Astm stp 497 1972

ASTM STP 497 Composite Materials Testing and Design (Second Conference) COMPOSITE MATERIALS: TESTING AND DESIGN /SECOND CONFERENCE) A conference sponsored by the AMERICAN SOCIETY FOR TESTING AND MATERIALS Anaheim, Calif., 20-22 April 1971 ASTM SPECIAL TECHNICAL PUBLICATION 497 H T Corten, conference and program chairman List price $36.50 04-497000-33 ~ l j ~ AMERICAN SOCIETY FOR TESTING AND MATERIALS 1916 Race Street, Philadelphla, Pa 19103 by American Society for Testing and Mater&Is 1972 Library of Congress Catalog Card Number: 70-180913 NOTE The Society is not responsible, as a body, for the statements and opinions advanced in this publication Printed in Baltimore, Md February 1972 Foreword The Second Conference on Composite Materials: Testing and Design was held 20-22 April 1971 in Anaheim, Calif Committee D-30 on High Modulus Fibers and Their Composites of the American Society for Testing and Materials sponsored the conference, in conjunction with the American Institute of Mining, Metallurgical and Petroleum Engineers H T Corten of the College of Engineering, University of Illinois, served as conference and program chairman Of the 53 papers presented at the ten sessions, 36 are included in the volume, which complements the first conference publication, ASTM STP 460, Composite Materials: Testing and Design Related ASTM Publications I nterfaces in Composites, STP 452 (1969), $16.50 Composite Materials: Testing and Design, STP 460 (1970), $31.00 Contents Introduction Structural Design and Optimization Design, Analysis, and Testing of an Advanced Composite F-111 Fuselage J E A s h t o n , M L B u r d o f f , and F Olson Design Philosophy for Boron/Epoxy Structures - R N H a d c o c k 28 Properties Characterization I A Brief Survey of Empirical Multiaxial Strength Criteria for Composites G P S e n d e e k y j 41 Design and Fabrication of Tubular Specimens for Composite Characterization - J M W h i t n e y , N J Pagano, and R B Pipes A Finite Element Analysis for Stress Distribution in Gripped Tubular Specimens - R R R i z z o and A A Vicario 52 68 Properties Characterization II Mechanical Behavior of Carbon/Carbon Filamentary Composites - C W B e r t and T R Guess 89 Mechanical Behavior of Three-Dimensional Composite Ablative Materials - IV R A d s i t , K R Carnahan, and J E Green 107 Thermal Expansion Characteristics of Graphite Reinforced Composite Materials - W T F r e e m a n and M D C a m p b e l l 121 Fatigue and Environment Fatigue of Composites - M J A New Theory to Predict Reinforced Plastics Toughening Mechanisms in inforced Composites 143 Salkind Cumulative Fatigue Damage in Fiberglass L J B r o u t m a n and S S a h u Continuous Filament Unidirectionally Re- E F Olster and R C J o n e s 170 189 Properties Characterization III Analysis of Short Beam Bending of Fiber Reinforced Composites - C A Berg, J Tirosh, and M Israeli The Engineering Properties of Polymer Matrix Composite Materials by a Pure Moment Test - R G Hill and W E A n d e r s o n The Influence of Local Failure Modes on the Compressive Strength of Boron/Epoxy Composites - J A Suarez, J B Whiteside, and R 206 219 237 N H a d c o c k vii Copyright by ASTM Int'l (all rights reserved); Sun Jan 22:07:06 EST 2016 Downloaded/printed by University of Washington (University of Washington) pursuant to License Agreement No further reproductions authorized The Embedded Strain Gage Technique for Testing Boron/Epoxy Composites - L M Daniel, J L M u l l i n e a u x , F J A h i m a z , and T 257 Liber The Preparation and Testing of Miniature Carbon Fiber Reinforced Composites - J R M c L o u g h l i n 273 Acoustic Emission, Toughness, and Impact Acoustic Emission from Composite Materials - R G L i p t a i Toughness of Filamentary Boron/Aluminum Composites - J R H a n c o c k and G D S w a n s o n Impact Behavior of Unidirectional Resin Matrix Composites Tested in the Fiber Direction - R C N o v a k and M A D e c r e s c e n t e