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PRATHAM

IIT BOMBAY STUDENT SATELLITE

Performance Verification Review

Structures Subsystem

By

Pratham Team

Department of Aerospace Engineering,

Indian Institute of Technology, Bombay

May, 2012

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Contents Chapter 1 .................................................................................................................................................... 5

Introduction ............................................................................................................................................ 5

1.1 Overview of report ........................................................................................................................ 6

Chapter 2 .................................................................................................................................................... 7

Requirements & Constraints ................................................................................................................... 7

2.1 System requirements: ................................................................................................................... 7

2.2 Sub-system requirements: ............................................................................................................ 7

Chapter 3 .................................................................................................................................................. 11

Configuration & Interfaces .................................................................................................................... 11

3.1 Full Configuration Layout: ........................................................................................................... 14

Chapter 4 .................................................................................................................................................. 19

Design ................................................................................................................................................... 19

4.1 Design Approach: ........................................................................................................................ 19

4.2 Structural Elements:.................................................................................................................... 20

4.3 Factors of safety: ......................................................................................................................... 20

4.5 Material Properties: .................................................................................................................... 20

Chapter 5 .................................................................................................................................................. 22

Analysis ................................................................................................................................................. 22

5.1 Modeling: .................................................................................................................................... 22

5.2 Simulations & Results: ................................................................................................................ 22

Acknowledgements................................................................................................................................... 34

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List of Tables

Table 1: Static loading levels during launch............................................................................................ 8

Table 2: Sine sweep test levels for qualification ..................................................................................... 9

Table 3: Stiffness requirements ............................................................................................................... 9

Table 4: Elements used to mesh satellite structure ............................................................................. 23

Table 5: Static analysis results ............................................................................................................... 24

Table 6: First 15 fundamental frequencies ........................................................................................... 26

Table 7: Harmonic response of middle antenna ................................................................................... 32

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List of figures

Figure 1: Internal configuration layout ................................................................................................. 11

Figure 2: External configuration layout ................................................................................................ 11

Figure 3: Solid Works model .................................................................................................................. 13

Figure 4: Solid Works model: Inside view ............................................................................................. 13

Figure 5: Solid Works model: Inside view ............................................................................................. 14

Figure 6: Zenith ....................................................................................................................................... 14

Figure 7: Nadir ......................................................................................................................................... 15

Figure 8: Sun-side.................................................................................................................................... 15

Figure 9: Anti sun-side ............................................................................................................................ 16

Figure 10: Leading side ........................................................................................................................... 16

Figure 11: Lagging side ........................................................................................................................... 16

Figure 12: Satellite model showing the co-ordinate axes .................................................................... 22

Figure 13: von Mises stress – Static analysis ........................................................................................ 25

Figure 14: 1st mode - antenna mode ..................................................................................................... 26

Figure 15: 3rd mode - antenna mode .................................................................................................... 29

Figure 16: 7th mode - panel mode ......................................................................................................... 29

Figure 17: 10th mode - panel mode ....................................................................................................... 29

Figure 18: 14th mode - panel mode ....................................................................................................... 30

Figure 19: First 25 modes ....................................................................................................................... 30

Figure 20: Normal stress (Y) vs. frequency (longitudinal loading) ..................................................... 32

Figure 21: Y deformation vs. frequency (longitudinal loading) .......................................................... 32

Figure 22: Normal stress (Y) vs. frequency (lateral loading) .............................................................. 33

Figure 23: Y deformation vs. frequency (lateral loading) .................................................................... 33

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Chapter 1

Introduction

Pratham is student satellite designed and to be fabricated by students of Indian Institute

of Technology Bombay. Pratham is a nano-satellite with weight 9.57 Kg and dimensions

of the main body 254mm x 254mm x 253mm. There are 3 monopole antennae with

lengths 183mm each. The scientific mission of Pratham is to measure the total electron

count of ionosphere and perform the tomography of ionosphere.

The requirement that the structure must satisfy is that it must house all the components

and ensure that it does not fail during the lifetime of the satellite. During the lifetime, the

satellite will undergo various types of loading. Structural subsystem should ensure that

satellite structure is able to withstand the loading and no component should fail due to

structural loading.

The design of satellite structure is dependent on many factors. Some of the most

important factors are placement of components and the material properties of

components. The design approach that is followed in this satellite is first system

engineering team prepares a configuration layout and structures team analyses it and

decides the parameters like material of satellite body and thickness of the material. The

analysis is given to system engineering team and there is some flaw in the system,

system engineering team prepares a modified configuration layout. This iterative

process is followed till a design satisfying both system engineering and structural

requirements is obtained. The various parameters that can be changed by structural

subsystem are material properties of structure, geometric parameters like thickness of

the structure and joining mechanisms.

Finite element software ANSYS is used for structural analysis. A structural model

is prepared and meshed to get the finite element model. Various static as well as

dynamic analyses are performed on this finite element model to get the response of the

structure to various types of loading. ANSYS can be used to perform both static as well as

dynamic analyses on the structure. Validation of the results obtained using ANSYS is

done by first validation of the finite elements and geometry by comparing results with

theoretical results and then by analysis of individual structural element in ANSYS

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isolated. This approach was giving a good match between the results obtained from

finite element analysis and theoretical results as well as between individual element and

entire structural model. The results obtained using FEA suggest that satellite will be

above the failure stresses and strains when specified loads are applied and no

component on the satellite fails due to these loads.

1.1 Overview of report

In chapter 2, requirements and constraints are given. Chapter 3 describes the

configuration and interfaces. In chapter 4, the design is presented and in chapter 5, all

the simulations that have been performed are described.

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Chapter 2

Requirements & Constraints

The following section details the requirements imposed on the Structural Subsystem by

the other subsystems of the satellite and the constraints under which the Structural

Subsystem needs to design the satellite. It gives a broad look at the various tasks

handled by the subsystem and their significance in the overall scheme of the satellite.

