40207_47

AEROSPACE EQUIPMENT AND INSTRUMENT STRUCTURE

47

Gay C. Krumweide and E d d y A. Derby

47.1 INTRODUCTION

It is necessary to discuss GFRP (graphite fiber reinforced plastics) thoroughly in order to understand the importance of composites in the manufacture of aerospace equipment and instrument structures. Before composite materials such as GFRP became viable candidate materials for aerospace primary structures, or even for sporting equipment, they were first used for aerospace equipment and instrument structures. Because composites exhibited both unique and superior properties, designers were willing to pay the prevailing high prices to achieve their design goals. For space hardware, where a pound of weight saved was worth thousands of dollars, designers were motivated to characterize composite materials suitable to their applications. Understandably, composites for primary and secondary structure (e.g. launch vehicles, aircraft frames, wing spars and skins, etc.) were a 'hard sell', and temperature extremes were too severe for 'thermosetting' GFRP to be used on missiles. The quantity of GFRP required for a particular piece of aerospace equipment or instrument structure was usually minimal, so relative costs were low, and composite materials' superior properties compared to heavy Invar, or high coefficient of thermal expansion (CTE)

Handbook of Composites. Edited by S.T. Peters. Published in 1998 by Chapman & Hall, London. ISBN 0 412 54020 7

aluminum, or very expensive beryllium fur- ther justified the material selection.

Consequently, numerous mirror bezels, telescopes, optical benches and reflectors were designed and built from GFRP in the early 1970s. As a result of these efforts, today more and more structures are being fabricated from GFRP materials, principally because specific materials matching specific property requirements are now available.

Although GFRP has dominated composite materials applications, DuPont's Kevlar-49 has been found to be ideal for antenna reflectors because of its extremely light weight and RF transparency. Many communication satellites utilize this type of Kevlar reflector, such as the SatCom-F, Telstar, ANIK-E, SpaceNET, and Superbird SCS (Fig. 47.1).

Fig. 47.1 SuDerbird SCS Kevlar dual-shell reflector.

Historical perspective and progress 1005

For aerospace equipment and instrument structures, generally, thermoplastic versions of GFRP have not seen as much application as the thermosets due to high investment costs in tooling and facilities. Considering the relatively small quantities that are usually bought, thermoplastic applications are not often cost effective. Also, the required high temperature cures subject the laminate to microcracking instabilities.

Metal matrix composite applications have been limited, and knowledge of material properties and processes has been restricted due to their use on classified programs. Materials like silicon carbides and carbon-carbon find limited, but important, use in aerospace structures, particularly mirrors. Design selection criteria plays an extremely important role in determining what type of material is used on any particular component of aerospace equipment and instrument structures. The primary reasons for choosing a composite material, rather than a metal, are weight, dynamic stability and thermal stability.

Table 47.1 categorizes some of the desirable and undesirable characteristics that con- fronted the early users of GFRP. Table 47.2 addresses the undesirable properties listed in Table 47.1 and indicates how the aerospace equipment and instrument designers have rec- ognized the barriers or problems with GFRP and found work-around techniques to allow their usage. Table 47.3 illustrates several structural applications where GFRP have been used

and/or are going to be used for aerospace equipment and instrument structures.

Table 47.4 compares the mechanical and thermal properties of candidate materials for aerospace equipment and instrument structures. One could ask, from the obvious property advantage of beryllium, why all such structures are not made from beryllium? If not beryllium, why not metal matrix composites or carbon-carbon (C/C) or silicon carbide (Sic)? The answer is that raw material cost, fabrication cost, practical size, and other critical properties all come into play for any specific application. Table 47.5 shows typical design requirements for various applications and indicates which materials typically satisfy the critical requirement.

The primary reason that GFRP is a material candidate for most applications is the wide range of material systems that are currently available that can compare with Kevlar, aluminum, Invar, beryllium, metal matrix composites, silicon carbide, and carbon-carbon.

47.2 HISTORICAL PERSPECTIVE AND PROGRESS

From a historical perspective, and to illustrate where significant progress has been made in the use of composites for aerospace structure, the following areas of interest will be addressed:

Table 47.1 GFRP properties

Desirable Undesirable

Low density Coefficient of thermal expansion (CTE near zero) High material cost High specific strength High specific stiffness Readily formable Crack growth resistant relief) Adaptable laminate properties Easily repairable

Low short transverse properties Hygroscopic

High fabrication cost Low impact strength Subject to microcracks (translaminar stress

