INDIAN INSTITUTE OF SPACE SCIENCE AND TECHNOLOGY,

THIRUVANANTHAPURAM

TEMPERATURE CONTROL OF SPACECRAFT

INTERNAL PROJECT REPORT

Submitted By:

Dheeraj Agarwal A. John Vivian Prashant

Nitin Kumar Mishra

IN

THERMAL SYSTEMS GROUP

AT

ISRO SATELLITE CENTRE, BANGALORE

JULY ’09 – AUGUST ’09

Yashwanth Kumar Nakka

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ACKNOWLEDGEMENT

A small step in the field of technology requires great support and expert guidance. Our

successful project is a faithful culmination effort by many people, some directly involved, some

enlightening us with their concealed presence. We at this juncture would like to extend our

profound gratitude to Dr. B.N. Suresh, Director IIST without whom, this endeavor would not

have been a reality.

We extend a warm gratitude to Mr. B. Nagaraju (Scientist/ Engineer – D) of Thermal System

group who has been an integral part of our project and for helping us with almost everything

including our visits to various labs. We also thank him for his enormous support in achieving and

successfully completing this project. His guidance has been a major thrust for our project.

Our hands on session at Thermal Systems Lab were very educative. We thus, thank Shri.

Sundaresan, Engineer SF, Mr. Suresh Babu, Engineer SE and Shri Gaikwad Engineer SD,

who helped us with almost everything including our experimental procedure, visits to various

labs and lastly for spending his valuable time with us.

It was a very rich experience for us at ISRO Satellite Centre (ISAC), where we gained a lot

about Thermal control of Spacecrafts and many more things. We thus, use this platform to thank

Dr. Badari Narayana, GD, TSG, Dr. Alok Shrivastava, Head, TDS, Shri S. G. Barve, Head,

TDAD, for spending their valuable time with us.

We take this opportunity to thank the administrative team of ISAC as well as that of IIST who

made arrangements for our comfortable stay and accessing ISRO’s facilities. Our sincere thanks

to Dr. Surendra Pal, Associate director (ISAC), Mrs. Sheila Iyer, Head HRD Division

(PPEG), Mr. Prakash K S (Sc. /Engg. SE), HRD Division (PPEG).

Lastly, we would like to thank all those who had silently been behind us but may have left our

notice for making this project a success.

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ABSTRACT

Thermal Protection of Spacecrafts is the most vital system of a satellite, which can have no

margin of error atoll as it has a major role in ensuring the complete success of the mission as a

whole. The report presents a brief idea about the materials used and fabrication of some of the

most widely used temperature control systems and also suggests some new materials and

techniques which require more research and experimentation to achieve very effective and

accurate desirable outputs. While planning any satellite mission the design and analysis of heat

loads on the spacecraft are worked out using suitable software.

Based on the temperature requirements and the nature of heating or cooling required by the

components on the spacecraft the most appropriate temperature control devices and methods are

implemented. This is very necessary because the electronic and experimental packages etc

perform at their best only when they are operated within the prescribed temperature ranges. If

this is not achieved the components may fail and can also lead to the complete failure of the

mission.

Thermal control should also be efficient, reliable and cheap. Mass minimization also poses a

challenge to the designers. The temperature control systems have to be put through rigorous

testing procedure so as to ensure maximum reliability. This is achieved by testing the individual

components. In addition to this, satellites are tested, when fully integrated inside a thermo

vacuum chamber. The various materials used for thermal protection of spacecrafts are

manufactured with high level of accuracy so as to meet the standards and also the handling of

these materials is done very carefully so as not to cause damage to the optical properties of these

materials.

Thermal control of spacecrafts is a field, which requires high level of accuracy which is achieved

only through a lot of experience and insight in the field. It is impossible to describe all the

intricacies and complications involved in this field. This report aims to provide a birds eye view

on all the basic concepts and techniques involved in the field of spacecraft thermal control.

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CONTENTS

1. INTRODUCTION.................................................................................................1-9

1.1 ABOUT THE THERMAL SYSTEMS GROUP

1.2 EARTH ORBITS

1.3 TEMINOLOGY

1.4 THERMAL LOADS ON SATELLITE

1.5 INSULATION

1.6 RADIATORS

1.7 HEATERS

1.8 HEATPIPES

1.9 LOUVERS

1.10 PHASE CHANGE MATERIALS

1.11 ELECTROCHROMIC MATERIALS

2. MULTI LAYER INSULATION BLANKETS...............................................10-16

2.1 INTRODUCTION

2.2 CHARACTERISTICS OF MLI

2.3 MATERIAL SELECTION FOR MLI

2.4 FABRICATION TECHNIQUES OF MLI

3. OPTICAL SOLAR REFLECTORS...............................................................17-19

3.1 INTRODUCTION

3.2 STRUCTURE OF OSR

3.3 FABRICATION OF OSR

4. LOUVERS...................................................................................................20-25

4.1 INTRODUCTION

4.2 ASSEMBLY

4.3 TYPES OF LOUVERS

5. HEAT PIPES.................................................................................................26-43

5.1 INTRODUCTION

5.2 LIMITATIONS ON TRANSPORT CAPACITY

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5.3 THERMODYNAMIC CONSIDERATIONS