2.1 System requirements:

1. All components should be properly housed in the satellite structure. Placement of all

components should be proper and there should be no interferences due to other

components.

2. Satellite structure should be able to withstand the loads during launch. All the

components should be safe and working after the launch.

3. Satellite structure should be able to withstand the thermal loads arising in the orbit.

2.2 Sub-system requirements:

2.2.1 Weight budget and constraints:

The launch cost of any satellite is directly dependent on the weight to be launched. To

keep launch costs to a minimum, the total weight of the satellite must be kept to a

minimum.

1. The maximum weight has been fixed at 16kg. This is purely based on the limit

imposed on the CubeSATs which allow a maximum weight of 1.0kg for a cube

of side 100mm, carried over to PRATHAM which has approximately sixteen

times the same volume as the CubeSATs.

2. The structure of the satellite itself must not take up more than 30% of the total

weight of the satellite.

3. The detailed description of the weight budget and its related issues can be found in the System Engineer’s documentation.

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2.2.2 Launch Vehicle Placement Requirements:

The satellite is launched into Low Earth Orbit by the Polar Satellite Launch Vehicle. The

launch vehicle interface to be used is the IBL230V2, to be provided by VSSC.

1. Launch vehicle interface requires 8 no’s M6x1, 9mm long helicoil inserts at 230mm

PCD on bottom deck of the satellite.

2. There should not be any interference in the joint from the satellite to the launch

vehicle body.

2.2.3 Launch Loading Requirements:

The satellite is carried to its orbit by a launch vehicle in a flight lasting about 17

minutes. During this period, the vehicle experiences high levels of acceleration,

vibrations and shocks which are transmitted to the payloads attached to the flight decks

of the vehicle. Launch loads experienced include static loads, vibration loads, acoustic

loads and shocks and impose certain strict requirements on the structure of the

satellite.

The loading specification for which the launch vehicle interface is tested is assumed to

be the loading data for the satellite during launch.

2.2.3.1 Static Loading:

Static loading occurs on the satellite during launch as a result of the accelerations

experienced during flight. The static loads that are used for testing to verify the design

to ensure that structure meets the safety margins are as listed in Table 1.

Direction Loading

Longitudinal ± 11g

Lateral ± 6g

Table 1: Static loading levels during launch

Lateral loads are considered to act simultaneously with longitudinal loads. Earth’s

gravity is also included in the above levels. All loads apply at the center of gravity of the

satellite as body forces.

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2.2.3.2 Harmonic Loading:

The levels defined for qualification and acceptance in the sinusoidal vibration sweep

test are as given in Table 2. The satellite is tested in conjunction with its launch

interface on the shaker table. Here, the ‘Qualification level’ is used to prove that the

structure can withstand loads, which are a factor greater than the expected flight

limit loads (qualification loads), while the ‘Acceptance level’ is used to discover

production deviations, not discovered during inspection. The applied loads are

equivalent to the flight limit loads.

Frequency range

(Hz)

Qualification level Acceptance level

Longitudinal axis 5-10

10-100

10mm (DA)

3.75g

8mm (DA)

2.5g

Lateral axis 5-8

8-100

10mm (DA)

2.25g

8mm (DA)

1.5g

Sweep rate 2 oct/min 4 oct/min

Table 2: Sine sweep test levels for qualification

These levels are defined at the interface of the satellite with the deployer. The test

is to be carried out along all three axes of the satellite, on the flight model.

2.2.4 Stiffness Requirements:

The requirements for stiffness of the satellite are such that there should be no

component onboard the satellite which is free to vibrate at a natural frequency below

the specified limits. This implies that all extended structures must thus comply with the

stiffness requirements, as well as the structure as a whole.

Global fundamental frequency in longitudinal mode >90Hz

Global fundamental frequency in lateral mode >45Hz

Local fundamental frequency >175Hz

Table 3: Stiffness requirements

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2.2.5 In-orbit Requirements:

The satellite structure should be able to withstand the cyclic thermal loading in orbit.

The orbit will give rise to differential thermal expansion and contraction. This

differential expansion and contraction gives rise to thermal stresses. The cyclic nature

of the stresses induces fatigue in the satellite which imposes constraints on the size of

microscopic defects present in the structure and the lifetime of the satellite. The

structure should be able to withstand this loading.

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Chapter 3

Configuration & Interfaces

The internal and external configuration of the satellite is shown in figures 1 and 2

respectively.

Figure 1: Internal configuration layout

Figure 2: External configuration layout

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The entire satellite structure was modeled in CAD software Solid Works 2009. The

entire model is shown in Figure 3 and the inside view is shown in Figures 4 and 5.

Figure 3: Solid Works model

Figure 4: Solid Works model: Inside view

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Figure 5: Solid Works model: Inside view

3.1 Full Configuration Layout:

3.1.1 Zenith

The zenith is the side of the satellite which always faces away from the earth. The Zenith

side of PRATHAM incorporates one of the solar panels and the GPS antenna along with

GPS circuit and a magnetorquer.

Figure 6: Zenith

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3.1.2 Nadir

The Nadir side always faces towards the earth. The Launch Vehicle interface which in

our case is the IBL-230 V2 will be attached to Nadir side. It will also have a battery.

Figure 7: Nadir

3.1.3 Sun-side

As the name suggests, the sun-side faces the sun. Our sun-side has a solar panel

mounted on it. From the inside, it has the power circuit and a magnetorquer.

Figure 8: Sun-side

3.1.4 Anti sun-side

It is the side opposite to the Sun-side. The monopoles are attached to the anti-sun side.

It has the monopole circuit and the beacon circuit.

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Figure 9: Anti sun-side

3.1.5 Leading side

The Leading side is normal to the direction of orbit of the satellite and has a solar panel

mounted on it and has a magnetorquer and the OBC circuit.

