Thermal Analysis of Passive Radiators for Interplanetary Space Applications

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THERMAL ANALYSIS OF A PASSIVE RADIATOR

FOR

INTER-PLANETARY SPACE APPLICATIONS

A PROJECT REPORT

Submitted by

In fulfillment for the award of the degree

of

BACHELOR OF ENGINEERING

in

Mechanical Engineering

S.N. PATEL INSTITUTE OF TECHNOLOGY & RESEARCH

CENTRE,

Umrakh, Bardoli

Gujarat Technological University, Ahmedabad

2015-16

Dave Yash Jayeshbhai (120490119029)

Patel Dipakkumar Sureshbhai (120490119061)

Modi Bijankkumar Krishnakant (120490119067)

Rajput Shailesh Santosh Kumar Singh (120490119093)

Dorik Abhishek Deepak (120490119094)

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S.N. PATEL INSTITUTE OF TECHNOLOGY AND RESEARCH

CENTRE,

UMARAKH, BARDOLI.

MECHANICAL ENGINEERING

2015-16

CERTIFICATE

Date: 31/05/2016 This is to certify that the dissertation entitled “THERMAL ANALYSIS OF

A PASSIVE RADIATOR FOR INTER-PLANETARY SPACE

APPLICATIONS” has been carried out by:





Dorik Abhishek Deepak (120490119094) Under our guidance in fulfillment of the degree of Bachelor of Mechanical

Engineering (8th Semester) of Gujarat Technological University, Ahmedabad

during the academic year 2016.

Guides-: Head of the Department:

____________________ ____________________

Mr. Harshal T. Shukla Dr. Piyush S. Jain

(Asst. Prof.)

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EXTERNAL EXAMINAR APPROVAL

This is to certify that the project work embodied in this report entitled

“THERMAL ANALYSIS OF A PASSIVE RADIATOR FOR INTER-

PLANETARY SPACE APPLICATIONS” was carried out by:





Dorik Abhishek Deepak (120490119094) at SITARAMBHAI NARANJI PATEL INSTITUTE OF

TECHNOLOGY AND RESEARCH CENTRE, UMARAKH (049) is

approved for award of the degree of B.E. Mechanical Engineering by

Gujarat Technological University.

Date: 31/05/2016 Place:

Examiner(s): 1.

2.

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Acknowledgement

The project associates wish to thank all of the authors whose collective insights from their

research papers and publications have made this study possible; David G. Gilmore the

editor of Spacecraft Thermal Control Handbook Volume I, Jose I. Rodriguez, Howard

Tseng and Burt Zhang from Jet Propulsion Laboratory of California Institute of Technology

for their research paper on Thermal Control System of the Moon Mineralogy Mapper

Instrument.

We would like to thank our project guide Prof. Harshal Shukla [Project Guide] for

enlightening the work and his constant encouragement despite of oncoming hurdles

throughout the project and making this work possible.

I express my gratitude to Dr. Piyush Jain [HOD MECH.] for co-operation and support

and also thank all the people who have contributed in their own way in making this project

successful.

Yash Dave

Dipak Patel

Bijank Modi

Shailesh Rajput

Abhishek Dorik

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Abstract

An elaborate growth has been observed in the use of Satellites for various domestic,

military and navigational applications. Satellites carry various Infrared instruments and

Electronic Packages in them collectively called Payloads. The Payload can function

properly only if it is maintained within specified temperature ranges. The Thermal Control

System (TCS) of a Satellite keeps the equipment temperature within the specified operating

range. It is broadly divided into two classes namely, Passive Thermal Control System

(PTCS) and Active Thermal Control System (ATCS).

The current study aims to appraise the merits of using Passive Radiators for Interplanetary

Space Applications as it draws no power from the satellite system, and measuring its

Effectiveness in Dissipating the heat developed inside the payload to space against

Environmental Backloads incident over its surface from the Celestial Surroundings. It

maintains the desired temperature range by Controlling Conductive and Radiative Heat

Paths through the selection of Geometrical Configurations and Thermo-Optical Properties

of the surface in addition to savings in Mass and Power respectively which has always been

a crucial element in spacecraft design and configuration. A Parametric study is conducted

to explore the scopes of using Passive Radiators. The entire system is Modelled and

Simulated in FEA software UG NX 7.5 with a Flat Plate Radiator used in the initial Space

Thermal Analysis.

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List of Figures

2.1 Radiator energy balance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5

2.2 Satellite thermal environment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6

2.3 Solar Spectral Distribution . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7

2.4 Ideal representation of four basic passive-control surfaces . . . . . . . . . . . . . . . . . . . 10

2.5 Second-surface mirror thermal finish . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

4.1 Package Model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17

4.2 Dissipator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18

4.3 Thermal Strap . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .18

4.4 Package and Radiator with MLI. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18

5.1 Boundary conditions for Simulation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20

5.2 Orbital Heating Parameters for Simulation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21

5.3 Case 1-Radiator. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22

5.4 Case 1-Dissipator. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22

5.5 Case 1-Thermal Strap . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22

5.6 Case 1-Package . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22

5.7 Case 3-Radiator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .23

5.8 Case 3-Dissipator. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 23

5.9 Case 3-Thermal Strap . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24

5.10 Case 3-Package . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24

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List of Tables

2.1 Thermo-optical properties of common surfaces. . . . . . . . . . . . . . . . . . . . . . . . . . . .12

3.1 Spacecraft Specification for case study. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14

4.1 Specification of satellite package model. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17

5.1 Boundary conditions for Simulation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19

5.2 Orbital Heating Parameters for Simulation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20

5.3 Time Interval for Transient Condition. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21

5.4 Case 1 Result. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22

5.5 Case 2 Result. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 23

5.6 Case 3 Result. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 23

5.7 Case 4 Result. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24

5.8 Case 5 Result. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24

5.9 Case 6 Result. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25

5.10 Case 7 Result. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25

5.11 Case 8 Result. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25

5.12 Case 9 Result. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26

5.13 Case 10 Result. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26

5.14 Case 11 Result. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26

5.15 Case 12 Result. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27

5.16 Case 13 Result. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27

5.17 Case 14 Result. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27

5.18 Case 15 Result. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28

5.19 Case 16 Result. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28

5.20 Case 17 Result. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28

5.21 Case 18 Result. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29

5.22 Case 19 Result. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29

5.23 Case 20 Result. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29

5.24 Case 21 Result. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30

5.25 Case 22 Result. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30

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Nomenclature

TCS Thermal Control System

PTCS Passive Thermal Control System

ATCS Active Thermal Control System

AFT Allowable Flight Temperature

FEA Finite Element Analysis

FEM Finite Element Method

IR Infrared

Albedo Sun-light reflected off planet

MLI Multi-layer Insulation

SSM Secondary Surface Mirrors

OSR Optical Solar Reflector

M3 Moon Mineralogy Mapper

BOL Beginning of Life

EOL End of Life

�̇� Radiation power of Radiator

A Surface Area

Ε Emittance

Ε Stefan-Boltzmann Constant

T Temperature

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Contents

Acknowledgements iv

Abstract v

List of Figures vi

List of Tables vii

Nomenclature viii

1. Introduction 1

1.1 Project Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2

1.2 Objective. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3

1.2.1 Problem Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3

2 Literature Review 4

2.1 Radiators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

2.2 Spacecraft Thermal Environment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5

2.3 Thermal Control Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7

2.3.1 Passive Thermal Control System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

2.3.2 Active Thermal Control System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

2.4 Thermal Surface Finishes. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

2.4.1 Common Thermal Surface Finish. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

2.5 Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

3 Case Study to Design Flat-Plate Radiator 13

3.1 Case Study of 3-axis stabilized Spacecraft. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13

3.1.1 Spacecraft Specification. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .14

3.2 Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15

4 Modeling & FEM of Flat-Plate Radiator 16

4.1 Modeling of Satellite Package. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16

4.1.1 Specification of Satellite Package model. . . . . . . . . . . . . . . . . . . . . . . . . . 16

4.2 FEM of Satellite Package. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17

5 Simulation & Results 19

5.1 Simulation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19

5.1.1 Boundary Conditions for Simulation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19

5.1.2 Orbital Heating Parameters for Simulation . . . . . . . . . . . . . . . . . . . . . . . .20

x

5.1.3 Time Interval for Transient Condition. . . . . . . . . . . . . . . . . . . . . . . . . . . . 21

5.2 Results. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21

5.2.1 Radiator Dimension – 300x150x2mm . . . . . . . . . . . . . . . . . . . . . . . . . . . 22






6 Inferences & Future Work 31

6.1 Inferences. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31

6.2 Future Work. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32

References 33

Appendix I 34

Appendix II 35

Appendix III 38

Appendix IV 45

1

Introduction Chapter 1

This chapter gives brief introduction to satellite system, satellite configurations, and

thermal control subsystems along with its activities. It puts emphasis on importance of

satellite in communication system. The chapter contains overview and objective of the

project.

