NASA / TM--1999-209434

Quantification of Processing Effects on

Filament Wound Pressure Vessels

Robert A. Aiello

Modern Technology Corporation, Middleburg Heights, Ohio

Christos C. Chamis

Glenn Research Center, Cleveland, Ohio

December 1999

https://ntrs.nasa.gov/search.jsp?R=20000017929 2018-05-14T14:05:31+00:00Z
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NASA / TM--1999-209434

Quantification of Processing Effects on

Filament Wound Pressure Vessels

Robert A. Aiello

Modern Technology Corporation, Middleburg Heights, Ohio

Christos C. Chamis

Glenn Research Center, Cleveland, Ohio

National Aeronautics and

Space Administration

Glenn Research Center

December 1999
NASA Center for Aerospace Information

7121 Standard Drive

Hanover, MD 21076

Price Code: A03

Available from

National Technical Information Service

5285 Port Royal Road

Springfield, VA 22100

Price Code: A03
Table of Contents

Summary. .................................................................................. 1

Introduction ................................................................................ 1

Composite Analysis ...........................................................................

Coordinate Systems ...........................................................................

Mesh Generation ............................................................................ 2

Tape Pretension ............................................................................. 3

Linear Generation ............................................................................ 3

Polar Wrap Generation ........................................................................ 3

Hoop Wrap Generation .............................................. 3

Helical Wrap Generation ...................................................................... 4

Property Input ............................................................................... 4

Conclusions ................................................................................ 4

Appendix A--Input Deck Setup for Polar-Wound Pressure Vessels .................................... 5

Appendix B--Pressure Built-up Due to PRetension in a Polar-Wound Pressure Vessels ..................... 7

References ................................................................................. 8

Figures .................................................................................... 9

NAS A/TM-- 1999-209434 iii

Quantification of Processing Effects on Filament Wound Pressure Vessels

Robert A. Aiello

Modern Technology Corporation

Middleburg Heights. Ohio 44130

and

Christos C. Chamis

National Aeronautics and Space Administration

Glenn Research Center

Cleveland. Ohio 44135

Summary

A computational simulation procedure is described which is designed specifically for the modeling and analysis

of filament wound pressure vessels. Cylindrical vessels with spherical or elliptical end caps can be generated automatically. End caps other than spherical or elliptical may be modeled by varying circular sections along the x-axis

according to the end cap shape. The finite element model generated is composed of plate type quadrilateral shell

elements on the entire vessel surface. This computational procedure can also be sued to generate grid, connectivity

and material cards (bulk data) for component parts of a larger model. These bulk data are assigned to a user designated file for finite element structural/stress analysis of composite pressure vessels. The procedure accommodates

filament would pressure vessels of all types of shells-of-revolution. It has provisions to readily evaluate initial

stresses due to pretension in the winding filaments and residual stresses due to cure temperature.

Introduction

Filament wound pressure vessels, though simple to fabricate, have relatively complex internal structure especially in the domes. Additional complexities are introduced because of the presence of initial stresses. These stress

result from ( 1) the pretension in the filament tape (wet or prepreg) and (2"j the curing process. Accurate description

of the local filament location/direction and initial stresses are required for credible evaluation of vessel structural

integrity and durability. Computational methods ['or composite structural stress analyses are available (refs. 1 and 2).

These methods have recently been augmented to simulate the local filament location/direction and respective initial

stresses. A specific code for that purpose is identified as COBSTRAN for Composite Blade Structural Analysis

(ref. 2). The objective of the present article/report is to provide an overview of the COBSTRAN application to filament would pressure vessels and present some initial results.

The COBSTRAN computer code was developed and used successfully at NASA Glenn Research Center for the

design and analysis of composite turbofan (ref. 3) and turboprop blades (ref. 4) and also for composite wind turbine

blades (ref. 5). Its application has been extended to include more general models of three-dimensional solid and shell

structures and is currently being expanded to include filament wound tank-type structures and/or pressure vessels.