Figure 10: Leading side

3.1.6 Lagging side

The Lagging side is opposite the leading side. It contains the magnetometer and a solar

panel.

Figure 11: Lagging side

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Chapter 4

Design

4.1 Design Approach:

The design approach followed is given in a Figure 12.

Configuration layout

prepared by

integration team

Configuration

modeled in Ansys

Static and dynamic

analysis performed

in Ansys

No failure observed and NO all results are satisfactory for both integration and

structures

YES

The configuration is

accepted as final

configuration

Figure 12: Design approach

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4.2 Structural Elements:

1. The Satellite essentially consists of six structural members: the panels

which make the box of the body. These sides are named Leading-side,

Lagging-side, Sun-side, Anti-Sun-side, Zenith and Nadir depending on

their position during orbit.

2. The zenith, leading, lagging and sun-sides have solar panels mounted on

them. The solar panels are connected to Aluminium honeycomb which in

turn are connected to the main satellite body.

3. The zenith, leading and sun-sides have magnetorquers mounted on them,

which are the primary actuators of the satellite.

4. The anti-sun-side has 3 monopole antennae mounted on it.

5. The zenith consists of a GPS puck.

6. The battery and the battery box are mounted on the inner surface of the

nadir.

7. The electronics consists of 3 communication circuits, 1 OBC board,1 power

circuit, 1 sun-sensor board, a GPS circuit and a magnetometer.

4.3 Factors of safety:

The criteria for qualification of satellite for flight aboard the launch vehicle is

passing of various tests for static loading, sinusoidal vibrations, and random

vibrations and testing for shock and impact. The load factor for the satellite is

specified by the Interface Control Document to be 1.25, that is, the ultimate load

that can be withstood by the satellite is up to 1.25 times the actual qualification

levels specified.

4.5 Material Properties:

4.5.1 Material for the satellite body, solar panel backing plates and monopole

antennae:

Various alloys of aluminium were studied to decide which material to use

for the satellite body. These included Al 7075, Al 6061 and Al 2024.

Aluminium 6061-T6 alloy was chosen as the material for the satellite body.

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It is ductile, lightweight, and easily machinable, can have complex

structures milled out of whole ingots, easily procurable, cheap, and

sufficiently stiff and is also characterized for space applications, thus

making it a very attractive option. Due to its favourable properties and

space heritage, aluminium 6061-T6 alloy was decided upon as the material

of choice and hitherto there has been no cause, either due to weight

constraints or structural requirements, to revise this decision.

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Chapter 5

Analysis

5.1 Modeling:

The entire satellite structure, including the body, panels and other components onboard,

was modeled in Solid Works 2009. The joints were modeled using the “mate” function in

Solid Works. In this model, the joints are assumed to be at discrete points, where the

screws are placed and thus, provides a highly accurate representation of the structure.Only

the boards of the PCBs were modeled. The components on the PCBs were not modeled as

their design is not finalized.

.

Figure 12: Satellite model showing the co-ordinate axes

5.2 Simulations & Results:

Simulations were performed on:

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1. Vibration model: Consisting of the 6 sides, solar panels with washers, antennae and

antennae holder and FE ring. The vibration model was simulated mainly to validate

the Qualification model (QM) and to gain an understanding of vibration concepts.

2. Qualification model (QM): Consisting of all components onboard the satellite,except

the fasteners. However, the joints were appropriately mated, to accurately represent

the final model

3. Circuit board: A dummy board was simulated to launch loads to qualify its structure

for flight

5.2.1 Qualification model:

The co-ordinate system used in the analysis is the Cartesian co-ordinate system with origin

at the base of the Nadir of the satellite. The longitudinal axis used for this analysis is along

Y axis (axis joining Nadir and Zenith). Elements used to mesh the structure:

Generic element type name ANSYS name Description

10 node quadratic tetrahedron Solid187 10 node tetrahedral structural solid

20 node quadratic hexahedron Solid186 20 node structural solid

20 node quadratic wedge Solid186 20 node structural solid

Quadratic quadrilateral target Targe170 3D target segment

Quadratic triangular contact Conta174 3D 8 node surface to surface contact

Quadratic triangular target Targe170 3D target segment

Table 4: Elements used to mesh satellite structure

Boundary conditions:

The part of launch vehicle interface attached to the satellite is constrained in space. This

is the correct assumption as in actual launch; it will be constrained using ball lock

mechanism constraining all the degrees of freedom.

5.2.1.1 Static analysis of QM:

1. Aim of analysis: To determine the stress developed and total deformation when a

static load is applied.

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2. Type of analysis: Static

3. Material properties: As per data

4. Constraints applied: The FE ring, attached to the base of the Nadir was constrained

for movement in all directions.

5. Loads applied: 11g acceleration in +Y direction and 6g accelerations in +X and +Z

directions, as specified in table 1.

6. Transcript of results:

Maximum von Mises stress (occurs on a washer) 20.51 MPa (4.6% of failure stress)

Maximum total deformation (at the tip of the

middle antenna)

0.214mm

Table 5: Static analysis results

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7. Screenshots of contour plots obtained:

Figure 13: von Mises stress – Static analysis

8. Interpretation of results: The maximum stress obtained in the analysis is far less than

the yield strength. Also, the total deformation is far too less for any contact between 2

surfaces to take place.

9. Conclusion: After being loaded with maximum static loading levels possible during

actual launch, we can say that the structure will maintain its integrity. Hence the

structure will not fail under static loads even in such a worst case analysis.

5.2.1.2 Modal analysis of QM:

1. Aim of analysis: To determine the fundamental frequencies of the structure.

2. Type of analysis: Modal

3. Material properties: As per data

4. Constraints applied: The FE ring, attached to the base of the Nadir was constrained

for movement in all directions.

