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(A Q NBS SPECIAL PUBLICATION 563

U.S. DEPARTMENT OF COMMERCE / National Bureau of Standards

Advanced Composites:

Design and Applications

QC

100

,U57

no i 563

1979

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MFPG

29th Meeting

• si si mbQpft

N )V t 3 1979

MFPG

Advanced Composites: Design and Applications

Proceedings of the 29th Meeting of the

Mechanical Failures Prevention Group,

held at the National Bureau of Standards,

Gaithersburg, Maryland, May 23-25, 1979.

Edited by

T. Robert Shives and William A. Willard

Center for Materials Science

National Measurement Laboratory

National Bureau of Standards

Washington, DC 20234

The 29th meeting of the MFPG and these proceedings were sponsored by

the Center for Materials Science of the National Bureau of Standards,

Washington, DC 20234; the Office of Naval Research, Department of the Navy, Arlington, VA 22217; the Naval Air Systems Command, A. J. Koury,

Washington, DC 20360; the National Aeronautics and Space Administration, Goddard Space Flight Center, Greenbelt, MD 20771; and the Department of Energy - Fossil Energy, Washington, DC 20545.

U.S. DEPARTMENT OF COMMERCE, Juanita M. Kreps, Secretary

Luther H. Hodges, Jr., Under Secretary

Jordan J. Baruch, Assistant Secretary for Science and Technology

NATIONAL BUREAU OF STANDARDS, Ernest Ambler, Director

Issued October 1979
Library of Congress Catalog Card Number: 79-600151

National Bureau of Standards Special Publication 563

Nat. Bur. Stand. (U.S.), Spec. Publ. 563 , 304 pages (Oct. 1979)

CODEN: XNBSAV

U.S. GOVERNMENT PRINTING OFFICE

WASHINGTON: 1979

For sale by the Superintendent of Documents, U.S. Government Printing Office, Washington, D.C. 20402

Stock No. 003-003-02120-2 Price $7.00

(Add 25 percent additional for other than U.S. mailing)
FOREWORD

The 29th meeting of the Mechanical Failures Prevention Group was held

May 23-25, 1979, at the National Bureau of Standards in Gaithersburg,

Maryland. The program was organized by the Design Committee under the

chairmanship of Jesse E. Stern of the Goddard Space Flight Center. The

committee, the session chairmen, and especially the speakers are to be

commended for the excellent program. Special appreciation is extended

to Harriet Gordon of the Goddard Space Flight Center for her expert

coordination and preparation of the many items of correspondence that

were generated in order to develop the technical program into its final

form.

The symposium opened with a welcome to the National Bureau of Standards

by Dr. Elio Passaglia, Deputy Director of the Center for Materials

Science. Dr. Passaglia's comments centered on the continuing need for

standards and accuracy of measurements especially in the design and

application of new materials such as those to be discussed at the

symposium.

Appreciation is extended to the following members of the NBS Fracture

and Deformation Division: T. Robert Shives and William A. Willard for

their editing, organization and preparation of the proceedings, and

Joel C. Sauter for photographic work. Most of the papers in the

proceedings are presented as submitted by the authors on camera ready

copy. Some moderate editorial changes were required.

Gratitude is expressed to the following members of the NBS Center for

Materials Science: Marian L. Slusser for handling financial matters,

Delores J. Covey for typing, and Larry W. Ketron for drafting work.

Special thanks is extended to Jo Ann Lorden of the NBS Public Information

Division for the meeting, hotel and banquet arrangements.

HARRY C. BURNETT

Executive Secretary, MFPG

Center for Materials Science

National Bureau of Standards

III
TABLE OF CONTENTS

Page

III

FOREWARD

SESSION 1: AEROSPACE AND AIRCRAFT APPLICATIONS AND DESIGN

A. Characterization and Application of Advanced 3

Composite Materials (abstract only) . Max Waddoups

B. Physical and Mechanical Response of Graphite/ 4

Polyimide Materials to Long Term Exposure in a

Space Environment; Material Preparation and

Control Testing. J. H. Powell

C. Effects of Room-Temperature Aging on Composite 17

Prepregs. H. C. Nash, C. F. Poranski, Jr.*, and

R. Y. Ting

SESSION 2: AEROSPACE AND AIRCRAFT APPLICATIONS AND DESIGN

A. Application of Instrumental Techniques to the Study 25

of the Cure of Phthalocyanine/Graphite Composites.

Robert Y. Ting

B. Moisture Diffusion Analysis for Composite Micro- 32

damage. C. L. Leung* and D. H. Kaelble

C. Recent Work on Techniques and Applications of 47

Moisture Barriers to Graphite Epoxy Composites.

James F. Haskins

SESSION 3: AEROSPACE AND AIRCRAFT APPLICATIONS AND DESIGN

A. In-Service Ultrasonic Inspection System for 59

Composites (abstract only). F. H. Chang*,

J. R. Bell, A. H. Gardner, C. P. Fisher and

G. P. Handley

B. Design Considerations for Graphite -Epoxy Laminates 60

of Low Thermal Expansivity. Dennis D. Smith

IV
C. Aircraft Composite Materials Selection and Appli- 72

cation. V. L. Reneau

D. Composite Structures in Rotors and Propellers. 80

David F. Thompson

SESSION 4: AEROSPACE AND AIRCRAFT APPLICATIONS AND DESIGN

A. Development of a Library Module for the Analysis of 91

Advanced Composite Materials. Karyn T. Knoll

B. Structural SMC - Material, Process and Performance 108

Review. Ralph B. Jutte (presented by Joseph D.

Feko, Owens-Corning Fiberglas Corporation)

C. Development of a Graphite-Epoxy Spacecraft Precision 130

Mounting Platform (abstract only). Kathleen S.

Budlong and Alvin W. Sheffler

SESSION 5: AUTOMOTIVE APPLICATIONS AND DESIGN

A. Design Analysis of Automotive Composite Structures 133

(abstract only). M. F. Kowalski, D. S. Fine*, and

R. K. Herrman

B. Fiberglass Reinforced Plastics Use in Transporta- 134

tion Vehicles (abstract only) . R. D. Pistole

C. Designing Automotive Components with Continuous 135

Fiber Composites. H. T. Kulkarni* and P. Beardmore

SESSION 6: INDUSTRIAL APPLICATIONS AND DESIGN

A. Fabrication of Large Composite Spars and Blades. 155

O. Weingart (presented by Robert Gordon, Structural

Composite Industries, Inc.)

B. The Effect of Fiber Orientation on the Design 173

Parameters of Reinforced and Filled Thermoplastic

Compounds (abstract only) . Peter J. Cloud* and

Robert T. Alvord

V
C. Quality Assurance of Manufactured Components 174

(abstract only). R. D. Margolis

SESSION 7: INDUSTRIAL APPLICATIONS AND DESIGN

A. Degradation Data of Kevlar Pressure Vessels. 177

W. D. Humphrey

B. Design Assurance of a Leak Failure Mode for 198

Composite Overwrapped Metal Tankage. W. W. Schmidt

C. Composite Materials in Recreational Equipment 208

(abstract only) . David Ratchford

SESSION 8: INDUSTRIAL APPLICATIONS AND DESIGN

A. Holographic Nondestructive Evaluation of Spherical 211

Kevlar/Epoxy Pressure Vessels. D. M. Boyd* and

B. W. Maxfield

B. The Air Force "Manufacturing Cost/Design Guide 226

(MC/DG) " (abstract only). Bryan R. Noton* and Captain

Dan L. Shunk

C. Tensile Strength and Failure Modes of Boron-Epoxy 228

Composites with a Notch. Charles E. S. Ueng

SESSION 9: FAILURE MODES IN ADVANCED COMPOSITES

A. Failure Criteria for Composites under Complex 241

Loading (abstract only). P. W. Mast*, L. A. Beaubien,

D. R. Mulville, S. A. Sutton, R. W. Thomas, J. Tirosh

and I. Wo lock

B. Failure Analvsis of an Idealized Composite Damage 242

Zone. Carl T. Herakovich* and David A. O'Brien

C. Interlaminar Failure in Epoxy Based Composite 255

Laminates. A. S. Wang*, F. W. Crossman and

G. E. Law, Jr.

VI
SESSION 10: MARINE APPLICATIONS AND DESIGN

A. Current Developments in Advanced Composites for 267

Naval Ship Application. J. J. Kelly* and

H. H. Vanderveldt

B. Environmental Effects on Fiber Reinforced Organic 276

Matrix Composites. M. Silvergleit* , A. B. Macander,

F. A. Brauer, and H. P. Edelstein

C. Composite Technology for Marine Application. 286

W. P. Couch

APPENDIX

List of Registrants for the 29th MFPG Meeting. 301

MFPG Publications 307

* Indicates speaker when a paper had more than one author.

VII
ABSTRACT

These proceedings consist of a group of thirty one submitted

entries (twenty one papers and ten abstracts) from the 29th meeting of

the Mechanical Failures Prevention Group which was held at the National

Bureau of Standards in Gaithersburg , Maryland, May 23-25, 1979. The

subject of the symposium was the design and application of advanced

composites. Special emphasis was directed to aerospace, aircraft, automotive, marine, and industrial applications and design. Failure modes

in advanced composites were also discussed.

Key words: Advanced composites; applications of composites; composites;

composite structures; design using composites; fiber composites; fiber

glass; filament wound composites.

UNITS AND SYMBOLS

Customary United States units and symbols appear in some of the

papers in these proceedings. The participants in the 29th meeting of

the Mechanical Failures Prevention Group have used the established units

and symbols commonly employed in their professional fields. However, as

an aid to the reader in increasing familiarity with the usage of the

metric system of units (SI), the following references are given:

NBS Special Publication, SP330, 1977 Edition, "The International System

of Units."

ISO International Standard 1000 (1973 Edition) , "SI Units and Recommendations for Use of Their Multiples."

E380-76 ASTM/IEEE Standard Metric Practice (Institute of Electrical and

Electronics Engineers, Inc., Standard 268-1976).

Disclaimer

Certain trade names and company products are identified in order to

adequately specify the experimental procedure. In no case does such

identification imply recommendation or endorsement by the National

Bureau of Standards, nor does it imply that the products are necessarily

the best available for the purpose. Views expressed by the various

authors are their own and do not necessarily represent those of the

National Bureau of Standards.

/

VIII
SESSION" 1

Clia,i3?mgLxi: «J. HS. Stern

Goddard Space Flight Center
I

I

j

I
Characterization and Application of Advanced Composite Materials

Max Waddoups

General Dynamics, Fort Worth Division

Fort Worth, Texas 76101

Abstract

A technical understanding of the structural characteristics of composite

materials and the critical issues in aircraft structural component

development have been simultaneously generated. It is established that

the strengthening concept, structural characteristics, and fabrication

techniques are distinct from metals. These differences and the resultant

structural implications are the subject of this lecture.

For practical purposes, composites are linear elastic. Tensile failure

modes are brittle fracture governed, and compressive failure modes are

micro and macro stability governed. Even though composites have no

classical ductility, they are surprisingly tough materials. In most

full-scale structures currently developed, tensile fracture constitutes

the critical strength failure mode.

While generally tensile fracture governs strength, an equivalent-tocrack coalescence and monotonic growth in tension are not observed in

composites (except glass-fiber-reinforced plastics). In fact, tensiondominated cracking is not of serious consequence in advanced laminates.

Weak plane (intralaminar and interlaminar) flaw propagation, predominantly

in compressive and shear loadings, yields the lift limiting failure modes

for composites. When modeled as strain-energy release-rate-driven flawgrowth processes, different growth rates and initial strengths generate

competing failure modes in full-scale structures.

Simple structural forms that control the life-limiting failure processes

will become the rule for successful structures. Only on stiffness

critical or fairing type of structure will the complex shape moldability

advantages of composites be used.

Because of the scale effects in the fracture process zone and the ability

to react significant out-of-plane loads, simple mechanical fastening will

dominate highly loaded, strength-critical structure. Secondary adhesive

bonding will not be very successful in highly loaded fracture critical

structure.

3
i

PHYSICAL AND MECHANICAL RESPONSE OF

GRAPH I TE/ POLY I MI DE MATERIALS TO LONG TERM EXPOSURE

IN A SPACE ENVIRONMENT; MATERIAL PREPARATION AND

CONTROL TESTING

J. H. Powell

Rockwell International

Tulsa, Oklahoma 74151

Abstract: Graphite/polyimide (Gr/PI) composite laminates are currently

being prepared for deployment on NASA Langley Research Center's Long

Duration Exposure Facility (LDEF). The experiment is intended to provide scientific and engineering design information on the response of

Gr/PI composites to an actual exoatmospheric environment. Material

systems to be exposed are a bismaleimide PI (Hexcel F- 178) on Thome!

300 graphite fibers and a condensation PI (Monsanto Skybond 703) on

HT-S graphite fibers, including both "precured" and "cocured" laminates.

Initial physical characteristics determined on both exposure and simultaneously processed control panels include flatness, microcracking

,

void content, dry weight, and general appearance. Mechanical tests

performed to date (on control panels only) are the basis for strength

and elastic properties in tension, compression, and (rail) shear at

ambient and elevated temperatures. Following in-space exposure of test

panels and return to earth, identical physical and mechanical determinations will be made. Data generated before and after exposure will be

compared to identify any changes in physical or mechanical characteristics, particularly any unexpected changes. The ultimate output of

the experiment will be either confirmation of predicted material behavior or the basis for revised prediction methods.

Key Words: Exoatmospheric exposure of composites; polyimide/graphi te

mechanical properties; polyimide composite stability in space; Long

Duration Exposure Facility (LDEF); ultrasonic inspection; cocured and

precured laminates.

The essentially universal organic matrix material currently used in

aerospace advanced composite structures is that based upon epoxide

chemical polymerization mechanisms -- for several excellent reasons.

As a class, epoxy resin systems are well characterized, allowables are

established for many of them, processing procedures are essentially

trouble free, well developed, reproducible and compatible with existing

industry processing equipment. The Space Shuttle Vehicle employs

epoxy/graphi te composite in two major structural components -- the

payload bay doors and orbital maneuvering system ( 01 IS ) pods, both

k
applications resulting in substantial weight savings even though requiring external insulation. Epoxies are, however, approaching their

upper temperature capability for efficient operation; in addition to

which their capability is further limited by absorbed moisture at elevated temperatures.

Appreciable gains in high temperature performance must therefore be

sought in the higher temperature resistant systems, the more promising

current materials belonging to the polyimide family. Increasing temperature capability by 100 to 250F over the present 250 to 350F maximum

operating range of epoxies potentially saves critical weight through

higher operating allowables or reduced insulation requirements. Under

discretionary company funding, Rockwell International's Tulsa Division

had previously investigated two interim, conventionally processible

polyimide systems, one suitable for 450F use, the other for 550 to 600F

use. When the Long Duration Exposure Facility (LDEF) project solicited

proposals for experiments, Rockwel 1 -Tul sa responded and was awarded a contract 1 to test these two materials.

Objective: The experiment is intended to accumulate design and scientific data associated with long term exposure in a space environment,

including any evidence of changes in mechanical properties, weight loss,

dimensional instability, microcrack growth, delamination , or other degradation due to hitherto "unknowns". An addition polyimide (API) and

a conventional condensation polyimide (CPI) will be evaluated both in

the "precured" and "cocured" laminate form. By comparing post-exposure

with control data, previous assumptions and prediction tools will be

verified or corrected as necessary.

Experiment description: Essential elements of the experiment are described sequentially below under the enumerated headings: (1) test

vehicle, (2) general experiment flow, (3) materials, (4) panel design,

(5) processing, (6) non-destructive examination , (7) specimens preparation and testing, (8) test evaluation and analysis, and (9) loading and

shipment of experiment.

(1) The test vehicle is the LDEF structure (Figure 1) which is currently estimated to be launched between late 1980 and late 1981 during

one of the regular Space Shuttle missions. Test panels are mounted in

one of two trays, the arrangement of that for 3 panels shown in Figure

2. The LDEF is released from the orbiting Space Shuttle at an extremely low acceleration and put into its own independent orbit at an average distance of 235 nautical miles above the earth's surface. The

orbital period will be about 92 minutes, or 15.6 revolutions per 24

hours. In stabilized flight, the LDEF will maintain a fixed attitude,

the projection of its longitudinal axis passing through the center of

the earth. Initial oscillatory motions experienced during release will

2 NASA Langley Research Center Contract NASI- 15309, March 1978.

5
\

DIRECTION

OF ORBIT

SECOND

TRAY

TOWARD

EARTH

LOCATION OF TRAY

BEARING PART OF

PI EXPERIMENT.

OTHER PANELS

DIAG OPP POSITION

Figure 1. Long Duration Exposure Facility (LDEF);

Dimensions 30 Feet Long x 13.5 Feet In

Diameter

Figure 2. LDEF Tray Conceptual Sketch Showing

3-Panel Arrangement

6
dampen out within a few days. After 6 to 12 months in orbit, the LDEF

will be retrieved during a subsequent Shuttle mission and returned to

earth. During this period, the test panels will be exposed to alternating low and high temperatures, calculated to range between -109F and

+236F, and various types and levels of radiation endemic to the space

environment. Ambient pressure is assumed to be on the order of 1 x 10" 10 torr. While the Gr/PI experiments are completely passive, the

LDEF vehicle will be equipped to acquire and record certain information

,

e.g. temperatures and impinging radiation at various locations on the

structure. The position of each experiment will also be ascertained

after the fact. Through use of applicable analysis techniques, the

exposure environment for each experiment can be reconstructed.

(2) General experiment flow is shown in Figure 3. Two matched sets of

panels are subjected to an identical series of operations, excepting of

course that one set is exposed on the LDEF. Both "precured" and "cocured" laminates are prepared for each material type (condensation and

addition). All panels are nondestructi vely inspected, employing ultrasonic through transmission, microscopy, and physical measurement techniques. Finally, panels are sectioned and mechanically tested. Data

from the two sets of panels are analyzed and compared, and the effects

of exposure to the space environment assessed.

SET NO. 1 (CONTROL)

LAMINATES

NO. 1 (PRECURED) CONDENSATION

NO. 2 (COCURED) TYPE

NO. 3 (PRECURED) ADDITION

NO. 4 (COCURED) TYPE

NDT

INSPECTION

TEST SPECIMEN FABRICATION

EVALUATE

RESULTS

ESTABLISH

CRITERIA SET NO. 2 (EXPOSURE)

LAMINATES

NO. S (PRECURED) CONDENSATION NDT

NO. 6 (COCURED) TYPE INSPECTION

NO. 7 (PRECURED) ADDITION

NO. 8 (COCURED) TYPE

SHIP TO

TEST AGENCY

EXPOSE PANELS TEST SPECIMEN FABRICATION

NDT

INSPECTION

Figure 3. General Experiment Flow

(3) Material systems used in the experiment are representative of two

polyimides, the standard vacuum bag/autoclave processible condensation

Pi's and the more recent intermediate temperature vacuum bag/autoclave

processible addition Pi's. Neither system requires pressure in excess

of 100 psi nor curing temperature in excess of 350F. The condensation

system (CPI) selected is Ferro Corporation's CPI - 2249 (Monsanto Skybond 703)/HT-S collimated tape, molding to a ply thickness of .0055

inch at 63% fiber volume. This system was selected on the basis of

past successes in IR&D projects where both tape and fabric systems

7
I

were conventionally autoclave processed, yielding good mechanical properties in tension and compression at room temperature and 600F. The

system is considered an interim material, ultimately to be replaced by

advanced systems, but still having many potential applications at an

economical cost. The addition system (API) is Hexcel 1 s F- 178 bismaleimide on Thornel 300 fibers, supplied in tape form and molding to a ply

thickness of .005 inch at 65% fiber volume. This system processes much

like epoxies, exhibits superior moisture resistance and offers about

100F more temperature resistance. It and similar emerging API systems

offer excellent potential for application in the 400 to 450F temperature regime at an economical processing cost.

(4) Panels were designed to represent a commonly applied fiber orientation, i.e. quasi-isotropic (0, +45, 90)

n construction. The specific panel geometry, i.e. planform dimensions and thickness were analyzed to verify ability to withstand environmental requirements of the

governing NASA document. 1 Table I summarizes the key elements of the

dynamic environment that were considered in designing panels. The

critical design condition was a natural frequency of 50 Hertz, minimum,

for each panel when simply supported along its four edges. Overall

panel sizes were determined by LDEF tray geometry and the number of

specimens needed to provide 5 replicates of each specimen type and test

condition (Figure 4). Thickness was determined by minimum panel stiffness requirements.

Panels are mounted in trays on specially designed support structures

which are unique to these particular experiments. Figure 5 conceptually illustrates tray support structure and test panel. Both tray and

support are constructed from aluminum alloy, thus have high coefficients of thermal expansion. Graphite laminates, on the other hand,

have very low thermal expansion. Panels are therefore attached by

clamping, allowing surrounding structure to expand and contract freely

with temperature change without loading or buckling test panels. A

thermal isolation septum is placed between the experiments and interior

of the LDEF to minimize interaction between the internal and external

envi ronments

.

(5) Test panels are processed in a conventional autoclave at 85-100

psig and 350F. Both polyimide types are prebled prior to final cure,

using part of the calculated bleeder material during prebleeding and

the remainder during cure. The API is processed much like epoxy, being

essentially a volatile-free prepreg. The CPI, on the other hand, requires a relatively long, 200F dwell under full vacuum to release a large part of the N-methyl pyrollidone (NMP) solvent. Autoclave pressure is applied at the beginning of heat up, and curing completed at

350F. The API is postcured without restraint in an oven at 450F for

16-20 hours, the CPI for 96 hours at temperatures progressing from

J Long Duration Exposure Facility (LDEF) Experimenter Users Handbook,

LDEF No. 840-2, LDEF Project Office, Langley Research Center, Hampton,

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1-1/3 TRAYS REQD

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igure 4. Test Laminates for Graphi te/Polyimide Experiments

Showing Layouts of Test Specimens

Figure 5. Support Structure Conceptually Shown

Including Interface With LDEF Tray

10
350F to 600F. "Precured" panels are molded directly against a smooth

platen, while "cocured" panels are molded over honeycomb core with a non-porous fluorocarbon film between core and laminate. The pattern

of the honeycomb is thus imparted to one surface of the cocured laminate.

(6) Non-destructive examination is performed using ultrasonic through

transmission, visual examination, photomicroscopy , and physical measurement. Ultrasonic measurements are made with the panel fully immersed in a water tank, and panel quality is assessed by transmitting

an ultrasonic signal through the panel normal to its surface. Attenuation of the input signal indicates voids, discontinuities, or other

anomalies and is evidenced by loss of received signal or blank regions

in the closely spaced pen lines drawn by the recorder. Input signal

can be adjusted to accommodate different panel thicknesses or uniformly

dispersed voids by varying the signal frequency and gain. The inherent

void containing CPI system can thus be measured to detect blistered and

delaminated areas. Panels will be examined ul trasonical ly both before

and after space exposure. Laminate surfaces and edges are examined

with and without the aid of a low-powered lens for general uniformity

in appearance, and for evidence of resin starving, roughness, fiber

splintering, cracks, and other anomalies. Specimens representing

typical regions of laminates will be potted in resin with edges normal

to the surface, ground and polished, then examined under high magnification. Photo-micrographs of sound, microcracked , and porous areas

will document the typical condition. Physical measurements determine

cupping, bowing, and twisting. With three corners of a panel touching

a surface table (i.e. in one plane), twist is easily determined by

departure of the fourth corner from plane. With the panel placed concave upward on a surface, cupping and bowing are directly measurable

with a straight edge.

(7) Specimens are cut from test panels according to the layout schemes

shown in Figure 4. Test specimens have the geometries shown in Figure

6. In all panels, the zero degree (0) reference direction is defined

by the fiber direction of surface plies. Hence, tension specimens are

laid out with their long axes in the (0) direction, while rail shears

and compression specimens have their long axes parallel to the (90)

direction. Layout directions were dictated by specimen and panel

dimensions, thus options for specimen orientations were limited. The

grip regions of high temperature tensile specimens are tabbed with

aluminum to produce the desired type of failure. A "gripper" strip

about one and one-half inch wide is bonded to each surface of the long

edges of rail shear specimens. This has been found necessary to prevent slippage of specimens in the test fixture. Slippage promotes

premature bearing failures in specimens, thus not demonstrating full

shear strength.

11
(ASTM 0695 SPECIMEN) INTRALAMINAR SHEAR

(RAIL SHEAR SPECIMEN)

Figure 6. Mechanical Test Specimen Configurations

Table II presents the test matrix for control panels. The elevated

temperature test condition for API (addition) specimens is 450F, for

CPI (condensation) specimens, 600F. API control panels, laid out

for specimen cutting are shown undergoing warp and twist measurement

in Figure 7. Subsequent to space flight, exposure panels will be

sectioned and tested in an identical manner.

Testing is performed in a 20,000 pound capacity Instron Testing Instrument equipped with automatic load vs. strain plotting capability and a high temperature Conrad-Missimers test chamber.

(8) Upon completion of control testing for each panel type, test data

will be analyzed, compared with predictions made during the experiment

design task, and compared with requirements. Upon verification of compliance with requirements, exposure panels will be ready for installation in LDEF trays.

(9) Exposure test panels will be trimmed to final dimensions on assembly in LDEF trays. The support structures will have been assembled in

situ in trays prior to installing test panels. Edge retainers which

clamp test laminates in place are made from aluminum shim stock, which

can be stripped to accommodate variable panel thicknesses. Trays were

received from NASA installed, shock mounted in reusable shipping containers, and will be returned in like manner. Figure 8 shows one of

the two trays provided and support structure, which will contain the

cocured CPI panels. This tray also contains graphi te/epoxy test panels

made under the "sister" experiment to the polyimide experiment, in

addition to the two smaller cocured CPI panels.

Test results to date: The control testing phase of the experiment was

incomplete at the time of this paper's issue; thus, only part of the

control test data to be generated is reported herein. Exposure tests

are not expected to be completed before the 1981-1982 time period.

12
Table II. Test Matrix for Control and Exposure Specimens

LAMINATE NUMBER

CURE PROCESS

PRECURED COCURED SIZE

NO. REQD* TEST

TEST

TEMP

RT

MATERIAL TYPE

ET COND ADDITION INSTRUMENTATION

60-001 60-005 60-009

60-003

8.50 X 1.0 8.50 X 1.0 8.50 X 1.0 8.50 X 1.0

10

10

10

10

TENSION TENSION

TENSION

TENSION

EXTENSIOMETER

EXTENSIOMETER EXTENSIOMETER

EXTENSIOMETER

60-001 60-005

60-009

60-007

0.5 X 3.13

0.5 X 3.13 0.5 X 3.13

0.5 X 3.13

1010

10

10

COMPRESSION

COMPRESSION

COMPRESSION COMPRESSION

COMPRESSIOMETER COMPRESSIOMETER

COMPRESSIOMETER

COMPRESSIOMETER

60-001 60-005

60-009 60-003 \ 60-007 f

4.0 X 8.0

4.0 X 8.0 4.0 X 8.0

4.0 X 8.0

10

10

10

10

**** SHEAR

SHEAR

SHEAR

SHEAR

BI-AXIAL S/G

BI-AXIAL S/G

BI-AXIAL S/G

BI-AXIAL S/G

NOTES: MO SPECIMENS PER LAMINATE - 5 PER TEST TEMPERATURE

**ITRI SPECIMEN

***ASTMD695

****RAIL SHEAR

Figure 7. Measurement of Warp and Twist

in API Control Panels

13
Figure 8. LDEF Tray and Support Structure

Thus, final evaluations will probably not be available until 1982. Control test data generated thus far are reported in Tables III and IV, and

Figure 9.

While the major significance of the experiment will not be known until

exposure panels are returned from space, some general impressions result

ing from API panel fabrication and testing may be cited. First, the bis

maleimide/graphi te prepreg molded readily without incident. Non-destruc

tive testing indicated freedom from significant porosity and other flaws

Microscopic studies have not been performed to date, and will now probable be accomplished simultaneously with those done on exposed panels.

Similarly, part of the control testing will be performed at this latter

time, to accommodate possible diversion of some testing to other more

significant investigation in the event of unanticipated exposure effects

The API system retained about 80% of matrix critical room temperature

properties at 450F. Bending (plate) stiffness of panels is high enough

to meet natural frequency requirements during Space Shuttle launch and

landing. Dimpling associated with cocuring tends to reduce properties

3-13% as shown in Figure 10.

Condensation polyimide control specimens will be tested in modes similar

to those used in API testing upon availability of test panels. The

elevated temperature will, of course, be 600F.

Acknowledgments: The author is indebted to Mr. D. W. Welch, co-investigator, for direction of the analytical activity associated with this

experiment; to Mr. G. T. Smith for panel processing and test specimen

preparation; and to Mr. R. E. Haak for direction of support structure

fabrication and mechanical testing.

Ik
Table III. Physical and Mechanical Properties of

Bismaleimide/Graphi te Modified Quasi-Isotropic

Laminates [0(0/+45/90) 2s 0] , Hexcel

F-178/T-300, Precured

ROOM TEMPERATURE 450 F

PROPERTY NUMBER

OF

SPECIMENS

TEST VALUE NUMBER

OF

SPECIMENS

TEST VALUE

LOW HIGH AVG LOW HIGH AVG

TENSILE STR, ULT (0). KSI 5 61.2 71.6 66.5 3 69.7 73.6 71.7

TENSILE MOD [0], MSI 5 7.37 8.12 7.85 3 7.50 8.26 7.78

COMPRESSIVE STR, ULT [01, KSI 4 62.8 70.1 67.2

COMPRESSIVE STR, ULT [90] , KSI 4 44.5 51.8 47.7 3 37.7 45.4 40.9

COMPRESSIVE MOD [0], MSI 4 7.57 8.22 7.98

COMPRESSIVE MOD [90], MSI 5 5.92 6.67 6.29 3 6.51 6.74 6.63

RAIL SHEAR STR. ULT [90/0], KSI 2 27.4 28.5 28.0 1 22.3

RAIL SHEAR MOD [90/0], MSI 2 2.21 2.40 2.30 1 2.38

FLEXURAL STR, ULT [0), KSI 4 152.6 159.2 156.9

FLEXURAL MOD [0], MSI 4 11.20 11.59 11.32

BENDING STIFFNESS. El, LB-IN.2 4 345 416 368

INTERLAM SHEAR STR. ULT [01. KSI 5 5.46 6.50 5.99

PHYSICAL PROPERTIES SP GR 1.57, FIBER VOL FRAC 0.61, MOLDED PLY THICKNESS 0.0053 IN.

Table IV. Physical and Mechanical Properties of

Bismaleimide Polyimide/Graphite Modified QuasiIsotropic Laminates [0(0/+45/90)2s 0] , Hexcel

F-178/T-300, Cocured

ROOM TEMPERATURE 450 F

PROPERTY NUMBER

OF

SPECIMENS

TEST VALUE NUMBER

OF

SPECIMENS

TEST VALUE

LOW HIGH AVG LOW HIGH AVG

TENSILE STR. ULT [0], KSI 5 65.2 71.6 68.4 2 61.6 72.7 67.1

TENSILE MOD [0], MSI 5 7.18 7.95 7.64 3 7.90 8.54 8.20

COMPRESSIVE STR, ULT [0], KSI 3 52.2 60.7 56.3

COMPRESSIVE STR, ULT [901, KSI 5 44.6 48.8 46.3 3 32.6 42.0 38.3

COMPRESSIVE MOD [0], MSI 3 9.11 9.25 9.18

COMPRESSIVE MOD [90], MSI 5 5.57 6.45 6.03 3 6.17 7.17 6.57

RAIL SHEAR STR, ULT [90/0], KSI 2 24.7 24.8 24.75 1 19.4

RAIL SHEAR MOD [90/0], MSI 2 2.040 2.042 2.041

FLEXURAL STR, ULT [01 , KSI 4 121.3 155.3 139.7

FLEXURAL MOD [0], MSI 4 10.53 11.82 11.07

BENDING STIFFNESS, El, LB-IN.2 4 349 393 370

INTERLAM SHEAR STR, ULT [0], KSI 5 5.51 6.39 5.93

PHYSICAL PROPERTIES SP GR 1.56, FIBER VOL FRAC 0.60, MOLDED PLY THICKNESS 0.0053 IN.

15
Figure 9. Typical Stress-Strain Curves for F178/T-300

Bismaleimide/Graphite Laminates [0(0/+45/90)2S 0]

Figure 10. Summary of "Cocured" API Laminate Properties

in [0(0/+45//90)2c0] Construction, Expressed as Percent

of Precured" Properties

16
EFFECTS OF ROOM-TEMPERATURE AGING ON COMPOSITE PREPREGS

H. C. Nash, C. F. Poranski, Jr. and R. Y. Ting

Chemistry Division

Naval Research Laboratory

Washington, D. C. 20375

Abstract: Prepreg materials of organic-matrix composites have very

limited shelf-life, even under the required cold-storage conditions. A

prepreg aging study has been carried out by storing prepreg materials at

room temperature under three different controlled humidity levels. Exposed samples were examined periodically by proton NMR spectroscopy, and

unidirectional 16-ply laminates were fabricated. Prepreg appearance,

curing behavior, and laminate properties were found to change with

prepreg out-time.

Key words: Aging of polymer; prepreg; composite; NMR; mechanical

properties; fiber; moisture.

As fiber-reinforced composite materials become increasingly important

in many aerospace applications, the storage, handling and processing of

composite prepreg materials are also receiving greater attention. A

major difficulty in handling many prepregs is that the material must be

stored at very low temperatures. The shelf life of these prepregs is

generally only a few months at freezer temperatures. Presumably,

continued B-staging occurs at room temperature and, at the same time,

the resin may degrade by absorbing moisture from the atmosphere. Also,

since the prepreg must warm up to room temperature before lay-up for

fabrication, it could absorb moisture through condensation.

An aging study was therefore carried out to examine the effects of

out-time and humidity at room temperature on the properties of a

polyimide/graphite prepreg. Laminates were fabricated from material

which had been exposed to various humidities at room temperature and

the mechanical properties were evaluated as a function of prepreg outtime. Extracts of the aged prepreg were also analyzed using proton

NMR spectroscopy to see if this technique could reveal any chemical

change resulting from the aging process.

Experimental

Material: The polyimide/graphite prepreg (F-178/T300) was obtained

directly from the manufacturer, Hexcel. The resin is described as a

bismaleimide type of polyimide which cures through a free-radical

addition mechanism. The glass transition temperature, determined by

IT
torsional pendulum measurements, was 360°C for this resin compared to

260 C for an epoxy. Fresh prepreg material was cut, and stored at room

temperature in three environmental chambers controlled at 16%, 50% and

95% relative humidity, respectively. Aging extended over a period of

twenty-eight weeks. A 16-ply laminate was prepared every four weeks.

An additional sheet of aged prepreg was used for proton NMR analysis.

Unidirectional 16-ply laminates were fabricated in a hydraulic

press using conventional vacuum bag techniques. The processing included a full vacuum cycle at 121 C (250 F) and a pressure cycle at

177°C (350°F) with a pressure of 6.9 x 10 5 Pa (100 psi) . Post cure

o o was carried out at 249 C (480 F) for 10 hours. The curing cycle was

kept the same for processing all aged samples.

Samples were cut from the laminate and tests were carried out to

determine if any change occurred in their mechanical properties relative to composites prepared from unaged prepreg. The short-beam shear

test (ASTM-D 2344) was chosen to determine the interlaminar shear

strength, a resin-dominated property. The flexure test (ASTM-D 790)

was selected to measure the laminate flexural strength and modulus,

which are fiber-dominated properties. For proton NMR analysis, weighed

2 cm x 2 cm samples were extracted using acetone-d. for ~2 hours at

room temperature. The resulting solutions were filtered directly into

dry NMR sample tubes. The residues were dried and weighed. The spectra

were obtained on a JOEL FX60 Q spectrometer system.

Results

Figure 1 shows the flexural strength of the graphite/polyimide

composite plotted as a function of the prepreg aging period in weeks.

It can be seen that laminate flexural property steadily decreases as

aging continues. The effect is more pronounced at higher humidity

levels. After twenty weeks of aging at 95% relative humidity, the

flexural strength is reduced by nearly 30% from that of the original

control sample. The same trend is followed by the flexural modulus

data.

Figure 2 presents the interlaminar shear strength data as a

function of prepreg aging period. In this case, the results show a

rapid increase in shear strength initially as aging proceeds. Then it

becomes constant at a level approximately 55% higher than that of the

initial control. The increase seems to be independent of the humidity

level at which prepreg samples were aged.

At first glance these results seemed surprising because any resin

degradation taking place due to moisture absorption in the prepreg would

lead to a poorer resin-dominated property such as interlaminar shear

strength. The fact that interlaminar shear strength was improved and

flexural properties deteriorated may be resolved by considering some

physical changes to the prepregs which occurred during aging.

18
First of all, during the aging period the prepreg material clearly

showed a steady loss of tackiness. As aging continued, the prepreg

gradually became dry and brittle, making it difficult to keep all plies

properly aligned and adhered during lay-up. Secondly, the bleeding

during curing was greatly reduced for aged samples. Wetting of the

bleeder plies became less and less, particularly around the corners

of the square plies, as aging continued. These changes seem to suggest

that the amount of resin flow was greatly reduced due to aging.

Figure 3 shows the proton NMR spectrum of the resin extracted from

the control sample. Analysis of the spectrum showed that the extracted

resin had two major components. The lines at 4, 7 and 7.2 ppm were

assigned to the methylene, olefinic and aromatic protons of the

bis-maleimide of methylene dianiline (see Figure 4A) . The multiple

peakes in the region from 4.1 to 6.1 ppm were due to the allyl protons

in triallyl isocyanurate (see Figure 4B) . However, it was surprising

that there was no noticeable difference in the appearance of the

spectra of the aged samples compared to that of the control sample.

The most probable reason for this is that the initial products of the

aging process were not soluble in acetone-d^, or that they did not reach

a high enough concentration, before reacting further, to be detected.

It was noticed, however, that the amount of extractable materials

decreased with aging.

These observations suggest that exposure to high humidity caused

little or no chemical degradation in the resin system. But room temperature storage produced a prepreg material that was excessively B-staged.

The aged prepregs essentially had a higher degree of crosslinking than

a fresh sample at the beginning of the fabrication process. This change

in degree of crosslinking was not detected by proton NMR, but the

material was essentially one with higher molecular weights having

greater melt viscosities. This, therefore, lowered the flow when the

sample was subjected to the processing cycle originally designed for

the control sample. The result was a composite laminate having higher

resin content, but poorer flexural properties. Laminate density

measurements and fiber/matrix volume determination confirmed this. The

result showed that the resin content in the laminates increased with

increased age of the prepreg, resulting in decreased laminate density.

Reference

1. W. D. Bascom, J. L. Bitner and R. L. Cottington, ACS Org. Coat,

and Plast. Preprint, 38, 477 (1978).

19
in

o, m

Io

2.8

2.6 HO

®

2.4

M

2 .2

PSu

3

X

0) 2.0

1.8

6

O

X

o

-L

F178 T300

Room Temp. Aging

o

x

o

9

o

o

x

o

R.H.

16%

50%

95%

8 12 16 20

Aging Period (week)

24 28

Figure 1. Laminate flexural strength as a

function of aging period.

20

am

io 18

H

+Jtm 16

c

0

tl M

CO

Sh

<0 14

d) £1

COua

in 12

E

csu

a

•pc 10

F178/T300

Room Temp. Aging

5

o

9

_L _L

o

6

0 §

R.H. o 16%

X 50%

0 95%

8 12 16

Aging Period (week)

20 24 28

Figure 2 . Interlaminar shear strength of aged

samples

.

20
J I u J I L

ppm

Figure 3. 60 MHz proton nmr spectrum of material extracted

from unaged F-178/T-300 prepreg. The line at 0

ppm is that of the reference, tetramethylsilane.

The multiplet at 2 ppm is due to the residual

protons in the solvent, acetone -d .

CH

2

-CH=CH2

H

2C = CH-CH2-N N-CH2

-CH=CH2

B

Figure 4. Chemical structures of components in Hexcel polyimide

resin

.

I 21

SESSION 2

Chairman: S. E3. J. Johnsen

Detroit Diesel Allison Division,

GMC

APPLICATION OF INSTRUMENTAL TECHNIQUES TO THE STUDY OF

THE CURE OF PHTHALOCYANINE/GRAPHITE COMPOSITES

Robert Y. Ting

Chemistry Division

Naval Research Laboratory

Washington, D. C. 20375

Abstract: Work on resin evaluation in the NRL V/STOL aircraft program

is directed toward the development of high temperature resins for application as potential composite matrix materials. As successful prepregging of the polyphthalocyanine/graphite-f iber system has recently

been demonstrated, the processability of this material is now under

study. Instrumental techniques including differential scanning calorimetry (DSC) , thermogravimetric analysis (TGA) and dynamic

dielectric analysis are employed to examine the thermal properties of

unidirectional prepregs and the fabrication of unidirectional

laminates. Results of these studies have led to the development of a cure cycle for this new resin system.

Key words: V/STOL, polyphthalocyanine, graphite-fiber, DSC, TGA,

dielectric analysis, cure, prepreg.

Introduction : Organic composite materials having high temperature capabilities are being evaluated for structural applications in advanced

aircraft such as future vertical and short take-off and landing (V/STOL)

aircraft. The basic concept is to replace much of the metallic

structural components with organic composities for weight-saving

purposes and consequently to increase the range and payload of the

aircraft.

For such applications the presently available composite materials

suffer many shortcomings, including moisture pick-up, chemical degradation and poor storage stability. Most of all, current matrix materials

(epoxy resins) have a limiting use temperature around 135°C. Resins are

therefore being sought which will meet operational requirements at

temperatures in excess of 200°C, but will be relatively insensitive to

moisture attack and will have the ease of fabrication of state-of-theart epoxies. Phthalocyanine resins recently developed at NRL (1) seem

to provide the desired features when properly cured (2) . Figure 1 shows

the network structure of these polymers. The phthalocyanine polymer

based on the C-10 diamide resin is now being evaluated as a potential

matrix material in fiber-reinforced composites.

25
Experimental ; Thornel 300 graphite fibers were used in producing unidirectional prepreg tapes. Prepregging of the C-10/T-300 system was

carried out by both the usual hot-melt method and a resin/solvent

slurry method. In the hot-melt procedure the resin was staged at about

200°C to give a liquid intermediate state during which the fiber reinforcements were easily impregnated. The prepreg had the general

appearance of being dry and brittle; however, the material could be

briefly heated at 130°C to develop sufficient amount of tack for

complex lay-up. In the slurry technique, dimethyl formamide (DMF) was

used as the solvent, and the final product had an average resin content

of ca. 40% and volatile content of ca. 6%. Prepregs prepared by the

hot-melt technique, on the other hand, had ca. 33% resin content with

practically no residual volatiles.

A Perkin-Elmer DSC-2 unit was used for the DSC cure study. The

scanning was performed over the temperature range of 310°K (37 °C) to

500°K (227°C) at a heating rate of 10°C/min under a dry nitrogen

atmosphere. A DuPont 900 Series unit was also used to perform the TGA

work on prepreg samples 20 mg) over the temperature interval of 20°C

to 350°C. Sample weight loss was determined as a function of temperature at a heating rate of 5° C/min, while purging with dry nitrogen.

A dynamic dielectric technique (3) was employed to monitor the

curing of composite samples using an Audrey III Dielectrometer

.

Sixteen-ply unidirectional samples (15.2 cm x 15.2 cm in size) were

laid up by hand between two aluminum-foil electrodes, 5.1 cm x 5.1 cm

in size, for the dielectric monitoring. Sample capacitance and dissipation factors were continuously measured ^s a function of cure time

and temperature at a fixed frequency of 10 Hz. Conventional vacuumbag technique was used for sample cure in a hydraulic press (4)

.

Results ; The results of DSC scans on uncured polymer is shown in

Figure 2. Trace B is a "null-reference" baseline. Trace E Is from a

pure, uncured C-10 phthalocyanine resin, which exhibits a melting peak

at 180°C, and an exotherm at 197°C, related to the onset of cure.

Trace D is that for the prepreg material prepared by using the slurry

technique. The two sharp endotherms at low temperatures are perhaps

related to the release of some unknown, low boiling volatiles. A peak

at 145°C may be identified with the release of DMF added during prepregging. The other peaks at 134 °C and 174 °C are believed to be

related to the melting of resin oligomers. These temperatures are

lower than the melting point of pure C-10 resin. This phenomenon of

melting-point depression by B-staging is well-known (4) . Trace C in

Figure 2 is that for a different batch of prepreg, presumably prepared

according to the same procedure as the material of Trace D. It is

noticed, however, that the thermal behavior of the C-10 resin is quite

sensitive to its thermal history. Apparently, the two batches of prepregs were B-staged differently. In Trace C there is no endo therm

peak at 124°C, but the solvent peak and higher temperature melting

26
peaks are more pronounced when compared with those in Trace D. Trace

A in Figure 2 gives the record of hot-melt prepreg. Since no volatile

release is expected from this material, the endo therm peaks again

manifest the melting of oligomers of separate entities. This multiple

melting behavior of B-staged phthalocyanine was also observed by

Gillham recently in a torsion braid analysis of this resin (5)

.

In order to further identify the nature of the observed DSC

endo therm peaks, TGA was also performed; Figure 3 shows this result for

a slurry prepreg sample. Substantial weight loss is found in the

temperature range of 50°C to 90 °C, suggesting that the corresponding

DSC endotherms are indeed related to release of volatiles and/or

moisture. The further weight loss observed near 150°C shows DMF evaporation; the amount is nevertheless small. Rapid weight loss begins to

take place beyond 275 °C. This suggests a limit of initial cure

temperature to no more than 260°C.

Figure 4 shows the dielectric dissipation curve of a slurry prepreg sample subjected to an isothermal cure at 225°C. Small peaks at

ca. 100°C and 150°C show the release of moisture and DMF, respectively.

As the temperature exceeded 200°C, the material quickly melted and

polymerization to phthalocyanine was also activated thermally. This

caused the dissipation to increase and reach a maximum, which was

believed to be the gelation point (6) . Beyond this point the sample

gradually solidified, but a distinct cure-peak was not observed as in

the case of epoxy-matrix materials (7).

Similar techniques may be applied to laminate samples during the

fabrication in a press. In Figure 5, an example for in situ dynamic

dielectric monitoring is given. In this case, a 16-ply laminate sample

was placed in a vacuum bag, and initially heated to 145°C under vacuum.

