface damage up through 25,000 hours. After 36,000 

hours a few loose boron fibers were visible in the heated zones of the flight simulation 

specimens. Thermal aging results for 50,000 hours of exposure are shown in Figure 13-5. The 

data indicate that 50,000 hours with no applied loads or temperature cycling at 394 K (250°F) 

degrades the tensile strength to the same degree as 36,374 hours of flight simulation exposure 

at 373 K (212"F), Le., a loss of slightly over 20%. 

For the B/E system, lowering the maximum temperature to 373 K (212°F) resulted in 25,000 

hours of flight simulation exposure with negligible surface damage and good retention of tensile and fatigue properties. However, a loss of compressive and shear strength prgperties of 

about onethird is a major concern. After 36,000 hours of exposure, surface erosion sufficient 

to expose boron filaments, combined with significant tensile strength losses and severe compressive and shear strength degradation, rules out the likelihood of use for 50,000 hours of 

service at even the low temperature of 373 K (212°F). 

13 -6 
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G/E - PHASE I 

G/E, [OO -C 450],, unnotched. The baseline mean strength of the G/E unnotched specimens at 

408 K (275 OF) to 436 K (325 F) (Figure 13-6) was about 496 MNlm2 (72 ksi) with a calculated 

Weibull CY of 12.7. This low value of shape parameter indicates poor reproducibility that impedes the clear interpretation of test trends. 

The short-term specimens were cycled to a peak stress of 338 MN/m2 (49.1 ksi) at 422 K 

(300 OF), and typical wearout was observed, including two early fatigue failures. Failure mode 

was net tension, the same as the baseline tests. 

The wearout analysis was used to select the peak long-term stress of 305 MNlm2 (44.2 ksi), and 

after 1000 hours, the cruise temperature was reduced to 408 K (275°F). Multiple fatigue 

failures occurred, principally by progressive dehmination The few residual strength results 

(Figures 13-6 and 13-7) showed only a slight reduction in the mean tensile strength but a 

drastic reduction in the mean compressive strength. The combination of severe oxidation and 

delamination growth during flight simulation exposure was a major impact to the program. 

In like manner to the B/E system, thermal exposure without load at 450 K (350OF) produced 

severe degradation of the matrix and a very large decrease in tensile strength. No compressive 

tests were performed, but the appearance of the failed specimens indicated that compressive 

properties would also be poor. Thed aging at 394 K (250 OF) and at 450 K (350 OF) at reduced 

pressure, however, gave no degradation in properties after at least 10,000 hours of exp0sur.e. 

The post long-term test value of 0107 after R = 0.1 fatigue at room temperature was down to 

241 MNlm2 (35 ksi) from the 448 MNlm2 (65 ksi) baseline. The low value of the baseline 0107, 

386 MNlm2 (56 ksi), after R = 0.1 fatigue at 450 K (350°F) is probably attributable mostly to 

material variability. Constant amplitude fatigue at R = -1 during baseline testing was shown 

to produce severe reductions in strength. Compressive failures are suspected for these tests 

that produced 0107 values of 172 MN/m2 (25 ksi) at 297 K (75 OF) and 172 MNlm2 (25 ksi) at 450 

K (350 OF). In hindsight, the test program, expecially for the graphite systems, should have emphasized compressive effects. 

G/E, [OO f 450],, notched. Broad scatter was again evident in the baseline notched G/E with an 

CY value of 12.0 and a mean strength of about 414 MNlm2 (60 ksi) (Figures 13-7 and 13-8). 

The short-term tests, cycled to a peak stress of 249 MNlm2 (36.1 ksi) at 422 K (300°F), produced no fatigue failures, but showed decreasing strength and increasing scatter. The residual 

strength failure mode was net tension. 

After wearout analysis, the long-term specimens were cycled to a peak stress of 232 MNlm2 

(33.6 ksi). The temperature was again reduced from 422 K (300 OF) to 408 K (275 OF) after about 

1000 hours. The badly oxidized residual tensile strength specimens retained a surprising 

amount of their original notched strength, exhibiting values from 358 MN/m2 (52 ksi) to 469 

MNlm2 (68 ksi). 

Again tensile fatigue testing did not discriminate the effects too well, with the post long-term 

testing 0107 value (R = 0.1) at 297 K (75 OF) being 283 MNlm2 (41 ksi), compared to the baseline 

values of 379 MN/m2 (55 ksi) and 345 MN/m2 (50 ksi) at 297 K (75 OF) and 450 K (350 OF). 

13 -9 
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LONG TERM I BASELINE FATIGUE 

STRENGTH, 107 CYCLES 

65321 1.299 

Figure 13-8 Phase I Notched [Oo +- as0], G/E 

13 * 12 
The stress ratio of R = -1 showed the much more severe effects of compression with 6107 

values of 124 MNIm2 (18 ksi) and 152 MNIm2 (22 ksi) at 297 K (76 OF) and 450 K (350 OF) respectively, for the baseline tests. 

G/E - PHASE II 

G/E, [OO f dso],, unnotched. A second set of unnotched flight simulation specimens was also 

fabricated for the GIE system following the poor behavior of the material when exposed at a 

maximum cruise temperature of 408 K (275 OF). The same lower temperature of 373 K (212 OF) 

as used for the BIE system was selected, and testing was halted near 36,000 hours to ensure 

sufficient material for residual strength evaluations. 

Data for the Phase I1 thermal exposures are summarized in Figures 13-9 and 13-10. These data 

are for specimens fabricated from two different batches of GIE prepreg. All the data in Figure 

13-9 are for the second batch of material (Phase 11). The tensile strength values were higher for 

the Phase I1 material while the R = 0.1 fatigue properties were the same as those of the Phase 

I materid In Figure 13-10 the thermal aging data are from Phase I material, and the flight 

simulation data are from Phase I1 material. 

The GIE specimens performed exceptionally well at the reduced flight simulation temperature. 

No fatigue failures occurred during the 36,000 hours of exposure. Residual tensile strength, 

both unnotched and notched, was not reduced after 36,000 hours. Losses in R = 0.1 fatigue 

strength were not seen until after 25,000 hours. The value of 6107 decreased from the 448 

MNIm2 (65 ksi) baseline to 402 MNlm2 (58 ksi) at 25,000 hours and only slightly lower to 386 

MNIm2 (56 ksi) after 36,000 hours. The matrix dominated properties also held up quite well 

with only negligible losses after 25,000 hours. There was, however, a significant decrease in 

residual shear strength after 36,000 hours from a baseline value of 74 MNlm2 (10.7 ksi) to 53 

MNIm2 (7.7 hi). In like manner to the BIE specimens, surface damage to the G/E was not 

observed until 36,000 hours of exposure and then was limited to slight loss of resin and some 

exposed fibers in the heated zones. 

The data in Figure 13-10 appear to indicate that thermal aging at temperatures as low as 394 K 

(250 OF) can seriously degrade the tensile strength of GIE, while flight simulation exposure at a 

slightly lower temperature and slightly shorter time has no effect. However, the two data sets 

are from specimens fabricated from different batches of prepreg obtained several years apart. 

The Phase I1 batch (flight simulation) in addition to showing superior strength retention over 

the Phase I batch (thermal aging) also exhibited considerably better thermooxidative 

resistance. Much less loss of resin and consequently fewer exposed graphite fibers were found 

at the fracture areas of failed residual strength specimens after 36,000 hours of flight simulation exposure than for similar test specimens after 25,000 hours of thermal aging at 394 K 

(250°F). The unfortunate necessity of using different batches of G/E prepreg makes comparison of the thermal aging and flight simulation data sets impossible. 

Again, as with the BIE system, lowering the maximum temperature to 373 K (212 OF) greatly 

improved the performance of the G/E material. Excellent results were obtained after 25,000 

hours of flight simulation, and, with the exception of some surface oxidation and erosion and a 

28% loss in shear strength, the material remained in good condition after 36,000 hours of exposure. The possibility of 50,000 hours of usage at a maximum temperature of 373 K (212OF) 

for the A433501 G/E system cannot be ruled out. 

13 -13 
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Figun 13-9 Phase II Unnotched [Oo f 45O], G/E 

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13-15 
HTW710 G/PI, [Oo f eo],, unnotched. The HT-SI710 GlPI data are summarized in Figures 

13-11 through 13-13. Baseline testing showed poor reproducibility, as exhibited by a mean 

strength at 505 K (450 OF) of 469 MNlm2 (68 ksi) and an CY value of 12.15. Notice that one of the 

baseline specimens reached only 296 MNlm2 (43 ksi). These panels had measured void contents 

up to 13%. so that defect sensitive failures modes could be greatly influenced by individual 

specimen quality. 

The 505 K (450°F) short-term peak stress was originally selected to be 331 MNlm2 (48 ksi) but 

nine out of ten specimens failed the proof load at 283 MN/m2 (41 hi); consequently the stresses 

were reduced to a peak 207 MNlm2 (30.1.ksi). Tests of the short-term panels showed that they 

were about 69 MNlm2 (10 ksi) weaker than the baseline paneL Eventually, the short-term 

specimens exhibited typical wearout indications of decreasing strength and increasing scatter 

with test time. 

The long-term test peak stress was reduced to 153 MNlm2 (22.2 ksi), and at 505 K (450°F) maximum cruise temperature all Phase I specimens achieved 10,000 hours. Residual tensile 

strengths at 505 K (450 OF) ranged from 228 MNlm2 (33 hi) to 414 MNlmz (60 ksi) and at 297 K 

(75 OF) ranged from 228 MNlm2 (33 ksi) to 510 MNlm2 (74 ksi). Reproducibility was quite poor, 

even though the failure mode remained tensile, and no aging or delamination affects were 

observed. Residual compressive strength was reduced by approximately the same degree as 

the tensile strength. This was in contrast to the epoxy systems where compressive strength 

was considerably more affected by flight simulation exposure than the tensile strength. The 

shear strength exhibited the greatest decrease with a loss of more than 50% after 10,000 hours 

of flight cycling. This was an indication that GlPI flight simulation failures could be expected 

early in the Phase I1 test program if the failure mechanism proposed for the epoxy systems 

was operative for the polymide system as well. 