5. Loads applied: NIL

6. No. of modes extracted: 25

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7. Transcript of results:

The first 15 natural frequencies are presented in Table 15.

Mode Frequency (Hz)

1 152.46

2 154.45

3 155.21

4 155.46

5 155.90

6 156.14

7 168.04

8 168.41

9 170.47

10 195.41

11 296.94

12 297.20

13 297.24

14 342.06

15 342.53

Table 6: First 15 fundamental frequencies

8. Screenshots of contour plots obtained:

Figure 14: 1st mode - antenna mode

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Figure 15: 3rd mode - antenna mode

Figure 16: 7th mode - panel mode

Figure 17: 10th mode - panel mode

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Figure 18: 14th mode - panel mode

9. Screenshots of graphs plotted:

Figure 19: First 25 modes

10. Interpretation of results:

The first 6 modes are antenna modes. The first panel mode is the 7th mode of the

structure. The first side mode (Zenith) is the 20th mode (=416.48 Hz) of the satellite. Of

the first 25 modes extracted (25th free-free mode = 572.93 Hz), no PCB modes were

observed.

11. Conclusion:

The fundamental frequency of the structure is greater than 90 Hz as demanded by

global stiffness requirements. We can say that the structure will not resonate with the

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Launch Vehicle Interface and will survive launch without any deformation. Moreover, the

absence of PCB modes is reassuring for the survival of onboard PCBs.

5.2.1.5 Harmonic response:

1. Aim of analysis: To obtain the frequency response of the structure when subjected to

sinusoidal loads.

2. Type of analysis: Harmonic

3. Material properties: As per data

4. Constraints applied: The structure was assembled on a block having a weight 100

times that of the satellite. This was done so as to make the structure of the satellite

flexible and not rigid.

5. Loads applied: 2 analyses were performed. In the first case, a pressure equivalent to an

acceleration of 3.75g in the Y direction was applied to the base of the block. In the second

case, a pressure equivalent to an acceleration of 2.25g was applied in the X direction. (As

specified in Table 2)

6. Frequency range: 10Hz – 200Hz

7. Cluster results: No

8. Solution method: Mode superposition

9. Constant damping ratio: 1%

10. Spatial resolution: Maximum

11. Transcript of results:

The following data was obtained for the middle antenna. Graphs are also plotted.

Maximum

normal stress

in Y (Pa)

% of

yield

stress

Maximum

directional

deformation in

Y (mm)

Frequency

(Hz)

Phase angle

(maximum

stress) (o)

Longitudinal 1.59e7 5.76 3.5 152.5 89.31

Lateral 7.90e6 2.86 0.17 154.4 106.21

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Table 7: Harmonic response of middle antenna

12. Screenshots of graphs obtained:

Figure 20: Normal stress (Y) vs. frequency (longitudinal loading)

Figure 21: Y deformation vs. frequency (longitudinal loading)

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Figure 22: Normal stress (Y) vs. frequency (lateral loading)

Figure 23: Y deformation vs. frequency (lateral loading)

13. Interpretation of results:

A peak is observed in each of the graphs at a frequency near about the first mode of

the structure and thus, it is consistent with the modal analysis. Stresses are far less

than the failure values and deformation in lateral mode is less. Deformation in

longitudinal mode is slightly higher, but since it is the displacement of the antenna tip,

it is acceptable.

14. Conclusion:

Within the frequency range (10 – 100 Hz) specified by the Interface Control

Document of the LVI, the satellite does not undergo failure in any conceivable mode.

We can say that the structure will survive sinusoidal loading.

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Acknowledgements

We thank Department of Aerospace engineering, IIT Bombay for all the support that

they have provided. We would like to thank Prof Mujumdar for his immense help in all

kinds of problems. We would also like to thank Mr. Mandar Kulkarni for his help in

structures lab. Last we would like to thank entire satellite team for all the support that

they have provided in all scenarios.

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PRATHAM IIT BOMBAY STUDENT SATELLITE

Performance Verification Review

Thermals Subsystem

By

Pratham team

Department of Aerospace Engineering,

Indian Institute of Technology, Bombay

May, 2012

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Contents 1. INTRODUCTION . ................................................................................................................................ 4

2. Requirements & Constraint Analysis . ............................................................................................... 5

2.1 Requirements from Power Sub-System to Thermals Sub-System . .................................................. 5

2.2 Requirements from Communication and Ground Station Sub-System to Thermal subsystem . ...... 5

2.3 Requirements from On Board Computer Sub-System to Thermals Sub-System .............................. 5

3. Thermal Modeling . ............................................................................................................................. 6

3.1 Geometry . ........................................................................................................................................ 6

3.1.1 Satellite body . ............................................................................................................................. 6

3.1.2 Solar Panels. ................................................................................................................................ 7

3.1.3 Multilayer Insulation . ................................................................................................................. 8

3.1.4 Battery Box . ................................................................................................................................ 8

3.1.5 GPS . ............................................................................................................................................ 8

3.1.6 Magnetorquer . ............................................................................................................................ 9

3.1.7 Magnetometer ........................................................................................................................... 9

3.1.8 Printed Circuit Boards (PCB) . .................................................................................................... 10

3.1.9 Monopole . ................................................................................................................................ 10

3.2 Grid . ................................................................................................................................................ 11

3.3 Boundary conditions and Couplings ............................................................................................. 11

3.3.1 Solar Fluxes . ............................................................................................................................. 11

3.3.2 Heat dissipation from PCBs . . ................................................................................................... 12

3.3.3 Thermal coupling between solar panel and satellite sides . ................................................. 13

3.3.4 Thermal coupling between MLI and satellite sides .............................................................. 13

3.3.5 Radiation coupling between MLI and solar panels . .............................................................. 14

3.3.6 Conduction coupling between the satellite sides. ................................................................. 14