Low peel strength

1006 Aerospace equipment and instrument structure

Table 47.2 GFRP design barriers and work-around techniques

Undesirable properties Barrier to design

Anisotropic Low strength in behavior angles

Thermal instability of angles

High through- thickness expansion of laminate

Hygroscopic nature Expansion/distortion

Material cost/ fabrication cost

High basic part cost

Susceptibility to Micro properties microcracks affected

Susceptibility to Low impact strength impact damage

Low peel strength Low joint allowables

Work-around techniques Reference

1. Radius blocks Stumm, 1981 2. Substitute material Campbell, 1981

(metals) locally Krumweide, 1988 3. Use mortise and

tenon joints 1. Use mortise and

tenon joints Campbell, 1981 2. Butt edges/insert Krumweide, 1988 3. Back-to-back splice 1. Egg crate joint Stumm, 1981 2. Mortise and tenon Krumweide, 1988 3. Add local Krumweide, 1991

compensation 4. Maximize in-plane

material orientation 5. Fitting Boss through

Skin 1. Moisture barrier 2. Define exposure/

3. Cyanate ester Telkamp, 1990

1. Proper material Krumweide, 1977 selection Krumweide, 1988

2. Tooling/fabrication techniques

3. Assess cost/weight ratio

4. Ovencure 5. Minimize pieces 6. Eliminate molded

7. Tailor joints/load 1. Uselow Stumm, 1981

temperature Krumweide, 1991 curing resin systems Brand, 1992

Stumm, 1981

Hertz, 1977; Stumm, 1979 Walrath, 1979; Levy, 1984

drying scenario Krumweide, 1989

Krumweide, 1991; Brand, 1992

parts

2. Low stacking angle 3. Small ply thickness 4. Reduce fiber modulus 5 . Thermal cycle parts 6. Cyanate ester 1. Protect surface with

2. Material design

1. Use mortise and

2. Fasteners and angle

Dunbar, 1978 Kevlar/honeycomb Herrick, 1984

(thickness, orientation)

tenon joints Krumweide, 1979

clips

Dunbar, 1978

Stumm, 1981


0 problems encountered/work-around

0 construction methods; 0 tooling development; 0 major milestones.

ites, principally GFRP, and indicates a variety of work-around techniques which have been implemented successfully to address these design concerns. To help explain and clarify these techniques, some examples of specific design solutions are discussed below.

techniques;

47.2.1 PROBLEMS ENCOUNTERED/WORK- AROUND TECHNIQUES Moisture effects

Aerospace designers have constantly been For GFRP, it has been determined that the pre- seeking the ideal material - the one material dominant mechanism of water penetration which meets design requirements with the into the laminate is through the resin by a dif- fewest compromises. Of course, few materials fusional process. The moisture diffusion is are completely ideal for a given set of require- governed by Fick's second law which is simi- ments and design trade-offs are almost always lar to Fourier's equation for thermal necessary. conductivity:

(47.1) 6 C 6% Table 47.2, above, summarizes typical prob-

'barriers to design' associated with compos- lems, presents some design concerns, and/or sf = Dzsz'

Table 47.3 GFRP Applications for aerospace structures

Equipment structures Instrument structures

RF Systems: -Reflectors (Viking, Nimbus-G, ACTS) -Feed horns and waveguides (INTELSAT) -MUX cavities -Diplexers -Phased arrays

Large segmented reflectors

Solar panel (substrates) (BS-3, RADARSAT, Mars Observer)

Booms, Stardust (Shuttle RMAB, M-SAT)

Bus (Mighty Sat M I , SMEX/WIRE, Indostar Forte)

Electronic chassis (EO1, SMTS-OBC, MARS 98, Stardust)

CCA cardguides (NAWC Coldplate)

Solar Concentrators

Metering structures (LANDSAT, Thematic mapper; SOHO, UVCS; HEAO-B, COSTAR and Hubble Telescopes)

Camera housings (Mars Observer Camera (MOC), Hubble WFOV Camera)

Optical benches (MAGSAT, UARS HRDI, Hubble FGS Optical Benches)

FPA or relay optics (Hubble Telescope; THEMATIC MAPPER)

Mirror bezel

Mirrors (Microwave Limb Sounder, NGST, INM Scan Mirror)

Lens holders

Support benches (Hubble FGS)

Startrackers (Hubble Equipment Shelf)

Submillimeter reflector

Helicopter mast mounts

Laser comm gimbals


where c = moisture concentration; t = time; Dz = moisture diffusivity; and z = thickness coor- dinate.

Diffusivity is dependent on resin type and temperature but is independent of moisture concentration or laminate orientation.

Figure 47.2 illustrates the importance of temperature in determining the diffusivity of a laminate. Figure 47.3 illustrates the moisture absorption behavior as a function of temperature for this same laminate. Note that this is for a 100% relative humidity (RH) exposure. Typical moisture absorption values (primarily by the resin) are 3.54% by weight for epoxy and 1-1.3% by weight for cyanate ester. The

maximum moisture pick-up for the composite (Mcm)may be calculated by the following equation:

MCm = (Mrn)[ x wr (47.2)

where Wr = resin weight YO and (MJr is the maximum moisture pick-up for the neat resin for a given relative humidity condition.