5.4 TYPES OF HEAT PIPES

5.5 WORKING FLUIDS

5.6 MATERIAL COMPATIBILITY

5.7 TESTING OF HEAT PIPES

5.8 HEAT PIPE APPLICATION AND PERFORMANCE

6. THERMAL CONTROL PAINT………………………………………….............44-45

6.1 INTRODUCTION

6.2 BLACK PAINTS

6.3 WHITE PAINTS

7. HEATERS………………………………………………………………….............46-47

7.1 INTRODUCTION

7.2 TYPES OF HEATERS

8. TEMPERATURE MEASUREMENT DEVICES………………………................48-50

8.1 INTRODUCTION

8.2 THERMISTOR

8.3 PLATINUM RESISTANCE THERMOMETER

9. THERMO VACCUM CHAMBER………………………………………...............51-55

9.1 INTRODUCTION

9.2 MAJOR COMPONENTS OF THERMO VACUUM CHAMBER

9.3 POWER SUPPLY AND POWER MEASUREMENT

9.4 DESCRIPTION

9.5 THERMAL TESTS

10. SCOPE FOR FUTURE STUDY…………………………………………................56-77

10.1 INTRODUCTION

10.2 ELECTROCHROMIC MATERIALS

10.3 EMMISIVITY MODULATION DEVICES

10.4 PHASE CHANGE CONCEPT

10.5 MEMS BASED THERMAL CONTROL METHODS

CONCLUSION...........................................................................................................78

REFERENCE..............................................................................................................79

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1. INTRODUCTION

1.1 About the Thermal Systems Group

Spacecraft thermal protection design & fabrication is essentially a multi-disciplinary exercise.

Besides application of a large number of modern scientific and engineering techniques. High

degree of quality and reliability has to be ensured so as to ensure desired functioning of the

electronic packages. Satellites in space have to be protected and maintained at desirable

temperatures to ensure the proper functioning of the electronic packages. There have been many

technical advances in the field of thermal protection of spacecrafts. But the techniques should not

only have to be innovative but also cost effective & reliable.

The calculations regarding the thermal maintenance of a satellite depend on various factors like

external loads, internal heat dissipation, radiation heat exchange & conduction heat exchange.

External loads consist of Solar load, EarthShine & albedo.

Solar load is the radiation incident directly from the sun, while EarthShine refers to the radiation

emitted by earth and albedo load is the radiation reflected by the earth surface. Thermal-control

subsystem (TCS) hardware used to control temperatures of all spacecraft components includes

surface finishes, insulation blankets, heaters, and refrigerators.

Establishing a thermal design for a spacecraft is usually a two-part process. The first step is to

select a thermal design for the body, or basic enclosures, of the spacecraft that will serve as a

thermal sink for all internal components. The second step is to select thermal designs for various

components located both within and outside the spacecraft body.

ISRO has contributed significantly in the field of thermal control systems. Indian Remote

sensing satellites & communication satellites have been very successfully designed, fabricated &

launched by ISRO. ISRO has indigenously developed state of the art active & passive thermal

control systems. The thermal control systems of all satellites launched by ISRO have performed

well & provided the required environment for the satellites. The development of the thermal

systems group right from scratch has catered to the needs from ARYABHATA to Space Capsule

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Recovery Experiment. The excellent thermal performance of all ISRO satellites speaks volumes

of the engineering calibre & leadership skills of the engineers of the Thermal Systems Group of

the ISRO.

Thermal Systems Group of the ISRO has successfully achieved the indigenisation of all the

thermal control equipments, devices & softwares. It has achieved expertise in passive radiant

coolers, aluminium ammonia grooved heat-pipes, thermo foils, aluminised mylar, polyester nets,

mechanical cryorefrigerators , optical solar reflectors, Multi layer insulation, tape heaters &

surface treatment techniques. It has out smarted many brilliant space organizations abroad by

focussing on producing cheap & reliable technology for thermal control of spacecrafts. Its

phenomenal success has been the output of quality work done by the engineers of the thermal

systems group.

1.2 Earth Orbits

A variety of orbits are used for different types of Earth-oriented missions. The most common

orbits, are low Earth (LEO) and geosynchronous (GEO). Orbits whose maximum altitudes are

less than approximately 2000 km are generally considered low Earth orbits. They have the

shortest periods, on the order of an hour and a half. Some of these orbits are circular, while

others may be somewhat elliptical. The degree of eccentricity is limited by the fact that the orbit

is not much larger than Earth, whose diameter is approximately 12,760 km.

The inclination of these orbits, which is the angle between the plane of the equator and the plane

of the orbit, can vary from 0 deg to greater than 90 deg. Inclinations greater than 90 deg cause a

satellite in LEO to orbit in a direction opposite to Earth's rotation. Low Earth orbits are very

often given high inclinations so that the satellite can pass over the entire surface of Earth from

pole to pole as it orbits.

This coverage is important for weather and surveillance missions. One particular type of low

Earth orbit maintains the orbit plane at a nearly fixed angle relative to the sun. The result of this

is that, on every orbit, the satellite passes over points on Earth that have the same local time, that

is, the same local sun-elevation angle. Because Earth rotates beneath the orbit, the satellite sees a

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different patch of Earth's surface on each revolution and can cover the entire globe over the

course of a day. The ability to see the entire surface of Earth at the same local Sun angle is

important for weather observation and for visual surveillance missions. This type of orbit is

known as Sun-synchronous. It may be positioned so that satellites always see points on Earth at a

specific time, anywhere from local sunrise/sunset to local noon. They are often known as "noon"

or "morning" orbits.

1.3 Terminology

Several orbital parameters are commonly used in analyses of environmental heating. The same

parameters are used to describe the spacecraft orbit, and their use simplifies the process of

getting the inputs necessary to conduct the thermal analysis for any given program. The most

important parameters are:

Equatorial plane: The plane of Earth's equator, which is perpendicular to Earth's spin axis.