3.3.7 Conduction Coupling between PCB and satellite sides . ....................................................... 15

3.3.8 Conduction coupling between PCB and electrical components . ......................................... 15

3.3.9 Conduction coupling between magnetorquer surfaces . ...................................................... 16

3.3.10 Other couplings ..................................................................................................................... 16

3.4 Internal radiation ........................................................................................................................... 16

3.5 Transient Analysis.......................................................................................................................... 16

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4. Results . ............................................................................................................................................. 18

List of Figures

Figure 1: Magnetorquer . ............................................................................................................................ 9

Figure 2: Magnetorquer attached to side . ................................................................................................ 9

Figure 3: Pratham in Orbit ....................................................................................................................... 12

Figure 4: Side of Pratham ........................................................................................................................ 14

Figure 5: Final view of Pratham (outside) in IDEAS. ............................................................................. 17

Figure 6: Final View of Pratham (inside) in IDEAS . .............................................................................. 17

List of Tables

Table 1: Heat dissipations ....................................................................................................................... 13

Table 2: Panel Connections...................................................................................................................... 15

4 | P a g e

1. INTRODUCTION

The environment up to a height of around 500km above the Earth is strongly affected by

the presence of the denser layers of the atmosphere and by interaction with the

hydrosphere, and thus experience relatively mild conditions of temperature. However,

above 500km, the atmosphere becomes too thin to moderate the ambient thermal

conditions and consequently, the environment above such a height is prone to extremes of

temperature, being directly exposed to solar radiation and deep space.

Spacecraft, which are primarily intended to perform scientific missions which require them

to be positioned far away from the Earth’s surface, spend their entire lifetime in the harsh

environments of the upper atmosphere or interplanetary space. Being electromechanical

systems, however, these spacecraft also have certain requirements of temperature ranges

outside which their systems cannot function viably. Hence, every such spacecraft must

incorporate a thermal subsystem to achieve an optimal thermal environment for its smooth

functioning, through either active or passive means.

The requirements of thermal control and the challenges involved in realizing these in case

of Nano satellites in Low Earth Orbit (LEO) are unique. The IIT Bombay Student Satellite,

PRATHAM, belongs to this class of satellites, having a mass of less than 10kg and orbiting in

a LEO of approximate altitude 817 km. The Thermals Subsystem of PRATHAM aims to

maintain a cycle of temperature within the satellite which has a narrow range and lies

around the normal terrestrial temperatures, minimize spatial gradients of temperature,

dissipate excess heat generated by other subsystems and maintain sensitive components

within their specified functional ranges of temperature.

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2. Requirements & Constraint Analysis

In this part we will discuss the requirements of other sub-systems which need to be

addressed by the thermals subsystem.

2.1 Requirements from Power Sub-System to Thermals Sub-System

Thermal sub-system shall protect the solar panels from heating above 70O C

The optimal operating range of the battery is from 5o to 20o C. The acceptable

operating range is 0o to 30o C. The thermals sub system shall try and maintain the

battery within its optimal operating range. If that is not possible the battery must be

maintained in the acceptable operating range.

They shall remove the excessive heat from Power Circuits, since the temperature

range of the components (industrial grade) is -40o to +85o C

2.2 Requirements from Communication and Ground Station Sub-System to Thermal

subsystem

Thermal sub-system will protect the monopoles from heating above 100o C

They shall remove the excessive heat from the 2 monopole circuits, since the

temperature range of the components (industrial grade) is -40o to 85oC

2.3 Requirements from On Board Computer Sub-System to Thermals Sub-System

They shall remove the excessive heat from the OBC circuits as the temperature

range of the components(industrial grade) -40o C +85o C

The aim of the thermal subsystem is to satisfy all these requirements. The most stringent

requirement is maintaining the temperature of the battery. In the next chapter we will look

into the design procedure and performing the thermal analysis of the satellite

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3. Thermal Modeling

This chapter deals with the thermal modeling and simulation of the satellite, which would

help us in determining the temperature of the various parts of the satellite. The thermal

modeling and analysis has been done using the I-DEAS TMG thermal analysis software.This

software makes it easy to model nonlinear and transient heat transfer processes including

conduction and radiation. This software has a module for orbit and attitude modeling

which would help us in determining the solar fluxes acting on the sides of the satellite.I-

deas TMG uses finite element models (FEM) and highly accurate numerical methods for

solving the element based thermal models. ISRO has been using this software consistently

in the thermal analysis of their satellites. Therefore, the various models that are being used

in this analysis have been already validated. The thermals engineer in ISAC associated with

small satellites, Mr.Nagaraju helped us in the thermal modeling. The modeling involves

specifying the geometric model of the various parts of the satellite, couplings and boundary

conditions associated with it and the solution parameters required for solving the problem.

The whole modeling process is explained in this chapter.

3.1 Geometry

3.1.1 Satellite body

The Pratham satellite is a cube with each edge having a length of 253 mm. The satellite has

six faces with the following terminology:

Side facing the Sun – Sunside

Side opposite to the sun facing side – ASS

Side facing the earth – Nadir

Side opposite to Nadir – Zenith

Side whose normal is aligned with the velocity vector – Leading

Side opposite leading – Lagging

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There are various element properties that are available within I-deas which are used to

specify the geometry like the shell, solid and beam properties. In our analysis, we have used

only shell and solid element properties. All the satellite faces are modeled using the shell

element property. Shell property implies that the face is modeled as a two dimensional

plate with a very small thickness. In thermal point of view, the temperature gradient will be

only along the plane of the face and there would not be any temperature gradient along the

thickness of the face. The thickness of the face is specified as an input, so that volume is

calculated exactly, which will result in accurate calculation of the mass and thermal

capacity. The sides have a thickness of 3mm. A separate element property was applied on

both sides of the satellite body for the purpose of defining the radiation boundary

condition.

