Composite Space Component 29jun2015

8/19/2019 Composite Space Component 29jun2015

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this such temperature condition, thermo-elastic properties of the composite will vary with the change oftemperature [3]. The research of fiber reinforced composites is still focused on mechanical properties,

while only a few investigations on their thermal expansion behavior and dimensional stability [4].

For majority of spacecraft components CFRP composites are widely accepted because they

provide highest trade-off metrics based on weight and performance [2]. The LEO space environment

constituents consist of high vacuum, UV radiation and thermal cycles [5]. Thermal cycling is one of

environmental effects of space that which is known to induce environmental degradation on satellite in

LEO, (between 100 and 1500 km above the Earth’s surface) as it passes in and out of the earth’s shadow.

As a result tThe exterior surface is exposed to long-term periodic sharp temperature changes [6]. In

general, temperature of space components varies between -101°C to 125°C [7]. Near to zero coefficient of

thermal expansion is desired for dimensional stability of space structures, especially those employed for

image sensing. For fabrication of ultra-stable space components, dimensional stability of CFRP is

concerned because of thermal loadingimportant. The dimensional stabilityis is directly related toclosely

connected with morphologycomposition, structure and chemistry of fiber/resin of the composite used [8].

However, the limited understanding of composite structure’s dimensional stability remains a serious

problem against the selection of CFRP composites for this class of space structures. Carbon/epoxy

composites are example of high-end materials, known for their low density and considerable stiffness and

strength to weight ratio. Dimensionally stability of CFRP is influenced by factors like fiber and resin

used, fiber volume fraction and the number and orientation of individual layers. HoweverAmongst these,

the most influential factor is the layup sequence [9]. As the composite analysis methods continue to

improve, designers are learning to utilize composites consisting of laminas (plies) of oriented fibers to

obtain unique material properties. Typically, this is accomplished by orienting unidirectional laminae at

various angles to obtain a laminate with the desired properties.

In engineering applications, structural parts made of carbon/epoxy materials frequently work at

high temperatures which leads to thermal aging. Recently several experimental results are reported on the

influence of thermal loading on composite strength and aging properties [9]. However,

there is still a lack of FEM software resultsbased studies, regarding the effect of thermal loading on

dimensional stability. Composite materials can be tailored in order to obtain a minimal thermal

deformation. Furthermore, composites, with their high specific stiffness (E/ρ) and thermal stability ( K /α),

can avoid a weight penalty (E is modulus of elasticity, ρ is density, K is thermal conductivity and α is

coefficient of thermal expansion) [10]; E is modulus of elasticity, ρ is density, K is thermal conductivity

and α is coefficient of thermal expansion. Therefore, composites have become more employed for

satellite structures and optical space components.

Dimensional and alignment errors among constituting subcomponents can cause serious degradation

of space components’ in functional performance. There has been some study on the degradation due to

dimensional instability of space components [10]. In order to evaluate such dimensional instability, an

approach based on finite element analysis program (such as, ANSYS® APDL, ) may be employed.

Thermal expansion coefficient is an important property of composite material which affects deformations

and thermally induced stresses in a composite component. The matrix material typically exhibits

significantly different coefficient of thermal expansion than fibers. An accurate determination of thermal

expansion coefficient taking into account combined effect of matrix and fiber is necessary. In laminated

composites, this coefficient depends largely on the orientation of fibers, the fiber fraction, type of resin

and reinforcement.

The objective of this work is to select suitable composite material in order to design a typical

dimensionally stable bench table for space application. A bench table is a support structure which houses

antenna and allied electronic and mechanical components to obtain digital images from a space satellite.

The influence of thermal environment on dimensional stability of carbon/epoxy composite based bench

table is investigated by varying following composite material propertiesparameters: combination of

fiber/matrix material, fiber orientation and layup sequence. The most suitable composition of composite

material for bench table is established through this study.


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2. Theoretical Basis for Analysis

For isotropic bodies, the coefficient of thermal expansion is the same in all directions. For composite

materials, the coefficient of thermal expansion, like other properties, changes with direction. Because of

the structure of the composite, a unidirectional composite shows different CTE in the longitudinal and

transverse directions. Thus the unidirectional composites are orthotropic with the axes 1 (longitudinal), 2

(transverse) and 3 (thickness) as the axes of symmetry. Because of the random fiber distribution in the

cross section, material behaviour in directions (2 and, 3) is nearly identical. Therefore a unidirectional

composite or ply can be considered to be transversely isotropic that is, it is isotropic in 2-3 plane [11].