The results of the 10,000-hour thermal aging exposures, Figure 13-12, clearly show the effect of 

aging temperature, with very little decrease in tensile strength when aged at 505 K (450 OF) and 

almost a 50% reduction at 561 K (550°F). 

The post long-term fatigue tests at R = 0.1 and 297 K (75 OF) produced a 0107 value after 10,000 

hours of flight cycling of only 193 MNlm2 (28 ksi) as compared to the baseline values of 310 

MNlm2 (45 ksi) at 297 K (75 OF) and 269 MNlm2 (39 ksi) at 505 K (450 OF). Panel-to-panel scatter 

may have influenced these results. 

Again the effect of compressive stresses was demonstrated by the baseline fatigue tests at 

R = -1. Values of 0107 were 152 MNlm2 (22 ksi) at 297 K (75°F) and 124 MN/m2 (18 ksi) at 

505 K (450°F). 

As expected, long term failures began to occur early in Phase I1 with all specimens eventually 

failing prior to 40,000 hours. The combined effect of fatigue loading, surface oxidation, and 

thermal cycling stresses contributed to these wearout failures. Residual tensile strength of the 

Phase I1 specimens showed a continuing decrease with exposure time, as well as wide scatter 

about the mean. The failure mode was tensile but with a contribution through delamination by 

the combined action of 1) exposed outer ply fibers from surface oxidation, 2) degraded shear 

and compressive strengths with time, and 3) application of stresses from thermal cycling and 

compressive loading. 

13 - 16 
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13 -17 
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13 -19 
As shown in Figure 13-13 the 30,000 hour residual compressive and shear strengths were also 

reduced to less than half of the baseline values. It should be noted that the reduction in shear 

strength had already been observed after only 10,000 hours of exposure. Also shown in Figure 

13-13 are the Phase I1 thermal aging results for 50,000 hours at 505 K (450°F) and 25,000 

hours at 561 K (550OF). Again, the effect of aging temperature is clearly shown with a 90% 

reduction in strength in only 25,000 hours at the higher temperature versus a 20% loss in 

50,000 hours at the lower temperature. Comparison of the 30,000 hour flight simulation 

residual strength with the 50,000 hour thermal aging residual strength, both at the same maximum temperature, shows the importance that fatigue loading and, perhaps, thermal cycling 

have on property degradation - a loss of twethirds compared to only 20%. 

HTW710 G/PI, [OO f 450],, notched. Even these notched specimens suffered from voids, 

porosity, and panel-tepanel variability. As shown in Figure 13-14, their mean strength at 

505 K (450°F) was about 338 MNIm2 (49 hi). The Weibull CY was 12.15. 

The short-term peak stress was 141 MNlmz (20.5 ksi) at 505 K (45OoF), and all specimens had 

about the same strength after 100 and 200 hours. Net tension failures were obtained from ail 

specimens. The long-term peak stress was 125 MNIm2 (18.2 hi). All specimens survived 10,000 

hours at this relatively low stress leveL Residual notched tensile strengths at 505 K (450°F) 

ranged from 255 MNIm2 (37 ksi) to 400 MNlm2 (58 ksi), and at 297 K (75 OF) they ranged from 

248 MN1m2 (36 hi) to 428 MNlm2 (62 ksi). No particular strength shift was seen. but the scatter was clearly increased. The decrease in the average notched tensile strength after 10,000 

hours of flight simulation exposure was quite small compared with that observed for unnotched specimens (Figure 13-12). Similar differences in properties of notched and unnotched 

specimens of both BIE and GIE were also observed after 10.000 hours of flight simulation. 

The post long-term notched fatigue test at R = 0.1 and 297 K (75°F) produced a 0107 of 276 

MNlm2 (40 hi) as compared to a baseline value of 303 MNlm2 (44 ksi) at 297 K (75 OF) and 345 

MNlm2 (50 hi) at 505 K (450OF). Panel variability is evidently operating here to confuse the 

trends. 

The baseline reversed loading tests at R = -1 produced much lower a107 values of 83 MNlm2 

(12 ksi) at 297 K (75 OF) and 90 MNIm2 (13 ksi) at 505 K (450 OF), indicating the severe effect of 

compressive loading. 

In like manner to the unnotched long term HT-SI710 Phase I1 specimens, all of the notched 

specimens also failed prior to reaching 50,000 hours. These wearout failures began as early as 

18,000 hours and by the 28,000 hour point all had failed or had been removed from test because 

of their poor condition. The residual tensile properties, Figure 13-13, of the notched specimens 

were not as severely degraded as those of the unnotched specimens but did decrease with increasing flight cycling time and also exhibited considerable scatter. 

Celion 6000/LARC-160 G/PI, [OO .C 450],, unnotched. The Celion 60001LARC-160 GlPI data are 

summarized in Figures 13-15 and 13-16. The average baseline tensile strength at the flight 

simulation maximum exposure temperature, 450 K (350 OF), was 696 MNlm2 (101 ksi). The CY of 

22 is an indication of the excellent reproducibility of these tests. For the combined 35 baseline 

tensile tests at seven temperatures a pooled CY of 28 was calculated. 

13- 20 
Figure 13-14 Phase I and Phase II Notched [O'J 45O], HT-S/710 GPI 

13 -21 
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Figure 13-15 Unnotched [OO f 4501, Celion 6000/LARC-160 G/PI 

13 -22 
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13 -23 
No short term testing of this system was conducted because of the poor results obtained with 

the HT-SI710 GIPI and because of the delays that would have occurred in flight simulation 

testing of the other systems. The procedure for selecting the long term loads was similar to 

that used for selecting short term loads for the other systems. An initial value of 444 MNIm2 

(64.4 ksi) resulted in an early fdure. With a recalculated peak stress of 243 MNIm2 (34.8 ksi) all 

the specimens completed 10,000 hours. The condition of the specimens was excellent with 

almost no visible surface erosion or attack except for minor surface microcracking. Tensile 

properties were reduced 10 to 15% with increased scatter. The failure mode remained tensile, 

and no aging or delamination effects were found. Again, as with the HT-SI710 GIPI system, 

the compressive strength was not greatly affected, but the residual shear strength was reduced by almost 50% after 10,000 hours. 

Results of residual fatigue tests showed losses in 0107 for tests at R = 0.1 and 297 K (75 OF) of 

about the same magnitude as the residual tensile tests, Le., 10 to 15%. Baseline values of 0107 

for R = -1 were again considerably below those of the R = 0.1 tests. Baseline tests at an R 

value of 10 gave results intermediate to the R = 0.1 and R = -1 values. 

The thermal aging data, Figure 13-16, are not in agreement with the flight simulation data. 

Again, panel to panel variability may be playing a role with these results. In general, the 

residual flight simulation data, tensile, compression, and shear, are superior to the post thermal aging data at equal and higher aging temperatures and for equivalent times. Additionally, 

the material aged at the higher temperature has shown better strength retention than that 

aged at the lower temperature after 10,000 hours. A post-cure effect at the higher aging 

temperature may be contributing to these results. 

Celion 6000/LARC-160 G/PI, [OO f 450],, notched. As shown in Figure 13-17 the mean strength 

of the notched Celion 60001LARC-160 GIPI at 450 K (350°F) was 475 MNlm2 (68.9 ksi). The 

Weibull CY was the same as that for the unnotched specimens, 22. 

, 

The long term peak stress of 191 MNIm2 (27.3 ksi) was calculated using a procedure similar to 

that used for the unnotched specimens. All specimens survived the 10,000 hours in excellent 

condition. Residual notched tensile strengths at 297 K (75OF) ranged from 663 MNlm2 

(96.2 hi) to 389 MNIm2 (56.4 ksi), and at 450 K (350°F) they ranged from 592 MNlm2 (85.9 ksi) 

to 332 MNIm2 (48.2 ksi). The data scatter was significantly increased over that of the baseline 

tests, but the mean values were almost unchanged. This type of behavior after 10,000 hours of 

flight simulation, i.e., larger effects for unnotched specimens than for notched, was also 

observed for both epoxy systems and the other polymide system. 

Residual fatigue testing at R = 0.1 at both 297 K (75 OF) and 450 K (350 OF) gave values of 0107 

approximately 25% lower than the baseline values. 

The significant effect of compressive loading on fatigue strength was again demonstrated with 

0107 values for R = -1 tests well below those for R = 0.1 tests of the baseline material. 

B/Al, [Oo f 4501,. Trend conclusions for B/Al were somewhat hampered by panel variability 

and, in Phase 11, by batch variability. Recall that the Phase 11, 25,000 hour and 40,000 hour 

thermal aging exposures were conducted on a different batch of material than that used for the 

Phase I thermal aging and flight simulation tests. Because the second batch of BIAl was far 

superior to the first batch during thermal aging, comparisons of these data with the flight 

simulation results are of little value and the data are not included in this summary. 

13 -24 
Y 

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100 

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653.217 472 Figure 13-17 Notched [Oo f Go], Celion 6000/LARC-160 G/PI 

13 -25 
100 2001 200 

STATIC SHORTTERM 

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Figure 13-18 Phase I and Phase II Unnotched [Oo f 45O], B/Al 

13 -26 
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13 -27 
Figures 13-18 and 13-19 summarize the data for the crossply lapp. Static strength at 505 K 

(450°F) was about 455 MNlm2 (66 ksi) with an a of 13.7. 

The short-term specimens were cycled to a peak stress of 345 MNIm2 (50 ksi) at 505 K (450 OF). 

Fatigue failures occurred in both sets of specimens, and the two had quite different mean 

strengths. Note the one residual strength value of 323 MN/m2 (46.8 ksi), which was below the 

peak cyclic stress. 

