CHAPTER 3 PRELIMINARY DESIGN OF MICRO …shodhganga.inflibnet.ac.in/bitstream/10603/9567/13/11_chapter 3.pdf32 CHAPTER 3 PRELIMINARY DESIGN OF MICRO-SATELLITE STRUCTURE Structural

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CHAPTER 3

PRELIMINARY DESIGN OF MICRO-SATELLITE

STRUCTURE

Structural design of a satellite is a complicated iterative process that

contains selection of configuration, materials, design, analysis and testing.

Structural design is dependent on the design requirements set by the other

subsystems like power, propulsion, communications and thermal. The design

process starts at a conceptual stage with design specifications which are based

on mission requirements. In general the specifications include the

accommodation of payload and subsystems, launch requirements,

environmental protection, thermal and electrical paths, good stiffness, mass

efficiency and high reliability. The structural design process also provides a

interface to each individual unit or equipment in order to aid the integration

sequence.

3.1 METHODS OF SATELLITE DESIGN

Today, three separate methods exist within the field of satellite

design. These practiced methodologies are:

Conventional subsystem by subsystem design, unique to each

mission.

Modular design for multiple missions.

Low cost/high risk design.

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3.1.1 Conventional subsystem by subsystem design

Conventionally, satellites are designed uniquely for each mission.

After defining the requirements and constraints of the mission, each one of

the subsystems of the satellite namely power, propulsion, attitude

determination & control, thermal control, communications, command & data

handling, and the structure - are designed separately and iteratively. This

method of satellite design probably produces the most reliable spacecraft for a

particular mission but also makes the spacecraft most expensive.

3.1.2 Modular design for multiple missions

In the second methodology - modular designs for multiple missions

- the satellite subsystems are developed in the same manner as before.

However, the requirements on which the designs made are not for a single

unique mission, but for an anticipated range of missions. Thus, the

development costs are significantly reduced because each satellite is not

designed from the scratch. Each bus contains a nearly identical 1) modular

power system, 2) command & data handling module, 3) attitude determination

& control system, 4) modular structure, and 5) payload interface. Only the

payload instruments differ from mission to mission. The modular satellite

design provides the advantage of reducing the development costs at the

expense of not providing the most optimal design for a single given satellite

mission.

3.1.3 Small satellites with low cost/high risk design

“Small Sat Revolution” a new satellite design methodology

emerged to produce satellites with less production cost and with a high-risk

design. In addition to developing satellites weighing thousands of kilograms

and costing hundreds of millions of dollars, engineers began designing

"smallsats" weighing 200 kg or less and costing only a couple million dollars.

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The size of these small satellites also reduces the operational costs. This

methodology asserts that launching of many small, less capable, high risks,

low cost satellites to perform a mission will in the long run prove cheaper

than launching a few large, highly capable, overly redundant, lower risks, and

a very high cost satellites.

Therefore on the positive side, smallsats are cheaper than the

conventional satellites and afford the space flight opportunities for groups that

would otherwise be unable to afford one aboard a conventional satellite. On

the negative side, smallsats cannot carry as many instruments as, they have a

shorter lifetime than, and are more susceptible to single point failures than

conventionally designed and sized satellites.

In the process of design and development of a micro-satellite

structure within the specified envelope and also to accommodate 40 numbers

of subsystems a wide collection of existing spacecraft configurations, mass,

mission and structural concepts data was analyzed and the commonalities

were identified. The idea behind this is to identify the most common type of

structure like central tube, truss, frame, equipment or instrument box,

equipment platform, solar array panel and antenna support structure that can

be used in the satellite for various missions, the advantages and disadvantages

of various structural elements.

