published paper: thermal analysis of passive radiators for interplanetary space applications

International Conference on: “Engineering: Issues, opportunities and Challenges for Development”

ISBN: 978-81-929339-3-1

THERMAL ANALYSIS OF PASSIVE RADIATORS FOR INTER-

PLANETARY SPACE APPLICATIONS

Shailesh Kumar Singh Rajput1, Yash Dave

2, Abhishek Dorik

3, Prof. Harshal T. Shukla

4,

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.

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

Thermal Control, Thermo-optical properties, Space Thermal Analysis, UG NX.


ISBN: 978-81-929339-3-1

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).

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 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


ISBN: 978-81-929339-3-1

• Heat Incident on the Radiator = α×S×A×SinΘ

• Instrumental Heat Dissipation = Q

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.

Figure 1: Layout of Package Assembly


ISBN: 978-81-929339-3-1

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 [], 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 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


ISBN: 978-81-929339-3-1

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


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


ISBN: 978-81-929339-3-1


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 watts 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

Figure 4: Thermal Strap Figure 5: Package


ISBN: 978-81-929339-3-1

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