The long-term specimens were cycled to only 208 MN/m2 (30.1 ksi), and no fatigue failures were 

observed during the first 10,000 hours. The 505 K (450 OF) residual tensile strengths were scattered from 262 MN/m2 (38 ksi) to 414 MN/m2 (60 hi), and the 297 K (75 OF) residuals scattered 

from 303 MNlm2 (44 ksi) to 407 MNIm2 (59 ksi). The aluminum surface appeared rough, and 

the fibers had been degraded by the 10,000 hours at 505 K (450°F). Residual compressive tests 

were not successful because of premature shear failures in the doubler bonds. However, the 

specimens did sustain quite high compressive stresses (well in excess of the tensile strength) 

before the doubler bond failure. 

A significant portion of the degradation experienced by the BfAl d,uring flight simulation exposure appears to be directly related to thermal aging effects and probably not a consequence 

of the fatigue loading and thermal cycling. While no 505 K (450°F) thermal aging data were 

available for the first batch of material, interpolation of the 450 K (350°F) and 561 K (550°F) 

data, Figure 13-19, indicates that tensile strength losses equivalent to those determined after 

10,000 hours of flight simulation could also have occurred during thermal aging at 505 K 

(450°F) for 10,000 hours. Similar effects were not observed for the resin matrix systems. 

The post long-term fatigue at R = 0.1 and 297 K (75°F) produced a 0107 value of 276 MNlm2 

(40 ksi) after 10,000 hours as compared to the baseline values of 379 MN/m2 (55 ksi) and 221 

MN/m2 (32 ksi) at 297 K (75°F) and 561 K (550°F) respectively. Again, the baseline reversed 

loading fatigue tests at R = -1 produced reductions in performance. The 0107 values were 

200 MNlm2 (29 ksi) and 173 MNIm2 (25 ksi) at 297 K (75 OF) and 505 K (450 OF) respectively. 

Wearout fatigue failures began during Phase I1 after 17,000 hours of flight simulation cycling, 

and by 32,000 hours all but the three 25,000 hour specimens had failed. Surface roughening 

was severe and extensive longitudinal cracking of the aluminum accompanied by fracture of 

the outer ply boron filaments was visible on all specimens. 

Residual tensile strength decreased with time from a mean of 290 MNlm2 (42.1 ksi) at 17,707 

hours to 222 MNlm2 (32.2 ksi) after 27,402 hours. The limited material available did not allow 

residual strength testing other than tensile. 

Early fatigue failures and a greater than 50% loss in residual tensile strength accompanied by 

severe surface damage after 25,000 hours does not make BlAl an attractive material for 50,000 

hour use in a 505 K (450°F) flight cycle environment. 

B/Al, [O0l6. Figure 13-20 summarizes the trend data for the unidirectional BIAL The residual 

properties after flight simulation and thermal aging exposures are summarized in Figure 

13-21. 

13 - 28 
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STRENGTH, 107 CYCLES 

PHASE II 

LONG TERM 

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Figure 13-20 Phase I and II Unnotched [OO], WAl 

13 -29 
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13 -30 
As no baseline tensile tests were conducted at 505 K (450 OF), the 561 K (550 OF) baseline data, 

shown in.Figure 13-20, were used for selection of the short-term loads. A mean strength of 1300 

MNlm2 (188 ksi) was accompanied by a Weibull (Y of 17.55. Panel variability was again 

observed. The short-term specimens, cycled to a peak stress of 956 MNlm2 (138.8 ksi) at 505 K 

(450 OF), seemed to get stronger with time. One exception was an individual specimen reaching 

only 882 MNlm2 (128 ksi) after 100 hours. 

The long-term specimens were cycled to a peak stress of 747 MNlm2 (108.3 ksi). The residual 

tensile, notched tensile, and shear strengths were degraded approximately 25% by the 10,000 

hours at 505 K (450 OF). Again surface roughness and fiber damage were apparent. 

Again, as observed for the crossply BIAI, the decrease in unnotched tensile strength during 

10,000 hours of flight simulation could be accounted for by thermal aging effects. In fact, this 

loss in strength was somewhat less than would be predicted for a 10,000 hour age at 505 K 

(450 OF). The effect of thermal aging on matrix dominated properties, transverse tensile and interfiber shear, was nil except at 700 K (800 OF) where oxidation of the boron fibers resulted in a 

60% reduction of the interfiber shear strength. 

'Post long-term fatigue at R = 0.1 and 297 K (75OF) achieved a 0107 value of 772 MNlm2 (112 

ksi) after 10,000 hours compared with the 827 MNlm2 (120 ksi) baseline value at 297 K (75 OF) 

and the 455 MNlm2 (66 ksi) baseline at 561 K (550 OF). Scatter is again operating. The R = -1 

baseline fatigue data again demonstrated compression degradation effects. The a107 values 

were 552 MNlm2 (80 ksi) at 297 K (75 OF) and 441 MNlm2 (64 ksi) at 505 K (450 OF). 

The Phase I1 unidirectional BlAl flight simulation exposures were concluded with a wearout 

failure of the last surviving specimen just short of 32,000 hours. As with the Phase I1 

crossplied specimens, only residual tensile strength testing could be performed b&use of the 

limited material available. The loss in tensile strength after 10,000 hours was 25%, whereas a 

loss in tensile strength'of only 30% is indicated for the specimen that failed near 32,000 hours. 

However, surface degradation, including roughening, splitting, and filament fracture, as well 

as edge cracking were greatly increased over that observed after 10,000 hours. 

Again, the BlAl system, either crossplied or unidirectional material, has not shown promise for 

extended usage in a themallfatigue environment with a maximum service temperature of 505 

K (450 OF). 

13 - 31 
SECTION 14 

CONCLUSIONS 

This report describes a program to characterize the properties of several types of advanced 

filamentary reinforced composite material systems before and after exposure to simulated 

supersonic cruise environments for real times of up to 50,000 hours. Because usage of the composites was intended for the lower wing surface of a supersonic aircraft, exposure loads and 

subsequent residual property measurements were largely tensile. However, during the latter 

portion of the program a study was added to specifically examine the effect of thermal exposures on compressive and shear properties. Moisture effects were generally avoided by 

utilizing careful bakeout and storage procedures. Material behavior during these exposures 

and post exposure residual property tests provide exposure effects and reveal material failure 

modes and degradation mechanisms. In addition to the flight simulation exposures, the program provides extensive thermal and ambient aging data for times out to 50,000 hours. This 

report has also presented considerable baseline fatigue, creep, fracture, and tensile data for all 

the systems. In several instances, however, analysis of the test data has been partially 

hampered by panel-tepanel variability of each of the composite systems. Additionally, in two 

cases, G/E and B/Al, where material was purchased at widely spaced intervals, batch-tobatch 

variability prevented direct comparisons of results. The observations and conclusions relating 

to these baseline measurements and the environmental exposure tests are summarized briefly 

below. 

B/5505 B/E and A-S/3501 G/E should be limited to. a temperature of 373K (212°F) for 

cumulative exposures of 25,000 hours for B/E and 36,000 hours for G/E because of: 

1. Early flight simulation test failures (caused by compressive loading and thermal cycling 

combined with oxidation induced matrix degradation). 

2. Loss of residual strength during flight simulation and thermal aging exposures (caused by 

oxidation induced matrix degradation). 

3. Moisture effects on elevated temperature strength (caused by matrix degradation). 

HT-S/710 G/PI should be limited to 505K (450°F) for exposures of 15,000 hours because of: 

1. Loss of residual tensile strength during thermal aging (caused by oxidation induced matrix 

degradation). 

2. Loss of residual strength, primarily matrix dominated, during flight simulation exposure 

(caused by combined compressive and thermal stressing in conjunction with oxidation induced matrix degradation). 

Celion 60001LARC-160 G/PI should be limited to 450K (350°F) for exposures of 10,000 hours 

because of: 

1. Loss of residual tensile, compressive, compressive fatigue, and shear strengths during 

thermal aging (caused by oxidation induced matrix degradation). 

14-1 
2. Loss of residual strength, primarily matrix dominated, during flight simulation exposure 

(caused by combined compressive and thermal stressing in conjunction with oxidation induced matrix degradation). 

B/6061 BlAl should be limited to 505K (450OF) for exposures of 15,000 hours because of: 

1. Loss of residual tensile strength during thermal aging (caused by interface diffusion induced fiber degradation). Some loss in strength can occur at temperatures as low as 450K 

(350 OF). 

2. High temperature fatigue effects (matrix surface cracking and fracture of outer ply boron 

filaments). 

3. Loss of residual strength during flight simulation exposure (caused by interface diffusion 

induced fiber degradation, matrix surface cracking, and fracture of outer ply boron 

filaments). 

B/P 105 AC B/PI is not suitable for this application because of severe matrix degradation during thermal aging and short-term flight simulation exposures at 505K (450 OF). 

For the two epoxy systems the baseline and short-term flight simulation results were used in a 

wearout model to establish the long-term loading conditions. For the GlPI and B/Al materials, 

however, the wearout theory could not be applied because of inconsistent short-term data. The 

wearout model was based on a tensile failure mode and made no consideration for thermal effects. Phase I results for the epoxy systems indicated that a more comprehensive model was 

needed. This new model would be required to account for failure mechanism changes and thermal degradation during the course of the flight simulation exposures in order to more accurately predict complex real-time exposure effects. An alternate approach was found during the 

Phase I1 long term tests of the epoxy systems by simply reducing the maximum temperature 

to a level that precluded any sigruficant thermal effects during the long exposure periods. This 

temperature was determined from a log time versus 1/T plot of thermal aging time 

temperature combinations at which significant epoxy matrix degradation began. Extrapolation of the curve to 50,000 hours gave the revised test temperature. 

The parameters necessary for surface microcracking of the resin matrix composites during 

thermal aging or flight simulation exposure varied for the different systems. For both epoxies 

the combination of thermal exposure and fatigue cycling was required. No surface microcracking was observed for HT-9710 GlPI on either thermal aging or flight simulation specimens. 