3.2 PRIMARY STRUCTURAL DESIGN

Primary structures are designed using several criteria that depend

on the mission requirements. Conventional spacecraft incorporate 4 basic

primary structural designs:

1) Skin-frame structures

2) Truss structures

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3) Monocoque cylinders

4) Skin-stringer structures

3.2.1 Skin-frame structures

The skin-frame structural design uses an interior skeletal networkof axial and lateral frames to mount exterior skin panels using fasteners or

rivets. The frames support bending, torsion, and axial forces. The skinreinforces the structure by supporting the shear forces introduced by the

interior member connections. The skin is sometimes minimized to save mass,even though the thin skin leads to some structural instability. When the skin

buckles due to shear, it transfers all additional shear loads to in-plane tensionforces at 45° which must be supported by the connections. The buckling

modes of the skin exhibit large deformations that make it insufficient forexterior mounted components such as solar cells. The buckling strength of the

assembly is typically increased by adding intermediate members.

3.2.2 Truss structures

Truss structures use an array of members that can only support the

axial loads. Truss members are produced independently and arrangedtypically in an array of triangles for stability. The members are manufactured

using extruded tubes made of composite, metallic, or sheet metal materials. Astable truss is statically determinate and has no excess members to introduce

the alternate load paths. Trusses are generally mass efficient when themembers are configured into rectangular or triangular cross-sectional

assemblies. However, they become less efficient as the cross-section becomesmore circular or hexagonal. Also, the design of the structure creates an

inherent stress concentration at the interface mounting points, such asseparation systems. Components may be mounted both internally and

externally and the absence of shear panels enables easy access to a payload.

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However, this absence of shear panels is not helpful to spacecraft requiringbody mounted solar cells.

3.2.3 Monocoque cylinders

Monocoque cylinders are axisymmetric shells that do not contain

any stiffeners or frames. The shells are manufactured by using metallic or

sandwich panels with curved sections formed by rolling. Typically, two or

three curved sections are fabricated and assembled into the cylindrical

configuration. The strength of monocoque cylinders is usually limited by its

buckling strength. The shells are most efficient when the loads are distributed

evenly throughout the structure. Components are typically mounted to the

walls using fasteners; however care must be taken not to overload the shell

and cause local failures. The monocoque cylinder design is applicable to

spacecraft with body mounted solar cells and relatively lightweight

components.

3.2.4 Skin-stringer structures

Cylindrical skin-stringer structures are designed using axial and

lateral frame members attached to an outer skin. These designs are similar to

skin-frame structures; however, this class of structures refers to circular

cylinder configurations. The skin is sometimes minimized to save mass, even

though the thin skin leads to some structural instability. The Post-buckling

behavior of the skin transfers the additionally applied shear loads to torsion by

the diagonal tension phenomenon described above. The skin and members

must attach uniformly to enable the assembly to act as a continuous structure.

Typical connection methods include fasteners and/or rivets. Interior

components are usually mounted to the walls at locations along the stringer

assembly. This method is more efficient than monocoque cylinder component

mounting at introducing local loads. The skin must be designed sufficiently

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stiff to enable sound mounting of exterior entities such as body mounted solar

cells.

The absence of shear panels in truss structures cannot be used for

body mounted solar panels. The monocoque cylinder structure is difficult to

assemble and integrate as it is generally made of two or three curved sections

and components are mounted on the wall as it does not contain stringers or

frames. The interior components in the skin-stringer structure are usually

mounted to the walls at locations along the stringer assembly and hence

difficult to mount more number of components. As the skin-frame structures

use interior skeletal network of axial and lateral frames to mount exterior skin

panels using fasteners which is further used to fix the body mounted solar

panels. The axial and lateral skin-frame structures are used to accommodate

more number of subsystems than any other structures.