Ecliptic plane: The plane of Earth's orbit around the sun. From the point of view of Earth, the

sun always lies in the ecliptic plane. Over the course of a year, the sun appears to move

continuously around Earth in this plane. Because of the tilt of Earth's spin axis, the equatorial

plane is inclined 23.4o

from the ecliptic plane.

Sun day angle: The position angle of the Sun in the ecliptic plane measured from vernal

equinox. At vernal equinox this angle is 0o, at summer solstice 90

o, at autumnal equinox 180

o

and at winter solstice 270o.

Orbit inclination: The angle between the orbit plane and the equatorial plane. Orbit inclinations

vary from 0o to 98

o. For inclinations less than 90

o the satellite appears to be going around its

orbit in the same direction as Earth's rotation. For inclinations greater than 90o, it appears to be

going opposite Earth's rotation.

Apogee/perigee: Apogee is the point of highest altitude in an orbit & perigee the lowest.

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Ascending node/descending node: The ascending node is the point in the orbit at which the

spacecraft crosses Earth's equator while travelling from south to north. The descending node is

the point crossed during the southbound portion of the orbit.

Right ascension of the ascending node (RAAN): The position angle of the ascending node

measured from vernal equinox in the equatorial plane.

1.4 Thermal loads on satellite

Satellite thermal control process of energy management depends on environmental heating. The

principal forms of environmental heating on orbit are direct Sunlight, Sunlight reflected off Earth

(albedo), and infrared (IR) energy emitted from Earth. During launch or in exceptionally low

orbits, there is also a free molecular heating effect caused by friction in the rarefied upper

atmosphere.

The overall thermal control of a satellite on orbit is usually achieved by balancing the energy

emitted by the spacecraft as IR radiation against the energy dissipated by its internal electrical

components plus the energy absorbed from the environment. Atmospheric convection is absent

in space, as there is no air.

Direct Solar Sunlight is the greatest source of environmental heating incident on most spacecraft

in Earth orbit. Sun is a very stable energy source. The intensity is at its minimum value of 1388

W/m2 & the intensity is at its maximum of 1420 W/m

2.

Albedo is the Sunlight reflected off a planet or moon. A planet's albedo is usually expressed as

the fraction of incident Sunlight that is reflected back to space. Earth IR consists of incident

Sunlight which is not reflected as albedo, and is absorbed by Earth and eventually reemitted as

IR energy. The intensity of IR energy emitted at any given time, from a particular point on Earth

can vary considerably depending on factors such as the local temperature of Earth's surface and

the amount of cloud cover. The IR energy emitted by Earth has an effective average temperature

of around 255 oK.

Free Molecular Heating (FMH) is also a significant form of environmental heating. This kind

of heating is a result of bombardment of the vehicle by individual molecules in the outer reaches

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of the atmosphere. For most spacecraft, FMH is only encountered during launch ascent just after

the booster's payload fairing is ejected. Charged-Particle Heating is weak compared to the four

principal environmental heating sources & generally not significant in the thermal design of

room-temperature systems. But at cryogenic temperature charged-particle heating can become a

significant factor in thermal design because of the high sensitivity of such systems to

environmental heat loads. The near-Earth trapped charged particles known as the Van Allen belts

lie about the plane of the geomagnetic equator and feature relativistic electrons and protons. The

spatial characteristics of the Van Allen belts and the spectral properties of the trapped particles

within them undergo both regular and irregular variations with time.

Aerobraking maneuvers are used to make large changes in orbit altitude or inclination and they

are especially useful in slowing down a spacecraft on an interplanetary trajectory to the point

where orbital capture by a planet is possible. Aerobraking occurs when a portion of the orbit

enters a planet's atmosphere, creating aerodynamic drag on the spacecraft. This drag slows the

spacecraft, there by gradually lowering the altitude or changing the orbital plan and it can also

rapidly warm the spacecraft because of friction in the atmosphere. The advantage that

aerobraking provides is placement of the spacecraft into the desired orbit at reduced mass and

cost.

1.4.1 Thermal Surface Finishes

All visible surfaces on the inside and outside of unmanned spacecraft are thermal-control

finishes. 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 is to 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 layer of insulation blankets, radiator coatings, and paints.

Electronics boxes located inside the spacecraft and the structural panels to which they are

attached, are usually painted to achieve a high emittance. Internal temperature sensitive

components that do not dissipate much heat, such as propellant lines or tanks, often have a low-

emittance finish of aluminum or gold.

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1.4.2 Thermal Surface Degradation

Thermal Surface Degradation occurs in orbit by charged particles, ultraviolet (UV) radiation,

high vacuum, and the contaminant films that deposit out on almost all spacecraft surfaces. The

degradation is a problem because of the solar load & change in load. The general result of these

processes is an increase in solar absorbtivity with little or no effect on IR emittance. This is

normally undesirable from a thermal-control standpoint because spacecraft radiators must be

sized to account for the substantial increase in absorbed solar energy that occurs because of

degradation over the mission.

1.5 Insulation

Multilayer insulation (MLI) is the most common thermal-control elements on spacecraft. MLI

blankets prevent both excessive heat loss from a component and excessive heating from

environmental fluxes, rocket plumes, and other sources. Most spacecraft flown today are covered

with MLI blankets, with cutouts provided for areas where radiators reject internally generated

waste heat. MLI blankets are also used to protect internal propellant tanks and propellant lines.

But MLI is not very effective in the presence of a gas.