The moisture content as a function, Iw, can be represented by:

Mcm = A (WB (47.3)

where A and B are coefficients obtained from empirical data. Figure 47.4 depicts the percent of moisture pick-up for various levels of

Table 47.4 Typical mechanical and thermal properties of isotropic/quasi-isotropic materials ~~ ~

Material Modulus E Density CTE CP K (GPa) (Mg/m3) ( X 1 O-6/K) (kJhgK) (WImK)

Copper 117 8.86 16.6 0.38 398

Aluminum 70.3 2.68 23.8 0.96 138-237

Stainless steel 193 8.03 16.5 0.50 16.26 Super Invar 144.7 8.03 0.18 0.50 13.8 Fused silica 73 2.19 0.5 0.71 1.33 Ule fused silica 67.5 2.19 0.029 0.75 0.86 Zerodur 90.3 2.54 0.108 0.92 1.7 P75S/EP 103.3 1.72 -0.14 0.88 55.3 P100S/EP 143.3 1.80 -0.72 0.88 116 P120S/EP 170.9 1.83 -0.85 0.88 190 PlSOS/EP 186 1.83 -1.08 0.88 334 T300/EP 55.1 1.58 2.7 0.88 2.4 T50/EP 82.7 1.58 0.72 0.88 20.8 AS-4/EP 51.7 1.58 2.52 0.88 5.2 XN50A/EP 103.4 1.80 -0.072 0.88 48.4 XN70A/EP 144.7 1.83 -0.54 0.88 98.6 M55J/EP 103.4 1.66 -0.27 0.88 20.8 M60J/EP 120.6 1.69 -0.45 0.88 24.2 FT500/EP 82.7 1.77 0.36 0.88 41.5 FT/ 700/EP 137.8 1.80 -0.72 0.88 81.3 GY70/EP 110.2 1.66 -0.18 0.88 NA Boron/EP 89.6 2.05 5.4 0.88 NA S-GL/EP 26.9 2.02 9.7 0.71 0.34 KV49/EP 29.6 1.38 6.3 NA 1.04 KV-149 40.6 1.38 1.8 NA 1.04 NA = Not Available

Beryllium 289 1.83 11.5 1.88 179-207

Titanium 113.7 4.43 9 0.54 16.9-20.8


Table 47.5 Material selection criteria

Critical requirements

Mass

Dynamic stability

Thermal stability

Dynamic loads (Gs)

Temperature extremes

Hygrostability

High thermal conductivity

Low thermal conductivity

Aerospace application Equipment structure Instrument structure

-

RF Solar panel Bus Booms Stable Mirrors reflectors substrates structures structures

GFRP(1) GFRP GFRP GFRP (5) GFRP (6) Kevlar Aluminum Aluminum Beryllium Beryllium

GFRP (1) Kevlar

GFRP (1) Kevlar

GFRP

GFRP (1) Kevlar

GFRP (2) Kevlar

GFRP (2)

GFRP (3)

GFRP (4)

RF transmissibility Kevlar

cost GFRP Kevlar

GFRP (1) Kevlar Aluminum

GFRP (1) Kevlar



GFRP (2) Aluminum

GFRP (3) Aluminum

GFRP (4) Kevlar

Kevlar


M/M

GFRP Aluminum M/M

GFRP (5) Invar

GFRP Aluminum M/M

GFRP (2) Aluminum M/M



GFRP (4)

-

GFRP Aluminum

Beryllium M/M

GFRP Aluminum Beryllium M/M

GFRP Beryllium M/M

GFRP Beryllium M/M Aluminum

GFRP (2) Beryllium M/M Aluminum



GFRP (4)

-

GFRP Aluminum

GFRP (5) Beryllium Invar

GFRP (5) Invar Beryllium


GFRP (2) Invar Beryllium



GFRP (4)

-

GFRP Invar

Si/C, C/C

GFRP (6) Beryllium Si/C, C/C Aluminum

GFRP (6) Invar Beryllium Si/C, C/C

GFRP (6) Invar Beryllium Aluminum

GFRP (2x6) Si/C, C/C Beryllium Aluminum

GFRP (2)(6) Invar Beryllium Aluminum Si/C, C/C

GFRP (3x6) Invar Beryllium Aluminum Si/C, C/C )

GFRP (4x6)

-

Aluminum Invar

~

(1) Combinations of Kevlar and GFRP is used when mass and dynamic stability is important. (2) Cyanate resins have been shown to handle temperature extremes (T, high, no microcrackhg). (3) Pitch fibers (especially ultra, ultra high modulus have high thermal conductivity). (4) PAN fibers have low thermal conductivity. (5) Metals shown may be applicable if a small size structure. (6) GFRP has been used successfully as mirror substrates and for some submillimeter reflectors core and skin.


E

I.WOE4 4- '- ~ -jIijIlj*l 0.0024 0.00211 0.0032 O.WJ6 0 . W

TEMPERATURE ( 1 I K )

Fig. 47.2 Diffusivity against temperature for P75S/cyanate; P75S/epoxy system.

relative humidity and shows the significance between an epoxy and cyanate ester resin. Table 47.6 provides values for the swelling coefficient @) for typical GFRP materials used to fabricate aerospace structures, where p is calculated from:

or A L / L = PAM

Once /3 is known, the expected strain can be calculated for a given percent moisture content in the laminate.