3.3 CONVENTIONAL SATELLITE STRUCTURES

3.3.1 Sandwich Structures

Sandwich structures are often used in skin-frame designs. A

sandwich structure consists of two thin face sheets attached to both sides of a

lightweight core. The design of sandwich structures allows the outer face

sheets to carry the axial loads, bending moments, and in-plane shears while

the core carries the normal flexural shears. Sandwich structures are

susceptible to failures due to large normal local stress concentrations because

of the heterogeneous nature of the core/face sheet assembly. Component

mounting must therefore use potted inserts to distribute the point loads from

connections. Sandwich panel face sheets are commonly fabricated using

aluminum or graphite/epoxy composite panels. The core is typically

fabricated using a honeycomb or aluminum foam construction. Honeycomb

sandwich paneling is the lightest option for compressive or bending loading

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specific applications. Honeycomb sandwich cores are manufactured using

thin strips formed into honeycomb cells. The honeycomb geometry is non-

isotropic with greater stiffness in the longitudinal direction. However, the core

acts nearly isotropically for in-plane loads when assembled in a sandwich

configuration. The disadvantages of using honeycomb cores are the potted

inserts required for mounting and the thermal inefficiencies. These

inefficiencies stem from the low thermal conductivity of the adhesive layers

used in construction and make use of honeycomb prohibitive in optical and

mirror aerospace application.

Figure 3.1 Sandwich panel

3.3.2 Multifunctional Structures

Multifunctional structure (MFS) technology shown in Figure 3.2

incorporates several functions into the primary structure of a spacecraft. The

main goal of these members is to minimize unwanted mass by incorporating

chassis, cables, connectors and thermal control components into the satellite

primary structural walls. The walls are typically constructed out of fiber

composites or sandwich panels, and the electrical components are embedded

during manufacturing. The design allows for an easily accessible, removable,

and modular electrical system. The benefits of this technology include a 70%

reduction in electronic enclosures and harnesses, a 50% reduction in

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spacecraft volume required for these conventional components, a reduction in

labor required for spacecraft assembly and an extremely robust system with

wide applicability to several missions.

Figure 3.2 Multi functional structural panels with integral electronic,

structural and thermal control

3.3.3 Isogrid Structures

An Isogrid panel is a plate (or face sheet) with triangular integral

stiffening ribs (often called stringers). The isogrid pattern of triangular shape

shown in Figure 3.3 was found to be very efficient because it takes the

advantage of the fact that triangular trusses are very efficient structures. The

term isogrid is used because the structure acts like an isotropic material. The

pattern may be manufactured by machining a metallic panel, or it may be

constructed using the fiber composite materials.

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Figure 3.3 Isogrid pattern

It is much necessary to keep the C.G and M.I values as per the

launcher requirement so that it may be required to move any of the 40

subsystems or to add dummy mass inside the micro-satellite structure. It is

very difficult to satisfy the C.G and M.I constraint given by the launcher with

more number of subsystems using the Multifunctional Structure or with

isogrid structures and also very difficult to accommodate the subsystem as per

its requirements. Hence it was decided to consider the sandwich panel

structures for the development of the micro-satellite under consideration.

3.4 EXTERNAL SHAPE DESIGN

The geometrical shape of the external satellite structure is important

in sending a small compact and light weight structure into space and should

have the capacity to contain all the subsystems required and easy access to all

the subsystems and instruments during the assembly and testing. It is also

necessary to design the external spacecraft geometry to keep the

manufacturing cost as low as possible. The shape and size of the external

structure of the satellite are closely related to the solar cell mounts and is

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required to attach solar cells to the main body. The cylindrical, spherical,

cubical, hexagonal and pentagonal shaped external structures are shown in

Figure 3.4 to Figure 3.8 respectively.

Figure 3.4 Cylindrical shaped structure

Figure 3.5 Spherical shaped structure

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Figure 3.6 Cube shaped structure

Figure 3.7 Hexagonal pyramid shaped structure

Figure 3.8 Octagonal pyramid shaped structure

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The Table 3.1 gives the surface area exposed to the sun and the

volume available inside the micro-satellite external geometry in the envelope

of the 600*600*600 mm3.

Table 3.1 Surface area and volume of micro-satellite external shape

configurations

Shape Surface area (mm2) Volume * 108 (mm3)

Cylindrical

Spherical

Cubic

Hexagonal

Octagonal

565486

565486

360000

421728

596400

1.69

1.12

2.16

1.789

1.788

For the design of the external spacecraft geometry, it is necessary to

keep the manufacturing cost as low as possible. This leads to the selection of

flat surfaces only in order to avoid the growing of costs necessary for tooling

the curved surfaces.