MLI is composed of multiple layers of polished radiation shield The simplest MLI construction

is a layered blanket assembled from thin (1/4-mil thick) embossed Mylar sheets, each with a

vacuum-deposited aluminum finish on one side. As a result of the embossing, the sheets touch at

only a few points and conductive heat paths between layers are thus minimized. The layers are

aluminized on both sides, so that the Mylar can act somewhat as a low-conductivity spacer.

Higher-performance construction is composed of metalized Mylar films (with aluminum or gold)

on both surfaces with silk or Dacron net as the low-conductance spacers.

Tapes are used to close blanket edges and cutouts to protect seams stitched with organic thread

to provide reinforcement of interior or cover layers in local areas and to aid in electrical

grounding of the blanket. The tape should have the same optical properties as the surface to

which it is applied, but different optical properties are acceptable if they do not significantly

affect the performance of the entire blanket and do not cause temperatures in the vicinity of the

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tape to exceed the allowable material limit or adhesive limits. Surfaces to be taped must be clean

and free of oils to ensure a good bond. Adhesives that are silicone-based are commonly used to

bond fasteners to structure in low-to-moderate-temperature applications. These silicone

adhesives are not used on surfaces directly exposed to the low Earth-orbit environment or on

surfaces expected to reach temperatures of 200°C or more.

1.6 Radiators

Spacecraft excess heat is rejected to outer space by radiator surfaces. Radiators occur in different

forms such as spacecraft structural panels and panels deployed after the spacecraft are on orbit.

Radiators reject heat by infrared (IR) radiation from their surfaces. The radiating power depends

on the surface's emittance and temperature. The radiator must reject both the spacecraft waste

heat plus any radiant-heat loads from the environment or other spacecraft surfaces that are

absorbed by the radiator. Radiators are given surface finishes with high IR emittance (e > 0.8) to

maximize heat rejection and low solar absorptance (a < 0.2) to limit heat loads from the Sun.

1.7 Heaters

Thermal control of a satellite or component is not completely achieved through passive

techniques. Heaters are sometimes required in a thermal design to protect components under

cold-case environmental conditions or to make up for heat that is not dissipated when an

electronics box is turned off. Heaters may also be used with thermostats or solid-state controllers

to provide precise temperature control of a particular component. Another common use for

heaters is to warm up components to their minimum operating temperatures before the

components are turned on.

1.8 Heat Pipes

Heat pipes use a closed two-phase liquid-flow cycle with an evaporator and a condenser to

transport relatively large quantities of heat from one location to another without electrical power.

It can change the flux density of a heat flow and it can function in various ways as a thermal-

control device. The driving mechanism in a heat pipe is capillary pumping, which is a relatively

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weak force that is provided by a wick. But the pipe may be susceptible to severe performance

degradation when operating in a gravitational field. Planning is therefore needed to facilitate the

ground testing of systems that include heat pipes.

1.9 Louvers

Louvers are active thermal-control elements, which are placed over external radiators. They may

also be used to modulate heat transfer between internal spacecraft surfaces or from internal

surfaces directly to space through openings in the spacecraft wall. Louver in its fully open state

can reject six times as much heat as it does in its fully closed state, with no power being required

to operate it. The most commonly used louver assembly are bimetallic, spring-actuated,

rectangular-blade (venetian-blind) type, Hydraulically activated louvers and pinwheel louvers.

Louver reliability can be improved at the design stage by making each louver blade

independently actuated by a bimetallic clock spring. Thus a single-point failure is associated with

one blade, not the entire assembly. It consists of five main elements: baseplate, blades, actuators,

sensing elements, and structural elements.

1.10 Phase-Change Materials

Phase change materials are used to prevent overheating of high-power-dissipating components.

Also the components that are only activated occasionally are prevented from cooling to

temperatures below operational level. The simplest form of PCM thermal control for electronic

components is the one that is used for short-duty-cycle components. The generated heat is

absorbed via latent heat of fusion by the PCM without an appreciable temperature rise of the

component. This kind of system is totally passive and very reliable. A more general application

of PCM thermal control for electronic components is for cyclically operating components.

1.11 Electrochromic material

Electrochromic materials have the peculiarity to change reversibly the optical properties i.e

emissivity or absorbtivity upon insertion and extraction of small ions after application of small

voltages. Electrochromic devices (ECD) have the ability to absorb or reflect light, when a

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voltage is applied between the electrodes of the device. Polycrystalline tungsten oxide (WO3) is

the best suited material for an electrochromic layer for infrared Emissivity modulation of

spacecraft.

These electrochromic devices may be very effectively used to develop the thermal protection

system of the spacecraft. These electrochromic devices can be used to vary the emissivity of the

spacecraft surface and regulate the heat transfer rate from the spacecraft surface to the outer space.

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2. MULTI LAYER INSULATION BLANKETS

2.1 Introduction

Multilayer insulation, also called super-insulation is the best insulation available as far as

spacecrafts are concerned. It has a particular application wherever it is required to minimise the

heat flow to or from a component or to reduce the amplitude of temperature fluctuations because

of time varying solar radiation flux and to minimise temperature gradients. MLI consists of

several layers of alternating low emissive radiation shields (6µm aluminised Mylar) and low

thermal conductivity spacers (75µm polyester net). The spacer minimises contact area thus

minimising conduction across it. It provides with a thermal conductivity of the order of 10-6

W/cm2-k. The top and bottom sheets are kapton sheets aluminised on the inside. These provide

abrasion resistance and can withstand temperatures of upto 350°C without catching fire etc.