Unidirectional composites have two principal coefficients of thermal expansion, the longitudinal

coefficient of thermal expansion αL and the transverse coefficient of thermal expansion αT. Schapery [12]has derived the following expression for the longitudinal and transverse coefficient of thermal expansion

in terms of fiber and matrix material properties.

αL = αf Ef Vf αEV .................................................................................................................... (1)α

T = ( 1 µ

f )α

f Vf ( 1 µ

)α

V α

Lµ

LT ....................................................................................... (2)

where

αf and α are coefficient of thermal expansion of fibers and matrixEL is the elastic modulus of the composite in the longitudinal directionEf and E are elastic modulus of fibers and matrixVf and V are volume fraction of fibers and matrixµLT is the major passion ratio of the composite

In this work the values of αL and αT for composite materials AS/H3501 and T300/N5208 aredirectly available obtained from bookpublished data [11].

Procedure to calculate laminate stress and strain in composite materials (Fig.1) is explained by

Agarwal, Broutman and Chandrashekhara et al. [11].

In ut

Laminate construction Laminate elastic constants Applied loads

Laminate stiffness matrices [Q]

[Q] For different orientations

Laminate stiffness matrices [A], [B] and [D]

Mid- plane strains (ε) and plate curvatures (k)

Global laminae strains (εX, εY, γXY)Local laminae strains (εL, εT, γLT)


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Fig.1: Flow chart for laminate stress analysis [11].

In unidirectional fiber form, longitudinal direction is fiber direction and transverse direction is

perpendicular to fiber direction. Subscript for fiber is denoted by ‘f’ and for matrix is denoted by ‘m’.

For transversely isotropic material values of EL, ET, GLT, µLT are generally available for standardmaterials. In order to calculate laminate stiffness matrix [Q], following additional relations are used [13].

[Q] =

−µµµ−µµ 0

µ−µµ−µµ 0

0 0 GLT

............................................................................................................ (3)

Orthotropic material with tTransversely isotropic materials are characterized by are having following

relations.: [?].

E = E3 = ET ............................................................................................................................................ (4)G = G3 = GLT ....................................................................................................................................... (5)µ = µ3 = µLT ....................................................................................................................................... (6)µ3 = µf Vf µ 1 Vf +µ−µ

−µ +µµ

............................................................................................ (7)G3 = +µ............................................................................................................................................ (8)where

EL Young’s modulus in longitudinal direction ET Young’s modulus in transverse direction GLT

Shear modulus

µLT Major poisson’s ratioµTL Minor poisson’s ratioµ3 Poisson ratio in 2-3 planeG3 Shear modulus in the 2-3 planeVf Volume fraction for fiberµf Poisson ratio for fiberµ Poisson ratio for matrixE Young’s modulus for matrixUsing After formulating [Q] then formulation the stresses and strains are obtained by employing

procedure shown in Fig. 1. The stresses and strains are evaluated by appropriate failure theory to

determine design acceptability. is obtained.

The present work employs Tsai-Wu failure theory to check the sufficiency of composite

component design. This theory has more general applicability than other failure theories such as,Maximum-stress failure theory, Maximum-strain failure theory, Tsai-Hill failure theory, because it

Local laminae stresses (L, σT, τLT)

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distinguishes between the compressive and tensile strengths of a lamina [14]. This theory provides a

single criterion to predict the failure of lamina.

The Tsai-Wu failure criterion states [15]: under plane stress condition the failure will occur when the

following inequality is not satisfied.

0 < ξ


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selection ofng appropriate composite material to achieve dimensional stability and also to protectionagainstsatisfy the Tsai-Wu failure criterion.

Fig.2: Schematic representation of bench table on satellite structure

In bench table, the cylinder is mounted on sandwich structure using bolted connection. The

sandwich structure consists of metallic honeycomb structure of aluminium7075-T6 material which is

covered on both flat surfaces using carbon-epoxy composite materials. Dimensions for sandwich structure

base are given in Table 1.

Table 1: Dimension for sandwich structure base.

Dimension for base (mm)

Length 830

Breadth 814

Height of honeycomb structure 98Thickness of honeycomb foil 0.07

Height of Top and Bottom faceplate 2.75

Fig.3: Model for sandwich structure base

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Formatted: Left


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Carbon-epoxy is used as face plate to provide better strength for withstanding bending stresses.The Ccylinder part of is used in bench table acts as support structure on which antenna is to be mounted.

Cylinder is made up ofconstructed with Carbon-epoxy material. Dimensions for cylinder are given in

Table 2.