Impact Resistance of Unidirectional Fiber Composites - C C Chamis, M P H a n s o n , and T T S e r a f i n i Pendulum Impact Resistance of Tungsten Fiber/Metal Matrix Composites - E A Winsa and D 14/ P e t r a s e k 285 299 311 324 350 Components and Vibration B ehavior Stress Concentrations and Failure Criteria for Orthotropic and Anisotropic Plates with Circular Openings - L B G r e s z c z u k Reinforced Cutouts in Graphite Composite Structure - L H K o c h e r and 363 382 S L Cross The Strength of Bolted Connections in Graphite/Epoxy Composites Reinforced by Colaminated Boron Film - G E P a d a w e r Vibration Characteristics of Unidirectional Filamentary Composite Material Panels - R R Clary 396 415 Static and Fatigue Strength of Metal Matrix Composites Analysis of an Improved Boron/Aluminum Composite - R C J o n e s and 439 J L Christian Off-Axis and Transverse Aluminum Alloys The Initiation and Growth Aluminum Alloys - Tensile Properties of Boron Reinforced G D S w a n s o n and J R H a n c o c k of Fatigue Cracks in Filament Reinforced 469 J R H a n c o c k 483 Creep, Rupture, and Radiation Creep Behavior of Elastic Fiber/Epoxy Matrix Composite Materials - J B K o e n e m a n and T P K i c h e r Required Critical Aspect Ratio of Fibers Intended for Stress Rupture Applications - R W J e c h Neutron Irradiation Effects on a Metal Matrix Composite - K C Kvarn and R C J o n e s 503 516 525 viii Copyright by ASTM Int'l (all rights reserved); Sun Jan 22:07:06 EST 2016 Downloaded/printed by University of Washington (University of Washington) pursuant to License Agreement No further reproductions authorized Tailoring of Fibers and Composite Materials The Transverse Strength of Boron Fibers - K G K r e i d e r and K M P r e w o Evaluation of Unidirectional Glass-Graphite Fiber/Epoxy Resin Composites - L L K a l n i n Filament Misalignment and Composite Strength - W D Claus, Jr Tensile Behavior of Bead Filled Composites - L Nicolais, M Narkis, and 539 551 564 575 R E L a v e n g o o d Carbon/Carbon Composites - Solid Rocket Nozzle Material Processing, Design, and Testing - R C L a r a m e e and A l a n C a n f i e l d 588 ix Copyright by ASTM Int'l (all rights reserved); Sun Jan 22:07:06 EST 2016 Downloaded/printed by University of Washington (University of Washington) pursuant to License Agreement No further reproductions a STP497-EB/Feb 1972 Introduction Fibrous composite materials composed of stiff, strong fibers in a matrix have sparked the imagination of material scientists and engineers as the structural materials of the future Translation of the concept of fibrous composites into a primary load carrying structure has been and remains a challenging process Experience with ductile quasi-isotropic metals available in standard forms (plate, bar, forging, etc.) has lead to established procedures for selecting and testing materials, designing structures for stiffness and strength, fabricating and joining members, etc Conversely, fibrous composite materials exhibit little ductility in the usual sense and a degree of anisotropy in stiffness and strength that allows design of the material as well as the structure in which it will be used New and unorthodox fabrication methods are used in which the material and the structural component frequently are fabricated simultaneously In addition, joining methods can involve extensive use of adhesive bonding The designer of fibrous composite structures is presented with numerous degrees of freedom and an opportunity to exercise ingenuity totally unavailable to him with conventional materials However, this new freedom requires that the designer integrate into a single design step many aspects of the conventional step-by-step design process Conceptually, it is possible to design a structural component by starting with the properties of the constituent materials, the fibers and the matrix A parallel filament lamina is designed that provides the necessary stiffness and strength