2.2 Characteristics of MLI

As mentioned above, MLI provides an effective thermal conductance of the order of 1 X 10-5

W/cm2-k. It is very lightweight with areal mass density being as low as 500 g/m

2. It is flexible

enough to be mounted/attached with any curved surfaces or extended surfaces. The kapton layer

on both sides shows good resistance to fire. This external layer can with stand temperatures of

upto 350°C. While the outer surface may be at temperatures as high as 150°C or as low as -150

°C, the inside is maintained at a comfortable 0 oC to 40 °C. MLI undergoes minimal degradation

on exposure to solar and other radiation fluxes with the TML (total mass loss) being less than 1%

and CVCM (cumulative volatile condensable materials) being less than 0.1%.

2.3 Material selection for MLI

The main components of MLI are radiation shields and spacer. The function of the radiation

shield is to attenuate the radiation and that of spacer to decrease solid conduction. The other

materials required for its fabrication are low conductivity washer and thread for assembling the

MLI components, joining fasteners etc. Electrically conductive tapes for grounding, pressure

sensitive tapes for edge sealing etc.

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The primary requirements of these surfaces and materials are that they should posses the required

optical and physical characteristics and should have good mechanical resistance to shock,

abrasion etc. Also they should withstand the space environment exposure such as high vacuum,

high temperatures, UV radiation and particle bombardment etc. The following gives the

description for the selection of various MLI components and materials, used in our applications.

2.3.1 Radiation shields

The radiation shields are low emittance and highly reflecting foils used to attenuate the incoming

radiation from hot to cold boundaries as it forms the major percentage of total heat transfer

through MLI.

The number of shields and type of shields are selected depending upon the performance and

working environment. The criteria for the selection of radiation shields are as follows.

• Minimum weight.

• The shield should be flexible enough for it to bend or fold it to any required shape

(cylindrical, spherical etc).

• Emissivity should be low.

• It should be effective in the operating environment (-150 to 150o C).

• Thickness should be minimum

• High folding endurance.

• High tensile and shearing strength.

• High di-electric constant.

It is found that gold, silver and aluminium are low emittance materials that can be used either as

coatings for radiation shields or to form the films.

The vapour deposited gold films appear more attractive for reusable vehicle radiation shields,

because of lower emittance and no degradation effects after exposure to moisture. Although gold

coated surfaces exhibit excellent thermal properties, their use is not so widely extended owing to

problems of cost and manufacturing process control. This increased cost is due to the cost of

manufacturing and not due to the gold, which is deposited in very thin layers.

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The silver coated films tarnishes in air. These are coated with SiOX to protect the silver coating.

Although silver coated plastic films with SiOX overcoatings are commercially available, they are

very sensitive to structural changes and are expensive owing to problems presented by

evaporation techniques. Similarly, copper coated films are prone to oxidation in the presence of

atmospheric air. Because of this reason silver and copper coatings are not used commonly.

The aluminium and aluminium coated plastic films are most frequently selected for radiation

shields because the emissivity of aluminium is only slightly higher than clean silver but silver

tarnishes in air, while aluminium forms a very thin protective layer of aluminium oxide which

prevents further degradation of the surface. Also aluminium is inexpensive and readily available

as metal foils of various thicknesses and as a coating on a variety of metallic and non-metallic

surfaces.

In applications, aluminium coated plastic films (polyester, polyamide etc) are used as compared

to aluminium foil of the some thickness. This is because plastic films are light and have very low

thermal conductivity and offers shear strength far superior to that of still thicker metal foil. Also,

the parallel or lateral thermal conductivity of plastic coated films is very low. Most polyester

films, however, are limited to an operating temperature below 150o

C, the point at which the film

begins to deteriorate (Although kapton and Teflon films can be used for temperatures as high as

400o

C) while aluminium foils can be used up to 540o

C.

In the present application, vapour deposited aluminium (about 500 angstroms) on both sides of

0.25 mil Mylar was selected as the radiation shield fulfils all of our requirements. It was

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perforated to 0.5% of the total areas with holes of 1.5 mm diameter spaced 25mm apart in zigzag

fashion.

The 1.0 mil and 3.0 mil kapton, aluminised on one side was used as the top and bottom cover

shields of the blankets. The cover shields provide the blanket the required strength against the

mechanical loads during fabrication, handing and launch, shock and vibration loads and also has

the high resistance to temperature (up to about 400o

C). And it also does not degrade in space

environment. They also provide the required optical properties to the blanket surfaces which are

facing the space. These shields are also perforated to 0.5% of the total surfaces.

2.3.2 Spacers

The spacer is a low thermal conductivity material sandwiched between the radiation shields to

decrease the contact area between parallel neighbouring radiation shields. The commonly used

spacers in multilayered insulation can be broadly classified into four categories, multiple

resistance, point contact, single component and composite. These are described below.

Multiple resistance spacers

Particles in contact exchange heat through their points of contact. Therefore, a mat of fibres

arranged in a parallel array, in which the heat must pass from each fibre to reach the next

radiation shield, can be used as an effective spacer.

Glass fibres, quartz fibres and plastic fibres are used as spacers in MLI systems. Problems in

connection with use of these spacers arise from their very poor dimensional stability.

Point contact spacers

The dimensional stability of the MLI system can be improved by using silk fibre glass or nylon

screens, shoes knots, which can be assimilated to isolated spheres, to space the neighbouring

radiation shields. The mesh size of screen, or distance between the spheres should be as large as

possible. But, an upper limit exists, determined by the requirements that direct contact between

radiation shields produced by their sagging must be avoided.

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Single component MLI

Radiation shields can be separated from each other without using any spacer material.