Table 2: Dimension for cylinder

Dimension for cylinder (mm)

Thickness 2

Height 1000

Cylinder inner diameter 698

Cylinder outer diameter 700

Flange inner diameter 700

Flange outer diameter 800

circular face blend radius 5

Fig.4: Cylinder

Twelve holes of 10mm radius at angular pitch of 30 degreess are drilledprovided in the base plate

and flange of cylinder for their assembly using threaded bolts. Both the components are joined with the

help of M10 size stainless steel (SS) bolt and nut. Cylinder and base are modelled in UG-NX software

version 8.5.

3.1 Material selection:

Different types of CFRP materials are availablemay be used for design of cylinder and base, such

as, T300/N5208, AS/H3501, AS4/APC2, IM6/epoxy and T300/Fiberite 934. In this work AS/H3501 and

T300/N5208 are selected as possible design options as since both are having very low CTE. Combination

of fiber and matrix leads to formation of lamina. Differences in the CTE in the longitudinal and transverse

directions indicate that laminate CTEs are strong functions of layup. Material Properties for carbon/epoxy

(AS/H3501) and (T300/N5208) are as follows [11].

Table 3: Material properties for carbon/epoxy (AS/H3501)

Elastic constants Strength

V 0.66 σLU (MPa) 1447ρ (Kg/3) 1600 σ′LU (MPa) 1447

EL (GPa) 138

σTU (MPa) 51.7

ET (GPa) 8.96 σ′TU (MPa) 206GLT (GPa) 7.10 τLTU (MPa) 93

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µLT 0.3 Thermal expansion coefficients (10−6/˚C) µ3 0.3 αL -0.3 G3 (GPa) 3.4461 αT 28.1

Table 4: Material properties for carbon/epoxy (AS/H3501)

Elastic constants Strength

V 0.7 σLU (MPa) 1500ρ (Kg/3) 1600 σ′LU (MPa) 1500EL (GPa) 181 σTU (MPa) 40ET (GPa) 10.30 σ′TU (MPa) 246

GLT (GPa) 7.17

τLTU (MPa) 68

µLT 0.28 Thermal expansion coefficients (10−6/˚C) µ3 0.28 αL 0.02 G3 (GPa) 4.0234 αT 22.5 Material properties of aluminium 7075-T6 and structural stainless steel (SS) are as follows:

Table 5: Material properties of aluminium 7075-T6

Aluminium 7075-T6 Values

Density (kg/m3) 2810Young’s modulus (GPa) 71.7

Poisson’s ratio 0.33

Coefficient of thermal expansion (10−6/˚C) 23.6Yield strength (MPa) 503

Table 6: Material properties of structural stainless steel

Stainless Steel (SS) Values

Density (kg/m3) 7800Young’s modulus (GPa) 200

Poisson’s ratio 0.3

Coefficient of thermal expansion (10−6/˚C) 12Yield strength (MPa) 250

3.2 Finite Element Discretization:

The bench table has been discretized using 2D shell elements SHELL 181 of ANSYS ®. Quad

elements are generated by mapped meshing using HYPERMESH v12.0 software. Washer split or circular

partition is made around all the holes. Element size used for base and cylinder is 20 mm and 10 mm

respectively. The FE mesh is checked and cleared for element connectivity, duplicates, jacobian, warpage,

aspect and skewness.



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Fig.5: Discretization details for Base.

Base and cylinder are aligned and assembled. Surface to surface contact has been established

between the two components using contact elements. The bench assembly is bolted using 12 numbers of

M10 bolts which are modelled in FE using 1D bar elements, BEAM 188. Assembly is show below in

figFig.7.

Fig.6: Discretization details for Cylinder.


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Fig.7: FE model for assembly of bench table structure.

3.3 FE Analysis:

ANSYS® 15.0 is used to conduct the FE analysis of the bench table. The number/thickness of

layers considered for composite material was 8/0.125 mm for cylinder and 8/0.2 mm for base

respectively. Before carrying out the actual analysis, preliminary evaluation of 7 layup sequences (Table

7) were evaluated forwas carried out with the bench table subjected to thermal loading of 100˚C, -100˚C

as load cases. Based on the Tsai-Wu criterion, two of these layup sequences were shortlisted for design of

bench table (Table ?). for actual load cases and analysis which

is shown in results section.

Table 7: Different combinations of layup sequences studied

Combination no. Layup sequence for cylinder Layup sequence for base N1 [90/180/135/45]s [0/90/45/-45]s

N2 [150/120/60/30]s [60/30/-30/-60]s

N3 [90/90/150/30]s [0/0/60/-60]s

N4 [90/90/165/15]s [0/0/75/-75]s

N5 [90/120/60/180]s [0/30/-30/90]s

N6 [90/90/180/180]s [60/30/-30/-60]s

N7 [90/90/90/90]s [0/0/0/0]s

For aluminium and SS materials, isotropic material model was employed.