The significance of Fig. 47.4 cannot be over- emphasized. Up until 1989 designers of aerospace structures had to utilize the work-around techniques 1 and 2 mentioned in Table 47.2 in order to handle the excessive expansion and the corresponding distortion associated with the hygroscopic nature of epoxy based GFRP material systems. Table 47.7 indicates work-around techniques utilized on various programs before cyanates were introduced. Cyanate ester resin systems now being used (Brand et al., 1992) in GFRP materials offer increased stability at reduced cost for aerospace structures, primarily due to the possible elimination of expensive moisture barriers (10% of manufacturing costs) and the even greater cost of facilities for dry- out processing.

p = (E&) (47.4)

I W

ERLWW RES4N r m 0 40

( worn I - E

080 P76S I ERLISBZ g (EwxIl

5 080

P

0 30

w

P765 I ERLlOsO URE ' J I CYUUTE I

I L

s z 0 10 020

OW

0 00

RELATIVE HUMIDITY I X ) 0 00 2 00 4 00 8 00 8 00 1000 TIME ( H O U R S ~ 7 I P )

Fig. 47.3 Percentage moisture change against time (h1Iz) for 0 . 5 m thick P75S/cyanate at 100% RH exposure.

Fig. 47.4 Percentage moisture pick-up against relative humidity (Yo). (a) ERL1999 epoxy resin; (b) P75S/ERL1962 epoxy; (c) P75/ERL1999 cyanate.


Table 47.6 Typical material CMEs for pseudoisotropic laminates

Material

T300/EP T50/EP P75S/EP P1 00s / EP P120S/EP P75S/CY

p (PPM/%M) M , @ 100% RH

400 215 162 103 86

105

1.13 1.13 1.03 1.03 1.03 0.32

Microcracking

Microcracking of a composite which causes dimensional instability should no longer be an issue if the proper material selections are made. A need to thermal cycle to gain dimensional stability (causing translaminar cracking or microcracking to progress to completion) was a requirement in the past, e.g. in Hubble Telescope tube components. This was due, primarily, to of the relative brittleness (non-toughened) of the resins and the unavail- ability of thin cure ply thickness prepregs. Also, at the time, the phenomenon was not fully understood.

Figure 47.5 illustrates the effect of composite microcracking. As the temperature is lowered and reaches the threshold of microcracking, as evidenced by an abrupt strain change, the curve changes to a new slope. Dimensional instability manifests itself in the following ways:

0 A hysteresis effect is produced in the structure.

0 Changes in CTE occur (CTE becomes more negative with increased thermal cycling).

0 Moisture response rates and moisture levels increase.

0 Irreversible expansion of the GFRP material.

Table 47.7 Prior work-around technique for addressing the hygroscopic nature of graphite/epoxy

Instrument Work-around technique Annotations (Ref.) (from Table 47.2 -

Hygroscopic nature)

Teal ruby 1 Stumm, 1979. Cryogenic application, indium/bismuth deposited metal-type moisture barrier - space application

Thematic Mapper

Mars Observer camera

2 Walrath, 1979. Two flight units operational, bake-out and environmental controls during assembly and test very well controlled and monitored - space application

Telkamp, 1990. Pre-flight unit test provides dimensional change data, compensated for by preset of optics - space application

2

HRMA Cylinders HEAO-B 1 Hertz, 1977. Aluminum foil-type moisture barrier; space application


I 2%

Fig. 47.5 Interpretation of common events occur- ring during measurements illustrating microcracking of pseudoisotropic GY-70/930 (0" /45" /90° / 135")%. 12mm/ply.

Today the following procedures can be used by designers to select appropriate fiber and resin systems to eliminate or minimize microcracking.

0 Establish the thermal range of the exposure environment.

0 Determine the stiffness, CTE, and weight requirements of the application.

0 Use a thinner cure ply thickness (CPT) prepreg, if high modulus materials are required (e.g. 0.0063 mm or less CPT) or a woven fabric.

0 Select a resin system compatible with the thermal and mechanical requirement (e.g. cyanate esters such as Fiberite 954-3 or YLA RS-3 do not microcrack when used with UHM pitch fibers (Amoco P75S) even at cryogenic temperatures).

0 Establish the optimal fiber orientation (e.g. (0"/45°/900/1350)s).

Of course, adequate testing of the material laminates before they are employed in the structure will ensure that correct material selections have been made.

Anisotropic behavior

The anisotropic behavior of GFRP offers a challenge in the design of aerospace structures. Some typical areas of concern are joints, springback, cutouts in cylinders and bowing of panels and assemblies.

Joints

Early in the 1970s, designers discovered that the through-the-thickness (translaminar) properties exhibited by GFRP reduced strength and increased CTE values for joints, i.e. the CTE in-plane may be 0.10X10-6/oC and 34X104/"C through the thickness. In the following years, however, improved mechanical/ thermal joint designs have evolved which optimize joint characteristics relative to thermal distortion/ cost/weight requirements. Figure 47.6 illustrates this evolution for both fixed and removable joints with the development path moving from A to D.

@ WAL OCUBLE CLIP @ MllY 6 TENON

Fig. 47.6 GFRP joint design evolution.