Finally, for Celion 6000lLARC-160 GlPI thermal exposure alone was sufficient to cause 

microcracking while fatigue cycling alone was not. Surface cracking of the metal matrix composite, BIAI, was observed on specimens subjected to the combination of fatigue loading and 

thermal exposure. 

For the epoxy systems, thermal aging at 394K (250 OF) and one atmosphere pressure produced 

no effects for the first 10,000 hours. Matrix degradation began between 10.000 and 25,000 

hours and was severe after 50,000 hours. The fiber-controlled tensile properties, however, 

showed almost no change. Aging at 450K (350°F) and one atmosphere pressure was more 

damaging, with matrix degradation beginning between 1000 and 5000 hours. After 5000 

hours, the matrix was severely embrittled and crumbled away during tensile testing, leaving 

14-2 
many bare fibers. Tensile properties were considerably reduced for aging times of 5000 hours 

or longer. Reduced pressure exposures at 450K (350°F) delayed the effects but did not 

eliminate them. 

The HT-S/710 G/PI system survived 25,000 hours of thermal aging at 505K (450 OF) with no effects. Some decrease in tensile strength was measured after 50,000 hours, but matrix degradation was not observed. Raising the aging temperature to 561K (550°F) reduced the time at 

which tensile strength decreases were observed to 10,000 hours. In like manner to the 505K 

(450°F) exposures, this initial fall-off in tensile strength was not accompanied by matrix embrittlement. After 25,000 hours the material was partially delaminated and showed high 

weight loss. The tensile strength was greatly reduced and severe matrix embrittlement had occurred. Degradation after 50,000 hours was such that the specimens could not be tested. 

Again, aging effects were less severe for exposures conducted in a reduced pressure environment. 

The effects of 10,000 hours of thermal aging at 450K (350°F) and 505K (450°F) in one atmosphere air on Celion 60001LARC-160 GlPI were significant losses in tensile strength, beginning as early as 500 hours, and some matrix degradation. Tensile strength losses were higher 

for the 450K (350 OF) exposures, 35%, than for those at 505K (450 OF), about 20%, and were accompanied by greater matrix degradation. Weight loss during aging, however, was higher at 

the higher temperature. For the reduced pressure exposures no matrix degradation was 

observed, and strength losses were only of the order of 10% after 10,000 hours. The 10,000 

hour performance of the Celion 60001LARC-160 GlPI in one atmosphere air at 450K (350OF) 

was no better than that of G/E and at 505K (450°F) was considerably worse than that of 

HT-S/710 GIPI. 

The Bl6061 borodaluminum used in the Phase I1 portion of the thermal aging evaluation 

(batch 2) was much superior to that used during Phase I (batch 1). The Phase I data indicated a 

maximum use temperature of only 450K (350°F) for exposures longer than 10,000 hours. The 

Phase I1 data, on the other hand, indicate that a maximum use temperature of 505K (450OF) 

for 40,000 hours of exposure is possible. Some precautions may be required, however, to protect the material edges to prevent oxidation of the boron fibers. 

Results of the thermal agingkompression study showed that 10,000 hours of thermal aging in 

one atmosphere air at 373K (212°F) and 394K (250°F) produced no significant reduction in 

compressive strength or compressive fatigue strength of GIE. The major effect of the thermal 

exposure was a considerable decrease in shear strength of 6 ply material. These reductions 

were observed as early as 1000 hours and increased to 40% or more by 10,000 hours. In contrast to the GIE system, thermal aging of Celion 60001LARC-160 GlPI at 450K (350°F) and 

505K (450 OF) produced significant reductions in compressive, shear, and fatigue strengths 

after 10,000 hours of exposure. In like manner to the thermal aginghension results for this 

material, the compressive and 6 ply shear strength losses were higher for the 450K (350°F) 

age. The other properties, however, showed greater decreases at the higher aging temperature. 

Weight loss was also greater at the higher temperature. For the 6 ply specimens these losses 

were quite large, greater than 8%. Again, as with the GIE system, the property most affected 

was the shear strength of the 6 ply material. Losses of 65% were found after 10,000 hours at 

450K (350 OF). Extensive edge cracking, particularly in the 24 ply specimens, was fovnd as early as 1000 hours into the exposure. After 10,000 hours these cracks extended throughout the 

matrix. One set of 6 ply specimens was coated with a solution of a polyphenylquinoxaline 

(PPQ) in an effort to improve the thermooxidative stability of the GlPI material. Preliminary 

14-3 
results were promising, with slight improvements over the uncoated specimens in almost 

every case. 

For the resin matrix systems, matrix degradation by oxidation was shown to be a major contributor to mechanical property losses during thermal aging. For GIE and Celion 

6000lLARC-160 GlPI the extent of oxidation could readily be detected by metallographic 

techniques, especially with the SEM. No evidence of time dependent oxidation similar to that 

for the GIE and Celion 6000lLARC-160 GlPI systems was observed for either the BIE or 

HT-SI710 GIPI thermal aging specimens. Examination of the BIE samples gave almost no information, while studies of HT-SI710 GlPI revealed increased porosity and fiber-matrix separation accompanied by numerous fine cracks at the fiber-matrix interface as thermal exposure 

time increased. 

For the BlAl metal matrix system, thermal aging strength decreases were primarily fiber 

related rather than caused by degradation of the aluminum matrix by oxidation. Reactions at 

the boron filament/aluminum matrix interface during high temperature exposure led to 

changes in the transverse tensile and shear failure modes from primarily matrix to increased 

contributions from interfacial fracture and, for the Phase I batch, filament splitting. It was 

this observation that indicated that the boron fibers in the Phase I1 BIAl had a considerably 

improved thermal stability over that of the Phase I fibers. 

Fatigue strength was found to be very dependent on the stress ratio used in testing. Results 

for tension-compressive loading (R =-1) and compressivecompressive loading (R = 10) were 

considerably lower than those obtained for tension-tension loading (R=0.1). The R=10 S-N 

curves were just slightly higher than those at R=-1 for Celion 6000/LARC-160 G/PI, the only 

system tested at all three R values. The effect of a notch was generally to reduce the fatigue 

strength. However, some exceptions were observed at elevated temperature. Severe degradation of the aluminum matrix was found after fatigue testing at 561K (550°F). Testing in an 

argon or nitrogen atmosphere or lowering the temperature to 505K (450 OF) significantly 

decreased the matrix degradation. 

Absorption of moisture by the epoxy systems caused a significant decrease in short time 

elevated temperature strength. Moisture effects were considerably less damaging for the GiPI 

systems, particularly for the Celion 6000ILARC-160. Loss of resin from the surface sufficient 

to expose a large number of graphite fibers was experienced by both epoxy and polyimide 

matrix composites when exposed for 50,000 hours to weather and atmospheric contaminants. 

Significant losses in elevated temperature flexural strength, probably as a result of absorbed 

moisture, were also found. 

Uncoated BlAl is subject to pitting and subsequent intergranular corrosion cracking when exposed for long periods in an industrial-seacoast environment. These corrosion cracks, many extending to a depth of 0.00018m (0.007 in.) after 50,000 hours of exposure, can seriously degrade 

the matrix controlled properties by acting as crack starters. An aluminized polyimide coating 

developed for high temperature corrosion protection was found to prevent this pitting and intergranular corrosion cracking during the 50,000 hours of outdoor exposure. 

Boron or graphite reinforced composite materials with 0" plies aligned in the direction of 

loading undergo very little creep even at stress levels, which result in facture in short exposure 

times. Because of the low creep strains observed in these materials it is suggested that stress 

rupture tests be substituted for creep measurements. 

14-4 
A major goal of this program was to develop a predictive capability such that by properly setting the experimental conditions (loads and temperatures) a number of advanced composite 

materials could complete 50,000 hours of real-time exposure to a supersonic cruise environment with an 80% survival rate. This goal was not accomplished. Two factors were largely 

responsible: 

1. The inadequacy of the wearout theory because no consideration had been given to thermal 

effects and the attendant failure mechanism changes during the course of the exposures. 

2. Data scatter of the composite systems, including in-panel variability, panel-to-panel 

variability, and batch-to-batch variability. 

One material, however, did show a potential for successfully surviving 50,000 hours of flight 

simulation. The GIE system completed 36,000 hours at a maximum flight simulation 

temperature of 373K (212°F) with the only significant effect being a 28% loss in shear 

strength. As the tests were halted prematurely (to prevent further BIE failures) the possibility 

exists that the G/E specimens may have lasted the full 50,000 hours of exposure. 

14-5 
SECTION 15 

1. 

2. 

3. 

4. 

5. 

6. 

7. 

8. 

9. 

10. 

11. 

12. 

13. 

Hoffman, S.: Bibliography of Supersonic Cruise Aircraft Research (SCAR) 

Program From 1972 to Mid-1977, NASA RP 1003, 1977. 

Hoffman, S.: Bibliography of Supersonic Cruise Research (SCR) Program 

From 1977 to Mid-1980, NASA RP 1063, 1980. 

Hoffman, S.: Bibliography of Supersonic Cruise Research (SCR) Program 

From 1980 to 1983, NASA RP 1117, 1984. 

Imig, L. A.; and Illg, W.: Fatigue of Notched Ti-8AI-lMo-lV Titanium Alloy 

at Room Temperature and 550 OF (560 OK) With Flight-by-Flight Loading' 

Representative of a Supersonic Transport, NASA TND-5294, 1969. 

Imig, L. A.: and .Garrett, L. E.: FatigueTest Acceleration With Flight-byFlight Loading and Heating to Simulate Supersonic-Transport Operation, 

Halpin, J. C.; Jerina, K. L.; and Johnson, T. A.: Analysis of the Test 

Methods for High Modulus Fibers and Composites, ASTM STP 521, 

American Society for Testing and Materials, 1973, pp. 5-64. 