A Satellite is an artificial object which orbits around a planet to collect useful information.

It consists of payload mounted on it to fulfill objectives of satellite mission. An instrument

working on a satellite to perform various tasks is called a Payload. The satellite serves as

foundation block in the structure of communication system. During the past 40 years,

hundreds of satellites have been built in support of scientific, military, and commercial

missions. Each can be broadly categorized as either three-axis-stabilized, spin stabilized or

pallets; these types are distinguished by their configuration, internal equipment and thermal

control designs.

The most common satellite configuration today is three-axis-stabilized. It is characterized

by a body which is roughly box-shaped and by deployable solar-array panels. As mentioned

earlier the satellite system consists of “payload” and “bus”. The bus consists of sub-system

that typically include thermal control subsystem (TCS). TCS consists of surface finishes,

insulation blankets, heaters and refrigerators.

Thermal control sub-system consists of following different activities:

Environment Interaction: It comprises the interaction of the external surfaces of the

satellite to the surrounding space. Either the surfaces need to be protected from the

environment or there has to be improved interaction. Two main goals of environment

interaction are the reduction or increase of absorbed environmental fluxes and reduction or

increase of heat loses to the environment.

Heat Collection: It involves the removal of dissipated heat from the equipment in which it

is created to avoid unwanted increase in the satellite’s temperature.

Heat Transfer: It includes taking the heat from heat collection device to a radiating device.

2

Heat Rejection: After the heat has been collected and transported, it has to be rejected at

an appropriate temperature to a heat sink, which is usually the surrounding space

environment. The rejection temperature depends on the amount of heat involved, the

temperature to be controlled and the temperature of the environment into which the device

radiates the heat.

Heat provision and storage: It is to maintain a desired temperature level where heat has

to be provided and suitable heat storage capability has to be foreseen.

1.1 Project Overview

In a satellite there are various sensors, computers, instruments and communicating devices

which are used for specific application, commonly defined as payload. During their

operation, massive heat is generated within the satellite package which in turn increases the

overall inside temperature of the satellite. Every payload has its own definite operational

temperature range inside which it functions efficiently called the allowable flight

temperature (AFT).

In order to maintain the temperature range, the excess amount of heat should be radiated.

For this, the heat is collected in a device called dissipater. The heat is then conducted to

radiator using thermal strap. The radiator is in direct contact to environment, which is the

surrounding space where all of the conducted heat is radiated. The whole dissipater and

thermal strap assembly are contained in a metallic package and the radiator is exposed to

the environment.

The software used for design and thermal analysis of the radiator-dissipater assembly is

UG-NX v7.5. First of all, a model of the assembly with calculated dimensions is created in

modeling section of the software. The model is then meshed for finite element analysis

(FEA) in the NX space system thermal-fem section. The meshed element in this part is

provided with appropriate material as well as thermo-optical properties. The prepared fem

of the assembly is then given the thermal loads and temperature constraints acting on it in

NX space system-simulation section. Here, all the working constraints are provided and a

result is generated, which includes detailed information of nodal and elemental

temperatures as well as elemental heat load and elemental heat flux. A real time simulation

considering all orbital condition is obtained at the end.

3

The obtained results of simulation are helpful in determining project objectives such as

shape, size and orientation of the radiator and the end conclusion.

1.2 Objective

1.2.1 Problem Definition

The payload of satellite operates in certain definite temperature range in which it performs

most efficiently. Thus, increase or decrease in inside temperature of the satellite outside of

AFT results in malfunction of the equipment.

In case if the inside temperature range is exceeded, in such situation the satellite is thrown

in survival mode, where all the instruments are shut off and the only function of the satellite

is to maintain its orbit, thereby securing the acquired data from any damage.

A satellite radiator is acted upon by three thermal radiation loads namely solar flux, albedo

flux and planet-shine flux. As a result, certain temperature constraint is developed on

radiator surface. It is difficult to radiate the heat generated by the payload inside the satellite

due to these incident thermal loads. In order to overcome this problem, the orientation of

the radiator can be adjusted such that view-factor is maximized while the incident radiations

are minimized.

The objective of project work is to maintain AFT using PTCS, thereby assuring maximum

efficiency of the payload and to maximize view-factor while simultaneously minimizing all

calculable heat radiations incident on it.

4

Literature Review Chapter 2

This chapter briefs the information fetched from research papers and publications and

detailed information of spacecraft thermal environment, thermal control systems and their

type and thermal surface finishes. At the end of the chapter the entire project work is

summarized.

The research papers used for literature survey are Thermal Control System of the Moon

Mineralogy Mapper Instrument by Jose I. Rodriguez, Howard Tseng and Burt Zhang from

Jet Propulsion Laboratory of California Institute of Technology, The Moon Mineralogy

Mapper Imaging spectrometer for lunar science Moon Spacecraft Thermal Control

Handbook by David G. Gilmore.

2.1 Radiators

The basic function of radiator is to radiate the heat generated inside the satellite to

surrounding atmosphere. They are of different forms, such as spacecraft structural panels,

flat plate radiators, radiators mounted on side of satellite or panels deployed after satellite

is in orbit. Despite different configuration, all radiators reject heat in form of infrared (IR)

radiation from their surfaces. The radiator power depends on surfaces emittance and

temperature. Therefore, most radiators are given surface finishes with high IR emittance to

maximize rejection and low solar absorptance to limit heat loads from the sun. E.g. quartz

mirrors, silvered or aluminized Teflon and white paint. The figure 2.1 illustrates the energy

balance between the environmental loads, internal energy of satellite and the radiated

energy.

The radiating power of a radiator is a strong function of temperature given as follows:

�̇� = 𝐴 ×∈× 𝜎 × 𝑇 (1)

Where A is surface area, ε is emittance, and σ is Stefan-Boltzmann constant (5.669 ×

10−8W/m2K4), and T is Temperature.

5

2.2 Spacecraft Thermal Environment

For a satellite, environmental heating plays a major role in the process of energy

management. The significant environmental heating availing in satellite orbit are direct sun-

light, sun-light reflected off planet (albedo), and IR energy emitted from the planet (refer

Figure 2.2). Apart from that, during launch or in exceptionally low orbits, friction causes

free molecular heating effect in rarefied upper atmosphere.

Solar flux:

The direct sunlight incident on the radiator surface is called solar flux. It is the most evident

source of environmental heating for most satellites. The IR energy emitted from sun is of

much shorter wavelength than that emitted by body at room temperature. This distinction

allows for selection of thermal control finishes that are very reflective in solar spectrum but

whose emissivity is high at room temperature. These finishes minimize solar loads while

maximizes satellites ability to reject heat.

The energy distribution is approximately 7% ultraviolet, 46% visible, and 47% near (short-

wavelength) IR, with the total integrated energy equal to the 1322 to 1414 W/m2 (refer

Figure 2.3).

Figure 2.1: Radiator energy balance [1]

6

Albedo flux:

Sunlight reflected off a planet or moon is called albedo. A planet’s albedo is usually

expressed as the fraction of incident sunlight that is reflected back to space, and is highly

variable. The albedo heat flux reaching a satellite will decrease as the satellite moves along

its orbit and away from sub-solar point (the point on planet where sun is on zenith i.e.

directly overhead). This geometric effect is accounted for by the analysis codes used to

perform space thermal analysis.

Planet Shine:

All incident sunlight not reflected as albedo is absorbed by planet or moon and eventually

re-emitted as IR energy known as planet shine. The intensity of IR energy emitted at any

given time from a particular point on planet or moon can vary depending on local

temperature of planet or moon and the amount of cloud cover.

Figure 2.2: Satellite thermal environment [1]

7

2.3 Thermal Control Systems

In satellite design, the thermal control system (TCS) has the function to keep all the satellite

parts within acceptable temperature ranges during all mission phases, sustaining

The external environment, which can vary in a wide range as the satellite is exposed to deep

space or to solar or planetary flux, and rejecting to space the internal heat dissipation of the

satellite itself.

The thermal control subsystem can be classified as passive thermal control system (PTCS)

and active thermal control system (ATCS).

Both work in following two ways:

1. Protecting the equipment from extreme hot temperatures, either by thermal insulation

from external heat fluxes such as Solar, Albedo and Planet Shine.

2. Protecting the equipment from extreme cold temperatures, by thermal insulation from

external sinks, by enhanced heat absorption from external sources, or by heat release

from internal sources.

Figure 2.3: Solar Spectral Distribution [1]

8

2.3.1 Passive Thermal Control System

PTCS works on the principle of passive thermal control (PTC), which is achieved by

control of conductive and radiative heat paths through selection of proper geometrical

configurations, insulation blankets, sun shields, radiating fins, thermo-optical properties,

thermal coatings, heat sinks and phase change material (PCM).