Springback

Springback occurs when composite laminates are formed into angles during processing. Springback, or angle closure, occurs to the extent of about 1.5" upon removal of the laminate from the tool (General Dynamics Corporation, 1985). Exposure to moisture con- ditions will open this angle (via moisture absorption) to return to, or even exceed, the original angle. Through-the-thickness expansion of the laminate is the source of this dimensional instability. Using flat laminates and butt bonding at right angles is the work- around technique. Very thin clips can be used at the joint.

Thermally, structures can be very sensitive to this problem as was discovered in the late 1970s in the testing of the GEMS Forward Mirror Support (Campbell et al., 1981). This problem has often been overlooked by designers over the years and needs to be considered whenever composites are utilized on dimensionally stable structures. Back-to-back angles or more simply mortise/tenon joints have been the work-around technique.


Flat panel bowing/warping

When laying up a flat panel, the typical fiber angle tolerance from the desired angle is O G o . This manufacturing tolerance is a major con- tributor to the bowing or warping of a panel. The thicker the laminate, the more a problem these variations are, in that substantial forces are required to make the panel flat. The laminate can straighten upon moisture absorption, and drying the laminate to its original cured state will bring back the prior laminate geom- etry.

The best solution for ehinat ing bowed or warped panels is to use a 'rotate and fold' laminate technique, specifically developed, to produce essentially balanced laminates. This technique is depicted in Fig. 47.7. In essence, ply lay-up tolerances are canceled by becom- ing symmetrical, with the resulting laminate properties truly isotropic throughout the laminate. Note that this process is only applicable to pseudoisotropic laminate configurations.

Cut-outs in curved surfaces (cylinders)

Some design approaches require holes in cylinders. The degree of distortion when a large hole is cut in a cylinder is surprisingly extensive and the stresses generated can force the cylinder to take a different shape (oval or hour-glass). Even the ends of a cylinder that has a large hole in it will not be flat or parallel to each other. The thicker the cylinder wall, the more difficult it is to correct this condition. Stiffening rings usually correct the problem on thin walled cylinders. Springback and cutouts, in formed parts, may reduce the benefit gained by a reduced number of piece parts if tight dimensional control is necessary. Parts may not fit together. Again, through-the-thickness expansion of the laminate is the source of this dimensional instability.

Fig. 47.7 The rotate and fold laminate construction technique.


Assembly bowing

By maximizing the use of flat stock and mak- ing dual skin, eggcrate core structures, very stable assemblies are possible. This brings about an additional concern with through-the- thickness expansion in that when an egg-crate core is used in a structure, severe bending is to be expected unless a specific design technique is employed. Since through-the-thickness expansion is 1000-times the basic in-plane expansion of the laminate, bending of the structural assembly occurs. Figure 47.8 illustrates the correct and incorrect slotting method. If done correctly the structure (Fig. 47.9) stays relatively straight through thermal cycling and moisture absorption.

INCORRECT SLOTTING METHOD

I 1 1 n n n n I

WET OR T R.T.

lnnnnl ORY OR T = R.T.

CORRECT SLOTTING METHOD

I U n U n

WET OR T R.T.

DRY OR T : R.T.

CROSSING RIB THICKNESS ADHESIVE

Fig. 47.8 Correct and incorrect slotting methods.

Fig. 47.9 FGS keel. The structure remains stable because the correct slotting method was employed. Mortise and tenon joints, used instead of clips, also promote stability.

47.2.2 CONSTRUCTION METHODS

Some important design options for aerospace equipment and instrument structure are:

0 truss (tubes) or cylinders (for telescopes); 0 monocoque skins or honeycomb sandwich

0 bonded, or bonded and bolted (for typical

0 molded unibody or flat laminate/bond

(for reflectors);

joints);

assemblies (for typical structures).

Truss compared with cylinder (telescopes)

Material availability, size, weight, interface, loads and stiffness requirements or considerations may drive the designer to a particular configuration for a particular application. Designers who favored tubular truss structures in the 1970s might not choose such an approach today. Design techniques are different, material choices are many, and manufacturing methods have changed. Today, for example, the Hubble Telescope could be designed as a faceted, dual-shell structure with a ribbed core. The faceted dual-shell (skins) could have water-jet cutouts with a

Hisforical perspective and progress 1015

latticework appearance. Then, again, the telescope might look the same but be fabricated from a non-microcracking, low moisture absorbing, thin-prepreg, cyanate ester prepreg system material.

Monocoque skins compared with honeycomb sandwich (reflectors)

Large GFRP honeycomb reflectors were successfully fabricated in the mid-1980s for the Advanced Communication Technology Satellite (ACTS) spacecraft reflector. At the same time that large membrane reflectors were made for a Direct Broadcast Satellite (DBS) spacecraft reflector. They both met the desired design requirements. The tie-down configurations and weight constraints, in part, dictated which approach was utilized. Here again the requirements of size, weight, interface, loads, stiffness and material availability will influence the design. Designer experience and preference will also determine what concept is used in the future.