The laminated structural component employs as many fibers in selected locations and directions as are required to provide the stiffness and strength for the prescribed service loading conditions While modest successes (filament wound pressure vessels) have been achieved using simplified versions of this method, complicated structures subjected to multiple loading conditions exhibit prohibitive complexity To make the design process manageable, the concept of the "unidirectional lamina" was introduced as the "fundamental unit of material" in design This procedure employed test data obtained from a unidirectional lamina as the basis for design of laminated components and structures One step, the design of the material, was replaced by data obtained by testing the lamina This procedure shifted design attention from the micromechanics level of the material, to the macromechanics (anisotropic elasticity)level of the components and structures In retrospect, this concept of the lamina as the unit of material for use in design has restricted the freedom of the designer very little but has allowed him to get on with the job of designing and producing structural components Several years of experience with the concept of the lamina as the unit of material for design were reported in the First ASTIVl Conference on Composite Copyright*1972by ASTMInternational www.astm.org Copyright by ASTM Int'l (all rights reserved); Sun Jan 22:07:06 EST 2016 Downloaded/printed by University of Washington (University of Washington) pursuant to License Agreement No further reproductions authorized LARAMEE A N D C A N F I E L D ON ROCKET NOZZLE M A T E R I A L 595 TABLE 2-Carbon/carbonprocessing STEP I , RING OR CONE FORMING [ :MATERIAL GROUPS 1, 2, 3, & ] [ '.MATERAL G l ~ O ~ & ] REINFORCEMENT LAID UP C L O S E TO RING OR CONE NET REINFORCEMENT PLUS RESIN MOLDED AND CURED AT 200 TO 1000 PSI AND 300 F/POSTCURED AT 450 F STEP II ['i CARBONIZATION D LM.ENSIO NS I ['" MATERIAL, GROUPS 1, 2, 3, & MATERIAL GROUPS & ] I I CURED RING OR CONE ] PYROLYZED IN INERT ATMOSPHERE FURNACE AT 800 TO 1800 F A B STEP III REINFORCEMENT INFILTRATED WITH (CVD) PYROLYTIC CARBON IN INERT GAS OR VACUUM FURNACE AT 2500 TO 4000 F E ' PYROLYZED RING CARBONIZED IN INERT ATMOSPHERE FURNACE AT 2500 TO 4000 F [ iGRAPHITIZATION [ [ MATERIAL GROUPS 1, 2, 3, & I I.MATERIAL GROUPS &6 ] CARBONIZED BILLET GRAPHITIZED [ IN INERT GAS FURNACEAT 5000 TO 5500 F I STEP II CARBONIZED BILLET ORAPHITIZED OR STEP I REINFORCEMENT INFILTRATED WITH PYROLYTIC GRAPHITE (CVD) BOTH P R O C E S S E S IN INERT GASOR VACUUM FURNACEAT 5000 TO 5500 F STEP IV DENSIFICATION ] MATERIAL GROUPS 1, 2, 3, & I MATERIAL GROUPS &6 I REIMPHEGNATE CARBONIZED OR GRAPHITIZED RING OR CONE WITH RESIN AND CARBONIZE OR GRAPHITIZE, OR INFILTRATE WITH PYROLYTIC CARBON OR G RA PHITE REINFILTRATE CARBONIZED OR [ GRAPHITIZED RING WITH PYROLYTIC CARBON OR GRAPHITE ] I Copyright by ASTM Int'l (all rights reserved); Sun Jan 22:07:06 EST 2016 Downloaded/printed by University of Washington (University of Washington) pursuant to License Agreement No further reproductions authorized 596 COMPOSITE MATERIALS (SECOND CONFERENCE) five weeks Processing on most material products such as ~ the graphite cloth/graphite bond of group 1, the graphite fiber/graphite bond of group 3, the granular bulk graphite/pyrolytic graphite coating of group 4, the graphite yarn/graphite bond of group 5, and the graphite felt/pyrographite bond of group stops at this point In step IV for material of groups 1,2, 3, and 5, the carbonized or graphitized billet can be reimpregnated with a resin and the resin subsequently carbonized or graphitized In addition, all material groups with a carbonized or graphitized billet can be reinfiltrated with pyrolytic carbon or graphite There is no limit to the number of densification cycles, but generally they number between one and six cycles The cost, density, and process cycle vary directly with the number of