PTC focuses on heat gain control and heat dissipation to control thermal environment inside

a satellite, which means it works by preventing heat entering the interior (heat gain control)

and by natural heat removal from the package (natural cooling).

PTCS includes:

• Multi-layer Insulation (MLI)

• Coatings

• Thermal fillers at interfaces

• Mirrors (secondary surface mirrors-SSM, or optical solar reflectors-OSR)

• Radioisotope Heater Units (RHU).

2.3.2 Active Thermal Control System

An ATCS uses external forces using mechanical systems to control thermal environment,

more precisely mechanically pumped fluid in closed loop circuits to perform heat

collection, heat transportation and heat rejection. It is required to perform the heat rejection

function when the combination of environmental thermal loads and satellites internal heat

generation exceeds the capabilities of the PTCS to maintain temperatures.

ATCS employs cold plates and heat exchangers to remove large amount of heat, both of

which are cooled by a circulating ammonia loop. The heated ammonia circulates through

large radiators located on exterior of satellites, releasing the heat by radiation to space that

cools the ammonia as it flows through radiator.

ATCS includes:

• Thermostatically controlled resistive electric heaters

9

• Fluid loops to transfer the heat dissipated by equipment to the radiators

• Louvers

• Thermoelectric Cooler

2.4 Thermal Surface Finishes

For design consideration of a satellite, selection of thermal surface finishes plays a vital

role. According to various thermal control requirements, wavelength dependent coatings

are employed in satellite thermal design. For instance, solar reflectors, such as second

surface mirrors, white paints, and silver- or aluminum-backed Teflon, are used to minimize

absorbed solar energy, yet they emit energy almost as a blackbody would. To minimize

both the absorbed solar energy and infrared (IR) emission, polished metal such as aluminum

foil or gold plating is used. Black paint is commonly utilized on the interior of the vehicle,

to facilitate radiant heat transfer among internal components.

Thus the existing state of the art includes a rather wide variety of wavelength dependent

coatings. The problems of in-space stability, out gassing, and mechanical adhesion to the

substrate have all been resolved for most coatings.

The external surface of the satellite is in direct contact to environment and simultaneously

to sunlight, albedo and planet shine. Hence, to achieve an energy balance at the desired

temperature between satellite internal dissipation, external sources of heat, and re-radiation

to space, as illustrated in Figure 2.1, selection of radiative properties of surface coating is

important.

The two primary surface properties of importance are the IR emittance and the solar

absorptance. Sometimes, two or more coatings are combined in a checkerboard or stripe

pattern to obtain the desired combination of average absorptance and emittance.

Thermal-control surfaces can be categorized as solar reflector; solar absorber, flat reflector,

and flat absorber (refer Figure 2.4). The solar reflector reflects incident solar energy while

absorbing and emitting IR energy. Solar reflectors are characterized by a very low α/ε ratio.

Solar absorbers absorb solar energy while emitting only a small percentage of the IR

10

energy. Flat reflectors reflect energy throughout the spectral range (i.e. in both the solar

and IR regions), while flat absorbers absorb throughout the spectral range.

Radiators are typically coated with second-surface mirrors or white paint. The second

surface mirrors makes use of a visibly transparent material, such as quartz glass or Teflon,

to achieve a high emittance, along with a reflective silver or aluminum coating on the back

to minimize solar absorptance (refer Figure 2.5).

2.4.1 Common Thermal Surface Finish

Almost all visible surfaces on the inside and outside of unmanned spacecraft are thermal

control finishes; this reflects the fact that all physical objects absorb and emit thermal

energy in the form of radiation. The flow of heat resulting from absorption and emission

by these surfaces must be controlled in order to achieve a thermal balance at the desired

temperatures. The principal external surface finishes seen on most spacecraft are the outer

Figure 2.4: Ideal Representation of Four Basic Passive-Control Surfaces [1]

11

layer of insulation blankets, radiator coatings, and paints. Electronic boxes located inside

the spacecraft, and the structural panels to which they are attached, are usually painted to

achieve a high emittance. (While most paints have the required high emittance regardless

of color, black paints have been the conventional choice for internal applications.) Internal

temperature sensitive components that do not dissipate much heat, such as propellant lines

or tanks, often have a low-emittance finish of aluminium or gold. Common thermal finishes

and their optical properties are shown in Table 2.1

2.5 Summary

The whole project work can be briefly summarized as follows.

The satellite payload generates heat which is radiated with the help of passive thermal

control system. In passive thermal control system, Radiator is used to dissipate internal heat

of satellite to surrounding space. The function of the radiator of the satellite is to maintain

energy balance between environmental load, internal energy and radiated energy. To

achieve this balance, radiator material and surface coatings of required radiative properties

such as absorptance and emittance are selected.

Figure 2.5: Second-Surface Mirror Thermal Finish [1]

12

Table 2.1: Thermo-optical properties of common surfaces [1]

Surface BOL (Beginning of Life) EOL (End of Life)

α Ε α ε

Aluminum anodized (structures) 0.2 0.6 0.33 0.8 0.8 1

Aluminium (vessels) 0.15 0.2 0.75 0.1 0.1 1

Aluminized Kapton (al. inside) 0.4 0.8 0.5 0.4 0.8 0.5

Aluminized Kapton (al. outside) 0.15 0.05 3 0.15 0.05 3

Beta cloth 0.3 0.85 0.35 0.4 0.85 0.47

Black paint (insides) 0.95 0.9 1.06 0.9 0.9 1

GFRP (solar panels, structures) 0.85 0.85 1 0.85 0.85 1

Goldised Kapton (gold outside) 0.25 0.02 12 0.25 0.02 12

MLI (back aluminised Kapton) 0.3 0.6 0.5 0.6 0.6 1

OSR (radiators) 0.08 0.8 0.1 0.08 0.8 0.1

Silver paint (electrically cond.) 0.035 0.45 0.78 0.5 0.6 0.83

Solar cells 0.75 0.75 1 0.85 0.85 1

Titanium tiodize (apogee motor) 0.6 0.6 1 0.6 0.6 1

White paint (antenna) 0.2 0.85 0.24 0.6 0.85 0.71

Along with radiator, multi-layer insulation, surface coatings and mirrors or deflectors are

used to minimize the incident environmental load on radiators as much as possible. The

size and shape of the radiator package is calculated considering all availing constraints that

are applicable to the satellite package. Modeling and finite element analysis of the package

assembly is carried out using UG-NX software’s space system thermal component. Finally,

simulation, after providing all the possible conditions of a planetary orbit gives the final

thermal solution with exact temperature pattern of the whole assembly and end conclusion

is obtained.

13

Case Study to Design Flat-Plate Radiator Chapter 3

This chapter comprises of case study undertaken from one of the reference book which

provides data about sample specification of a 3-axis stabilized spacecraft. The solution of

heat balance equation gives the total surface area of a flat plate radiator which is required

to obtain specified thermal heat dissipation. Here the solution is obtained manually.

To get familiar with the design equations used for the passive radiator and to obtain sample

values of spacecraft specification, a case study from one of the reference book “Design of

Geosynchronous Satellites” by author Brij N. Agarwal for radiator design and Spacecraft

Thermal Control is considered. The case study gives sample values (Table 3.1) of a 3-axis

stabilized spacecraft viz. Total thermal dissipation, Allowable temperature range, Mass of

radiator including mountings, specific heat (Cp). The objective of the case study is to find

out the total surface area of a flat-plate radiator for various conditions such as hottest

condition, winter solstice at EOL and steady state condition because OSR is assumed to be

isothermal.

3.1 Case Study of 3-axis stabilized spacecraft

Heat balance equation for radiator is given as,

Heat radiated = Heat incident + Instrument heat dissipation

Which gives following mathematical formula:

𝜺𝝈𝑻𝟒𝜼𝑨 = 𝜶𝒔𝑨𝑺 𝐬𝐢𝐧 𝜽 + 𝑷 (2)

Making Area of radiator (A) subject of equation, above equation can be further simplified

as follows:

𝑨 = 𝜶𝒔𝑨𝑺𝐬𝐢 𝐧 𝜽+𝑷

𝜺𝝈𝑻𝟒𝜼 (3)

14

3.1.1 Spacecraft Specification

Table 3.1: Spacecraft Specification for Case Study [2]

Total Thermal Dissipation (Q) 300 W

Allowable Temperature Range

T1 = 278 K (5 ºC)

T2 = 310 K (37 ºC)

Mass of Radiator (incl. mounting) 85 kg

Specific Heat (Cp) 900 J/kg K

Emissivity (ε) 0.8

Solar Absorptance at EOL (αs) 0.21

Solar intensity at winter solstice

(S)

1397 W/m2

Solar aspect angle (θ) 23.5º

Efficiency (η) 0.9

Substituting the values from Table 3.1, the value of A can be calculated from above

equation which is,

𝑨 ≅ 𝟏. 𝟏𝟔 𝒎𝟐 (4)

15

3.2 Summary

The case study can be summarized as follows.