Bonded joints compared with bonded and bolted joints

Experience with all bonded structures, cost considerations, and successful application (i.e. hardware on-orbit) are driving designers of aerospace structures away from fasteners alto- gether. The short cyclic load duration for spacecraft today are such that little, if any, benefit is gained from bolts or fasteners and their use should be minimized. Aircraft applications for equipment and instruments may be another story and each application needs to be reviewed thoroughly, assessing cost, weight, risk and so on.

Molded (unibody) compared with flat laminate bondedassemblies

The use of fewer parts may not be beneficial if they do not match up well at assembly. Also, if the cost of tooling and touch labor for

a molded part is excessive on a particular design, a molded part may not be better. Then again, flat laminate construction may mean that more parts have to be made, inspected and handled, and if this cannot be done efficiently by the manufacturer, then it may not be the best approach. Some manufacturers can handle flat laminate construction very efficiently, if they have experience of assess- ment of cost, weight, risk. Many flat laminate 100% bonded assemblies for aerospace equipment and instrument structure have been fabricated and successfully flown, both on aircraft and spacecraft. Apparently the previ- ous high risk, thought to exist with this approach, has not been proven true in practice. Here again, designer experience and preference will determine the course to fol- low.

47.2.3 TOOLING DEVELOPMENT

The number of cures, the size of the part and facilities available influence the choice of tooling to such a degree that each product fabricated must be thought-out thoroughly to select the proper tooling.

An advantage of flat laminate construction is that it eliminates a need for fabricating production molds. A designer may want to look at this approach first in order to cut costs, if experience and fabrication techniques support this method of fabrication.

The high cost of production molds can be significantly reduced through the use of alter- nate techniques. For example, for cylinders, thin-walled rolled and welded aluminum molds create dramatic savings as their light weight permits envelope bagging which eliminates the need to withstand autoclave pressures.

Monolithic (bulk graphite) molds are expensive and heavy (storage and heat-up rate concerns), but their CTE match to GFRP or Kevlar/epoxy is a great advantage. Greater accuracy and better replication of parts is possible.

1016 Aerospace equipment and insfvument structure

The tooling concept and materials used to fabricate the tooling can be as important to the success of the hardware as the hardware design itself. The aerospace structural designer must stay in touch with technological developments in tooling and tooling materials.

ment and their application to aerospace equipment and instrument structures (Tables 47.8 and 47.9). The photograph of the Hubble Space Telescope (Fig. 47.10) only shows the metering truss structure for the telescope; substantially more GRFP was used in the fabrication of the Ford Plane Assembly (FPA) and the many

47.2.4 MAJOR MILESTONES other components. The milestones cited repre- sent major advances and point the way for

Milestones in the use of composite materials have been in the areas of materials develop-

continued use of composites for a wide variety of aerospace structures.

Table 47.8 Milestones for GFRP materials

Material Achievement

Thin prepregs Reduced weight of hardware components Minimized microcracking Supported mirror technology Increased thermal stability of laminates

Toughened epoxies

High modulus PAN fibers

Increase impact resistance of laminate Increase bond strength (interlaminear) of joints Increase compressions strength of laminate

Reduce hardware weight Allow high stiffness, thermal stable, cyanate laminate Allow high strength application

Ultra-ultra high modulus pitch fibers Increase thermal conductivity Minimum weight, stiff structures Increased EM1 capability

Cyanate ester resins Reduced moisture levels Reduced strain response to moisture level Reduced microcracking Increased use temperature (high TE)

Table 47.9 Milestones for GFRP applications

Application Date Achievement

Hubble telescope

Nimbus-G

Shuttle RMAB

1970-1980

1970-1980

1970-1980

Design and fabrication of a very large (2.4 m diameter x 5.2 m length) and thermally stable structure

No fasteners - membrane shell 100% bonded microwave reflector; Qual unit qual-tested twice. Flown in 1990 as TOPEX.

Reusable structure. Multiple shuttle missions.

Current applications 1017

-

Fig. 47.11 ACTS reflector.

Fig. 47.10 Hubble space telescope metering truss assembly.

47.3 CURRENT APPLICATIONS

47.3.1 ACTS REFLECTORS

The Advanced Communications Technology Satellite (ACTS) reflectors were manufactured by Composite Optics, Inc. (COI), for General Electric/Astro-Space Division under contract with NASA Lewis Research Center (Fig. 47.11).

These large reflectors (2.2 and 3.3 m aper- ture) had measured surface RMS accuracies of approximately 0.071 mm over their entire surface after thermal cycling. These reflectors were launched into orbit in September 1993 and are functioning well. Their GFRP skins (P75S/ERL1962) and Kevlar honeycomb core (Kevlar 49/934) resulted in a 3.44 kg/m2 weight for the completed reflector (Rule, 1989).