Structural Design Guide for Advanced Composite Applications, Second 

Edition, Advanced Composites Division, AFML, 197 1. 

Hertz. J.; Christian, J. L.; and Varlas, M.: Advanced Composite Applications for Spacecraft and Missiles, AFML-TR-71-186, 1972. 

Christian, J. L.: Development and Evaluation of Advanced Metal-Ma trix 

Composite Materials, Rep. GDC-ERR-1571, General DynamicsiConvair, 

1970. 

NASA TND-7380, 1973. 

Forest, J. D.; and Christian, J. L.: Development and Application of AluminudBoron Composite iMateria1, AIAA Paper No. 68-975, October 1968. 

Christian, J. L.; and Jones, R. C.: Development and Evaluation of Improved 

BorodAluminum Composite Material, Rep. GDC-ERR-1450, General 

DynamicslConvair, 1970. 

Forest, J. D.; and Christian, J. L.: Development and Application of High 

Matrix Strength AluminudBoron, presented at ASM Metals Engineering 

Congress, October 1969. 

Christian, J. L.: Material and Fabrication Development and Application 

Studies of AluminudBoron-S tainless S tee1 Composites, presented at 

ASME 15 th Annual International Gas Turbine Conference, Brussels, 

Belgium, May 1970. 

15-1 
14. 

15. 

16. 

17. 

18. 

19. 

20. 

21. 

Miller, M. F.: Schaefer, W. H.: Weisinger, M. D.; et al: Development of 

Improved Metal-Matrix Fabrication Techniques for Aircraft Structures, 

AFML-TR-7 1- 18 1 , 197 1. 

Shockey, P. D.: Hofer, K. E.; Kaminski, B. E.: Lemon, G. H.; and McKague, 

E. L.: Development of Engineering Data for Advanced Composite 

Materials, VoL I-Static and Thermophysical Properties, Vol. 11-Fatigue and 

Load Rate Characteristics, AFML-TR-70-108, 1972. 

Shockey, P. D.; Hofer, K. E.; and Wright, W. D.: Structural Airframe Application of Advanced Composite Materials, VoL IV-Mechanical PropertiesStatic, AFML-TR-69- 10 1. 1969. 

Development of Engineering Data on the Mechanical and Physical Properties of Advanced Composites Materials, AMFL-TR-72-205 - Part I, 1972. 

Advanced Composites Data for Aircraft Structural Design, Report No. 

NA-71-679-9, North American Rockwell, 1972. 

Beeler, D. R.; and Chase, V. A.: Advanced Polyimide Composites, SAE Conference, Oct. 1968. 

Vaughan, R. W.: Glass Reinforced P13N Polyimide Laminates Having 

Improved Processing Characteristics, 15 th National SAklPE Conference, 

April 1969. 

Vaughan, R. W.; Jones, R. J.; Sheppard. C. H.; and Burns, E. A.: Development of Low Void Polyimide Resin for Autoclave Processing of Glass and 

Graphite Reinforced Composites, National SAMPE Conference, Oct. 197 1. 

22. Petker, I.; et al: High-Strength, High-Modulus Filament-Reinforced Composi tes, AFML-TR-67-395, 1968. 

23. Birchfield, E. B.; and Kollmansberger, R.: Develop Fabrica tion/Processing 

Techniques for High Temperature Advanced Composites for Use in Aircraft Structures. AFML-TR-72-91, 1972. 

24. Scheck, W. G.: Development of Design Data for Graphite Reinforced Epoxy 

and Polyimide Composites, Rep. GDC-DBG70-005, General Dynamics/ 

Convair, 1974. 

25. Stuckey, J.; and Scheck, W. G.: Development and Evaluation of Graphite 

and Boron Polyimide Composites, National SAMPE Technical Conference, 

Oct. 1972, Palo Alto, California. 

26. Prunty, J.: Design, Fabrication, and Test of a Graphite/Epoxy Metering 

Shell (GEMS), Rep. CASD-NAS75-015, General DynamicsiConvair, 1975. 

27. Chase, V. A.; and Harrison, E. S.: Develop and Demonstrate Manufacturing 

Processes for Fabricating Graphite Filament Reinforced Polyimide (Gr/Pl) 

Composite Structural Elements, NASA CR-172497, February 1985. 

28. Peterson, R. E.: Stress Concentration Design Factors. John Wiley and 

Sons, Inc., 1953. 

15-2 
29. 

30. 

31. 

32. 

33. 

34. 

35. 

36. 

37. 

38. 

39. 

40. 

41. 

42. 

Waddoups, M. E.; Eisenmann, J. R.; and Kaminski, B. E.: Macroscopic 

Fracture Mechanics of Advanced Composite Materials. J. Composite 

Materials, 5, 1971, p. 446. 

Konish, H. J., Jr.: A Study of Fracture Phenomena in Fiber Composite 

Laminates, AFML-TR-73-145, Vol. 111, 1973. 

Adsit, N. R.; and Waszczak, J. P.: Fracture Mechanics Correlation of Boron/ 

Aluminum Coupons Containing Stress Risers, ASTM Symposium on Fracture Mechanics of Composites, 1974. 

Paris, P. C.; and Sih, G. C.: Stress Analysis of Cracks. Fracture Toughness 

Testing and its Applications, ASTM STP 381, American Society for Testing and Materials, 1970, pp. 30-83. 

Chase, V.; and Beeler, D.: Manufacturing Methods for Large High-Temperature Sandwich Structures, AFML-TR-70-211, 1970. 

Jones, R. L.: Evaluation of Electron Scanning Microscopes, Rep. ERRFW-1188, General DynamicslConvair, 1972. 

Metcalfe, A. G., ed.: Interfaces in Metal Matrix Composites, Academic 

Press, 1974. 

Hertz, J.: Investigation Into the High Temperature Strength Degada tion 

of Fiber-Reinforced Resin Composites During Ambient Aging, NASA 

CR-124290. 1973. 

Shirrell, C. D.: Halpin, J. C.; and Browning, C. E.: Moisture - An Assessment of Its Impact on the Design of Resin Based Advanced Composites, 

Third Conf. on Fibrous Composites in Flight Vehicle Design. November 4-6, 

1975, NASA TMX-3377, April 1976. 

Schaefer, W. H.; and Christian, J. L.; et al.: Evaluation of the Structural 

Behavior of Filament Reinforced Metal Matrix Composites, AFRILTreadway, D. G.; Sutherland, W. M.; and Keller, E. E.: Aerospace Corrosion 

Prevention Studies, Rep. GDCA-ERR-16 12, General Dynamics/Convair. 

1971. 

TR-69-36, 1969. 

Shockey, P. D.; Anderson, J. D.; and Hofer, K. E.: Structural Airframe 

Application of Advanced Cpmposite Materials - Vol. V, Mechanical 

Properties-Fa tigue, AFML-TR-69- 10 1. 1970. 

Shockey, P. D.; and Hofer, K. E.: Development of Engineering Data for 

Advanced Composite Materials, Vol. 11, Fatigue and Load Rate Characteristics, AFML-TR-70-108, 1972. 

Military Standardization Handbook, Metallic Materials and Elements for 

Aerospace Vehicle Structures, MIL-HDBK-SB, Department of Defense, 

Washington, D. C., 1975. 

15-3 
43. Wolff, R. V.; and Lemon, G. H.: Reliability Prediction for Adhesive Bonds, 

AFML-TR-72-121, 1972. 

44. Waddoups, M. E.; Wolff, R. V.: and Wilkins, D. J.: Reliability of Complex 

Large Scale Composite Structure-Proof of Concept, AFML-TR-73-160, 

1973. 

45. Halpin, J. C.; Johnson, T. A.; and Waddoups, M. E.: Kinetic Fracture 

Models and Structural Reliability, International Journal of Fracture 

Mechanics, Vol. 8, 1972. 

46. Wolff, R. V.; and Lemon, G. H.: Reliability Prediction for Composite JointsBonded and Bolted, AFML-TR-74-197, 1974. 

47. Wolff, R. V.; Wilkins, D. J.; Lemon, G. H.; Orringer. 0. R.; and Shinozuka, 

M.: Life Assurance of Composite Structures, Vol. I1 - Technology, Vol. I11 

-Element Characterization, AFFDL-TR-77-50, 1977. 

15-d 
APPENDIX A 

AN ANALYSlS OF THE B/PlOSAC BORON POLYIMIDE SYSTEM 

Approximately midway through Phase I of the program the decision was made to cancel all 

testing of the BlPlOBAC borodpolyimide system. The removal of the BlPI was based on very 

low tensile properties found for the majority of the panels and extensive thermal degradation 

observed during thermal aging and short-term flight simulation testing. The following is a 

brief description of the problems experienced with this material system, and what are believed 

to have been the reasons for the poor results. 

In mid 1973, 28 kg (62 lb) of BlPlOBAC prepreg was ordered from AVCO Corporation with 

delivery expected in one to two months. Because of difficulties encountered by AVCO in 

obtaining the polyimide resin from the manufacturer, Ciba-Geigy, delivery was delayed until 

the end of 1973. Upon receipt of the prepreg, quality assurance testing was performed. The 

material was found to be acceptable. Two,quality assurance panels were then fabricated. 

Before quality assurance testing could be completed, however, Ciba-Geigy informed us that 

they would no longer produce the P105AC resin (see letter on Page A-2). and as they were the 

only source of this material, testing was put on hold. 

After a four-month delay, permission was received from NASA to resume fabrication and testing as it was thought that AVCO Corporation would obtain a license to produce the resin. The 

quality assurance panels were ultrasonically C-scanned, specific gravity and resin, fiber, and 

void contents were determined, and quality assurance mechanical property data were 

obtained. The results (see Tables 6-8 and 6-9) indicated the materials to be acceptable. Only the 

room temperature shear strength of the unidirectional material was on the low side, 55 MN/m2 

(8 ksi), compared to an expected value of about 69 to 83 MNlm2 (10 to 12 ksi). Based on the 

quality assurance tests, the required number of full s'ize panels were fabricated. 