A set of sample specification were taken from one of the literature. Substituting different

values viz. Total Thermal Dissipation, Allowable Temperature Range, Specific Heat,

Emissivity, Solar Absorptance at EOL, Solar intensity at winter solstice, solar aspect angle

and Efficiency of the system in Heat balance equation, the surface area of flat-plate radiator

is calculated. The calculations are performed manually.

16

Modeling & FEM of Flat-Plate Radiator Chapter 4

This chapter includes the models and FEM of satellite package containing Dissipator

block, Thermal strap, and Flat-plate radiator enclosed in package body with MLI. It

contains the specification of materials and case study dimensions. Models and its FEM

with varying radiator dimensions for seven different cases are created for further

simulation.

A satellite package is shown in figure 4.1 consisting of electronic packages which are

connected to some other heat dissipating devices. In order to remove the heat generated by

these packages a flat plate radiator is kept which is exposed to space. All the other things

including packages and cables are insulated by MLI (Multi Layered Insulation). Also

heaters are provided with micro-controller chips in order to maintain the temperature. The

heater gets ON when it gets feedback signal from the sensors if the temperature limit goes

beyond 10 ºC and it gets OFF when the temperature reaches to 20 ºC. Figure 4.2 gives

inside view of the satellite package.

4.1 Modeling of Satellite Package

Heat generated in Dissipator block is transferred to thermal strap attached to the Dissipator

block (refer figure 4.2 and figure 4.3). The heat gets conducted through thermal strap to

flat-plate radiator where it is to be dissipated to surrounding environment. Surface finishes

are provided having high Infrared emittance to maximize rejection and low solar

absorptance to limit heat loads from the sun, thereby increasing radiating power of the

radiator.

4.1.1 Specification of Satellite Package Model

A package consisting of the Dissipator block (Heat dissipating device) which is attached to

a thermal strap (Heat carrying strip) responsible for transferring heat by that Dissipator

block to the radiator which is at the top of the package exposed to space. The whole package

except radiator is wrapped and insulated with MLI (Multi Layered Insulation). The Table

17

4.1 provides information on Dimension, Material and Thermo-Optical Properties of each

component of Satellite Package.

Figure 4.1: Package Model

Table 4.1: Specification of satellite package model [2]

Device Material Dimension (mm)

Thermo-optical Properties

IR

Properties

Solar

Properties

Radiator Aluminium-6061 300x150x2 ε = 0.7 α = 0.45

Dissipator

Block Stainless Steel 30x40x50 - -

Thermal Strap Copper - - -

Package Aluminium-6061 300x150x300 - -

Insulation MLI - ε = 0.85 α = 0.40

4.2 FEM of Satellite Package

The models of each component are divided into small segments which is called meshing of

the parts. The meshing is useful in finite element analysis and simulation of the parts to

obtain accurate temperature range of each of the parts when thermal loads are applied on

them.

18

Figure 4.2: Dissipator Figure 4.3: Thermal Strap

Figure 4.4: Package and Radiator with MLI

19

Simulation & Results Chapter 5

This chapter contains details of simulation performed for seven different cases by changing

dimensions of flat-plate radiator i.e. thickness and surface area to study relation between

dimensions of the radiator and heat dissipation occurrence. The results shows the

temperature range of each components, keeping same boundary conditions for a moon

orbit.

5.1 Simulation

The meshed components are provided with material properties, thermo-optical properties.

The simulations are performed by providing boundary conditions and orbital heating

parameters as mentioned in Table 5.1 and Table 5.2 respectively. The simulations are

performed for 3.75 W and 15 W heat dissipation for both Steady State condition and

Transient condition. For Steady State condition, the payload is assumed to dissipate heat

throughout the orbital period, while for transient condition it is assumed that heat

dissipation takes place in time interval as shown in Table 5.3 later.

5.1.1 Boundary Conditions for Simulation

Table 5.1: Boundary conditions for Simulation [2]

Temperature at bottom surface of package 20 ºC

Thermal Coupling between Dissipator and Base of Package 60 ºC/W

Thermal Coupling between Dissipator and Thermal Strap 300 W/m2 ºC

Thermal Coupling between Thermal Strap and Radiator 300 W/m2 ºC

Coupling between MLI and Package 0.03 W/m2 ºC

20

The figure 5.1 shows Boundary conditions applied for simulation in software.

Figure 5.1: Boundary conditions for Simulation

5.1.2 Orbital Heating Parameters for Simulation

Table 5.2: Orbital Heating Parameters for Simulation [2]

Minimum Altitude 100 km

Sun Flux 13676 W/m2

Orbital Period 7076.14 sec

Orbital Inclination 90º

Eccentricity 0

Argument of Periapsis 0

Satellite Position Local noon to ascending

21

Figure 5.2: Orbital Heating Parameters for Simulation

5.1.3 Time Interval for Transient Condition

Table 5.3: Time Interval for Transient Condition [2]

Time (sec) Q = 3.75 W (W) Q = 15 W (W)

0 3.75 15

600 3.75 15

601 1.875 7.5

1800 1.875 7.5

1801 0 0

6476.1452 0 0

6477.1452 3.75 15

7076.1452 3.75 15

5.2 Results

The results are obtained for each component in the form of minimum and maximum

temperature of the component and temperature distribution pattern along the surface of the

component. The results can be interpreted as, lower the temperature range, more is the heat

dissipation occurrence and vice versa. Following are the different cases with difference in

radiator dimensions with figures showing temperature range and temperature distribution

on the surface.

22

5.2.1 Radiator Dimension – 300x150x2mm

Case 1: Q = 3.75 W Transient Condition

Table 5.4: Case 1 Result

Sr. No. Part Min. Temperature (ºC) Max. Temperature (ºC)

1 Radiator 20.00 208.02

2 Dissipator 20.00 189.18

3 Thermal Strap 20.00 197.53

4 Package 20.00 72.17

Figure 5.3: Case 1-Radiator Figure 5.4: Case 1-Dissipator

Figure 5.5: Case 1-Thermal Strap Figure 5.6: Case 1-Package

23

Case 2: Q = 15 W Transient Condition



1 Radiator 20.00 412.61



4 Package 20.00 72.17

Case 3: Q = 3.75 W Steady State Condition



1 Radiator 596.70 602.43

2 Dissipator 514.15 544.05

3 Package 20.00 59.00


Figure 5.7: Case 3-Radiator Figure 5.8: Case 3-Dissipator

24

Figure 5.9: Case 3-Thermal Strap Figure 5.10: Case 3-Package

Case 4: Q = 15 W Steady State Condition



1 Radiator 1299.19 1304.88

2 Dissipator 1182.87 1254.07

3 Thermal Strap 1255.83 1283.72

4 Package 21.52 59.00

5.2.2 Radiator Dimension – 350x200x2 mm




1 Radiator 20.00 240.82


3 Package 20.00 71.60


25




1 Radiator 725.24 733.14

2 Dissipator 612.27 653.20

3 Package 20.00 58.30






1 Radiator 20.00 417.74


3 Package 20.00 70.85





1 Radiator 20.00 573.73


3 Package 20.00 70.85


26




1 Radiator 1394.21 1420.11

2 Dissipator 1192.14 1295.43

3 Package 20.00 57.46

4 Thermal Strap 1306.80 1339.83




1 Radiator 2091.63 2117.53

2 Dissipator 1861.90 1988.10

3 Package 20.00 57.46

4 Thermal Strap 2003.83 2037.25





1 Radiator 20.00 218.95



4 Package 20.00 72.17

27




1 Radiator 20.00 401.73



4 Package 20.00 72.17




1 Radiator 596.85 600.64

2 Dissipator 514.15 544.05

3 Package 20.00 59.00





1 Radiator 1299.33 1303.12

2 Dissipator 1182.86 1254.06

3 Thermal Strap 1248.06 1283.70

4 Package 20.00 59.00

28





1 Radiator 20.00 229.62


3 Package 20.00 71.60





1 Radiator 20.00 403.00


3 Package 20.00 71.60





1 Radiator 725.44 730.664

2 Dissipator 612.26 653.19

3 Package 20.00 58.30


29




1 Radiator 1426.30 1431.52

2 Dissipator 1281.50 1360.46

3 Thermal Strap 1357.05 1404.97

4 Package 20.00 58.30





1 Radiator 20.00 374.09


3 Package 20.00 70.85





1 Radiator 20.00 515.14


3 Package 20.00 70.85


30




1 Radiator 1394.75 1412.03

2 Dissipator 1192.14 1285.43

3 Package 20.00 57.46

4 Thermal Strap 1306.80 1339.82




1 Radiator 2092.17 2109.45

2 Dissipator 1861.90 1988.10

3 Package 20 57.46

4 Thermal Strap 2003.82 2037.24

31

Inferences & Future Work Chapter 6

This chapter contains the conclusions obtained from results of previous chapter and future

work.