Temperature was subsequently raised to 260°C rapidly and kept for an

extended period of time. The dissipation curve showed at first a broad

melting (Figure 5A) , followed by a rapid increase to reach the gel

point. This would allow one to determine the exact time at which high

pressure should be applied to the sample in order to fabricate

laminates with very low void contents. Figure 5B shows the dissipation

of a prepreg lay-up with raw resin powder sprinkled between the plies

in an attempt to increase the sample resin content. As resin melting

temperature was reached, one observed an excessive amount of resin flow

accompanying the melting, as indicated by the first pronounced dissipation peak. It suggests that the control of this resin flow would be

difficult since the melt viscosity of unstaged C-10 resin is apparently

low.

In summary, it has been shown that studies of prepreg materials

using DSC, TGA and dynamic dielectric techniques prior to laminate

fabrication enable one to develop information concerning the thermal

processes such as volatile release, melting and gelation. In situ

27
dynamic dielectric analysis during fabrication provides sample

capacitance and dissipation data, which are related to the chemorheological changes in the curing of a thermosetting system. This

information is useful in identifying the melting and gelation of

laminate samples. Proper pressure application with respect to the

gel point led to cured samples with desired resin content and minimum

void content, which is critical for applications where fracture

processes dominate in composite failure.

References

1. T. R. Walton, J. R. Griffith and J. G. O'Rear, ACS Advances in

Chemistry Ser., 142 , 458 (1975).

2. T. R. Walton, J. R. Griffith and J. G. O'Rear, ACS Org. Coat.

Plast. Preprint, 34, 446 (1974).

3. N. G. McCrum, B. E. Read and G. Williams, Anelastic and Dielectric

Effects in Polymeric Solids , John Wiley, New York, 1967.

4. J. V. Gauchel and H. C. Nash, in High Performance Composites and

Adhesives for Advanced V/STOL Aircraft , ed. L. B. Lockhart and

W. D. Bascom, NRL Memo . Report 3433, Washington, D. C, 1976.

5. J. V. Gillham, ACS Org. Coat. Plast. Preprint, 40, 866 (1979).

6. J. K. Gillham, J. A. Benci and A. Noshey, J. Polym. Sci., C46 , 279

(1964)

.

7. C. A. May, D. K. Whearty and J. S. Fritzen, SAMPE Ser., 21, 803

(1976) .

28
HEAT

Fig. 1: Polyphthalocyanine network

0 50 100 150 200 250 300 350 (°C)

Temperature

3: Percent weight loss of C-10/T-300 prepreg

as a function of temperature

29
LU

opug

30
TIME

Fig. 4: Dielectric dissigation record of a prepreg

sample under 225 C cure

Fig. 5: Dynamic dielectric results of in-situ

monitoring of a 16-ply C-10/T-300 laminate

31
MOISTURE DIFFUSION ANALYSIS FOR COMPOSITE MICRODAMAGE

C. L. Leung and D. H. Kaelble

Rockwell International Science Center

Thousand Oaks, California 91360

Abstract: Because of their favorable performance characteristics, advanced composites using graphite fibers as reinforcement are increasingly being used in aerospace and commercial applications. There has

been much concern, however, about the extent and mechanisms of strength

degradation of these composites when exposed to moisture for long periods of time. In this paper, methods to obtain directional diffusion

coefficients, i.e., D

x , D

y

and D

z for the principal axis of a unidirectional fibrous composite are described by moisture absorption and effusion measurements. Additionally, analytical models have been developed

to calculate, using the previously established diffusion coefficients,

the moisture profile of the composite after periods of moisture exposure. The feasibility of the integration of the effusion rate measurement, moisture profiling, and strength prediction into a computer-aided

analyzer for composite microstructure damage or defect is discussed.

Key words: Diffusion coefficients, iteration, moisture profile modeling, nondestructive evaluation, statistical estimation.

INTRODUCTION

A number of reports show that the mechanical properties of graphiteepoxy composites are degraded by moisture [1-4]. Hydrothermal stresses

generated by hydrat ion-dehydration can cause microcracks and del ami nations to occur within the composite. This paper outlines an iterative

calculation method to isolate the coefficients of diffusion for the

three axes of rotation. Using these coefficients, as well as the rates

of water effusion from the composites, an inversion method which solves

for the original moisture profile is also reported.

THEORY

1. Directional Diffusion Coefficients

Fick's second law of diffusion can be represented in terms of three

principal axes of diffusion by the following relationship [5-7]:

3c

3t

= D

x + D

y

+ D

z (1)

32
The schematic of Fig. 1 identifies the orientation of the x, y, z coordinate system for Eq. (1) as centered within a unidi rectional ly

reinforced composite. Recent studies show that for undamaged composites, Eq. (1) is obeyed and appears to provide a valid basis for diffusional kinetics analysis [3]. Uniaxial ly reinforced composites can

be rotated relative to the x, y and z coordinates to isolate the diffusion coefficients D

x , D

y

and D

z as shown in Fig. 2. In the limit of

short times for moisture absorption, the following relationship for

the initial $ vs t ' was derived [5]:

t- o

oo o

4

1/2 1/2 1/2 1/2

(2)

In Eq. (2), it is assumed that each of the six sides of the composite

absorbs and desorbs moisture independently. An alternative expression

for Eq. (2) is [5,7]:

i = x,y,z

4_

x. 7T

1/2

(3)

where is the apparent experimental diffusion coefficients as deter

mined from the initial slopes of the <(>,• 1 vs t ' curves.

This averaged experimental diffusion coefficient D- can be reexpressed

as [5,7]:

D, = D 1 +

x

i D

i h

D. (4)

where by definition, the following axis transformations would apply:

when i = x, then j = y, k = z

when i = y, then j = x, k = z

when i = z, then j

= x, k = y

Inspection of Eq. (4) shows that the measured values of _ „ y z must be further analyzed by an iterative calculation to isolate precise values of D

x

, D and D

z

. Rearrangement of Eq. (4) provides the

following

33
(D. ) = D. 1 + -!-

(D.) i x.

(D,

)

i

; n-l ( DVn-l

(5)

where n = number of iterations. This iterative calculation continues

until an appropriate convergence test condition is satisfied such as

[8]:

0<

{< D~i>n- (°~i>n-l}

:

< e

{< D~i>n " < D~i>n-2

}

When Eq. (6) is satisfied, the diffusion coefficient will be:

(6)

D. = (D.) + e

i = x,y,z

v l 'n n (7)

For the first iteration where n=l, the initial inputs for iteration are

the measured coefficients of diffusion in the three principal axes as

fol lows:

< D i>n-l D

i = x,y,z

< D~n>n-l

= 0.

J = y,x,z

(8)

< D

~k>n-l -\ = z,x,y

Equation (5) is thus solved for the principal axis i = x,y,z to compute

(D x

) n= 2> (D

y

) n= 2 (D

z

) n= j. Successively higher iterations of n=2, 3,

4... can thus be calculated using the appropriate (n-1) values.

Moisture absorption and desorption experiments in the principal axis

i = x,y,z of a uniaxial composite followed by iterative calculations

leading to Eq. (7) are therefore proposed as a useful approach to isolate the principal axis components of moisture diffusion kinetics in

these uniaxial composites.

2. Moisture Profile Modeling

Subsequent to the determination of the directional diffusion coefficients as discussed above, the depth/concentration profile of moisture

3^
in the composite can be obtained by a series of mathematical analyses

and experiments. The experiment involves the heating of the composite

to a desired temperature followed by measurement of the rate of moisture effusion as a function of time. The analytical problem is then

to determine the initial distribution of water, i.e., the concentration c Q (x) = c(x,o) as a function of x (where x is the distance into

the thickness of the composite), from the subsequent history of effusion rate measurements. The attempt to treat this as a deterministic

problem leads to difficulties. One of these difficulties is associated with the inevitable incompleteness of the experimental data:

any real experiment can yield only a finite set of numbers. A particularly convenient way is to apply statistical estimation theory to

the analysis. This analytical methodology is termed inversediffusion.

The inverse-diffusion system is illustrated more explicitly in Fig. 3.

The material is assumed to be a rectangular block of composite occupying a region of space defined by the inequalities

0 < x < L x

0 < y < L

y

(9)

0 < z < L

z

where x, y, z are the usual Cartesian coordinates. The principal

values of the diffusion tensor are denoted by D

x , D

y

and D

z

. The

initial concentration is assumed to depend only on "x, i.e.,

c(x,y,z, t=0) = c 0(x) (10)

It is further assumed that the measurements involve the total diffusion rate from the entire surface at a specific set of later times.

The inversion problem, therefore, involves a theoretical solution of

the above diffusion problem with an arbitrary c Q (x) and dry boundary

conditions (c^ = o) [9]. The matrix l_ mnshown in Fig. 3 is then used

to obtain from the measured J m

(the experimental total diffusion rate)

at times tm estimates a n of the coefficients of sin (nfr x/L ) , i.e.,

c Q

(x) = c, +

1 si n-nirx

(11)

n=l

This process of going from the experimental J

m

to the estimates a„ is

incorporated in the box labelled "estimator" in Fig. 3. The final

result of the entire estimation process is presented as a plot of the

original moisture concentration at time t=0 on the ordinate and the

depth of moisture penetration on the abscissa.

35
To establish the validity and accuracy of the estimator, called the

Inverse Diffusion Model, a Direct Model is independently developed

[10], which is based on a numerical solution of Fickian diffusion

equations. The Direct Model gives data for:

(a) Moisture profile at any specific time,

(b) Moisture diffusion rate at each time interval,

(c) Moisture content at each time interval.

Moisture effusion rates obtained from the Direct Model, item (b), can,

therefore, be analyzed by the Inverse Diffusion Model, yielding moisture profiles which should correlate to those generated from Fick's

Law, item (a), thus validating the Inverse Diffusion Model.

EXPERIMENTAL

A 48-ply unidi rectional ly reinforced composite panel (Hercules 3501-5

epoxy reinforced with Hercules Type AS graphite fiber) was fabricated

and cured by standard production procedures. The composite has a fiber volume Vf = 0.60 and void volume V

v

< 0.01.

Moisture absorption kinetics were evaluated by measurements of weight

change as a function of moisture exposure using an analytical balance

sensitive to 10 g. Moisture effusion kinetics were evaluated on the

Moisture Evolution Analyzer (DuPont Model 902H) supplemented by weight

change measurements.

Thin plate specimens were cut to-provide dimensions of length * width

x thickness = (20 * 5 * 1) * 10" 3 m. They were completely dried in a desiccator prior to moisture exposure.

RESULTS AND DISCUSSION

The results of moisture diffusional analysis for successive cycles of

absorption (D)n and desorption (D)p are summarized in Table I. The

results in Table I show that each additional cycle of full hydration

and subsequent drying produces large increases in diffusional coefficients for moisture absorption (D^)^. However, the coefficients for

moisture desorption (D,-)q are essentially constant over the three

hydration-dehydration cycles. The results are indicative of a damage

mechanism in the composite microstructure which enhances moisture

uptake from the dry state, but does not affect moisture release from

the wet state.

Using the coefficients of diffusion tabulated in Table I, a series of

the profiles generated by the estimation process of the Inverse

Diffusion Model are compared against the profiles from a Direct

(Fickian) Model, as discussed previously. For the sake of argument,

36
the composite was assumed to be completely saturated with water prior

to effusion measurements and the moisture profiles were calculated at

times of 20 and 100 minutes after effusion started. The moisture profiles as calculated from the Inverse Diffusion Model were compared

against a numerical solution of Fick's Law, as shown in Fig, k,

A much more severe test of the estimator of the Inverse Diffusion

Model is for the case of varying boundary concentrations, in which

case the specimen is hypothetical ly partially saturated, removed from

moisture and the moisture penetration mapped after varying lengths of

drying times. Again using the coefficients of diffusion tabulated in

Table I, analytical solutions of moisture profiles solved for 20 and 100

minutes of absorption and various drying times are shown in Fig. 5.

Several important conclusions can be drawn from the results shown in

Figs. 4 and 5. In each case, there is excellent agreement between the

Direct and Inverse solutions, thus validating the Estimator of the

Inverse Diffusion Model. Particularly for Figs. 5a and 5b, after dehydration has begun, a ridge of high moisture content is present immediately beneath the surface which is dry. It is feasible that there

exist different states of stress to the right (into the center) and to

the left (on the surface) of the ridge of water beneath the surface.

If a thermal spike is applied to the composite at this time, the different stress states can cause microcracks or delamination beneath the

surface of the composite.

Such subsurface microcracks can be detected by measuring the moisture

effusion rates and analyzing the data by the Inverse Diffusion Model.

Figure 6 shows the estimated moisture profile of a partially saturated

composite which contains two layers of air voids, as shown in Fig. 7.

The discontinuities at X/L values of 0.1 and 0.9 as shown in Fig. 6

are correlated to defects within the sample.

SUMMARY AND CONCLUSIONS

This report describes and applies a quantitative methodology to the

diffusion kinetics analysis for moisture degradation in graphi te-epoxy

composites. An Inverse Diffusion Model has been developed by an estimation process. This model has been shown to give correct moisture

profile at t=0, given a set of effusion rates at later times. It can

also serve as a sensitive, nondestructive analytical methodology for

subsurface microdamage and defects.

ACKNOWLEDGEMENTS

This research was sponsored by the Center for Advanced NDE operated by

the Rockwell International Science Center for the Advanced Research

Projects Agency and the Air Force Materials Laboratory under Contract

F33615-74-C-5180.

37
REFERENCES

(1) Transactions on the Workshop on "The Effects of Relative Humidity

and Elevated Temperature on Composite Structures," Ed.

J. R. Vinson, R. B. Pipes, W. J. Walker and D. R. Ill rich, Air

Force Office of Scientific Research Tech. Report AFOSR 770030.

(2) C. E. Browning and J. T. Hartness, "Effect of Moisture on the

Properties of High Performance Resins and Composites," ASTM STP

546, American Society for Testing and Materials, Philadelphia,

1974, pp. 284-302.

(3) D. H. Kaelble, P. J. Dynes and L. Maus, "Kinetics of Environmental

Degradation in Graphite-Epoxy Laminates," ASTM STP 580, American

Society for Testing and Materials, 1975, pp. 247-262.

(4) E. L. McKague, J. E. Halkais and J. D. Reynolds, J. Comp.

Materials, _9, 2 (1975).

(5) W. Jost, "Diffusion in Solids, Liquids and Gases," Academic Press,

New York, 1960.

(6) J. Crank, "The Mathematics of Diffusion," 2nd Ed. Oxford

University Press, 1975.

(7) C. H. Shen and G. S. Springer, J. Composite Material s, _10, 2

(1976).

(8) A. Ralston, "A First Course in Numerical Analysis," McGraw Hill,

New York, 1955, p. 328-332.

(9.) J. M. Richardson, "The Inverse Diffusion Problems," Internal

. Letter: "Diffusion With Arbitrary Boundary Concentrations," Feb.

1978.

(10) J. M. Richardson, "Diffusion with Arbitrary Boundary Concentrations," Feb. 1978.

38
TABLE I. Moisture Diffusion Kinetics

O Q

Experimental Diffusion Coefficients (cnr/sec) * 10°

Adsorption Resin Transf ibrous Transl aminar Inter! ami nar

Cycle

1 2.37 1.62 0.68 0.63

2 3.4 4.89 1,32 1.36

3 7.14 1. 77 1. 75

Desorption

Cycle

1 3.14 22.0 5.18 4.90

2 5.03 16.0 4.67 4.90

3 18.6 5.18 4.90

Iteratively Calculated Diffusion Coefficients, (cm^/sec) * 10^

Absorption Resin Transf ibrous Transl aminar Inter! aminar

Cycle

1 2.08 1.30 0.44 0.39

2 2.98 3.94 0.81 0.85

3 5.81 1.08 1.07

Desorption

Cycle

1 2.75 18.0 3.18 2.95

2 4.87 12.9 2.87 3.06

3 15.0 3.54 3.39

39
FIGURE 1

Orientations of fibers and plies within a unidi rectional ly

reinforced composite (fibers parallel to x-axis, plies

parallel to xy plane).

ko
x - 20 mm

x

x

= 1 mm

y

x = 20 mm

Figure 2.

Specimen orientations and geometry for kinetics of

adsorption and desorption in interlaminar , transfibrous, and translaminar directions of uniaxial

reinforced composite (layers represent plies

and dots fiber ends).

kl
SOLN OF

DIFFUSION EQUATION

SYSTEM

MODEL

ON LINE

OFF-LINE

ESTIMATOR

c°(x)

I M,N a. &, e.

M = NUMBER OF MEASUREMENTS

N = NUMBER OF ESTIMATIONS

c-i = CONCENTRATION OF WATER IN THE COMPOSITE CORRESPONDING

TO THE CHEMICAL POTENTIAL 1 a = A PRIORI BIAS AGAINST LARGE AMPLITUDE c°(x)

(j = A PRIORI SMOOTHNESS WHICH IS APPLIED TO c°x IN THE RANGE ( 7T. )ir

i = NUMERIC RANGE FACTOR 0 c 0.5

a \ = OBSERVATIONALLY CONDITIONED AVERAGE FOR an n J

FIGURE 3

Schematic representation of the estimation process.

h2
20 MINUTES DESORPTION

1 I — ===^==

(a)

0.05 0.1 0.15 0.2 0.25 0.3 0.35 0:4 0.45 0.5 X/L

IO0 MINUTES DESORPTION

1i

X/L

FIGURE 4

Moisture profiles via analytical modeling of a fully saturated

composite after (a) 20 minutes desorption, (b) 100 minutes

desorption.

h-5
20 MINUTES ABSORPTION

0.9 -

0.8 DIRECT FICKIAN SOLUTION

ESTIMATOR INVERSE SOLUTION

1/2 MINUTE DESORPTION

20 MINUTES DESORPTION

0.05 0.15 0.2 0.25 X/L

100 MINUTES ABSORPTION

0.35 0.4 0.45 0.5

FIGURE 5

Moisture profiles via analytical modeling of partially saturated

composite after (a) 20 minutes absorption and (b) 100 minutes

absorption.
45000

40000

TRANSFIBROUS

35000

0.1 0.2 0.3 0.4 0.5X/L

0.6 0.7 0.8 0.9 1.0

FIGURE 6

Moisture profile of composite containing air voids.

^5

RECENT WORK ON TECHNIQUES AND APPLICATIONS

OF MOISTURE BARRIERS TO GRAPHITE EPOXY COMPOSITES

James F. Haskins

General Dynamics Convair Division

P.O. Box 80847

San Diego, CA 92138

Abstract: A number of space systems such as Thematic Mapper, MAGSAT and other space optical devices require that the components be dimensionally stable during both ground and space

operation. It has been found that certain graphite/epoxy composite laminates can provide extremely low coefficient of thermal expansion (a < 0.05 x 10"^ in./in.F) and because of this are ideally

suited to optical benches, telescope housings, and mirror substructures. However, the epoxy must

be protected from moisture because it is hygroscopic and does change dimensions as it absorbs or desorbs moisture. Typically it may change as much as 100 microinches/inch over a period of

several months when exposed to high relative humidity. This becomes a problem because most of

the equipment will be manufactured, assembled, and operated in a humid environment before it is launched into space. Holding the humidity at a low level lessens this problem but does not

eliminate it.

A real advance in optical equipment for use in space can be made if graphite/epoxy can be sealed

from moisture without substantially changing its weight and coefficient of expansion. The objective of this work was to test the effectiveness, develop inspection procedures, and check the

reliability of a moisture barrier developed previously at General Dynamics Convair Division.

This paper gives the results of a critical evaluation of the effectiveness and reproducibility of the

eutectic coatmg. Fifteen-inch long graphite/epoxy coated specimens were used to study processing

variables, inspection techniques, and dimensional changes due to moisture pickup.

Key words: graphite/epoxy; advanced composites; dimensional stability; space optics; moisture

barrier; eutectic coating.

Graphite composite materials have been successfully applied to a variety of satellite structures.

Weight saving resulting from excellent specific mechanical properties has been realized in addition to the attainment of stringent dimensional stability. Research and development efforts are now directed toward challenging new applications. During 1978 this effort was focused on the

development of application processes for an existing moisture barrier for graphite/epoxy and

the experimental verification that such a coating does indeed seal the composite from moisture.

General Dynamics discovered the moisture problem in 1970 while evaluating graphite/epoxy

composites on both Air Force and NASA programs (References 1 and 2).

^7
Graphite/epoxy composites absorb moisture from the atmosphere. This absorption causes two

basic problems; i.e. (1) decrease in matrix-dominated mechanical property values and (2)

changes in physical dimensions. The effects are reversible in that heat and/or vacuum will drive

off the moisture and will return the composites to their dry properties and dimensions. The absorption of moisture primarily occurs in the epoxy 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. Epoxy

matrix composites reinforced with boron, glass, and Kevlar filaments behave similarly to

graphite/epoxy composites.

This work is concerned with sealing GY-70/X-30 graphite epoxy from moisture in its isotropic

layup [(0/45/90/135)s]2. This material was chosen because of its low coefficient of expansion.

This makes it ideal for optical devices to be used in space. The problem with moisture absorbed

by the material is that it will give up the moisture when subjected to the vacuum of its space en- vironment. Certain optical systems will detune as the moisture is removed by the low pressure

26 30 34 38 42 50 54

Year

Figure 1. Example of length change with time underfixed environmentfor two sets of nominally similar

1-inch steel gage blocks (Emerson 1957).

hQ
of the working environment. Since ground time for many of these devices is as much as a year

in 50% relative humidity, one finds that the dimensions of the optical system can change by

amounts which are unacceptable. For this reason it is important to be able to provide a

moisture barrier for the composite material which will seal it completely from humidities up to

100%.

Dimensional stability of materials is not generally well understood. When speaking of holding

dimensions to one part in a million for long periods of time one must realize the difficulties involved when even the most stable metals are used, for example, gauge blocks, such as Johanson Blocks, made of a specially heat treated, fine grained steel which has been thermally cycled

many times. Even when blocks of this type are held at constant temperature they may change in

length over a period of years.

Work done by the Bureau of Standards is shown in Figure 1 for 8 blocks which were observed

for 25 years. This background allows one to appreciate the extreme demands placed on a

material such as graphite/epoxy. A material of this type, made up of two components of vastly

different coefficients of expansion, certainly contains residual stresses because of its method of

manufacture in addition to dimensional changes caused by moisture absorption. In Figure 2

both weight and length changes are shown for a typical bare laminate of the type being studied

here. The barrier being studied must seal the surface of the material so that no weight or length

change occurs over long periods of time.

200

AL/L,

MlN./IN.

160 -

140 -

120 -

100 -

Thickness 0.080 in. environment RT & 52% RH

Measured data: AW - 2 1 0 days

AL/L - 1 50 days

AL/L

1.8

1.6

1.4

- 1.2

- 1.0

- 0.8

- 0.6

- 0.4

- 0.2

0

AW, %

0 100

Figure 2. Environmental effects on [(0/45/90/135) s\ 2 GY-70/X-30.

200

DAYS

300

Objective: A real advance in opticl equipment for use in space can be made if graphite/epoxy

can be sealed from moisture without substantially changing its weight and coefficient of expansion. The objective of this work was to test the effectiveness, develop inspection procedures,

and check the reliability of a moisture barrier developed previously at General Dynamics Convair Division. The specific objectives for 1978 were:

h9
1 . Develop a better understanding of the processing variables and their sensitivity to support the development of a production process specification.

»

2. Develop quality control procedures and accept/reject criteria for production hardware.

3. Demonstrate Items 1 and 2 above through fabrication and test of six 15-inch specimens

using the established procedures.

4. Scale-up to approximately full-sized cylinders and demonstrate the effectiveness of the

coating through test of simulated full-scale Thematic Mapper hardware.

5. Modify, if required, our specification for applying the coating to production hardware.

Approach: The effectiveness of the coating was studied by preparing samples of

graphite/epoxy and coating them using the process then currently in use and measuring their

dimensional stability when exposed to a humid environment.

The then current process was examined step by step and comparisons were to be made to the

original process by the inventor to determine if any changes had occurred during the four years

of use since its conception in 1974. Then, additional specimens were prepared using the technique representative of the original process. These specimens were carefully inspected during

processing to determine if performance correlated with an exacting visual inspection.

Repeatibility was studied once the process had been perfected. This was done by preparing

several batches of specimens using the standardized process.

Eddy current techniques were developed to determine whether coating thickness could be

evaluated by nondestructive means. Long-term evaluations were made to study the stability of

the coating by observing the length of a number of specimens exposed to high humidity.

Length changes were measured and recorded to the nearest microinch per inch using a special

laser dilatometer. Weight gain of each of these specimens was also recorded so that any observed weight gain could be correlated with the corresponding length change. For those specimens

where length and weight changes were recorded, a check of the inspection records was made to

determine whether deflects in the coating correlated with these changes.

A twelve-inch diameter cylinder was to be fabricated and coated. This cylinder was exposed to

accelerated aging (95% relative humidity) and weight gain determined.

Progress: As part of a 1974 IRAD program, a moisture barrier coating was developed for sealing graphite/epoxy. This coating was a very special tin-indium eutectic which could be applied

over the composite surface after it was copper plated. The coating was applied to a few small

laboratory specimens and was checked by observing them in a humid environment. Their

dimensions did not seem to change in this humid environment and the moisture barrier was

considered a real breakthrough. Since its invention the requirements for dimensional stability

have increased, the size of the components have increased, the plating and coating tanks have

been enlarged, and the original chemicals have been replaced. As is usually the case in newly in50
vented processes, more research is needed to completely understand and document the process.

The following progress was made for each of the specific objectives for 1978.

1 . Processing Variables — Starting with surface preparation, it was found that the process is extremely sensitive to surface preparation and the amount of material removed by glass

beading had to be carefully controlled. It was further found that contamination existed in the

glass beading unit itself and a separate beading unit was recommended.

Generally speaking, standard plating practices were not being followed. Current settings were

estimated rather than calculated. Bath control ws poor and had to be greatly improved.

The eutectic coating is applied by spraying the eutectic in a stearic acid bath which is maintained at a temperature above the eutectic melting point. This stearic acid bath was contaminated

and had to be cleaned. The spray nozzle used in applying the eutectic had been modified and

had to be replaced. A brushing process used during the application of the coating had been

eliminated, thus resulting in some localized bare spots; in these areas, the copper plating was

visible. The excess eutectic is removed by a slinging process after the entire surface is covered.

This slinging process had been changed to eliminate a bubbly appearance which later proved to

affect the sealing of the composite surface.

2. Accept/Reject Criteria — To develop accept/reject criteria a number of specimens were

prepared and their acceptability was measured by observing their dimensional stability.

Careful visual inspection of each specimen requires a great deal of time and the use of a lowpowered microscope. A trained observer with proper light and the correct magnification was

able to check the coated surface for pinholes, blisters, uncoated areas, and contaminates. A

system was developed during the investigation for cataloging the imperfections and recording

the data for the entire surface of each specimen. The performance of each specimen could be

measured by observing changes in length during moisture exposure and comparing this to the

number of imperfections found during visual inspection.

A method for measuring coating thickness was developed by using eddy current techniques.

This method was able to provide a quick way of covering the entire coated surface. The instrument was adjustable and the impedance setting was varied to obtain the greatest sensitivity.

Weight gain attributable to the coating was also used as a method of measuring coating

thickness. Since the moisture barrier is made up of electroless copper, copper plate, electroless

tin and the eutectic, the specimens were weighed after each process step so as to measure the

thickness of each layer of the coating. It was found that the copper must be thick enough for

eutectic coverage. The eutectic thickness was about half as thick as the copper. These

thicknesses amount to a weight change of several percent. For example, the surface area of our

specimens was about 75 square inches and they weighed about 85 grams. For this case, 0.0001

inches of uniformly distributed copper would increase the weight by approximately one gram.

The eutectic would increase the weight approximately 0.8 grams per 0.0001 inches of thickness.

Coating thicknesses for composites can be checked very accurately by weight if the coating is uniform.

51
To check coating thickness and uniformity, a number of specimens of known performance in

the area of moisture absorption were sectioned, mounted, and polished prior to microscopic

examination. These mounts showed the uniformity of the coating and confirmed the calculated

average thicknesses of each component. The variation in coating thickness was recorded and

compared to previously recorded eddy current readings for the same areas. It was found that

the eddy current measurements were recording the copper thickness and that variation in the

tin-indium coat made little change in the eddy current readings. This provided a method of inspecting the copper thickness both before and after application of the eutectic. The eutectic

thickness was checked then by weight and visual observation.

Scanning electron microscope studies were made of the coated surface. Extremely small

pinholes and uncoated areas could be detected using this instrument. Such observations are not

useful on large components. Studying very small pinholes and their cause did help us confirm

the critical fact that the coating bath temperature must be below the temperature at which the

epoxy is post cured. If this is not the case, further curing of the epoxy can cause bubbles and

pinhole formation in the coating, rendering it less effective.

3. Demonstration Through Fabrication and Test — The composite material evaluated was GY70/X-30 graphite/epoxy, 16 plies thick and laminated in what is known as the isotropic layup

Figure 3. Hygroscopicity dimensional measurement test.

52
Material: [(0/45/90/1 35)s]2 GY-70/X-30

Test conditions: RT&95% R.H.

01 2345678

Exposure time in (days) 1/2

Figure 4. Dimensional stability study ofsix graphite/epoxy specimens (January 30 to February 8, 797.

MATERIAL: [(0/45/90/1 35) s]2 GY-70/X-30

TEST CONDITIONS: RT & 95% R.H.

DAYS

WT. GAIN TOTAL LENGTH CHANGE

SPECIMEN NUMBER (%) [fx IN./IN.)

o H2538 0.005 1.9 a H2539 0.006 1.0 H2540 0.007 0.4 • H2541 0.005 -0.2

* H2542 0.004 1.2 H2543 0.006 0.5

AVERAGE 0.006 0.8

Figure 5. Dimensional stability study of six coated graphite/epoxy specimens (January 30 to February

16, 1978).

53
[(0/45/90/1 35(s]2- This layup gives a nominal thickness of 0.080 inches. The specimens were 15

in. long and 2

lA in. wide. This large size was chosen to increase both the accuracy of the length

change measurement and the surface to be coated. The specimens used in the original work

were about one third this size. The type of equipment used to make the measurements is shown

in Figure 3. Length change was measured by comparing the specimen length to a calibrated

piece (ULE) Ultra low expansion glass. This was done with a tilt mirror using a laser light

source and an optical lever of 20 ft. This data was recorded on a stationary glass plate. The tilt mirror was supported on knife edges formed by beveling one end of both the ULE and the test specimen. The specimen was mounted against stops on the ULE and a precise spacing is established between the knife edges. This spacing and the length of the optical lever determined

the magnification factor used in obtaining the length change. The length change was determined by reflecting the small laser beam from both the tilt mirror and the reference mirror and

measuring the separation between these projected small spots. This difference can be measured

to the nearest 1/64 of an inch so that with a knife edge spacing of one tenth inch and the reflection projected 20 ft from the mirror one can easily measure one half a microinch/inch using a

15 inch specimen.

After the process variables were brought under control by making and testing several batches

of six specimens it became difficult to reduce dimensional changes much below what is shown

in Figure 4. Figure 4 also shows the difference between uncoated specimens and coated

specimens when exposed to 95% relative humidity for a number of days.

WT. GAIN TOTAL LENGTH CHANGE

SPECIMEN NUMBER (%) fa IN./IN.) o H2555 0.013 2.4 a H2558 0.010 1.8 H2556 0.011 1.8 * H2557 0.015 1.3 • H2554 0.009 2.4 H2559 0.007 2.4 AVERAGE 0.011 2.0

Figure 6. Dimensional stability study ofsix coated graphite/epoxy specimens (February 17 to March 9,

1978).
c

CDCD

C

CO szo

U)

c

CD

1,000 2,000 3,000 4,000 5,000

Time (hr)

Figure 7. Exposure of eutectic coated [(0/45/90/135)^2 GY-70/X-30 specimens (average of six

specimens).

Note: Eddy Current Meter Settings Bare GY-70/X-30:0 Meter Reading

Standard Coated Specimen: 80 Meter Reading Reactance Setting X440 R479

Outside Inside Top Baffle Bottom Baffle

Top Middle Bottom Top Middle Bottom Top Bottom Top Bottom

1 86 72 82 73 44 62 53 42 40 47

2 88 76 86 85 39 58 58 47 46 49

3 87 78 86 82 41 58 58 43 40 44

4 97

over 78 83 67 46 65 56 45 40 43

5 100

over 80 93 93 44 70 56 42 40 43

6 100 83 75 63 44 58 52 44 42 47

7 97 77 77 58 39 60 57 45 42 44

8 88 70 73 70 40 54 54 42 40 43

9 80 60 72 61 40 60 56 44 39 43

10 82 62 76 77 39 59 60 51 47 47

1

1

84 73 86 69 43 66 56 42 45 54

12 100

over 84 97 63 40 63 56 47 43 57

13 100

over 97 100 70 45 86 56 44 47 57

14 100 92 85 72 42 65 56 43 47 58

15 93 85 88 69 40 66 56 45 44 50

16 83 80 80 75 42 61 56 43 42 49

Bottom

Table 1. Eddy current measurements for eutectic-coated BY-70/X-30 cylinder.

55
Figures 5 and 6 show studies made on two succeeding sets of six coated specimens. The

moisture absorption did not correlate with the length change for the twelve specimens shown in

the figures. When they were taken as two sets of six, and averaged both in weight gain and

length change, Figure 5 shows a smaller weight gain and a smaller change than the set of Figure

6. However, in both cases the length change exceeded what would have been expected for an

uncoated material in that environment.

The set of six specimens given in Figure 6 were kept in the 95% relative humidity environment

for 2500 hours (approximately 3Vz months) at which time they came to an equilibrium length.

This is shown in Figure 7. At this time the moisture was removed from the chambers and the

specimens were evacuated for another 3,000 hours. As can be seen in Figure 7, the lengths of

the specimens remained constant. This led to the conclusion that the recorded length change

was not due to moisture. The specimens had all been thermally cycled a number of times to

stabilize them and relieve residual stresses, however, at this low level of length change, such

changes could very well be additional stress relief. More work needs to be done in this area to

determine if this is a real effect. It could be a problem if the optics were final tuned earlier than

four months after the graphite/epoxy is laminated.

4. Simulated Hardware Component Test — To determine the problems associated with coating

large components a 12-inch diameter cylindrical tube containing two baffles was coated using

the improved process. The coated cylinder ws inspected both visually and by the eddy current

method. The cylinder passed the visual inspection but the eddy current readings were low on

the inner wall of the cylinder and the baffles. These low readings were caused by improper

anode placement during copper plating. An eddy current reading of 75 or greater was determined to be satisfactory for this coating. Table I shows the eddy current readings for various

locations on the cylinder passes but the inside surface failed to meet the required 75 reading.

The eddy current method was being developed during this program and was not fully

understood when the cylinder was prepared. The measurement of length change for the

cylinder would be very expensive and so it was decided to measure the effectiveness of the

coating by weight gain in a humid environment. The cylinder ws kept at 95% relative humidity

at room temperature. In 269 days its weight increased from 532.363 gms to 534.118 grams

which is a weight gain of 0.33%. This is equivalent to approximately 20 microinches/inch

length increase. If the entire weight gain was due to moisture this would mean the coating was

about 75% effective. This cylinder is still being observed for further weight gain.

References:

1. Advanced Composite Applications for Spacecraft and Missiles, Phase I Final Report,

AFML-TR-71-186 March 1922.

2. Development of Design Data for Graphite Reinforced Epoxy and Polyimide Composites,

NAS 8-26198, May 1974.

56
SESSION 3

Chairman: C. 33. S. tleng

Georgia Institute Of Technology

In-Service Ultrasonic Inspection System for Composites

F. H. Chang, J. R. Bell, A. H. Gardner,

C. P. Fisher and G. P. Handley

General Dynamics, Fort Worth Division

Fort Worth, Texas 76101

Abstract

Advanced composite structures are presently used on high performance

U.S. Air Force aircraft to save weight and increase performance. These

structures are often difficult to inspect because of their complex/

multi-bondline construction and because of the general lack of experience in the inspection of composite structures in the field environment.

Ultrasonic techniques have generally been employed for the detection of

service induced discontinuities such as disbonds and delaminations.

They are not considered cost effective for inspection of large areas for

small flaws or for monitoring areas containing known small flaws for

propagation during service, primarily because of the lack of flaw position locating and permanent inspection result recording devices. An

ultrasonic technique was developed under an AFML contract to record the

position and flaw indication in a permanent manner and minimize

operator dependence to obtain a high level of inspection reliability.

The technique is simple to apply in the field environment by personnel

without advanced training. In the program, specimens representing

typical composite structures and containing implanted flaws simulating

service induced or production type defects were fabricated. Improvements were made to the state-of-the-art ultrasonic inspection techniques

to detect these defects. A follow-on program is being conducted to

establish the producibil ity of the semi -automated ultrasonic inspection

system for aircraft composite structure, based on the preliminary design

developed in the previous contract, to provide a complete inspection

system from inception to obsolescence.

59
DESIGN CONSIDERATIONS FOR GRAPHITE-EPOXY

LAMINATES OF LOW THERMAL EXPANSIVITY

Dennis D. Smith

Composite Design Group

Boeing Aerospace Company

Seattle, Washington 98124

Abstract: The development of graphi te-epoxy laminate designs for space

missions requiring structural components with near zero coefficient of

thermal expansion (CTE) is examined. The selection of fiber systems

and laminate designs is approached by evaluating CTE sensitivities to

material and process variabilities. Uniaxial and biaxial thermal stability laminates are considered. From this design approach, material

and process specification tolerances can be determined to control

laminate CTE within design objectives. Design examples are presented

to illustrate this laminate design approach.

Key words: Advanced composite materials; graphite-epoxy laminate

design; laminate design sensitivities; thermal expansion.

INTRODUCTION

The structural efficiency of recent aerospace vehicles and vehicle components has significantly benefited from the application of advanced

composite materials. In particular, graphite-epoxy has come into the

mainstream of spacecraft design activities. Increased strength and

stiffness, reduced weight, improved thermal stability, greater fatigue

resistance, and lower cost have been exhibited as inherent characteristics of this material system. For specific applications, any one

of these particular design goals may be optimized by proper

consideration of the orthotropic nature of individual graphite-epoxy

plies or lamina. To accomplish this, the composite designer combines

several such layers to develop a multiple ply laminate with the desired

mechanical and physical properties.

Designing with graphite-epoxy materials involves the weighing of

various material related compromises. For example, selection of a graphite fiber and orientation of the lamina to achieve a specific

property such as high stiffness may well prove to be detrimental to

one or more of the other laminate properties, such as CTE. This paper

considers the development of a laminate of near-zero thermal expansion

and the interactive design compromises associated with its use.

60
MATERIALS

There are many fibers presently in use in advanced composite structures.

Some of the more common of the graphite fibers fall into three main

categories: 1) high strength, 2) high modulus, and 3) ultra-high

modulus. Table 1 shows typical properties of these fiber systems.

Most graphite filaments are made from a graphitization of the organic

precursor polyarcylonitrile (PAN). A pitch precursor is still in the

development stage and shows excellent cost advantages over the PAN

precursor filaments.

The graphite fibers (except GY-70) are supplied to processors in tow

form, (small bundles of fiber) or in woven fabric form. Tows are combined with a matrix material such as epoxy and supplied to fabricators

in preimpregnated (prepreg) sheets or rolls of various widths and areal

weights (weight of fiber per unit area). The fabricators, such as

Boeing, combine the tapes into various laminate configurations and cure

the laminates. Epoxy matrix materials are thermosetting at temperatures

ranging from room temperature to 400°F. Most aerospace laminates are

cured from 250°F to 350°F in a pressurized autoclave, although oven

cured and press cured laminates are also made.

The combination of graphite fiber and epoxy matrix in prepreg form is

the building block for the designer. Typical properties of three

major graphite fiber types combined with an epoxy matrix are shown in

Table 2. These are typical values which do not necessarily represent

a specific graphite fiber/epoxy system and are intended to represent

the material type. It can be seen from the material properties that

a ply of graphite fibers and epoxy in prepreg tape form is highly

orthotropic, that is, the longitudinal and transverse properties

differ greatly. The material is generally considered homogeneous even

though the two-phase material is obviously heterogeneous. The simple

homogeneous orthotropic representation is usually necessary for

designer's (macromechanical ) laminate analysis. It will be seen,

however, that a micromechanical analysis which recognizes the inhomogeniety of the ply but ignores the internal structure of the fiber and

matrix will be required to assess specific laminate performance.

DESIGN REQUIREMENTS

When the design of a vehicle component requiring graphite epoxy is

initiated, the first task is to identify the design objectives. This

may be readily determined, but more likely only a higher level assembly

performance is known or specified. Before the individual component

performance requirements can be isolated, an investigation of component

contribution to assembly or vehicle performance may be required. Such

a study on a thermal stability critical 2.4-meter graphite epoxy

optical support (metering) structure was performed and reported in

reference 1. The metering truss structure component thermal stability

6l
sensitivity study resolved the individual member contributions of three

types of thermal deformations: despace, decenter and tilt. Figure 1 shows the truss structure and the deformation vectors. The relative

distortion figure for each major component of the trusr, assembly is

shown in Table 3. A nominal value for the coefficient of thermal

expansion (CTE) is assigned to a component and all other component CTE's

are made zero. For this linear system, this allows the total truss

thermal performance to be a function of a singular component type.

With this data, the truss tube members could be readily identified as

the critical component of the metering truss structure. Successive

iterations, as described in the reference, were made to ascertain the

nominal CTE values and allowable CTE tolerance ranges for the individual

components. The truss component thermal influence analyses were performed by utilizing a deterministic analysis approach. In deterministic

analyses, the uncertainty that exists between predicted and measured

performance is considered by the utilization of a safety factor. A

probabilistic approach, which recognizes that performance variability

is possible, is more consistent with composite materials. There are

not deterministic, single-valued properties for graphite epoxy lamina

and laminates. This does not mean that a probabilistic design approach

incorporating a Monte Carlo random-number-generator with appropriate

probability distributions functions is always required to properly

assess specific laminate characteristics. However, a designer must be

aware that advanced composite material characterization is statistical

in nature and that certain material properties may, at times, show wide

variations from nominal. The material system in lamina or laminate

form has inherent variabilities which should be understood regardless

of the fact that simplifying assumptions have allowed deterministic

analyses to quite adequately predict these properties.

Laminate Design . The designer, once he has established the graphiteepoxy component CTE, is ready to select the fiber/epoxy system and

configure the lamina plies to develop a laminate with the desired

thermal stability. As discussed earlier, the laminate CTE can be

determined by specification or deduced from a thermal sensitivity study.

Determination of a laminate form which will result in the desired CTE,

however, requires the review of potentially many candidate laminate

systems. There are two basic laminate categories: isotropic laminates

and anisotropic laminates. An isotropic laminate results in mechanical

and physical properties which are essentially constant in any plane

which passes through a point in the laminate. Such laminate properties

as CTE, strength, and stiffness are essentially constant in any inplane direction. This laminate can be more accurately described as

pseudo-or quasi-isotropic since truly isotropic material behavior

would be difficult to achieve. A symmetrical anisotropic (orthotropic)

laminate has three perpendicular planes of material property symmetry.

This laminate results in CTE, strength, and stiffness differences in

the orthogonal laminate axes. The amount of difference is a function

of the laminate design, but an extreme would be a unidirectional

62
laminate (all lamina fibers parallel). The laminate properties would

be the same as the lamina properties in such a layup.

In general terms, shell structures, cylindrical assemblies and bulkheads

that have structurally or thermally interactive components (rings,

stiffeners, etc.) are potential applications for isotropic laminates.

Tubes, rods, beams, rings, flexures and other components which do not

have direct thermal stability coupling may be fabricated from

orthotropic laminates. For some components, this distinction may be

difficult to resolve. A general design consideration is to utilize an

isotropic laminate unless there is some reason against its use. These

reasons could be to increase stiffness, strength, microyield or hygroscopic stability, to name a few.

The requirement for an isotropic laminate creates a unique design

condition and effectively dictates the design. Three laminate families

which are commonly utilized to produce in-plane isotropy as well as

out-of-plane balance are:

1) (0°/±60°) s 6 ply laminate

2) (0°/±45°/90°)_ 8 ply laminate

3) (0°/±30O/±60°790o

) s 12 ply laminate

Note: s represents symmetrical laminate

The more plies in the laminate, the closer the laminate form represents

true isotropy. The 8-ply laminate is in general use as a balance between obtaining reasonable isotropy and minimizing fabrication costs.

Symmetrical laminates are used to prevent in plane warpage due to

unbalanced thermal or mechanical strains.

Orthotropic laminates may take many forms, such as:

(0

X /*)S (0x /±ft/90)

s (Ox/±y*e*)

t

Note: -9- represents crossply of undefined angle

x represents undefined number of 0° plies

Obviously, hundreds of orthotropic laminates could be designed of

a graphi te-epoxy material, of hybrid graphi te-epoxy systems and of

combinations of graphite and fiberglass, boron, or Kevlar. The designer

needs to assess the design objectives and develop a thermally stable

laminate design which also meets strength and stiffness requirements.

In order to achieve laminate design with essentially zero coefficient of

expansion (CTE) in the plane of the laminate, different design

approaches must be utilized for isotropic and anisotropic laminates.

63
For an isotropic thermally stable laminate design, very high modulus

fibers are required. Hal pin and Pagano (2) developed a solution for the

pseudoisotropic thermal expansion coefficient in terms of the lamina

properties:

(a

a = a =

x y

l a

+

( a 1 - a2 ) (E-| - E

2

)

1 + (1 + 1

where a =a = laminate thermal expansion (PPM/ F) in X & Y directions

a

a

y

i

12

= lamina 0° thermal expansion (PPM/°F)

= lamina 90° thermal expansion (PPM/°F)

= lamina 0° modulus of elasticity (PSI)

= lamina 90° modulus of elasticity (PSI)

= Poisson's ratio

In general, this equation for CTE requires that ot

^ (longitudinal

lamina CTE) be negative and the ratio of fiber modulus to matrix

modulus be greater than 100 in order to achieve zero CTE in an isotropic laminate. As an example, laminate analyses of commercially

available graphite fiber/epoxy systems allows a comparison of pseudoisotropic laminate thermal expansion.

TABLE 4

Isotropic Laminate Thermal Expansion

Fiber in Laminate

T300/AS

HMS/T-50(PAN)/M0DM0R I GY-70/P-75S

Isotropic Laminate CTE

(PPM/QF)

1.00 - 1.20

0.13 - 0.50 ~ 0

Only the ultra-high modulus fibers E-, £ 70-75 MSI result in an isotropic

laminate of near-zero CTE. The negative fiber CTE is satisfied

( a F s -.7 PPM/°F) and the fiber to matrix modulus ratio is greater

than

h

100 (Ef/Em 5 114).

For orthotropic laminates, many different laminate configurations are

capable of achieving near-zero CTE. The screening between laminates

of this type generally considers laminate sensitivity to material

variabilities as well as cost and secondary design requirements such

as strength and stiffness.