Several considerations like easier manufacturing, the

accommodation of satellite into the launcher in the permitted envelope of

600mm×700mm ×800mm, reduced effort to fit 40 subsystems mostly of

cuboids shape, inside the spacecraft volume points out the best shape to be a

regular prism, composed by flat surfaces. Since it is very difficult to use much

space because of uncommon angles and to avoid additional seams and joints

which will add more stress concentration in hexagonal and octagonal design

as far as the number of sides is concerned cubic is preferred over the other

designs.

The following Table 3.2 shows the advantages and disadvantages

of the different spacecraft external geometry.

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Table 3.2 Advantages and disadvantages of micro-satellite external shape

configurations

Shape Advantages DisadvantagesCylindrical Maximized volume

Maximized surface areafacing toward the sun

Complex designDifficult tomanufacture

Difficult to attach solarcells

Spherical Maximized volumeMaximized surfacearea facing toward thesun

Complex designDifficult tomanufactureIncreased cost fortooling curved surfacesDifficult to attach solarcellsDifficult to usemaximum volume

Cubic Simplistic designEasy to manufactureEasy to attach solarcellsMaximum use ofavailable space orvolumeReduced effort to fitall the subsystemsinside the spacecraftvolumeLess seams and joints

Does not expose muchsurface area towardsthe sun

Hexagonal Increased surface areaper unit of volumeexposed to sunImproved capacity

Additional seams andjointsIncreased complexity

Octagonal Greatly increasedsurface area per unitof volume exposed tosunGreatly improvedcapacity

Many additional seamsand joints add stressconcentrationsIncreased complexityDifficult to use spacebecause of uncommonangles

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3.5 SOLAR PANEL MOUNTING

The solar cell mounts are closely related to the shape and size of the

structure. Deployable solar panels would be required if the body mounted

solar cells are inadequate to generate power to perform the satellites mission

while staying within the maximum dimensions of the launch vehicle. Both

body mounted and deployable solar panels have been considered on need

basis.

Body mounted solar panels shown in Figure 3.9 is the simplest and

most dependable way of attaching solar cells to the satellite. These cells are

simply fixed to the sides of the satellite’s main body and absorb energy when

that side of the satellite faces towards the sun. This design carries a lower cost

than deployable cells due to the fact that it requires no moving parts or special

equipment. The primary drawback of this system is the highly limited amount

of power these cells can acquire due to the small size of the solar arrays and

inability to point them all towards the sun at once.

Deployable solar panels shown in Figure 3.10 to Figure 3.12 would

result in large increase in power output by comparison. This benefit would be

offset by the complicated structural support and mechanisms necessary for

deployment of the solar cells. The final design was made based on whether

body-mounted cells will supply the necessary power for the satellite or not.

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Figure 3.9 Body mounted solar panel

Figure 3.10 Flowering mounted solar panel

Figure 3.11 Unwrapping mounted solar panel

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Figure 3.12 Telescoping solar panel

For the micro-satellite considered the power system generates,

stores and provides power for the subsystems and the payloads of the satellite

as required. Store and Forward, Micro Electro Mechanical System and the

Global Positioning System are the planned payloads for the satellite. The

subsystems are supplied power both in the continuous and peak modes.

Power is generated through body-mounted solar panels on four

bellybands. The power generated varies periodically with the period of the

spin. During sunlight period excess energy is stored in batteries and during

eclipse period power is drawn from the batteries. The description of the body-

mounted solar panel fixed to the micro-satellite is given in Table 3.3.

Table 3.3 Panel description

Panel Name Area availablefor mounting(mm2)

Comments

Solar panel 1

Solar panel 2

Solar panel 3

Solar panel 4

600*600

600*600

600*569

600*569

20% area is used for mountingscrews and routing.20% area is used for mountingscrews and routing.20% area is used for mountingscrews and routing.20% area is used for mountingscrews and routing.