This code combines composite mechanics and laminate theory with a resident data bank of fiber and polymer

matrix properties. COBSTRAN is versatile in that it can be used as a preprocessor or post processor with general

purpose structural analysis computer codes or stand-along with its resident general purpose finite element structural

analysis module. As a preprocessor for several general purpose finite element structural analysis programs. (COSMIC NASTRAN/MSC/NASTRAN and MHOST (refs. 6 to 8)) COBSTRAN, generates a model with anisotropic

integrated laminate material properties. Stress or strain output form the analysis program is provided as input tot the

COBSTRAN post-processor which uses the composite mechanics and laminate theory routines to calculate individual ply stresses, strains, interply stresses, through-the-thickness stresses and failure margins (ref. 1). As an exception, COBSTRAN may utilize nodal l`orce output from MHOST for post-processing. MHOST is the resident general

purpose finite element structured analysis module in COBSTRAN.

NASA/TM-- 1999-209434 1
COBSTRAN, written in FORTRAN 77, is currently configured and tested for operation on various UNIX based

platforms including CRAY X-MP, Silicon Graphics and Hewlett Packard. The COBSTRAN/MHOST code is now

available on the PC for automatic pre- and post-processing. The various features of COBSTRAN for filament wound

pressure vessels are described in subsequent sections. Input deck setup sample case for a polar winding pressure

vessel is illustrated in Appendix A. The corresponding pressure build up due to pretension in the winding filaments

is illustrated in Appendix B.

Composite Analysis

A COBSTRAN filament winding processing diagram is depicted in figure I. COBSTRAN is designed to carry

out the many linear analysis required to efficiently design and analyze structural components made of multilayered

multidirectional fiber composites. Inputs to the COBSTRAN composite analysis are constituent fiber and matrix

material properties, factors reflecting the ply fabrication process, and ply orientation and location geometry.

COBSTRAN performs the micromechanics, macromechanics and laminate mechanics of these fiber composites.

See the COBSTRAN Theoretical/Programmers Manual (ref. 9). The orientation angle of the 1I-axis of a ply is

shown in figure 2. For each ply designation this orientation angle is internally calculated for each wrap that exists

at a finite element grid point. This angle (measured in degrees) is defined in a local composite coordinate system in

which the 0.0 degree axis is tangent to the tank surface and whose projection along the structural x-axis which is

parallel to the axial axis of the pressure vessel. The composite 33-axis is positive inward normal to the surface.

Outputs are the various ply and composite properties, composite structural response (accounting for bendingstretching coupling) and composite stress analysis including the results of combined-stress 'strength failure criteria'

and Hoffman's failure criteria. The output from COBSTRAN post-processing is designed to make use of information

generated in the preprocessor phase and stored in temporary binary files. Therefore, the COBSTRAN pre- and postprocessing problem must be run in sequence on the same machine.

Coordinate Systems

COBSTRAN utilizes a fight hand rectangular coordinate system for structures (vessel) in which the vessel lies

axially along the x-axis (fig. 2). For the filament wound tank modeling the input coordinates are assumed to be on

the internal surface of the model which corresponds to the mandrel external surface. These coordinates are automatically transferred to the mid-thickness position in accordance with the thickness of the liner and number of wrap ply

generated by COBSTRAN.

Mesh Generation

The mesh generation capability in COBSTRAN is designed to create two-dimensional plate type meshes. All

meshes are created generally with two-dimensional NASTRAN QUAD4 elements. Under the FWTANK option only

quadrilateral elements are created. Tank geometry is defined with user specified circular or elliptical cross-sections

by specifying radii and x-axis location. The density of the mesh is controlled by the user input in which the number

of circular sections is specified and also the number of elements about the circumference. COBSTRAN calculates

the surface geometry and generates a mesh (finite element model) based on this input.

The coordinates of the mesh are first generated on the radii specified by the user input parameters. These dimensions must be the internal dimensions of the vessel since the final thickness of the wraps is not initially known.

Following the ply generation and designation of which plys exist at each node, the position of each node is transformed half of the ply stack thickness and normal to the internal surface of the vessel.

NASA/TM-- 1999-209434 2
Tape Pretension

Filament winding pretension is used to keep the fibers taut and in place in order to prevent fiber wash and fiber

wrinkling. However, as the plys build-up the pressure generated by the vessel wall against the mandrel may be high

enough to cause mandrel buckling, or liner damage or fiber damage and substantial inner plies wrinkling after mandrel removal. To prevent any or all of those t¥om occurring the winding tension through the ply thickness build-up

is varied. This is usually referred to as filament winding or tape programmed pretension.