Lamina Sensitivity . Determinate laminate analyses can establish the

nominal coefficient of thermal expansion (CTE) of a graphite epoxy

laminate. This type of analyses, however, will not resolve the influence of material and process variabilities of laminate CTE. In

order to provide a designer some visibility into laminate CTE scatter,

6k
lamina sensitivity studies are often needed. For critical components

requiring stringent CTE control, they are essential.

All graphite fibers have variabilities associated with the predominate

fiber property, which is fiber modulus. In general, as the modulus of

a fiber increases, so does the statistical scatter. Therefore a designer could expect more fiber modulus variations from ultra-high

modulus fibers than from AS or T300 (high strength) fibers. Increased

fiber variability manifests itself as greater scatter in laminate CTE,

the amount dependent upon the level of fiber dominance in any particular

laminate.

For example, a lamina sensitivity study recently completed on a thermally stable pseudoisotropic GY-70/934 (fiberite 1534) laminate will be

reviewed. The lamina sensitivity is determined by analyses which

considers the influence of a particular variable, such as fiber modulus

or fiber volume, on lamina properties. The variability of the Celanese

GY-70 fiber and its influence on lamina properties is shown in Figure

2. Figure 2 shows the relationship of fiber modulus to lamina

mechanical and thermal properties. The only significant lamina property changes due to fiber modulus variability is in the longitudinal

(0°) lamina modulus (E-|) and in the longitudinal lamina CTE. That is,

as fiber modulus increases, the 0° lamina modulus (E-j) significantly

increases. The 90° lamina modulus (Eo) is not influenced by this

variability, however. The shear modulus, Poisson's ratio and transverse (90°) lamina CTE as well as the transverse modulus are essentially

insensitive to fiber modulus fluctuations. Actual fiber modulus test

data, on material prepreged by Fiberite for Boeing shows the mean

GY-70 fiber modulus to be 74.4 x 10° PSI with a standard deviation of

2.1 x 106 PSI.

Lamina properties are also influenced by the volumetric relationship

of the GY-70 graphite fiber to epoxy matrix. Figure 3 shows how this

relationship influences lamina properties. It can be seen that lamina

properties and therefore laminate properties may be strongly influenced

by the fiber to epoxy volume ratio. All lamina properties show a direct response to changes in fiber volume. Small variations in lamina

fiber volume may result in significantly different lamina properties

than nominal properties utilized in design.

Laminate Sensitivity . The lamina property fluctuations noted necessarily influence laminate CTE. Figure 4 shows how these changes affect

the isotropic (0/±45/90) s GY-70 laminate CTE. From this figure, it can

be seen that a ± 2 MSI lamina fiber modulus range will result in a ± 0.031 PPM/°F laminate CTE variation, while a ± 2% lamina fiber

volume tolerance will vary laminate CTE over a ± 0.047 PPM/°F range.

Also shown is the laminate sensitivity to the angled ply layup errors.

The bottom chart of Figure 4 shows that a one degree error in layup of

the ± 450 ply Would result in a CTE fluctuation of 0.028 PPM/°F.

65
Figure 5 shows laminate sensitivity of an orthotropic T50 (PAN) laminate

for comparison. The laminate change in CTE due to fiber modulus varaition and angular errors in the cross ply lamina is less for the orthotropic laminate while the influence of fiber volume is about the same.

Orthotropic laminates are generally a better laminate design selection

where primary component material requirements are essentially uniaxial.

When biaxial CTE control is specified, the isotropic laminate is then

required.

In order to assess the performance characteristics of various orthotropic laminates, several sensitivity analyses would be required. This

type of sensitivity analysis, performed either on a deterministic or a probabilistic basis, becomes, along with cost and mechanical property

studies, a necessary tool for making rational design selections to

meet a given set of mechanical property requirements as well as low

thermal expansivity.

DESIGN VERIFICATION

From the laminate CTE sensitivity to the materials and process parameters, a designer can determine the material specification requirements as well as fabrication controls that must be imposed. Depending

upon the allowable laminate CTE tolerance, the designer has a method by

which to constrain only those variables which are critical and to

tolerate other variables since their influence may not be significant.

Material specifications and in-process controls can then be tailored

toward a specific low expansion laminate.

Another consideration is the method of design verification. A trade

must be made on the imposition of costly material and process controls

compared to specifying design verification test requirements. If hand

picked fibers, close fiber volume control and tight tolerances on ply

angle are utilized, it is entirely possible that these additional

program expenses could be more than offset by reduced verification

testing.

SUMMARY

A composite designer has many material and laminate options when

developing a low expansion component or assembly. A most important

step in the development of a satisfactory design is a clear understanding of the design requirements. A vehicle or assembly sensitivity

study similar to the Boeing Space Telescope truss study can be utilized

to determine laminate CTE requirements at the component level.

Once the design objectives are available, a trade of candidate laminates

investigating cost, weight, sensitivity to material and process variations, resultant mechanical properties, and joint or interface constraints can be made. Other considerations, not addressed here include

66
microyield, hygroscopic sensitivity, susceptibility to microcracking

,

fatigue and visoelestic effects. Many of these parameters are difficult

to determine analytically without the benefit of an existing data base

of many such tests.

Of primary importance to a composite designer is the knowledge that

while he is utilizing a material that is quite capable of obtaining

laminates with thermal expansions in the ± 0.025 PPM/°F range, modulus

values to 30-35 MSI, strengths to well over 120 KSI, weights of 40%

less than aluminum, and costs which are competitive with conventional

metal structure, he must always be aware of the statistical nature of

advanced composite materials. Unless a designer's "homework" is done,

the same material that has such great potential may well prove that

Murphy's law was right all along.

REFERENCES

II] Smith, D. D. and R. E. Jones, "A Statistical Evaluation of a Space

Stable Optical Support Structure," SAMPE Journal, Vol. 14, No. 5

(Sept. /Oct 1978), p 4.

[2] Halpin, J. C. and N. J. Pagano, "The Laminate Approximations for

Randomly Oriented Fibrous Composites," J. Comp. Mat., Vol 3

(1969), p. 720

67
©ISPACt

• PERFORMANCE REQUIREMENTS

KVACf - *!. OECfNTER - *|Qj| Tilt • *3 wc <n FIGURE 1 r OPTICAL SUPPORT TRUSS DESIGN

TABLE 1 : TYPICAL GRAPHITE FIBER PROPERTIES

Graphite Fiber Type

(KSI) (MSI)

High Strength AS/T300 400 32

High Modulus HMS/T-SO(PAN) 340 S3

U1tr*-Mgh Modulus GY-70 250 75

TABLE 2 : TYPICAL GRAPHITE EPOXY LAMINA PROPERTIES

Lamina Htter1*1

(KSI)

E 2

(MSI)

G

12

(MSI)

CTE

X (PPM/°F)

CTE

y

(PPM/°F)

High Strength/opoxy 21.0 1.5 .66 -.10 14.5

High Modulut/epoxy 32.0 1.3 .60 -.30 14.8

U1tr«-h1gh todului/epoxy 45.0 1.2 .70 -.54 15.8

TABLE 3 : THERMAL STABILITY SENSITIVITY TO TRUSS COMPONENTS

Truss Conponeirt^.

—TSTTnal CTE Value)

(10- 6 1n/in/ f)

JWal At

(axHy«metr1cal)

Lateral AT URtlsyBketrtcal)

Despace

(I0*1n.)

Decenter

(10-*1r..) (art-sec)

Despace

(i6-«in.)

Decenter 1 *,ut

O0-«1«.\\ (arc-sec)

Trial A

Strut ,20/0 400 • • • -444 -1.60

Ring .20/0 -40 • • * .?64 0.15

Spider 6eam .20/0 n • « • -102 0.0<

Hub .20/0 7 • • « 2 0.C1

Total 378 > 118 -280 < 394 -1.40 < 2 •NeqIIMe

68
-3 -2 -1 0 +1 +2 +3

CHANGE IN FIBER MODULUS

AEp, , MSI

G12

-3 -2 -1 0 +1 +2 +3

CHANGE IN FIBER MODULUS

AEFl , MSI

£ -.600r

I O

a

.300

3-2-1 0 +1 +2 +3

CHANGE IN FIBER MODULUS

A Ef| , MSI

^12

2

o

COo

.29C-

;.28d

-3 -2 -1 0 +1 +2 *3

CHANGE IN FIBER MODULUS

AEpj , MSI

-3 -2 -1 0 -fl *2 +3 -3 -2 -1 0 +1 +2 +3

CHANGE IN FIBER MODULUS CHANGE IN FIBER MODULUS

AEp,. MSI AEp| . MSI

FIGURE 2 : LAMINA PROPERTIES VS GY-70 FIBER MODULUS

69
50 60 70 50 60 70

PERCENT FIBER VOLUME PERCENT FIBER VOLUME

FIGURE 3 : LAMINA PROPERTIES VS GY-70 FIBER VOLUME

70
42

CTE (10" 6/IN/IN/°F) CTE (10' 6/IN/IN/°F)

FIGURE 4 : GY-70 ISOTROPIC LAMINATE FIGURE 5: T50(PAN)/T300 ORTHOTROPIC

CTE SENSITIVITY . LAMINATE CTE SENSITIVITY

71
AIRCRAFT COMPOSITE MATERIALS SELECTION AND APPLICATION

V. L. Reneau

The Cessna Company

Wichita, Kansas 67277

Abstract: Composite aircraft structure requires proof of safety and

durability similarly to metallic structure. The proofs of structural

analysis methods, reliable material properties, and of assembly methods

are similar to those required for metal structure. The effects of environment, time and fatigue damage are different than for metal and

must be compensated for when design properties (allowables ) are developed. Design allowables are not as readily available for composites as

for metals and frequently must be obtained by testing.

Key words: Composites; design allowables; aircraft.

The Citation III airplane is the aircraft on which Cessna has chosen to

introduce load carrying graphite and Kevlar fiber reinforced composite

structure. It is a 10 to 15 place jet aircraft capable of operating at

an altitude of 51 > 000 ft. Its flaps, spoilers, wing tips, engine nacelles, radome and most of the fairings are made from composite materials. Graphite and/or Kevlar are used in the flaps, nacelles and

spoilers. The environmental exposures anticipated for the airplane were

used in the tests conducted to develop design allowables for the

composites

.

The analysis of composite parts has one more phase than that of a metal

part in that the internal structure of the parts is designed and analyzed along with their external shape and size. Fiber reinforced

materials have directional properties. In some cases, it is advantageous to design the internal structure of parts so that most of the

fibers are parallel to the load directions to maximize stiffness and

strength where they are needed.

The rigorous analysis of crossplied fiberous composite materials is not

simple and requires either a computer or a lot of someone's time. In

general, tedious matrix calculations are required to obtain crossply

properties from fiber direction design allowables. Several computer

programs are available which will handle either simple or sandwich

panel structure. If a few numbers only are required the calculations

can be approximated on a programmable desk calculator.

These calculations must necessarily interact with the design calculations for overall part stress and deflection. Once calculated or

72
measured property values have been obtained for a given crossplied

composite structure they can be used in normal analysis of the part.

There are some precautions necessary to prevent twisting of laminated

parts from temperature changes or loading. For example, a crossplied

part should be assembled so that its internal structure is mirror

imaged across its center plane. If the part is a sandwich, each skin

should be balanced across its center plane and each skin should be a

mirror image of the other.

Structural composite materials are glass, Kevlar, graphite or boron

fiber reinforced resins. For some unknown reason they have been divided into glass fiber reinforced materials and "advanced composites"

which includes the rest.

They are selected in theory for a combination of cost, strength, stiffness, density and environmental suitability that will give a functional

part at the least weight and cost. Actually, the availability of certifiable property data may select the material and will certainly limit

use to a few well characterized materials.

Design allowables are reduced mechanical properties. They are obtained

from the means of the test data by subtracting a multiple of the standard deviations from them. The multiplier is usually three or four depending on the number of specimens tested and the degree of reliability

required.

This value is further reduced by factors for temperature extremes,

moisture and other anticipated environments, aging and fatigue damage.

The final design allowables are noticeably smaller than the corresponding test averages. The method is taken from MIL-HBK-5 "Metallic

Materials and Elements for Aerospace Vehicles" Section l.U, or from

"The U.S.A.F. Design Guide for Advanced Composites Materials".

The multiplier for the standard deviation is a function of the number

of specimens tested and the degree of assurance required. Ultimately,

it derives from the standard non-skewed "bell" distribution curve.

For example a "B" basis design allowable is an allowable above which

the strength, etc., of 90% of the specimens can be expected to fall 95%

of the time. An "A" basis allowable is one above which the strength of

the specimens will fall 95% of the time. An "A" allowable is the set

of properties required for primary structure. A "B" allowable is the

set of properties required for secondary load carrying structure or experimental structure.

73
The factors for calculating these reduced values can be obtained from

standard tolerance factor plots. For example if 6 specimens are

tested and a "B" allowable is required, the standard deviation of

the 6 data points is multiplied by about 3 and subtracted from the

mean of the 6 data points. If 10 specimens had been measured, the multiplier is about 2.3. For "A" allowables, the multiplier is higher.

The effects of moisture, temperature, fatigue, and exposure to fuel,

cleaning solutions, lubricating oils and hydraulic fluids were obtained

by testing several groups of coupons for each exposure. The difference

between the means of the "unexposed" coupons and those of the "exposed"

coupons was subtracted from the reduced properties obtained from the

standard deviation reductions. When all the anticipated effects of

aging and environmental damage had been subtracted the residual mechanical properties were used for design allowables.

A graphite fiber composite was selected for the aircraft spoilers and

the flap substructure because of its high strength, stiffness and long

fatigue life. The specific material was Narmco 5208 epoxy resin reinforced with Union Carbide T-300 graphite fiber. It was selected because Kenneth Hofer at the Illinois Institute of Technology Research

did a thorough characterization of it for the Air Force (T) and because the Air Force, Boeing, Lockheed and McDonnell-Douglas were using

it in major flight test articles under NASA contract. It was sufficiently characterized that design allowables could be calculated from

available data.. We would have preferred a 250° F cure rather than the

350°F required for 5208, but, the desired moisture resistance and data

were not available.

The Kevlar composite used in the Citation III nacelles and in the skin

panels on the flaps was selected because of its low density, high

stiffness and good impact resistance. Kevlar itself is less dense than

either glass or graphite and has very high tensile strength. Its compressive strength is low, especially in warm humid conditions. In

stiffness and impact strength controlled designs, such as the Citation

III nacelles, a Kevlar composite is an advantageous choice.

The specific Kevlar fabrics (Hexcel F-155 & F-l6l - Kevlar) and resin

combinations were selected partly because they were being used on the

Lockheed L-1011 and on a Sikorsky helicopter. The full gamut of test

data for calculating design allowables was not available and Cessna did

the coupon testing required.

(1) AFML TR-7U-266 "Development of Engineering Data on the Mechanical

and Physical Properties of Advanced Composite Materials" Hofer, K.E.

et. al.

7^
After components are designed and the first "units are built of composite materials with reliable design allowables, it is necessary to

assure that the parts are acceptable by component testing. This testing determines if they were cured properly, if the joining and assembly

methods are adequate and if the assembly has adequate fatigue life and

damage tolerance to operate safely over the anticipated life of the

airplane

.

For example: a Citation III prototype flap was loaded to 150$ of the

expected design load both before and after environmental exposures

simulating the service environment expected. Sixty thousand design

limit load fatigue cycles were imposed on a second unit after saturation

exposure to hot water. It also survived 150$ of design limit load.

The ability of the flap to carry such a load after environmental

exposure indicates that mechanical property reductions for design

allowables were reasonable. Finally a third unit was damaged by

successively cutting the front and rear spar caps, making a lU" saw

cut through a skin panel and finally cutting a spar loose from the

load fitting at one end. It survived a limit load test after each

injury and survived 15,000 limit load fatigue cycles with spar cap and

a 2.5" length, skin panel cut at the same place. (Each cut was repaired after the test and before the next damage and test cycle).

All this work is part of the certification procedure. The Federal

Aviation Administration (FAA.) requires that a manufacturer show by

test or analogy with a similar design that an airplane will be safe

to fly over its expected life

The details of FAA requirements peculiar to composite structure are

contained in FAA Advisory Circular #20-107, "Composite Aircraft

Structure" and in FAR 25 sections 6l3 & 853. In general it requires

proof of design allowables, static strength and safe life.

75
FIGURE 1

THE CITATION III AIRPLANE

76
-65°F TO + l60°F

WATER

OILS AND FUELS

CLEANING SOLVENTS

FIGURE 2

ENVIRONMENTAL EXPOSURE

Ep = COS^a + Ep SIN a + SIN2 2a ( E Q - 2 a0 ) Ea 1^ H (E90 )

FIGURE 3 ELASTICITY

EQUATION

77
o

o

o\ CO

00

w

PC<o

-J

<:o

oMH

<!

CO 9w

S3

CO

CO2

COao

CM

CO o

II II

COo

II CM

COwoHw

S3

PSwHH<U

CO

PSo

COQM

1=3 W

PSH

COMo

oMo

H

COonH

CO

PiPH

S3

QWU

S3ow

PSH

COwH

O LO

CN rH

Csl CM

S3HOW CO

PS

W

H

CO

CO

78
76

STRENGTH

KSI

23

13

WOVEN STYLE 281 KEVLAR

TENSILE

E = 3.4 MSI

G = 0.35 MSI

P = .05 ///IN3

COMPRESSIVE

SHEAR

TEST MEAN REDUCED FOR: SCATTER, HEAT

MOISTURE, FLUIDS

AGING, FATIGUE

FIGURE 5

KEVLAR DESIGN ALLOWABLES

79
COMPOSITE STRUCTURES IN ROTORS AND PROPELLERS

David P. Thompson, President

T M Development, Inc.

Chester, Pa. iy013

ABSTRACT

A variety of composite rotor blade and propeller designs are

reviewed. Desireable blade composite material qualities are

enumerated. Various failsafe design features are shown for

propeller blades. Multiple load path design features are

discussed. Some manufacturing and quality control aspects

of composite blade construction are discussed. Composite

material properties versus cost factors are discussed.

Key words: Composite propellers and rotors; propeller

materials requirements; blade aeroelastic behaviour; propeller fatigue behaviour; failsafe blade design; blade manu

facture and quality control.

INTRODUCTION

This paper includes a partial review of composite propeller

and rotor blade work of D. P. Thompson and John A. McAfee

and their company, T M Development, Inc. of Chester, Pa.

EXAMPLES OP COMPOSITE BLADE DESIGNS

Pig. 1 Fiberglass Main Rotor Blade

80
This was an early parallel-filament glass-epoxy prepreg

blade design flown on two experimental small helicopter designs in the early 1960 f s. The metal rootend fitting was

bonded to, and also captured by continuous filament-tape

wraparounds, step-tapered outboard. A rigid foam core stabilized the structural shell. A taper ratio of two and a

nine degree, non-linear twist distribution provided excellent hovering efficiency. Airfoil is symmetrical 632~015

section.

A set of these blades survived, undamaged, when a sudden

transmission seizure stopped the coaxial rotor from 400 rpm

in 1% revolutions! Films revealed that large whipping deflections of the blades stored energy, preventing destructive peak loads.

Fig. 2 Tail Rotor Blade

This linear filament glass/epoxy tail rotor

blade weighs only 1.2 lb. per blade. Airfoils are 23 series cambered sections. Rootend connection is similar to Fig. 1 main

rotor blade using a wrapped fitting. Rotor

diameter is 4 feet.

Fig. 3 Fiberglass Propellers, One Piece

A series of propellers was developed in the early 70' s for

81
small, home built aircraft. Diameters of 48 to 60 inches

were built for engines of 60 - 70 hp (VW, etc. ) turning at

up to 4000 rpm. These propellers were of a one-piece,

twisted-plate design format (similar to early aluminum pro

pellers). Material was solid parallel-filament E-glass/

epoxy prepreg and the process was single-stage compression

molding in matched, closed dies. An .018 stainless steel

leading edge wrap was cobonded during the molding cycle.

These achieved considerable technical success in terms of

aerodynamic performance, structural toughness and reliab.il

ity.

Fig. 4 Stacked Four-blade

Propeller

This illustrates how four

(and six blade) propellers

may be assembled by stacking

two-blade units. This was

an air-boat propeller design

to absorb 130 hp at 3500 rpm

with a 54 inch diameter.

Fig. 5 Variable Pitch Propell

Blades

These fiberglass/epoxy blades

were designed with an internal

"coke bottle" metal fitting

molded into the rootend which

mates with a pitch-change retention bearing. A foam core

filler stabilized the composit

structural shell. A formed

stainless steel leading edge

abrasion cap was molded-in and

cobonded to the blade.

82
Fig. 6a and b Main Rotor Blades-Glass/Epoxy

These straight, eleven inch chord rotorcraft blades feature

a bonded-plus-wraparound filament root retention and a

molded-in trailing-edge strut attach fitting. Airfoil is

symmetrical 13% thick section. Replacing 35 foot diameter

wood blades on a two place rotorcraft, these have demonstrated excellent dimensional stability, consistent tracking

and smooth flight behaviour with little maintenance

required. These FRP blades were designed to closely duplicate the mass and stiffness distributions of the original

wood blades so as to avoid introducing new undesired

dynamic effects.

83
Fig. 7 Turbo-Propeller Blades

Developed for, and in cooperation with, Hartzell Propeller,

Inc. are these composite, wide-chord "paddle-blades". They

are designed for efficiency at greatly reduced tip speeds

to provide quiet operation. Propeller diameters to 9 feet

have been developed, for turbo-prop engines in the 1000 hp

class, fitting standard Hartzell constant-speed hubs. Both

fiberglass/epoxy and Kevlar/epoxy designs have been built.

The Kevlar blades have been designed to very thin tip airfoil sections (t/c of 1.5%). The first Kevlar bladed propeller design has now received PAA certification and production go-ahead.

8^
24©

gO qo 160 UO JXO

Diameter, inches

Fig. 8 Propeller Weight Trends for Metal and Composite

Blades

The above chart illustrates the substantial weight savings

achievable with Kevlar blades in turboprop propellers.

Pig. 9 Mini-RPV Propellers - Keviar/Epoxy

This is a 30 inch diameter ground adjustable-pitch propeller

designed by T M Development, Inc. for an RPV 2- stroke engine

test program, absorbing 'lb hp at 7500 rpm. Hub is aluminum

85
alloy and blades are of Kevlar/epoxy, oriented filament

construction.

BLADE MATERIAL REQUIREMENTS

Aerodynamic shape factors include planform, airfoil section

progression along the blade, twist distribution and tip

shape. Ideally these factors should be selected by the

aerodynamicist and designer without undue limitations imposed by the materials and processes of manufacture. Modern

composites do, in fact, allow the propeller and rotor designer extensive freedom to configure for performance without sacrificing practical structural and manufacturing

features. This occurs partly due to the unique ability

with composites to create a considerable variety of structural density and stiffness qualities within a selected

aerodynamic envelope. Their high specific strength and

stiffness qualities are advantageous in this. Considerable

latitude is available, even after the tooling is built, to

modify the propeller structural and vibrational characteristics. This sort of modification is possible by several

means. First, by changes of type of filament and resin

system (prepreg system; for greater or less stiffness,

including possible hybrid combinations of reinforcements,

with glass, graphite and Kevlar as obvious candidates.

Secondly, by changing the fiber orientation, structural behaviour may be changed. For example, it might be desired

to soften the blade bending stiffness while increasing the

torsional stiffness. This could readily be done by increasing the proportion of biasply filaments to spanwise filaments. Third, if the propeller design embodies a lowdensity foam, honeycomb core, or a hollow cavity within the

composite structural shell, the design thickness of the

shell may be readily changed without disturbing the blade

aerodynamic envelope design. By the methods described, a

basic composite rotor blade or propeller design, within a

fixed external envelope, can be adapted to have a variety of

structural, aeroeiastic and/or mass characteristics.

FATIGUE PROPERTIES OF COMPOSITE BLADES

Starting in the 19bO f s, laboratory tests began to reveal the

remarkably favorable fatigue properties of properly designed

glass filament/epoxy reinforced plastic blade structures.

These favorable fatigue properties are even more strongly

evident in today's composite blade structures, notably in

Kevlar/epoxy applications. These properties include: (1)

Non-notch sensitivity due to multiple filament load paths,

with related zero or slow propagation of structural damage,

86
(2) Appreciable internal damping which tends to limit peak

values of resonant vibration stress levels when they occur,

(3) Absence of the stress-corrosion-fatigue destructive

sequence sometimes experienced by lightweight metal structures, (4) Visual evidence of fatigue damage which eases

inspection techniques and permits "failsafe" operational

characteristics, (5) Generally high fatigue-strength to

weight ratios as compared with traditional blade materials.

MANUFACTURING & QUALITY CONTROL CONSIDERATIONS

The primary manufacturing method selected for these products

was compression molding to positive stops. In this molding

approach, the molding pressure is, in part, a function of

the layup thickness, hence carefully designed and accurately

performed layups are required.

An air conditioned clean assembly room is necessary to provide controlled conditions for blade layup.

Individual blade logs are kept by blade serial numbers and

include element weights, material batch numbers, recorded

cure cycle charts and test results from co-molded test

coupons.

FAA process and facility approvals must be obtained in order

to manufacture production run blades under an Approved Type

Certificate. Process and quality control manuals are required to be prepared and observed. All materials are purchased under strict control specifications.

FAILSAFE BLADE DESIGN FEATURES

As described above, certain inherent "failsafe" tendencies

exist in some composite filament reinforced-plastic structures. Additional failsafe benefits are achieved by designing the blade with alternate load paths. This is accomplished in the blade rootend retention area by allowing for:

(1) High quality adhesive bonding of the composite blade

shell to a metal rootend fitting, plus, (2) Mechanical

"capture" gripping action of blade shell onto the root fitting by means of a flared "coke bottle" effect in a round

type (propeller) fitting, or by wraparound of a flat block

type (rotor blade) fitting by the spar filament tapes, (3)

Reinforcement of fitting mechanical capture by filament

overwraps locally.

Outboard of the critical rootend connection, failsafe behaviour is obtained by designing the basic blade to moderate

87
stress levels with a spread-out distribution of the primary

tensile filaments. Thus, even a substantial sized damage

tends to leave enough material to prevent catastrophic

failure. Then, the tendency for slow damage propagation

allows the blade structure to survive until inspection and

maintenance procedures correct the problem.

COMPOSITE MATERIAL PROPERTIES VERSUS COST FACTORS

Generally, the blade designer must attempt to fulfill rotor

or propeller blade performance and safety requirements at

minimum cost. This would dictate the use of the least

costly composite material system for any specific blade

design program.

As the structural and/or aeroelastic rigorousness of the

blade specification increase, so the composite material

selection may be forced into the more expensive materials.

In some propeller designs, reduced blade weight with the

more exotic fibers will partly alleviate the higher cost per

pound of material. Helicopter rotor blades, however, more

usually demand discrete mass properties to control coning

angle and blade dynamics.

If an E-glass/epoxy prepreg meets the technical requirements

of a particular blade design it is the obvious choice at,

perhaps, $ 5.00/1 b. However, in a higher performance application, an appreciably greater material stiffness/weight

ratio might prove desir&ble to avoid flutter and divergence

tendencies in an unusually thin blade configuration. In

such a case the appropriate selection might be a Kevlar/

epoxy prepreg at, perhaps, $ 25.00/lb. , or some ingenuously

engineered hybrid combination of E-glass with graphite

fibers at $ 50/1 b. to give an "optimum" product performance/

cost factor.

Also, the straight material performance/cost factors must

be biased by the influences of material selection on blade

process costs. Handling and layup manhours will differ for

the various materials. Probable scrap and spoilage rates

may differ. Potential differences in product development

and certification costs due to the perceived risk level of

material "unknowns" must also be taken into account when

planning an advanced blade program.

88
SESSION 4

Cliairman: cJ. «J. Bcialdone

Gi-ocicieLrci Space IFliglit Centex*
!
DEVELOPMENT OF A LIBRARY MODULE FOR THE

ANALYSIS OF ADVANCED COMPOSITE MATERIALS

Karyn T. Knoll

Non-metallic Materials Division

Air Force Materials Laboratory

Wright-Patterson AFB, Ohio 45433

Abstract: Because of their orthotropic nature, advanced

composite materials present some special problems to the engineer. The relation between the elastic constants of an

orthotropic material is more complex than that of an isotropic material. In addition, composites exhibit coupling

between normal and shear deformations as well as between

bending and stretching load responses. The more complicated

behavior of advanced composites is usually handled using

techniques of matrix mathematics, but the required bookkeeping often overwhelms the isotropic-trained engineer who is

used to working with fewer terms and less complicated equations. This paper discusses the development of a module, designed to be used with the Texas Instruments TI-58 and TI59 desktop calculators which can be used to handle much of

the required "bookkeeping". Given such input as the elastic

constants of a single ply of composite material, the stacking

sequence of a laminate, and the applied loads (or strains)

,

the module will output desired information such as the physical properties of a laminate, the strains (or loads) in

both the whole laminate and the individual plies, and the

values of the laminate properties as the laminate is rotated.

The module will also provide "strength ratios" for a particular laminate and load vector to indicate how close the material is to failure. It is hoped that this module will make

the analysis and use of advanced composites seem less formidable to both research and developmental engineers.

Key words: Advanced composite materials; composite materials

analysis; orthotropic materials; programmable calculators;

solid state software.

The orthotropic behavior of advanced composite materials can

be described using techniques of matrix mathematics. However, the bookkeeping required by such techniques can soon

become overwhelming. The Air Force Materials Laboratory at

Wright-Patterson Air Force Base has recently completed a set

of calculator programs capable of handling much of the

91
required bookkeeping. These programs are presently being

fabricated into a library module which can interface with a

standard TI-5 8 or TI-59 calculator. (A library module is a

solid state, plug-in set of prewritten, read-only programs.)

See Figure 1. This paper discusses the structuring and potential applications of the composite materials module.

The module consists of nine user-accessible programs which

calculate material properties, initial stresses and strains,

mechanical properties, strength ratios, and translated moduli for both single plies and laminates of a composite material. Because the calculator's data storage registers are

unaffected by the call for a program, the results of one program can easily be used as input for another. Several of

the programs in the composite materials module are designed

to be used in conjunction with one another.

Program 1 of the composite materials module provides data

which is essential to all remaining programs. This program

partitions the calculator to insure the proper number of

data storage registers, clears the calculator memory, and

performs a test to determine whether or not the calculator

is connected to the PC-100A printer (since all programs

have been designed to work on or off the printer) . It then

allows the user to input the engineering constants - the

longitudinal, transverse, and shear moduli and the major

Poisson's ratio - and the thickness of a single ply for the

desired material. Program 1 also allows the user to convert

his input data into English or SI units. This conversion

can be performed on an individual entry or on the entire

list as desired. Upon completion of input data entry,

Program 1 calculates the reduced stiffness moduli (Q's), the

compliances (S's), and the invariants required to calculate

rotated values of moduli for a single ply. These results

are then used by the remaining eight programs in the module.

Programs 2 and 3 deal with initial stresses and strains induced by changes in temperature and moisture concentration.

Program 2 is concerned with stresses and strains in a single

ply of composite material. The user inputs four environmental coefficients - one each in the longitudinal and transverse directions for thermal and for moisture effects. Again,

these values can be converted into English or SI units either

singly or as a list. Next, the user must input temperature

change, moisture concentration, and curing strains. The

program will then calculate initial stresses and strains for

the ply. Program 2 can provide the user with single-ply

initial stresses and strains in both the ply-axis system and

in a coordinate system rotated thru any angle from the ply

92

axes. Program 3 is used to calculate non-mechanical stresses and strains in a laminate. Non-mechanical stresses and

strains are caused by the fact that plies at different orientations want to expand different amounts but are prevented from doing so by their presence in the laminate. See

Figure 2. After the user inputs the stacking sequence of

the laminate of interest, the program calculates the nonmechanical loads and strains for the given laminate. Program 3 is also capable of providing the stresses and strains

in each ply of the laminate in both the ply- and laminateaxis systems. The initial strains from Program 2 as well as

the non-mechanical strains from Program 3 are retained

thruout the execution of Programs 5 and 6 so that they are

available for use in the calculation of the strength ratios.

Program 4 calculates the mechanical properties of a single

ply of composite material. This program calculates both the

on- and off-axis values of the stiffness moduli and compliances. Rotation of stresses and strains is also accomplished by Program 4. In addition, multiplication of stiffness moduli and strains or compliances and stresses can be

performed for both the on- and off-axis cases.

Program 5 is used to handle in-plane properties of laminates.

Elements of the in-plane moduli (A-matrix) and their inverses can be calculated for any given laminate (symmetric or

unsymmetric) . Program 5 also provides values of the A- and

inverse A-matrices rotated from the original laminate axes

thru any desired angle. For symmetric laminates, this program can be used to multiply the A-matrix and the in-plane

strains or the inverse A-matrix and the applied loads. Program 5 can also provide the stresses and strains in each ply

of the laminate in both the ply- and laminate-axis systems.

Program 6 is similar to Program 5 except that it deals with

the flexural rather than the in-plane case. This program

can be used to calculate elements of the flextiral moduli

(D-matrix) and their inverses for any given laminate

(symmetric or unsymmetric) . Values of the D- and inverse

D-matrices rotated thru any desired angle from the original

laminate axes can also be calculated. For symmetric laminates, Program 6 can be used to multiply the D-matrix and

laminate curvatures or the inverse D-matrix and the applied

moments. Stresses and strains in each ply of the laminate

in both the ply- and laminate-axes systems at any given distance from the laminate midplane can be calculated by Program

6.

9h
Uncured ply - no environmental effect

ooooooooooo

When heat or moisture is applied, the

expands. The amount of expansion is

related to the ply orientation.

For example

:

Zero-degree ply wants to expand an

amount equal to e

90-degree ply wants to expand an

amount equal to e^

€ -H

oooooooooo

However, when the two are placed

together in a laminate, each acts

on the other and they "compromise"

so that both expand an amount gqual

to the non-mechanical strain £

FIGURE 2: Non-mechanical strains in a laminate

> Hew /jxis

FIGURE 3: Translation of moduli by a distance "d'

95
Program 7 calculates values for the coupling moduli (B-matrix) and their inverses for unsymmetric laminates. This

program can also rotate the values of the B- and inverse Bmatrices thru any desired angle from the original laminate

axes. Although Program 7 does not perform any multiplication, its output can be used as input for programs (written

by the module user) dealing with unsymmetric laminates.

Program 8 uses the strains generated in previous programs

(both mechanical and non-mechanical) to calculate values

for the strength ratios of each ply in a given laminate.

The user should use Program 1 to provide the necessary material properties, Programs 2 and 3 to calculate non-mechanical strains, and Program 5 for in-plane mechanical loads

or Program 6 for bending. It is then necessary to input

ultimate strengths for the material of interest. Program 8

allows the user to convert the input ultimate strengths into

English or SI units either singly or as a complete list.

After all of the ultimate strengths for the material have

been input, the values of the F's and G's - coefficients

for the strength ratio equation - are calculated and output.

Program 8 then calculates the strength ratio for any ply of

the given laminate when supplied with the ply orientation

(for the in-plane case) or the ply orientation and distance

of the ply from the laminate midplane (for the bending

case). When the non-mechanical strains are negligible, the

user may skip their calculation and proceed directly from

Program 1 to Program 5 or 6.

Program 9 calculates the values of the A- , B- , and D-matrices and their inverses when translated a distance, d,

from the midplane of the laminate. See Figure 3. After

using Program 1 to calculate the needed material properties,

the user inputs the stacking sequence and the distance from

the midplane of the laminate. Program 9 then calculates the

translated values of the A- , B-, and D-matrices using the

following relationships:

A. . = A . . id i:

B . = dA . . + B . . ID ID ID

D. \ = d

2

A. . + 2dB. . + D .

.

ID ID ID ID

where the primed variables ai ^nslated values and

96
unprimed variables are those which have been calculateu from

the laminate midplane. Program 9 will also invert the translated matrices. This program can be used to calculate moduli for stiffeners or sandwich beams.

To demonstrate the use of the module, I'll describe a sample problem done with the programs used in the development

of the module. Figure 4 illustrates the keyboard of a standard TI-59 calculator. (A TI-58 is the same except that there

is no "write" command) . Figure 5 shows the output produced

by the PC- 10 OA printer. (The symbols typed next to the

printed output have been added for this explanation. They

are not produced by the PC-100A printer) . If this problem

were run off the printer the numerical answers would appear

in the display of the calculator and the user would be able

to step thru them using the Run/Stop (R/S) key.

The problem solved in Figure 5 is to find the mechanical

stresses in the center of each ply Cin both the ply- and

laminate-axes systems) in a I0./+45] s laminate of AS/3501

when a bending moment of 1000 Nm is applied in the 1-direction. After inserting the composite materials module and

turning the calculator on, the user should press 2nd ENG

to put the calculator in engineering notation. If another

type of output (e.g. FIX 3) is desired, the user may set

the calculator in the desired notation and the numerical

output will be printed Cor displayed) in the desired notation. However, some of the labels may be incorrectly

printed if the calculator is in other than engineering notation.

To solve the sample problem, the user must first press 2nd

PGM 01 to call the first program. This program allows him

to input the engineering constants and ply thickness for

AS/3501 and to calculate the material properties which will

be required by the moment/curvature program. After calling

Program 1, the user presses E which partitions and clears

the calculator memory, tests for the printer connection,

prints "ENT" (to tell the user to enter his data) , and

places a "1" in the display to inform the user that the calculator is ready for the first engineering constant (El) . The user must then key in the value of EL for AS/3501 and

press R/S to enter this value. The calculator then prints

the value of EL and places a "2" in the display indicating

that it now requires the value of the second engineering

constant (Em) . The user continues entering the engineering

constants and the ply thickness for AS/3501. Following entry of the fifth value, the calculator advances the paper

97
DISPLAY

A' B' c D' E'

A B c D Ld

log CP

2nd INV lnx CE CLR

PGM P+R sin cos tan

LRN x%t

2

X 1/x

Ins CMs Exc Prd Ind

SST STO RCL SUM

X

y

Del Eng Fix Int |x|

BST EE ( )

Pause x=t Nop Op Deg

1 GTO 7 8 9 X

Lbl x>t E + X Rad

SBR 4 5 6 r -

|

Stflg D.MS Grad

| RST 1 2 +

Write Dsz Adv Prt List

R/S 0 +/-

FIGURE 4: Keyboard of TI-59 calculator

98
PROGRAM 1

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figure 5a: sample problem

99
PROGRAM 6 PROGRAM 6

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6. 00 N 2. 5 00 z

0. 0 0

~ "T •_' « 0 0

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n n r

i

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M(2)

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d(ll)

d(22)

d(12)

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M(l)

K(2)

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FIGURE 5B:

SAMPLE PROBLEM

E

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82. 127254-03 €{1)

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100
PROGRAM 6 PROGRAM 6

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1 6 . 3 2 9 3 9 9 6 - 0 3 £ ( 1 )

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2. 77299 09 Q(12)

4. 4816 09 Q(66)

0. 00 Q(16)

0. 00 Q(26)

Q

128. 24325 09 Q(ll)

11. 09196 09 Q(22)

b- . i I C 99• Q(12)

4. 4316 09 Q(66)

0. 00 Q(16)

0. 00 Q(26)

E'

b. ^yyyyb -u:

1 8 . 8 9 7 0 14- 0

i -_• i_ O u_ 7 j. u

£(D ' £(2)'

£(6)'

E

4

6. 299004

2. 1097999

•24. 442097 -i i

€(1)'

£(2)'

£(6) ?

364. 104

227. 15629

ih ad)

ih a(2)

•

328. 61911 06 °"(6) '

L

1 £ cr tr - • y l y i 4 u . y 6 y y 9 y

-109. 5397

06 cr(l) ' 06 c(2) ? 06 CT(6) '

L

874. 24925 06 a (1)

217. 01 104 06 q- (2)

•3 18. .47335 06 <T(6)

L

317

-* lit":

• I I• II I

• j q o •* q 0 A

CCD

ad)

a(6)

FIGURE 5C: SAMPLE PROBLEM

101
iind awaits further instruction. If the user wishes to convert the units of any entry, he must key in the entry and

press C (to convert from English to SI) or 2nd C (to convert

from SI to English) before he presses the R/S key; to convert all five entries, the user should press C or 2nd C after

entering the fifth value. The calculator will then print

out the converted values for all five input values. Upon

completion of data entry for AS/3501, the user presses key

A and the calculator calculates and prints out the reduced

stiffness moduli (Q) , compliances (S) , the U's (constants

required for the calculation of off-axis values of Q) , and

the invariants (I)

.

After using Program 1 to calculate the required values of

the material properties for AS/3501, the user presses 2nd

PGM 06 to call Program 6 which deals with mechanical bending

of a laminate. To enter the laminate's stacking sequence,

the user presses A and a zero is displayed. The user keys

in the total number of plies in the laminate (in this case,

6 ) and presses R/S. The total number of plies is printed,

the paper is advanced, and a "6" is displayed to indicate

that the ply orientation of the sixth (top) ply is required.

The user keys in a "0" and presses R/S to enter this value.

The zero is printed and a "5" is displayed indicating that

the fifth ply orientation is now needed. The user continues entering all ply orientations in the laminate (as the

display counts down) . Upon entry of the last angle, the

calculator advances the paper and awaits instruction.

To enter the applied bending moments, the user presses SBR

STO D and a "1" is displayed. The user keys in 1000 and

presses R/S to enter the value of M]_. A "2" is then displayed to indicate that the value of M2 should now be entered. Following entry of all three bending moments (Mi,

M2

, and Mg) , the user may press SBR RCL D to output the

stored values thereby allowing him to check his input

data.

Once the stacking sequence and bending moments have been

entered, the laminate curvatures can be calculated by

pressing SBR RCL SBR INV B SBR X SBR RCL D. SBR RCL SBR

INV B calculates and outputs the values of d^j (the inverse of the D-matrix) . SBR X tells the calculator it

will be performing a matrix multiplication. SBR RCL D

outputs the stored bending moments. Upon completion of

this sequence, the matrix multiplication dj_jMj, which results in the values of Kj_, is automatically performed. The

user then presses SBR RCL E to output Kj_. If the user

102
wishes to simply perform the matrix multiplication d^jMj

without outputting the values of dji and M-: , he can omit the

3BR RCL before SBR INV B and before D.

Once the laminate curvatures have been calculated, the stresses and strains in each ply of the laminate (in both the plyand laminate-axis systems) can be found. Since all ply

stresses and strains are calculated in the same manner, only

the -45 degree calculations (Figure 5C) will be explained

here. To describe a desired ply the user presses 2nd C and

a zero is displayed indicating that the ply's orientation

is required. The ply angle should be entered using the R/S

key and a "1" will be displayed to tell the user that a

value of Z (distance from the laminate midplane to the point

of interest) is needed. Since we are interested in the

stresses and strains in the center of the upper -45 degree

ply, a 1.5 (the value of Z in units of ply thickness) is

keyed in and R/S is pressed. Now that the ply of interest

has been identified, the ply strains may be recalled by

pressing SBR RCL E for ply strains in the laminate axes and

SBR RCL 2nd E for ply strains in the ply axes. Ply stresses

can be calculated by multiplying the on-axis values of Q

and the ply strains in the ply-axes system. This multiplication is performed by pressing SBR RCL C SBR X SBR RCL 2nd E.

The resulting ply stresses can be output by pressing SBR

RCL D for ply stresses in the laminate axes and SBR RCL 2nd

D for ply stresses in the ply axes. Once the stresses and

strains in the -45 degree ply have been output, the user

may calculate the ply stresses and strains in any other ply

by pressing 2nd C, identifying the new ply, and repeating

the sequence. The results of all three ply orientations of

our sample laminate are shown in Figures 5B and 5C. Once

again, if the user did not want to output the values of Q-j_j

and £j when he performed the matrix multiplication to calculate the stresses, he could have omitted the SBR RCL before C and before 2nd E.

As can be seen in Figures 5B and 5C, the symbol "L" is used

to identify force-related parameters (stress, loads, and moments) and "E" is used for displacement- related parameters

(strains and curvatures) . This symbology is used in all

programs in the module. Also, in programs dealing with

laminates, primed values indicate "ply axes" and unprimed

values mean "laminate axes" where stresses and strains are

concerned

.

Though calculation of the ply stresses in the sample problem

may seem complicated the first time thru, Figure 6

105
KEYS NOTES

2nd PGM 01

E

2nd PGM 06

SBR STO D

SBR INV B SBR X D

SBR RCL E

2nd C

SBR RCL 2nd E

SBR RCL E

C SBR X 2nd E

SBR RCL 2nd D

SBR RCL D

Calls Program 1

Enter E^, E^ ^,0 Lf \

(Use keys C and 2nd C for unit conversion)

Calculates Q, S, U, I

Calls Program 6

Enter laminate stacking sequence

Enter M , M , Mg

Multiplies d. .M. (Result: k.

)

ij 3 1

Recalls curvatures (k.

)

1

Enter ply orientation and z

Recalls ply strains (ply axes)

Recalls ply strains (laminate axes)

Multiplies Q. .£\ (Result: CT\ )

Recalls ply stresses (ply axes)

Recalls ply stresses (laminate axes)

FIGURE 6: Summary of solution to sample problem to find stresses

in the center of each ply of a [0/-U5] laminate to

which a moment (M =1000 Nm) is applied

10k
illustrates that the problem solution is actually quite

straightforward. To simplify program use, labels identifying the user-defined keys will be provided for each program.

A user who has become familiar with the composite materials

module can calculate ply stresses in each ply in less than

fifteen minutes. When I solved the same program using a

non-programmable calculator, the solution required several

hours and much more paper. In addition, the possibilities

for error were limitless.

As mentioned earlier, the programs for the composite materials module have been turned over to Texas Instruments

Incorporated for fabrication of the module. The library

modules are scheduled for delivery sometime in September

of 1979.

The Air Force Materials Laboratory foresees several potential

applications for the composite materials module. Though

computers will still be needed for detailed analyses, the

calculator/module combination can provide the engineer with

a fast, economical way to perform preliminary calculations

which can give him a "feel" for various trends. Armed

with this information, he can then make more efficient use

of his computer resources.

By relieving the user of the necessity of having to set up

all of the necessary calculations on his own, the module

enlarges the group of personnel able to perform composite

material calculations. Calculations which were once performable only by a small number of highly paid engineers or

consultants can be done by junior engineers or technicians.

Figure 7 lists some of the calculations carried out by the

composite materials module.

Once the number of people able to do composite material analyses has increased, the need for many of the expensive tests

being done today in order to obtain empirical results for

use in design specifications may decrease. The module may

also provide a quick and easy means for technicians to check

test results as testing is being done. In addition, engineers will have a readily available method for obtaining information concerning proposed designs during the course of

their planning sessions.