47.3.2 HRDI OPTICAL BENCH (UARS)

This 1.24 x 1.8 x 0.11 m thick, 49.9 kg optical bench for NASA's Upper Atmosphere Research Satellite (UARS) mission was nearly 100% bonded (Fig. 47.12). The only fasteners were anti-peel fasteners at the four comers where graphite/epoxy (T50/ERL1962) panel sides terminate at titanium fittings. The bench carried 3.2 times its own mass and had a CTE of 0.36 X 10-6/oC (Dodson, 1989).

47.3.3 WCS(SOH0)

The Ultraviolet Coronagraph Spectrometer (WCS) for the European Space Agency's Solar and Helospheric Observation (WHO) spacecraft (1995 launch) is 100% bonded graphite/epoxy (P75S/ERL 1962) (Fig. 47.13). This truss structure is the first of its kind in that the basic truss panels are bonded assemblies that have an I beam cross section. Eighty


Fig. 47.12 HRDI bench (UARS).

ophcalstabilitynquktnmts(CTE -0.09 X lW/OC) had to be satisfied along with strength and frequency (70 Hz) requirements. The structure will hold 90.7 kg of equipment and only weighs 21.8 kg itself (Kilpatrick et al., 1990). Most of the surfaces on the structure were clearable and inspectable (minimal closed areas).

47.3.4 COSTAR (HUBBLE TELESCOPE CORRECTIVE OrrrCS)

The Corrective Optics Space Telescope Axial Replacement (COSTAR) was developed by Ball Corporation by NASA/GSFC. This structure (Fig. 47.14) supports a Deployable Optical Bench (DOB) that serves as the corrective optic for the Hubble Space Telescope. Built by Hercules Aerospace Company, this composite structure has to have a low CTE (<0.15) and

:ig. 47.13 W Coronagraph Spectrometer (UVCS) telescope structure assembly (SOHO).

Predictions 1019

Fig. 47.14 Corrective Optics Space Telescope Axial Replacement (COSTAR).

high resonant frequency (100 Hz for the DOB). The long-term stability effects due to moisture absorption (ground) and desorption (space) were addressed by optimizing the composite laminate orientation. The long-term CTE concerns were more important in meeting line-of-sight requirements (>15 arc-sec) than was a need to optimize CTE. A removable cover and mounting of aluminum electronic box required very unique design solutions (Neam and Gerber, 1992).

47.4 PREDICTIONS

The dominance of GFRP as the preferred choice composite material in the fabrication of aerospace equipment and instrument structures will continue into the next century because of a number of reasons:

0 new materials; 0 new manufacturing processes; 0 new applications; 0 economics.

This, of course, does not mean other composite materials will not be used, but their use will be in similar proportion as is the current practice.

47.4.1 NEW MATERIALS

GFRP is a constant evolution of material variations. Fiber makers continue to change their fiber manufacturing processes to create a new fiber to satisfy the needs of the user. Properties like fiber strength, modulus, and thermal/elec- trical conductivity are continuing to be adjusted for various application requirements. The demand for properties changes also are requested of the resin manufacturers and/or


prepreg suppliers. Cyanate ester resin formula- tions are in their infancy, current GFRP cyanates have levels of moisture absorption less than one-third the GFRP/epoxies. Soon moisture absorption for GFRP/cyanate may be one-tenth that of GFRP epoxies. In the future, we may have thermoset resins that will have no moisture absorption and their operating temperatures may be over 240°C (464°F).

New manufacturing processes

It is hard to imagine new manufacturing methods for processing GFRP that are not in use today, but the demand for less expensive materials, and hardware may bring developments. One method currently being developed by Composite Optics, Inc. in San Diego, CA, called SNAPSAF is predicted to reduce the manufacturing cost to less than half current levels.

New applications

New applications stimulate the need for new materials, and new materials, in turn, create new applications. Future applications for GFRP are mirrors of all sizes, surface figure and accuracies in the IR and visible range, large antenna reflectors (9.1-27.4 m diameter), electronic chassis and cardguides, plated RF components (MUX Cavities, Diplexers), phased arrays, to name a few. The list is on- going but it is hard to tell whether new GFRP materials bring about these applications or the applications bring about the material. For instance, Amoco’s KllOOX fiber is being exper- imented with for electronic chassis, but the original demand was for high thermal conduc- tive material for heat sinks (thermal straps). Ln another example, thin prepregs were developed to make lighter skins for honeycomb panel structures, but in the future mirrors for visible range optics will also mostly use these thinner prepregs.

47.4.2 ECONOMICS

A 1993 conference held in Logan, UT, spon- sored by the AIAA and Utah State University, had a basic theme for small satellite structures of ‘cheaper, faster, better’. The ’better’ was in some cases interpreted as lighter for better performance.

One of the primary messages to the indus- try from that conference was that economics dictates that satellites become smaller and lighter so that the ’on-orbit’ cost for a particular satellite is much less, can be launched by a small, low-cost launch vehicle (e.g. OSC Pegasus or Lockheed LLV1). GFRP is currently the most logical choice to satisfy the requirement of cheaper, faster, and better.