densification cycles for this process step With an average number of graphitization densification cycles (one to three for a graphitized ring billet), the final billet production run cost could range from $100 to $200/lb and the process time could be increased by two to ten weeks to a total process time of five to thirty weeks The final density would range from a specific gravity of 1.30 to one of 1.90 Carbon/carbon composites can be tailored to meet numerous and varied design requirements by varying the type of reinforcement, the ply or grain direction of the reinforcement, the type of resin, the number of processing steps, and many other processing variables Generally, the CVD densification of pyrolytic carbon or graphite in a carbonized or graphitized billet results in a higher final density than a liquid resin or pitch impregnation densification However, a variable density cross section may result with this process owing to a high density buildup at the inside surface When a graphite or carbon (felt or yarn) is CVD infiltrated, a more uniform density is usually obtained because the porosity of the felt or yarn is high A typical cross-sectional view of a graphite yarn/graphite bond (Carbitex 715) composite is shown in Fig The transverse view shows the graphite yarn ends in a graphitized resin matrix, while the longitudinal or axial view also shows the continuous graphite yarn in the graphite matrix In general, the processing steps include the processing technology of the standard carbon cloth phenolic reinforced plastic, bulk graphite, and pyrolytic graphite materials For example, the bulk graphite process includes forming, baking, pitch impregnation, rebaking, and graphitization steps The maximum baking and graphitization temperatures are 1740 and 5430 F Material Properties Material design properties such as ultimate tensile, compressive, and shear strengths, Young's modulus, coefficient of thermal expansion, density, and thermal conductivity are available at room temperature through 5000 F, but more data are still required for complete application of the material The same Copyright by ASTM Int'l (all rights reserved); Sun Jan 22:07:06 EST 2016 Downloaded/printed by University of Washington (University of Washington) pursuant to License Agreement No further reproductions authorized L A R A M E E A N D C A N F I E L D ON ROCKET NOZZLE M A T E R I A L FIG 597 l-Carbitex 715-graphite yarn/graphite bond (artist duplication) statement is also true to a lesser degree for the standard nozzle liner materials of carbon or graphite cloth phenolic, bulk graphite, and pyrolytic graphite Carbon/carbon mechanical-thermal properties are determined in laboratory tests by the material manufacturer, by contract laboratories such as the Southern Research Institute, or by Thiokol Chemical Corp The variation of maximum ultimate tensile, compressive, and shear strengths with the grain, ply, or plane are shown in Table both at room temperature and at 5000 F for the six carbon/family groups and four standard materials Also two future carbon/carbon composites with high modulus, high strength yarn and cloth are shown with preliminary data to show potential material properties Generally, the carbon/carbon composite properties of group 1, graphite cloth/graphite bond, and group 5, graphite yarn/graphite bond, are equal or better than those of the standard liner materials The carbon/carbon materials all used a standard strength yarn and cloth of low ultimate tensile strength and low Young's modulus The substitution of a high Young's modulus, high ultimate tensile strength graphite yarn and cloth in a graphite bond matrix, as shown in the future carbon/carbon materials, offers excellent strength and modulus increases over the standard graphite cloth phenolic for future design applications while exhibiting the same range of density as the standard materials, except for the standard pyrolytic graphite (specific gravity of 2.20) The