The external pressure generated by the tape pretension at any ply including the mandrel surface is calculated in

accordance with the paper: Residual Stresses in Filament Wound Laminates and Optimum Programmed Winding

Tension by C.Y. Liu and C.C. Chamis (ref. 10). Currently the algorithm programmed in COBSTRAN assumes an

infinitely stiff mandrel and the pressure is calculated only for hoop plys. The pressure generated by each plys pretension is accumulated as an external pressure in the finite element model when the input request includes the pneumonic 'PRESSURE LOAD'.

Liner Generation

A liner, usually a thin metal layer, is used at the inner surface of the shell to prevent seepage of the pressure

vessel contents through the composite wall thickness. In COBSTRAN the liner is designated by the key word

'LINER' in the wrap stack-up section of the input deck and is treated as another ply. Tile ply designation number is

provided on this input line. The thickness of the linear is indicated on the 'CIRCLE." 'ELLIPSE." "DSPHERE" or

"DELLIPS' lines of the parameter section of the input deck. The number of ply layers required to fill this specified

thickness is determined by the code. A sample case input deck setup is illustrated in appendix A. The corresponding

pressure build up is illustrated in appendix B. The initial residual stresses are evaluated as described (ref. 11). Both

of these must be added to those due to applied internal pressure.

Polar Wrap Generation

A polar wrap is designated by the key word 'POLAR' in the wrap stack-up section of the input deck. The angle

of the polar wrap is determined by calculating the offset of wrap over the total length of the vessel. The offset is

equal to the sum of the radii of both bosses plus the tape width. The polar wrap angle is equal to the arctangent of

the offset divided by total vessel length. The number of polar wraps per cycle is determined by dividing the maximum circumference of the tank by the tape width, adjusted for tape overlap and polar wrap angle. The starting

angle for a single polar wrap is defined from zero on the positive y-axis incrementing about the x-axis in the positive direction using the right hand rule. If the number of polar pairs is zero only one full circumferential set of wraps

will be made. If the number of polar pairs is one, the wrap will be designed as one wrap starting at the designated

starting angle and second wrap starting 180 ° from the starting position. This feature provides wraps at alternating

angles over the vessel for each polar pair indicated. If the number of polar pairs is greater than one. the number of

pairs are divided into 360 ° each pair will be designated as one wrap with the starting angle incremented from the

designated starting angle and a second wrap starting 180 ° from the incremented starting position (fig. 3).

Hoop Wrap Generation

A hoop wrap is designated by the key word 'HOOP' in the wrap stack-up section of the input deck. The hoop

wrap generated by this designation exists only over the cylindrical portion of the vessel. The HOOP ANGLE is

measured from a normal to the cylinder x-axis and the tape center line. If the HOOP ANGLE is given as 0.0. the

angle of the wrap will be calculated as the arctangent of the tape width, adjusted lor overlap, divided by the tank

circumference. The number of wraps for each HOOP call is equal to the tape width divided into the circumference.

One hoop wrap consists of one wrap in the positive x-direction over the length of the cylindrical section and a wrap

in the negative x-direction over the cylindrical section thus generating alternating ply angles. The starting angle for

a single hoop wrap is defined from zero on the positive y-axis incrementing about the x-axis in the positive direction

NAS A/TM-- 1999-209434 3
usingtherighthandrule.Thesubsequent wrapsareincrementedbytheratioof thenumberofwrapsdividedinto

360°.Thetapetensiongeneratesanequivalent external pressureonthevesselwhichwillexistinthefiniteelement

deckif thePRESSURELOADoptioniscalledintheCOBSTRANinputdeck(fig.l). A mandrel pressurebuilt-up

foraspecificvessel isillustratedinfigure4,

Helical Wrap Generation

A helical winding, designated by the key worked 'HELICAL' in the wrap stack-up section of the input deck. is

generated for a given circuit. A circuit is defined as one complete wrap starting at the boss going around the opposite

boss and returning to the original boss. The final position at the starting boss is determined by the tank shape and

helical angle used, This final position is the starting position for the next circuit, The incrementing of each circuit

and the tape width is included in the program. The helical wrap generated by this designation exists over the entire

vessel. The HELIX angle is measured from the cylinder x-axis and the tape centerline. One circuit consists of one

wrap in the positive x-direction over the length of the vessel from boss to boss and a continuous wrap in the negative

x-direction from boss to boss thus generating alternating ply angles. The starting angle for each helix call is defined

from zero on the positive y-axis incrementing about the x-axis in the positive direction using the right hand rule

(fig. 5). COBSTRAN/FWTANK can accommodate filament would shells with other shapes-of-revolution (ref. 12)

(see also fig. 6).