Embossing and crinkling the shields produces random small area contact that creates contact

resistance high enough to reduce conduction heat transfer to a value comparable to the amount of

heat transferred by radiation through the same assembly of radiation shields. This method of

spacing can be used only with radiation shields of low thermal conductivity such as the sheets of

polyesters films (0.25mil thick) with the metal coating on one side. Also, the embossing pattern

should be deep enough to allow to for material memory; otherwise, the geometrical

characteristics would change with time.

Composite spacers

In composite spacer systems there are several spacer materials, being used to provide some

specific characteristics. For example, a thin film may be used to provide the required flexibility

to the insulation so that it can be mounted properly on curved surfaces etc.

polyester spacer layer

This material may not be important from the point of view of thermal characteristics. The second

material is used to impart the required thermal properties.

Depending upon the required performance any of the above type of spacer materials can be used.

Also, a number of layers of spacers can be accommodated between neighbouring radiation

shields.

The selection of spacer material is guided by the following criteria:

• Insulation effectiveness in working environment.

• System weight.

• Thickness of insulation.

• Method of attachment.

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• Number, shape and size of perforation.

• Assembly procedure.

• Pre-operational storage and shipping conditions.

• Cost effectiveness.

The spacer used for ISRO satellites is 100% polyester screen. It is lightweight, flexible, easy to

handle and gives dimensional stability. Because of the knots provided on the screen, it acts as a

point contact spacer and reduces the heat transfer by solid conduction to the minimum.

2.3.3 Electrostatic discharge grounding of blankets

The MLI blankets fixed on the spacecraft are exposed to charged particles in space. Because of

these charged particles, the external dielectric surfaces of the spacecraft can change to a potential

as high as 11000 V. Although, the voltage gradient at the blanket surface is not known, it is

directed normal to the spacecraft surface, and thus a significant potential is impressed across any

exposed thermal blanket. The highly capacitance nature of blanket shield, high dielectric

constant of aluminized kapton and aluminized mylar, then allows it to store such impressed

potentials. The discharge of the potential gives rise to spark which radiate energy to the

surrounding surfaces of the spacecraft.

2.3.4 Fabrication of Multi Layer Insulation Blankets

• Measure the solar absorbtance and emissivity of the radiation shields samples cut from

each roll.

• Cut the required number of layers of aluminized Mylar and aluminized kapton and 100 %

polyester screen to rough size.

• Perforate the radiation shield using punching fixtures.

• Assemble the layers, by putting first layers of aluminized kapton, kapton side facing

outside and fixing its edges by kapton adhesive tape. The spacer layer are kept over this,

followed by aluminized Mylar layer, likewise the aluminized Mylar layers are put one by

one interspaced by spacer layers. On the top again aluminized kapton layer is put. The

16

assembled layers are held together by stapling the pins at few points on the edges which

are removed later on.

• The strip of aluminium foil with electrically conductive adhesive tape is taken and fixed

on both sides of each layer, on the preselected location of the blankets. The spacer in the

vicinity is cut by scissor. Each of the layers is fixed with tape on both sides.

• Edge of the blanket is sealed

• Mark the locations of velero tape fasteners, cut outs or openings to accommodate sensors,

subsystem packages or any other openings/projections.

• Make the cut outs on blankets using sharp knife edge and seal the edges by folding one of

the top or bottom aluminised kapton layer.

• The stitching is done on preselected positions, using PTFE Teflon washers and nylon

thread at a pitch of 75mm.

• Fix the loop type velero type of 16 mm width on the blanket, using suitable adhesive.

• Rivet the grounding points using aluminium rivets and washers. The electrical wire lead

which is used for grounding is also fixed below head while riveting.

• Pack the blanket separately in polythene bags and put a bag of silica in each bag to

absorb the moisture trapped in the polythene bags.

• Fill the bag with dry nitrogen or argon gas to remove the atmospheric and seal the bags.

• Store the blankets in the clean room.

17

3. OPTICAL SOLAR REFLECTOR

3.1 Introduction

The thermal control of any spacecraft is achieved by totally isolating it from the cold space using

multi-layered insulation (MLI) and providing radiators to allow the dissipation of excess heat

from within spacecraft.

These radiating windows are covered with optical solar reflectors (OSR) or secondary surface

mirrors (SSM). These are also called cooling windows because there function is to minimise the

solar energy input and to emit maximum amount of the excess thermal energy generated within

the spacecraft. This is achieved due to the combination of suitable optical properties namely high

IR emittance (0.78) and low solar absorptance (0.08).

3.2 Structure of OSR

In OSRs, the high reflectance to solar energy is due to the second surface (silver) which is highly

reflecting, whereas the high IR emittance is provided by the quartz substrate which is transparent

to solar wavelength region and opaque in the IR region beyond 4.5 micrometres. Silver coating is

used because of their high spectral reflectance.

However the problems associated with it are its poor

adhesion characteristics to the substrate and

degradation in atmospheric and space environment. The

adherent layer should be transparent and should have

little effect on the reflectance of silver. Tantalum oxide

is used for this purpose as it meets the above

requirements for adherent materials. The ITO coating is used to provide electrical conductivity

and is also sufficiently transparent at the same time. These properties, namely electrical

conductivity and transparency are mutually opposing properties. So a combination of 90%

Indium oxide and 10% Tin oxide is used. This combination provides the best trade off between

these properties as per our requirement.