3.4 Boundary conditionThermal Loadings, Element coordinate system and Boundary conditions:

Cylinder and base are coupled to each other using beam elements at the centre of 12 bolt holes.

For coupling purpose, master nodes are created at top and bottom surfaces of assembly (Fig.8). Master

node of bottom of each beam is constrained in all degree of freedoms.






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For the bench table components in space environment, the temperature variation is in the range of

-101˚C to 125˚C. The tThermal loads are induced due to thermal gradient and/or the difference in CTE of

different layers. Reference temperature for thermal loading is the temperature at which assembly of

satellite is carried out and at this temperature zero thermal strains exist. The reference temperature

considered in this study is 25 deg. C.

When the satellite is launched in space, temperatures vary in the range of -101 ˚C to 125˚C.

Fig.9?: Thermal Lload cases.

The temperature states corresponding to each load cases is as shown in Table 8.

Table 8: Thermal load cases

Load cases Temperature (˚C) Case:1 100

Case:2 -126

Case:3 Half 100 & half -126

Element Coordinate system:

For correctness obtaining correctof FE results, all composite elements should have same type of

element coordinate system and the flow of element coordinate system should be smooth. X-axis, Y-axis

and Z-axis, is in black, green and blue colour respectively for element coordinate system. The alignment

of element coordinate systems for cylinder and face plates is shown below in figFigs.11 ? and fig.12 ?

respectively. X-axis, Y-axis and Z-axis, axes are represented is in black, green and blue colour s

respectively in these figures.for element coordinate system.

Formatted: Font: (Default) Times New Roman, 1



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Fig.10?: Element coordinate system

Fig.11?: Cylinder element coordinate system for surface layer .

Fig.12?: Face plates element coordinate system for surface layer .

Structural boundary conditions:


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Cylinder and base are coupled to each other using beam elements at the centre of 12 bolt holes.For coupling purpose, master nodes are created at top and bottom surfaces of assembly (Fig.?). Master

node of bottom of each beam is constrained in all degree of freedoms.

Fig.?: Structural boundary condition for bench table assembly.

Surface to surface contact is given provided between flange of cylinder and top face of sandwich

structure (Fig.13?). Both the surfaces are allowed to make contact or separate but these cannot penetrate

each other.

Fig.13?: Target and Contact nodes showing surface to surface contact.

FE discretization for components of base Ssandwich structure, i.e., base consist of honeycomb

structure, faceplates and bolts as is shown in Figs. 14, 15 and 16. Discretization for entire assembly is

shown in Fig. 17 shows entire assembly.

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Fig.14: FE discretization for Hhoneycomb structure

Fig.15: FE discretization for Ttop and bottom plate of honeycomb structure

Fig.16: FE discretization for Bbeams acting as bolts with displacement boundary condition (constraint at

bottom.)


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Fig.17: FE discretization for entire Bbench table after assembly of all components

4. Results and discussion

AS/H3501 has carbon fiber with negative CTE along fiber direction while T300/N5208 has

positive CTE in fiber direction. Response of the structure to space conditions was studied for two layup

sequences (Table ?). Minimum deformation of the structure and safety in Tsai-Wu failure criteria are the

basis for the evaluation of material/layup sequence. Out of seven different combinations only first and

second combinations are safe against Tsai-Wu criteria for AS/H3501 material which is further considered

for analysis. And then, the number/thickness of layers considered for composite material was 8/0.25 mm

for cylinder and 10/0.275 mm for base respectively . Different

combinations of layup sequence were studied for real thermal load cases are as shown in Table 9.

Table 9: Layup sequence considered for design of bench table.

Combination no. Layup sequence for cylinder Layup sequence for base

N1 [90/180/135/45]s [0/90/45/45/-45]s

N2 [150/120/60/30]s [60/-60/0/30/-30]s

Out of two material types T300/N5208 is eliminated due to less dimensional stability (Table ??) and

relatively less strength properties as compared to AS/H3501.

Table ??: Comparison of dimensional stability between T300.N2508 and AS/H3501 materials

The results for Maximum value of Tsai-Wu criterion is as shown in Table

9?.