Property Input

Two options are available. Option one uses user-supplied ply properties. This option indicates that the unidirectional layer properties are known and will be user-supplied input. Plies are identified by ply designation number.

Option two uses constituent material properties and composite micromechanics to determine ply properties: The

appropriate ply designation, the type of fiber and matrix must be specified. Fiber and matrix types from the

COBSTRAN resident DATABANK that are most similar to those of the actual ply should be selected. COBSTRAN

will use the ply properties exactly as they are entered, but any other material properties will be calculated from the

fiber and matrix properties as given in the DATABANK (ref. 1).

Conclusions

An overview is described of a computational simulation procedure to evaluate filament wound pressure vessels.

The procedure uses composite mechanics with finite element structural analysis. Unique features of the procedure

include internally calculated filament/location direction, dome built-up thickness and pressure build-up due to pretension winding. Polar, helical and other filament wound patterns are accommodated as well as all types of sells-ofrevolution. The initial stresses due to pre-tension and cure temperature are evaluated from the pretension pressure

built-up and the vessel cure temperature. Sample input deck is illustrated in appendix A and pressure built-up in

appendix B for a polar wound pressure vessel for a 40-ply vessel the average pressure built-up is 30 psi.

NASA/TM-- 1999-209434 4
Appendix A

Input Deck Setup for Polar-Wound Pressure Vessels

The following deck setup demonstrates the application of tape pretension for a 5 in. radius cylindrical tank with

no end domes. The wall thickness applies only to the liner, the final thickness is determined following tile application of all wraps. The grid locations are then translated to the mid-thickness point of the ply stack-up. The 'TRANS"

and 'BOUND" parameters apply' only to a request lbr a "MHOST" model and are ignored by the 'MSC

NASATRAN' model selection in the option deck.

TITLE= DEMONSTRATION PROBLEM NO. 22B

TITLE= MODELING OF CYLINDER WITH NO DOMES

TITLE= USING EZREAD OPTION AND 'CIRCLE' FORMATS

TITLE- ALUMINUM LINER 0.02 THICK + HOOP WRAPS

TITLE= FIRST HOOP STARTING AT 0.0 DEGREES Y-AXIS

PREPROCESSOR

$DIAG 16 54

SDIAG 98 $ FILL MDES ARRAY WITH WRAP NUMBER NOT COMPOSITE ID

$ FOR DEBUGGING ONLY: NOT FOR MATERIAL ANALYSIS

DIAG. 59 $ DO NOT PRINT INTERNALLY GENERATED CIRCULAR CONTOUR

PRTOUT $ PRINT ALL ELASTIC PROPERTIES

FWTANK $ FILAMENT WOUND TANK MODELING

MHOST $ FORMAT OF THE MODEL GENERATED

$ICAN $ THE ICAN AND MICRO OPTIONS ARE REQUIRED

SMICRO $ FOR USING THE ICAN CODE FOR PRE- AND POST-PROCESSING

ECHO $ DISPLAY THE INPUT DECK

PROPERTY INPUT $ SPECIFY PROPERTIES FOR INNER LINER

PRESSURE LOAD $ GENERATE PRESSURE LOADING DATA

ENDOPTIONS $ ......... DELIMITER .......................................