At about this time in the program, mid 1974, the start of short-term testing of B/PI had fallen 

well behind schedule. Since only crossplied specimens were to be used in the short-term tests, 

the main effort in baseline testing was, therefore, on the crossplied material. Baseline tensile 

and notched tensile tests were performed. Results are given in Tables A-1 and A-2. The data 

were typical of good material; consequently short-term specimens were prepared and installed 

in the flight simulation apparatus. 

Baseline testing of the unidirectional material, which previously had been given low priority, 

was now started. Also during this time period, the end of 1974, results from the thermal aging 

tests were becoming available, and proof testing of the short-term specimens was begun. All 

three of these tests gave indications of a potential problem with the BlPI material system. 

Unidirectional tensile strengths, Table A-3, were only half of what was considered normal. The 

notched unidirectional tensile strength, Table A-4, were somewhat higher but still unacceptable. The three notched specimens tested at room temperature gave particularly low values. 

These specimens and all of the unnotched specimens were cut from panel LRC-32. The remaining notched specimens were cut from panel LRC-34. 

A- 1 
Plastics Operations 

CIBA-GEIGY Corporation 

Ardsley. New York 10!502 

914 470331 

CIBA-GEIGY 

November 15, 1973 

Dear Customer, 

It is necessary for us to discontinue one of our polyimide 

products, PlOSAC, when present stocks are exhausted. This 

is occasioned by the unanticipated withdrawal from commerce 

of one of the monomers required, thiodianiline. P105A, an 

earlier version, was replaced by PlOSAC early in 1973. The 

former product also was based on this monomer and cannot be 

re-introduced. Adhesive composition, LSU-1106, also containing thiodianiline is similarly affected. 

Our P13N polyimide family contains no thiodianiline and is 

not affected. We encourage use of P13N in your applications. 

If changes in processing are reauired due to different flow 

characteristics, our Development & Applications Laboratory 

stands ready to render necessary assistance. 

We are sorry for any inconvenience that this reluctant but 

unavoidable decision may have caused. CIBA-GEIGY appreciates 

your valued business and looks forward to providing continuing 

requirements for high quality engineering resins. 

very truly yours, 

v 

Stephen S. Hirsch 

Manager, Business Development 

SSH:jk 

A-2 
Tablo A-1. Basolino Tonsilo Proportks [OO f 45O], B/PI 

Temporuturo Tonsilo Strength Tensile Modulus 

Numbor K (OF) MN/d (kri) GN/d (Msi) 

Spocimon 

cco1-1 

-2 

-3 

CCO1-4 

-5 

-6 

CCO1-7 

-8 

-9 

-22 

-23 

-24 

-25 

-26 

-27 

-28 

-29 

-30 

-3 1 

-32 

-33 

-34 

-35 

cco1-10 

-1 1 

-12 

CCO1-13 

-14 

-15 

CCO1-16 

-17 

-18 

cco1-19 

-20 

-21 

218 -67 450 

478 

513 

av 480 

297 75 510 

505 

565 

av 527 

505 450 490 

484 

547 

539 

518 

478 

538 

565 

549 

494 

443 

544 

525 

555 

562 

522 

494 

av 520 

533 500 

561 550 

589 600 

616 650 

a Strain gage measurement 

50 1 

465 

483 

av 483 

446 

437 

447 

av 443 

415 

452 

463 

av 443 

444 

468 

482 

av 465 

65.3 

69.4 

74.4 

69.7 

74.0 

73.2 

81.9 

76.4 

71.1 

70.2 

79.4 

78.2 

75.2 

69.4 

78.1 

81.9 

79.6 

71.6 

64.2 

78.9 

76.2 

80.5 

81.5 

75.7 

71.7 

75.5 

72.6 

67.5 

70.1 

70.1 

64.7 

63.4 

64.8 

64.3 

60.2 

65.6 

67.1 

64.3 

64.4 

67.9 

69.9 

67.4 

- 

- 

- 

- 

- 

- 

- 

12.7 

11.5 

12.2 

12.1 

a15.3 

a16.0 

a13.2 

14.8 

a 9.5 

a 9.5 

a 9.0 

- 

- 

- 

- 

- 

- 

- 

- 

- 

- 

- 

- 

- 

- 

- 

- 

- 

9.3 

- 

- 

- 

a 8.4 

a 8.0 

a 8.1 - 

8.2 

- 

- 

- 

- 

- 

- 

A-3 
Tablo A-2. Barolino Notehod Tonsil. Pr0porti.r of [Oo -C 45O], B/PI 

Tomproturo Spuimon 

Numbor K (Of) 

cco2-1 

-2 

-3 

CCO2-4 

-5 

-6 

CCO2-7 

-8 

-9 

-22 

-23 

-24 

-25 

-26 

-27 

-28 

-29 

-30 

-3 1 

-32 

-33 

-34 

-35 

cco2-10 

-11 

-12 

CCO2-13 

-14 

-15 

CCO2-16 

-17 

-18 

218 -67 

av 

297 75 

av 

505 450 

av 

533 500 

av 

561 550 

av 

589 600 

av 

444 (a) 

432 

458 

445 

503 

424 

404 

444 

396 

379 

377 

383 

380 

342 

381 

379 

399 

372 

367 

372 

358 

369 

361 

356 

372 

373 

416 

395 

423 

411 

392 

398 

381 

390 

396 

388 

388 

391 

CCO2-19 616 650 437 

-20 425 

-21 459 

av 440 

64.4 

62.7 

66.4 

64.5 

73.0 

61.5 

58.6 

64.4 

57.5 

54.9 

54.7 

55.5 

55.1 

49.6 

55.2 

54.9 

57.8 

54.0 

53.3 

54.0 

52.0 

53.5 

52.4 

51.6 

54.0 

54.1 

60.4. 

57.3 

61.3 

59.7 

56.9 

57.7 

55.3 

56.6 

57.5 

56.3 

56.3 

56.7 

63.4 , 

61.7 

66.5 

63.9 

a Net section strength, Le., based on total width less hole diameter. 

A-4 
Tablo A-3. Barelino Tensile Propertior [0°]6 B/PI 

Spwimon 

Number 

Temporaturo 

K (OF) 

Tonrile Strongth 

MN/d (kri) 

mol-22 

-23 

-24 

-25 

-26 

-27 

-28 

-29 

-30 

mol-10 

-11 

-12 

mol-16 

-17 

-18 

mol-19 

-20 

-21 

505 450 

533 500 

561 550 

589 600 

848 

765 

717 

800 

772 

779 

945 

572 

657 

av 761 

630 

627 

910 

av 722 

510 

606 

33 1 

av 482 

421 

690 

745 

av 619 

123 

111 

104 

116 

112 

113 

137 

82.9 

95.3 

110 

91.4 

.91.0 

132 

105 

74.0 

87.9 

48.0 

70.0 

61.0 

100 

108 

89.7 

- 

- 

- 

- 

Results of thermal aging exposures at 505 K (450" F) and 561 K (550' F) for times to 1000 

hours were also very poor. Tensile testing of specimens aged for 500 hours gave low strength 

values with the specimens literally falling apart during the tests. Specimens aged 1000 hours 

failed by delamination during the doubler bonding operation and could not even be tested. 

The third indication of unsatisfactory material came during the short-term tests. At 75% of 

the proof load used for the short-term exposures, four of the specimens failed. The stress levels 

were approximately 350 MN/m2 (50 ksi), well below the average F, of 520 MN/m2 (76 ksi) 

obtained from the baseline tests. The appearance of the failed specimens, which prior to proof 

loading had been baked out for 72 hours at 478 K (400' F) to remove any absorbed moisture, 

was similar to those from the thermal aging tests. The resin had crumbled away in the region 

of the failures leaving bare boron filaments. 

At the beginning of the short-term testing of the G/PI system, a few of the B/PI specimens 

that had survived the proof test were included in with the G/PI specimens. One of these failed 

after only seven hours, while the remaining ones were badly damaged and about to fail. The 

appearance was similar to those described above with the resin breaking away from the boron 

filaments. Figure A-1 shows the heated area of one of the unfailed specimens. It is held 

together by the boron filaments with only a small amount of resin remaining. 

A-5 
tablo A-4. Bardino Notehod tonsilo Proportior of [0°]6 B/PI 

tompomturo Spoeimon 

Numbor K (OF) 

CUO2-4 

-5 

-6 

CUO2-7 

-8 

-9 

-22 

-23 

-24 

-25 

-26 

-27 

-28 

-29 

-30 

-3 1 

-32 

-33 

-34 

-35 

CUo2-10 

-1 1 

-12 

CUO2-13 

-14 

-15 

CUO2-16 

-17 

-18 

CUo2-19 

-20 

-21 

297 75 717 (a) 

557 

469 

av 581 

505 450 669 

896 

862 

1250 

1320 

1190 

703 

1230 

1080 

1100 

1110 

1120 

903 

910 

882 

827 

862 

av 995 

533 500 

56 1 550 

589 600 

616 650 

1240 

1200 

1070 

av 1170 

1190 

1130 

952 

av 1090 

979 

1110 

1040 

av 1040 

910 

1010 

979 

av 966 

104 

80.8 

68.0 

84.3 

97.1 

130 

125 

181 

191 

173 

102 

179 

156 

159 

161 

162 

131 

132 

128 

120 

125 

144 

180 

174 

155 

170 

173 

164 

138 

158 

142 

161 

151 

151 

132 

146 

142 

140 

a Net section strength, i.e., based on total width less hole diameter. 

A-6 
Ftgure A-1 Hd Ana of WPI Shor)-Term Specimen Aftor Only 7 hr of Flight Simulation Exposure at 

505 K (4500 F) 

To check the extent of the material problem and determine the cause, the following tasks were 

performed: 

a Measured tensile properties of the remaining panels. 

b. Measured specific gravity, resin content, and fiber content of all tested panels. 

c. Conducted metallographic and SEM evaluations. 