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Series and parallel circuits are formed with the solar cells to obtain

the required voltage and current levels. The cell dimension, bus voltage,

current and size of the satellite decide the number of series and parallel

circuits. The following points are to be considered while designing the solar

strings.

Cells arranged in series in each panel are called sub-string.

Each sub-string has fixed number of cells to meet the bus

voltage.

A main string is formed by connecting at least one sub-string

from each panel in parallel.

The main strings are in the switchable form to regulate the

power input to the bus.

Number of main strings was selected to provide maximum bus

demand even at the event of on string failure.

The power required by the micro-satellite considered is 53W and

the power requirement is met with the body-mounted Gallium-Arsenide cells

as given in Table 3.4.

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Table 3.4 Solar cell-String comparison

Cell description Silicon cells Gallium-Arsenide cells

Size of cell 20mm*40mm 20mm*40mm

Number of cells persub-string

39 23

Power per cell(W) 0.125 0.2

Voltage per cell(V) 0.45 0.78

% area occupied by onesub-string

10% 6%

Average power of onemain string (W)

4 4

Number of main stringspossible

8 14

Total power (W) 32 56

3.6 INTERNAL STRUCTURAL CONFIGURATION

The internal mounting faces serve two purposes: i) as mounting

surfaces for various components and ii) as additional structural support for the

spacecraft. A strong internal structure is desirable which is normally achieved

at the cost of additional mass. Four concepts were considered for the internal

mounts as given in Table 3.5

Table 3.5 Summary of internal structure configuration

Concepts Advantages DisadvantagesPlatform Improved lateral support No added vertical

supportCrossed Improved lateral and

vertical supportLoss of usable space

Trunk Much improved lateraland vertical support

Higher mass

Open Very low mass Loss of symmetry

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Even though the crossed internal structural configuration adds mass

and loss of usable space still it is better to accommodate more number of

subsystems within their requirements, improved lateral and vertical support

compared to other configurations and easy achievable of C.G and M.I values

according to launcher requirements. For selecting a suitable crossed internal

structural configuration a preliminary free vibration analysis was made for

some 11 possible options without subsystems. The shape, the mass and the

natural frequency for the first three modes of all the possible options are given

below.

3.6.1 Option 1

Full-length vertical cross webs between the top deck and the

bottom deck with a height of 600mm between them as in Figure 3.13.

Figure 3.13 Configuration 1

3.6.2 Option 2

Full-length vertical cross webs between the top deck and the

bottom deck with a height of 450mm between them as in Figure 3.14.

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3.6.3 Option 3

Framework cuboids of 600*600*600 mm3, a bottom deck, a middle

deck and a top deck connected by angle sections between them and without

any cross webs as in Figure 3.15.


3.6.4 Option 4

The structure consists of a deck plate placed in-between the top

deck and the bottom deck exactly at the center, two cross webs passing

between them along X-direction and one cross web passing full length along

Y-direction. The overall dimension of the structure is 600*600*600mm3. In

the same way there are two cross webs passing along X-direction and one

cross web passing full length along Y-direction between the middle deck and

the bottom deck as in Figure 3.16.

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3.6.5 Option 5

The structure consists of top deck and bottom deck placed at a

distance of 600mm with the middle deck placed in-between them at the

center. There are diagonal cross webs between top deck and middle deck,

middle deck and bottom deck as in Figure 3.17.


3.6.6 Option 6

The structure consists of diagonal cross webs between bottom deck

and middle deck and vertical cross webs passing full length between middle

deck and top deck. The overall dimension of the structure is

600*600*600mm3 as in Figure 3.18.

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3.6.7 Option 7

Full diagonal cross webs pass between the bottom deck and middle

deck and half Diagonal cross webs between middle deck and top deck as in

Figure 3.19.