Finally, the module can be an aid to a short-course in composite materials. By relieving students of the need to

laboriously perform all of the long and somewhat tedious

105
m = l/(l-(v

LTrEL

/E

T

) Qu

= mE

L

Q10 = vTTQ LTX 12 22

Ql6

= 0

h - ur

u

s h ' u

s R

l ' l u2

I R

2 = |U3

Q22

= ma,

^26 = 0

u

l

U, =

= OjAT+Oj,c+e,p

°L

= Qll

e

L+Q12

eT

°- = Ql2e

L

+Q22e

T

(1/8)(3Q11+ 3Q22+2Q12+ 4Q66 )

(1/2)(Q11-Q

22)

(1/8)(Q 11+Q

22 -2Q12-4Q

66 )

(1/8) (Q n

+Q22

+6Q12 -4Q

66)

(1/2) (Ur

U

4

)

< pXa+^ia

< - -^2A

ji

11

= i/(x

L

x^) G

ll

F

22

= 1/CXjXJ.) G

22

F

12 = Ff2

v^ll!?22G

12

F

66

= i/(xLT) z G

66

F

l

G

l

F

2 = l/Kp-l/XJ, G

2

a = G

66 (£6

} +G11^1^

2 + 2G

12 (

b

= G

1 (e^)+G

2

(e^+2G

F

llQjl+2Fl 2QllQl2+F22

(5j2

-Qi

7

+2Fi2Qi2^22+F22^22

2

11X11X12

+F

12Q11Q22

+F

12Q12

+F

22X12X22

= F-Q7

lly12

66x66

= FlQn+F

2Q12

= ^12^22

N

M, M.2

N Ms „ N

, or, r Ms r N >., nn f Ms r Ns

+ 2G

22

(e

2) (e

2

-e

T

)+2G

66

(e

6

)(e

6

)

11', (e^-eT ) 'V

nr ( N s r N s ^ , N s.2^ f Ns2^,N s 2G

12

(e l"

e

L)(£ 2" eT

)_G

22

(e 2" e

T

) " G

66 (£

6

} " G

l

(e r

e

L

}

N

1 "L^ ^2

S

M

_= (b/2a)-/(b/2a)

2

-(c/a)

S^

1 = (b/2a)+/(b/2a)2

-(c/a)

X

il

X

22

X

12

X

66

X

16

X

26

A! . 13

U

1

V

0

+U

2

V

1

COS 2<J)+U2V

2

S in2(i)+U

3

V

3

C0S 4<|)+U3

V

4

s in44>

U

1

V0~ U

2

V

1

cos 2(^ _U

2

V

2

S iri2(J)+U

3

V

3

C0S 4(})+U3V

4

s in4^ U

4

V

Q

-

U

3

V

3

cos 4cb - U

3

V

4

sin4cf>

U

5

V0 " U

3

V

3

cos4*

" U

3

V

4

S 1x1 ^

$U

2

V

2

cos2<|)+32U2V

1

s in2<()- U

3

V

4

cos 4<f>+U3V

3

sin4cf>

= - yJ

2

V

2

cos 2^+^U^s in2c})+U3V

4

cos 4<j) - U^sin4<j)

= A.. B!.=dA..+B. D.' . = d

2

A. .+2dB. .+D. . 13 !J !] !J 13 !J 1J

1 = _ 1

Q

11

'22

12

^66

U

l

+U

2

+U

3

uru2+u

3

U4"U

3

VU

3 = 0 <26

FIGURE 7

106
matrix mathematics required to solve composites problems,

more time can be devoted to grasping concepts without

sacrificing the valuable experience of obtaining numerical

solutions. The possibility of arithmetic error in the calculation of results is also greatly reduced.

It is hoped that with the aid of the composite materials

module, the orthotropic material calculations which right

now seem so formidable to many will eventually become as

familiar and straightforward as those for isotropic

materials.

ACKNOWLEDGEMENTS

The author gratefully acknowledges Stephen W. Tsai, Frank

Huber, and Marvin Knight of AFML, Roy Chardon of Kirtland

AFB, and the personnel of Texas Instruments in Lubbock,

Texas for their aid and advice in the development of the

composite materials module.

REFERENCES

1. "Personal Programming", Texas Instruments Incorporated

(1977)

.

2. S. W. Tsai and H. T. Hahn, "Composite Materials Workbook", Technical Report AFML-TR-78-33 (1978).

3. S. W. Tsai and H. T. Hahn, "Introduction to Composite

Materials, Volume 1: Deformation of Unidirectional

and Laminated Composites", Technical Report AFML-TR78-201, Volume 1 (1978).

4. S. W. Tsai and H. T. Hahn, "TI-59 Magnetic Card Calculator Solution to Composite Materials", Unpublished

(1979)

.

107
STRUCTURAL SMC -- MATERIAL, PROCESS AND PERFORMANCE REVIEW

Ralph B. Jutte

Owens-Corning Fiberglas

Technical Center

Granville, Ohio 43023

Abstract: Sheet molding compound (SMC) has been used for functional or

fascia automotive parts such as front and rear panels for several years.

When properly formulated, processed, and molded, SMC has a potentially

substantial role in the manufacture of more structurally demanding

automotive parts.

Through variations in the materials, equipment and processing techniques, different types of SMC can be produced. These are random fiber

SMC (SMC-R); continuous fiber SMC (SMC-C); and directional fiber SMC

(SMC-D).

Representative formulations show how static mechanical properties

are dependent upon glass content, glass orientation, and on the test

temperature

.

Based on equivalent performance, these systems are lighter in weight

than steel and are generally less costly, on a material basis, than

al uminum.

Key Words: Glass fiber reinforcement; polyester resin, structural

sheet molding compound.

The increasing costs of energy and the drive to conserve energy from

conventional resources have caused major changes in materials technology. Attention is now being given to materials, processes, and

products which combine low energy for manufacturing and reduced

energy of operation.

This attention has caused design engineers and FRP materials suppliers

to extend the use of FRP from typical applications in which appearance

is the prime requirement to so-called "structural" applications, where

parts are required to carry loads in addition to those imposed by

fasteners and their own weight. Coupled with this extension of FRP

use is the added requirement that designs in FRP be of maximum efficiency; that is, that the parts most efficiently utilize the properties of the material.

108
Owens-Corning Fiberglas has undertaken a program to address these

issues. We have developed an easy-to-manufacture , versatile family

of structural sheet molding compound materials, denoted SMC-R, SMC-C/R,

SMC-D, and SMC-D/R, and have committed significant effort to characterizing these materials relative to their engineering mechanical

properties. Additional work is underway to develop constitutive relationships among composition, process parameters, molding conditions,

and engineering properties.

This report will define the OCF structural SMC materials, identify the

advantages of each, and report on static mechanical properties as well

as process methods.

SMC-R

SMC-R (Figures 1 and 2) is defined as RANDOM FIBER SMC. It contains

reinforcements with a fiber length of three inches or less in a random pattern in two dimensions. Nominal glass content is denoted by

the number after the letter R, e.g., SMC-R50 would contain 50 percent

glass fiber by weight. Most SMC used today is SMC-R. The usual reinforcement is chopped roving, although chopped strand mat also is suitable. The reinforcement can be cut to the uniform lengths, or a combination of lengths can be used. The usual length used in industry

today is one inch. Although nominal glass content for structural

applications ranges from 30 to 70 percent by weight, the majority of

interest is at 50 percent or higher glass content.

SMC-R

Random

Fiber SMC

SMC-D

i Directional

Fiber SMC

SMC-C

Continuous

Fiber SMC

Fig. 1 - SMC type designates reinforcement configuration used

109
SMC-R SMC-C/R SMC-D/R

Fig. 2 - Combination of fibers can be used to expand basic systems.

• Materials - a variety of resins such as polyesters, vinyl esters

and epoxies have been evaluated and used during prototyping. This

report covers data based on the Owens-Corning Fiberglas (OCF) E980

polyester resin, a high-reactivity isophthalic polyester resin

specifically selected for use in structural SMC. Low shrink additives

can be incorporated into the resin for such applications as wheels

where superior dimensional and warpage control is desired. In the

majority of structural applications, surface appearance characteristics

and dimensional control are adequate without the use of low shrink additives.

Glass roving at glass concentrations below 45 percent can be either

OCF 951 or OCF 433. Above 45 percent, OCF 433 glass is recommended

because its more soluble size system allows it to be processed into a denser, more uniform sheet. The greater f il amentation associated with

this product allows better glass distribution and carrying characteristics during molding at higher glass contents. Usually, the more

highly soluble systems also give better strength but lower impact

resi stance

.

All the above fibers are E glass. Other candidate reinforcements include high-strength S-2 Glass®, graphite, and aramid.

The wide variety of SMC catalysts, promoters, and mold releases used

in present SMC production parts are applicable for use in structural

SMC.

Fillers, used as extenders to reduce costs, normally are added at

levels as high as tolerable. The highest loading ratios can be

achieved with ground limestone (CaCO x it is commonly used for

applications that do not require fl rdancy. Specific resin

to filler ratios depend largely on the glass content and process

no
considerations; glass wet-out on the SMC equipment being of greatest

importance. Typical filler to resin ratios are 1.5 to 1 at 30 percent glass content and 0.5 to 1 at 50 percent glass content. No filler

is recommended for the 65 percent glass formulation.

A variety of alkaline earth metal oxides and hydroxides can be used to

thicken SMC paste. Selection of a specific thickener depends on the

resin type, filler content, and desired matured viscosities. As filler content decreases, thickening rates become slower and viscosities

plateau at a lower level. At lower filler levels, magnesium oxide

(MgO) thickener works well. It should be carefully stored prior to

use or used in slurry form to prevent hydrolysis.

Specific formulations for SMC-R50 and SMC-R65 are shown in Table A.

SMC-R50

Resin Paste

Resin OCF-E980 100 Parts

Inhibitor Benzoquinone .0005 phr

Catalyst T-Butyl Perbenzoate 1 phr

Mold Release Zinc Stearate 2.5 phr

Filler Calcium Carbonate 50 phr

Thickener Magnesium Oxide 1.5 phr

Compound

Resin Paste - 50 Parts Glass - OCF 433AB - 50 Parts SMC-R65

Resin Paste Resin OCF-E980 100 Parts

Inhibitor Benzoquinone .0005 phr

Catalyst T-Butyl Perbenzoate 1 phr

Mold Release Zinc Stearate 2.5 phr

Thickener Magnesium Oxide 1.5 phr

Compound

Resin Paste - 35 Parts Glass - OCF433AB - 65 Parts Table A

• Processing - conventional SMC machines are used to make SMC-R. A

paste containing the resins, fillers, catalyst, mold release and

thickening agents are uniformly metered onto the carrier films

(Figure 3).

Care must be taken to produce an SMC product which thoroughly wets

out at the higher glass contents. This may require a slower line

speed and a reduced weight per square foot sheet than used at the lower

glass contents. At the 60 to 70 percent glass range, sheet weights

up to 20 ounces per square foot have been produced at line speeds to

10 to 20 feet per minute.

ill
Chopper

Continuous

Strand Roving

Resin/Filler

Paste

Chopped Roving -

,,Vx

Resin/ Filler Paste _

Carrier Film

Chain Link Compaction Belt

Carrier

Film Take-Up Roll

Fig. 3 - SMC-R production process

Molding conditions vary with part complexity. In general, 800 to 1500

pounds per square inch (psi) molding pressures are used. Mold temperatures are typically 290 to 310°F. The higher glass content

materials mold quite satisfactorily under the same conditions, commonly used to mold SMC-R30 compounds.

The maturated paste viscosity of SMC-R30 polyester compounds is

commonly 40 to 60 million centipoise as measured by a Brookfield HBT

Viscometer. SMC-R50 and R65 compounds (Table A) are generally molded

at a lower viscosity. A paste viscosity of five to twenty-five million

centipoise appears satisfactory. Lower viscosities result in a tacky

sheet and separation of the resin and glass during molding. Higher

viscosities require proportionally higher molding pressures.

• Advantages and Limitations - SMC-R has the greatest molding versatility. Parts of varying cross sections including ribs and bosses are

moldable even at the high glass content. However, presence of deep

ribs and bosses may require shorter fibers and lower glass contents to

maintain good glass distribution and fill out.

SMC-R has application in parts where strength uniformity is important.

It is considered isotropic in an unmolded state, but becomes anisotropic upon flowing in the mold. To minimize its anisotropic properties,

a charge with the maximum mold coverage should be used. This will minimize flow and resultant knit lines. The presence of fiber orientation and knit lines can be measured, to some degree, by molding multicolor compounds, by burnoff of molded parts and X-ray techniques.

These techniques are valuable for determining a proper charge pattern.

112
• Mechanical Properties - Strength Properties at Various Glass

Contents - SMC-R has no inherent directional characteristics. However,

this isotropy is sacrificed if extensive flow occurs.

Figures 4 through 6 present tensile, flexural, and impact strength

properties from a variety of compounds at varying glass contents for

polyester resin based systems. These figures are based on data from

flat laminates made with a variety of polyester matrix formulations

that contained 20 to 65 percent glass by weight and tested at 23°C.

These properties generally vary linearly with glass content.

Fig. 4 - General range of average Fig. 5 - Dependence of flexural

tensile strengths at various glass strength on glass content for

contents for SMC-R systems SMC-R systems

10 20 30 40 50 60 70 % Glass Content

Fig. 6 - Effect of glass content

on notched izod impact for SMC-R

systems

113
Properties at Elevated Temperatures - it is necessary to consider performance data at, or near, the operating temperatures encountered in

a specific application.

Static tensile and flexural properties of laminates were measured at

elevated temperatures after short term exposure conditioning -- 30

minutes as specified by ASTM test methods D790 and D638. Figures 7

through 14 show these properties for SMC-R50 and SMC-R65 systems.

The graphs include the mean of the data and the predicted 95 percent confidence limits of the mean based on ten specimens, five

taken normal and five parallel to the major axis of the molded plaque.

These specific limits may vary if different molding parameters are

used.

The properties presented are for systems formulated as shown in Table

A. Test plaques were 30.5 centimeter (12 inch) by 45.7 centimeter

(18 inch) by 2.5 + 0.25 millimeter (0.1 +0.01 inch) thick and molded

at 149°C + 5.5°C "C300 + 10°F) for two minutes under 5.5 MPa (800 psi).

The SMC charge was dimensioned to cover 80 percent of the mold for the

SMC-R65 and 45 percent for the SMC-R50. To minimize random flow

effects, laminates were molded with all plies of the same dimension

and to a nominal thickness of 2.5 millimeter (0.1 inch). To include

data variation from other sources, such as change in raw materials,

process set-up and operation, laminate testing samples were prepared

and tested over a six month period. The number of specimens per

test condition ranged from 40 to 70.

The increased variation in the SMC-R50 data as compared to the SMC-R65

data can be largely contributed to the material molding charge used,

45% and 80% coverage respectively. This resulted in additional flow

induced fiber orientation in the SMC-R50 system, thus increased

mechanical property variability.

Tensile Strength vs. Temperature

forSMC-R65

Temperature "F 280 320

- Average

- 95% Confidence Limit (N = 10)

80 100 120 140 160 Temperature "C

Fig. 7 - Tensile strength versus

temperature for SMC-R65

Tensile Strength vs. Temperature

for SMC-R50

Temperature "F

68.0 1O4.0 140 0 176 0 212 0 248 0 284 0 320 0

- Average

95% Conlidence Limil (N = 10)

20 40 60 80 100 120 140 160 Temperature °C

Fig. 8 - Tensile strength versus

temperature for SMC-R50
Tensile Modulus vs. Temperature

forSMC-R65

Temperature *F

68.0 104.0 140.0 176 0 212.0 248.0 284.0 320.0

8

5

80 100 Temperature °C

Fig. 9 - Tensile modulus versus

temperature for SMC-R65

Tensile Modulus vs. Temperature

for SMC-R50

Temperature °F 120 160 200 240 280 320

(/, 20O00 - _ 3 ~~ ~~ —

60 80 100 Temperature °C

Fig. 10 - Tensile modulus versus

temperature for SMC-R50

Flexural Strength vs. Temperature

forSMC-R65

Temperature °F

68 0 104 0 140 0 176.0 212 0 248.0 284 0

' 320.0

- Average

95% Conlidence Limit (N = 10)

80 100 Temperature °C

Flexural Strength vs.Temperature

for SMC-R50

Temperature °F

68.0 104.0 140.0 176.0 212.0 248 0 284.0 320 0

& 300

- Average

95% Conlidence Limit (N = 10)

80 100 Temperature 'C

Fig. 11 - Flexural strength versus Fig. 12 - Flexural strength versus

temperature for SMC-R65 temperature for SMC-R50

Flexural Modulus vs. Temperature

forSMC-R65

680

25000

1040 1400

Temperature °F

176.0 212.0

d.5

3 15000

248.0 2840

Average 95% Confidence Umit

(N = 10)

60 80 100 120 Temperature 'C

0.8

160

a

Fig. 13 - Flexural modulus versus

temperature for SMC-R65

Flexural Modulus vs. Temperature

for SMC-R50

Temperature °F

68.0 104.0 140 0 176 0 212.0 248.0 284.0 320.0

- 95% Confidence Limit (N = 10)

80 100 120 Temperature "C

2

Tig. 14 - Flexural modulus versus

temperature for SMC-R50

115
SMC-C and SMC-C/R

SMC-C (Figure 1) is defined as CONTINUOUS FIBER SMC. It contains continuous reinforcing fiber deposited in a unidirectional fashion.

Continuous Fiber SMC often is produced in combination with random

fibers as SMC-C/R (Figure 2).

In these two systems, the reinforcement content can be adjusted up to

70 percent by weight to produce compounds tailored to specific applications .

Examples of specific types would be SMC-C60, e.g., 60 percent

unidirectional fibers; and SMC-C30/R20, e.g., 30 percent unidirectional fibers combined with 20 percent random chopped fibers.

• Materials - continuous reinforcement input used could be from numerous ends of roving or possibly a unidirectional mat, sometimes

called a "beam". When continuous reinforcement is used with chopped

random fibers (SMC-C/R), one might consider hybrid combinations of

various types of fibers, for example, S-2 Glass and E glass, graphite

and E glass, and aramid with E glass. The majority of our work has

been directed toward using glass roving input, generally E glass.

Other material considerations for SMC-C are similar to those discussed

previously for SMC-R.

« Processing - a belt type SMC machine (Figure 15) can be used to

produce SMC-C or SMC-C/R, although some additions and modifications

are necessary. When continuous roving is used as the input reinforcement, a twist will occur if the roving is pulled from inside the

package. This can be minimized by using a low-yield roving pulled

from the outside. Yield is defined as yards/pound of reinforcement.

Continuous Strand Roving

Fig. 15 - Schematic of SMC-C/R production process showing introduction

of continuous and chopped reinforcement

116
Provisions must be made for additional creel space and equipment to

pull from the outside of the roving package. A creel for this purpose

is shown in Figure 16. A roving spacing device must also be added

to the SMC machine to evenly distribute the rovings across the sheet

width. One technique using a hinged center point tube holder allows

this bank width to be varied (Figure 17).

To enhance the wet-out without strand pattern disturbance, a single

flexible chain link belt was added to the compaction section (Figures

18 and 19). This prevents the roving strands from falling into and

following the serrations found in typical compaction rollers. A further refinement to the Owens-Corning development machine is a double

chain link belt compaction section for even better wet-out (Figures

20 and 21).

Fig. 16 - Creel for SMC-C set up for outside pull of roving

117
!

i

Fig. 17 - View of the

adjustable roving spacing device and the

continuous roving entering the SMC compaction

section

Fig. 18 - SMC compaction

section showing a single

steel mesh belt that has

been added to an existing unit

118
Fig. 19 - Side view of

an SMC compaction section

showing single mesh belt

around existing serrated

compaction rollers

Fig. 20 - End view of

double mesh bel t SMC

compaction unit

119
9 Advantages and Limitations - SMC-C - this system is capable of

producing the maximum mechanical properties in the direction of the

fibers, but with low values when tested perpendicular to the fiber

di recti on.

Combining plies at various angles to one another could be done to

minimize the severe anisotropy of mechanical properties, and to "build"

specific property performance characteristics as needed particularly

modulus and impact (Figure 22).

Strategic placement in parts of SMC-R can be used to strengthen

specific areas.

Creel and machine modifications are necessary to produce the product.

Sheet weight may be somewhat restricted compared to SMC-R.

Because of unidirectional ity of reinforcement, the handling integrity

of SMC-C sheet is not as good as SMC-R.

Flow limitations would require practically 100 percent mold coverage

in the fiber direction. Some degree of flow in the lateral direction

is possible depending on part shape and resin system used.

Incorporation of bosses, ribs or any three-dimensional design feature

in the part is not desirable.

SMC-C/R - the incorporation of random chopped fiber into the SMC-C

system improves its processing and charge handling prior to molding.

There is an improvement in the transverse mechanical properties as the ratio of chopped to continuous fibers is increased.

Higher sheet weight products can be more easily produced for the

same total glass content than with unidirectional fiber above.

The same charge pattern and flow limitations mentioned in SMC-C

system will prevail but to a lesser degree. The feasibility of incorporating minor bosses and ribs will be somewhat improved, particularly when the proportion of the chopped fiber is increased.

Parts may be molded with either the continuous or chopped random fibers

at the surface. Random fibers produce a more uniform appearance,

while the continuous fibers on the surface produce somewhat better

flexural properties.

The process considerations for 100 percent continuous unidirectional

fiber also apply here.

120
• Mechanical Properties - limited mechanical properties have been

generated to-date on SMC-C and SMC-C/R systems (Figures 23 through 26).

The data clearly shows excellent ultimate strength properties as well

as the - high directionality associated with materials systems employing

continuous fibers. Use of continuous fibers significantly increases

the modulus values, above those achievable with SMC-R systems at comparable glass contents.

Undirectional Orthoganal

Examples of laminate ply construction

when using SMC-C

Continuous

Random

Random

Continuous

Random

Continuous

Continuous

Random

or

1 Ply

1 Ply

1 Ply

1 ply

For laminates requiring more than two plies,

balance of p!y construction is recommended.

The flatter the laminate requirements, the more

critical this balance ply construction becomes to prevent warpage.

Fig. 22 - Examples of laminate ply construction when using SMC-C

121
MPa psixicr

900-

120- I I Property in Predominant

1 J Fiber Direction BProperties Perpendicular IB to Predominant Fiber Direction

Note: Polyester Based Systems

• Test Temperature; 23°C

• Random Fiber Length; 1 1nch

SMC - SMC - SMC -

C30/R20 C60/R5 C60 SMC - SMC - SMC - C30/R20 C60/R5 C60

50-

v> AOB

2 30H

"inc

0) *" 20

10 1

- Unidirectional - - Orthoganal

SMC - SMC - SMC - C30/R20 C60/R5 C60

Unidirectional

Properties in Predominant

Fiber Direction Properties Perpendicular

to Predominant Fiber Direction Note:

• Polyester Based Systems

• Test Temperature; 23°C

• Random Fiber Length; 1 1nch

SMC - SMC - SMC - C30/R20 C60/R5 C60 —— Orthoganal —

Fig. 23 - Tensile strengths as effected by different types of

SMCs, laminate construction,

and fiber orientation

Fig. 24 - Tensile modulus values

as effected by different types of

SMCs, laminate construction, and

fiber orientation

MPa psi x 10°

1000

£ 800 cs? 600-

160-

140 -

120-

100-

60-

60-

40

20-

SMC - SMC - SMC - C30/R20 C60/R5 C60

Properties in Predominant

Fiber Direction

I Property Perpendicular lo Predominanl Fiber Direction

Note:

. Polyester Based Systems

. Test Temperature; 23°C

• Random Fiber Length; 1 1nch

SMC - SMC - SMC - C30/R20 C60/R5 C60

-Orthoganal -

GPa psi x 10°

Ilk SMC - SMC - SMC - C30/R20 C60/R5 C60

-Unidirectional

Properties in Predominant

Fiber Direction Properties Perpendicular

i to Predominant Fiber Direction Note:

. Polyester Based Systems

Test Temperature; 23°C

• Random Fiber Length; 1 1nch

SMC - SMC - SMC - C30/R20 C60/R5 C80

-OrlhoganolFig. 25 - Flexural strengths as

influenced by different types of

SMCs, laminate construction, and

fiber orientation

Fig. 26 - Flexural modulus values

as influenced by different types

of SMCs, laminate construction and

fiber orientation

SMC-D and SMC-D/R

SMC-D (Figure 1) is defined as DIRECTIONAL FIBER SMC. It contains

fibers of four inches or longer in length in a discontinuous unidirectional pattern. Handling and performance considerations, however,

generally require the Directional Fiber SMC to be made in combination

with some random fibers as an SMC-D/R system (Figure 2).

122
• Materials - the resin, filler, reinforcement, and other material

considerations for SMC-D are the same as those discussed in the previous SMC-R and SMC-C sections.

• Processing - there are two SMC manufacturing approaches in use to

produce the SMC-D and SMC-D/R material systems; on-line and off-line.

Both use a similar long fiber cutting arbor to cut the roving into the

long, unidirectional pattern. The blades in this arbor are set-up to

cut fibers into long lengths on a random cut pattern (Figure 1).

Using the on-line approach, introduced by OCF-Europe in support of the

FRP bumper programs, the long glass fiber cutter is placed on the SMC

machine. By synchronizing the peripheral speed of this long cutter

with the belt speed, long fibers are cut onto the moving SMC web so

that the roving ends butt against one another (Figures 27 and 28).

The alternate approach to producing the SMC-D/R material is an off-line

operation using an SMC-C/R material matrix as described previously.

Figure 29 shows the off line making of SMC-D/R charge pattern, and

molding results. The sheet is allowed to thicken to three million

centipoise (as measured by the Brookfield HBT Viscometer) or above in

paste viscosity before being processed through the long cutter.

The blades pierce the carrier film and cut the continuous rovings into

predetermined lengths. Thus, SMC-C/R is converted into SMC-D/R.

This operation could be incorporated into the conventional SMC cutter/

slitter machines.

• Advantages and Limitations - the major reason for considering the

SMC-D approach is its improved molding characteristics relative to

SMC-C. Cutting the continuous fibers allows the material to flow in

the fiber direction as well as in the transverse direction. Thus,

this system provides a compromise between the good flow characteristics

of SMC-R and the superior mechanical properties of SMC-C.

Two process approaches are available. An additional cutter must be

utilized for either approach, The use of a secondary piercing operation may provide the most versatility since all three SMC materials

can then be produced on one machine.

SMC-D has a definite direction bias relative to mechanical properties;

therefore, it should be used where this characteristic contributes to

part performance. Specific material selection is based on part requirements and molding limitations.

Random wi ndows occur in the molded laminate. Their size and effect

are controlled by the degree of flow occurring, the length of the

fibers used, ratio of long and short fiber contents, and the overall

glass content.

123
Similar to SMC-C requirements, additional roving creel space is necessary.

As in SMC-C, a balanced ply construction is recommended in the molded

part to minimize residual stresses and warpage.

0 Mechanical Properties - Owens-Corning is currently reviewing test

procedures to permit representative measurements of SMC-D properties

at high glass contents. The concern is that since some localized

discontinuities exist (windows) the relatively small test specimens

currently specified in standard ASTM test methods may not provide a true definition of composite properties.

Mechanical properties of SMC-D systems are expected to fall between

those of SMC-R and SMC-C systems of comparable glass content with

values approaching those of SMC-C as the length of the long unidirectional fibers increase.

Continuous Strand Roving

Fig. 27 - Schematic of SMC-D/R on-line production process utilizing

two roving choppers

I2h
Fig. 28 - On-line SMC-D/R process showing long and random fiber

cutters on an SMC machine

Off-Line Production Process

Cutting Arbor

Fig. 29 - Schematic of SMC-D/R off-line production process showing

secondary piercing operation used to convert SMC-C/R and SMC-D/R

125
Cost and Performance Comparisons

• Raw Material Cost - SMC formulations for use with 30 percent to 65

percent glass content by weight were established and composite raw

material cost per pound calculated (Table B). These formulations were

based on the use of polyester resin, calcium carbonate filler and

other commonly used components. Raw material cost varied from $0.38

per pound for 30 percent glass content to $0.56 per pound for 65 percent glass contact composition.

• Process Cost - Processing costs to make the sheet product are expected to vary depending on the type of SMC produced and glass content

desired. As indicated in previous processing discussion, some equipment modifications are necessary to produce the SMC-D and SMC-C materials. Generally, higher glass contents mean a lower weight per

square foot product at a slower line speed. This reduces productivity

thus increasing costs. Additional experience is needed before process

cost variables can be accurately assessed.

• Performance Comparisons - the performance data previously shown

indicates the relatively broad range of properties that can be obtained by varying the SMC composition and type. A multitude of

different compositions and additional comparisons can be made in each

category. In Tables C and D, representative SMC systems were selected

and equivalent performance calculations and comparisons made. For

comparative purposes, representative values for high strength low

alloy steel (HSLA 950) and aluminum (7021-T6) currently being used

for bumpers have been included.

Volume Costs - cost per cubic inch takes into consideration both

the material cost per pound and the material density. Since the

densities of SMCs we have presented are nearly constant at 0.066

pounds per cubic inch, volume costs are directly proportional to the

material's cost per pound. The higher density of steel and aluminum

make them significantly more costly than SMC on simple volumetric

basis

.

Flexural Strength - when comparing the different SMC-R systems, an

increase in the glass content provides a weight savings but no appreciable cost savings for equivalent flexural strength performance.

The thinner laminates at the higher glass contents can provide weight

savings but the increased material costs negates any cost saving.

When the SMC-D and SMC-C systems are used effectively with the stress

applied in the direction of the predominant fiber direction, additional weight and cost savings over the SMC-R systems can be expected.

Stiffness - comparing the SMC systems, it is apparent that the much

higher modulus values of SMC-C systems allow for decreased laminate

126
thickness, thus weights, over the SMC-R systems.

When compared with the HSLA steel, the SMC systems are lighter in

weight, but more costly. When compared with the aluminum, the SMC

systems are less costly and for SMC-C the weight is also less (Table D)

These values for SMCs, however, are conservative since the comparisons

are based on flat plate theory and do not take into account the ease

of molding SMC into complex shapes, addition of ribs and changing

cross sectional geometry, and the commonly used techniques for adding

stiffness

.

Tensile Strength - when comparing the SMC systems to steel and aluminum the same general trend in cost and weight as noted in stiffness

apply.

Comparing the various SMC systems, the value of increased glass loading and fiber orientation is evident. The higher tensile strengths

occurring from use of higher glass contents allow a reduction in

thickness that more than offsets higher material costs. As a result,

there is a significant weight and cost savings potentially available

where higher glass type SMCs are applicable.

SMC Raw Material Costs (Cost Calculations based on prices in effect March 1, 1979)

Resin Filler

Reinforce- ment Catalyst Mold Release Thickener

Glass

Content

Polyester

Calcium

Carbonate

Fiberglas

Roving

T-Butyl

Perbenzoate Zinc Stearate Magnesium

Oxide Total

$/lb. .49 0.05 0.58 $2.16 0.94 0.88 $/lb.

30% Parts lbs. 35 35 30 0.35 088 0.53 101.8

Cost$ 17.15 1.75 17.40 0.76 0.83 0.47 38.36 0.377

40% Parts lbs. 35 25 40 0.35 0.88 0.53 101.8

Cost$ 17.15 1.25 23.20 0.76 0.83 0.47 43.66 0.429

50% Parts lbs. 35 15 50 0.35 0.88 0.53 101.8

Costs 17.15 0.75 29.00 0.76 0.83 0.47 48.96 0.481

60% Parts lbs. 35 5 60 0.35 0.88 0.53 101.8

Costs 17.15 0.25 34.80 0.76 0.83 0.47 54.26 0.533

65% Parts lbs. 35 0 65 0.35 0.88 0.53 101.8

Costs 17.15 0.0 37.70 0.76 0.83 0.47 56.91 0.559

Table B

127
Cost-Performance Comparisons for Raw Materials (Cost Calculations based on prices in effect March 1. 1979)

Flexural Strength C) Stillness < 1 > Tensile Strength (1)

Materials

Cost

S/lb.

Density

lb./in. 3

Volume

Costs

$/ln. 3 103

psi

ThirLnocc 1 MIL 1. 1 It i. > Inches (2)

Cost

$/sq. ft.

Flexural FAOQUIUS

106 psi

1 1 i.uMUsa

Inches (3)

Cost

S/sq. ft. 103 psi

1 illijrl ItJob Inches (4)

Cost

S/sq. It.

SMC-R30 0.377 0.066 2.49 23 0.114 0.41

• 1.9 0219 0.79 1

1

0091 0.33 SMC-R40 0.429 0.066 2.83 30 0.100 0.41 2.0 0.215 0.88 17 0059 0.24 SMC-R50 0.481 0066 3.17 37 0.090 0.41 22 0209 0.95 23 0.043 0.20 SMC-RS5 0.559 0.066 3.69 48 0.079 0.42 2.3 0.206 1.09 30 0033 0.18

SMC-C30/R20

unidirectional

0.481 0.066 3.17 QCOO uuuy O.J U- 0.83 33 U.U l o 0.08

SMC-C60/R5

orthoganal 0.559 0.066 3.69 68 0.066 0.35 4.5 0.164 0.87 40 0.025 0.13

SMC-C60/R5

unidirectional 0.558 0.066 3.69 100 0.055 0.29 5.3 0 156 0.83 64 0.016 0.09

SMC-C60

unidirectional 0.533 0.066 3.52 130 0 048 0.24 5.4 0 154 0.78 81 0012 0.06

Steel (HSLA950)

0.192 0.271 5.20 50 0078 0.58 30 0087 0.65 50 0.020 0.15

Aluminum

7021-T6

0.87 0.090 7.83 50 0.078 0.88 10 0126 1.42 50 0.020 0.23

Notes: 1. Properties from predominant fiber direction in laminate.

2. Thickness to support 100 pounds on a 1 inch wide sample on a 2 inch span at indicated stress.

3. Thickness to support 100 pounds on a 1 inch wide sample on a 2 inch span with 0.01 inch deflection C4.Thickness to support 1000 pounds on a 1 inch wide sample at indicated stress.

Material

Pounds

tor Equivalent Performance

Dollars

lor Equivalent Performance Systems Basis - 10 Pounds Steel Basis -S1.00 Steel

Volume Flexural

Filling Strength Stiffness

Tensile Strength Volume Flexural

Filling Strength Stiffness

Tensile Strength SMC-R30 2.4 4.5 6.1 11.1 047 0 72 1.23 221 SMC-R40 2.4 3.1 60 7.2 0.52 0.72 1.36 1 64 SMC-R50 24 28 59 52 058 0.72 1.48 1.36 SMC-R65 24 2.5 5.8 4.0 067 0 74 1.69 1.21

SMC-C30/R20

unidirectional 24 18 5.1 22 0.58 0.48 1 30 057

SMC-C30/R20

orthoganal 24 2.1 46 30 0.67 0.61 1 34 0.93

SMC-C60/R5

unidirectional 24 1.7 4.4 20 0.67 0.50 1 28 0 57

SMC-C60

unidirectional 24 15 4.3 1.5 0.64 043 1.21 0.43

Steel (HSLA 950) 100 100 100 10.0 1.00 1 00 1.00 1.00

Aluminum

(7021 T6)

3.3 3.3 48 33 1 40 1 41 203 1 43

Table D

128
Summary

A versatile family of high-strength SMC materials has been developed.

The basic systems are designated SMC-R, SMC-C, and SMC-D, where R,

C, and D stand for "Ramdon", "Continuous", and "Directional", and

indicate the configuration of the reinforcement used. Combination

of fibers allow other systems to be considered also, such as: SMC-C/R

and SMC-D/R. These materials are processable on conventional SMC

machines with minor modifications, and thus provide the compound with

the ability to produce structural grade SMC with a minimum capital

investment.

The molder is provided a materials system with which he can deliver

mechanical properties and selective directional reinforcement in

order to comply with special part requirements. This flexibility in

reinforcement will lead to improved efficiency of design, which will

result in minimum part cost and weight for any given application,

Data reported in this paper have been for conventional static mechanical properties. Although such data are necessary for any design

and are sufficient for those applications where cyclic loading is not

significant, in order for the use of these materials to be extended

to the broad range of parts for which they are suitable, fatigue information as well as data on shear properties and Poisson's ratio is

necessary. At this time, Owens-Corning is well underway with a program to develop these properties for SMC-R, SMC-C, and SMC-D,

which will lead to an even wider range of applications than those

presently considered.

129
Development of a Graphi te-Epoxy Spacecraft Precision Mounting Platform

Kathleen S. Budlong and Alvin W. Sheffler

RCA Astro-Electronics Division

Princeton, New Jersey 08540

Abstract

A wide variety of spacecraft applications require structure which is ultralight in weight while satisfying tight stiffness (resonant frequency)

criteria, along with critical pointing accuracy and thermal distortion

requirements. This paper presents the design, analysis, and test of a graphi te-epoxy precision mounting platform intended to replace the current

aluminum platform being flown on an Air Force satellite program.

The Precision Mounting Platform (PMP) provides an instrument mounting

surface which is defined and controlled for each instrument to within

0.01 degree. It maintains a mutual alignment of the instruments to within 0.003 degree through the full range of thermal gradients occurring on

orbit. While the aluminum PMP uses active thermal controls to minimize

thermal gradients and maintain alignment, the composite design takes full

advantage of the near zero coefficient of thermal expansion of the GY70/

X-30 graphi te-epoxy , thus eliminating active control. The ultra-high

stiffness of the composite is also critical in maintaining the resonant

frequency range defined by overall spacecraft dynamics for the aluminum

PMP. The graphi te-epoxy PMP provides in addition a weight savings of 32

percent while fulfilling all the requirements of a direct replacement

for the current aluminum version.

130
Olaetirmgtn: "W\ "W*. Gunkel

General Motors Technical Center
i

:-;<i
Design Analysis of Automotive Composite Structures

M. F. Kowalski, D. S. Fine, and R. K. Herrman

Chevrolet Engineering Center

Warren, Michigan 48090

Abstract

Advanced composite materials offer potential for structural mass reduction in automotive structures due to their high strength-to-weight

ratios. In addition, they provide styling freedom to reduce aerodynamic

drag forces and part consolidation advantages for vehicle system design.

Consequently, they may play an increasing role in the reduction of fuel

consumption in the next generation of automobiles.

Traditionally, prevention of mechanical failure has been a key automotive design goal. This is necessitated because our products are often

subjected to abuse for lack of required periodic inspections and maintenance. Failures lead to product liability claims, high warranty repair

costs, and dissatisfied customers. Failure prevention is difficult

because of the wide variety of operating conditions and environments our

product experiences, and because the product must be produced by a mass

production technology at a cost our customers can afford. Considering

the fact that there are over 15,000 parts in an average automobile, our

components must be designed for very high reliability.

Because of the nature of advanced composite materials, particularly their

anisotropy and their multiple failure modes, the design and analysis

procedures currently employed to minimize failure on metal structures

are not adequate for composite structures. In anticipation of the need

to address the automotive design possibilities of composite materials, a study has been conducted to determine what modifications are needed in

the existing finite element structural analysis procedures. It has been

found that a composite material analysis program, SQ5, now widely used

in the aerospace industry, offers potential for automotive structural

analysis.

SQ5 is based on classical laminate plate theory and can be used in two

modes. The first is an input pre-processor to compute equivalent orthotropic properties from a laminate description. The second is an output

post-processor to predict first ply failure and safety margins from

laminate strength data. The program has been modified to employ the

interactive Tsai-Wu tensor polynomial lamina strength failure criterion.

Limited comparisons have been performed with theoretical solutions,

based on classical laminate plate theory, and with test data for static

and dynamic response of composite panels. Predicted failure loads have

also been compared with measurements.

153
Fiberglass Reinforced Plastics Use in Transportation Vehicles

R. D. Pistole

GTR Reinforced Plastics Company

South Field, Michigan 48037

'

Abstract

The use of Fiberglass Reinforced Plastics (F.R.P.) in ground vehicles has

steadily increased in recent years.

In 1975 one of F.R.P.'s largest penetrations into exterior passenger car

panels was grille opening panels or front headers. This use was cost

effective due to the consolidation of many metal parts by one F.R.P.

molded panels.

Today, with the need for lighter weight automobiles to improve fuel

economy, F.R.P. is again the answer with its additional advantages.

One piece truck hoods are becoming more popular as they too turn to fiberglass composites. Many new truck designs utilizing F.R.P. are appearing

on our highways.

The manufacturing of F.R.P. car and truck components will be discussed

in this presentation.

13k
DESIGNING AUTOMOTIVE COMPONENTS WITH CONTINUOUS FIBER COMPOSITES

H. T. Kulkarni and P. Beardmore

Research Staff - Ford Motor Company, Dearborn, Michigan 48121

ABSTRACT:

The primary characteristics of continuous graphite fiber reinforced

plastics as structural materials are high specific strength and modulus

and exceptional fatigue strength. All these properties are a function

of fiber orientation in the individual layers and the lay up sequence.

The advantages of continuous fiber reinforced composites relative to

the conventional chopped fiber systems currently utilized in automobiles

lies in the ability to apply these composites in major load/stiffness

critical components. The application of a composite materials design

methodology to prototype automotive components is described. Examples

of different types of components are given, specifically wheel and

body structures. Both the design and the fabrication of these structures in composite materials are described and any specific problems are

outl ined.

Key words: Automotive component design; continuous fiber composite

materials; fabrication; finite element stress analysis; front end;

graphite; part integration; wheel.

I. INTRODUCTION:

Current automotive usage of fiber reinforced plastics is limited primarily to the application of chopped fiber reinforced materials which can

at best be applied in semi-structural applications. The application of

fiber reinforced composites to structural, safety related components

requires the introduction of continuous fibers into these systems in

order to withstand the structural loads imposed on the components.

Continuous graphite fiber reinforced composites exhibit high specific

strength and modulus in combination with excellent fatigue strength,

making these materials excellent candidates for load bearing and stiffness critical applications. These materials offer the incentive of

large potential weight savings over conventional materials and thus are

attractive from the viewpoint of minimizing overall vehicle weight with

the associated increase in fuel efficiency. The application of GrFRP

materials in the aerospace industry has developed a background of baseline data and design procedures which might be utilized in the design

135
of automotive components in these materials.

Drawing on the recent Ford experience in the design and fabrication of

a full size GrFRP 1979 Ford experimental vehicle, the present paper

attempts to illustrate the differences in approach for the design of

composite components relative to conventional steel components. As

part of this vehicle program a composite component design methodology

was established for automotive components by drawing together all the

best parts of design systems available through the aerospace industry

into one integrated design program specifically tailored for automotive

applications. This design methodology was used for all GrFRP components

and a short summary of the procedure is given in Section 2. The design

methodology is described in detail elsewhere (1) and the current description of the methodology is only intended as a summation. Illustrations of the usage of the methodology are given in Sections 3 and 4.

The design of a strength critical component, namely the GrFRP wheel is

detailed in Section 3, whereas by way of contrast a stiffness critical

design, namely the GrFRP front-end structure, is outlined in Section 4.

This procedure is followed in order to provide a comparison for the

basic design philosophy used based on the differing critical requirements of the component under consideration. Again, neither of these

illustrations is intended to be completely detailed since the complete

integrated design cycle for both of these components has been the subject of separate papers (2) and (3). The overall intent of this presentation is to provide a capsule summary of design techniques with illustrations which can be used in understanding the application of continuous fiber reinforced composites in automotive or other types of ground

transportation. For extensive details on the procedures the references

quoted above should be studied.

2: DESIGN METHODOLOGY FOR COMPOSITE COMPONENTS

A typical automotive component design cycle is shown in Figure 1. There

are three distinct phases involved and the overall design cycle is independent of the materials used . Phases 1 and 3 are independent of material and therefore are common to both composite components and conventional isotropic material components. However, the details of the

design Phase 2 are dependent on the material and it is in this aspect

of the design cycle that special consideration must be given when

designing with composite materials. Phases 1 and 3 are self explanatory

and reference 1 should be used to obtain further details of these

aspects of the design cycle.

Phase 2 of the design cycle consists of five basic elements as shown in

Figure 1. These are (a) laminate design, (b) stress analysis, (c)

fabrication methods, (d) cost analysis and (e) failure analysis. Of

course, the impact of the last three of these factors on the design of

composite components is similar to the effect on conventional components.

136
There are significant differences between composite materials and conventional materials in the specific details of these latter three

factors but the overall methodology of consideration in the design

phase is very similar. As a distinct example of the differences in

these factors, consider the influence of fabrication methods. The

utilization of plastic based composite materials allows the opportunity

for integration of several parts into a single fabricated piece. This

technique of component integration is one of the key factors in application of all plastics into automotive components and represents an area

where fabrication cost savings and advantages in overall stiffness can

be obtained based on the unique nature of the material relative to steel . The cost advantage, of course, accrues from the elimination of assembly

costs at later stages and the increased stiffness developed by the

elimination of joints. This example of the differences in fabrication

techniques between composite materials and conventional materials

illustrates the interplay of the five factors listed above in arriving

at the best overall design for given functional requirements. The two

areas of Phase 2 which require a unique approach for the application of

composites are laminate design and stress analysis and these will be

discussed separately below.

(a) Laminate Design

The stacking of individual composite layers to form a part involves the

selection of the individual layers to yield the required laminate properties. Consequently, the basic design parameters going into the design

of the laminate are the layer properties and the laminate parameters.

These are as follows

:

. Layer parameters - type and orientation of fibers, type of matrix,

form of the layer (cloth or tape).

. Laminate parameters - stacking sequence of the layers.

The design of the initial laminate must be based on a selective predominant design criterion. The criterion might be one of the following:

. Hatch the stiffness and/or strength of the steel component or some

given percentage thereof.

. A specified strength and/or stiffness criterion with no reference to

a steel component.

Once the choice of the predominant design criterion is made, an initial

laminate is developed by following the iterative technique as shown in

Figure 2. Based on simple hand calculations a reasonable initial laminate design is developed by this iterative procedure prior to the full

stress analysis

.

(b) Stress Analysis

Except for very simple components, all laminate designs are verified in

detail by the application of a finite element stress analysis procedure.

137
In this technique a composite finite element model is developed for the

component. The stress analysis differs from such analysis for conventional materials in that it involves lamination theory, that is the

ability to examine the stresses in the individual layers which make up

the laminate. The normal criterion for failure is over stressing relative to the design allowables in any layer in the structure. Obviously,

failure could occur in a subsurface layer which is not visible on

external examination and thus the failure criterion must be applied to

every layer to insure integrity of the total structure. This is in

distinct contrast to isotropic materials where failure normally initiates on the surface and is, therefore, evident by visual examination.