The faces of the satellite were not connected with each other in the model. Instead of

equivalency and keeping common nodes for edges of faces which are connected to each

other, we kept a distance of 1mm between the edges and two separate nodes were created

to the these edges for purpose of applying the coupling condition between them

3.1.2 Solar Panels

Solar panels are present on four faces of the satellite. Three solar panels present on leading,

lagging and Sunside faces of the satellite have a length of 260 mm and a width of 219 mm.

The other solar panel present on the zenith surface has a length of 260 mm and a width of

164 mm. A distance of 12 mm must be kept between the solar panel and the satellite body.

This space will be used for applying the Multilayer insulation (MLI). The solar panel is also

modeled as a shell element with a thickness of 3mm. It is attached to the satellite faces by a

steel screw of 3mm diameter at the four corner points of the solar panels. Therefore a

conduction coupling condition is applied between the solar panels and the satellite faces at

these points. Again, a separate element property must be applied on both sides of the solar

panels for the purpose of radiation boundary condition

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3.1.3 Multilayer Insulation

The empty space between the solar panels and the satellite face is filled with multilayer

insulation (MLI). MLI is modeled as a shell element with a thickness of 10 mm covering the

entire area of the satellite sides. Here, a conduction coupling is applied between the entire

face of the satellite face with entire face of the MLI. Conduction coupling is not applied

between solar panel and MLI, as the MLI is not in contact with the solar panels. here.

Nevertheless, a radiation coupling is applied between MLI and solar panel.

3.1.4 Battery Box

The battery box present in the Nadir side consists of a hollow box with a dimension of 80 x

74 x 42 mm. The box is filled with a six batteries inside. Therefore, the best way to model

the battery box is to model it as a solid element. In thermal point of view, the solid property

element implies that the temperature gradient is present in all three dimensions of the

battery box. Although the thermal property of Al T-6061 will be specified for the battery

box, it would not be accurate as the volume inside the box is filled with batteries and

should ideally include the thermal properties of battery. But it is safe enough to assume the

battery box to be solid cubical box with properties of Al T-6061. All the outer faces of the

battery box have a separate element property for the purpose of radiation boundary

condition as well as conduction coupling for the face attached to the nadir side. A thermal

insulation of 5mm thickness is applied between the battery box and nadir side. But, the

battery box is assumed to be in direct contact with the zenith side and thermal insulation is

modeled as a coupling between the nadir side and the battery box and is attached using

stubs which have a hole through which screw is inserted. The conduction coupling between

the respective face of the box and the projected area of the nadir side is applied. The

conduction through the stub and the screw is ignored.

3.1.5 GPS

The GPS present in the zenith side consists of a box with a dimension of 125 x 55 x 40 mm.

The box consists of circuits and electronics inside. Therefore, the best way to model the

GPS is to model it as a solid element. The mass of the GPS is calculated and the

corresponding density is estimated. The thermal properties for the GPS are defined in later

sections. It is not in direct contact with zenith, but is attached to the zenith using screws

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inserted through aluminium stubs. These stubs and screws are modeled as conduction

coupling between GPS surface and zenith.

3.1.6 Magnetorquer

Magnetorquer is present in the zenith, leading and Sunside faces of satellite. The figure

shows the magnetorquer configuration. It is in direct contact with the satellite sides. The

magnetorquer is modeled by creating a shell element from the given shape of torquer with

a thickness of 2mm. Then, the shell element of the same shape is created at a distance of 15

mm. The aluminium strip which joins these two surfaces is modeled by conduction

coupling, which is applied between these two surfaces, the details of which will be

explained in later sections. In addition to this, a coupling between the satellite sides and the

surface of torquer is applied.

Figure 1: Magnetorquer

Figure 2: Magnetorquer attached to side

3.1.7 Magnetometer

The modeling of the magnetometer is similar to that of GPS. It has dimension of 107 x 38 x

22 mm. It is modeled as a solid element. Thermal properties and density is specified in later

sections. It is not in direct contact with lagging side, but is attached to the lagging side using

screws inserted through aluminium stubs.

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3.1.8 Printed Circuit Boards (PCB)

There are total of six PCBs present namely, the telemetry, beacon, uplink, power, OBC and

the sun sensor board. All these boards have similar geometric configuration. These PCBs

have a thickness of 1.75 mm and are attached to the satellite sides using screws inserted

through stubs having dimensions of 10 x 10 x10 mm. The PCBs have electrical components

that dissipate heat. The PCB is modeled as a shell element with a thickness of 1.75 mm. The

stubs and screws are modeled by conduction coupling. The components are modeled as a

solid element and volumetric heat generation boundary condition is applied to the

electrical components. The components are in direct contact with the PCB and conduction

coupling with appropriate contact resistance was applied. Like every other part, separate

elements are defined on both sides of the PCBs for the purpose of radiation boundary

conditions.

3.1.9 Monopole

Three monopoles are present on the antisunside face of the satellite. The monopole has a

holder, which is placed on a square base, which is attached to antisunside using screws.

The monopole is modeled as a cylindrical solid element with 7 mm diameter. The holder is

also modeled as a solid element with an inner diameter of 8mm and an outer diameter of

20mm. The 1mm gap between the monopole and the holder is filled with heat shrinkable

tubes. These tubes are modeled by conduction coupling applied between holder and

monopole. The square base is modeled by conduction coupling between the holder and the

antisunside. Separate elements are defined on outer side of the monopole and holder for

the purpose of radiation boundary conditions.