Table 9?: Values of Tsai-Wu criterion a

Orientations Maximum value of Tsai-Wu Criteria (ξ)

Notations AS/H3501

100˚C -126˚C Both Half

N1 0.27 0.85 0.89

N2 0.25 0.92 0.94








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Table 13 ? shows Both N1 and N2 are safe for all three load cases. Factor of safety for N1 and N2nearly equal to 1.1 as required recommended for space application [?]. Maximum


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The overall least values for deformation are obtained for layup sequence N2. Maximumdeformation with N2 orientation at 100˚C occurs in honeycomb core and at -126˚C occurs in bottom face

plate. In Bench table, total deformation is very less and symmetric along axis of symmetry which is along

the length of the cylinder. This gives a dimensionally stable bench table. Figures 19(A), 19(B) and 19(C)

show total maximum deformation.

Fig.19 (A): Deformation with N2 orientations at 100˚C

Fig.19 (B): Deformation with N2 orientations at -126˚C


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Fig.19 (C): Deformation with N2 orientations both half (-126˚C and 100˚C)

Stresses in aluminium and steel materials are checked for failure as per Von Mises criteria. Thesechecks for failure for N2 are shown in Table 15.

Table 11: Von Mises stress

Aluminium 7075-T6 Structural steel (SS)

Von Mises stress for

honeycomb

Allowable (503)

Von Mises stress for bolts

Allowable (250)

100˚C -126˚C Both

half100˚C -126˚C

Both

half

381.8 479.7 475 240 240 240

Von Mises stresses in metallic materials are maximum at in the regions corresponding to the

boltswasher split area.. However, Iin actual assembly, the real stresses would be much less less as

compared to the simulation results due to the presence of washers which will distribute the stresses over a

larger surface. Figures 20(A), 20(B) and 20(C) show Von Mises stress.

Fig.20 (A): Honeycomb Von Mises stress at 100˚C


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Fig.20 (B): Bolts Von Mises stress at 100˚C

Fig.20 (C): Honeycomb Von Mises stress at -126˚C

The stresses in composite and metallic materials are within acceptable limit. Hence the support bench design with AS/H3501 as composite material and layup sequence N2 is selected as the best choice

against both requirements of dimensional stability and strength.

5. Conclusions

The dimensional stability for space borne bench table structure of composite material has been

successfully established using finite element based study. The choice of material and selection of layup

sequence is observed to play an important role in achieving the dimensional stability as well as in meeting

the Tsai-Wu strength criterion for the components of structure. Composite material with positive and

negative coefficient of thermal expansion for fiber and matrix materials provided the optimum design .

The layup sequence is found to have significant influence on the deformations of the bench table .

Bibliography

[1] Chang HH and Tarn JQ, International Journal of Solids and Structures 2007, 44, 1409-1422.

[2] Mini KM, Lakshmanan M, Mathew L, and Kaimal G, Steel and Composite Structures 2012, 12, 117-


20/20

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127.

[3] Chen CS, Tsai TC, Chen WR and Wei CL, Steel and Composite Structures 2013, 15, 57-79.

[4] Hui SM, Hui WG, Qin CG and Shu YW, Transactions of Nonferrous Metals Society of China 2009,

20, 47-53.

[5] Han JH and Kim CG, Composites structures 2004, 72, 218-226.

[6] Park SY, Choi HS, Choi WJ and Kwon H, Composites: Part B 2011, 43, 726-738.

[7] Kim RY, Crasto AS and Schoneppner GA, Composites Science and Technology 2000, 60, 2601-

6082.

[8] Abusafieh A, Federico D, Connel S, Cohen EJ and Willis PB, Composite structures 2002.

[9] Mlyniec A, Korta J, Kudelski R and Uhl T, Composite Structures 2014, 118, 208-216.

[10] Yoon JS, Kim HI, Han JH and Yang HS, Journal of Aerospace Engineering 2012.

[11] Agarwal BD, Broutman LJ and Chandrashekhara K, Analysis and Performance of Fiber

Composites, 3rd ed., Wiley: U.S, 2006.

[12] Schapery RA, Journal of Composite Materials 1963, 2, 280-404.

[13] Peters ST, Handbook of Composites, Chapman and Hall: London, 1988.

[14] ANSYS APDL, Tsai-Wu failure criteria 2015.

[15]

[16]

[17]

Barbero EJ, Finite Element Analysis of Composite Materials, 2nd ed., CRC PRESS, Taylor and

Francis Group: U.S, 2008.

VELMURUGAN PR , “Composite Materials,” IIT Madras, Madras, 2012.

K. K. Rao, K. J. Rao, A. G. Sarwade and M. S. Chandra, International Journal of Engineering

Research and Applications 2012, 2, 365-374.

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Composite Space Component 29jun2015

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