6.1 Inferences

It is evident from comparison of the results of the cases 1, 2, 5, 7 and case 8 for transient

condition having 2 mm radiator thickness with cases 11, 12, 15, 19 and case 20 for transient

condition with 3 mm radiator thickness as well as cases 3, 4, 6, 9 and case 10 for steady

state condition having 2 mm radiator thickness with cases 13, 14, 17, 21 and case 22 for

steady state condition with 3 mm radiator thickness that as the thickness of the radiator

increases, the temperature range of the radiator also increases which can be justified as with

increase in thickness, thermal radiance decreases and heat transfer co-efficient increases

that leads to the first conclusion.

1. As the thickness of the radiator increases, the heat dissipation rate decreases i.e.

thickness of radiator is inversely proportional to heat dissipation rate.

From Stefan-Boltzmann equation of Radiating Power of radiator, it is seen that radiating

power is directly proportional to surface area of the radiator. This prompts to possibility of

increase in surface area of radiator being an effective tool for heat removal. On comparing

this relation with result obtained from case 3 and case 6, it is seen that temperature range

of radiator increases w.r.t case 1 and case 3 which indicates increase in surface area affects

heat dissipation process to decrease. But actually this happens because with increase in

surface area above certain extent the amount of incident thermal loads (Albedo, Solar) also

increase. This leads to the second conclusion.

2. The increase in surface area increases thermal loads incident on it which nullifies the

increased radiating power and as a result the radiator temperature is increased.

Results of cases with 15 W heat dissipation from dissipator suggests following conclusion.

32

3. Flat-plate radiator would be highly effective in case the amount of heat to be removed is

low, but the same cannot be said for higher amounts of heat removal.

6.2 Future Work

Future work includes overcoming the concluded limitations of flat-plate radiator.

1. To select best radiator profile which can reduce or nullify incident environmental loads

on radiator surface as well as maintain AFT for higher heat dissipation from the dissipator.

2. To select and design best radiator modification which can reduce the amount of

environmental thermal load incident on it and efficient enough to radiate higher power

dissipation.

33

References

[1] Spacecraft Thermal Control Handbook, Vol. 1: Fundamental Technologies,

Chapter 1-6, David G. Gilmore, Page 1-222.

[2] Design of Geosynchronous Satellites, Chapter 5, Brij N. Agarwal, Page 283

[3] “Thermal Control System of the Moon Mineralogy Mapper Instrument” by Jose I.

Rodriguez, Howard Tseng and Burt Zhang; Jet Propulsion Laboratory, California

Institute of Technology.

[4] “The Moon Mineralogy Mapper (M3) on Chandrayaan-1” by Alok Chatterjee,

Padma Varanasi.

[5] “The Moon Mineralogy Mapper (M3) for lunar science” by A. Chatterjee, Padma

Varanasi, A.S.K Kumar.

[6] http://ocw.mit.edu/courses/aeronautics-and-astronautics/16-851-satellite-

engineering-fall-2003/lecture-notes/l23thermalcontro.pdf

[7] https://www.nasa.gov/pdf/473486main_iss_atcs_overview.pdf

https://www.nasa.gov/pdf/473486main_iss_atcs_overview.pdf

34

Appendix I

Work Plan (7th Sem)

Work Plan (8th Sem)

35

Appendix II

Business Model Canvas

Business Canvas Report

KEY PARTNERS:

NASA is the partner which not only would apply the model but also take

part in promoting it for other applications.

ISRO

SIEMENS PLM acting as technical partner for the software.

Brown university is compatible with verifying the results of the model.

KEY ACTIVITIES:

Design Parameter: Selection of area and thickness of satellite package and

radiator with proper thermo-optical properties and materials.

36

Space thermal simulation: Simulation of satellite package with orbital

parameters and boundary conditions.

Interpretation: The simulation results yields the results in the form of

temperature range of each component which can be interpreted in to

conclusions for each case.

KEY RESOURCES:

Research papers

Personal Computers

UG-NX 7.5 software

Reference Books by Brij N. Agarwal and David G. Gilmore

VALUE PROPOSITIONS:

Heat Dissipation: The model works on control of heat dissipation by the

dissipator component of the satellite, so value of the project mainly depends

on efficiency of the radiator to dissipate this heat in the surroundings.

Radiator parameters: For the radiator to be efficient in simulating and

getting desired outputs the dimensions of the radiator are modulated by trial

and error methods and the most effective parameter is selected for execution

hence resulting in efficient heat radiation and economic design for

increasing its value proposition.

Economic material selection: The dimensional simulations would bring

upon the desired solution while simulating on the provided materials, these

materials are such that while actual manufacturing would be available on

economic revenues and hence increases the value propositions of the

model.

CUSTOMER RELATIONSHIP:

Weather forecasters: The weather forecasters can use the satellite model

for fetching weather conditions prevailing in an area due to certain

changes.

Military Organisations: The military organisations can know enemy line

activities and other weather related important constraints using satellite

model.

37

NASA: NASA being one of the key partners also applies the model in its

use for simulating the space thermal environment considering all the

constraints and heat loads.

CHANNELS:

Research Journals

Online Marketing

COST STRUCTURE:

Manufacturing cost

Material cost

Development cost

Verification cost

Software licensing costs

Research cost

Maintenances cost

38

Appendix III

Research Paper

THERMAL ANALYSIS OF PASSIVE RADIATORS FOR

INTER-PLANETARY SPACE APPLICATIONS Shailesh Kumar Singh Rajput1, Yash Dave2, Abhishek Dorik3, Prof. Harshal T.

Shukla4,

Dr. Rajesh N. Patel5 Research Scholar, Mechanical Dept., SNPIT & RC, Umrakh, Surat, Gujarat, India.1

Research Scholar, Mechanical Dept., SNPIT & RC, Umrakh, Surat, Gujarat, India.2

Research Scholar, Mechanical Dept., SNPIT & RC, Umrakh, Surat, Gujarat, India.3

Assistant Professor, Mechanical Dept., SNPIT & RC, Umrakh, Surat, Gujarat, India.4

HOD, Mechanical Dept., Nirma University, Ahmedabad, Gujarat, India.5

Abstract: An elaborate growth has been observed in the use of Satellites for various

domestic, military and navigational applications. Satellites carry various Infrared

instruments and Electronic Packages in them collectively called Payloads. The Payload

can function properly only if it is maintained within specified temperature ranges. The

Thermal Control System (TCS) of a Satellite keeps the equipment temperature within the

specified operating range. It is broadly divided into two classes namely, Passive Thermal

Control System (PTCS) and Active Thermal Control System (ATCS).

The current study aims to appraise the merits of using Passive Radiators for Interplanetary

Space Applications as it draws no power from the satellite system, and measuring its

Effectiveness in Dissipating the heat developed inside the payload to space against

Environmental Backloads incident over its surface from the Celestial Surroundings. It

maintains the desired temperature range by Controlling Conductive and Radiative Heat

Paths through the selection of Geometrical Configurations and Thermo-Optical

Properties of the surface in addition to savings in Mass and Power respectively which has

always been a crucial element in spacecraft design and configuration. A Parametric study

is conducted to explore the scopes of using Passive Radiators. The entire system is

Modelled and Simulated in FEA software UG NX 7.5 with a Flat Plate Radiator used in

the initial Space Thermal Analysis. Correlations between Heat Transfer Capacity,

Thermal Backloads, Radiator Area and the Operating Temperature are investigated to

provide Design Guidelines for Consistent and Predictable Performance with minimum

Degradation in a thermally stable orbit.

39

Keywords: Heat Dissipation, Modelling, Payload, Passive Radiator, Package,

Simulation, Thermal Control, Thermo-optical properties, Space Thermal Analysis, UG

NX.

I. INTRODUCTION

The objective of the thermal design is to provide proper heat transfer between all

spacecraft elements so that the temperature sensitive components can remain within their

Allowable Flight Temperature (AFT) limits. These could be achieved through the use of

either Passive Thermal Control System (PTCS) or Active Thermal Control System

(ATCS)[2,3].

The Active Thermal Control System is used in applications where the equipment has

a narrow specified temperature range and there is a great variation in equipment power

dissipation. It involves use of Mechanical and Electrical equipments like heaters, coolers,

piping, etc., adding to the mass and power requirements of the satellite resulting in increase

in cost.