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3.2 Grid

The model has been meshed using a structured grid. Hex8 elements were used. The mesh

sizes were different for each application region and has been built considering the

computation time without sacrificing on accuracy of the solution

3.3 Boundary conditions and Couplings

Boundary Conditions define known thermal conditions in the model. The Thermal Boundary condition form sets up a variety of heat loads and fluxes, fixed temperatures, and joule heating. Thermal coupling are applied to specify conductance between two dissimilar meshes or to specify conductance between two surfaces that are connected through screws or stubs. In this section, we explain in detail all the boundary conditions and the coupling applied in the model

3.3.1 Solar Fluxes

We used Orbit/Attitude Modeling to request solar, albedo, and view factor calculations for

selected elements within a defined orbit. Heat loads resulting from direct solar flux, albedo

and planet IR are automatically modeled, using a radiosity approach to determine the

reflection and absorption of the incident flux throughout the model. Eclipses are detected

and modeled accurately. The satellite orientation can also be modeled. We had a preview of

the defined orbit prior to solving the model. The steps undergone in modeling these are

1. Building a large radiative enclosure around the model. This creates a non-geometric element modeling the ambient environment for radiation simulation.

2. Defining the elements that are subjected to radiation. These include the solar panels, the part of the MLI applied on satellite body which is open to space, monopole and its holder and LVI.

3. Setting up the satellite orbit and orientation. Our satellite has a 10.30 sun synchronous orbit at an altitude of 817 km from earth. The velocity vector is from lagging to leading. These input data were needed for modeling the orbit.

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Figure 3: Pratham in Orbit

3.3.2 Heat dissipation from PCBs

The PCBs are major source of heat dissipation in the satellite. The electrical components in

the PCB generate heat. The components are modeled as solid element. Therefore, all

components that have heat dissipations above 0.1 watt were defined. We have the average

as well as maximum dissipation values. The time period for which the dissipation is

maximum, is also known. Considering these facts, a heat generation boundary condition

was applied to these components. A Heat Generation boundary condition defines the heat

flux or heat load per unit volume. The heat flux values were specified such that the

maximum heat generation value is present for the given time period and for the remaining

time, the heat generation values were given such that average heat generation becomes

equal to the given average heat dissipation. These are heat dissipation values for different

PCBs

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PCB Average heat

dissipation(W) Maximum heat dissipation(W) Tmax(s)

Telemetry 0.165 1 420

Beacon 0.35 1 600

Uplink 0.1 0.1 Full orbit

Power

0.7W over 4 DC-DC convertors, 0.4W in 4 output diodes, 0.16W in 4 input diodes

3 W total theoretical peak over 4 DC-DC convertors, Practical 1.5 W, peak of input and output diodes is 1.8 W 1000

OBC 0.4W for 2muc and 0.15W for 3ICs

0.4W for 2muc and 0.15W for 3ICs Full orbit

Sun Sensor 0.33 0.33 Full orbit

Table 1: Heat dissipations

3.3.3 Thermal coupling between solar panel and satellite sides

The solar panels are attached to the satellite sides using screws of 3mm diameter. The solar panels are at a distance of 12mm from the satellite sides. The conduction through the screws is modeled by coupling condition.We apply conductive coupling which creates conductance G = (K x A) / L from the primary elements to the nearest secondary elements, where L is the distance to the secondary element along the primary element's surface normal, which is 12mm in this case. The element mesh of the satellite sides should be prepared such that the element size has the approximate area of the screw. Then the element, that holds the approximate location of the screw should be selected as the primary element and the element in solar panel, nearest to the screw should be selected as the secondary element. The area A is calculated from the area of the primary element and conductivity, K of the steel is specified.

3.3.4 Thermal coupling between MLI and satellite sides

The 10 mm gap between the satellite side and solar panel is filled with MLI. The MLI is modeled by a shell element with 10mm thickness, which is placed at an equal distance from panel and side. Here, the conduction coupling is specified between the entire surface area of the MLI and the entire surface area of the satellite sides. .We apply conductive coupling which creates a conductance G = (K x A) / L from the primary elements to the nearest secondary elements. All the elements of the satellite sides are specified as primary elements and all the elements of the MLI are specified as the secondary elements. The

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conductivity of MLI is equal to 0.04 and is specified as the value of K in the conductance term.

3.3.5 Radiation coupling between MLI and solar panels

It creates radiative conductance between the primary and secondary elements. The magnitude of conductance is equal to:

· GBVF · 1 · A1 (T12+ T2

2)(T1 + T2)

is the Stefan-Boltzmannconstant. GBVF is the specified gray body view factor,

1 is the emissivity of the primary elements, A1 is the area of the primary elements, T1 is the absolute temperature of the primary elements, and T2 is the absolute temperature of the secondary elements.

The gray body view factor is the net amount of radiation emitted by the primary elements and absorbed by the secondary elements including all intermediate reflections. The separate element defined on reverse side of the solar panel for the purpose of radiation is selected as the primary element. The element on the front side of the MLI is selected as the secondary element.

3.3.6Conduction coupling between the satellite sides

Each body surface of the satellite has a flange as shown in the figure which is attached to other sides of the satellite using screws. The material of the flange is also aluminum substrate.

Figure 4: Side of Pratham

260 mm length

Flange

20 mm width

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Constant h conductive coupling is specified, which creates a conductanceG = h x A from

each PrimaryElement to the nearest Secondary Elements .Here the conduction is

contributed by the contact resistance and the conduction through the 3mm thickness of the

flange. It has been observed that thermal contact resistance is very high compared to

thermal resistance of the aluminium flange. Therefore, only the contact resistance is

defined by specifying the value of h. The thermal resistance term is given by

Thermal resistance = 1/hA + L/KA

1/hA is the contact resistance and h is taken to be equal to 300 W/m2K. This value is determined experimentally and the same value of h will be used for all contact resistances. Let us compare these values. 1/h is equal to 0.0033 and L/k for 3mm thickness is equal to 1.9354 x 10-5.It is clearly evident that the thermal resistance due to aluminium substrate can be neglected. The description of the flange attachment to the satellite sides are given below

The elements in the edge where the flange is located are specified as the primary element and the elements in the edge of its attached side, which is given in the second column in the above table, are specified as the secondary elements.