3.6.8 Option 8

Half diagonal cross webs between bottom deck and middle deck,

also between the middle deck and top deck. The overall dimension of the

cuboids is 600*600*600 mm3 as in Figure 3.20.

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3.6.9 Option 9

Half vertical cross webs between bottom deck and middle deck and

also between the middle deck and top deck. The overall dimension of the

cuboids is 600*600*600 mm3 as in Figure 3.21.


3.6.10 Option 10

Full vertical cross webs in between bottom deck and middle deck -

half vertical cross webs in between middle deck and top deck. The overall

dimension of the cuboids is 600*600*600 mm3 as in Figure 3.22.

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3.6.11 Option 11

This configuration consists of bottom deck, middle deck having a

small hole at the center and the top deck as the dimensions mentioned in

options 10 which is shown in Figure 3.23. The vertical cross webs between

the bottom deck and the middle deck have cutouts at the top and the bottom.

These cutouts, holes are provided for the purpose of integration. Also it has 4

numbers of half-length vertical cross webs between the top deck and the

middle deck.


The mission needs i) the structure of less mass and stress, ii)meeting the constraints of Center of Gravity, Moment of Inertia andFrequency and iii) accommodation of all the subsystems inside the structure.Eleven (11) best possible configurations of the micro-satellite were createdand analyzed using MSC PATRAN/MSC NASTRAN software.

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On the basis of mass, stress and first three frequency values a studywas made on the eleven possible structural configurations for the micro-satellite considered with body-mounted solar panels. The structuralconfigurations with very high mass were rejected first and further theconfigurations with very low frequency values and high stress values werealso not considered because of their low stiffness and low strength values.Then the remaining configurations were considered to find out whether theyaccommodate all the possible subsystems and also provide easy integration.

The variation of structural mass, stress values and lateral andlongitudinal frequencies with structural configurations are shown in Figures3.24, 3.25 and 3.26 respectively. From the Figures 3.24, 3.25 and 3.26 it isestablished that the best possible structural configuration is configuration 11which has i) less mass, ii) good stiffness and strength values, iii) goodprovisions for integration and iv) enough volume to accommodate all thesubsystems needed as per the general requirements. So configuration 11 hasbeen used for all the analyses and in the development of the satellite.

Figure 3.24 Variation of structural mass for various configurations

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Figure 3.25 Variation of stress values for various configurations

Figure 3.26 Variation of frequency for various structural configurations

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3.7 MATERIAL SELECTION

In order to meet the mission requirements many lightweight and

durable materials are utilized in the micro-satellite structural design. Material

strength is integral to the structural survival of the spacecraft. The materials

used need to be lightweight with high yield strength and rigidity. The

important characteristics that are demanded for the development of micro-

satellite structure are strength, stiffness, and low density, uniform coefficient

of thermal expansion between hull and frame, low specific heat, low cost and

easy machinability.

3.7.1 Steel

Most structural steels are penalized by their magnetic properties,

although austenitic stainless steels are non-magnetic. Unfortunately, the

stiffness to the density ratios for austenitic steels is lower than those of the

aluminum alloys. Nonetheless, these materials can be still utilized for their

high strength in instances where the titanium is not desirable, perhaps for

machining or temperature reasons.

3.7.2 Titanium

Titanium, a non-magnetic material is used in many applications

where aluminum structure does not possess the required strength. Though

somewhat more difficult to machine, it has a substantially greater yield

strength, a higher stiffness to density ratio, and is particularly suitable for low-

temperature applications (e.g., cryogenic fuel storage). At high-temperatures,

titanium is surpassed by steels, though it still exceeds aluminums capabilities.

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Research may greatly increase the ease of manufacturing complex titanium

components.

A relatively new class of titanium based materials is the

intermetallic titanium alumides, or tialuminides. These low density materials

exhibit high strength at temperatures greater than 700° C and resist oxidation

at all temperatures. Primary uses are as a composite matrix material and in

honeycomb structures. Unfortunately, tialuminides react poorly with

hydrogen and become brittle. This limits their utility in vehicles that use

hydrogen to actively cool the airframe unless protective coatings are used.