Passive Radiators on the other hand reject heat into space without drawing any external

power from the satellite system. This makes it an attractive and feasible option for

instrument cooling. It does not draw any power from the satellite system. There are no

vibrations and electromagnetic interference produced in this system. It is a highly reliable

system with no moving parts, moving fluids or electric power input other than the power

dissipation of spacecraft functional equipment.

The passive radiators radiatively couple the satellite to the space and helps in

dissipating the heat generated inside into the space. The heat transfer in space is mainly

governed by Radiation heat transfer.

II. THERMAL LOADS ON PACKAGE

For the spacecrafts operating in orbits above the planetary atmosphere, it absorbs heat

from sources like direct sunlight incident over its surface, reflected sunlight from the

celestial bodies (Albedo) and planet emitted radiation. Electrical and electronic components

40

of a spacecraft produce heat which is rejected by spacecraft by infrared radiation from

external surfaces.

The heat balance equation is given by:

[Heat Radiated] = [Heat Incident] + [Instrumental Heat Dissipation]

• Heat Radiated from the Radiator = ε×η×σ×A×T4………………………………………...(1)

• Heat Incident on the Radiator = 𝛼 × 𝑆 × 𝐴 × sin 𝜃………………………………………(2)

• Instrumental Heat Dissipation = Q …………………………………………………….(3)

Direct solar flux is the only dominating source of environmental heating incident on

most of the spacecrafts in planetary orbits. The intensity of sunlight at Earth’s mean

distance from the sun (1 AU) is known as solar constant and is equal to 1367 W/m2. Albedo

is the sunlight reflected off a planet’s surface. It is usually expressed as the fraction of

incident sunlight that is reflected back to space. This light when incident on the spacecraft

becomes a thermal load on it. All incident sunlight that is not reflected as Albedo is

absorbed by planet is eventually emitted as IR energy and when incident over the

spacecraft, heat up its outer surface.

These loads can be reduced over the radiator plate by controlling the IR emittance and

solar absorptance properties of the coatings on the radiator plate.

III. CONSTRUCTION AND INPUT PARAMETERS

A schematic diagram for the Package Radiator assembly is as shown in Fig. 1. The

package consists of the Dissipator (Heat dissipating device) which is attached to a thermal

strap (Heat carrying strip) conductively. Thermal strap is responsible for transferring the

heat from the Dissipator to the Radiator which is placed at the top of the Package and is

exposed to the outer space. The heat is than spread over the radiator surface and dissipated

radiatively in the space. The whole package except the radiator is wrapped and insulated

with MLI (Multi Layer Insulation). The material properties of all the components of the

Payload are as shown in the Table 1.

41

Figure 1: Layout of Package Assembly

Table 1: Thermal-Optical Properties of Package subsystem

Device Material Dimensions Thermo-optical Properties

IR (ε) Solar (α)

Package Aluminium 6061 300 × 150 × 300 - -

Dissipator Stainless steel 50 × 50 × 40 - -

Radiator Aluminium 6061 300 × 150 × 2 0.85 0.6

MLI (Null Shell) 320 × 170 × 320 0.7 0.45

Thermal Strap Copper - - -

White paint is typically used as Radiator coating. The coating consists of a visibly

transparent material such as quartz glass or Teflon to achieve high emittance. A reflective

silver or Aluminium coating is used on the back to reduce solar absorptance. Ending-of-

life (EOL) absorptance and emittance values are selected to account for the stability of the

radiator coating through its entire operational life span. The entire Package-Radiator

assembly is covered with a Multi-Layer Insulation (MLI) blanket. MLI prevents excessive

heat loss from the spacecraft components as well as excessive heat gain from the celestial

surroundings.

Based on the Thermal Control System of M3 Instrument [4], the spacecraft interface

temperature for operating conditions is selected as 20 °C (acceptable range being between

0 to 40 °C) for the analysis purpose. The thermal analysis has been conducted considering

two different equipment power dissipations of 3.5 W and 15 W in Steady State Conditions.

In steady state conditions the thermal loads over the system (both external as well as

internal) do not vary over time and remain constant throughout the mission life. The main

objective of this work is to analyze the parametric relationship between total heat

42

dissipation and radiator area. The various boundary conditions have been tabulated in Table

2.

Table 2: Boundary Conditions

Bottom face of Package = 20 °C

Thermal coupling between Dissipator and Base of Package, R= 60 °C/W

Thermal coupling between Dissipator and Thermal Strap, h= 300 W/m2 °C

Thermal coupling between Thermal Strap and Radiator, h= 300 W/m2 °C

Thermal coupling between MLI and Package, h= 0.03 W/m2 °C

IV. RESULTS

Table 3: Table of Results

Sr. No. Radiator Area Groups Steady State Conditions

(Temperature °C)

Heat Dissipation

(watts)

Minimum Maximum

1. 300 × 150 × 2 Radiator 596.69 602.433

3.75 Dissipator 514.15 544.05

Package 20 59

Thermal Strap 547.22 581.22

2. 350 × 200 × 2 Radiator 280.5 788.409

3.75 Dissipator 659.45 706.40

Package 20 20


3. 400 × 250 × 2 Radiator 982.16 994.95

3.75 Dissipator 794.81 855.05

Package 20 20


4. 500 × 300 × 2 Radiator 1394.21 1420.11

3.75 Dissipator 1192.14 1285.43

Package 20 57.46


5. 300 × 150 × 2 Radiator 1299.19 1304.877

15 Dissipator 1182.87 1254.07

Package 21.52 59


43

6. 350 × 200 × 2 Radiator 1481.761 1489.684

15 Dissipator 1330.87 1413.56

Package 19.584 23.527


7. 400 × 250 × 2 Radiator 1680.78 1693.56

15 Dissipator 1464.92 1559.36

Package 20 20


8. 500 × 300 × 2 Radiator 2091.63 2117.53

15 Dissipator 1861.90 1988.10

Package 20 57.46


The simulation result for the 500 × 300 × 2, 15 watt heat dissipation has been shown

in the figure. It can be seen that increasing the radiator area does not bring the Dissipator

temperature to the required acceptable range.

Figure 2: Dissipator Figure 3: Radiator

44

Figure 4: Thermal Strap Figure 5: Package

V. CONCLUSIONS AND DISCUSSIONS

It can be seen from the Stefan Boltzmann radiative heat transfer equation that the heat

transfer increases with the increase in surface area of the flat plate radiator,

Heat Radiated = σ×A×T4 watts

But with increase in the radiator area, the environmental load on the surface becomes

more dominating, resulting in decreased heat transfer and accumulation of heat in the

package leading to excessively high equipment temperatures. Such high equipment

temperatures can destroy the package components. Hence increasing the radiator area alone

is not a solution to increasing the heat transfer from the package. Conversely, it decreases

the heat transfer because the environmental backloads become more dominating.

VI. REFERENCES

[1] Burt Zhang, Melora Larson, Jose Rodriguez, “Passive Coolers for pre-cooling of JT loops for deep space

infrared imaging applications”, Cryogenics 50 (2010) 628-632, Available:

www.elsevier.com/locate/cryogenics,

[2] Brij N. Agrawal, “Design of Geosynchronous Spacecrafts”, PRENTICE-HALL, INC., Englewood

Cliffs, NJ 07632

[3] David G. Gilmore, “Space craft Thermal control handbook”, Vol. 1 Fundamental Technologies, The

Aerospace Press, El Segundo, California.

[4] Jose I. Rodriguez, Howard Tseng and Burt Zhang, “Thermal Control System of the Moon Mineralogy

Mapper Instrument”, 2008-01-2119.

http://www.elsevier.com/locate/cryogenics

45

Appendix IV

Plagiarism Report

46

47

48

49

50

20/04/2016 Periodic Progress Report (PPR) Details

1/2

PRINT BACK

Enrollment No : 120490119093 College : Vidyabharti Trust, Institute Of Technology &

Research Centre, UmrakhBardoli

Student Name : Rajput Shailesh Santosh Kumar Singh Department : Mechanical Engineering

Mobile No : Discipline : BE

Email : Semester : Semester 8

PPR Details

Time Interval :

Periodic Progess Report : First PPR

Project

:

Thermal Analysis Of A Passive Radiator For InterPlanetary Space Applications

Status : Reviewed (Freeze)

1. What Progress you have made in the Project ?

Simulations are being conducted by varying the parametric relations between thermal loads and area. All the news

results are being analysed for meaningful conclusions.

2. What challenge you have faced ?

Understanding the complexity of the software has been the biggest challenge. With every set of simulation requiring

its own modelled and FEM part generation, the procedure is bit time consuming.

3. What support you need ?

To draw in some robust meaningful conclusions we need more data published in the same domain of study.

Standard data is hard to get as it is mostly published into paid journals with limited accessibility. Limitations in

software expertise is also one of the core problems due to inherent complexities of the software. Stiff college

schedule makes time a major constraint.