IMAT 1

NDES 8

NYSPACES

NOPLY = 8

KSMF = 2

14

$ MATERIAL PROPERTY TYPE

$ NUMBER OF PLY DESCRIPTIONS

$ NUMBER OF CHORDWISE SPACES NOT NECESSARILY EQUAL

$ STRAIN MAGNIFICATION OPTION FOR POSTPROCESSING

NAME SECTION SPACES TO X-STA HELIX RADIUS WALL

NUMBER NEXT SEC ANGLE THICK

GEODESIC 1 8 0.0 25,0 5.0

NAME SECTION SPACES TO

I/UMB_R NEXT SEC

CIRCLE 1 1 0

CIRCLE 2 1 2

CIRCLE 3 1 4

CIRCLE 4 1 6

CIRCLE 5 1 8

CIRCLE 6 0 I0

$

$ NAME SECTION SPACES TO

$ NLD4B E R NEXT SEC

$

$ GEODESIC 15 8

S

$ NAME SECT ROT RADIAL

$ NUM AXIS DIREC

$

TRANS 1 1 2

TRA/4S 2 1 2

TRANS 3 1 2

PRESS

INT/EXT

0.04 i000.0 0.0

X-STA RADIUS WALL PRESS

THICK INT/EXT

0 5

0 5

0 5

0 5

0 5

0 5

02

02

02

02

02

02

O0 0 0

O0 0 0

O0 0 0

O0 0 0

O0 0 0

O0 0 0

X-STA HELIX RADIUS WALL PRESS

ANGLE THICK INT/EXT

12.0 25.0 5.0 0.04 I000.0 0.0

NASA/TM--1999-209434 5
Appendix A--Continued

TRANS 4 1 2

TRANS 5 1 2

TRANS 6 1 2

$

$ NAME SECT

$ NUM

$

DOF(MAX

BOUND 1 1

$ BOUND 7 1

ENDPARAMETERS $

$

$ NAME WRAP

$ STK-ORDER

$

LINER 1

HOOP 2

HOOP 3

HOOP 4

HOOP 5

ENDWRAPS $-

1

2

3

4

5

6

7

8

2

3

139

5

30

T300 EPOX .005

T300 EPOX .005

T300 EPOX .005

T300 EPOX .005

ALT6 ALT6 .010

T300 EPOX .005

T300 EPOX .005

T300 EPOX .005

0.6E+00

•5E÷03 117.0E+03

0.01E-01

•0E+03 30.0E+03

6 WITH A SPACE BETWEEN EACH)

DELIMITER .....................................

PLY TAPE PERCENT START HOOP TAPE

DESIGNATION WIDTH OVERLAP ANGLE ANGLE TENSION

5

8 1.0 0.0 0.0 0.0 45.0

7 1.2 0.0 45.0 0.0 35.0

8 1.4 0.0 90.0 0.0 25.0

1 1.6 0.0 135.0 0.0 15.0

DELIMITER ....................................

0.0 .60 0.0

0.0 .6O 0.0

0.0 .60 0.0

0.0 .60 0.0

0.0 .001 0.0

0.0 .60 0.0

0.0 .60 0.0

0.0 .60 0.0

5.595E-02 2.88E+07 1.26E+06 5.94E÷05 0.30E+00

5.62E+03 21.94E+03 7.2E+03 0.011E-06 12.5E-06

0.100E+00 1.00E÷07 1.00E÷07 5.94E÷05 0.30E+00

30.0E+03 30.0E+03 7.2E+03 0.011E-06 12.5E-06

NASA/TM-- 1999-209434 6
Appendix B

Pressure Built-Up Due to Pretension in a Polar-Wound Pressure Vessels

Tape Tension Analysis per Chris Chamis paper

a = inside radius of mandrel

b = outside radius of mandrel

c = outside radius of laminate

B = 0 if mandrel is infinitely stiff

E = Young's Modulus

e = Tange tial directio n

r = Radial direction

F = Modulus of orthotropic material

Ee = 1.94E+07 Er = 1.38E+06 F = 7.33E+05

S = Sqrt (Ee / Er) = 3.74614

Mandrel outside rad 5

a = S-(Ee/F) - B = -2.27E+01

= S + (Ee/F)+ B = 3.02E÷01

Layer Ply thick Radius Tension Pressure Total Pressure

1 0.01 501

2 0.01 5.02

3 0.005 5.025

4 0.005 5.03

5 0.005 5,035

6 0.005 5.04

7 0.005 5.045

8 0.005 5,05

9 0.005 5.055

10 0.OO5 5O6

11 0.005 5.065

12 0,005 5.07

13 0.005 5.075

14 0.005 5,08

15 0.005 5.085

16 0.005 509

17 0.005 5.095

18 0005 5.1

19 0005 5.105

20 0,005 5.11

21 0,005 5.115

22 0,005 5.12

23 0.005 5.125

24 0.005 5.13.