maximum compressive modulus, coefficient of thermal expansion, and thermal conductivity with grain, ply, or plane at 500 and 5000 F are shown in Table for standard materials and carbon/carbon materials present and future The carbon/carbon and standard materials compare on an equal basis with two Copyright by ASTM Int'l (all rights reserved); Sun Jan 22:07:06 EST 2016 Downloaded/printed by University of Washington (University of Washington) pursuant to License Agreement No further reproductions authorized NA 88.0 45.0 14.0 NA 3.0 3.7 18.0 3.0 35.0 5.5 13.0 NA 1.0 8.3 39.0 14.0 est 11.0 14.0 7.0 3.0 3.0 40.0 8.0 5000 F 5000 F 8.8 NA 3.0 21.0 NA NA 12.5 2.5 est 11.0 3.2 est NA NA NA NA Future Carbon~CarbonMaterials 10.0 36.0 7.6 12.0 Standard Liner Materials 12.0 15.0 8.0 13.0 27.0 10.0 est Carbon~Carbon Materials Room Temperature Ultimate Compression, ksi NA NA 0.4 1.5 3.0 1.5 3.5 2.0 4.0 1.3 2.0 1.5 Room Temperature NA NA 0.5 1.0 est 3.0 0.6 3.5 NA NA 3.3 2.0 NA 5000 F Ultimate Shear, ksi Copyright by ASTM Int'l (all rights reserved); Sun Jan 22:07:06 EST 2016 Downloaded/printed by University of Washington (University of Washington) pursuant to License Agreement No further reproductions authorized Group - h i g h modulus cloth Group - h i g h modulus yarn Graphite cloth phenolic Carbon cloth phenolic Bulk graphite Pyrolytic graphite Group Group Group Group Group Group Material Room Temperature Ultimate Tension, ksi TABLE 3-Carbon/carbon properties 1.52 1.46 2.20 1.40 1.90 1.40 1.45 1.38 1.30 1.90 1.46 1.80 Room Temperature Specific Gravity m m Z "1'1 m o z o z ro F- m E m o O~ "o r r O0 L A R A M E E A N D C A N F I E L D ON ROCKET NOZZLE M A T E R I A L 599 TABLE4-Standard and carbon/carbon materials, present and future Young's Modulus, Ec psi Material 500 F 5000 F Coefficient of Thermal Expansion, % in./in./deg F 500F Thermal Conductivity, Btu k, ft-h-deg F 5000F 500F 5000F 2.8• 10- NA NA 4.0 NA 5.2 34.0 28.0 12.0 NA 2.0 NA 72.0 20.0 est NA NA 80.0 19.0 Carbon/Carbon Materials Group Group Group Group Group Group 2.9• 106 1.6 1.6 1.2 est 4.0 2.2 NA 2.2• 10-6 NA 0.9 NA 1.2 0.2 X 106 est 2.6 NA 0.7 NA 1.0 Standard Liner Materials Graphite cloth phenolic Carbon cloth phenolic Bulk graphite Pyrolytic graphite 1.2 2.3 1.0 3.2 0A 0.2 0.2 1.7 1.3 6.0 1.2 0A0 1.5 2.0 3.1 1~ 1.0 3.0 0.5 4~ 70.0 19~ 160.0 46~ Future Carbon/Carbon Materials Group 1-high modulus cloth 17.0 Group 5-high modulus yarn NA NA NA NA NA NA NA NA NA NA NA exceptions The first exception is that the group thermal conductivity, k, is considerably higher at room temperature and 5000 F than that of the standard graphite and carbon cloth phenolics The other exception is that the Young's modulus, Ec, of future group composites is considerably higher than the E c of standard materials (17.0 vs 3.2 • 106 psi) or of existing carbon/carbon materials at room temperature The existing mechanical and thermal properties defining the carbon/carbon composites are good, but considerably more design data are needed before the materials can be as well characterized as the standard nozzle liner materials This material property comparison is good only for the with grain or ply direction Further evaluation of against grain or ply properties is necessary for full evaluation of the materials, as they differ considerably in comparison to with grain or ply properties Material Ablative Performance The 101 carbon/carbon composite rings were tested in 17 different solid rocket motors with as many different nozzle geometry configurations The motor burning time, chamber pressure, gas temperature, and nozzle internal diameters varied during the testing to establish material erosion loss curves under many varied conditions Copyright by ASTM Int'l (all rights reserved); Sun Jan 22:07:06 EST 2016 Downloaded/printed by University of Washington (University of Washington) pursuant to License Agreement No further reproductions autho 600 COMPOSITE MATERIALS (SECOND CONFERENCE) Nozzle