4. Which literature you have referred ?

Moon Spacecraft Thermal Handbook by David G. Gilmore has helped abundantly in increasing our knowledge about

various thermal finishes and coatings which can be applied to reduce the absorption from the surroundings and

increase the emittance from the spacecraft.


2/2

Comments

Comment by Internal Guide :

Try to gather some more research paper..


1/2

PRINT BACK






PPR Details

Time Interval : 0 days, 0 hours, 13 minutes, 36 seconds

Periodic Progess Report : Second PPR

Project

:




The results obtained from ongoing simulations are been analysed to derive some meaningful conclusions as well

study is being conducted to find some alternative to the Passive Flat Plate Radiator Design so as to make the system

more viable for use with optimum interaction with the space and meeting the Allowable Flight Temperature (AFT)

requirements..


Availability of resources in the study domain is a persistent problem.


Slowly we have been gaining hands over the software and are now able to understand the procedures more

accurately. Any further resource to nourish our grip over the software is helpful.


Along with D G. Gilmore's Handbook on Space Craft Design, we are also referring to Brij N. Agarwal's "Design of

Geosynchronous Spacecraft" to strengthen our knowledge of Thermal Control Sysytem.


2/2

Comments


Good Work


1/2

PRINT BACK






PPR Details


Periodic Progess Report : Third PPR

Project

:




Till now we had been modelling, Meshing and simulating the system for steady state conditions wherein the load

conditions do not change with time. Now we are putting the completed Steady State conditions to transient

environment conditions wherein the loads over the satellite vary with time and celestial conditions.


The excessive time taken by each Transient condition simulation makes the work slow and requires more graphics at

the computer end to produce viable results. Lack of Research papers in this domain are slowing our progress.


If some source of Research Papers can be obtained, it would be very beneficiary for us to derive some coherent

conclusions.


The "Thermal Control System of the Moon Mineralogy Mapper Instrument" paper by Jose I. Rodriguez has helped us

in understanding the Transient conditions and environment of space and proceed to the Transient Simulation stage.


2/2

Comments


Proceed further with some more results.


1/2

PRINT BACK






PPR Details


Periodic Progess Report : Forth PPR

Project

:




We are simultaneously working on the generation of results in both Steady State and Transient Environment of the

spacecraft and comparing the results of these categories. The result would be helpful in providing design guidelines

for both the conditions. Meanwhile we have also been working on other tasks related with the project like BMC sheet

and its report.


The Transient conditions simulation puts a great demand at the computer graphics and RAM requirements. Apart

from that it is a very time consuming procedure. Meeting, learning the software and Simulating in both transient and

steady state conditions in a stiff time constraint is a bit demanding with other course work in progress simultaneously.


The college has been very helpful in providing us with computers with high specifications which helps us in

obtaining results of Transient conditions. But that again can be accessed only in the college hours and after

completion of study curriculum.,


With the limited availability of Research Papers in this domain, we have been referring to various available papers in


2/2

Design of Thermal Control of Spacecraft and books by Brij N. Agarwal and D G. Gilmore to nourish our knowledge of

the topic.

Comments


Good now compare the results of steady and transient state.

GIC Patent Drafting Exercise Team ID: 48225

GTU Innovation CouncilPatent Drafting Exercise (PDE)

FORM 1 (FOR OFFICE USE ONLY)THE PATENTS ACT 1970 Application No:

(39 OF 1970) Filing Date:& Amount of Fee paid:

THE PATENTS RULES, 2003 CBR No:___________________APPLICATION FOR GRANT OF PATENT

1. Applicant(s) :

ID Name Nationality Address Mobile No. Email

1 Rajput ShaileshSantosh Kumar Singh

Indian Mechanical Engineering ,Vidyabharti Trust, Institute OfTechnology & Research Centre,Umrakh-Bardoli ,Gujarat Technologycal University.

9426175099 [email protected]

2 Dave YashJayeshbhai



3 Dorik AbhishekDeepak



4 Patel DipakkumarSureshbhai



5 Modi BijankkumarKrishnakant



2. Inventor(s):

ID Name Nationality Address Mobile No. Email

1 Rajput ShaileshSantosh Kumar Singh



Note : This is just a mock Patent Drafting Exercise (PDE) for semester 8, BE students of GTU.These documentsare not to be submitted with any patent office.

Page 1

2 Dave YashJayeshbhai



3 Dorik AbhishekDeepak



4 Patel DipakkumarSureshbhai



5 Modi BijankkumarKrishnakant



3. Title of Invention/Project:Thermal Analysis Of A Passive Radiator For Inter-Planetary Space Applications

4. Address for correspondence of applicant/authorized patent agent in india

Name: Rajput Shailesh Santosh Kumar SinghAddress: Mechanical Engineering , Vidyabharti Trust, Institute Of Technology & Research Centre, Umrakh-Bardoli ,

Gujarat Technological University.Mobile: 9426175099Email ID: [email protected]

5. Priority particulars of the application(S) field in convention country

Country Application No. Filing Date Name of the Applicant Title of the Invention

N/A N/A N/A N/A N/A

6. Particulars for filing patent co-operation treaty (pct) national phase Application

International application number International filing date as alloted by the receiving office

N/A N/A

7. Particulars for filing divisional application

Original(First) Application Number Date of filing of Original (first) application

N/A N/A

8. Particulars for filing patent of addition

Original(First) Application Number Date of filing of Original (first) application

N/A N/A


Page 2

9. DECLARATIONS:(i) Declaration by the inventor(s)

I/We, the above named inventor(s) is/are true & first inventor(s) for this invention and declare that the applicant(s).herein is/are my/our assignee or legal representative.Date : 20 - April - 2016

Name Signature & Date

1 Rajput Shailesh Santosh Kumar Singh ______________

2 Dave Yash Jayeshbhai ______________

3 Dorik Abhishek Deepak ______________

4 Patel Dipakkumar Sureshbhai ______________

5 Modi Bijankkumar Krishnakant ______________

(ii) Declaration by the applicant(s) in the convention countryI/We, the applicant (s) in the convention country declare that the applicant(s) herein is/are my/our assignee or legalrepresentative.applicant(s)

(iii) Declaration by the applicant(s)I/We, the applicant(s) hereby declare(s) that:-

I am/We in possession of the above mentioned invention.

The provisional/complete specification relating to the invention is filed with this aplication.

The invention as disclosed in the spcification uses the biological material from India and the necessarypermission from the competent authority shall be submitted by me/us before the grant of patent to me/us.

There is no lawful ground of objection to the grant of the patent to me/us.

I am/we are the assignee or the legal representative of true & first inventors.

The application or each of the application,particulars of each are given in the para 5 was the first applicatinin the convention country/countries in respect of my/our invention.

I/we claim the priority from the above mentioned applications(s) filed in the convention country/countries &state that no application for protection in respect of invention had been made in a convention country beforethat date by me/us or by any personMy/Our application in india is based on international application under Patent Cooperation Treaty (PCT) asmentioned in para 6

The application is divided out of my/our application(s) particulars of which are given in para 7 and pray thatthis application may be treated as deemed to have been filed on ___________under section 16 of the Act.

The said invention is an improvement in or modification of the invention particulars of ehivh are given in para8.

10. Following are the attachments with the application:

(a) Provisional specification/Complete specification

(b) Complete specification(In confirmation with the international application) / as amended before theinternational Preliminary Examination Authority (IPEA),as applicable(2 copies),No.of pages.....No.ofclaims.....


Page 3

(c) Drawings (In confirmation with the international application)/as amended before the internationalPreliminary Examination Authority(IPEA),as applicable(2 copies),No.of sheets....

(d) Priority documents

(e) Translations of priority documents/specification/international search reports

(f) Statement and undertaking on Form 3

(g) Power of Authority

(h) Declaration of inventorship on Form 5

(i) Sequence listing in electronic Form

(j) ........................................ Fees Rs.XXX in Cash /Cheque/Bank Draft bearin No.XXX Date: XXX on XXXBank.

I/We hereby declare that to the best of my /our knowledge, information and belief the fact and mttersstated herein are correct and I/We request that a patent may be granted to me/us for the said invention.

Dated this 20 day of April , 2016

Name Signature & Date

1 Rajput Shailesh Santosh Kumar Singh ______________

2 Dave Yash Jayeshbhai ______________

3 Dorik Abhishek Deepak ______________

4 Patel Dipakkumar Sureshbhai ______________

5 Modi Bijankkumar Krishnakant ______________


Page 4


FORM 2THE PATENTS ACT, 1970

(39 OF 1970)&

THE PATENTS RULES, 2003PROVISIONAL SPECIFICATION

1. Title of the project/invention :

Thermal Analysis Of A Passive Radiator For Inter-Planetary Space Applications

2. Applicant(s) :

Rajput Shailesh Santosh Kumar Singh ( Indian )Address : Mechanical Engineering , Vidyabharti Trust, Institute Of Technology & Research Centre, Umrakh-Bardoli , Gujarat Technologycal University.