ITO COATING (1OOO A)

TANTALUM OXIDE COATING (700 A)

SILVER COATING (1000 A)

TANTALUM OXIDE COATING (500 A)

QUARTZ SUBSTRATE

ITO COATING (200 A)

18

3.3 Fabrication of OSR

As is obvious from the structural composition explained above, the process of fabrication of

OSR is basically consists of depositing the coatings of silver (Ag), ITO and tantalum oxide

(Ta2O5) on a suitable substrate such as quartz glass. This is accomplished by the process of RF

sputtering. The hot pressed targets of tantalum oxide (Ta2O5), silver (Ag) and ITO with 99.99%

purity are used for deposition. But before sputtering, the quartz substrate is ultrasonically

cleansed in reagent grade acetone for 10 mins and baked in an oven at 150°C for 30 min.

The chamber used for sputtering is evacuated to a pressure of 6 -8 E-6 torr and flushed with

argon gas (99.99% pure). Argon gas pressure of about 2E-2 torr and substrate temperature of

20°C is maintained throughout the process. The deposition rate of the coating is monitored using

a quartz crystal oscillator. The front surface coatings of Ta2O5, Ag, Ta2O5 and ITO are deposited

as per the diagram given in the previous section. After this is done, the substrate is reversed for

depositing the second surface coating of ITO (200 angstroms). The chamber is maintained under

high vacuum for about 10min after completion of each stage of deposition.

After sputtering, the film is annealed at 200°Cfor 30 min in an oven with hot air circulation

facility.

3.4 Characteristics of OSR

• Effect of annealing on the optical properties: The transmission characteristics of ITO

change as a result of annealing at 225°C for 30 min. While Transmissibility improves

from 89% to 92% after annealing, sheet resistance decreases.

• Effect of radiation: the solar absorptance is found to increase on exposure to radiation

under test conditions. Although the spectral reflectance decreases, but the decrease is a

minimal of around 2-3%. The emittance value of the OSR remains a constant as it an

inherent property of the quartz substrate.

• Environmental stability of OSRs: to evaluate the environmental stability of the coating

during the pre and post environment, the OSRs are subjected to humidity, thermal

cycling, thermovacuum performance, UV and particle irradiation testing. The humidity

19

test reveals the resistance to corrosive environment. Thermal cycling and thermo vacuum

tests are done to observe the effect of on-orbit temperature cycling on the physico-optical

properties of OSR. The OSRs survive severe thermal cycling (900 cycles of -50 to 65°C

per 10 mins), humidity (RH 95%at 50°C for 96 hrs) and thermovacuum performance (10

cycles of -45 to 70°C with soaking time of 2hrs in a vacuum of 10-6

torr ) tests without

any significant degradation in vacuum.

Test Solar absorptance

(before)

Solar absorptance

(after)

Humidity (96 hrs) 0.044 0.045

Thermal cycling(900 cycles) 0.042 0.043

Thermovacuum performance 0.041 0.041

20

4. LOUVERS

4.1 Introduction

Louvers are active thermal-control elements that have been used in different forms on numerous

spacecraft. The louvers have gained a wide acceptance in the Aerospace industry as highly-

efficient devices for controlling the temperature of a satellite. The first louvers were flown in

1960s. Since then, louver units have flown on numerous satellites, including NIMBUS-4, 5, 6 &

7; Landsat-2, 3, 4 & 5; OAO A2 & A4; ATS-6, Viking-1 & 2; Voyager-1 & 2; NAVSTAR/GPS

series; Solar Maximum Mission; AMPTE, SPARTAN, Space Telescope, Magellan, GRO,

UARS, EUVE, TOPEX, GOES, MGS, MSP, MTSAT and TRMM.

In general, a louver in its fully open state can reject six times as much heat as it does in its fully

closed state, with no power being required to operate it. The most commonly used louver

assembly is the bimetallic, spring-actuated, rectangular-blade (venetian-blind) type.

Louver reliability can be improved at the design stage by making each louver blade

independently actuated by a bimetallic clock spring. Thus a single-point failure is associated with

one blade, not the entire assembly. The spring can be integrated with a heater/controller to

decrease the passive closed-to-open temperature range of 10-17°C to as little as 1°C. The

bimetal/heater assembly drives the blade from fully closed to fully open over only a 1 °C

temperature change. The louver begins to open passively (by conduction from the mounting plate

to the bimetallic spring) at about 10°C. This passive opening provides backup if the active

controller fails off.

The main features of louvers include:

• Lightweight.

• Self-contained.

• Available in many proven geometric configurations.

• Maintains temperature control over wide spectrums

• Uses no power.

• Configurable for mission-unique requirements.

• Fully-qualified for various satellite requirements.

• Sun shielded configuration available.

21

4.2 Assembly

Louver radiator assemblies consist of five main elements:

• baseplate,

• blades,

• actuators,

• sensing elements,

• Structural elements.

The baseplate is a surface of low absorptance-to-emittance ratio that covers the critical set of

components whose temperature is being controlled. Blades, which are driven by the actuators,

are the louver elements that give variable-radiation characteristics to the baseplate. While closed,

louvers shield and decouple the baseplate from the surroundings, but while open, they allow a

radiative coupling between the baseplate and the surroundings.

The radiation characteristics of the baseplate can be varied over the range defined by these two

extreme positions. As the temperature increases, the bi-metallic sensor, or actuator, contracts and

applies torque to rotate the blades toward an open position, thereby allowing heat to dissipate. As

the temperature decreases, the actuator expands, causing the blades to rotate to a closed position

so that heat from the baseplate/radiator can be reflected by the highly polished blade surfaces.

The opening and closing of the louver blades continues throughout the orbital flight to maintain

thermal control within a narrow temperature band. Since a pair of louver blades is driven by

independent sensors, local thermal control across the emitting base is afforded.