25 0.005 5.135

26 0005 5.14

27 0.005 5.145

28 0.005 5.15

29 0.005 5,155

30 0.005 5.16

31 0.005 5.165

32 0.005 517

33 0.005 5,175

34 0005 5.18

35 0.005 5.185

36 0.005 5.19

37 0.005 5.195

38 0.005 5.2

39 0.005 5.205

40 0.005 5.21

0 0

0 0

46 2.071602 2.07

46 2.083424 4.15

44 2.004071 6.15

44 2.015228 8.17

42 1.934205 10.10

42 1.944714 12.05

40 1.862052 13.91

40 1.871931 15.78

38 1.78766 17.57

38 1.796925 19.37

36 1711072 21.08

36 1.719739 22.80

34 1.632333 24,43

34 1.640417 26.07

32 1.551484 27.62

32 1.558999 29.18

30 1.468565 30.65

30 1.475525 32.12

28 1.383615 33.51

28 1.390034 34.9

26 1.296672 36.20

26 1,302562 37.50

24 1.207771 38.7

24 1.213146 39.92

22 1.116949 41.0

22 1.121819 42.16

20 1.024238 43.18

20 1.028616 44.21

18 0.929672 45.14

18 0.933568 46.07

16 0.833282 46.91

16 0.836707 47.74

14 0.735099 48.48

14 0.738063 49.22

12 0.635153 49.85

12 0.637666 50.49

10 0.533472 51.02

10 0.535544 51.56

Press : (To/p)'LN

Ee

B=

Em(b*b-a'a)

0"¢°¢ + a'b'b

2.0°b'b

(1-p.m)'b*b + (1 + p,m)'a'a

60F

i _ Total pressure

I I 1 I

10 20 30 40 50

Number of plies

AVERAGE 5.112375 26.6 1.28909 30.03

NASA/TM--1999-209434 7
References

1. P.L.N. Murthy, and C.C. Chamis: "Integrated Composite Anlayzer (ICAN),'" Users and Programmers Manual.

NASA TP-2516, March 1986.

2. R.A. Aiello and C.C. Chamis: "Large Displacements and Stability Analysis of Nonlinear Propeller Structures,"

NASA TM-82850. May 1982.

3. C.C. Chamis: "Vibration Characteristics of Composite Fan Blades and Comparison with Measured Data."

NASA TM X-71893, May 1976.

4. C.C. Chamis, and J.H. Sinclair: "Analysis of High Velocity Impact on Hybrid Composite Fan Blades," NASA

TM-79133, April 1979.

5. C.C. Chamis and T.L. Sullivan: "Free Vibrations of the ERDA-NASA 100 Wind Turbine." NASA TM

X-71879, March 1976.

6. Anon. NASTRAN User' Manual, NASA SP 222(03), 1978.

7. W.J. Anderson, "MS NASTRAN" Interactive Training Program, John Wiley and Sons, New York, 1983.

8. "3-D Inelastic Analysis Methods for Hot Section Components" (Fourth Annual Status Report), Volume I--

Special Finite Element Models, NASA CR-180893, Nakazawa, May 1988.

9. R.A. Aiello and C.C. Chamis: "Composite Blade Structural Analyzer (COBSTRAN)" Theoretical/

Programmer's Manual. NASA TM-101958, August 1989.

10. C.Y. Liu, and Chamis, C.C.: "Residual Stresses in Filament Wound Laminates and Optimum Programmed

Winding Tension" presented at the 20th Anniversary Technical Conference of the Reinforced Plastic Division

of The Society of the Plastics Industry, Inc., pp. 1-10, 1965.

I 1. C.C. Chamis: "Residual Stresses in Angleplied Laminates and Their Effects on Laminate Fracture." NASA

TM-78835, April 1978.

12. R.A. Aiello, "COB STRAN/FWTAN" Final Report and User's Manual, NASA Contract Report, in publication,

1999.

NASA/TM-- 1999-209434 8
I Structure to be

wound

geometry

• composite

specification Linear an0 system winding 0e°H pattern Types J__]

of

domes

I Mandrel

geometry

and type H _e,ect,r.,ension ___.tPerformfinite___

constant/ element structural

program med analyses Evaluate

pressure

built-up

I initial Eva uae ply I -= { e ove mandrel

stresses Evaluate ply }

stress --

redistribution

I Apply service

loading

conditions t-[,vauae t ro0uce o uer structure file to control

durability winding machine

Figure 1 .--COBSTRAN filament winding process diagram.