Dave Yash Jayeshbhai ( Indian )Address : Mechanical Engineering , Vidyabharti Trust, Institute Of Technology & Research Centre, Umrakh-Bardoli , Gujarat Technologycal University.

Dorik Abhishek Deepak ( Indian )Address : Mechanical Engineering , Vidyabharti Trust, Institute Of Technology & Research Centre, Umrakh-Bardoli , Gujarat Technologycal University.

Patel Dipakkumar Sureshbhai ( Indian )Address : Mechanical Engineering , Vidyabharti Trust, Institute Of Technology & Research Centre, Umrakh-Bardoli , Gujarat Technologycal University.

Modi Bijankkumar Krishnakant ( Indian )Address : Mechanical Engineering , Vidyabharti Trust, Institute Of Technology & Research Centre, Umrakh-Bardoli , Gujarat Technologycal University.

3. Preamble to the description :

The following specification describes the invention.


Page 1

4. Description :

a. Field of Application / Project / Invention :

The project finds its applications in the field of Mechanical Engineering as well as Aerospace Engineering.

b. Prior Art / Background of the Invention / References :

The project aims at optimising the heat transfer capacity of Passive Radiators in space environment. These can beachieved by minimising the effect of space environmental loads on the Radiator plate by using suitable surfacecoatings and plate orientations.

c. Summary of the Invention/Project :

The project through software based analysis aims at understanding the relationship between the change in area ofradiator and its effect on the heat transfer capacity of the flat plate radiator. It has been established by the StefanBoltzmann relation that in Radiation heat transfer which is dominant in space environment, increasing the surface areaof an object increases its heat transfer capacity. But through the analysis it is evident that increasing the radiator area,the environmental thermal load on the surface increases significantly. this leads to reduced heat transfer andaccumulation of heat in the Package resulting in failure of the satellite system.

d. Objects of the Invention/Project :

For the analysis purpose, from the previous experience a package Radiator system has been assumed consisting of:1) Dissipator -- heat dissipating device which in ore case dissipates 3.5 watts and 15 watts for analysis.2) Thermal Strap -- It is a Heat Carrying Strip that transfers heat from the dissipator block and spreads it over theradiator surface.3) Radiator -- Flat plate radiator is responsible for the heat transfer from the satellite system to the surroundings.4) Package -- It contains all the components of interest, Dissipator and Thermal Strap.5) MLI -- The whole Package is Insulated with Multi Layer Insulation (MLI) that prevents excessive heat loss from thepackage to the surroundings and excessive heat gain from the space surroundings to the package.

e. Drawing(s) :

48225_1_Package-Radiator Assmebly

f. Description of the Invention :

Passive Radiators reject heat into space without drawing any external power from the satellite system. This makes itan attractive and feasible option for instrument cooling. It does not draw any power from the satellite system. There areno vibrations and electromagnetic interference produced in this system. It is a highly reliable system with no movingparts, moving fluids or electric power input other than the power dissipation of spacecraft functional equipment. Thepassive radiators radiatively couple the satellite to the space and helps in dissipating the heat generated inside into thespace.

The present work through the space thermal analysis aims at appraising the scopes of use of Passive radiator in thespace exploration and imaging industry. In our study, heat is developed at the dissipator, carried by the Thermal Strapto the radiator and the radiator transfers this heat to the space to complete the heat transfer process. But the totalthermal control is obtained by balancing both the heat to be dissipated in the space and the one that is incident on theradiator surface from the surroundings. We in our study have selected a range of radiator areas to analyse the effect ofradiator area variation over the heat transfer capacity.

Part file creation, Modelling and Simulation has been conducted for various radiator areas for two states ofenvironmental conditions namely, Steady state and Transient Conditions. The final conclusion is derived by analysingthe results of these set of simulations and providing suitable design guidelines.

g. Examples :

h. Unique Features of the Project :

1) Thermo-optical properties have been considered for the radiator surface in the simulations.2) Numerous set of Radiator areas have been considered in the analysis.3) For various set of Radiator areas, different thickness values have been considered to analyse the effect of increasein heat capacity of radiator over the heat transfer.4) Increase in heat transfer through use of Passive Radiators results in savings in Power and Mass requirements of thesatellite as the entire assembly is very light and robust.5) With reduced mass requirement for Radiator, the payload carrying capacity can be increased.6) It is very cost effective and offers stable and consistent operation in a thermally stable environment.7) It is highly reliable system.

5. Date & Signature :

Date :20 - April - 2016


Page 2

____________________ ____________________Sign and Date Sign and Date

Rajput Shailesh SantoshKumar Singh

Dave Yash Jayeshbhai


Dorik Abhishek Deepak Patel Dipakkumar Sureshbhai

____________________Sign and Date

Modi Bijankkumar Krishnakant

6. Abstract of the project / invention :An elaborate growth has been observed in the use of Satellites for various domestic, military and navigationalapplications. Satellites carry various Infra-red instruments and Electronic Packages in them collectively called Payloads.The Payload can function properly only if it is maintained within specified temperature ranges. The Thermal ControlSystem (TCS) of a Satellite keeps the equipment temperature within the specified operating range. It is broadly dividedinto two classes namely, Passive Thermal Control System (PTCS) and Active Thermal Control System (ATCS).

The current study aims to appraise the merits of using Passive Radiators for Interplanetary Space Applications as itdraws no power from the satellite system, and measuring its Effectiveness in Dissipating the heat developed inside thepayload to space against Environmental Back-loads incident over its surface from the Celestial Surroundings. It maintainsthe desired temperature range by Controlling Conductive and Radiative Heat Paths through the selection of GeometricalConfigurations and Thermo-Optical Properties of the surface in addition to savings in Mass and Power respectively whichhas always been a crucial element in spacecraft design and configuration. A Parametric study is conducted to explore thescopes of using Passive Radiators. The entire system is Modelled and Simulated in FEA software UG NX 7.5 with a FlatPlate Radiator used in the initial Space Thermal Analysis. Correlations between Heat Transfer Capacity, Thermal Back-loads, Radiator Area and the Operating Temperature are investigated to provide Design Guidelines for Consistent andPredictable Performance with minimum Degradation in a thermally stable orbit.

Drawing Attachments :

48225_1_Package-Radiator Assmebly


Page 3


Page 4


FORM 3THE PATENTS ACT, 1970

(39 OF 1970)&

THE PATENTS RULES, 2003STATEMENT AND UNDERTAKING UNDER SECTION 8

1. Declaration :

I/We, Rajput Shailesh Santosh Kumar Singh ,Dave Yash Jayeshbhai ,Dorik Abhishek Deepak ,Patel Dipakkumar Sureshbhai ,Modi Bijankkumar Krishnakant

2. Name, Address and Nationality of the joint Applicant :

Rajput Shailesh Santosh Kumar Singh ( Indian )Address :Mechanical Engineering , Vidyabharti Trust, Institute Of Technology & Research Centre, Umrakh-Bardoli , Gujarat Technologycal University.

Dave Yash Jayeshbhai ( Indian )Address :Mechanical Engineering , Vidyabharti Trust, Institute Of Technology & Research Centre, Umrakh-Bardoli , Gujarat Technologycal University.

Dorik Abhishek Deepak ( Indian )Address :Mechanical Engineering , Vidyabharti Trust, Institute Of Technology & Research Centre, Umrakh-Bardoli , Gujarat Technologycal University.

Patel Dipakkumar Sureshbhai ( Indian )Address :Mechanical Engineering , Vidyabharti Trust, Institute Of Technology & Research Centre, Umrakh-Bardoli , Gujarat Technologycal University.

Modi Bijankkumar Krishnakant ( Indian )Address :Mechanical Engineering , Vidyabharti Trust, Institute Of Technology & Research Centre, Umrakh-Bardoli , Gujarat Technologycal University.

Here by declare :(i) that I/We have not made any application for the same/substantially the same invention outside India.(ii) that the right in the application(s) has/have been assigned to,

Name of theCountry

Date ofApplication

ApplicationNumber

Status of theApplication

Date ofPublication

Date of Grant

N/A N/A N/A N/A N/A N/A

(iii) that I/We undertake that up to the date of grant of patent by the Controller , I/We would keep him inform in writing thedetails regarding corresponding application(s) for patents filed outside India within 3 months from the date of filing ofsuch application.

Dated this 20 day of April , 2016

3. Signature of Applicants :


Page 1


Rajput Shailesh SantoshKumar Singh

Dave Yash Jayeshbhai


Dorik Abhishek Deepak Patel Dipakkumar Sureshbhai

____________________Sign and Date

Modi Bijankkumar Krishnakant

ToThe Controller of PatentThe Patent Office, at Mumbai.


Page 2