The results of the tensile tests are listed in Table A-5. Only two panels had acceptable tensile 

properties. LRC-33 and LRC-35. It was from LRC-33 that the crossplied baseline tensile 

specimens were cut. 

Values of specific gravity, resin content, fiber content, and percent voids are listed in Table 

A-6. There appears to be no correlation between these values and the tensile data of Table A-6. 

Two specimens were submitted for metallographic examination, one that had been thermal 

aged at 561 K (550" F) for 1000 hours, and a baseline tensile specimen. The aged specimen 

could not be sectioned or mounted as the resin matrix crumbled away during preparation. The 

tensile specimen, shown at lOOX magnification in Figure A-2, revealed several points of interest and offered possible explanations for the poor behavior of the material, First, the spacing of 

the filaments within the plies was very nonuniform with many of the filaments touching each 

A-7 
table A-5. tensile Data for B/PI Panels tested at 297 K (750 F) 

Panel 

Number Oriontation 

tensile Strongth Intmdd 

Use 

LRC-32 [0oi6 1100 159 0, 7 

965 140 

896 130 

av 

-- 

987 143 

LRC-33 [0 " 2 45 "1, 510 74.0 

505 73.2 

565 81.9 

av 

-- 

527 76.4 

LRC-34 [o 0i6 1360 198 0, 7 

1190 172 

1160 169 

av 

-- 

1240 180 

LRC-35 

LRC-36 

LRC-37 

LRC-38 

LRC-39 

LRC-40 

LRC-41 

[0" f 45'1, 

[O" +: 45'1, 

[O" f 4501, 

[0" A 45'1, 

[o 0i6 

[O" f 4501, 

[Oo f 45'1, 

530 76.8 

512 74.2 

512 74.2 

av 

-- 

518 75.1 

415 60.2 

410 59.4 

av 

-- 

412 59.8 

461 66.8 

486 70.5 

av 474 68.6 

387 56.2 

389 56.4 

av 388 56.3 

917 133 

917 133 

-- 

-- 

876 127 

av 

-- 

906 131 

412 59.7 

415 60.2 

432 62.6 

av 420 60.8 

360 52.2 

395 57.3 

av 378 54.7 

-- 

-- 

0 

0. 7 

8 

8 

8 

1. 5 

0, 5 

1 

A-8 
Tablo A-5. Tonsilo Data for WPI Pads Tostod at 297 K (750 F) - Concludod 

Pand Tensilo Strongth Intondod 

Us. 

Numbor Oriontatlon MN/d (kri) (4 

LRc-42 [O" 2 45"], 436 - 63.2 9 

LRc-43 [0" f: 45'1, 398 57.7 9 

437 63.4 

av 

-- 

436 63.3 

412 59.7 

av 

-- 

405 58.7 

a0 

1 

2 

5 

7 

8 

9 

Baseline 

Thermal Aging 

Ambient Aging 

Fatigue 

Short-term Exposure 

Long-term Exposure 

tablo A-6. Spuiflc Gravitlos and Resin, Fibor, and Void Contonts of B/Pl Panels 

~ 

Fibor Content 

Pond Spocific Resin Content (wt %) Void Content 

Numbor Oriontation Gravity (wt %) (4 (vol. %) 

LRC32 

LRC-33 

LRC34 

LRC-35 

LRC-37 

LRC38 

LRC-39 

LRC-40 

LRC-4 1 

[o 0i6 

[o 0i6 

[O" f 4501, 

[O" f 45"], 

[0" f 45O], 

[0 O f 45 "Is 

to 0i6 

[0" f 45'1, 

[O" f 4501, 

1.98 28.4 

1.97 32.5 

1.93 26.9 

1.94 30.6 

1.91 33.5 

1.98 26.3 

1.98 27.0 

1.90 37.2 

1.91 31.0 

71.6 

67.5 

73.1 

69.4 

66.5 

73.7 

73.0 

62.8 

69.0 

3.6 

0.8 

7.0 

3.8 

3.4 

4.7 

4.2 

1.0 

4.9 

a Includes approximately 9% glass fiber from scrim cloth. 

A-9 
Figun A-2 Photomicrograph of a Transverse Sodion of B/PI Showing Nonuniform Filamont Spacing 

and an Indication of Poor Bondlng of tho Rosin to tho Boron Filaments 

other. For example, in this photomicrograph, as many as six touching filaments in a row can be 

seen. Efficient distribution of stresses within a composite and, hence, high strength, particularly in the transverse directions, requires good bonding of the matrix to the filaments with a 

minimum of filament-to-filament contact. The second, and more important feature of the 

photomicrograph, was the apparent lack of contact or bonding between the fibers and the resin 

matrix. This is difficult to show conclusively with a light microscope with its poor depth of 

field. Because of the significance of this apparent lack of bonding, additional examinations 

were made using a scanning electron microscope (SEM), an instrument with excellent depth of 

field. 

The photomicrographs in Figures A-3 and A-4 clearly show examples of the separation 

between filaments and the matrix in both unidirectional and crossplied specimens. Further evidence of poor bonding of the resin to the boron is shown in Figures A-5 and A-6 by the pullout 

of the filaments from the matrix and the fact that the filaments are smooth with almost no 

resin adhering to the surface. 

It is this lack of bonding of the resin to the filaments that is believed to have been responsible 

for the poor tensile behavior of the B/P105AC system. Whether the problem was caused by the 

filaments (dirty or contaminated surface, for example) or by the resin is not known. There had 

been reported during this time period, however, instances of the P105AC resin going bad in 

storage, and this may well have been the case for the batch of material evaluated on thi5 

Pro@- 

The reason for the unsatisfactory performance of the specimens during thermal aging and 

short-term flight simulation was apparently a combination of two effects. These were the poor 

bonding of the matrix to the filaments discussed above combined with a resin matrix material 

that was not capable of extended service at 505 K (450" F). 

__ - --_ - - __ ~ 

----- - _- ~ 

A-10 
ORIGINAL PAGE IS 

QE l3OR QUNm 

Figun A4 Fractun Surface of [00l6 B/PI Tonsilo Spocimen Showing (arrowj Poor Bonding Botwoon Boron 

Filamont and Resin Matrix 

Figun A-4 Fracturo Surfaco of [P f 4PI1 B/PI Tonsile Spociman Showing (arrows) Poor Bonding 

Botwoon Filamonts and Rosin Matrix 

A-11 
ORIGINAL PAGE IS 

QE POOR QUALITY, 

, 

Flguro A-5 Frochrro Surfaco of [PI6 B/PI Tonsilo Spocimon Showing Pullout of the Boron Filamonts from 

tho Matrix (an indication of poor matrix-to-filamont bond) 

Flgun A-6 B/PI [PI6 Spoeimon Tonsilo Tostod at 505 K (450° F) After a 500-hr Thormal Ago at 505 K 

(4500 F). (Most of tho resin has broken away, lewing bar. boron filamontr and scrim cloth) 

.. .. 

A-12 
APPENDIX B 

QUALITY ASSURANCF TEST PROCEDURES 

B.1 PREPREG VOLATILE CONTENT 

Cut three specimens of the material, each approximately 0.051 by 0.051 m (2.0 by 2.0 h), from 

an area of the sample away from the edges. The specimen may be any shape that will fit into a 

preweighed and fired crucible or disposable aluminum dish but may not be rolled up or more 

than one ply. Weigh each crucible on an analytical balance to the nearest 0.1 mg ( W2). Place the 

test specimen in the crucible and record the total weight to the nearest 0.1 mg (Wl). Place the 

crucible containing the specimen in a preheated air circulating oven (for epoxy resin use 

436 K (325 O F) for 152; minutes and for polyimide resin use 450 K (350" F) for 202; minutes), 

cool to room temperature, and weigh to the nearest 0.1 mg (W3). The mean value of the three 

volatile determinations calculated as follows was reported: 

x 100 

w1-w3 Volatile Content, weight percent = 

Wl -w2 

where: 

Wl = Weight of crucible plus specimen before volatile removal 

W2 = Weight of crucible 

W3 = Weight of crucible plus specimen after volatile removal 

8.2 PREPREG RESIN SOLIDS AND FIBER CONTENT 

Cut three specimens of the material, each approximately 0.051 by 0.051 m (2.0 by 2.0 in.), from 

an area of the sample away from the edges. Weigh a specimen to the nearest 0.1 mg (Wl), place 

in a 300 ml tall-form beaker, and add 25 ml of dimethyl formamide (DMF) technical grade. 

Place the beaker on a hot plate and heat to boiling for 10 to 15 minutes. Remove the beaker 

from the hot plate and allow to cool to room temperature. Pour off the DMF, being careful not 

to lose any fibers. Rinse the fiber residue two or three times in acetone. After removing most of 

the acetone, place the fibers in a tared aluminum cup (throwaway type) and dry for 30 minutes 

minimum in an oven at 436 K (325" F). Determine the weight of fibers (W,) to the nearest 0.1 

mg and record. The mean value of three fiber and resin solids content determinations 

calculated as follows was reported: 

Fiber Content, weight percent = - x 100 w2 

Wl 

Resin plus volatiles, weight percent = 100 - - x 100 

w2 

Wl 

Resin Solids Content, weight percent = 100 - - X 100 - Volatile Content w2 

Wl 

B- 1 
where: 

Wl = weight of specimen 

W2 = weight of fibers 

Volatile Content = See Section B.l 

8.3 PREPREG RESIN FLOW 

Cut six pieces of style 181 glass fabric, 0.10 by 0.10 m (4.0 by 4.0 h) square, for bleeder and 

two squares of Teflon coated glass fabric of the same size, for separator. Weigh to the nearest 

0.1 mg (W,) on an analytical balance. Cut two pieces of the material to be tested 0.051 by 0.051 m 

(2.0 by 2.0 in.) and weigh with the separator and bleeder to the nearest 0.1 mg (W2). Cross ply 

the two specimens and sandwich them between the separator and the bleeder with the bleeder 

on the outside. Place in a preheated press and cure for 152; minutes under 0.69 MNlm2 (100 

psig) at 436 K (325 O F) for the epoxy system and at 450 K (350" F) for the polyimide system. 