Ply,A Transverse

22 _f' f F- Fiber

6

=====================================

A33 J:-Y_-:::::::::::::::::---::-:;_7 --"

T J.._.:.-::::::::::::-:----::::::--.---_L)_l_ Longitudinal

._.:_:?_!:::..___.......................... -)_'jl __ Matrix

1;i::;:::::::......................... _;Tf Shear

I_ Shear 11

Angleplied laminate

Ply j

1

2

3

4

,_Y),_ r-- 900 ply

(/ i-- -0 ° ply

_r_._,r.r_/1 / I-- +0 ° ply

" " ';' '"" ';" '"' ';' / 0

!."/.:;.,:.;;..:..,:::/,_ _- o ply

," ,;,' ,;' ;,,' ;,,' ,:,' ,;' /

z ."".;:".;::'.;';".;;;'.';:".;':"

_ .'".::'".::' .;.:".:Y.:;'".;.;' ..... _i

_"_.'_'_'.._'._'_:-"-"-"7- / _-- Fiber

___;'_____- v...xc._'._.'______;___._..... _........ D x_ •

................... :?:i_!i_!_!:_;:.y

Matrix

COBSTRAN x-axis

Figure 2._Fiber composite geometry relative to

COBSTRAN coordinate system.

NASA/TM--1999-209434 9
Figure 3.--Tank with Spherical Endcaps.

-- .h==

==.,=

I

..=._

60--

"_ 45

_ ao

'1o

o I

0 10 20 30 40 50

Number of plies

Figure 4.mMandrel pressure built-up due to filament

winding pretension-resulting in initial stresses.

NASA/TM--1999-209434 10
Figure 5.--Tank with Elliptical Endcaps.

_J

_s

Figure 6.mEIliptical shell used to verify interply shear.

NASA/TM--1999-209434 11
REPORT DOCUMENTATION PAGE Form Approved

OMB No. 0704-0188

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1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

December 1999 Technical Memorandum

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

Quantification of Processing Effects on Filament Wound Pressure Vessels

6. AUTHOR(S)

Robert A. Aiello and Christos C. Chamis

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

John H. Glenn Research Center at Lewis Field

Cleveland, Ohio 44135-3191

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington. DC 20546-0001

11. SUPPLEMENTARY NOTES

WU-523-22-13-00

8. PERFORMING ORGANIZATION

REPORT NUMBER

E-11899

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA TM--1999-209434

Robert A. Aiello, Modern Technology Corporation, Middleburg Heights. Ohio 44130, and Christos C. Chamis, NASA

Glenn Research Center. Responsible person, Christos C. Chamis. organization code 5000, (216) 433-3252.

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified - Unlimited

Subject Category: 39 Distribution: Nonstandard

This publication is available from the NASA Center for AeroSpace Information. (301) 621-0390.

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

A computational simulation procedure is described which is designed specifically for the modeling and analysis of

filament wound pressure vessels. Cylindrical vessels with spherical or elliptical end caps can be generated automatically.

End caps other than spherical or elliptical may be modeled by varying circular sections along the x-axis according to the

end cap shape. The finite element model generated is composed of plate type quadrilateral shell elements on the entire

vessel surface. This computational procedure can also be sued to generate grid, connectivity and material cards (bulk

data) for component parts of a larger model. These bulk data are assigned to a user designated file for finite element

structural/stress analysis of composite pressure vessels. The procedure accommodates filament would pressure vessels of

all types of shells-of-revolution. It has provisions to readily evaluate initial stresses due to pretension in the winding

filaments and residual stresses due to cure temperature.

14. SUBJECT TERMS

Composite mechanics; Filament wound; Pre-tension: Cure stresses; Finite element:

Computer codes: Pressure built-up

17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION

OF REPORT OF THIS PAGE

Unclassified Unclassified

NSN 7540-01-280-5500

19. SECURITY CLASSIFICATION

OF ABSTRACT

Unclassified

15. NUMBER OF PAGES

18

16. PRICE CODE

A03

20. LIMITATION OF ABSTRACT

Standard Form 298 (Rev. 2-89)

Prescribed by ANSI Std. Z39-18

298-102