Manufacturing processes are developed that will allow lightweight GFRP to be utilized for bus structures, solar panel substrates, reflectors and instruments. Weight reductions of 5040% are possible with GFRP. Package size may be the only limitation to allowing three or four satellites to be launched rather than one or two. Creative packaging and miniaturization will solve this problem and launch booster cost will then be minimal (25-35% of the current cost level of a single satellite).

This same idea (cheaper, faster, better) can also apply to aircraft equipment and instruments, namely, airborne avionics, tracking and targeting instruments, and phased arrays. Today, a military or commercial aircraft maker may not be able to sell a new aircraft but he could redesign the avionics (lightweight) to improve the performance of that aircraft (including helicopters).

47.5 CONCLUSIONS

The technology of composite materials, especially GFRP, is constantly evolving to allow increased aerospace application. This increased use of GFRP is for both current type applications as well as genuinely new applications where GFRP was not previously used. The future for composites, in general, looks

References 1021

Krumweide, G.C., Brand, R.A. 1991. Attacking Dimensional Instability Problems in Graphite/Epoxy Structures. Composite Design, Manufacture and Application. ICCM/8. Honolulu. July 15-19.

Krumweide, G.C. 1977. Development of a graphite/ Epoxy Reflector: A design-to-cost project. SAMPE Quarterly. 8(3).

Krumweide, G.C., Chamberlin, D.N., Rule, J.E. 1988. Adaptation and innovation in high-modulus graphite/epoxy composite design: Notes on Recent Developments. SPIE 0 - E LASIE 1988. Los Angeles.

Krumweide, Derby, E.A., Chamberlin, D.N. 1989. The performance of effective moisture barriers for graphite/epoxy instrument structures. SAMPE. Atlantic City, NJ.

Krumweide, G.C., Hoste, J.H. and Staats, J.R. 1979. Structural Development of the Thematic Mapper Optical Metering Structure. The Enigma of the Eighties: Environment, Economics, Energy. S A M P E J. 24(2), 1343-1355.

Levy, D.J. and Arnold, C.R. 1984. Metal Moisture Barriers for Composites. 29th Nat. SAMPE Symp. April 3-5.

Rule, J.E. 1989. Thermal Stability and Surface Accuracy Considerations for Space-based Single-and-Dual Shell Antenna Reflectors. E S A ESTEC. Noordwijk, Netherlands.

Stumm, J.E., Pynchon, G.E. Krumweide, G.C. 1981. Graphite/Epoxy Materials Characteristics and Design Techniques for Airborne Instrument Applications. 309. Airborne Reconnaissance. V, SPIE.

Stumm, J.E., Pynchon, G.E., Pepi, J.W. and Bovenzi, F.G. 1979. Low Temperature/High Stability Applications of Composites. The Teal Ruby Experiment. Conf. Advanced Composites. El Segundo, CA.

Telkamp, A.R., and Derby, E.A. 1990. Design Considerations for Composite Materials used in the Mars Observer Camera. Advances in Optical Structure Systems. 1303. Orlando, FL: SPIE.

Walrath, D.E. and Adams, D.F. 1979. Moisture Absorption Analysis of the Thematic Mapper Graphite/Epoxy Composite Structure. Modern Developments in Composite Materials and Structure. ASME Winter Meeting.

promising because of the need for aerospace equipment and instrument structures to be lighter which increases performance and saves costs. Manufacturing methods are being developed that allow GFRP hardware, in particular, to be processed less expensively with reduced cycle times. It is not hard to imagine GFRP material replacing most metallic material applications for aerospace equipment and instrument structure once 'old paradigms' are given up and awareness of new GFRP material technology is the norm.

REFERENCES

Brand, R.A., Derby, EA., Chamberlin, D.N. 1992. Evaluation of high-modulus pitch/cyanate material systems for dimensionally stable structures. Design of Optical Instruments, Orlando, FL: SPIE.

Campbell, M.P., Hoste, J.H., Kedward, K.T., Krumweide, G.C. 1981. Designing Composite Structures for Thermally Stable Applications. Fifth DOD/NASA. Conf. Fibrous Composites in Structural Design. New Orleans, LA. January 1981.

Dodson, DJ., Rule, J.E. 1989. Thermal Stability Considerations for Space Flight Optic Benches. TomorrowS Materials Today. Vol. 34.

Dunbar, D.R., Robertson, A.R., Kenison, R. 1978. Graphite/Epoxy Booms for the Space Shuttle Remote Manipulation. ICCM 11, Toronto, Canada.

General Dynamics Corporation (GDC Manual). 1985. Design for Cost and Quality Manual.

Herrick, J.W. Multi-Directional Advance Composites for Improved Damage Tolerance. Composites in Manufacturing 3. Anaheim. January 10-12,1984.

Hertz, J. Moisture Effects on Spacecraft Structures. 1977. The Enigma of the Eighties Environment: Environment, Economics, Energy. SAMPE, 24(2).

Kilpatrick, M.C., Girard, J.D., Dodson, K.J. 1990. Design of a Precise and Stable Composite Telescope Structure for the Ultraviolet Coronagraph Spectrometer (UVCS). Advances in Optical Structure Systems. Vol. 1303. Orlando, FL: SPIE.

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