Remove the crossply test specimen from the separator and bleeder. Weigh the separator plus 

bleeder to the nearest 0.1 mg (W3). The mean value of three flow determinations calculated as 

follows was reported: 

Percent Flow = A x 100 

w2-Wl 

where: 

W, = Weight of glass fabric plus Teflon coated glass fabric 

W, = Weight of glass fabric, Teflon coated glass fabric, and specimens 

Ws = Weight of glass fabric plus Teflon coated glass fabric after cure 

8.4 LAMINATE SPECIFIC GRAVITY 

The method used for determining the specific gravity of the laminates was by displacement of 

water. The test procedure was in accordance with Federal Test Method Standard No. 406, 

Method 5011. Cut three pieces of the laminate to be tested 0.013 by 0.013 m (0.5 by 0.5 in.). 

Using an analytical balance and a specific gravity bridge, weigh the specimen in air (W1) and in 

distilled water (W2) at 296 2 1.1 K (73.5 2 2" F). The mean value of three specific gravity 

determinations calculated as follows was reported: 

Specific Gravity = A 

Wl -w2 

where. 

W1 = weight of specimen in air 

W2 = weight of specimen in water 

8 
B.5 LAMINATE RESIN AND FIBER CONTENT 

The test specimen was approximately 0.013 by ~013 m (0.5 by 0.5 in.) by laminate thickness. 

The cured resin and fiber content was determined by acidlperoxide digestion as follows: 

Weigh the test specimen to the nearest 0.1 mg (Wl), place in a 300 ml tall-form beaker, 

and add 20 ml of concentrated sulfuric acid. Place the beaker on a hot plate and heat the 

acid until vigorous fuming occurs. 

When the composite is visibly disintegrated and resin particles and fibers are dispersed 

throughout the sulfuric solution, carefully add the hydrogen peroxide (50% strength) 

dropwise down the side of the beaker. Rubber gloves and a fume hood with appropriate 

safety glass shield shall be used throughout the addition, and precautions shall be 

taken as recommended by the applicable safety regulations and procedures for 

handling hydrogen peroxide. 

The reaction is considered complete when the hot sulfuric acid solution below the fibers 

becomes clear and colorless. At this point add two more ml of hydrogen peroxide to the 

solution, and heat the solution to fumes for another 10 minutes to ensure complete 

decomposition of the polymer. Remove the beaker from the hot plate, and allow to cool 

to 294 to 300 K ('70" to 80" F) and then place in an ice bath. 

Collect fibers by vacuum filtration through a medium-porosity, sintered-glass crucible 

that has been weighed to nearest 0.1 mg (W2). After the sulfuric acid has been filtered 

off, wash the fibers in the crucible thoroughly with 600 ml of distilled water, adding a 

few milliliters at a time. Verify removal of sulfuric acid traces by checking pH of the 

filtrate drops. 

Remove the crucible from the filtering system and place in an open beaker in an oven at 

422 K (300 O F) for 45 minutes. After drying, cool the crucible in a desiccator and weigh 

to the nearest 0.1 mg (W3). 

The resin and fiber content were calculated according to the following equations: 

W1 -(W3-W2) 

Wl 

Resin Content, weight percent = x 100 

x 100 

w3-w2 

Fiber Content, weight percent = 

Wl 

where: 

W1 = Weight of test specimen 

W2 = Weight of crucible 

W3 = Weight of crucible plus fibers after acidlperoxide digestion 

8-3 
8.6 B/Al LAMINATE FIBER VOLUME 

The test specimen was approximately 0.013 by 0.013 m (0.5 by 0.5 h) by laminate thickness. 

The fiber volume was determined as follows: 

1. Wash, dry, and weigh the test specimen to the nearest 0.1 mg (Wl). 

2. Leach the fibers from the samples in a 50% solution of sodium hydroxide at approximately 339 K (150' F). 

3. Wash, dry, and weigh the collected fibers to the nearest 0.1 mg (W2). 

Assuming a density of 2700 kg/m3 for the aluminum and a density of 2510 kg/m3 for the 5.6 mil 

boron, the volume percentages of fibers were calculated according to the following equation: 

QF Fiber Volume, volume percent = x 100 

WrW2 + w2 

where: 

W1 = Weight of test specimen 

W2 = Weight of fibers 

QF = Density of fibers 

eA = Density of aluminum 

8.7 LONGITUDINAL FLEXURAL STRENGTH 

The test method used for determining logitudinal flexural strength was generally in accordance with ANWASTM D 790-71 and Federal Test Method Standard No. 406, Method 1031. 

Specimens were cut with the fibers aligned parallel to the longitudinal axis. The specimen configuration with the loading geometry is shown in Figure B-1. Testing was done with three 

point loading using a span to thickness ratio of 32 to 1. The specimens, with the smooth side 

up, were loaded to failure at 0.0025 m (0.10 in) per minute crosshead speed in standard test 

machines (Instmn, Tinius-Olsen, or Baldwin). The longitudinal flexural strengths were 

calculated according to the following formula: 

3PL 

2 bd2 

FL =- 

where: 

FL = Ultimate longitudinal flexural strength 

P = Maximum load carried by the specimen 

8-4 
I 4 

t 

d 

Specimen Dimensions 

Length (1) = 0.076 m (3.0 in.) 

Width (b) = 0.01270 m (0.500 in.) 

Thickness (d) = Approximately 0.0015 m (0.060 in.) 

Spad'l'hickness Ratio (-) = 32 to 1 

Loading head and reaction supports. are 0.00635 m (0.250 in.) diameter steel rod 

L 

d 

Overhang must be the same over each end 

65321 7-31 3 

Figure B-1. Longitudinal Flexural Specimen 

6 

L = support span 

b = Specimen width 

d = Specimen thickness 

8.8 SHORT BEAM SHEAR STRENGTH 

The test method used for determining short beam shear strength was generally in accordance 

with ANSIlASTM D2344-76. Specimens were cut with the fibers aligned parallel to the 

longitudinal axis. The specimen configuration with the loading geometry is shown in Figure 

B-2. Testhg was done with three point loading using a span to thickness ratio of 4 to 1. The 

specimens, with the smooth side up, were loaded to failure at 0.0025 m (0.10 in.) per minute 

crosshead speed in standard test machines (Instron, Tinius-Olsen, or Baldwin). The short beam 

shear strengths were calculated according to the following formula: 

r= - 3P 

4bd 

B-5 
t-'* 

"T d 

Specimen Dimensions 

Length (1) = 0.0152 m (0.60 in.) 

Width (b) = 0.00635 m (0.250 in.) 

Thickness (d) = Approximately 0.0015 m (0.060 in.) 

SpadThickness Ratio (-) = 4 to 1 

Loading head and reaction supports are 0.00318 m (0.125 in.) diameter steel rod 

L 

d 

. 

Overhang must be the same over each end 

6532 1 7-3 i 4 

Figure B-2. Short Beam Shear Specimen 

where: 

7 = Ultimate short beam shear strength 

P = MaKimllm load carried by the specimen 

b = Specimen width 

d = Specimen thickness 
1. Report No. 

NASA CR-178272 

7. AuthorM 

J. R Kerr and J. F. Haskins 

9. Performing Organization Name and Address 

2. Government Accession No. 3. Recipient's Catalog No. 

8. Performing Organization Report No 

GDCC RAD-85405 

10. Work Unit No. 

4. Title and Subtitte I 5. Report Date 

12. Sponsoring Agency Name and Address 

National Aeronautics and Space Administration 

Washington, DC 20546 

Tie-TemperatureStress capabilities of Composite Materials 

for Advanced Supersonic Technology Application 

Contractor Report 

14. SRomoring Agency Code 

505-63-0 1-06 

- 

17. Key Words (Suggested by Author(s)) 

Composite Materials, Thermal Aging, Environmental 

Aging, Fatigue. Creep, Tensile, Shear. Fracture, Flight 

Simulation Exposure, Boron, Graphite, Epoxy Resins, 

Polyimide , Resins, Aluminum, Supersonic Vehicles, 

SC R 

Convair Divisioti of General Dynamics Corporation 

5001 Kearny Villa Road 

San Diezo. California 92123 

18. Distribution Statement 

Unclassified - Unliiiiited 

Subject Category 24 

11. Contract or Grant No. 

NAS 1-1 2308 

13. Type of Report and Period Coverec 

19. Security Classif. (of this report) 

Unclassified 

20. Security Classif. (of this page) 21. No. of Pages 22. Price 

Unclassified 6 72 

I 

15. Supplementary Notes 

Langley Contract Monitors: B. A. Stein and D. R. Ruiiiinler 

16. Abstract 

Advanced composites will play a key role in the development of the technology required for the design and 

fabrication of future supersonic vehicles. However, implementation of the material into vehicle usage is contingent upon accelerating the demonstration of service capacity and design teclinology. Because of the added 

material complexity and lack of extensive service data. labontory replication of ttic tlight service will provide 

the most rapid method of documenting the airworthiness of advanced composite systems. Consequently. a 

laboratory program was conducted to determine the time-temperature-stress capabilities of several liipli 

temperature composite materials. Tests included in this study were thermal aging. environmental aging. fatigue. 

creep, fracture, tensile. and real-time flight simulation exposure. The prognm had two phases. The first included 

all the material ptopeny determinations and aging and sirnulation exposures up through 10.000 hours. Tlie 

second continued these tests up to 50,000 cumulative hours. "his report presents the results of the Phase 1 

baseline 3nd 10.000 hour aging and flight simulation studies. the Phase I1 50,000 hour aging studies. and the 

Phase 11 flight simulation tests. some of which extended to almost -10.000 hours. 

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