Environment Typical material test nozzles were used to evaluate the 101 carbon/carbon rings as shown in Fig Two types of nozzles, submerged and external, provided the material erosion loss data The nozzles are fabricated with steel shells to resist pressure loads and are protected from the exhaust gas temperature by the ablative liner rings and cones bonded to the steel shell The throat ring is at the smallest internal diameter, with the inlet, nose, and submerged liner rings and cones located upstream from the throat ring, towards the burning propellant The exit cone is located downstream from the throat ring The liner environment includes a rapid pressure and temperature rise (1000 psi/s and 2500 F/s) at motor ignition During the average motor burning time of 60 s, the ablative wall surface supports an average gas pressure ranging from 500 to 900 psi, an average gas temperature ranging from 5500 to 3500 F, and a gas exhaust velocity (10 percent oxygen species) ranging from 60 to 2240 mph through the nozzle inlet, throat, and exit areas The area ratio is commonly used to define various areas along the axial center line length of the nozzle The area ratio is defined as the internal diameter plane area at any point divided by the internal diameter throat plane area at B on Fig Thus at the exit plane, throat plane, and inlet plane, the area ratios are 8.0, 1.0 and 4.0 at points A, B, and C, respectively, in Fig Carbon~Carbon Test Rings The 101 carbon/carbon nozzle rings tested as shown in the material background section were grouped by type of carbon/carbon reinforcement The EXTERNAL NOZZLE f SUBMERGED NOZZLE THROAT INLET ABLATIVE LINER THROAT INLET THROAT B- TI NE ABLA VE LI R C EXIT V ~ p / R ~ LOCATION A INLET Ap NOSE B THROAT C EXIT l S UBM i ~AND N D IFLANGEL F~I~:: ~EXIT GAS PRESSURE {PSI) 500 485 398 10 GAB TEMPERATURE ! F) 5.500 5,480 5,050 3,150 STEEL SHELL AND FLANGE ONE DIMENSIONAL CONVECTIVE HEAT GAS VELOCITY TRANSFER COEFFICIENT + SPECIFIC HEAT (H/CP) (MPH) 0,08 60 0.45 160 0.75 735 2,240 0,05 FIG 2-Nozzle liner material test areas and environment Copyright by ASTM Int'l (all rights reserved); Sun Jan 22:07:06 EST 2016 Downloaded/printed by University of Washington (University of Washington) pursuant to License Agreement No further reproductions authorized L A R A M E E A N D C A N F I E L D ON ROCKET NOZZLE M A T E R I A L 601 nozzle location where the carbon/carbon rings were tested and the number tested is shown below Carbon/Carbon Group Type Submerged Liner Nose-Inlet Throat Exit No No No No No No Yes (40) No Yes (8) No No Yes (7) Yes (14) No Yes (4) Yes (1) No No Yes (5) Yes (2) Yes (20) No No No The orientation of the with ply or grain direction of most of the test rings is as shown with respect to the nozzle center line (Fig 3a) The inlet, throat, and exit rings were fabricated with the cloth, felt, or fiber ply direction perpendicular to the nozzle center line in molding dies as shown (Fig 3b) The nose rings were fabricated with the cloth ply direction parallel to nozzle center line The cloth reinforcement was fabricated by tape wrapping onto a turning mandrel as shown (Fig 3c) The carbon/carbon rings were to in long, with 65 rings up to in in diameter, 33 rings to 10 in in diameter, and rings 10 to 16 in in diameter Over one half of the rings were tested in the nose-inlet nozzle area, with 27 and 19 rings tested in the exit and throat areas, respectively Over one half of the rings tested were group 1, graphite cloth/carbonaceous bond, and one third of the rings tested were group 3, carbonaceous fiber/carbonaceous bond No long carbon/carbon inlet or exit cones were tested Test Ring Erosion Profiles Typical material loss profiles of a few carbon/carbon and standard ablative ring materials tested together in the same nozzle are shown in Figs and A submerged nozzle test design shown in Fig was subjected to a motor gas environment that had a longer burning time (>60 s), a lower gas temperature (