The final laminate design is fine tuned by iteration involving changes

in the stacking sequence and orientation until all layers are judged to

be satisfactory under the imposed load conditions. In addition, weight

minimization can be achieved by eliminating excessive overdesign. With

experience, the number of iterations involved in the final stress analysis will be limited, yielding an efficient and flexible design system.

In the following two sections, examples are given of the application of

the design methodology to a GrFRP wheel and a GrFRP front end structure.

The basic material properties used in the design of these components

are listed in Table 1. The components are currently installed on the

experimental GrFRP Ford and have to date successfully operated through

all vehicle evaluations.

3: STRENGTH CRITICAL COMPONENT - GrFRP WHEEL

(a) Design Requirements

There are two basic requirements which dictate the design of the composite wheel. Firstly, there are specific load carrying capabilities

which are evaluated by the following critical tests;

Rim roll fatigue - an examination of the fatigue strength of the

wheel rim.

. Rotary fatigue - a test of the fatigue life of the wheel spider under

a bending load.

Impact - a test of the integrity of the wheel rim under severe impact

load simulating curb impact vehicle operation.

The second critical requirement for the wheel is based on the axisymmetric nature of this component. This geometric constraint involves

a design such that the load input points will change as the wheel

rotates. The wheel must therefore have equal strength and stiffness

along any radial direction from the center of the wheel

.

138
(b) Laminate Design

Consideration of the three loading conditions noted above indicated that

the dominant critical load requirements were those required to pass the

impact tests. The basic design philosophy adopted was to design the

composite wheel to pass the impact tests, evaluate such design analytically and ensure that the design would also pass both the fatigue

tests. Based on the impact loads, the number of composite layers

required for satisfactory wheel performance were computed. Details of

the impact loads are given in reference 2. Laminates were designed with

alternate layers of GrFRP cloth at orientations of 0°/90° and + 45° to

give approximately equal strength and stiffness along the radial directions .

(c) Stress Analysis

The finite element model generated for the GrFRP wheel is shown in

Figure 3. The laminate properties for various elements in the model

were defined and the model was analyzed for the various critical load

cases discussed previously. The layer stresses computed for both the

spider bending and the rim roll fatigue loads were far below the design

allowables for the individual GrFRP cloth layers. However, the laminate

design as shown in Figure 4 was rejected since the induced layer

stresses under the impact loads were much higher than the design allowables. The localized high stress areas in this design are shown in

Figure 4. Modification of the laminate and subsequent stress analysis

reduced the high stress areas adjacent to the bolt holes to within safe

limits. The final laminate design and stress distribution is shown in

Figure 5.

(d) Fabrication and Testing

The wheels were fabricated by conventional aerospace techniques resulting in a final component as shown in Figure 6. Note the stainless

steel trim ring on the GrFRP wheel which was both a styling feature

and a functional feature. This ring helped to distribute the loads

both in the impact test and also in tire changing procedures. This

GrFRP wheel design (Figure 5 and Figure 6) was subjected to all the

critical tests mentioned above. As predicted by the stress analysis,

the wheel successfully passed all of the tests. It should be noted

that a prototype wheel of the initial laminate design as shown in

Figure 4 was also built and tested and the wheel failed in the impact

test. Since the stress analysis for this particular design had indicated that failure would occur in the impact test, the excellent

correlation between the stress analysis for both the initial and the

final design and the testing results gave great confidence in the

composite design methodology procedure which had been utilized for all

the other components.

139
The GrFRP wheels represent a 42.7 pound weight saving for the vehicle

(8.55 pounds per wheel) for approximately 45% weight savings. It should

be noted that the wheels represent one of the most difficult design

fabrication tasks in the vehicle. Thus, the weight saving of approximately 45% relative to steel is not as high as can be achieved in many

of the components with less stringent requirements. Typically in those

types of components weight savings up to 65% can be successfully obtained. The severe loading conditions and critical safety requirements

for the wheel impose greater restrictions on the design flexibility for

composites than in most other vehicle components and thus do not represent the kind of weight savings that might be expected from these

materials. On the other hand, the successful design and fabrication of

wheels represents a major achievement in the ability of the materials

to withstand very complex and severe loading environments.

4: STIFFNESS CRITICAL COMPONENT - GrFRP FRONT END STRUCTURE

(a) Design Requirements .

The dominant design factor for the front end structure was selected as stiffness based on the paramount need to meet critical NVH ride quality

requirements. Thus, the overall design philosophy was based on deflection requirements with the modification that the deflection could be

adjusted to allow for the lower input loads since the overall weight

of the vehicle was reduced by about one-third. (3) Based on this philosophy, the individual components of the front end structure were

divided into two different groups, namely bending critical components

and torsion critical components. In addition, local areas such as the

bolt attachment points were separately analyzed for strength critical

requirements to insure the integrity of the joints.

(b) Laminate Design .

The mechanical properties of the individual layers used in the design of

the front end structure are given in Table 1. For a component which is

considered to be primarily bending stiffness critical, the GrFRP laminate design was matched to the bending stiffness of the steel part.

Likewise, for torsional stiffness critical components, the torsional

stiffness of the GrFRP components was matched to the torsional stiffness of the corresponding steel part. As a consequence of choosing a critical stiffness parameter, the GrFRP components only matched the

steel in either torsional or bending stiffness and differed in the

second stiffness area. This follows the basic design philosophy of

composites in that the maximum weight reduction is obtained by using

the directionality of these materials to match the critical requirements,

and having lesser properties in other less critical directions.

The stiffness of the individual components of the front end structure

was matched by using the following equations:

1*4-0
Bench" ng Sti f fness .

3 n 3 2

E t =

s s 12

(1)

Torsional Stiffness.

G K

s s (2)

Where E = Young's Modulus, G = Shear Modulus

t = thickness, i = i layer

n = total number of layers

Y = distance form neutral axis

W = Weight of the car

K = torsional constant

Subscript, g

= GrFRP, s = steel

Based on these calculations, the thickness of each individual composite

part was computed and the parts are compared with the steel parts in

Table 2. One difference between the GrFRP front end structure and the

steel front end structure lies in the fender inner apron. In the steel

assembly, the fender inner apron is non-structural and is made of

polypropylene. In the GrFRP structure, the inner apron was made structurally effective by fabricating this fender in GrFRP.

It is clear from comparison of the data in Table 2 that only one of the

two critical stiffnessess are matched in the composite material relative

to the steel component. It should also be pointed out that whichever

of the design criteria are used, i.e. bending or torsional stiffness,

the GrFRP component only had to be 67% of the steel stiffness to

match the same deflection criteria. As mentioned previously all the

bolted joints between the radiator and fender and between the fender

and the body were examined for failure modes such as shear out, net

tension and bearing strength. A major integration of components was

achieved in the GrFRP front end. In each fender assembly parts in

steel were integrated into parts in GrFRP. Likewise, in the radiator

support assembly, parts in steel became parts in GrFRP. Such part

integration supports the point made earlier about fabrication advantages of GrFRP.

ihl
(c) Stress Analysis

Finite element models for both the steel front end and the composite

front end were generated as shown in Figures 7 and 8. To test the

adequacy of the structure, loads were applied to the assemblies at two

frame mount locations on the radiator support lower extension. Models

were analyzed for both bending and torsional load cases. The static

analyses were carried out using finite element codes which have the

capability of both isotropic and anisotropic material analysis. The

results of the analysis are summarized in Table 3. Note that based on

the weight of the steel vehicle, namely 3750 pounds, the GrFRP front end

structure is only 80% as stiff in both bending and torsion as the steel

structure. However, based on the 2500 pound weight of the composite

vehicle, the relative stiffness of the GrFRP front end structure is 20%

greater than that of the steel structure in both bending and torsion.

Thus, the analysis would predict that the composite front end structure

in the composite vehicle would act as a structure of increased stiffness

relative to vehicle performance when compared to the steel vehicle,

i.e., the analysis predicted that no problem should result in the NVH

characteristics of the front end from a stiffness viewpoint.

(d) Fabrication and Testing

The individual components of the composite frontend assembly were fabricated by conventional aerospace hand layup techniques as discussed before.

The individual components were assembled into the complete front end

structure and installed on the experimental GrFRP vehicle, Figure 9.

The complete vehicle is currently undergoing vehicle evaluation tests

and has successfully demonstrated that the design of the front end based

on stiffness critical criteria has not reduced the NVH characteristics

of the structure. The overall GrFRP front end assembly weighed 30

pounds versus 95 pounds in steel . This represents a weight savings of

68.4% and demonstrates the potential of these materials in reducing the

weight of conventional structures.

5: CONCLUSIONS

Current aerospace design techniques can be adapted for the application

of continuous fiber composite materials to conventional automotive

structures. Such a procedure has been used to design various parts of

an experimental GrFRP vehicle and the influence of different design

constraints on the successful functionality of the components has been

demonstrated. This demonstrates that highly anisotropic materials can

be successfully utilized for the design of complex geometric components

with highly complicated stress environments. Also, the potential of

these materials for successful weight savinq applications in these types

of structures is illustrated. Proviu j ther constraining faotors,

particularly, cost and high yield manuidLtu. ing techniques, can be

11+2
successfully resolved, the potential for widespread usage of continuous

fiber reinforcement composites in the transportation industry has been

shown to be technically feasible.

REFERENCES

1. H. T. Kulkarni and P. Beardmore, "Design Methodology for Automotive

Components Using Continuous Fiber Reinforced Composite Materials".

FORD TECH. REPORT NO. SR-79-04, To Be Published.

2. M. B. Motwani and H. T. Kulkarni, "Experimental Graphite Reinforced

Plastic Wheel for '79 GrFRP Ford Concept Vehicle". SAMPE Meeting,

May 8, 1979.

3. H. T. Kulkarni, G. Lim, R. C. Knight, P. Beardmore "Design and

Fabrication of Automotive Components in Graphite Fiber Reinforced

Composites, Part II Front-End", SAE Paper No. 790031, Feb. 26,

1979.
TABLE 1

GrFRP LAYER MECHANICAL PROPERTIES

Sti ffness

Property GrFRP Tape GrFRP Cloth

E]

i x 10° psi 21 .0 10.3

E

22

x 106

psi 1 .2 10.3

6n o x 106 psi 0.65 1 .0

Poisson 1 s Ratio v 0.25 0.09

Thickness 1 t

J inches 0.005 0.011

Strength

TnKsi 190.0 76.0

T^2 Ksi 5.0 74.0

Ksi 160.0 64.8

C

22

Ksi 20.00 61 .0

T = Tension, C = Compression, E = Young's Modulus, G = Shear Modulus.

1,2 = Along and perpendicular to the fiber axis.

Ikk
TABLE 2

COMPARISON OF THICKNESS, BENDING AND TORSIONAL STIFFNESS

OF THE GrFRP PARTS WITH STEEL PARTS

COMPONENT

Fender Outer Panel

Front Reinforcement

Top Front Reinforcement

Side Front Reinforcement

Rear Side Reinforcement

Wheel Opening Reinforcement

Upper Support

Upper Reinforcement

Lower Supports

Gussetts

Lower Extensions

Side Supports

Gr = GrFRP

St = Steel

ASSEMBLY

Fender

STIFFNESS RATIO

'9 r/tst

1 .42

1 .26

1 .63

2.31

1 .79

1 .57

Radiator support 1 .05

2.57

1 .43

1 .43

1 .48

3.09

BENDING TORSION

m

0.9 0.58

0.628 0.51

1.177 0.66

3.39 0.94

1.49 0.80

1 .067 0.72

0.44 0.15

5.5 0.72

0.945 0.41

0.945 0.41

1.04 0.34

6.40 1.17

1^5
TABLE 3

RELATIVE STIFFNESS OF THE FRONT END ASSEMBLIES IN PERCENTAGE BASED

ON THE FINITE ELEMENT ANALYSIS

Total

weight of

the car

in lbs.

Front End

fabri cated

with

Sti ffness

Bending

in Percentage

Torsion

3750 steel 100 100

3750 GrFRP 80 81

2500* GrFRP 121 121

* Input loads to the GrFRP front end assembly will be 2/3

(i.e. 2500 ) of the input loads to the steel front end assembly.

3750

11+6
DESIGN CYCLE FOR A COMPONENT

LOADS

STYLING 8 PACKAGING CONSTRAINTS

• FUNCTIONAL REQ.

•FABRICATION

• MATERIAL PROPERTIES

• STRESS ANALYSIS

• COST ANALYSIS

• FAILURE ANALYSIS

NON-DESTRUCTIVE

TESTWG

COMPONENT TESTWG

ROAD TESTING VEHICLE RIDE QUALITY

TESTING

I INPUT

n DESIGN i

| m TESTING - •FAILUREMASS

EXP.

Fig. 1. Design cycle for a component showing the progressive

development of the design.

DESIGN FOR A COMPOSITE COMPONENT

COMPONENT DESIGN

CONFIGURATION

COMPOSITE

MATERIAL

PROPERTIES

LAMINATE

DESIGN

FABRICATION

METHODS

LOADS AND

BOUNDARY

CONDITIONS

REASONABLE

A

START STRESS

ANALYSIS

Fig. 2. Flow chart of the development of efficient laminate design

1^7
Fig. 3. Finite element model of the wheel.

LAY-UP DESIGN

Fig. 4. Initial lay-up and stress-distribution in the first GrFRP

wheel design due to 8g impact load.

IhQ
Fig. 5. Final lay-up and stress-distribution in the GrFRP wheel

design due to 9g impact load.

1^9
Fig. 7. Finite element model for the steel Front End assembly.

Fig. 8. Finite element model for the composite Front End assembly.

150
Fig. 9. GrFRP Ford experimental vehicle.

151

Glisiiraieixi: aJ\ US. Stern

(siaTDStituLting for* ^.zitoxi EEelixi)

Godciarci S;pa,oe Fliglit Center

FABRICATION OF LARGE

COMPOSITE SPARS AND BLADES

By

O. Weingart

Structural Composites Industries, Inc.

6344 North Irwindale Avenue

Azusa, California 91702

Abstract: A national wind energy program has been established to

develop the technology necessary to enable wind energy systems to be

cost-competitive with conventional power generation systems and

capable of rapid commercial expansion for producing significant

quantities of electrical power. Wind Turbine Generators being developed under this program will require rotor diameters of up to 300 ft.

Because of the considerable extension in the technology required to

design and construct blades of this size, it was highly desirable to

establish an early technology base. To that end, the Department of

Energy, through NASA-Lewis Research Center has awarded several

contracts for the design, fabrication and test of large low-cost wind

turbine blades built of composite materials.

This paper describes SCI's fabrication of a 20,000 lb. one piece,

composite spar for a 150 ft. long developmental blade. Also discussed

are a recent SCI design study of a composite blade for the 300 ft.

diameter Mod-2 Wind Turbine, and a new SCI contract from DOE/

NASA for design and possible fabrication of low cost composite blades

for the 125 ft. diameter Mod-OA Wind Turbine.

Key Words: Wind turbine; composite materials; filament winding;

large composite structures; rotor blades.

The Department of Energy, through NASA-Lewis Research Center,

awarded a contract to a team composed of Kaman Aerospace Corporation and SCI to design, fabricate, test and evaluate a potentially

low cost blade for a 300-foot diameter rotor.

Kaman has considerable experience in the design and construction of

composite rotor blades for helicopters. SCI has 20 years of background in the design and fabrication of large, low cost composite

structures, especially those produced by automated filament winding.

Kaman and SCI pooled their talents to propose a low cost blade of

unique composite construction. The primary structural element of

155
the blade is the one piece 20,000 pound spar. For fabrication of this

structure, SCI developed a new, patent pending, filament winding

process called the TFT* process. This process allows rapid spar

production, with low cost materials, and minimum labor. This low

cost spar - comprising 85% of the composite blade weight - is the key

to the low cost of the composite blades

»

Blade and Spar Descriptions The 150 ft. composite wind turbine blade

consists of a filament-wound E-glas s /epoxy leading edge D-spar,

afterbody panels of paper honeycomb with E-glas s cloth skins and a

pultruded trailing edge spline. The afterbody panels are joined to the

spar and spline by bonding. The blade is attached to the hub by a

steel adapter which is bolted to the composite spar,, A steel truss

carries trailing edge spline loads to the adapter. The general arrangement is shown in Figures 1 and 2.

The SCI spar is 135 ft. long (141 ft. as-wound prior to trimming) with

a

! D { shaped cros s - section, (Figure 3) which tapers from approximately 7 ft. chord by 4 ft. height at the root to 2 ft. chord by 7 inches

high at the tip. Minimum wall thickness tapers from three inches at

the reinforced root joint area to i inch at the tip. There is a 15°

linear twist from root to tip. This spar is one of the longest onepiece filament wound structures ever attempted.

The spar material is an E -glas s /epoxy composite in which ninety

percent of the structural fibers are oriented nearly parallel with the

longitudinal axis of the blade. Local reinforcement is provided at

the root by interleaving * 45° and 90° plies between the plies of

essentially 0° unidirection material to achieve an isotropic structure

for attachment of the root end adapter.

Manufacturing Considerations t Fabrication of such large monolithic

composite spars presents some interesting challenges:

Size: The most obvious challenge is the gross size of the structure.

Fortunately the filament winding process is not limited to any particular size. It is merely necessary to provide a sufficiently large

mandrel, winding machine and oven for the part in question. SCI had

previous experience with rocket motor cases and transportation containers up to 22-1/2 ft. diameter by 60 ft. long, so the spar represented an increase in length, but not in diameter, over our in-house

s tate-of-the-art.

Mandrel Deflection: The very long length and slenderness of the spar

was itself a problem, as mandrel deflection could be quite large, and

the rotation during winding could rapidly accumulate a large number

of reverse bending stress cycles. Furthermore the airfoil shaped

cross section of the mandrel caused a change in effective stiffness

* TFT stands for transverse filament tape

156
of roughly 5 to 1 as the chord line rotated from horizontal to

vertical. (This was christened the "flip flop" effect.) The solution

used in this project was careful mandrel design, combined with an

adjustable center support or 'steady rest' to reduce the unsupported

span. Linear finite element analysis, by the 'SAP' Program, was

used to predict the mandrel deflections and stresses, for various

stiffness distributions and support conditions, during the mandrel

design iterations (Figure 4).

Mandrel Extraction: Another problem area associated with the long

slender shape of the spar was that of mandrel extraction. Although

we felt that the natural taper of the spar would allow easy extraction

in production, in spite of the length and the twist of the mandrel, we

proposed a low-risk extraction method for the prototype, which is

described below under 'spar fabrication'. As a further precaution,

a multiple layer release system was used, consisting of alternate

coats of PVA film and fluorocarbon dispersion release agents. A

beneficial effect of the highly axially oriented composite is its ability

to accept high extraction forces.

Winding Pattern: The winding pattern had to produce a low-cost

composite, with high axial strength and stiffness, in a structure

where both wall thickness and circumference taper from root to tip„

The usual approach to this problem would be to use low angle helical

winding. However, this type of pattern requires that all filaments

passing any plane near the large circumference root must also pass

any plane near the small circumference tip. Thus the wall thickness

naturally tends to increase from root to tip, just the opposite to the

desired direction of taper, when helical winding is used. Also some

angle of winding must be used in helical winding, so truely axial

filament orientation is not achieved.

A tapered wall can be approximated by helical winding of layers of

gradually diminishing length. This technique requires movable

dummy head or disc for pattern turnaround part way down the

mandrel. The dummy is moved for each new layer length. The end

of each layer must be tied down and trimmed prior to proceeding

with the next layer. This slows the winding process and increases

labor hours and material waste.

The patent- pending process developed by SCI for the 150 ft. spar

allows high speed continuous winding of 0° filaments to produce

tapered wall on a tapered mandrel, without any trimming or dummy

heads. 90° or 45° filaments can also be applied as required. This

process minimizes the labor and material waste during spar winding.

Tooling and Equipment: The prototype spar was wound on a 150 ft.

long, 30,000 lb. low-cost steel mandrel, which is a stressed-skin/

stringer design similar to aircraft wing structures. An adjustable

steady rest was used to support the mandrel near ijt's center. The

157
mandrel has accumulated approximately 10,000 cycles of rotation to

date with no sign of fatigue failure.

The spar filament winding machine has a 150 ft. carriage stroke and

is capable of rotating the 50,000 lb. fully wound mandrel up to 10

RPMo It is one of the largest known filament winding machines.

An 80 ft. long hot air oven was built for curing the spar in two stages.

The oven is 10 ft. high by 10 ft. wide and is capable of 250°F temperatures. The spar mandrel, winding machine, and oven, shown

in Figure 5, and associated equipment were designed and built by

SCI during the latter half of 1977.

Spar Fabrication: The spar was wound during February 1978 at the

SCI plant in Mira Loma, California (Figure 6).

A half- inch thick shell was first wound on the mandrel, and cured.

The aft web was cut longitudinally and wedged open to free the mandrel. The cut shell was repaired with a spanwise hand-layup doubler

of glass clothe The spar wall was then wound to full thickness and

again cured. Because of the mandrel steady rest, each of the two

basic winding steps described above had to be done in two stages,

with a long scarfed area joining them together so that the total

winding process consisted of four steps as shown in Figure 7.

Mandrel Extraction Procedure: The mandrel tailstock was removed

and the spar was supported on heavy machinery dollies. Large screw

jacks, built into the root end of the mandrel and bearing on a metal

plate at the heavy root end spar buildup, were used to push the spar

off the mandrel until it was sliding freely. A truck mounted winch

was then used to pull the spar completely free of the mandrel and out

of the building. The mandrel was blocked to prevent it from sagging

as the spar pulled clear (Figure 8). During the final stages of extraction, a large forklift and a mobile crane were used to support the

spar.

Spar Handling and Shipping: The spar was lifted by the mobile crane

using a spreader bar and sling centered over the C. G. It was loaded

on a 60 ft. flatbed truck for a short trip to a nearby rail siding

(Figure 9), where it was loaded onto an 89 ft. flat car (Figure 10).

Two 60 ft. flat cars were used as idlers to protect the overhang at

each end of the 89 ft. flat car, but the spar was supported by the

center 89 ft„ car only.

During it's cross-country train trip from Southern California to

Connecticut, the spar was accompanied by the Kaman Program

Manager, who rode in the caboose. Figure 11 shows the 150 ft.

blade nearing completion at Kaman. The finished 150 ft. blade

withstood static flatwise bending loads equivalent to a 165 MPH wind.

158
Future Production: Future production spars will be wound in one continuous process, using a smooth "hard tooling" production mandrel,

which should eliminate the preliminary mandrel freeing step. The

final production process will not require the mandrel steady rest, thus

allowing full-length winding passes without interruption.

In order to achieve the lowest possible production cost, SCI feels

that the entire rotor blade., including the afterbody, must eventually

be built by filament winding. This approach was used by SCI in a

preliminary design study we recently conducted for Boeing Engineering and Construction Division, the prime contractor for the Mod-Z

Program.,

The SCI design approach differs from the previous 150 ft. blade

design described here in that all major composite structural elements

are filament wound. In addition, the final Boeing design for the Mod2 Wind Turbine requires a substantially different rotor configuration

than was anticipated at the start of the 150 ft. Blade Program 2-1/2

years ago. The basic TFT winding process developed by SCI is

employed in the new composite rotor design, but the orientation of

reinforcement in the composite is different, based on data gained

from the structural tests of the 150 fto blade. Figures 12 and 13

show construction and a typical cross section of the Mod-2 composite

rotor.

In April 1979 SCI was awarded a NASA contract to design and fabricate low cost composite rotor blades for the 1 25 -foot-diameter wind

turbine

The program's emphasis will be on developing the engineering and

manufacturing technology necessary to build such large windmill

rotors at low cost,, NASA will install the low-cost composite blades

on a 200 kW Mod-OA Wind Turbine for operational tests.

SCI received the low cost blade contract from NASA's Lewis Research

Center in Cleveland, Ohio, which is the developing agency for DOE

on large-scale windmills 0 This prime contract is expected to lead

to a major extension of rotor design and manufacturing technology.

The first phase consists of blade design and evaluation and the optional second phase involves fabrication of two blades o The SCI low

cost composite rotor blades will have an all-filament-wound structure and will weigh approximately 2500 pounds each. They will be

built by the SCI TFT process, which is also being used, under

license from SCI, to fabricate rotor blades for the DOE /NASA Mod-1

200 Foot diameter Wind Turbine Generator and for the DOE /Rockwell

International 40kW Small Wind Energy Conversion System.

159
i6o
161
162
163

165
166

168
169

171
_i <r o UJUt172
The Effect of Fiber Orientation on the Design Parameters

of Reinforced and Filled Thermoplastic Compounds

Peter J. Cloud and Robert T. Alvord

LNP Corporation

Malvern, Pennsylvania 19355

Abstract

Objective: To determine the limiting design stress and modulus value

which can be substituted into Hookeian stress-strain equations to enable

the engineer to safely design with discontinuous fiber reinforced

injection moldable thermoplastics.

I. Resins A.

B.

II. Additives: A.

B.

Compounds Investigated

6/6 Nylon

Polycarbonate

Glass Fibers - .0004" diameter x .050" long

electrical grade reinforcing fiber

Carbon Fiber - .0003" diameter x .035" long, high

modulus (30MM psi) highly graphitized reinforcing fiber

C. Glass Beads - .0015" - bead

0005" diameter solid glass

III. Compounds A. Nylon 6/6

1. RF-1006 - 30% Glass Reinforced

2. RB-1008 - 40% Glass Bead

3. RF-1006MG - 15% Glass Bead, 15% Glass Fiber

4. RC-1008 - 40% Carbon Fiber Reinforced

B. Polycarbonate

1. DF-1006 - 30% Glass Reinforced

173
Quality Assurance of Manufactured Components

R. D. Margolis

Rockwell International, Plastics Division

Thousand Oaks, California 91360

Abstract

This paper describes progress to date on a joint effort of Rockwell

Plastics and Ford Motor Company to apply non-destructive testing

techniques in the quality control testing of adhesive bonded plastic

truck hood assemblies. The program involves selection of a suitable

sonic testing device to measure adhesive bond integrity along with

laboratory trials to establish procedures and measure the instrument'

flaw detection capabilities. In addition, several plant trials are

underway to test the instrument's viability in an industrial environment and provide a data base for establishing quality control specifi

cations

.
SESSION V

CtLetir-metxi: C. O. Smith

University Of :N"eTDi*etslca,

DEGRADATION DATA OF KEVLAR PRESSURE VESSELS

W.D. Humphrey

Brunswick Corporation/Technical Division

Lincoln, Nebraska 68504

Abstract: Recent hydroburst tests of Kevlar/epoxy composite pressure

vessels showed losses in strength when subjected to extended periods of

high temperature and high humidity. These pressure vessels had served

to monitor the aging characterstics of Rocket Motor Cases during Engineering Development tests. Since Kevlar was a relatively new material

replacing fiberglass in this application, it was considered important

that these results be verified in a controlled experiment utilizing

accelerated aging methods.

To measure these effects, two series of tenth-scale models of Rocket

Motor Cases were fabricated, which included two different resin systems

and two very different design concepts. The accelerated degradation of

Kevlar composites from high temperature and humidity was evaluated by

means of water boil tests for periods of up to ten days.

Hydroburst tests of these aged vessels showed Kevlar to have much less

degradation than fiberglass. Actual values were determined to be

dependent on the resin system used and also dependent upon the dryness

of the composite at the time of test. Wet composites acted as if a

plasticizer was added to the resin system and in the configuration

tested actually regains virgin strength.

Key words: degradation; moisture; Kevlar 49; One-Tenth Scale Motor Case;

water boil; temperature effects; epoxy resins; S-glass.

Introduction: The impact of degradation from long term environmental

storage became a factor at the end of the MX Advanced Technology Program''- when two full-scale Kevlar 49^ units were burst tested at hoop

fiber stress values of 388 and 391 ksi. These units were tested after

approximately one year of uncontrolled warehouse storage. A 10 percent

This work was supported by SAMSO under Contract F04701-73-C-0423 through

the office of J.N. Mason, Major USAF, Propulsion Division, Directorate

of Engineering. Aerojet Solid Propulsion Company, Sacramento, California,

served as prime contractor and Brunswick Corporation, Lincoln, Nebraska,

served as the subcontractor of case development.

2 (r) duPont Kevlar^ Aramid Fiber, Type 969.

177
reduction of strength compared to a baseline hoop fiber stress of 430 ksi.

This latter unit had been tested within three months of fabrication. In

addition to these full-scale case results, significant degradation was

noted with a series of tenth-scale MX pressure vessels stored for periods

of up to two years at temperatures of 100 F and 150°F at 95 percent relative humidity.

This data was contradictory with data reported by other Rocket Motor Case

manufacturers, which showed 6 inch pressure vessels actually getting

stronger after one year of exposure to high temperature and humidity.

At the same time, duPont reported that moisture degradation was strictly

resin dependent. DuPont 's 72 hour water boil data of flex and shear

strengths indicated degradation to be no worse than for fiberglass and

that with "better" resin systems, degradation leveled off at 10 percent.

In order to understand this data, Brunswick initiated an IR & D study of

the effect of accelerated aging of Kevlar pressure vessels. This study

provided new insight and understanding of the degradation mechanisms of

Kevlar composites.

Literature Review: The in-house testing program included a brief

literature review of existing data, which hopefully would assist in

explaining the nature of the failure mechanism. Some of the more meaningful data are discussed in the following sections.

DuPont Design Manual (Reference 1) : The duPont Design Manual documented

several important characteristics of Kevlar 49, which are important to

this study. The manual reported that the effect of tap water on Kevlar

roving or yarn is zero after 24 hours at room temperature. The same

unimpregnated roving loses approximately 2 percent of its residual

strength after 100 hours of water boil.

The moisture gain of Kevlar is reported to be 3.5 to 4 percent at

equilibrium. The time to reach equilibrium is 22 hours at 55 percent

relative humidity and 72°F. The long term use temperature in air is

320°F and decomposition of Kevlar occurs at 930 F.

The loss in tensile strength in air at 100°C (212°F) is 12.5 percent

(460/525 ksi) and at 200°C (392°F) is 25 percent (395/525 ksi).

Similarly, the loss in elastic modulus at 100 C is 8 percent and at

200°C is 12 percent. These data are plotted in Figure i showing tensile

strength and modulus at temperatures for both air and nitrogen

environments

.

The adverse effect of ultraviolet radiation is well documented and is

becoming better understood. This ultraviolet degradation effect

initiates at the exposed surface and progresses inward. An unprotected

woven fabric loses half of its strength in only five weeks of Florida

sun; whereas, a three-strand rope (1/2 inch diameter) has lost 19 percent

of its strength after one year in the same sun conditions. See Table I.

178
Note that the effect stabilizes at 69 percent strength retention after

18 to 24 months.

MX Aging Vessels (Reference 2) : As a portion of the MX Advanced Technology Program, a series of one-tenth scale vessels were fabricated

concurrent with the full-scale rocket motor cases and served as witness

vessels. A long term aging program was conducted, which indicated an

excessive loss of residual strength with one of the resin systems tested.

The test vehicle used was a pressure vessel modeled as a scale version of

the full-scale baseline design and is discussed in detail later. Three

different design/material combinations were studied as follows:

1. All released Kevlar/LRF-209 3

2. All bonded Kevlar/LRF-216

3. Noninterspersed bonded polar and released hoops Kevlar/LRF-216

Where LRF-209 resin is a highly crosslinked epoxy aromatic amine resin

system with high heat distortion temperature (HDT) properties, LRF-216

resin is a low crosslinked epoxy anhydride using a proprietary catalyst

with a relatively low HDT property. The bonded versus release system

referred to a special coating utilized to limit the adhesive properties

of the Kevlar fiber in order to enhance its strength in biaxial strain

fields.

The vessels were proof tested to 3,320 psig and then exposed to 95 percent

relative humidity at either 100°F or 150°F and stored for various lengths

of time until ready for burst. Prior to burst test, the vessels were

dried for 96 hours at 150°F.

A complete tabulation of these data

Figure 2. A summary of degradation

1. All Released, Interspersed, LRF

150°F One Year

Two Years

100°F One Year

Two Years

are shown in Table II and plotted in

as measured follows

:

209

7.8 percent

10.4 percent

3.5 percent

6.0 percent

2. Bonded, Interspersed, LRF-216

150°F

100°F

One Year

Two Years

One Year

Two Years

13.2 percent

28.3 percent

9.2 percent

9.2 percent

LRF denotes Lincoln Resin Formula (Brunswick Corporation)

179
3. Hybrid, Noninterspersed, LRF-216

150°F 16 months 22.5 percent

Moisture Effects on Structural Resins: Many studies (References 3-5)

have shown that absorbed moisture results in significant degradation of

the matrix sensitive mechanical properties, particularly at higher

temperatures; whereas, fiber controlled composite properties are

relatively unaffected. Most investigations agree that this degradation

property is attributed to the plasticizing effect of moisture on the

resin modulus and lowers the glass transition temperature. Water behaving

as a plasticizing agent disrupts the strong hydrogen bond present in

highly polarized epoxy resin systems. This is reversible to some extent

and can promote a change in failure mode between wet and dry specimens.

High heat distortion resin systems are not as significantly affected

at high temperature moisture exposure as are the lower heat distortion

resins. It is interesting to note that one investigator (Reference 5)

attributed the increased strength in wet laminates to swelling of the

resin matrix, which (in certain configurations) relieves the residual

interlaminar shear stresses.

IR & D Test Program—Test Vessel: The use of pressure vessels, which

contain a biaxial stress field, has proven to be the most reliable means

of establishing valid design allowable data. Brunswick has developed

extensive burst strength data for Kevlar 49/epoxy composites using the

tenth-scale second stage MX vessel. The vessel is designed to fail at

mid-cylinder in a controlled manner. Design verification and reliability

testing indicate that the mode of failure and achieved level of burst

pressure are very repeatable. This vessel has been used on many programs,

which accounts for the large data base. Subtle processing parameters,

such as hoop winding terminations and dome reinforcement locations, have

been studied and their influence on performance is controlled. This

makes it possible to isolate effects of the variable under study with a

minimum number of replicate specimens. A design summary is shown in

Figure 3.

Test Plans (LRF-092 Resin System) : The initial program consisted of 21

vessels all wound with LRF-092 resin. The LRF-092 resin is a highly

crosslinked epoxy anhydride/ tertiary amine resin system with a good heat

distortion temperature. The resin is widely used on many of Brunswick's

pressure vessels and, thus, dictated a strong need for degradation data

for in-service applications.

Nine of these first 21 vessels were selected by random numbers for long

term storage tests, where six vessels were of the released configuration

and three of the bonded configurations. The remaining vessels were proof

tested and assigned for testing as follows.

180
Bonded Released

(2) virgin burst

(2) 72 hour water boil and burst

(2) 240 hour water boil and burst

(2) virgin burst

(2) 72 hour water boil and burst

(2) 240 hour water boil and burst

Test Results (LRF-092 Resin System) —Bonded Vessels: The burst results

of all LRF-092 vessels are summarized in Table III. The virgin strength

of bonded vessels was established from two vessels which demonstrated

identical strengths (4,490 psig and 346 ksi hoop stress); this was also

in excellent agreement with the hoop stresses 340 to 345 ksi established

by bonded vessels during the MX program.

The 72 hour water boil vessels tested at 4,490 and 4,590 psig, indicating

zero degradation in these two vessels, while the 240 hour water boil

vessel burst at 4,120 and 4,400 psig indicating 92 and 98 percent strength

retention. As' the results show, the degradation of Kevlar/LRF-092 vessels

was much less than expected based on the MX witness vessel tests at

Released Vessels: The burst test results of these vessels are also

summarized in Table III. Virgin strength was assigned at 5,600 psig

(431 ksi) , based on the results of S/N 762 and based on the mean value

of the MX statistical series (Reference 2).

The 72 hour water boil vessels burst at 5,690 and 5,630 psig, indicating

a net improvement in performance which parallels the results of the

three day water boil tests of bonded vessels.

An unexpected phenomenon occurred during cool-down from water boil with

S/N 747 and 748 released hoop vessels. A partial vacuum was created

inside as the vessels were cooled (the ports had been sealed for protection during boiling) . This caused one dome on each vessel to buckle

away from the released fiber cylinder. It was elected to test these

vessels and to evaluate the effects of buckled domes. Both vessels

failed prematurely in areas where the buckling had occurred (see Figures

4 and 5). Consequently, two vessels which had been selected for

environmental storage, were shifted to the boil tests as replacements.

Following ten days of water boil, with a modified vent plug installed,

vessels S/N 758 and 764 were burst tested. The vessels failed at

5,580 and 4,600 psig, indicating residual strengths of 100 and 82 percent

respectively (the dome/knuckle failure of S/N 764 is not a normal failure

mode and generally indicates nontypical failure levels). In any case,

the test results for released fiber vessels substantiate the results of

bonded vessels. That is, both showed degradation of Kevlar/LRF-092

vessels to be much less than expected.

Environmental Studies: Long term aging studies are currently underway

at Brunswick in an environmental chamber at 150°F and 95 percent

relative humidity. The vessel assignment was altered to provide two

Aerojet.

181
replacement vessels for the 240 hour water boil matrix. The dates of

installation and S/N of these vessels are as follows:

S/N Configuration Date Installed

753

755

757

Bonded - proofed & coated

Bonded - proofed

Bonded - not proofed

Released - proofed & coated

Released - proofed

1/10/78

1/10/78

1/10/78

1/10/78

1/10/78

761 - 763

759 - 560

In addition, one spool of bare Kevlar roving from the same lot used to

wind the above vessels was placed in storage January 18, 1978. An

additional spool of bare Kevlar roving was placed in storage September 18,

1978, and additional spools are planned for six month intervals.

It is planned to store these vessels and spools continuously for 18 and

24 months and then hydroburst and perform strand tensile tests to establis

degradation correlation with accelerated aging tests.

Test Plans (LRF-216 Resin System) : The accelerated aging tests of

Kevlar/LRF-092 vessels showed only minor degradation occuring after 240

hours of water boil and no effect at all was noted for 72 hour water

boil. These results were surprising and indicated that the LRF-092

resin system provided excellent moisture resistance for the pressure

vessels. The MX witness vessels with LRF-216 resin indicated excessive

degradation (up to 28 percent) after exposure to an environment of 150°F

and 95 percent relative humidity; it was important to determine if this

resin system was alone responsible for these poor results. Therefore,

twelve additional vessels (six released and six bonded) were added to the

program.

Test Results: A variable not previously considered was incorporated into

the study. This variable was the condition of the composite (wet or dry)

at the time of burst test. It was suspected that the drying operation of

96 hours at 150°F performed on the MX witness vessels may have altered

the performance of those vessels. This tendency was established by the

initial LRF-216 tests. S/N 774 and S/N 765 were tested shortly after

cooling to room temperature and achieved 95 and 96.5 percent of virgin

strength, while S/N 772 and 769 (tested following three days drying at

room temperature) achieved only 90 and 92.5 percent of virgin strength.

See Table IV summary.

Upon reviewing the significance of these results, it was elected to

deviate from the original test plan and delete the 72 hour water boil

tests and utilize the remaining vessels to screen the effect of the wet

or dried laminates. Hence, the next four vessels were subjected to ten

day water boil. Two of these vessels were tested wet and two tested

after three days of drying at 150°F. The dried vessels, S/N's 773 and

770, burst at 87 percent and 91 percent of ultimate, while the "wet"

vessels, S/N 771 and 776, burst at 102 and 98 percent of ultimate. The

182
relationship continued to hold true, indicating that "dried" laminates

tested lower than "wet" laminates.

The four remaining vessels were subjected to the same wet/dry screening

test, except the drying time was extended to match the four days actually

used by Aerojet. The dried vessels S/N 776 and 768 tested at 90 and 97

percent of virgin strength, while the "wet" vessels S/N 775 and 767 tested

at 100 percent and 87 percent of virgin values. The lower burst value,

4,870 psig, of S/N 767 was uncharacteristic, but the mode of failure

(dome/knuckle/cylinder) was nontypical indicating that the vessel was*

suspect

.

Summary data for all vessels fabricated and tested during the program

are shown in Table V. Apparent thicknesses at three stations in the

cylinder using pi tape measurements, total assembly weights prior to

environmental exposure, date of winding, and test results are all

summarized in this table.

Discussion: The results of this program showed, first, that Kevlar does

not degrade appreciably in pressure vessels following ten days of water

boil. Our literature search had prepared us somewhat for these results.

As mentioned, duPont measured only 2 percent degradation of bare Kevlar

fibers after 100 hours of water boil. Greater degradation had been noted

as a surface effect from ultraviolet radiation. This surface effect could

destroy thin laminates, but in thicker sections (ropes or cables), equilibrium was reached at about 70 percent strength retention. Composite

vessels are protected from sunlight by the resin matrix, and the effect

of ultraviolet exposure is expected to be low. Further, the vessels

tested by Aerojet were conditioned in a protected environment.

Returning to our original conclusion that Kevlar in pressure vessels does

not appreciably degrade, we found similar performances in both LRF-092

and LRF-216 resins. If one low value, 82 percent, can be excused from

the LRF-092 results, and 87 percent excused from the LRF-216 results

(both nontypical failures) , we have good correlation with less than ten

percent degradation. The vessels fabricated using LRF-216 resin indicated

that performance was affected by the test method. Dried vessels tested

at an average of 92 percent of virgin strength and "wet" vessels tested

at 99.5 percent of virgin strength. See Figure 6. The nontypical

failures are unique to water boil tests and did not occur in the virgin

tests of Browning and Whitney (Reference 3) . The dome failures appear to

be caused by water propagating between layers, perhaps originating at the

fitting interface. Pending tests with the high temperature/humidity

exposed vessels could resolve this question.

Our literature search also sheds light on the wet versus dry Kevlar

composites. Brunswick MX work (Reference 2) concluded that water acted

as a plasticizing agent, disrupting the bonding in high polar epoxy resin

systems. The reversibility of the water absorption effect and the resulting

change in failure modes from wet to dry were explained. The higher heat

183
distortion resins were shown to be less affected by moisture than were

the lower heat distortion resins where water boil affected weight gains.

It has been noted that LRF-209 vessels degraded 10 percent after two years

of 150°F and 95 percent relat ive humidity; but LRF— 216 vessels degraded

over 28 percent for the same tests. It follows that the heat distortion

temperature for the resins, as cured in these vessels, was 310°F for 1 LRF-209 and only 150°F for LRF-216. The higher heat distortion resins

actually are more densely crosslinked and, thus, have fewer reactive

sites where plasticizing can occur. This difference in heat distortion

temperatures helps explain the performance differences in vessels fabricated from these two resins.

The 1/10 scale MX vessel was designed specifically to fail in the midcylinder section by over-reinforcing the dome areas. By isolating the

Kevlar fibers using the released fiber mechanism, pure tensile loads were

achieved in the fiber while strained in a biaxial strain field. Without

this releasing mechanism, the Kevlar fibers would be subjected to biaxial

loads which could transversely pull the fibers apart. This is the primary

reason for the 25 percent performance gain in released fiber vessels.

It was demonstrated during the MX program that a very flexible, highelongation resin, acted very much like released fibers by allowing

deformation within the matrix rather than at the fiber/resin interface.

In fact, no additional improvement was measured by combining the above

"chicken fat" resin with released fibers. It is believed that the

differences measured in this study between wet and dry resins parallel

that observation and that "wet" vessels in fact represent a plasticized

resin system.

For comparison purposes, the results of this study are plotted in Figure 7

(previous Brunswick data, Reference 6) showing the improved performance

of Kevlar 49 over two types of S-glass. Note that the strength loss in

Kevlar vessels, regardless of the resin system used, is much less than

with fiberglass when utilizing the same test vessels and test procedures.

Conclusions

:

1. Degradation in Kevlar pressure vessels is much less than in fiberglass pressure vessels and is a function of the resin system used.

High heat distortion resins, such as LRF-092 and LRF-209, retain 95

percent of virgin strengths after water boil; Kevlar with a lower

heat distortion resin (LRF-216) still performed well, retaining 90

percent of virgin strength.

2. It was demonstrated by this study that for a Kevlar vessel designed

for membrane critical loading (i.e., mid-cylinder in the MX 1/10

scale) the dryness of the composite at the time of test can influence

performance. This is particularly true when the resin has a low

heat distortion temperature such as LRF-216, where water boil alters

181+
the characteristic of the resin. Original strengths were retained in

bonded LRF-216 vessels when testing, while the composite was still

wet (plasticized)

.

3. The literature search and a review of previous Brunswick data point

out several characterstics of Kevlar that should be considered in

establishing performance standards.

a. Water boil of bare fiber has little effect on performance

(2 percent after 100 hours)

.

b. Appreciable loss in strength occurs at the fiber surface from

ultraviolet radiation and/or oxidation. This is primarily a surface effect and its extent is thickness dependent.

c. The degradation of Kevlar from temperature alone is significant,

limiting the time of exposure and preventing high temperature

(+200°C) usages of this fiber.

d. The effect of moisture on a resin system is reversible and can

be utilized in certain applications to enhance performance.

REFERENCES

1. Kevlar 49 Data Manual, E.I. duPont and Company, Wilmington, Delaware,

revised October, 1977.

2. MX Program Chamber Development, Final Report - Phase II, Brunswick

Corporation to Aerojet Solid Propulsion Company, Contract S-117039.

3. Browning, C.E. and Whitney, J.M. , "The Effects of Moisture on the

Properties of High Performance Structural Resins and Composites,"

Fillers and Reinforcements for Plastics , Advances in Chemistry Series

134, 1974, pp. 137-148.

4. Tompkins, S.S., "Analysis of Moisture Absorption and Diffusion in Fiber

Reinforced Polymeric Resin - Matrix Composite Materials," Ph.D.

dissertation, University of Virginia, May, 1978.

5. Lee, B.L., Lewis, R.W. , and Sacher, R.E., "Environmental Effects on

the Mechanical Properties of Glass Fiber/Epoxy Resin Composites,"

Army Materials and Mechanics Research Center, June, 1978.

6. Reiners, A.E., "Characterize Type 449 Roving," Design and Development

Memo #289, Brunswick Corporation, Lincoln, Nebraska, September, 1978.

185
THE EFFECT OF TEMPERATURE ON THE

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KEVLAR® 49 ARAMID

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186
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189
190
191
FIBERGLASS DEGRADATION

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KEVLAR/LRF- 216

50

12 16 20 24 28 32 36 40

TIME IN TEMPERATURE/HUMIDITY, WEEKS

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TIME IN WATER BOIL DAYS

FIGURE 7

192
TABLE I

(FROM REFERENCE 1)

ULTRAVIOLET STABILITY OF KEVLAR 49 ARAMID FIBER

KFVI AR dQ PRfinilPT FORM

BREAK LOAD STRENGTH

RFTflTMFn

LBS/ IN (kN/m) %

FARRTf WOVFN FROM "3fiD DFNTFR

(420 DTEX) YARN

4.5 MIL (0.11 mm) THICK

UNEXPOSED 300 52

FLORIDA* SUN, 5 WEEKS 154

LB

27 51

3-STRAND ROPE, 1/2 IN.