3.7.3 Magnesium

Though its stiffness to density ratio is close to aluminum, this

material and its alloys are prone to brittle fracture. This reduces its

applications to those where its barely higher yield strength may be of some

use, or where its good low-temperature behavior is paramount. As already

mentioned, magnesium sublimes relatively quickly in vacuum (0.04 in/yr at

180° C). Exposure to temperatures less than 250° C can cause it to lose static

strength, although electroplating or coatings can prevent this.

3.7.4 Beryllium

With a density approximately 60% of aluminum and a stiffness to

weight ratio six times better than aluminum or titanium, this material has

many potential applications. Being stiffer than other materials it can be useful

in avoiding resonant frequencies that may occur between a satellite and its

booster during launch. It is non-magnetic, has a high elastic modulus and high

yield strength. Using beryllium instead of aluminum can reduce weight

significantly, and its high thermal conductivity makes it an excellent choice

for components that will conduct heat. It is also non-reactive with hydrogen.

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However, beryllium is extremely anisotropic and sensitive to

damage, and is twice as brittle as aluminum. Its relatively low fracture

toughness at cryogenic temperatures is a potential drawback, but development

of beryllium-aluminum alloys may improve this. It also needs special

facilities and tools for machining due to the toxicity of its dust, and is thus

very expensive.

3.7.5 Composites

Composite materials are quickly becoming the material of choice

for aerospace applications. They possess stiffness to weight ratios beyond all

metals, making them useful for damping unwanted vibrations. With a

negative axial coefficient of thermal expansion, they allow for structures that

will not deform in the temperature extremes of space. Their thermal

conductivity exceeds copper, and so they also provide lightweight thermal

management and heat sinks. Where stiffness is critical, as in telescopes,

antennas, and reflectors, carbon fiber composites are another natural choice.

Metal matrix, Carbon-Carbon, and Ceramic-Matrix composites are

best for high temperature applications, such as re-entry vehicle skins, since

they can withstand temperatures in excess of 2500° F with no active cooling.

However, there are some drawbacks to these materials. Effective oxidation

coatings must be developed, as well as manufacturing techniques for a large

scale structures. Grounding of electrical systems is done by adding conductive

strips, which increase the mass of the structure. Nicks and dents that can be

repaired or ignored in structural metals can destroy the integrity of the fibers

and render the composite unusable. In space, with little or no inspection and

maintenance, and where failure of primary structure can have devastating

consequences, composites are typically judged too unreliable for use as more

than secondary structure. But judicious use of composites in secondary

structure can still result in much mass savings. Another concern of laminated

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composites is their reaction to temperature changes. Uniform changes can

induce substantial internal stresses caused by different expansion rates

between the fibers and matrix. Temperature differentials can produce more

pronounced warping than in an isotropic material.

Composite materials also have been considered but it does not

appear to be a suitable solution because of the frequent changes in the

structure (due to iterative design process), which is not compatible with

composites technology.

3.7.6 Aluminum

A combination of high stiffness to density ratio, excellent

workability, non-magnetism, moderate cost, high ductility, high corrosion-

resistance, and availability in numerous forms makes aluminium the best

choice for micro-satellite considered. Its low yield strength is the only

appreciable disadvantage. Hence the material used for the micro-satellite is

considered as aluminium in the form of rods, plates and honeycomb panels.

Honeycomb aluminium is used for flat panels due to its extremely low density

in comparison to the aluminium plate of the same strength. The subsystems

are manufactured by shop machining process starting from the aluminium

block by which the unnecessary material is removed.

Documents

CHAPTER 3 PRELIMINARY DESIGN OF MICRO …shodhganga.inflibnet.ac.in/bitstream/10603/9567/13/11_chapter 3.pdf32 CHAPTER 3 PRELIMINARY DESIGN OF MICRO-SATELLITE STRUCTURE Structural