3.3.7Conduction Coupling between PCB and satellite sides

The PCB is attached to the satellite sides using screws inserted through the stubs. Again, the major contributor of thermal resistance is contact resistance. But here, thermal resistance will be equal to 2/hA because there are two contact interfaces, one between the stub and PCB and the other between stub and satellite side. Constant h conductive coupling is specified. Therefore the value of h specified is 150.The elements having the approximate location and area of the stubs at the satellite sides are specified as the primary element and the corresponding elements in the PCB are specified as the secondary elements.

3.3.8Conduction coupling between PCB and electrical components

The electrical components are attached to the PCBs by soldering which produces a thermal resistance such that the value of h becomes 1000 W/m2K. Then, a constant h conductive

Flanges of Connected to Nadir Leading, Lagging Zenith Leading Lagging

Leading Sunside, Anti-sunside Lagging Sunside, Anti-sunside Sunside Zenith, Nadir

Anti-Sunside Zenith, Nadir

Table 2: Panel Connections

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coupling is specified in way similar to that of other conductive couplings. This value of h can be varied by changing the method of soldering.

3.3.9Conduction coupling between magnetorquer surfaces

The two surfaces of the torque were modeled as shell elements at a distance of 15mm. The thin strip that connects these two surfaces is modeled by coupling. We apply conductive coupling which creates a conductanceG = (K x A) / L from the primary elements to the nearest secondary elements, where L is automatically calculated, which is 15mm here. The element meshes in the surface is built such that element area is equal to the area of the strip. The conductivity of aluminium is specified as the value of K.

3.3.10Other couplings

The coupling between the monopole and the holder was specified similar to the coupling between MLI and satellite side. The coupling between the monopole holder and the base was specified similar to the coupling between PCB and the stub .The coupling for the parts that are in direct contact with the satellite sides like the battery box and torquer is specified using the constant h conductance where the value of h is 300 W/m2K.

3.4 Internal radiation

All the surfaces inside the satellite radiate among each other due to the virtue of its temperature. Internal radiation is modeled by defining radiation coupling between two surfaces. The method behind the calculation of the radiation coupling was mentioned here. Here, all the surfaces within the satellite that can contribute to internal radiation were specified from which all possible dual combination of surfaces are chosen to which the radiation coupling is applied. All the above boundary conditions and coupling have been consistently used in earlier small satellite modeling and have been validated by Prof Nagaraju

3.5 Transient Analysis

For any transient simulation, we must define the time span for the solution. In addition, we should also specify the Time Step, which is the time mesh for the solution. A smaller time step will give more accurate results at a cost of increased computation time. A time span of 25 orbits was specified as it will take about 15 to 20 orbits for the temperatures to get stabilized.

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The final model of the satellite in IDEAS looks like

Figure 5: Final view of Pratham (outside) in IDEAS

Figure 6: Final View of Pratham (inside) in IDEAS

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4. Results

The simulations were carried out based on the modeling described above. After the first

round of simulations, temperatures on the various parts of the satellites were observed.

After analyzing the temperature of the various parts, it was decided as to which all parts

require thermal treatment. Thus, on the basis of the temperatures, we decided that the

thermal design that will be applied on the flight model. All the decisions related to thermal

design have been stated below

All the packages (including battery) which are mounted inside the spacecraft will be

covered with black paint.

All spacecraft panels will be black anodized from inside and covered with MLI

blanket (15 layers) from outside expect OSR window. The total OSR requirement for

the satellite is 8 cm2 on the Anti-sun side panel opposite to high heat dissipating

chips on the Telemetry and Beacon PCB.

All High dissipation components will be placed on the PCB using thermal filler

materials( Sty cast or Gap pads)

High dissipated chips (which dissipate~1W) will be placed on the Telemetry and

Beacon PCB using thermal filler materials along with heat sink size of 25*25*3MM.

Telemetry and Beacon PCB Heat sinks will be covered with black tape.

All solar panels will be isolated from the spacecraft panels conductively and

radiatively.

All solar panels back side will be covered with low emittance tape.

Interface ring will be polished from all sides

Payload antenna consists of one holder and one monopole. Total three antennas are

mounted on the spacecraft anti sun side panel. Monopole holders will be covered

with MLI blanket and all monopoles will be polished and left as it is.

The heater of the battery box has been removed as the simulations show that the

temperature of the battery box is comfortably within the specified ranges.

The next simulations were performed by incorporating the above design in the model.Two

cases have been considered.

Case 1: Considered all solar panels with all strings connected (GA-AS cells with 26%

efficiency).

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Case 2: Considered all solar panels with all strings not connected (GA-AS cells with 0%

efficiency).

Case 2 represents the worst possible case as all the heat that is incident upon the solar

panels is assumed to be converted into heat energy, i.e. , the heat radiated by the solar

panel to the MLI will be Total Solar Flux*absorptivity of panel .

The temperature of the Telemetry chip, beacon chip, DC-DC converters and the diodes are

so high because the amount of heat dissipated by them is high as mentioned in Table 1.

So based on the results that have been obtained above the following conclusions have been

made.

Solar panels have high temperature fluctuations due to presence of the highly

fluctuating solar flux incident on it

Solar panels have been isolated from the satellite sides both conductively and

radiatively, so that the fluctuation in solar panels doesn’t get reflected in the satellite

body.

Optimum position of the monopole in the Antisunside where the solar flux variation

is also less and the flux values also have less magnitude.

Very less temperature variation for internal components including PCBs

Satellite body temperature has less variation over the orbit

Battery box with insulation has temperatures within the range of operating

temperature of battery

Electrical components in Power and OBC boards have high maximum values of

temperatures

Presence of heat sink has lowered the temperature of power amplifier in the

telemetry and the beacon board

So, it has been concluded that all components are within their required temperature

range.