(13 mm) DIAMETER

UNEXPOSED 14,400 64,050

FLORIDA SUN, 6 MONTHS 13,000 57,830 90

12 MONTHS 11,600 51,600 81

18 MONTHS 9,950 44 ,260 69

24 MONTHS 9,940 44,220 69

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TABLE III

KEVLAR AGING WITH LRF-092 IN 1/10 SCALE MX VESSELS

BONDED

240 HR. WATER BOIL

72 HR. WATER BOIL

VIRGIN BURST

STORED (W. PROOF

& COATED)

STORED (W. PROOF)

STORED (NO PROOF)

S/N

745-746

749-750

754-756

753-)

755

757

f "At "k

BURST PRES.

(PSD (KSI)

4120-4400 317-339

4490-4590 346-354

4490-4490 346-346

STR. RETENT

(1)

95

101

100

RELEASED CIRCS

240 HR. WATER BOIL (1ST) 747-748 4370**-3150* 337-243 N/A

240 HR. WATER BOIL (2ND) 758-764 5580-4600 430-354 100-82

72 HR. WATER BOIL 751-752 5690-5630 438-433 101

VIRGIN BURST 762 5600- 431 100

STORED (W. PROOF 761-763

& COATED)

STORED (W. PROOF) 759-760 1

***

BUCKLED DOME FROM VACUUM PRESSURE

**8UCKLED LINER FROM VACUUM PRESSURE

***T0 STORAGE (1/10/78) AT 150°F and 95° R.H.

195
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197
DESIGN ASSURANCE OF A LEAK FAILURE MODE

FOR COMPOSITE OVERWRAPPED METAL TANKAGE

W.W. Schmidt

Brunswick Corporation/Defense Division

Lincoln, Nebraska 68504

Abstract: Assuring a leakage failure mode requires that the outer composite shell be designed such that its minimum cyclic fatigue life is

approximately an order of magnitude greater than that of the metal liner.

This requires making an accurate fatigue life prediction for both the

composite shell and metal liner. It is further possible to determine

the critical flaw size for unstable crack growth within the metal liner

by application of fracture mechanics principles. A semi-empirical

formula has been developed for this purpose since the stresses induced

in the metal liner upon initial pressurization exceed the yield strength

of the liner material in all cases, which violates a basic assumption of

linear elastic fracture mechanics. It is often found that the critical

flaw depth is greater than the thickness dimension of the metal liner.

In these instances, leakage will occur prior to unstable crack growth.

For designs in which unstable crack growth can occur, a dynamic analysis

has been developed. This analysis provides a means of assessing the

magnitude of shock induced stresses in the composite shell due to the

sudden loss of the liner's load carrying capability. Details of these

design procedures are provided.

Key words: composite materials; filament wound tankage; leakage failure

mode; linear elastic fracture mechanics; critical flaw size; cyclic

fatigue life; metal liner.

Filament wound tankage often requires the use of a metal liner, i.e.,

permeation barrier, due to a requirement for a low allowable leakage

rate, fluid compatibility considerations, or exposure to extreme temperatures. A desirable feature of metal-lined composite tankage, in

addition to light weight and improved damage tolerance, is an inherent

benign failure mode under normal conditions which consists of leakage

of the contents without catastrophic rupture. Design procedures to

ensure that a new or existing design will possess this feature have been

developed.

The principal means of assuring a leakage failure mode is to design the

outer composite shell so that its minimum cyclic fatigue life is at

least approximately an order of magnitude greater than that of the metal

liner. This is facilitated by the fact that fiber reinforced composites

do naturally possess much greater fatigue lives than metals in the range

of strain amplitudes common to this class of structures. For example,

198
cycle lives of over 100,000 pressurization cycles have been demonstrated

for Kevlar* 49 reinforced elastomeric-lined pressure vessels which

cycled from zero percent to 55 percent of the rupture pressure. This

corresponds to a strain amplitude which is close to one percent.

Once a designer has performed a complete stress analysis of the tank

under study, assessment of its fatigue lives for both the composite and

metal liner are straightforward. Graphical data presentations for

commonly used composite overwrap materials are available in references

[1] and [2]. These empirically derived relationships require only that

the designer know the percentage of available strength which is being

utilized at the operating condition which is being evaluated. Low cycle

fatigue as it pertains to metals is a broad field for which an abundance

of theory and data presentations are readily available. Two comprehensive texts, references [3] and [4], will provide a designer seeking

knowledge of fundamentals a clear understanding of cyclic fatigue

phenomenon.

At the stage of the design effort at which the fatigue lives are known

to exhibit the desired ratio of composite life to liner life, the

designer may still not possess sufficient information to assure that a

liner fatigue crack will not initiate a "premature" composite failure.

In order to achieve confidence, the designer must either ensure that any

fatigue crack will propagate through the thickness without attaining

critical size, or ensure that the composite shell possesses sufficient

strength to contain pressure and shock loading released due to a sudden

complete liner failure.

All metal-lined composite tankage is subjected to a proof test as a

final operation in the normal manufacturing sequence. In addition to

demonstrating structural integrity, the proof test develops a permanent

compressive stress in the metal liner (and a usually quite small tensile

stress in the composite overwrap) which permits more efficient structural

utilization of the metal than in a monolithic tank. Because the proof

test plastically deforms the metal liners, the analysis of crack

behavior at the proof stress level is beyond the realm of linear elastic

fracture mechanics, thus requiring a semi-empirical approach, references

[5] and [6] , to assess this regime from a fracture perspective. It is

also quite common to utilize aluminum alloy liners which plastically

deform at the maximum expected operating pressure, MEOP, which

necessitates the use of the approach described.

(§) * duPont Kevlar^ Aramid Fiber, Type 969.

199
In the plastic regime near TUS, the fracture behavior of part-through

cracks can be represented by:

= TUS (1 He

2

) (1)

where,

and

S

f = TUS = H =

c = Kf =

liner stress

tensile ultimate strength

.112 (TUS/Kf

)

4

half-crack length

fracture toughness

.

The above can be rewritten as:

n "I 1/2

c = 1 (1 - -1 ) H

v TUS y (1A)

Application of the stated relationship can be used to estimate the

critical half-crack length associated with an inelastic liner stress

state.

If the liner stress at the pressurization condition being evaluated is

less than the tensile yield strength, an estimate of the critical flaw

size can be made by applying the following equation which defines plane

strain fracture toughness:

where,

K = plane strain fracture toughness

and S^ and c are as previously defined.

Usually, the liner thickness will not be sufficient to satisfy the

criteria for plane strain conditions to exist. In these instances when

the sample is thin and the degree of plastic constraint acting at the

crack tip is minimal, plane stress conditions prevail and the metal liner

will exhibit greater toughness. Therefore, the actual value of c which

is calculated is not valid; however, a conservative value is quickly

determined. This check is simplified since values of plane strain

fracture toughness, K.^.^, are generally available for most commonly used

metal liner materials since this is a material constant which is not

dependent on specimen geometry. The plane stress fracture toughness,

Kc

, is dependent on both the material and specimen geometry; and,

therefore, a single value of this parameter for a given material does

not exist.

200
The designer will usually determine as a result of the approximation of

the half-crack length, c, which is approximately equal to the crack

depth, a, that the dimension is greater than the metal liner thickness.

This implies that the critical crack size is relatively large in comparison to the thickness and that any crack which is experiencing growth

due to repeated loading will attain a breakthrough condition, i.e.,

leakage, prior to unstable crack growth. If the designer suspects

unstable crack growth could occur prior to breakthrough and subsequent

slow leakage, a dynamic analysis of the composite shell should be

performed.

The approach described requires that the period of natural vibration for

the composite shell be calculated. Next, the time required for the

pressure induced liner loading to be relieved is estimated. A comparison

of the time of energy release and the period of natural vibration for

the composite shell will allow determination of the magnitude of the

shock induced stress increase. Finally, the adequacy of the structure to

withstand the potential occurrence of sudden liner failure is assessed.

The period of natural vibration for a complete spherical membrane shell

can be found by applying the following equation:

where,

T = period of natural vibration

R = radius

E = tensile modulus of elasticity of composite

v = Poisson's ratio of composite

p = density of composite

and, g = acceleration of gravity

Relationships for cylindrical shell geometries as well as shells with

bending resistance are contained in reference [7]

.

The crack tip velocity approaches the speed of sound in the material

which is

:

c = 1/^

where, c = speed of sound

and, E, g, and p are as defined above

It is necessary for the designer to assume an approximate distance at

which the crack propagation would terminate. This will depend largely

on the amount of available strain energy which is stored in the metal

201
liner. It is usually discovered that the time of energy release is

substantially less than the period of natural vibration for the composite

shell. Therefore, the effect of this sudden shock is to induce a dynamic

load into the composite which is up to two times the amount of contained

pressure loading which is carried by the metal liner prior to failure.

The shock spectrum for a step-ramp function as a function of time, t,

,

divided by the period of natural vibration, x, is presented graphically

in reference [8], Once the amount of additional load is determined,

the designer can easily calculate the corresponding maximum stress in

the shell and demonstrate design adequacy.

The sequence of design events which have been described is shown in

Figure 1. Both possible redesign phases would undoubtedly require consideration of the addition of composite overwrap material, since in each

instance, lowered stresses within the composite shell would be needed to

permit confirmation of the desired leakage failure mode.

Finally, the application of the above described procedure will be

illustrated by considering a possible design. Representative performance

requirements and a physical description of the selected design are contained in Table I. Also contained in Table I are the liner and composite

stresses at proof pressure and at maximum operating pressure. Cycle

lives for both the composite and liner have been calculated and these

are also contained in Table I.

A comparison of the cyclic fatigue lives indicates that the structure's

life is limited by the metal liner.

Fatigue life composite _ 20 ,000 _

^ Fatigue life liner 550

The values of liner stress contained in Table I are not the principal

stresses, but are rather the values of stress intensity which are

determined by applying the following equation:

s

f -

2 2 2

(a - O + (a. - a ) + (a - a ) m 9 9 z z m

1/2

(5)

where,

a , a A , and a are stresses in the meridional, circumferential

m, 0,. , ,.z . . and radial directions, respectively

The use of an effective stress is warranted due to the presence of a

biaxial stress field. It will be noted that at the proof pressure level,

the liner stress exceeds the yield strength of the liner material. This

requires that equation (1A) be used to assess the ability of the proof

test to screen critical flaws. Application of equation (1A) yields:

H = .112 44 ksi/26.6 ksi -'tftn.

202
REDESIGN

~~r~

NO

REDESIGN NO

NO

1

CALCULATE

STRESS DUE

TO SHOCK

Eq. (3) (4)

[7-8]

STRESS

analysis/

life deterh

minati on

YES

ESTIMATE

c Eq. (2)

CALCULATE

c ~ a

Eq. (1)

Figure 1. Design Flow Diagram

205
TABLE I

DESIGN PARAMETERS

Example Tank

Physical Description

Shape

Composite material

Liner material

Radius (mean) , R

Composite thickness, t

Liner thickness, t^

Proof pressure, P'

Operating pressure, P

Liner Material Properties (typical)

Tensile yield strength, TYS

Tensile ultimate strength, TUS

Fracture toughness,

Modulus of elasticity, E

Poisson's ratio, v Density, p

Performance Characteristics

Liner stress/at proof, S^'

(at oper. , S

f

Composite stress /at proof

(at oper.

Percent load carried ( at proof

I

«

by liner

Lner cycl

Composite cycle life

Liner cycle life, O-oper.-O, N f0

-6,

at oper.

>er.-0,

0-oper

,

N

fc

Spherical

S-glass/ epoxy

6061-T6 aluminum alloy

5.14 inches (130 mm)

.460 inches (11.7 mm)

.050 inches (1.27 mm)

9,000 psig (62.05 MPa)

6,000 psig (41.34 MPa)

38 ksi (262 MPa)

44 ksi (303 MPa)

26.6 ksir-YinT (29.2 MPaYm"

)

10,000 ksi (68.9 GPa)

.33

.10 lb. /in. (2.77 gm/cc)

41.6 ksi (286.8 MPa)

32.1 ksi (221.3 MPa)

46.7 ksi (322.0 MPa)

30.7 ksi (211.7 MPa)

7.0 percent

8.5 percent

550 cycles

20,000 cycles

20k
= .838 in. 2 (.0013 mm

2

)

and

c = 838 in.

-2 1 - 41.6 ksi

44 ksi

1/2

= .255 in. (6.48 mm)

From the approximate flaw size calculated, it is evident that the proof

test will not screen defects with dimensions less than the wall thickness. Consequently, the assurance of safe life during subsequent

cycling is dependent on a realistic evaluation of crack growth. It is

also evident that for the operational cycles after a proof test, the

critical failure condition is crack breakthrough and subsequent liner

leakage.

If the previous design were altered according to the tabulation below,

a new assessment would be necessary.

Composite Material

Composite Thickness

Liner Thickness

Design

per Table I

S-glass/ epoxy

.46 in. (11 . 7 mm)

.05 in. (1.27 mm)

New

Design

Kevlar 49/ epoxy

.20 in. (5.08 mm)

.25 in. (6.35 mm)

The new performance characteristics are as follows

Liner stress^at proof, S^'

i at oper. , S

f

Composite stressfat proof

{at oper.

Percent load carried ( at proof

by liner v at oper.

Liner cycle life, O-oper.-O

Composite cycle life, O-oper.-O

42.0 ksi (289.6 MPa)

18.8 ksi (129.6 MPa)

74.3 ksi (512.3 MPa)

61.0 ksi (420.6 MPa)

35.7

20.8

2,000

40,000

Again, a comparison of the cyclic fatigue lives indicates that the

structure's life is limited by the metal liner.

Fatigue life composite _ 40 , 000 _ Fatigue life liner 2,000

Application of equation (1A) for this design yields:

1/2

= .233 in. (5.92 mm)

In this instance, the critical flaw depth (assumed equal to the halfcrack length) is seen to be approximately equal to the liner thickness.

c = .838 in.

-2 1 - 42 ksi

44 ksi.

205
To determine the condition in which a liner failure could cause rupture

of the composite, a dynamic analysis of the composite shell will be

performed by applying equations (3) and (4). The period of natural

vibration of the composite shell is:

, —11/2

(1 - .27) .050 lb./in. J

x = 2tt(5.14 in.)

2(5.14 x 10 b psi) (32.2 f t . /sec. 2

) (12 in. /ft.)

= .98 x 10 sec,

The crack tip velocity approaches the speed of sound in the liner

material which is

1/2

c = (10 x 10

6 psi)(32.2 ft./sec. 2

)(12 in. /ft.)

(.10 lb./in. J

)

= .197 x 10 in. /sec. (5.00 x 10 mm/sec.)

It is assumed that the pressure is released when the crack has traveled

one-fourth of the circumference from the site of failure initiation.

The time required is:

ttR 2 time, t, - — 1 c

tt(5.14 in.)

2(.197 x 10 6 in. /sec.)

= .41 x 10

-4

sec

.

Because the time of energy release is substantially less than the period

of natural vibration for the composite shell, the effect of this shock

is to induce a dynamic load into the composite which is up to two times

the amount which is carried by the liner prior to failure. The shock

spectrum for a step-ramp function for t^/x = .41/. 98 would indicate a

load factor of 1.8. Composite stresses at the proof pressure are

tabulated below:

Nominal Composite Stresses at 9,000 psig (62.05 MPa)

Prior to liner failure

At failure, no shock

At failure, maximum shock

Design allowable stress

74.3 ksi (512.3 MPa)

115.6 ksi (797.0 MPa)

148.6 ksi (1,024.6 MPa)

104.0 ksi (717.1 MPa)

M.S. = - 1 = -.30

148.6

Based on the above values of nominal composite stress, it is concluded

that composite rupture would be possible at a pressure less than the

stated proof pressure if the liner were to suddenly fail. Since this

design provides a low overall margin of safety, this tank would not be

considered safe and complete redesign would be deemed appropriate.

206
The above calculations have been included to familiarize the reader with

the stated design techniques. It has been the author's experience, that

all metal-lined composite tankage which has been designed and manufactured by the Brunswick Corporation for which similar analyses have been

performed, have clearly satisfied the design requirements which provide

assurance that the desired feature is present. These findings have

been confirmed in all cases by destructive testing of replicate test

articles. The described design approach is deemed highly valuable as a

means of providing quantitative substantiation of a design claim and a

means of predicting actual composite behavior at elevated stress levels

at which time a composite failure could be influenced by the presence

of a metal liner.

REFERENCES

1. Humphrey, W.D., and Foral, R.F., "Fatigue and Residual Strength

Tests of Fiberglass Composite Pressure Vessels," ASME paper

74-PVP-47, presented at the Pressure Vessels and Piping Conference,

Miami Beach, Florida, June 24-28, 1974.

2. Kevlar* 49 DP-01** Data Manual , E.I. duPont de Nemours and Company,

Incorporated, Textile Fibers Department, Wilmington, Delaware, and

updates.

3. Manson, S.S., Thermal Stress and Low-Cycle Fatigue , McGraw-Hill,

New York, 1966.

4. Sandor, B.I., Fundamentals of Cyclic Stress and Strain , The

University of Wisconsin Press, Madison, Wisconsin, 1972.

5. Bockrath, G.E., and Glasco, J.B., "A Theory of Ductile Fracture,"

MDC-G-2895, McDonnell Douglas Astronautics Company, August 1972.

6. Feddersen, C.E., "Fracture Behavior of Surface-Cracked Tension

Elements," AIAA Journal , Volume 9, No. 12, December 1971,

pp. 2309-2313.

7. Kraus, H. , Thin Elastic Shells , John Wiley and Sons, New York,

1967.

8. Thomson, W.T., Vibration Theory and Applications , Prentice Hall,

Incorporated, Englewood Cliffs, New Jersey, 1965.

* duPont trademark.

** Temporary designation for generic assigned by FTC.

207
Composite Materials in Recreational Equipment

David Ratchford

Marshall Vega Corporation

Marshall, Arkansas 72650

Abstract

The general requirements of materials for recreational equipment are

discussed, these being high strength-to-weight ratio, fatigue resistance,

the ability to store and release energy, durability, aesthetic acceptability, and manufacturing capability.

It is developed, by engineering analysis, that composites of glass

fibers, graphite fibers, or their combination, bonded into a resinous

matrix of thermo-setting polyester or epoxy are uniquely suited to these

purposes. These composites have extremely high strength-to-weight ratio,

excellent durability and excellent fatigue characteristics, all unaltered

through wide environmental extremes. Additional advantages are that the

flexural characteristics of the material can be designed, through stressdirected fibers, to suit the purpose; and that glass fibers are nonyielding, i.e., yield and ultimate strength are the same, meaning that

lower safety factors can be used. Such composites are ideally suited

for handling, for the storage and release of energy and for the durability required of rough treatment.

An important and large scale application is limbs for bows used in

archery. Design parameters would be light weight, high strength, fatigue

resistance and quick energy release.

The limbs may be solid FRP composites, or FRP composites in the form

of "backing" which is laminated to wood cores. In either case the limb

is formed to a taper in order to minimize weight and to reduce inertia

and present a low profile to wind resistance upon release. Load requirements are almost purely flexural, therefore fiber orientation is

100% parallel to the longitudinal axis of the limb. Glass or graphite

fiber content is usually in the order of 70%. The limb has a flexural

strength of approximately 150,000 psi, and a flexural modulus of 5.5 X

106 psi. High heat-distortion resins are used so that the bow can be

stored in automobile trunks where temperatures can go above 150°F. An

aesthetically attractive, abrasive-resistant finish is usually applied

to the limb.

Additional applications in archery are bows molded entirely in one piece,

and arrow shafts.

208
SESSION S

Chairman: IR, . El . Maringer

Battelle Columbus Laboratories

HOLOGRAPHIC NONDESTRUCTIVE EVALUATION OF SPHERICAL

KEVLAR/EPOXY PRESSURE VESSELS

D. M. Boyd and B, W. Maxfield

Lawrence Livermore Laboratory, University of California

Livermore, California 94550

Abstract: Three spherical kevlar/epoxy pressure vessels were evaluated using holographic interferometry . The holographic interferograms

provide information on the failure modes, displacement profile, and

possible fiber damage. The holograms show a symmetric anisotropic

displacement pattern even though the vessels failed due to a leak in

the aluminum mandrels. The presence of a biconvex fringe pattern

found during the testing of vessel three is believed to indicate

matrix crazing. This information would be useful for burst-type

failure analysis. Holographic interferometry has been shown to supply

information on the characteristics of kevlar/epoxy pressure vessels

under stress.

Key words: composites; holography; interferometers; matrix crazing;

nondestructive testing; pressure vessels.

Introduction

Although the properties of individual fibers and fiber bundles provide

the basic information needed to design fiber reinforced pressure

vessels, it is only through a combination of destructive and nondestructive evaluations of completely fabricated units that true

vessel performance can be assessed. Holographic interferometry is

one of the nondestructive testing techniques which has been used

successfully in evaluating composite pressure vessels (1,2). The

anisotropic expansion of these vessels is easily measured using the

wide field of view surface displacement information available from

holographic interferograms.

This paper reports on the use of holographic interferometry to examine

three thin walled kevlar/epoxy pressure vessels. The information so

obtained was evaluated both quantitatively and qualitatively; only

minor differences in surface displacement patterns were evident. It

is, however, demonstrated that both local matrix crazing and the

anisotropic expansion of the vessels are easily measured by holographic interferometry. These results solidify the potential of

holographic nondestructive testing for evaluating the strength and

overall quality of kevlar/epoxy pressure vessels.

211
Pressure Vessel Design Information and Test Procedure

The three thin walled pressure vessels that we examined were fabricated from the organic fiber Kevlar 49 using the epoxy system DER 322

(Dow Chemical )/Jeffamine T-403 (Jefferson Chemical) in a 100/44.7

ratio. The 114 mm outer diameter pressure vessels were wound over

1 mm thick 1100 aluminum (hydroformed) spherical mandrels. Each

mandrel has electron beam welds at the waist, the fill tube and the

opposite pole plug. Further details on the construction of these NASA

experimental gas storage vessels are contained in Ref. 3.

Holographic testing of the pressure vessels was performed in a high

pressure cell having a fully equipped holography system with remotely

operated laser shutter, multiple exposure plate holder, and gas and

hydro pressure supply equipment. The optical configuration used in

making the holograms is shown in Fig. 1. With this configuration,

two holograms are made simultaneously such that nearly 360 degrees of

the vessel can be examined. The multiple exposure plate holder

allows for up to twelve double exposure holograms to be made remotely

without changing the photographic plate (4).

All measurements described here were made using a double exposure

holographic technique; each exposure was made at a constant pressure,

P, with a small pressure increment, AP, between exposures. This

procedure was repeated as the vessel was pressurized. The reconstruction of each double exposure hologram consists of an interference

pattern superimposed upon an image of the object; we call this the

fringe pattern.

All three vessels evaluated in this experiment have a design burst

pressure of 34 MPa. In order to ensure that the fringe density in

each reconstruction is not too high, the incremental change is

limited to 170 kPa. Pressure vessels A and B had been pressurized

previously to 20 MPa and 31 MPa respectively for acoustic emission

testing prior to the holographic examination. Three different test

cycles to 4.8, 22 and 31 MPa maximum pressure were run on these two

vessels with several holograms being made during each pressurization

cycle. The resulting measurements were then compared and evaluated

seeking evidence for possible damage caused by the initial acoustic

emission pressure test. The third vessel, previously unpressurized

,

began leaking at 13.8 MPa.

Results

The first series of tests on vessels A and B, which went to only a maximum pressure of 4.8 MPa, showed no significant differences

between the vessels. As shown in Fig. 2a, the fringe pattern obtained

for these vessels is nonsymmetric around the fill tube. Figure 2b

is the fringe pattern of the same vessel after rotating 180 degrees

212
about the fill tube axis. The similarity between these two patterns

shows that the displacement profile is symmetric about the two poles.

The nonsymmetric fringe pattern is a simple consequence of vessel

positioning during testing and is not due to the presence of a flaw.

Figure 3 (a,b,c,d) gives typical fringe patterns for the second

pressure test series on vessels A and B up to a maximum pressure of

22 MPa with the vessel arranged to view the waist region. The location

of the waist weld in the aluminum liner is evident in Figs. 3a and 3c.

As the pressure increases, the aluminum bladder yields and the effect

of the bladder on the surface displacement is no longer seen, as shown

in Figs. 3b and 3d. Previous examination of several aluminum bladders

showed fringe patterns similar to Figs. 3a and 3c. Thus, the stronger

waist weld region is displacing less than the remainder of the bladder

prior to the onset of yield.

During this second test series, an LVDT displacement transducer was

positioned to measure the radial expansion of the pressure vessels.

Figure 4 is a plot of the relative displacement of the fill tube and

waist regions versus pressure for vessels A and B. In these displacement plots, the change in slope occurring at 5.1 MPa is due to yielding of the aluminum liner. Vessel A registers a smaller displacement

than vessel B because of plastic deformation following the higher

prepressurization test (31 MPa versus 20 MPa for vessel B),

Figure 5 (a,b,c,d) shows the fringe pattern obtained during the second

pressure cycle to 22 MPa for vessels A and B with vessel arranged so

that the poles regions could be examined. The displacements are

symmetric about the fill tube. The higher fringe density at the

higher pressures for the same pressure increment is entirely consistent with the LVDT displacement measurements.

Converting the fringe pattern to a surface displacement vector is

possible for a single view hologram only when the part geometry and

the direction of displacement are known. Under the assumptions of

radial expansion and collimated illumination, the surface displacement of a spherical pressure vessel is given by:

Ay^= nX

A/

2 cos (arcsin X

n /R)

where n is the fringe number, X is the wavelength of light, R is the

radius of vessel and Xn is the distance from center of vessel to

fringe n. These assumptions are reasonable for our vessel and optical

configuration; in particular, illumination by a diverging beam

produces less than a 3% error in A/"". Figure 6 shows the waist and

pole-equator displacement profiles calculated for pressure vessel B.

Note particularly that the fill tube region displaces twice as much

as the waist region and that the displacement field is symmetric

213
about the fill tube.

In order to test the vessels at pressures closer to where burst is

expected, a containment box with Lexan windows was placed over the

vessel during the next pressure series. All vessels were repressurized in an attempt to produce limited damage to the kevlar fiber.

Vessel A was pressurized to 31 MPa, held at pressure for 50 minutes

and then depressurized. The fringe patterns resulting from subsequent

pressurization to 7.0 and 27.7 MPa are shown in Fig. 7 (a,b,c,d).

Leaking of hydraulic fluid (which obliterates the fringe pattern)

from the pole opposite the fill tube is evident in Figs. 7c and 7d.

The vessel does, however, appear to be maintaining its strength as

indicated by the symmetric displacement.

Leaking occurred at 28.9 MPa during the testing of vessel B, Typical

fringe patterns shown in Fig. 8 are strong evidence that this vessel

is also maintaining its strength at these pressures. The location of

the leaks in vessels A and B was found by gamma radiography after

pressurizing the vessels with a radiographical ly opaque barium

solution. Cracks in the aluminum liner at the plug weld opposite the

fill tube was found to be the cause of leakage.

Pressure vessel C, which was also to have been tested to 31 MPa, began

leaking at 13.8 MPa and the test was stopped. We interpret the fringe

patterns obtained on a subsequent pressurization to indicate possible

matrix damage. For example, Fig. 9a shows a discontinuity in the

fringe pattern near the pole. At higher pressures, a biconvex

fringe pattern developes, as shown in Fig. 9b. The discontinuity

obvious in Fig. 9a turns out to be the starting site of the biconvex

fringe pattern shown in Fig. 9b. The long axis of the biconvex

pattern is aligned with the direction of the kevlar fiber winding

pattern. Thus it appears that the low pressure (7 MPa) discontinuity

may be due to matrix crazing. This is the subject of continuing

investigation

.

Discussion

This holographic evaluation of spherical kevlar/epoxy pressure vessels

has provided information on three significant aspects of the pressure

vessel design. First, the fringe patterns show that symmetric displacement patterns occur about the two poles. This symmetry is

related to the fiber winding pattern and at the lower pressures is

affected by the aluminum bladder (see Figs. 3 and 5). Although

there is displacement symmetry, the fringe patterns are distinctly

not symmetric. Displacement contours have been quantified as shown

in Fig. 6. All three vessels show the pole region to have greater

radial displacement than the weld region. Previous work shows a direct correlation between the higher displacement (stress) at the

poles and the site of failure during burst tests (3). This indicates

21k
that the fringe pattern around the poles might be a useful measure

of either composite vessel performance or strength.

The second aspect of design confirmed by these tests is that failure

of vessels may occur in two ways. Bursting, which was not studied in

this sequence of measurements, is the result of failure of the

kevlar/epoxy matrix under stress. In these tests, failure was due to

cracks in the aluminum liner which resulted in fluid leakage.

Although the vessels leaked during pressurization , the location of the

crack in the liner is not readily identifiable from the fringe

patterns. The visible portions of the fringe patterns indicate that

these vessels were maintaining strength (see in Figs. 7 and 8) even

though the liners were leaking.

Thirdly, we wish to point out the significance of the biconvex fringe

pattern found during the testing of vessel C. This biconvex pattern

is occurring at lower pressures where del ami nation and fiber breakage

do not normally occur. In addition, the pattern is- aligned with

the winding direction. Thus, the biconvex pattern may be associated

with matrix crazing produced by the initial pressurization (in this

case, to 13.8 MPa). The quantity and location of these patterns

might prove useful in predicting the stress resistance of vessels.

A combined test program of acoustic emission and holographic interferometry on parts having no previous stress (pressure) history should

now be made in order to have a more controlled and easily interpretable

set of measurements.

Conclusion

A comparison of fringe patterns showed only minor variation in the

displacement profiles of three spherical, kevlar/epoxy pressure

vessels. Three characteristics of the kevlar/epoxy pressure vessels

were obtained from the holographic tests. The failure of all vessels

was due to a leak in the aluminum liner and not by bursting. The

fringe patterns provide displacement profiles showing areas of high

stress and show that vessels expand with axial symmetry during

pressurization. The presence of a biconvex fringe pattern is believed

to be a result of matrix crazing. Further testing will be performed

using combined holographic interferometry and acoustic emission to

better characterize spherical kevlar/epoxy pressure vessels. The

results of further testing may find a means of nondestructively

evaluating composite pressure vessels during a proof test pressure

cycle.

Acknowledgment

This work was performed under the auspices of the U.S. Department of

Energy by the Lawrence Livermore Laboratory under Contract No.

215
W-7405-Eng-78. We thank Bill Billings, Ken Harrison, and W. W. Wilcox,

Lawrence Livermore Laboratories, for their assistance in performing

the experimental work.

References

1) Meyer, M.D. and Katayanagi , T.E. "Holographic Examination of a Composite Pressure Vessel," Journal of Testing and Evaluation, Vol. 5,

No. 1 , Jan. 1977, pp 47-52.

2) Erf, R.K."Solid Propellent Rocket Inspection", Holographic

Nondestructive Testing, edited by Robert K. Erf, Academic Press,

New York, 1974, pp 365-372.

3) Hamstad, M.A. "Variabilities Detected by Acoustic Emission from

Filament Wound Aromid Fiber/Epoxy Composite Pressure Vessels,"

UCRL-86268, Lawrence Livermore Laboratory, Livermore, California,

Nov. 30, 1977.

4) Multiple exposure plate holder designed by W. Jorgenson of Sandia

Corporation, Livermore, California.

216
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CD > 2o

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CO I ' CO

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217
218
Fig. 3 Fringe patterns showing the effect of the welded waist for Vessels A and B: (a) Vessel A; P = 3.9 MPa,

AP = 0.2 MPa, (b) Vessel A; P = 20.7 MPa, AP = 0.2 MPa, (c) Vessel B; P = 3.9 MPa, AP = 0.2 MPa

(d) Vessel B; P = 20.7 MPa, AP = 0.2 MPa.

219
0 7 14 21 28

Pressure (MPa)

Fig. 4 Relative radial displacement versus pressure for vessels A and B: (a) pole of Vessel B, (b) pole of

Vessel A, (c) waist of Vessel B, (d) waist of Vessel A.

220
(a) (b)

(c) (d)

Fig. 5 Fringe patterns showing the increase in fringe density at low and high pressures for the same pressure

increment of 0.2 MPa: (a) Vessel A; P = 3.9 MPa, (b) Vessel A; P = 20.7 MPa, (c) Vessel B; P = 3.9

MPa, (d) Vessel B; P = 20.7 MPa.

221

(a) (b)

Fig. 7 Fringe patterns from Vessel A after pressurization to 31 MPa showing evidence of leak near the pole

opposite the fill tube: (a) fill tube pole; P = 6.9 MPa, AP = 0.2 MPa, (b) fill tube pole; P = 27.6 MPa,

AP = 0.2 MPa, (c) opposite pole; P = 6.9 MPa, AP = 0.2 MPa (note leakage) (d) opposite pole; P =

27.6 MPa, AP = 0.2 MPa (note leakage).

223
22 k
(a) (b)

Fig. 9 Fringe patterns showing locations of possible matrix crazing after pressurization of Vessel C to 13.8

MPa: (a) opposite pole; P = 6.9 MPa, AP = 0.2 MPa with 4 X magnification of the central region

with discontinuity in the fringe pattern and (b) opposite pole; P = 10.3 MPa, AP = 0.2 MPa with

4 X magnification of central region showing the typical biconvex fringe pattern which starts at the

discontinuity described in (a) above.

225
The Air Force "Manufacturing Cost/Design Guide (MC/DG)"

Bryan R. Noton

Battell e Columbus Laboratories

Columbus, Ohio 43201

Captain Dan L. Shunk

Wright Patterson Air Force Base

Wright Patterson Air Force Base, Ohio 45433

Abstract

The purpose of the MC/DG is to enable designers at all levels to conduct,

firstly, urgently needed, quick, simple, cost-trade comparisons of

manufacturing processes, and, secondly, structural performance/cost

trade-offs on both components and subassemblies. The MC/DG will consist

of two vol umes--"Ai rframe Structure" and "Manufacturing Technologies".

The MC/DG is being developed for the Air Force by a team consisting of

Battelle's Columbus Laboratories (BCL) and the following aerospace

companies:

General Dynamics Corporation, Fort Worth Division

Grumman Aerospace Corporation

Lockheed-California Company

Northrop Corporation, Aircraft Group

Rockwell International Corporation, Los Angeles Division

The Boeing Commercial Airplane Company is participating in a critique

mode.

Two contracts have been awarded on the MC/DG. The first program (Contract No. F33615-75-C-5194) was completed in July, 1976, and is reported

in AFML-TR-76-227. The objectives were to:

Identify the data requirements for both conventional and emerging

manufacturing methods

Identify the basic format design criteria and the formats for each

MC/DG manufacturing technology section

Prepare a detailed model of the complete MC/DG for industry

examination

Prepare an MC/DG implementation plan to collect and develop data

in the identified formats.

The objectives of the second contract, a 15-month program, awarded in

September, 1977, are to develop certain sections of the MC/DG for demonstration. These sections are:

Phase I: "Sheet-Metal Aerospace Discrete Parts"

Phase 11(a): "First-Level Mechanically-Fastened Assemblies"

Phase 11(b): "Advanced Composite Fabrication"

Phase III: "Trade-Off Study on Fuselage Shear Panel".

226
This paper will present the results from each of the above phases and

the utilization of the MC/DG sections in design studies involving metallic

and non-metallic structures.

227
TENSILE STRENGTH AND FAILURE MODES

OF BORON-EPOXY COMPOSITE WITH A NOTCH

Charles E . S . Ueng

Georgia Institute of Technology

Atlanta, Georgia 30332

Abstract: The purpose of this paper is to study the tensile strength

and failure modes of boron-epoxy composites with an edged notch.

Through the use of finite element crack program, both heterogeneity and

orthotropy of material properties are brought into consideration in

this paper. For comparison purposes, tension tests of actual composite

samples with notches are also performed. Notch insensitivity and the

failure modes are discussed. The fracture toughness of the test

material is also determined. The results obtained here are important

for further understanding of the failure of unidirectional boron-epoxy

advanced composites.

Key words: Boron-epoxy; edge cracks; failure modes; finite element

method; tensile strength.

1. Introduction:

Due to the increased use of modern composite materials in structural

applications, there has been a growing interest in the study of

fracture of composites. Three basic approaches are generally available

(1) a statistical theory to predict the number of cut fibers in a simple composite structure and to investigate what effects these broken

fibers have on structural strength; (2) extension of linear-elastic

fracture mechanics theory to orthotropic or anisotropic materials and

then to fibrous composites; and (3) study of cracks in a heterogeneous

material or at an interface and the extension of that theory to the

case of a fibrous composite. Rosen and Dow [l] and Zweben [2] have

researched a statistical approach to investigate the fracture of composites. It was noted in [2] that a drastic reduction in composite

strength would take place when only a few fibers are cut in a theoretical and hypothetical model. In some early research on the fracture

of orthotropic materials, Sih, Paris and Irwin [3] noted that "for

problems involving self-equilibrating loads, the stress intensity factors are the same for both isotropic and anisotropic materials". But

this has been shown [4,5,6] not to be always the case. The investigation of a crack at or near an interface between dissimilar materials

has been reported in [4,7,8]. Erdogen [7] found that for cracks at or

near a bi-material interface, mostly aluminum-epoxy , the power of the

singularity at the crack tip ranged between zero and one, depending on

228
the crack direction and loading condition. In the case of an isotropic

material, there is always a square root singularity regardless of loading condition as noted in one of the earliest investigation on cracks

by Williams [9]. Olster and Jones [8] reported that for cracks perpendicular to the fibers in a composite, the toughness increased in proportion to the debonding, implying that a weak interface is desirable

for increased toughness. They also found that in a multidirectionally

reinforced composite, the toughness is primarily controlled by those

filaments most nearly perpendicular to the crack. Cooper and Kelly [lO]

experimentally investigated the conditions under which a notch-insensitive composite could be produced and attempted to predict the values of

the work of fracture for a notch-sensitive material. They noted that an

important function of the matrix lies in increasing the work of fracture

by plastic flow, or other means, rather than by acting as a physical

barrier to crack growth. They discovered that notch-insensitivity could

be accomplished if splitting parallel to the fibers at the notch is complete. Hoover and Allred [ll] observed that changes in crack length do

not significantly alter the delamination process and the consequent

notch insens itivity in single edge notch specimens of Borsic-Al composites with weak fiber-matrix bonds. Mandell, Wang and McGarry [12] investigated the size and character of the damage zone at the tip of sharp

notches in fiber reinforced laminates. The damage zone was found to increase in extent approximately in porpotion to up to fracture for

notch-sensitive laminates. But for notch- insens it ive laminates this

zone reaches a critical stage then it rapidly spreads completely across

the specimen.

The motivation of this research was to incorporate both heterogeneity

and orthotrophy into a study of an edge crack or an edge notch in a unidirectional boron-epoxy composite to see what effect heterogeneity

would have in the neighborhood of a crack. Tension tests were performed on notched specimens in order to determine the mode of failure

and the failure load. A finite element model of this composite was

made which took into account the effects of layering and heterogeneity

in the neighborhood of the crack and orthotropy away from the crack.

Certain preliminary results of this research project were previously

reported in [13].

2 . Experimental Work:

Test samples of Avco Rigidite 5505/4 boron-epoxy composite were provided by the Lockheed Georgia Company. Figure 1 shows all the dimensions. The percent of boron by volume in the specimens was 50.2. The

boron had a tensile modulus of 58(10) 6 psi (400 GN/m ) and a Poisson's

ratio of 0.208. The epoxy had a tensile modulus of 0.45(10)6 psi

(3.10 GN/m2

) and a Poisson's ratio of approximately 0.351. The composite had a tensile modulus of 30.9(10) 6 psi (213 GN/m ) and a Poisson's

ratio of 0.256 as given by the furnished data from Lockheed. A 3/8"

diameter hole (9.5mm ) was drilled through each end of the specimens in

order to load them in axial tension with dowel pins. But it turned out

229
to be unsatisfactory because the first piece tested failed by debonding around the upper hole (Fig. 2). Consequently, tension grips

were used for the remaining tests. 1/4" deep edge notches (6.4mm) were

cut by a carbon- phenolic abrasive wheel on a milling machine. Originally, it was planned to cycle all specimens in order to start the

crack growing from the notches. One piece was cycled between 100 lb.

(45.4 kg) and 1,200 lb. (544.3kg) at a rate approximately 36 cycles per

minute. Instead of starting a crack, the cycling served to debond the

specimen vertically at the notch tip (Fig. 3). After about 30,000 cycles, it showed a clear line of debonding for about 1 5/16" (33 mm)

long on both sides of the notch. This phenomenon indicated that the

specimen was notch- ins ens it ive , so the cycling plan was discontinued.

The specimens, including the cycled piece, were failed in simple tension using a universal testing mechine with tension grips. It is

interesting to note that the cycled piece failed in the same manner as

the uncycled pieces. The first failure noted on all the notched specimens was a complete vertical debonding through the thickness at the

notch tip (Fig. 4) at an average load of 3,200 lb. (1,470 kg). The

average load at complete failure was 11,900 lb. (5,410 kg). For comparative purposes, three unnotched specimens were also failed in the

same manner. Internal failures were seen at loads as low as 3,320 lb.

(1,510 kg). Complete failure was in the same manner as for the notched

specimens (Fig. 5) at an average load of 13,500 lb. (6,140 kg). Table

1 presents, for each specimen, the load at which notch debonding occurred

(for the notched specimens), the smallest load at which internal failure

(load-dropping point) was noted, and the load at total failure.

Table 1 Test Results

Piece Notch Debonding; Earliest Internal Failure Total Failure

1 4000 lb. (1814 kg) 6400 lb. (2903 kg) 12,580 lb . (5706 kg)

2 4000 (1814) 7000 (3175) 10,540 (4781)

3 3000 (1361) 6600 (2994) 12,320 (5588)

4 2400 (1089) 5400 (2449) 12,060 (5470)

5 2600 (1179) 7000 (3175) 11,940 (5416)

6* 9320 (4228) 13,800 (6260)

7* 4600 (2089) 13,120 (5951)

8* 3320 (1506) 13,620 (6178)

3. Finite Element Analysis

A finite element computer program capable of characterizing the cracktip singularities was used to investigate the failure of a unidirectional boron-epoxy composite with an edge crack. Due to symmetry of the

loading condition, only half of the specimen (2" x 4%" , or 55TM&114mm

)

230
needed to be modeled. This area was divided into constant strain triangles and material constants were assigned to each triangle. Away from

the tip, the material constants of the composite were assigned to the

triangle elements of the full thickness, 0.040 (1mm). In the vicinity

of the crack tip, the boron and epoxy were modeled as separate entities

and the layering was taken into account. This was accomplished by stacking eight triangles of 0.005" (0.13mm) thickness and 0.004" (0.1mm) width

one on top of another in the crack tip region and alternating layers between boron and epoxy in sort of a three dimensional checkerboard fashion. The fibers, in other words, were approximated by rectangular prisms

rather than circular prisms. This modeling was feasible because of the

relatively large fibers (0.004" or 0.1mm in diameter). The purpose in

modeling the composite in this fashion was to take into account heterogeneity in the crack tip region, the region in which it is felt most.

The finite element grid is shown in Fig. 6. In the shaded area, the boron and epoxy were modeled separately and layering was taken into account. The crack element (8 elements actually) is centered at the bottom of the grid in the detailed diagram. A unit displacement was imposed along the top of the specimen and the resulting stresses along the

top and the stress intensity factor were determined by the program. The

relationship between the average mode-one stress intensity factor (K^)

and the applied stress (a) was

for a 1/4 inch (6.4mm) crack. The

crack elements was

h

and in the epoxy elements it was

= 0.959 a

stress intensity factor in the boron

= 1.904 a

= 0.014 a

It should be noted that in practice, the strength of the composite will

be governed by the strength of the boron and hence the stress intensity

factor in the boron crack elements should be used as the governing

stress intensity factor.

The stress results around the crack tip were also documented. Table 2

shows such data, where the origin is placed at the left free edge.

Table 2 Averaged Stress Around the Crack Tip

x 0.2453" 0.2473 0.2486 0.2500 0.2513 0.2526 0.2546

(6.23mm) (6.28) (6.31) (6.35) (6.38) (6.42) (6.47)

a -290 psi -220 +5850 +3830 +6640 +4850 +3700

at y=0 (-2.00MN/ 2) (-1.52) (+40.3) (+26.4) (+45.8) (+33.4) (+25.5)

231
It is interesting to note that the maximum value of O"

, 6,640 psi (45.8

MN/ 2), due to an average load when the notch debonding was observed, is

about 2/3 of the tensile strength of the epoxy, 10,000 psi (69.0 MN/ 2).

Also, the zig-zag stress distribution would not have been obtained i?

the heterogeneity were not included.

4. Discussions and Conclusions:

Based on the results reported here, the following observations can be

made

:

(1) The failure of the boron-epoxy composite is not by transverse

fracture at the notch tip as in isotropic materials;

(2) The failure of a dowel pin as observed here may mean that the

stress concentration is higher there than at the notched tip.

From the failure prevention point of view, a concentrated load

in a hole would be of more concern than an edge notch;

(3) The notch insensitivity as noted during the cycling operation,

indicates that fracture criteria normally associated with isotropic materials does not adequately predict the failure of

the composite studied here. The randomness of the failure as

represented by Fig. 5 supports this assumption. It appears

that the notch insensitivity and failure by debonding are due

to the high fiber volume fraction and apparently low bond

strength between the fibers and matrix;

(4) The maximum stress o~ = 6,640 psi (45.8 MN/_2) occurred immediately to the right of the notch tip, and is about 2/3 of the

tensile strength of the epoxy. This deviation may be caused

by differences in the properties of the epoxy in bulk form

compared to a small amount located between two boron fibers.

From the fairly high value of o~Xj it is expected that longitudinal cracking is incipient in this investigation;

and

(5) Based upon the information on the average failure loads between the notched and unnotched specimens, 11,880 lb. versus

13,513 lb. (5,404 kg vs. 6,142 kg), it is interesting to note

that the former value can be obtained by a proportion if the

length of the crack is deducted from the width of the specimen.

In other words, it can be concluded that the presence of the

notch introduces no stress singularity in the remaining unnotched material.

Furthermore, a fracture toughness, or critical stress intensity factor,

of the specimens tested may be calculated. Using the stress intensity

factor in the boron cracked elements, K = 1.904a, and the average failure load of the notched specimens, 11,888 lb. (5,404 kg), the fracture

toughness obtained is

K = 283,000 psi TinT

C

232
for a 2" (51 mm) by 0.040" (1 mm) cross-section. Using the failure load

of the unnotched specimens tested, 13,513 lb. (6,132 kg), an ultimate

tensile stress of 168,900 psi is obtained for the material based upon a 2" (51 mm) by 0.040" (1mm) cross-section. Following the example of

Mandell et al [12], the ratio of fracture t oughness to ultimate tensile

stress for the material tested here is 1 . 6&7in • For the same ma ter ial

in a 90/0/0/90 ply configuration, they obtained a value of 0 .32Vin

.

(Table 3, p. 283 of [12]). They also found that the subcrack length is

generally greater for laminates with larger ratios of fracture toughness

to ultimate tensile stress. The work performed in this investigation

supports that finding, at least for the case of boron-epoxy composite

since the subcrack length at fracture obtained here was a full 9" (229

mm), the length of the specimen, compared to a value of 0.2" (5mm) obtained in [12].

References

[l] Rosen, B. W. and Dow, N. F., "Mechanics of Failure of Fibrous

Composites," in Fracture (H. Liebowitz, ed . )

, vol. 7, Academic

Press, 1972, pp. 611-674.

[2] Zweben, C, "On the Strength of Notched Composites," J. Mech.

Phys. Solids, vol. 19, 1971, pp. 103-116.

[3] Sih, G. C, Paris, P. C, and Irwin, G. R., "On Cracks in Rectilinearly Anisotropic Bodies," J. Frac . Mech., vol. 1, no. 3, 1965,

pp. 189-202.

[4] Corten, H. T., "Failure Mechanics of Composites," in Fracture

(H. Liebowitz, ed . ) , vol. 7, Academic Press, 1972, pp. 675-769.

[5] Walsh, P. F., "Linear Fracture Mechanics in Orthotropic Materials,"

Eng. Frac. Mech., vol. 4, 1972, pp. 533-541.

[6] Bowie, 0. L. and Freese, C. E., "Central Crack in Plane Orthotropic Rectangular Sheet," Int. J. Frac. Mech., vol. 8, no. 1,

March 1972, pp. 49-57.

[7] Erdogan, F., "Fracture Problems in Composite Materials," Eng. Frac.

Mech., vol. 4, 1972, pp. 811-840.

[8] Olster, E. F. and Jones, R. C, "Effect of Interface on Fracture,"

in Composite Materials (L. J. Broutman and R. H. Krock, eds.), vol.

1, Academic Press, 1974, pp. 245-284.

[9] Williams, M. L., "On the Stress Distribution at the Base of a Stationary Crack," J. Appl. Mech., March 1957, pp. 109-114.

233
[10] Cooper, G. A. and Kelly, A., "Tensile Properties of Fibre-Reinforced Metals: Fracture Mechanics," J. Mech . Phys . Solids, vol.

15, 1967, pp. 279-297.

[ll] Hoover, W. R., and R. E. Allred, "The Effect of Crack Length and

Bond Strength on the Delamination Process in Borsin-Al Composites"

J. of Composite Materials, January 1974, pp. 55-64.

[12] Mandell, J. F., Wang, S. S., and McGarry, F. J., "The Extension of

Crack Tip Damage Zones in Fiber Reinforced Plastic Laminates 1 ', J.

of Composite Materials, July, 1975, pp. 266-287.

[13] Ueng, C. E. S., Aberson, J. A., and Lafitte, B. A., "Tensile

Analysis of an Edge Notch in a Unidirectional Composites", J. of

Composite Materials, April 1977, pp. 222-234.

25I+
Fig. 1 Specimen Dimensions

235

Fig. 4 Complete Vertical Deboading

Fig. 5 Typical Failure

257
INSERTION ONE

Fig. 6 Finite Element Grid

238
SESSION &

Clia,ii*3ncia,xi: GS- . Kardos

Carleton University

Failure Criteria for Composites Under Complex Loading

P. W. Mast, L. A. Beaubien, D. R. Mulville, S. A. Sutton

R. W. Thomas, J. Tirosh, and I. Wolock

Naval Research Laboratory

Washington, D.C. 20375

*

Abstract

Studies are in progress to develop failure criteria for composites andbonded joints under a broad range of in-plane loads. Criteria for

predicting the conditions under which a defect will grow in composites

structural components, in such applications as Naval high performance

craft, will be demonstrated. Validity of a new technique for determining failure criteria for composites has been verified in laboratory

tests.
FAILURE ANALYSIS OF AN IDEALIZED COMPOSITE DAMAGE ZONE

Carl T. Herakovich and David A. O'Brien

Department of Engineering Science & Mechanics

Virginia Polytechnic Institute & State University

Blacksburg, Virginia 24061

Abstract: First failure of finite width laminates with an idealized

composite damage zone is analyzed through the use of a linear elastic

finite element stress analysis and the tensor polynomial failure

criterion. Crossply and angleply [±45] laminates are analyzed under

tension and compression loading. It is shown that a boundary layer

exists near the damage zone which is very similar to the boundary layer

which has been shown previously to exist along free edges. Comparisons

are made between predicted initial failure of damaged and undamaged

laminates. Transverse tension is the dominate mode of failure for

tensile loading, but compressive loading leads to a variety of failure

modes depending upon the laminate configuration.

Key words: Laminated composites, failure, tensor polynomial, finite

elements, stress concentrations, interlaminar stresses, tension,

compression, graph ite-epoxy.

INTRODUCTION

Failure of advanced composite materials has been the subject of many

recent papers. Numerous failure theories have been proposed [1], and

theoretical and experimental investigations have been conducted on

notched and unnotched specimens. However, very little, if any,

attention has been given to surface damage in which one or more layers

have "peeled off" due to impact or other causes. This paper investigates the stress distribution around an idealized surface damage zone

in composite laminates, and then uses the tensor polynominal failure

criterion [2] to predict the initiation of failure. The idealized

damage zone consists of a centered rectangular groove or cutout

extending the entire length of a long symmetric composite laminate

(Fig. 1). For symmetry reasons, both top and bottom surfaces of the

laminate are considered to be damaged. This idealized "damage zone"

was chosen as a tractable problem which would provide insight into the

more general problem of the stress distribution around an irregularly

shaped damage region.

Supported by The NASA-VPI&SU Composites Program, NASA Grant NGR

47-004-129.

2k2
The stress distributions are determined using a quasi-three-dimensional,

linear elastic finite element stress analysis. The problem formulation

includes the interlaminar stresses which are present in the vicinity of

the idealized damage zone as well as those in the boundary layer along

the laminate free edge [3-6]. Results are presented for [0/90] s

, [90/0] s and [±45] s graphi te-epoxy laminates under both tensile and

compressive loading. The different character of initial failure in

the three laminates is discussed.

PROBLEM FORMULATION

The Finite Element Problem

The laminates consist of layers of an orthotropic material with

the principal axis of each layer alined at an angle e with the x axis

(Fig. 1). Only balanced, symmetric laminates are considered. The

constitutive equation for each layer can be written, in condensed

notation, in the form

{a} k = [C]

k {e} k

(1)

where: - . [C] is the transformed 6x6 stiffness matrix

{o} is a 6x1 stress vector

k {e} is a 6x1 strain vector

and k refers to the k

th layer of the laminate

2U3
Since the laminate is long and the loads are applied at the ends,

the states of stress and strain are independent of the axial coordinate

x. The most general displacement field for this problem can be written

in the form

u = £x + U(y,z)

v = V(y,z) (2)

w = W(y,z)

where 5 is the applied uniform axial strain. Thus, the problem reduces

to the determination of the unknown displacement functions U(y,z),

V(y,z) and W(y,z) over a cross section of the laminate. Symmetry

allows the investigation to be limited to one-quarter of the cross

section.

The finite element method of solution was chosen because of its

versatility with irregular shaped geometries and material inhomogenities

(including laminated materials). The region of interest is subdivided

into a finite number of triangular elements as shown in Fig. 2. As

indicated in the figure, a higher density of smaller elements was used

in regions of suspected stress concentrations. The dimensions of the

smallest element are shown in the figure.

The displacement fields (U,V,W) over each element were represented by

linear interpolation functions expressed in terms of the unknown nodal

displacements. The nodal values of the displacements were then the

primary unknown of the problem. They were determined using minimum

potential energy to establish the elemental force-displacement relationships in the form

[K]

(A) {u} (£)

= {F} (£)

(3)

where

[K]^ is the elemental stiffness matrix

lo)

{u} K

' are the elemental nodal displacements

{F}( £

) are the elemental nodal forces

The elemental relationships were then assembled into a set of simultaneous equations relating all unknown nodal displacements and known nodal

forces. The simultaneous equations were solved for displacements after

modification for boundary conditions. The strains and stresses in

individual elements were then determined from the strain-displacement

relations and constitutive equations. The linear interpolation function

gives rise to constant stress and strain elements.

2kk
The problem formulation given here permits the determination of all six

components of stress whenever present. It has been coded in FORTRAN

for an IBM computing system and is documented more completely in

references 7 and 8. The program uses the efficient in-core equation

solver COLSOL [9].

Tensor Polynomial Failure Criterion

The availability of all six components of stress provides a unique

opportunity to study the application of a three dimensional failure

criterion. The tensor polynomial criterion was chosen as being representative of composite failure. The criterion postulates the

existence of a failure surface of the form

F.a. Fim - 1 (4)

where F-j and F

n-j are strength tensors whose components can be expressed

in terms of the material principal strengths [2]. When (4) is expanded

in principal material coordinates for an orthotropic material, it has

the form

F

1 a1 + F

2 a2 + F

3 a3

+ F^c* + F^o2 + F^aj

(5)

+ F

44

T

23 + F55

T

13 + F66

T 12 + 2F

12

a

l a2 + 2F

13

a

l a

3 + 2F

23

a

2 a

3

= 1

Alternatively, the polynomial may be expressed in laminate coordinates

It then has the form

F K + F 2°y

+ F

3°z

+ F6^xy

+ F il°x

+ F22°y+ F

33°z

+ F44V + FKi4

+ F

66

T

xy

+ 2F16Vxy + 2F26°yT

xy

(6 »

+ 2F 36Vxy + 2F

45

T

yz

T xz

+ 2F

i2

0 x°y

+ 2F 13Vz + 2F23Vz

= 1

where primed quantities are the transformed strength values.

The polynomial is evaluated for each constant stress element to

determine which element will be the first to fail as the load is increased. Examination of the individual terms of the polynomial (5)

2U5
at failure provides insight into the mode of failure. Thus the form

(5) is useful for failure mode determination. The form (6) is helpful

when considering the stress free boundary conditions since individual

terms may be set to zero. An interesting observation is the form of

the failure criterion (6) for the intersection of a free edge with a free surface where boundary conditions require that, for any laminate,

all components of stress with the exception of a are zero. Equation

(6) then reduces to:

F

i°x

+ F il°x = 1

< 7 >

Material Properties and Geometry

All results presented in this paper have been obtained using material

properties which are representative of Thornel 300/Narmco 5208

graphite-epoxy . The elastic properties used were:

= 1.05 x 10

6

psi

= 0.38 (8)

x 10

6

psi

The principal strengths were taken to be:

X

t

= 218 ksi Y

t

= Z. = 5.9 ksi F

]2

= -0. 58xl0" 10 /(psi

)

2

X = -218 ksi Y = Z = -36.3 ksi S, 0 = ST , = S 0

~ = 9.8 ksi

C C C 1 2 1 3 23

E

t

b

ll = 27.6 x 10

6

psi G

12

= G

13

= G

23

Ec = 18.9 x 10

6

psi v12

= v 13

= v 23

E

t

L

22

L

33

1.6 x 10

6

psi Ec

L = 22

EcL = 331.8

(9)

The resulting components of the strength tensors are:

F

] = 0 Fn

= 0.00002 F

44

= 0.0104

F

2 = 0.142 F

22

= 0.00477 F

55

= 0.0104 (10)

F

3

= 0.142 F

33

= 0.00477 F

66

= 0.0104

_2

where the F. and F.. have units of (psi)

J Prediction of initial failure is of course dependent upon the elastic

constants and principal strength values employed. As indicated above,

some assumptions were made regarding the elastic properties in the 2-3

plane and the strength values. Thus, these predictions may require

modification as more complete experimental results become available.

As indicated in Figure 1, the half-width of the damaged zone was taken

to be b/3 where b is the laminate half-width. The depth of the damage

zone was always equal to one ply thickness. The aspect ratio (b/H)

2k6
was 34.1 in all cases. Thus, the laminates under consideration are

very wide. As shown in reference [10], the boundary layer width

decreases with increasing aspect ratio.

RESULTS AND DISCUSSION

All three laminates were investigated under tensile and compressive

loading, and both damaged and undamaged laminates were considered. The

edge of the damage zone was sufficiently removed from the free edge to

eliminate any possible interaction. It was verified that the classical

laminate solution was recovered in the interior region between the free

edge and the damage region.

The results of the failure analysis are summarized in Table 1. The

table shows the contribution of the significant terms of the tensor

polynomial, the strain at first failure and the location of first

failure. In those cases where first failure occurred in the vicinity of

the damage zone, the results for undamaged laminates are also presented.

As indicated in the table, failure of crossply laminates is predicted to

occur at the free edge for every case except compression of the [90/0]

laminate. First failure for this case is predicted at the corner of

the damage zone. Unlike the crossply laminates, angleply [±45] s laminates exhibit first failure at the damage zone for both tensile and

compressive loading. The predicted failure strains are only a relative

measure as they do not include residual thermal stresses or nonlinear

material behavior.

All stress distributions presented in this paper were obtained using

the two different stress averaging schemes depicted in Fig. 3. The

interlaminar stresses az , x zxand x Zy must be continuous across any

line z = constant and hence the interlaminar stress at point H in the

figure was taken to be the average of elements 14, 15, 16 and 17 which

are on both sides of the interface. The inplane stress components

ax , ay and x Xy are not required to be continuous across an interface

and, therefore, they were averaged along lines such as BB and DD in

the figure. Thus, at point F, the stresses are the average of elements

8 and 9 while for point G elements 10 and 11 are averaged. The failure

analysis was based upon the actual stress values in each element.

Crossply Laminates

Tensile loading of these laminates always resulted in first failure in

the 90° ply at the intersection of the free edge and the 0/90 interface.

Failure is dominated by the transverse normal stress 02 with the linear

term $2°2 being the dominate term. The contribution from the interlaminar normal stress (03) is small for both laminates, but it is much

greater in the [0/90] s laminate which exhibits an 18 percent lower

failure strain. The variable influence of the interlaminar normal

stress is clearly shown by the stress distributions in figures 4 and 5.

The [0/90] laminate gives rise to a tensile a at the free edge whereas

2^7
<u

e

so

13 •— i— +J

V •i— to ~

ro

O

Li_

O

<L>

C CO LO CO lo LO LO CTi 00

r .__ 1 1 CTi

O LO LO CO

_3 ro U)ns LO CM 1 0 LO . C ^c>

ITM- cy\

V NJ uu LO

—* C J LO 1 •1TM 1 * s, * • * • 1—

("O 00

O 0 0 0 O

O 0 0 1 1 1 1 ENs:

UJ Nl _l ^ NJ I J CM CTi UJ CM CO LO ' 1 0 LO 1 O LO 1—J * * • UJ 1 0 O O O 1 t—1

<C

u_ ^ NJ 1 0 CM CO

1—* LXJ 1 CM 1—1

LO 1 O 1 O

LO • • • CO Li_ O O O O to o1—I

+^ I— -1—* ZD k NJ V NJ *o Oi 00 CO

CQ

c LO C\J • r— i—i 1 1 1 O 0 E 1 0 CO

C£

t

—

£= • • 03 • 1— u_ fO O 0 O I 0 0

—J

O— "O * ) 0 LO r--- CM C\J C1J CTi CO

/1 1

Li j C\J LO L7> LO CTi _J /1——\> CM 1 V '

rJ-< TO rv^ * J * LJ

f0 • m

E • t—1

u_

E 0 0 O

O (O

O

O s: 03 T3 o Q C

2: ZD

>-_i CM CO 0 UD LO 00 CO 00 CO 0 to CM LO O «* LO O 0

Q_ C\J

O CM

O 0

CM • • • CCO

U_

O

O

O 0 COO 0 CO

GO

UJ

O ' 00 CO LTt 00

1— C\J C\J LO CTi LO 0 1 00 1 1 LO

C\J * * * * • • 1— 1 t O O 0 CO1 0 CO

UJ

_J

00

<C CM 1— LO CO CO

hD LO 0 0 LO 1 1 LO 1 • • • U_ O O

"O

fC01— 0 1—

O l— 0 O 1— 0

_l

0)

+-> (/) to to to

ra 1—1 1—1

1-1 1 1 LO to 1—1

LO to

c 0 0 0 O 1 1 1—1 0 1—1 1—1

•1— cr> CT> LO LO LO LO E 0 O ^t" 0 <3- 0 O CTi cr> +1 +1 cn +1 +1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1—1

1—1

c0

sz

to

+-> 0c

toE

S-

<D

+->

EO

+J 1— E to

(OE to

OO <U

'1• 1

S-

<+- CL r~ E E E a> O

CT>|— <_) •r-to 11 II E1—1

h- <_>

248
it is negative there for the [90/0] s laminate. Close examination of

these figures shows that the stress distributions are not merely a sign

reversal and hence the laminates exhibit fundamentally different behavior. This result was first reported by Wang and Crossman [6].

(Results obtained in [11] with a much finer mesh show that a z is

actually tensile at the free edge for both laminates, but is much

higher for the [0/90] s laminate. These refined results did not alter

the location of first failure.) These figures also show that stress

concentrations exist in a boundary layer near the edge of the damage

zone. The stress distributions in the damage zone boundary layer are

similar in character to those in the free edge boundary layer, but

the stress concentrations are of different strength.

Compression of these laminates indicates variable location and mode of

failure. The [0/90] laminate fails at the top outside corner of the

laminate with failure being completely dominated by o] as predicted by

eqn. (7). The [90/0] s laminate fails in a combined mode with a-j , 02

and 02 all contributing significantly; however the major contribution

is again from an. The compression failure strain for the damage zone

failure of the [90/0] s laminate is 30 percent below the compressive

failure strain of the [0/90] s laminates and 8 percent lower than first

failure of an undamaged [90/0] s laminate. The primary reason for the

different failure modes and locations is the value of o

z near the

damage zone. It is tensile for [90/0] s and compressive for [0/90] s

.

2 1+9
(The stress distributions for compressive loading are reversed in sign

from the tensile results in figures 4 and 5.) The different tensile

and compressive failure stresses are due mainly to the different

strengths in tension and compression (Eqn. 9). The undamaged [90/0] s laminate fails at the intersection of the midplane and the free edge

in a combined a], 03 failure mode when loaded in compression (Table 1).

The interrelationships between stresses can be explained by considering

equilibrium of the free body diagrams shown in Fig. 6. The interlaminar

normal stress xyZ is related to cy by force equilibrium in the y

direction. Since xyZ is always zero along a center cut C-C (laminate

solution), the interlaminar normal stress a z must be equivalent to a couple when oy is nonzero and must be equivalent to nonzero force and

nonzero moment when oy = 0 (as in the case of the [±45] s laminate). The

interlaminar shear and normal stresses xyz and az are of course always

related through the point-wise equilibrium equation.

[±45]

s Angleply Laminate

As indicated in Table 1, first failure of damaged [±45] s laminates

occurs at the inside corner of the damage zone for both tensile and

compressive loading. Tensile loading gives rise to a transverse 02

failure mode whereas compressive loading results in a mixed mode

failure between 02 and all three shear components. The undamaged

CO

Q.

CO _ CO O

UJ

0AHA0E ZONE H1 1 lOorSOIf hJISOorO I J H

'• b •' Y

T

_[0/90],

yz [so/o], 0

0.0 o.s Y/B

V

0, if

^cutting pfcne

1 ^

6Z

B

\-y/b

= 0 667

Fig. 5 - at z=h

Q

for [0/90]

s Fig. 6 - Partial Free Body Diagram

and [90/0]

s U=0.1%)

250
laminate fails on the midplane at the free edge when loaded in tension

and on the ±45 interface at the free edge when loaded in compression.

The failure modes of the undamaged laminates are essentially the same

as the damaged laminates. The failure strains of damaged laminates

are approximately 10 percent lower than those of undamaged laminates.

Compression loading of both damaged and undamaged [±45] s laminates shows

an interesting result. The 02 terms of the tensor polynomial are both

quite large, but of opposite sign at failure. However, the sum

F2a2+Fp2^2 1S actually negative in both cases. Unlike the crossply

laminates, shear plays an important role in first failure of the angleply laminate.

Stress distributions along the interface of [±45] s laminates are presented in Figures 7-9. Again * we see that there are stress

concentrations in a boundary layer near the free edge and the damage

zone. Since ay is zero in interior regions of the [±45] s angleply

laminate, equilibrium considerations (Fig. 6) require that the integrals

of xyZ , az and yaz along the interface be zero in each boundary layer.

Figure 7 shows that the stress distributions for these two stresses

exhibit the proper number of oscillations about zero and, at least

approximately, satisfy the equilibrium conditions. Invoking the stress

boundary conditions and using the stress transformation equations gives

Fig.3 7 - a and x at z=h for

z yz 0

[±45]

s (g=0.1%)

Fig. 8 - x zxat z=h

Q

for [±45]

g

(€=±0.1%)

251
the following equations for the stresses at the free edge, in material

principal coordinates, in terms of the laminate stresses.

CT 1 = °2

= V2 CT 3 = a

z

T

13

=

" T

23

= T zx

//2 T

12

=

" a

x /2

(11)

These equations show that a x , a

z and x zx control failure along free

edges - including the free edges of the damage zone. As shown in Fig.

8, x zx exhibits a very high stress concentration at the corner of the

damage zone and at the free edge. (The signs of shear stresses are not

important since the shear stresses appeared only as "squared" terms in

the tensor polynomial.) The distributions of ax above and below the

interface z = H are shown in Fig. 9. Stress concentrations are again

present near the damage zone and the free edge. The highest concentration is near the damage zone where first failure occurs. Table 1,

equations (11) and figures 8 and 9 indicate the first failure of [±45]$

laminates is primarily due to 02 (i.e. ax ) for tensile loading, and is

a mixed mode failure dominated by x zx ana a

x when loaded in compression.

The finite element formulation used in this investigation does not

satisfy the stress-free boundary conditions exactly and hence equations

(11) are only approximately satisfied. However, additional results for

undamaged laminates with a much finer mesh have indicated that the

location of failure does not shift [11].

o.

I J

z

1 45 Ml -45 H

b H

Y 1

o-»-

z - h«*

» Z. h." 0

0.0

' 1• ' 0.5 1 . Y/B

Fig. 9 - a at z=h for [±45]

U=o.u)

252
CONCLUSIONS

Stress concentrations and interlaminar stresses are present in the

vicinity of damage zones in laminated composites. The stress distributions near the damage zone are very similar to those which have been

shown previously to exist along free edges. The tensor polynomial

failure criterion predicts first failure to be transverse tension (c^)

for tensile loading of all three laminates considered. Compressive

loading leads to different failure modes depending upon the laminate

configuration. Compressive failure of [0/90] s symmetric laminates is

completely dominated by a-| normal stress, but the [90/0] s laminates

fail in a combined mode involving all three normal stress components.

The [±45] s laminate fails in a combined mode when loaded in compression. Transverse 02 tension and all three shear components make

significant contributions to the tensor polynomial for this case.

With the exception of compression of the [90/0] s laminate, crossply

laminates fail at the free edge before failing near the damage zone.

The damage zone is more critical for [±45] s with first failure

occuring there for both tensile and compressive loading.

Future work in this area should include residual thermal stresses,

moisture effects and nonlinear material behavior.

ACKNOWLEDGEMENT

The assistance of Mr. A. Nagarkar with some of the computational work

is greatly appreciated as is the typing of Ms. Frances Carter.

REFERENCES

1. Inelastic Behavior of Composite Materials , (Herakovich, C. T.

,

editor), ASME AMD Vol. 13, Am. Soc. Mechanical Engineers, 1975.

2. Wu, E. M. , "Phenomenological Anisotropic Failure Criteria," in

Composite Material

s

, Vol . 2 , Mechanics of Composite Materials , (Broutman, L. J., Krock, R. H. and Sendeckyj , G. P. , editors )

,

Academic Press, 1974.

3. Pipes, R. B. , and Pagano, N. J., "Interlaminar Stresses in

Composite Laminates Under Uniform Axial Extension," J_. Composite

Materials , Vol. 4, 1970, 538-548.

4. Herakovich, C. T. , Renieri, G. D. , and Brinson, H. F., "Finite

Element Analysis of Mechanical and Thermal Edge Effects in Composite

Laminates," Army Symposium on Solid Mechanics, 1976, Composite

Materials : The Influence of Mechanics of Failure on Design , Cape

Cod, MA. , Sept. 1976.

253
5. Wang, A. S. D. and Crossman, F. W. , "Some New Results on Edge

Effects in Symmetric Composite Laminates," J_. Composite Materials , Vol. 11, 1977, 92-106.

6. Farley, G. L. and Herakovich, C. T. , "Influence of Two-Dimensional

Hygrothermal Gradients on Interlaminar Stresses Near Free Edges,"

Environmental Effects on Advanced Composite Material

s

, ASTM STP 658,

Am. Soc. for Testing and Materials, 1978.

7. Renieri, G. D. and Herakovich 5 C. T. , "Nonlinear Analysis of

Laminated Fibrous Composites," VPI-E-76-10, also NASA CR-148317,

June, 1976, 177 pages.

8. O'Brien, D. A. and Herakovich, C. T. , "Finite Element Stress

Analysis of Idealized Composite Damage Zones," VPI-E-78-6, also

NASA CR-155923, February, 1978, 141 pages.

9. Bathe, K. , Wilson, E. L. , Numerical Methods in Finite Element

Analysis , Prentice-Hall, Inc., Englewood Cliffs, N.J., 1976.

10. Hsu, P. W. and Herakovich, C. T. , "Edge Effects in Angle-ply

Composite Laminates," J_. Composite Materials , Vol. 11 (Oct.

1977).

IT. Herakovich, C. T., Nagarkar, A., and O'Brien, D. A., "Failure

Analysis of Composite Laminates with Free Edges," submitted for

publ i cat ion.

251+
INTERLAMINAR FAILURE IN EPOXY BASED COMPOSITE LAMINATES

A. S. Wang

Drexel University

Philadelphia, Pa. 19104

F. W. Crossman

Lockheed Palo Alto Research Laboratory

Palo Alto, Ga. 9^304

G. E. Law, Jr.

Drexel University

Philadelphia, Pa. 19104

Abstract: One of the common failure modes in fibrous composite

laminates involves interlaminar cracks which cause delamination. Several Questions arise in the course of a delamination analysis. Notably

these include (a) how is a delamination initiated? (b) under what condition will a delamination grow? and (c) what determines the stability

of a delamination growth? Owing to the unusual complexity of composite

laminate systems, there appears no simple way to answer these questions

sati s fac tori1 y

.

In this paper, a method is developed based on the energy principle of

the classical linear fracture mechanics, in order to describe the

various aspects of the delamination mechanisms. Specifically, free-edge

induced delamination failure in some epoxy based composite laminates

are investigated employing this method. And, the results obtained are

compared with experiments. It will be shown that the energy method

predicts well free-edge delamination in laminates under uniaxial tension, including the initiation, the growth and the growth stability

of the delamination.

Key words: Delamination mechanisms; energy release rate; finite element method; free-edge stresses; graphite epoxy composites; interlaminar cracks; stability of cracks; virtual crack closure.

One of the common failure modes in fibrous composite laminates may involve the nucleation and propagation of interlaminar cracks. This will

usually cause ply delamination. In some cases, delamination may lead to

immediate failure of the laminate; and in other cases it may induce

long-term property degradation resulting in a reduced structural reliability and durability. Delamination caused by free-edge stresses has

frequently been encountered in laminates that have straight edges[l,3]

or curved edges such as around cut-outs and holes[4,5]. It is known

255
that interlaminar stresses are highly localized near the laminate's free

edge [6, 7], and their sign and magnitude depend on the ply stacking sequence. Hence, some laminates may be more prone to del ami nation than

others, depending on how the plies are stacked together.

Since free-edge delamination may cause immediate laminate failure, the

ultimate strength of the laminate is therefore also sensitive to the

stacking sequence of the plies. For example, reference [2] reported test

results for a family of six different stacking sequences of T300/934

quasi -isotropic laminates that were tested under uniaxial tension. It

was found that the tensile strength of the six differently stacked

laminates ranged from 63 ksi to 87 ksi . When correlated with the freeedge interlaminar normal stress cj z [7], the laminate of the lowest

strength (±45/0/90) s , was found to have the largest tensile dz ; the

laminate of the highest strength (90/±45/o)s , was found to have the

largest compressive dz . Hence, in the case of the tensile dz , an opening mode delamination could be induced prematuredly resulting in a lower

laminate strength. This qualitative correlation however, could not be

used as a failure criterion for the strength prediction.

In addition to the stacking sequence, free-edge delamination is found to

depend on the thickness or the volume of the ply which is stressed inter

-

laminarly. The characteristics of this effect is that the occurrence of

ply delamination may be delayed by a decrease of the ply thickness, or

it may be accelerated by an increase of the ply thickness. For example,

reference [8 ] reported test results for a series of T300/5208 graphite

epoxy quasi -isotropic laminates tested under uniaxial tension; the experimental data are summarized in Table I below:

TABLE I

Lami nate Axial Stress At On-set of Delamination

( ±45/0/90) s 45.3 ksi

45.3 45.4 (ave.)

^5.5

(±452/02/902 )s 38.3

38.6 38.2

37.8

(±453/03/903)3 28.7 29.7

30.7

It is seen that the applied axial stress at the on-set of delamination

varies with the thickness of the material ply, even though the stacking

sequence in all cases is the same. Furthermore, if a free-edge stress

analysis [7] is performed for all the three laminates, one would find
an identical stress field for all three( under the same applied axial

stress) . Clearly, the observed ply volume effect cannot be explained

adequately by a criterion based on stress alone. In particular, the conditions governing the on-set of delamination and the manner of its

growth are not known.

In a recent paper by Rybicki, et al[9], the free-edge delamination problem was described by a stable crack growth procedure based on the

energy release rate concept in classical fracture mechanics [10] . The

quantity measuring the material resistance against ply delamination is

the critical strain energy release rate Gc . It is reasoned that the

stressed plies dissipate surface energy as new crack surface is created

under the applied load; the rate of the available energy dissipation

per newly created surface, G, is viewed as a driving force which can

extend the crack further. Since the value of G depends on the level of

the applied load, the crack can be extended when G exceeds the material

resistance force Gc , Otherwise, the crack would remain stable until the

applied load is increased, or the crack-tip conditi ons are a"! tered. The

crack-tip conditions can be altered at least by the size of the crack;

hence the magnitude of G is a function of the crack length. A knowledge

of the latter may determine the stability of the crack growth.

The energy release rate G is a quantity determined analytically, and it

is in fact a complicated function of crack location, crack size, ply

stacking sequence, ply properties, ply thickness and the applied load,

To calculate G, a finite element procedure is suggested in conjunction

with a crack-closure technique[9 ] . Specifically, G is computed from the

work done in closing a known crack-tip extension. Because of the linear

elasticity assumption, the work done in closing a crack is equivalent to

the energy released when the same crack is opened[ll]

.

In this paper, the energy release rate approach is further developed

with the aid of a special finite element routine which models the freeedge delamination process in graphite epoxy laminates. Several laminates

of practical interest are investigated by this method; and the results

are compared with experiments. It will be shown that the energy method

not only predicts the on-set of delamination, it also describes its

growth behavior. The observed ply volume effect on delamination is

explained explicitly by the analytical model.

Consider the laminate with a symmetric lay-up as shown in Fig. 1. Under

the applied axial stress 5X , interlaminar normal stress (j z and shearing stress Txz and ryz will be induced in the laminate free-edge region.

These stresses are proportional to the applied laminate stress d , according to linear elasticity. Owing to these stresses, a free-edge

delamination may be created as shown. It is assumed that some amount of

surface energy, denoted by AE, is released creating the crack of size

Aa. For elastic fracture, AE may be determined by the difference in the

total stored strain energies of the laminate before and after the small

257
crack of size Aa. Thus, the energy release rate associated with the

crack is given by:

As the del ami nation crack grows incrementally in a self-similar manner,

G is then calculated numerically as a function of the crack size a.

As is illustrated in Fig. 2a, the computational procedure is as follows:

the cross-section of the laminate is represented by a net-work of finite

elements with densely populated grids deployed near the intended interface crack region. To calculate G for a crack with crack-tip node k, see

Fig. 2b, the nodal displacement components uk

, and wk

are determined

first in terms of the applied stress dx . Next, by introducing a virtual

crack extension Aa while maintaining the same applied stress, node k is

then zipped open to two separate nodes k' and k" as shown. A finite

element solution then yields the displacements for nodes k' and k". And

finally, nodes k' and k" are forced to close so that the displacements

of k' and k" equal those of k. In this step, the finite element solution

gives the nodal force components Xk , Yk and Zk which are needed to close

the two open nodes. Thus, the energy release rate G for the crack of size

a is given by the work done in closing the virtual crack extension Aa:

G(a) = [xk(uk

« - uk

") + Yk

(vk, - vk

») + ZR

(wk

, - w

k„) ] /2Aa (2)

The actual finite element net-work used in the computation reported in

this paper differs from laminate to laminate. In general, the incremental crack extension Aa is set at 0.2 times the thickness of a single

ply. For example, a typical graphite epoxy tape is approximately 0.005

inch thick; hence the incremental crack extension Aa is only 0,001 inch.

The crack-tip elements typically contain about 10 (graphite) fiber

cross-sections which is probably the smallest element size for which the

effective continuum representation of the material ply is still applicable, see reference [12 ].

Since the finite element nodal forces and nodal displacements are solved

in terms of the applied laminate stress d

x

, the work done during the

virtual crack closure process is then expressible in terms of dx%/ Ex

. Accordingly, the energy release rate G is expressed as

G(a) = g(a/t) t #/ E

x

(3)

In the above expression, 5X and are laminate stress and stiffness,

respectively; t is a conveniently chosen characteristic length, such as

the thickness of a single ply; g is a non-dimensional function of (a/t),

a being the delamination crack size, as shown in Fig. 1. Note that g is

derived directly following Eq..(2).

Fig. 3 illustrates the general behavior of g as a function of (a/t). It

is seen that g increases sharply with (a/t) to reach a maximum, gj^g^i

258
at a = a* ; it then decreases slightly to a minimum, g_. , at a = a**.

After this point, g increases again slowly and monotomcally

.

The physical interpretation of this behavior is that, under the applied

stress d , the laminate is capable of releasing delami nation energy

depending on the ply interactions; there exists a unique maximum energy

release rate g^g^ , given the location of the delamination. It is then

theorized that gmax is associated with the on-set of delamination, an

action that pops open a small crack of size a*. Once this happens, the

available energy release is somewhat reduced and a slightly higher load

will be required to extend the crack further. This is a stable crack

growth; but the process is rather short. The growth becomes unstable

once the delamination crack reaches a size of a**.

Suppose the criterion G —

G

c for fracture is used. Then, it follows

from Eq., (3) that the critical stress at the on-set of delamination is

given by:

d ). = (G E /g t)* (4) x'i v c x ' °max '

v '

and the critical stress at unstable growth is given by:

d )_ = (G E /g . t)2 (5) x

y fv c x ' tomin ' K '

Numerical calculation of g indicates generally that the value of a* is

approximately equal to the thickness of the material layer that is

stressed the most by the interlaminar stresses. For instance, in the

case of (±30/902)s , a* is roughly the thickness of the 90-layer, or

2 ply thick; in the case of (±30/90^)s , a* is about 3 ply thickness. On

the other hand, the value of a** is usually 2 to 3 times the thickness

of the laminate.lt is also interesting to note that gmin ls about 90%

of gjjjg^* making tfx )j_ about 5% smaller than ^

x)f.

For experimental comparison, two cases will be discussed in the following:

The first case concerns a T300/934 (±^5/o/90)s laminate which is tensile

tested under step-loading. Mi orographic examination is then performed to

discern the damages that occurred at each load level. Detailed results

of this study have been reported in reference £2} . It is found that edge

delamination in the 90-layer first appeared at about 60 ksi ; excessive

delamination extending inside the laminate occurred at about 63 ksi.

The laminate final strength was about 67 ksi.

To correlate these results with the energy release rate method, the

finite element calculation for the function g(a/t) is performed for the

laminate using the following ply properties [2 ]

:

Ej^ = 23.7 msi E

T

= 1.72 msi

259
G

LT

= 0.94 msi GTT =0.5 msi

V

LT =0.3 vTT

= 0.5

The individual ply thickness is 0.0055 in. and the overall laminate

stiffness in the load direction is Ex = 9 nisi.

Fig. 4 displays the calculated g function for del ami nation along the 90/

90 interface. Here, gmax

= 0*365 an(i it occurs at about a* equal to the

ply thickness t. On the other hand, gmj_ n

= 0.33 which occurs at about

a** equal to 15t(not shown in the figure).

Now, assume the on-set of delamination occurred at 3X

= 60 ksi. It then

follows from Eq..(4) that the critical energy release rate Gc for ply

delamination is 0.8 in-lb/in2

, or 140 j/m2. This value is compared to

Gc = 150 j/m.2 found independently for similar material system [l 3] , by a

test method suggested in reference [14 ]

.

Using gmj_ n

= 0.33 and Gc = 0.8 in-lb/in^, the critical stress for unstable delamination 3X) ^ is found to be 63 ksi.

For the second case of experimental correlation, consider the test

results reported in reference 8 , where the ply-volume effect on edge

delamination was observed experimentally. The three laminates with different ply volumes are (±*+5/0/90)s, (l^A^/^) s and ( ^5j/0^/9°-}) s 5 they are all made of T300/5208 graphite epoxy unidirectional tapes. The

mechanical properties of the tape are given in the following[8 ]:

E^ = 19.5 msi E

T

= 1,48 msi

G

LT= GTT

= 0.8 msi \>LT= vTT

= 0.3

The individual ply thickness is 0.0057 in.

It is noted that the three laminates are identical in every aspect except the thickness of the material layer. In fact, under the same axial

stress 3 , all three should develop identical laminate stress and strain

fields. And, they all have the same axial stiffness E^ = 7.7 msi. Now,

if we choose the thickness of the 90-layer to be the characteristic t

in calculating the function g, then a common g can be generated. Fig. 5

displays this common function g for edge cracks along the 90/90 interface in all the three laminates. Here, g, = 0.445 and it happens at

about a* equal to t. Note that t equals to 0.0057 in., 0.0114 in. and

0.0171 in. for the three respective laminates.

Again, assume Gc= 0.8 in-lb/in2 . From Eq,.(4) the critical stresses at

the on-set of edge delamination for the three laminates are given by

49.3 ksi, 34.9 ksi and 28.5 ksi, respectively. These compare closely

with the experimental values of 45.4, 38.2 and 29.7 ksi as shown in

260
Table I.

It should be pointed out that Eq..(4) or (5) expresses explicitly the dependence of the critical stress on the characteristic length t. t, if

properly chosen, represents the ply-volume effect.

In conclusion, the energy release rate approach developed in this paper

appears to be useful for delamination analysis of composites. It not

only predicts the on-set of delamination, it also describes its growth

behavior. For all practical purposes, it gives a good estimate of the

laminate's ultimate strength. Moreover, the observed ply-volume effect

on delamination can be explained explicitly by the analytical model.

Reference:

1. K.L. Reifsnider, E.G. Henneke and W.W. Stinchcomb, 'Delamination of

Quasi -isotropic Graphite Epoxy Laminates, 8 Composite Materials: Testing and Design. ASTM STP 617(1977) , p. 93.

2. J.G. Bjeletich, F.W. Crossman and W.J. Warren, 'The Influence of

Stacking Sequence On Failure Modes In Quasi -isotropic Graphite Epoxy

Laminates,' Failure Modes In Composites-IV, AIME(l978) .

3. G.P. Sendeckyj and H.D. Stalnaker, "Effect of Time at Load on Fatigue

Response of (0/±45/90)2s T300/5208 Graphite Epoxy Laminate,' Composite Materials: Testing and Design. ASTM STP 6l ?(197?) , p. 39.

k. J.M. Whitney and R.Y. Kim, 'Effect of Stacking Sequence on the Notched

Strength of Laminated Composites,' Composite Materials: Testing and

Design. ASTM STP 617(1977), p. 229.

5. M.S. Rosenfeld and S.L. Huang, 'Fatigue Characteristics of Graphite

Epoxy Laminates Under Compressive Loading,' Jour. Aircraft, Vol. 15

(1978), p. 246.

6. R.B. Pipes and N.J. Pagano , 'Interlaminar Stresses in Composite

Laminates Under Uniform Axial Extension, ' Jour. Composite Materials,

Vol. 4 (1970), p. 538.

7. A.S.D. Wang and F.W. Crossman, 'Some New Results on Edge Effects in

Symmetric Composite Laminates,' Jour. Composite Materials, Vol. 11

(1977), P. 92.

8. B.T. Rodini, Jr. and J.R. Eisenmann, ' An analytical and Experimental

Investigation of Edge Delamination in Composite Laminates, ' 4th

Conf. on Fibrous Composites, San Diego, Cal., Nov. 14-17 (1978),

9. E.F. RybickL, D.W. Schmueser and J. Fox, 'An Energy Release Rate Ap26l
proach for Stable Crack Growth in the Free Edge Del ami nation Problem

Jour. Composite Materials, Vol. 11 (1977) , p. ^70.

10. A.S. Tetelman and A.J. McEvily, Jr. , 'Fracture of Structural Materials, ' John Wiley and Sons, New York (1967).

11. G.R. Irwin, 'Handbuch der Physik; ' Vol. 6 (1958), Springer-Verlag.

p. 551.

12. E.F. Rybicki and N.J. Pagano/A Study of the Influence of Micro

Structure on the Modified Effective Modulus Approach for Composite

Laminates,' Proc. Interntl. Conf. Composite Materials, AIMMPE (1976)

13. M.G. Bader and P.T. Curtis, 'The Micromechanics of Fibre CompositesCarbon Programme,' 1st Report, July (1977) 1 and 2nd. Report, Dec.

(1977), Dept. Met. and Matls. Tech., Univ. Surrey.

Ik. J.E. Srawley and W.F. Brown, 'Fracture Toughness Testing and Its

Applications,' ASTM STP 38I (19&0 .

Fig. 1 Isometric View of Free Edge Del ami nation

262
(a)

-*| a+Aa

k

u

(b)

Fig. 2 (a) Finite Element Grid-work in the Crack-tip Region.

(b) The Finite Element Virtual Crack Closure Procedure.

g

on-set of delamination

region of stable growth

on-set of unstable growth

i/t

a*

Fig. 3 Energy Release Rate Function g vs. Delamination Size a/t

263
.4

.3

.2

.1

0.

0 .2 .4 .6

T300/93^ (±^5/o/90)

s

E =9.0 msi

x

t = 0.0055 in.

8 1. 1.2 1.4 a/t

.4

Fig. 4 Energy Release Rate Function g for Crack Along 90/90

Interface. T300/93^.

.3

.2

T300/5208 (j45i /01 /901 )s

E = 7.74 msi for i = 1,2,3. .A. t = 0.00571 in. , i = 1,2,3.

.1

0.

0. .2 .4 .6 .8 1.0 1.2 1.4 1.6

i/t

Fig. 5 Energy Release Rate Function g for Crack Along 90/90

Interface. T300/5208.

264
SESSION lO

Cliairman: 3M„ Silvergleit

David "VST. Taylor Naval Slirp

Research & Development Center

CURRENT DEVELOPMENTS IN ADVANCED COMPOSITES

FOR NAVAL SHIP APPLICATION

J.J. Kelly

David W. Taylor Naval Ship R&D Center

Annapolis, MD. 21402

H.H. Vanderveldt

Naval Sea Systems Command

Washington, D.C. 20362

Abstract: The reasons for and the limitations on the use

of composites for ship applications are described. The

current exploratory development program in advanced composites for Navy ship applications emphasizes high performance craft such as hydrofoils. The future program will

complete these efforts and stress broader applications for

composites. Development of fire resistant composites, survivable structural concepts and materials which reduce

maintenance requirements will be emphasized.

Introduction

The purpose of this paper is to provide an overview of the

Navy's R&D Program in composite technology for ship application. Prior to describing the current program and its

future directions, the reasons for both the Navy's interest

and its hesitation in all-out adaptation of these materials

should be clearly understood. The major reasons for the

Navy's interest are:

• These materials offer the potential of reducing

structural weight with little or no sacrifice in

strength

.

t The materials can be "designed" to meet the

requirements of a specific application.

• Past experience in applications such as small

boats, submarine sonar domes, superstructure,

ladders and piping by the U.S. Navy has generally

been good.

• Use of composite materials by both foreign navies

and commerci al /pri vate ships has been extensive

and generally exhibited good, reliable performance.

267
• The capacity of these materials to "fail" gradually

rather than suddenly without warning is a highly

desirable characteristic.

t The noncorrosive nature of most of these materials

can lead to a significant reduction in shipboard

maintenance requirements.

e For some applications such as minesweepers, the

non-magnetic features of most of these materials

are important.

In spite of these desirable features, the Navy is cautious

in the use of these materials in ship applications. The

following factors have limited and delayed the use of composi tes

:

§ The effect of long-term marine environment

exposure on mechanical and physical properties

of composite materials and configurations is not

well known.

e The cost of the materials and subsequent fabrication are relatively high.

9 Composites require the addition of fire retardants

or protective coatings to survive and not contribute in a fire environment. Most of these

materials will emit large quantities of smoke and

toxic products in a fire, even if not contributing

to supporting the fire. This is a serious problem,

especially in the confined spaces of a ship at sea.

• Lack of detailed knowledge on all loads and

conditions which can effect a specific application.

• At the present time, the shipbuilding industry,

Navy shipyards and fleet support activities are

not geared to handle these materials.

0 In contrast to the Air Force, Army and the Navy's

Air Fleet, there are only a limited number of

ships available for testing new materials/concepts

in a real environment.

t The funding levels for R&D in composites do not

permit investigation for a large number of applicat i ons

.

268
Program Management and Selection

The R&D programs currently directed by the Naval Sea Systems

Command are all aimed at either eliminating or reducing the

impact of these limitations so that the fleet's performance

can be increased and/or the cost of ownership can be

reduced. These programs are carried out by in-house design

personnel, Navy laboratories and industrial contractors.

For clarity, the R&D programs can be viewed as being conducted in two distinct phases. In one case, the risk is

held to a minimum by using accepted design practices and

materials with well established properties and behavior.

In general, these efforts are in response to a specific

fleet need, such as design and evaluation of composite

antenna masts. In contrast, the exploratory development

phase not only addresses advanced materials such as graphite

and boron fiber reinforced epoxies, thermoplastics, metal

matrix composites but concepts such as isogrid structures.

The primary purpose of this latter R&D phase is to identify

areas where composites can have a significant impact, and

develop both the materials and/or data necessary to support

design for a marine environment application. In general,

with this information in hand, a test component is fabricated to develop cost data and for testing both in the

laboratory and on a ship if possible. The results of these

efforts are incorporated in a specification guidance document and report to support follow-on advanced development,

manufacturing technology or direct use by the fleet.

As mentioned earlier, the available resources for these

programs and especially the exploratory development efforts

are limited. In order to obtain maximum use of these

resources, developments successfully completed or currently

underway in the Air Force, Army, NASA and Navy Air Systems

are carefully reviewed for use in this program.

Current Programs

The current Exploratory Development is shown in Table I with the lead performing activity identified. For the most

part, the program shown here had its genesis back in 1973/

74 when there was a fairly high emphasis on advanced ships

such as hydrofoils. Since many of these ships and craft

were weight critical, the use of composites for structural

and equipment applications was a high priority item.

Studies were conducted to determine where composites could

be employed and how much benefit could be derived. 1 ' 2

Since most of these tasks are discussed in detail in other

269
papers today, only a brief synopsis of each will be provided here.

The composite flap is a control surface for a hydrofoil ship. The flap was designed and fabricated by Boeing Marine Systems and installed on

the Navy hydrofoil "High Point" in late 1978.

The flap is actually a hybrid consisting of a titanium structure to which graphi te/epoxy skins

are attached. These skins are overlaid with a thin titanium sheet for erosion protection.

Evaluation of the flap should be completed in early

FY1980.

• The hybrid Box Beam is a test element which represents a ship structure. The hybrid Box Beam consists of a HY-130 steel frame with graphi te/epoxy

skins attached. It was designed and fabricated by

the McDonnell Douglas Astronautics Company. This

element has been received by the Navy and examined

nondestructi vely by a pulse echo ultrasonic test

method. The element has failed in the area of

load application twice and an in-depth analysis of

the failure and the necessary corrective action is

currently underway.

• The all composite Box Beam is again a test element

and represents an extension of the technology for

the Navy's data base. McDonnell Douglas Astronautics Company is currently doing the preliminary

design and testing for this element.

• The Silicon Carbide/Aluminum metal matrix composite has been developed by the Naval Surface

Weapons Center, White Oak Laboratory. The composite can be readily made with 2000, 6000 and 7000

series aluminum alloys. Current investigations are

aimed at extending this capability to the 5000

series marine grade aluminum alloys. Investigations of joining methods, corrosion behavior and

mechanical properties are currently underway.

§ Failure Analysis and Nondestructive Evaluation

methods are currently being developed and evaluated

by the Naval Research Laboratory. In particular,

the analysis effort is investigating failure behavior of composites and bonded joints under both

static and fatigue spectrums representative of the

loading induced on box beam and hydrofoil structures. The use of ultrasonic test techniques with

the ability to provide permanent records is the

270
primary thrust of the NDE effort.

• Structural Evaluation is the Navy's planned effort

to evaluate box beams under full spectrum loading;

analyze the results of the shipboard tests of the

instrumented foil flap and develop the necessary

analytical methods for future evaluations and

materials developments.

0 Marine properties are determined for a number of

composite materials and configurations that are

being used or contemplated for use in box beams,

flaps, etc. In addition, studies of other promising composite materials such as fiber-reinforced

thermoplastics, metal matrix composites and joint

designs are also investigated.

• GRP piping systems for shipboard service investigation involved an extensive amount of testing of

actual pipe, joints, and complete assemblies for

fatigue, fire resistance, fouling behavior, erosion

resistance, and shock resistance. This program

will be completed this fiscal year and has already

produced a draft specification for GRP piping

systems that is currently being reviewed by industry.

Future Directions

As can be seen on Table II, for the near term, the emphasis

in the exploratory development program will be to complete

the current underway efforts directed to advanced ship

applications. As shown in Table III, the process of

identifying specific shipboard areas of potential composite

application will be continued and items which can contribute to reduced maintenance requirements will be stressed.

Development of fire retardant composite materials with low

smoke and toxic product evolution will be initiated. For

the longer term, the use of composite materials for applications such as propellers or impellers will be investigated. This effort will draw on a new independent in-house

Navy laboratory effort in this area. Isogrid or similar

concepts using composite materials will be investigated for

application to ship and especially advanced craft structures such as Air Cushion Vehicle propeller shrouds where

they can offer improved maintenance requirements. The combination of ceramic fibers and foam as well as organic

composites for structural ballistics and fire protection

271
will also be examined. The SiC/Al composites and other

metal matrix composites will be investigated for specific

ship application. These applications will be determined

when a proper trade-off between existing materials, organic

and metal matrix composites has been conducted.

References

1. Greszezuk, L.B. and Hawley, A.W.; Application of

Advanced Composites to Patrol Craft Hydrofoils; Final

Report, Contract No. N00024-72-C-5536 , McDonnell Douglas

Astronautics Company, April 1973.

2. Greszezuk, L.B. and Hawley, A.V.; Application of

Advanced Composites to Hydrofoil Strut; Final Report,

Contract No. N00024-72-C-5536 , McDonnell Douglas Astronautics Company, December 1973.

272
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275
ENVIRONMENTAL EFFECTS ON FIBER REINFORCED

ORGANIC MATRIX COMPOSITES

M. S i 1 v e r g 1 e i t

A.B. Macander

F.A. Brauer

H.P. Edelstein

David W. Taylor Naval Ship Research and Development Center

Annapolis, Maryland 21402

Abstract: A review is presented on the effects of the

marine environment on fiber reinforced organic matrix composites. In addition to the effects of extended immersion

in seawater on composite properties, the limiting utility

of composites due to cavitation erosion, flammability and

smoke spread evolution and damage due to impact are discussed. In general, high quality, low void content glass

reinforced plastics (GRP) and graphite epoxy composites

retain over 90% of their initial strength after three

years immersion in seawater. Organic matrix composites,

which are subjected to high speed fluid flow are degraded

by erosion due to the collapse of the cavitation cloud,

however, erosion resistance can be improved by using

elastomeric or thin metallic overlays. With the exception

of polyimides, organic matrix composites have poor resistance to flammability and emit large quantities of dense

smoke. Impact resistance of composites is a function of

fiber/resin properties and composite orientation. Incorporation of high toughness flexible matrices and/or high

strength, low modulus fibers will increase composite

impact resistance.

Key words: Graphite epoxy; glass reinforced plastics;

advanced composites; extended immersion; marine environment; cavitation erosion; flammability; impact.

Emphasis in the Navy is being directed to reduce ship subsystem weights while maintaining or increasing lifetime

reliability. Reductions in weight can be achieved

through the development and use of low weight, high

strength/stiffness fiber reinforced organic matrix composites. Glass reinforced plastics (GRP) which are

currently being used and graphite epoxy composites which

are contemplated for use in Navy ship structural applications are the two typical composites discussed. The

276
literature abounds with information on fabrication procedures and physical and mechanical properties of GRP and

graphite epoxy composites. However, for Naval structural

use information is required on the response of these

materials to the marine environment. It may be expected

that composites used in surface ship applications may be

subjected to prolonged exposure to seawater. Under these

conditions the effect of water on strength retention and

response of composites under fatigue loading conditions

will be of extreme importance. In sub-surface applications, similar information is required except that the

effect of pressure on composites would be the dominant

factor. Application of composites in high speed areas,

such as propellers, foils, etc... .may be subjected, to

degradation by cavitation erosion. Conventional organic

matrices support combustion and evolve large quantities

of smoke while burning. Information on the response of

composites to flammability is essential since recent

shipboard fires have placed severe restrictions on

materials which support combustion and evolve smoke.

Damage may be induced in composites by impact during

docking procedures or maneuvers. This type of damage can

result in cracking of the matrix which may permit ingress

of water into the structure, thereby reducing its efficiency, or in the main, cause significant loss in strength

or stiffness. These major aspects of the Navy's program

on the effects of the environment and flammability and

damage on fiber reinforced organic matrix materials are

delineated in this paper.

EFFECT OF EXTENDED IMMERSION

The effects of extended exposure to seawater on GRP laminates which are used in surface ship applications are

shown in Figure 1. The response is quite typical of GRP

materials, in that if degradation occurs it will occur in

a relatively short time after which the rate slows down

with no further loss in strength.' There have also been

periodic surveys conducted by the Navy to evaluate the

laminate properties and to assess the durability of Navy

small boat GRP hulls. A study 2 conducted for the Navy

showed no indication of degradation in GRP hulls due to

aging or extended immersion in the marine environment.

GRP materials have also been used in submarine applications. Submarine fairwaters which are used to minimize

277
turbulent flow have performed satisfactorily for over 11

years. Fried and GraneH report that after 11 years service there was no evidence of deterioration and the material met the original specification requirements, Figure 2.

Studies^ were conducted to assess the effects of extended

seawater immersion at various pressures up to 13,000 psig

(30,000 foot depth) on glass reinforced filament wound

epoxy materials. It was observed that these materials were

unaffected by hydrostatic pressure and retained over 90% of

their initial strength provided that the laminates were of

good quality and had void contents of less than M. Other

compositional factors, such as type of resin and hardener

system, could markedly influence response of the material

to extended immersion. It was also demonstrated that the

fatigue strength in the "wet" state was somewhat lower than

the "dry" condition. A concept^ was developed wherein the

fatigue life of fiber reinforced plastics was expressed as

a percentage of the ultimate static stress.

The weight critical nature of the Navy's high performance ships, i.e., hydrofoils, air cushioned vehicles,

and surface effect ships, impose design limitations on

ship size and operational efficiencies. Graphite epoxy

composites which are lightweight and have high strength

and stiffness characteristics meet the design requirements of high performance ships and are actively being

considered for structural applications. As in the case

of GRP, it was observed^ that the properties of a high

quality graphite epoxy composite (void content<l%) were

not adversely affected by extended immersion in seawater,

Figure 3. It was also noted that there was no relation

between amount of water absorbed and strength retention

and that the maximum rate of water absorption occurred

during the initial phase of the immersion period.

CAVITATION EROSION

The phenomenon of cavitation erosion is associated

with high velocity systems, such as propellers, pumps

and foils. It results in local damage and removal of

material due to the cavitating environment. The mechanism of cavitation erosion has been proposed as occurring when vapor bubbles form in a rapidly moving liquid

and grow until they reach a higher pressure and then collapse. Energy created by the collapsing bubbles damage

the material with resultant erosion. Studies of this

phenomenon have shown 7 that high yield steels, titanium

278
and stainless steel are most resistant to cavitation erosion, while fiber reinforced organic matrix composites

are affected to a greater degree, Figure 4. Measurements

were made using the vibratory cavitation erosion apparatus,

ASTM 6-32. The initial effect of the cavitation phenomenon on fiber reinforced organic matrix composites is

to erode and degrade the matrix. Continued implosions

crack the matrix and the exposed fiber becomes unstable.

In the extreme case the fiber is eroded and can be fractured. Elastomeric type coatings are available which are

resistant to cavitation erosion but difficulties have been

noted in adhering large sections to the coatings.

FLAMMAB I L I TY

Under exposure to a flame, polyester and epoxy matrices will support combustion and evolve large quantities of smoke. Recent shipboard fires have resulted in

placing severe restrictions on the utilization of materials which support combustion and evolve smoke. These

restrictions may severely limit the use of organic matrix

materials, particularly in ships habitability spaces.

The effect of resin, fiber and fire retardant additives

on the flammability characteristics of organic matrix

composites were evaluated. 8 Flame spread index was deter

mined by the radiant panel test, ASTM E - 1 6 2 , the amount

of smoke and products of combustion were determined using

the National Bureau of Standards Smoke Density Apparatus,

while the amount of oxygen required to support combustion

was determined using the Oxygen Index Method, ASTM D2863.

Polyimide/gl ass laminates, which have the highest thermal

resistance of fiber reinforced organic matrix composites,

had the greatest resistance to burning as exemplified by

a flame spread index of 1 . A comparison of the flammability characteristics of various composites is shown

be! ow

:

TABLE 1 Flame Spread of Fiber Reinforced

Organic Matrix Composites

Composite Flame Spread

(Matrix/Rei nforcement

)

I ndex

Polyimi de/Gl ass

Polyester/Glass 31

279
TABLE 1 (cont)

Composite Flame Spread

(Matrix/Reinforcement) Index

Epoxy/Glass 38

Epoxy/Graphi te 32

Addition of fire retardants to the polyester and

epoxy matrices resulted in significant reductions in

flame spread index. For example, the flame spread index

of the polyester glass composite, Table 1, was reduced

from 31 to 5, after the addition of antimony trioxide

and hydrated alumina fire retardants. However, this also

resulted in a substantial increase in the amount of smoke

evolved. The amount of smoke generated, as determined by

the National Bureau of Standards Smoke Density Apparatus

is shown below:

TABLE 2

Composite Maximum Specific

(Matrix/Reinforcement) Optical Density

Polyimide/Gl ass 29

Polyester/Glass 469

Polyester/Glass* 634

Epoxy/Glass 364

Epoxy/Glass* 565

Epoxy/Graphi te 313

Epoxy/Graphi te* 678

*Resin system contains fire retardants

The high thermal capabilities of the polyimide system is

evidenced by its having the lowest smoke generation (maxi

mum specific optical density). The addition of fire

retardants showed significant increases in smoke generation for both the polyester and epoxy systems. Ceramic

type barriers when added to the polyester or epoxy systems were found to effectively reduce the flame spread

index and increased the time required to develop maximum smoke accumulation.

IMPACT

Fiber reinforced composite materials used in Naval

structural applications may be damaged by impact induced

280
in docking procedures, by falling tools or equipment

during ship overhaul procedures and by striking submerged

logs or other debris during high speed maneuvers. Studies are being conducted at the Center to determine the

influence of fiber, matrix, matrix flexibility and fiber

orientation on the impact resistance of composites. GRP

materials, incorporating an elastomeric modified epoxy

resin, exhibited significantly higher impact resistance

than conventional polyester or epoxy resin systems. Similar studies on graphite epoxy composites have indicated

that normal impact strength is a function of fiber tensile strength and modulus, fiber orientation and resin

ductility.

A summary of the data for graphite epoxy composites

is shown in Table 3. For example, a composite fabricated with high tensile strength, intermediate modulus

graphite fiber subject to normal impact had greater impact resistance than a lower strength ultra-high modulus

graphite fiber. Use of a thermoplastic resin system,

which has higher ductility than the conventional rigid

epoxy resin, in a composite reinforced with graphite

fibers also produced improved impact resistance characteristics.

TABLE 3

Impact1 Strength of 1/4" Thick Composites

Mater i al

Impact Energy

(ft-lbs)

Ultra High Modulus

Graphite Fiber/Epoxy z Matrix (0°/90°

Fiber Orientation)

4

1 6

High Tensile Strength, Medium

Modulus Graphite Fiber/Epoxy 2

Matrix (0°90° Fiber Orientation)

18

High Tensile Strength, Medium Modulus

Graphite Fi ber/ Fl exi bi 1 i zed Epoxy

Matrix (0°/90° Fiber Orientation)

22

281
TABLE 3 (cont)

Impact Energy

Material (ft-lbs)

High Tensile Strength, Medium

Modulus Graphite Fiber/Thermoplastic

Matrix 3 (0°/90° Fiber Orientation) 32

T] Impact Normal to Surface

2) Rigid Epoxy Matrix

3) Composite Thickness - 1/8"

Observation by ultrasonic nondestructive evaluation of

damage induced by impact and correlation with mechanical

properties showed that damage is confined to a local area.

Additionally, impact resistance of graphite epoxy composites can be improved by incorporation of fibers of high

strain characteristics, such as, glass or Kevlar. Another

possible alternative to improve impact resistance may be

the use of an elastomeric or metallic overlay.

SUMMARY

o GRP and graphite epoxy composites, void content

less than 1%, have not been degraded by extended

immersion in the marine environment.

o Fiber reinforced organic matrix composites are

eroded and degraded by cavitation erosion. Effects

of cavitation erosion can be reduced by utilization

of an elastomeric coating or thin metallic overlay.

o Polyimide resin matrix composites exhibit the greatest resistance to flammability and evolution of

smoke. In general, polyester and epoxy matrices

have poor resistance to flammability and generate

large quantities of dense smoke.

o Ceramic type barriers are effective in reducing

flame spread and smoke evolution.

o Fiber/resin properties and composite orientation

influence impact resistance of composite materials.

282
References

:

1. Silvergleit, M., "Historical Review - Navy Programs,"

Center for Composite Material, University of Delaware,

Newark, Delaware, 30-31 March 1976,

2. Graner, W.R. and R.V. Delia Rocca, "Evaluation of U.S.

Navy GRP Boats for Material Durability," Society of the

Plastics Industry, Reinforced Plastics/Composites Division,

26th Annual Technical Conference, February 1971.

3. Fried, N., and W.R. Graner, "Durability of Reinforced

Plastic Structural Material in Marine Service," Marine

Technology, Vol. 3, Nov 3, 1966.

4. Silvergleit, M., J. Kaminetsky and N. Fried, "Filament

Wound Glass Reinforced Plastics for Deep Sea Vehicles - the

Present-State-of-the-Art , " Society of Aerospace Materials

and Process Engineers, 1967.

5. Freund, J.F. and M. Silvergleit, "Fatigue Characteristics of Glass Filament Reinforced Plastic Material,"

Society of the Plastics Industry, Reinforced Plastics

Division, 21st Annual Technical Conference, February 1966.

6. Macander, A. and M. Silvergleit, "The Effect of the

Marine Environment on Stressed and Unstressed Graphite/

Epoxy Composites," Naval Engineers Journal, August 1977.

7. Silvergleit, M., R. Deppa and H.P. Edelstein, "Potential

Application of Advanced Composites for High Performance

Craft," National SAMPE Technical Conference Series, Vol. 5,

October 1973.

8. Silvergleit, M., J. S. Morris, and C. N. La Rosa,

"Fl ammabi 1 i ty Characteristics of Fiber Reinforced Organic

Matrix Composites," Polymer Engineering and Science, Vol

18, No. 2. , February 1 978.

283
RETENTION — PERCENT

IMMERSION TIME — YEARS

Figure 1 - Effect of Extended Immersion in the Ocean on the

Compressive Strength of Glass Reinforced Polyester

Laminates

.

GLASS REINFORCED PLASTIC FAIRWATER

EFFECT OF 11 YEARS SERVICE ON USS HALF BEAK

ORIGINAL CURRENT DATA SPECIFICATION'" PROPERTY CONDITION DATA (11 YRS. SERV.) REQUIREMENT

FLEXURAL STRENGTH, DRY 52,400 51,900 50,000

PSI WET (b) 54,300 46,900 45,000

FLEXURAL MODULUS, DRY 2.54 2.52 2.50

PSI x 10* WET 2.49 2.37 2.30

COMPRESSIVE STRENGTH, DRY 39,100 33,000

PSI WET 35,600 28,000

(a) MIL- P- 17549

(b) TWO HOUR BOIL

Figure 2

281+
5 70 z

60 - <I

50

• VOID CONTENT <l%

VOID CONTENT 2 1/2%

J L

4 5 6 7 IMMERSION TIME, MONTHS

10 11 12

Figure 3 - Effect of Extended Immersion on the Interlaminar Shear Strength of Graphite Epoxy

Composites

EXPOSURE TIME. MINUTES

Figure 4 - Cavitation Erosion Rate for Various Materials

285
COMPOSITE TECHNOLOGY FOR MARINE APPLICATION

W. P. Couch

David W. Taylor Naval Ship R£jD Center

Bethesda, MD 20084

Abstract: Under the sponsorship of the Naval Sea Systems Command

(NAVSEA 035) , the David W. Taylor Naval Ship Research and Development

Center (DTNSRDC) has been conducting a program to assess the current

technology status of advanced composites, and to determine the feasibility of application to marine structures. The approach being utilized

to achieve this objective involves a coordinated structures/materials

community effort seeking input from private industry and other government agencies. As a result of feasibility studies performed by the

McDonnell Douglas Astronautics Company and the Grumman Aerospace

Corporation and a review of high payoff areas for structural application

of advanced composites by the Naval Ship Composites Community, the

hydro-foil strut/foil system was selected as the primary structural

element to evaluate the use of advanced composites for advanced naval

ship structures. Two subcomponents were selected for this evaluation;

a foil test component (tapered box beam), and a hydrofoil control flap.

Concurrently, analytical methods incorporating composite mechanics

developed by the aerospace composites program are reviewed, modified

and utilized when applicable to determine the structural behavior for

marine, load-carrying components. In addition, preliminary design

studies are conducted on other naval ship structures to determine the

feasibility of composite construction. The following text describes the

current effort under investigation related to the four program subtasks:

box beam validation, foil flap validation, analytical methods and design

application studies.

Key Words: Advanced composites, structural analysis, advanced ships,

nondestructive tests, fatigue tests, elastic properties, ultimate

strength

.

Box Beam Validation: The objective of the box beam subtask is to assess

the fatigue behavior of a typical graphite/epoxy (gr/ep) structure under

simulated sea loads, and compare its behavior with current metal designs.

The McDonnell Douglas Astronautics Company has designed and fabricated! > 2

two box beams consisting of T-300/GY-70 hybrid skins and HY-130 steel

spars. This design represents a strut/foil system having an approximate

weight savings of 25% compared to current metallic strut/foil systems

fabricated entirely of steel with a yield strength of 130 ksi. Figure 1 shows schematically the development of the box beam configuration from

the forward foil of the hydrofoil, PCH-1 (HIGH POINT).

286
During the Preliminary Design Phase, several concepts were developed and

the subject design was selected for further investigation. The laminate skins consist of 53 plies O 58%) of T-300/5208 at 0° orientation

and 32 plies O 42%) of GY-70/5208 at ± 45° orientation. The 85 plies

making up the 1/2 inch thick hybrid composite were dispersed according

to the following layup:

[0°, C0|, -45°) 8 , 01, (±45°, 0§) 8 , 0°]

During the Design Verification Phase, the laminate was characterized to

validate the theory used to predict the elastic and ultimate strength

properties for a hybrid laminate. All averaged experimental data agreed

with predicted properties within 7.4%. An in-depth joint evaluation was

conducted as the next step of design verification. Both the scarf transition joint and the steel spar to gr/ep skin joint were modeled and

evaluated under static and fatigue loading, which led to a 33% reduction

in the design ultimate strength of the hybrid laminate. However, this

reduction did not significantly sacrifice the efficiency of the overall

box beam design since the steel spars are loaded to 90% of yield strength

when the hybrid laminate is stressed to the reduced design ultimate

strength.

Following the static and fatigue testing of joint models, the box beam

design was finalized; see Figure 2. Even though the strength of the

adhesive was adequate to transmit the shear stress from the skins to the

stiffeners without failure of the adhesive, the design included bolting

of the skins to the internal metal structure. The bolted/bonded design

was based on reliability considerations of both the box beam and of the

full size foil or strut. A NASTRAN finite element analysis was performed

on the final design, and compared with the strength of materials analysis

performed on the preliminary design; Figure 3 shows the maximum fiber

stress in the composite skins and steel spars along the longitudinal

centerline of the box beam as a result of the NASTRAN and strength of

materials analyses.

The Fabrication Phase included cocurring the steel transition scarf and

composite skins, and a thru-transmission ultrasonic C-scan inspection;

machining, welding and radiographic inspection of the steel spars and

load plate; and bolting/bonding the composite skins to the steel spars,

and pulse-echo ultrasonic and Fokker Bond Inspection after assembly.

The nondestructive evaluation of the components and finished assembly,

see Figure 4, revealed no detectable defects as a result of laminate

fabrication or box beam assembly.

The first box beam was scheduled to undergo a static proof test to a

maximum load of 60,000 lbs which would induce an extreme fiber stress

in the composite skins of 44 ksi (Design Ultimate Strength/Factor of

287
Safety) . Although the box beam incurred delaminations in the load

application area at 80% of maximum operating load, elastic strains

measured by electrical resistance strain gages agreed with those predicted by the NASTRAN finite element analysis; see Figure 3. The area of

delamination was inspected using radiographic and pulse-echo ultrasonic

test techniques, and the suspected areas of defects repaired. The load

area was also "stiffened" in an attempt to help distribute the shear in

this area. A second static proof test was conducted; a similar failure

occurred in the same area at a load of 47,500 lbs. It is currently

awaiting repair and additional testing while a 3-dimensional finite

element analysis (SAP 4) is being performed on the load application

detail to explain the cause of failure.

Foil Flap Validation: The second structural component to be addressed

is one of the aft inboard control flaps on the PCH-1. The aim in

developing a composite flap was to obtain in-service performance data

relative to operational loads in the marine environment. An aft inboard

flap was chosen because it is the least likely of the flaps to see

extreme load conditions, such as impact, for which the program has not

yet developed design/analysis techniques. The original plan was to

fabricate two flaps; one to be evaluated in the laboratory under full

lifetime simulated service loading, and the second to be put into fully

instrumented service on the ship. However, the phase out schedule for

the PCH-1 prevented following this approach; the first flap will be

evaluated in service, to be followed by the laboratory test.

The Boeing Company was awarded a 3-phase contract to design and fabricate

an advanced composite flap completely interchangeable with the present

metal design. Seven conceptual designs were evaluated during the

Preliminary Design Phase; the baseline design was selected for further

study. This concept was composed of titanium-clad, variable thickness

gr/ep skins and a titanium crank-spar assembly. T-300/934 bi-directional

preimpregnated gr/ep fabric oriented at 145° to the flap span and 6A14V

ELI -grade Ti were incorporated into the design. The titanium cladding

(.010-inch thick on both sides of each gr/ep skin) was added to protect

the composite from water penetration, cavitation erosion, and small

particle impact. The ELI-grade Ti was selected for the crank-spar

assembly in order to protect against corrosion fatigue and stresscorrosion cracking. The gr/ep fabric was oriented at —45° to take

advantage of its torsional and shear characteristics in that layup. Since

the variable thickness skins appeared inadequate to carry the evenlydistributed 6300 psf design load, DTNSRDC conducted a NASTRAN analysis

of the baseline design. The results led to constant 0.5-inch thick

skins including the Ti cladding; see Figure 5. Once the skin design was

finalized, a full range of static and fatigue tests were conducted for

mechanical property and salt water exposure data; see Table 1. The

results of these tests indicated no significant degradation in material

performance after exposure up to 270 days.

288
In order to verify the design and fabrication processes and to evaluate

joints, a full-scale portion of the flap was built and tested. This

feasibility component incorporates the first 20 inches of the flap, end

joint and crank fitting. The feasibility component sustained a static

proof test to maximum operation load or 40% of design ultimate. The

component failed after approximately 12,000 cycles (safe life = 16 x 106

cycles) due to a debond along the leading edge, which changed the load

path and resulted in a catastrophic failure of the titanium end fitting.

A more detailed NASTRAN analysis was conducted, and the component was

repaired and beefed up in the leading edge joint and end fitting areas.

A second fatigue test resulted in catastrophic failure of the titanium

end fitting after 145,000 cycles. The results of these tests led to a removal of all blind fasteners due to their poor fatigue performance

and a redesign of the crank-spar assembly for the final design.

A full scale component has been fabricated using the techniques developed

during the fabrication of the feasibility component. The covers were

made by laying up 36 plies of gr-ep cloth and cured at 350°F and 100 psi

in an autoclave. The titanium cladding was bonded to the covers using

EA9628 adhesive at 250°F and 100 psi. The crank and spar details were

cut to approximate shape with a plasma torch, rough machined to their

approximate final configuration, electron beam welded into a single

assembly, and milled and bored to final configuration. The composite

flap details were stacked for final assembly, bolted with titanium

fasteners, and bonded using the EA9628 adhesive at 250°F for 90 minutes

and 50 psi. The 52 internal strain gages were calibrated in the

laboratory, the flap was installed and actuated on the PCH-1 while on

blocks (see Figure 6), the instrumentation was hooked up to the ship's

data acquisition system, and the gages again calibrated. Both calibrations indicated that the gages are operating accurately, and that no

major performance deficiencies currently exist. The ship is scheduled

to go back into the water this summer, with calm and rough water trials

scheduled to be completed by the end of 1979. Following the in-service

evaluation, the flap will be shipped to DTNSRDC for a full lifetime

fatigue evaluation in salt water.

Analytical Methods: Because of the anisotropic nature of fiber-reinforced

plastics, strength of material and classical fracture/fatigue methods for

analyzing their structural behavior are approximations at best. Instead,

a thorough understanding of composite mechanics is required to perform

an accurate analysis of complex structures. This includes micro and

macro mechanics, lamination theory and failure mechanics. Much of this

work has been developed by academia and industry under aerospace sponsored programs. Often, the application of this theory for the design/

analysis of marine structures subjected to the unique loads and environment of the ocean requires extension and modification.

A basic knowledge of the stress-strain and transformation relations of

an orthotropic, multi-directional laminate was required to develop the

289
extensional and bending stiffness matrices for the NASTRAN finite element

analyses of the box beam and foil flap. In addition to the stress-strain

and transformation relations, both the distortion energy and maximum

strain failure theories were utilized to predict the elastic and strength

properties of the hybrid laminate during the design phase of the box

beams. Both the material characterization of the box beam laminate and

the static strain-gage evaluation of the box beam structure agreed with

the theoretical methods described above. It should be noted that the . aforementioned analyses have all been 2-dimensional , i.e., the thruthickness elastic stiffness and strength properties were not considered.

However, the in-plane analyses have not explained the reasons why the

box beam has incurred two failures; both of which may well have been

interlaminar in nature as opposed to in-plane. It is for this reason

that a 3-dimensional analysis is under way; an analysis that is an

order of magnitude more complex to perform.

Because graphite/epoxy has demonstrated a very low resistance to impact

loading when compared to metals and even glass or kevlar/epoxy , a new

study has been initiated to develop design/analysis techniques for this

extremely complex dynamic phenomenon. As presently envisioned, a

combined material/structural dynamic response computer code will be

developed that would give the designer/analyst an effective tool in

predicting the behavior of a composite structure under a variety of

impact loads such as tools, logs, etc.

Design Application Studies: Although the hydrofoil strut/foil system

was selected as the primary structural element to validate the current

technology base of advanced composites for marine application, feasibility studies are conducted on other structures with respect to reduced

weight and life-cycle costs and increased reliability. Several potential

components have been evaluated on SES and ACV type advanced ships

including logitudinal hull girder reinforcement, masts, ramps and lift

fans. Unless a larger emphasis is placed on making these advanced

ships an important part of the fleet, the potential weight savings and

increased reliability do not warrant the increase in initial costs

associated with research and development of advanced composite materials.

References

:

1. Greszczuk, L. B., A. V. Hawley and W. P. Couch, "Application of

Hybrid Composites to Hydrofoil Structures," Presented to AIAA/SNAME

Advanced Marine Vehicle Conference, San Diego, CA, 17-19 April, 1978.

2. Greszczuk, L. B. and W. P. Couch, "Design, Fabrication, and Nondestructive Evaluation of an Advanced Composite Foil Test Component

(Tapered Box Beam)," ASTM STP 674, Composite Materials: Testing and

Design (Fifth Conference), June, 1979.

290
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291

Figure 2 - Schematic of Hybrid Graphite/Epoxy Skin-Skeleton

Box Beam

293
DISTANCE FROM BASE PLATE, IN.

Figure 3 - Comparison of Longitudinal Bending Stresses in Composite Skin

and Steel Spar between Strength of Materials and NASTRAN Analyses

29k
295
Figure 5 - Patrol Craft Hydrofoil (PCH-1) - Composite Flap

296
297



MECHANICAL FAILURES PREVENTION GROUP

ADVANCED COMPOSITES:

DESIGN AND APPLICATIONS

29th MEETING

May 23-25, 1979

ATTENDANCE LIST

Jamil A. Abbasi

Federal Aviation Administration

800 Independence Avenue, SW

Washington, DC 20591

N . Ahmed

Perkin Elmer Co.

100 Wooster Heights

Mail Stop 823

Danbury, CT 06810

Arnold Anderjaska

Federal Aviation Administration

800 Independence Avenue, SW

Washington, DC 20591

Alfred J. Babecki

NASA/Goddard

Code 313

Greenbelt, MD 20771

D. M. Boyd

Lawrence Livermore Labs.

P. O. Box 808

Livermore, CA 94550

W. H. Browne, Jr.

NASA/Goddard

Code 7 31

Greenbelt, MD 20771

Harry C. Burnett

National Bureau of Standards

Materials Bldg. , Room B264

Washington, DC 20234

James M. Candela

Naval Air Rework Facility

Naval Air Station

Norfolk, VA 2 3511

Frank J. Cepollina

NASA/Goddard

Greenbelt, MD 20771

F . H . Chang

General Dynamics, Ft. Worth Division

P. O. Box 748

Fort Worth, TX 76101

Stanley L. Channon/STD

Institute for Defense Analyses

400 Army-Navy Drive

Arlington, VA 22202

Rustam Chinoy

Mechanical Technology Inc.

968 Albany/Shaker Road

Latham, NY 12110

Peter J. Cloud

LNP Corp.

412 King Street

Malvern, PA 19355

J. A. Corrado

David W. Taylor Naval Ship

R&D Center

Code 1702

Bethesda, MD 20084

W. P. Couch

David W. Taylor Naval Ship

R&D Center

Annapolis, MD 21402

Clifford K. Deakyne

E. I. DuPont de Nemours & Co.

Chestnut Run, Bldg. 701

Wilmington, DE 19898

301
Jerry D. Deans

Naval Air Rework Facility

4 361 Gadwall PI.

Virginia Beach, VA 23462

Wallace J. Downs

General Electric Space Division

Valley Forge, PA 19481

Harold P. Edelstein

David W. Taylor Naval Ship

R&D Center

Code 2823

Annapolis, MD 21402

Gerard Emptoz

French Scientific Mission

2129 Wyoming Avenue, NW

Washington, DC 20008

Joseph D. Feko

Owens Corning Fiberglass Corp.

2075 W. Big Beaver Road

Suite 500

Troy, MI 48084

David S. Fine

Chevrolet Engineering Center

30003 Van Dyke, C-2-232

Warren, MI 48090

Aaron Fisher

NASA/Goddard

Code 313

Greenbelt, MD 20771

David R. Geise

Naval Air Engineering Center

GSED 92623

Lakehurst, NJ 087 3 3

Joseph G. Gerber

General Dynamics-Fort Worth Div.

P. 0. Box 748

Fort Worth, TX 76101

R. F. Gilby

E. I. DuPont de Nemours & Co.

Textile Fibers Dept.

Chestnut Run

Wilmington, DE 19898

Robert Gordon

Structural Composites Industries, Inc.

6344 N. Irwindale Avenue

Azusa, CA 91702

W. W. Gunkel

General Motors Inc.

Warren, MI 48090

Frank Guzzetta

Armco-Research Center

Curtis Street

Middletown, OH 45043

Steven G. Hackett

NASA/Goddard

Greenbelt, MD 20771

S. L. Hanlein

NASA/Goddard

Code 731

Greenbelt, MD 20771

James F. Haskins

General Dynamics, Convair Div.

P. 0. Box 80847

San Diego, CA 92138

Henry R. Hegner

Mantech of New Jersey Corp.

2 341 Jefferson Davis Highway

Suite 930

Arlington, VA 22202

Carl T. Herakovich

Virginia Polytechnic Inst, and

State University

120 Norris

Blacksburg, VA 24061

Robert R. Holden

Hughes Aircraft Co.

Bldg. 117, Mail Station 1

P. 0. Box 90515

Los Angeles, CA 90009

G. H. Honeycutt

NASA/Goddard

Code 7 31

Greenbelt, MD 20771

302
W. D. HumphreyBrunswick Corp. , Defense Division

4300 Industrial Avenue

Lincoln, NB 68504

S. E. J. Johnsen

Detroit Diesel Allison Division

General Motors Corporation

Plant 8, T9

Indianapolis, IN 46206

Michael David Judd

Materials Section

Product Assurance Division

European Space Technology Center

Noordwijk, The Netherlands

Geza Kardos

Carleton University

Ottawa, Canada

S . D . Karmarkar

Metallic Materials Branch

Naval Surface Weapons Center

Dahlgren, VA 22448

J. J. Kelly

David W. Taylor Naval Ship

R&D Center

Annapolis, MD 21402

Ronald Khol

Machine Design

Penton Plaza

Cleveland, OH 44114

E. Kirchman

NASA/Goddard

Greenbelt, MD 20771

E. E. Klaus

133 Chemical Engineering Bldg.

The Pennsylvania State University

University Park, PA 16802

Karyn Knoll

2/LT, USFF

Air Force Materials Lab.

Wright Patterson Air Force Base

Ohio 45433

Hari Kulkarni

Ford Motor Co.

Scientific Research Labs

P. O. Box 2053

Dearborn, MI 48121

Arthur Lazarow, Jr.

Naval Air Engineering Office

Naval Air Station, Code 343

Norfolk, VA 23511

W. A. Leavy

NASA/Goddard

Code 731

Greenbelt, MD 20771

C. L. Leung

Rockwell Intl Science Center

1045 Camino Dos Rios

P. 0. Box 1085

Thousand Oaks, CA 91360

Richard Margolis

Plastics Division, Rockwell Intl

1049 Camino Dos Rios

P. O. Box 1085

Thousand Oaks, CA 91360

R. E. Maringer

Battelle Columbus Labs

505 King Avenue

Columbus, OH 43201

Frank T. Martin

NASA/Goddard

Code 753

Greenbelt, MD 20771

P. W. Mast

Naval Research Laboratory

Washington, DC 20375

Robert H. McDonnell

NASA/Goddard

6407 McCahill Drive

Laurel, MD 20810

303
Ernest W. Mielke

NASA/Goddard

Code 313

Greenbelt, MD 20771

D. R. Mulville

Naval Research Laboratory

Washington, DC 20375

Robert L. Mumper

Trident Engineering Associates

48 Maryland Avenue

Annapolis, MD 21401

D. E. Murray

NASA/Goddard

Code 731

Greenbelt, MD 20771

H. C. Nash

Naval Research Laboratory

Washington, DC 20375

Earl Nicewarner

Fairchild Space & Electronics Co,

20301 Century Blvd. , A7

Germantown, MD 20760

Bryan R. Noton

Battelle Columbus Labs.

505 King Avenue

Columbus, OH 43201

Maurice A. Nygaard

NASA/Goddard

Code 741.2

Greenbelt, MD 20771

James S. O'Brien

NASA/Goddard

Greenbelt, MD 20771

Derek J. Palmer

Defense Equipment, Staff MatIs

Officer

British Embassy

3100 Massachusetts Avenue, NW

Washington, DC 20008

John J. Park

NASA/Goddard

Code 313

Greenbelt, MD 20771

Elio Passaglia

National Bureau of Standards

Materials Bldg. , Room B308

Washington, DC 20234

Chris Passerello

Michigan Tech. Univ.

Dept. of ME. -EM.

Houghton, MI 49931

William L. Phillips

Naval Research Laboratory

Washington, DC 20375

John A. Philosophos

Allis-Chalmers

P. 0. Box 512

Milwaukee, WI 53201

R. D. Pistole

GTR Reinforced Plastics Co.

P. 0. Box 966

Southfield, MI 48037

Chet Poranski

Naval Research Laboratory

Code 6110

Washington, DC 20375

J. Howard Powell

Rockwell International-Tulsa Div.

2000 N. Memorial Drive

Tulsa, OK 74151

Mark R. Probst

Harry Diamond Labs

U.S. Army, Branch 36200

2800 Powder Mill Road

Adelphi, MD 20783

Yapa D. S. Rajapakse

Naval Research Laboratory

Code 6371

Washington, DC 20375

30k
Marwood E . Rand

Wyman-Gordon Company

Worcester Street

North Grafton, MA 01536

James J. Rast

NASA/Goddard

Greenbelt, MD 20771

Verlyn Reneau

Cessna Aircraft Co.

Wallace Division

P. 0. Box 7704

Wichita, KS 67277

James Rinde

Lawrence Livermore Lab.

P. O. Box 808, L-338

Livermore, CA 94550

A. W. Ruff

National Bureau of Standards

Materials Bldg. , Room B264

Washington, DC 20234

Depsek Sanghani

Perkin Elmer Corp.

100 Wooster Heights Road

Danbury, CT 06810

Donald K. Saylor

Union Carbide Corp.

12900 Snow Road

Parma, OH 44136

W. W. Schmidt

Brunswick Corp.

Defense Division

4 300 Industrial Avenue

Lincoln, NB 68504

S. J. Schneider

National Bureau of Standards

Materials Bldg., Room B308

Washington, DC 20234

John J. Scialdone

NASA/Goddard

Code 725

Greenbelt, MD 20771

James N. Scott

NASA/Goddard

Code 302

Greenbelt, MD 20771

T. R. Shives

National Bureau of Standards

Materials Bldg., Room Alll

Washington, DC 20234

M. Silvergleit

David W. Taylor Naval Ship

R&D Center

Annapolis, MD 21402

Charles 0. Smith

Univ. of Nebraska

College of Engineering

60th & Dodge

Omaha, NE 68182

Joseph R. Soderquist

Federal Aviation Administration

800 Independence Avenue , SW

Washington, DC 20591

Lawrence R. Sparrow

COMSAT Labs.

P. 0. Box 115

Clarksburg, MD 207 34

Charles L. Staugaitis

NASA/Goddard

Greenbelt, MD 20771

Robert Stedfeld

Editor, Materials Engineering

1111 Chester Avenue

Clveland, OH 44114

B. H. Stephenson

NASA/Goddard

Code 731

Greenbelt, MD 20771

J. E. Stern

NASA/Goddard

Code 721

Beltsville, MD 20771

305
L. D. Strege

NASA/Goddard

Code 731

Greenbelt, MD 20771

John L. Tarpley

NASA/Goddard

Code 313

Greenbelt, MD 20771

David F. Thompson

TM Development Inc.

Chester, PA 19013

Robert Y. Ting

Naval Research LaboratoryChemistry Division

Washington, DC 20375

Charles E. S. Ueng

Georgia Institute of Technology

225 North Avenue, NW

Atlanta, GA 30332

Alfred F. Vargas

Acoustic Emission Technology Corp.

P. 0. Box 408

Stratford, CT 06497

Charles E. Vest

NASA/Goddard

Code 313

Greenbelt, MD 20771

J . B . Wachtman

Director, Center for Matls Science

National Bureau of Standards

Washington, DC 20234

Jim Wadin

Dunegan/Endevco

Rancho Vie jo Road

San Juan Capistrano, CA 92675

A . S . Wang

Drexel University

Dept. of Mechanical Engineering

and Mechanics

Philadelphia, PA 19104

Tony K. Wang

General Motors Corp.

GM Technical Center

Bldg. MD-A, #96

Warren, MI 48090

J. B. Webb

NASA/Goddard

Code 731

Greenbelt, MD 20771

W. A. Willard

National Bureau of Standards

Materials Bldg., Room B120

Washington, DC 20234

Irvin Wolock

Naval Research Laboratory

Washington, DC 20375

George Wuerch

Northwest Alaskan Pipeline Co

1801 K Street, NW

Washington, DC 20036

H. W. Zussman

CIBA-GEIGY Corporation

Ardsley, NY 10502

306
MFPG PUBLICATIONS

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analysis and control.

Meeting No. 6 "Detection, Diagnosis and Prognosis",

December 6, 1968.

Meeting No. 7 "Failure Mechanisms as Identified

with Helicopter Transmissions",

March 27, 1969.

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June 25-26, 1969.

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Failure Through Design Methodology",

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January 21-22, 1970

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July 8-9, 1970

307
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October 19-20, 1970

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in Detection, Diagnosis and Prognosis"

January 25-26, 1971

8. Proceedings of Meeting No. 15 AD 725 200

"Failure Mechanisms: Corrosion"

April 14-15, 1971

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"Mechanical Failure Prevention Through

Lubricating Oil Analysis"

November 2-4, 1971

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April 25-27, 1972

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November 8-10, 1972

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308
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<3 U. S. GOVERNMENT PRINTING OFFICE : 1979 0 - 301-071+
NBS-114A (rev. s-7B)

U.S. DEPT. OF COMM. 1. PUBLICATION OR REPORT NO. 2.GoVt. Accession No. Z. ftecifmsnt's Accession No. BIBLIOGRAPHIC DATA

SHEET sp 563

4. title AND subtitle Advanced Composites: Design and Applications.

Proceedings of the 29th Meeting of the Mechanical Failures

Prevention Group, held at the National Bureau of Standards,

Gaithersburg, Maryland, May 23-25, 1979

5. Publication Date Aof rvVion 1Q7Q

-• ..mm n in, ,„,i„i Vk,.,., 1

- 1 1 . 7.JRJ3XKSEBX) EDITORS

T. Robert Shives and William A. Willard

8. Performing Organ. Report No.

9. PERFORMING ORGANIZATION NAME AND ADDRESS

NATIONAL BUREAU OF STANDARDS

DEPARTMENT OF COMMERCE

WASHINGTON, DC 20234

10. Pmjacl/Task/Work itett No.

11. Contract/Grant No.

12. SPONSORING ORGANIZATION NAME AND COMPLETE ADDRESS (Street, City, state, zip)

Center for Matls Science, NBS, Washington, DC 20234; ONR,

Arlington, VA 22217; NAVAIRSYSCOM (A.J. Koury) , Washington,

13. Type of Report & Period Covered

Final

DC 20360; NASA, Greenbelt, MD 20771; DoE, Fossil Energy,

Washington, DC 20545

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15. SUPPLEMENTARY NOTES

Library of Congress Number 79-600151

I I Document describes a computer program; SF-185, FIPS Software Summary, is attached.

16. ABSTRACT (A 200-word or less factual summary of most significant information, ft document includes a significant bibliography or

literature survey, mention it here.) These proceedings consist of a group of thirty one submitted entries (twenty one

papers and ten abstracts) from the 29th meeting of the Mechanical Failures Prevention

Group which was held at the National Bureau of Standards in Gaithersburg, Maryland,

May 2 3-25, 1979. The subject of the symposium was the design and application of

advanced composites. Special emphasis was directed to aerospace, aircraft, automotive,

marine, and industrial applications and design. Failure modes in advanced composites

were also discussed.

17. KEY WORDS (six to twelve entries; alphabetical order; capitalize only the first letter of the first key word unless a proper name;

separated by semicolons)

Advanced composites; applications of composites; composites; composite structures;

design using composites; fiber composites; fiber glass; filament wound composites.

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