Louver assemblies have been designed to operate between fully-closed and opened positions in

either a 10°C or 18°C temperature differential. They are capable of operation within an

environmental range of -85°C to +120°C, with a minimum operation capability (open-to-

close/close-to-open) of well over 30,000 cycles with no degradation in performance. Louvers

are lightweight, self-contained, consume absolutely no power, and can be adjusted to maintain

temperature control over wide thermal spectrums.

22

The actuator drives the blade angle as determined by the baseplate temperature. A strong

conductive path between the actuator and baseplate is therefore sought to minimize the

temperature gradient between them. The thermal coupling between a bimetallic actuator and

baseplate is composed of both radiative and conductive paths.

4.3 Type of louvers

Two designs have been used prominently abroad many spacecraft and have performed according

to the requirement. They are:

• Vane louvers.

• Pin wheel louvers.

4.3.1 Vane Louvers

As noted earlier, the most widely used louver assembly is the bimetallic, spring actuated,

Rectangular-blade type “venetian-blind" or "vane" louvers.

Design, Assembly and working

The arrangement of actuators, housing, blades, and structure for a vane louver assembly is shown

in figure above. In most designs, blade rotation is effected by the expansion or contraction of a

spiral bimetallic actuator, by virtue of heat gained or lost in exchange with the equipment-

mounting plate.

23

One end of the actuator is attached to the frame structure and the other to the Teflon

spool. A square cutout in the spool supports the inboard louver-blade end. The actuator is coated

black in order to enhance radiative interchange. The conduction path is through the aluminium

housing. The actuator is adjusted relative to the frame to obtain the desired temperature range

between fully closed and fully open positions.

Each blade of the louver is supported inboard and outboard by a bearing assembly. Inboard, the

Teflon spool bears against and rotates with respect to the aluminium support structure. The

outboard end of the louver-blade shaft rotates within and is supported by a Teflon bearing, with

end play established by the distance between the Teflon thrust pad and the set screw. Each louver

blade consists of a central torque tube bonded to flanges. The louver-blade cross section forms a

hollow, thin-walled rectangle of high aspect ratio. Each louver assembly contains several

independently actuated blades, so a degree of redundancy is inherent in this design approach.

An old design approach employs active control of blade position though a bimetal/heater

assembly. Frame structures are used for the larger louver assemblies, while the smaller

assemblies are frameless. In the latter case, the actuator and the end-support bracket are aligned

and then attached to the equipment mounting plate with a foamed lose out used at the edges. The

blades are supported and centred inboard by the bimetal/heater assembly. The fibreglass shaft

with bonded-on, ball-end pivot is supported outboard by a bushing in the end-support bracket.

The blades, composed of a foam sandwich about the fibreglass quill, have a 1-mil, first-surface-

aluminized Kapton film on each side.

4.3.2 Pinwheel Louvers

Assembly

The pinwheel louver consists of a lobed louver blade, an actuator assembly, a guard ring, and a

special radiator pattern. This type of louver may be selected because of its low mechanical

profile (it is less than 1.28 cm tall) or its tolerance of solar loads. Pinwheel actuator assembly is

shown in detail in Fig. It consists of a bimetallic element, bimetallic heater strip, driveshaft

assembly, bimetallic housing, outer housing, clamp ring, stop element, and two bushings. The

bimetallic heater strip is bonded directly to the bimetallic element, which in turn is bonded into

the bimetallic housing. The driveshaft assembly is attached to the inner coil of the bimetallic

24

element and carries the stop arm and two bearing surfaces that ride in the bushings, one of which

is mounted in the bimetallic housing. The bimetallic housing mounts inside the outer housing,

which contains the stop element and an adjustable bushing.

The whole assembly mounts in a hole in the spacecraft honeycomb-panel external wall and is

held in place with a clamp ring. The louver opens passively through the action of a bimetallic

spring or is driven open by an electronic controller and a small heater on the spring. When fully

open, however, the radiator surface constitutes only 5% of the circular area.

An old pinwheel louver-blade design, shown in detail in Fig, which consists of a fibreglass hub,

foam sandwich blades, a fibreglass support framework, and a single aluminized-Kapton-film

outer shield. The latter shields the hub and blades from most of the external environment. This

protection is necessary to prevent wide variations in hub and blade temperatures, which would

affect the bimetal temperature and thus its response.

Working

The actuator passive set point is adjusted by loosening the clamp ring, rotating the bimetallic

housing, then retightening the clamp ring. The stop elements limit the blade rotation at the fully

closed and fully open positions (45 deg of angular rotation). The actuator operation is the same

as for a vane louver actuator. A temperature change of 15°C is required to drive the louver from

the fully closed position to fully open.

25

The bushings are adjusted at assembly to limit the driveshaft axial movement to 10 mils.

The radiator/guard-ring assembly is shown in detail in Fig. The radiator consists of a guard ring

for louver-blade protection and alternating radiator segments of second-surface aluminized

Teflon and aluminized Kapton.

The Teflon areas are the radiating areas and have a low solar-absorptance value (a < 0.2). The

aluminized Kapton areas are the low emittance areas, which act as insulation when the louvers

are closed. The louver blade covers the Teflon areas when in the closed position and the Kapton

areas when in the open position. Each pinwheel louver had a heat-rejection capacity of

approximately 25 to 30 W when open and a heat leakage of approximately 5 to 7 watts, when

closed. The heat-rejection rate is linearly proportional to the louver-blade position.

26

5. HEAT PIPES

5.1 Introduction

A heat pipe is a heat transfer mechanism that can transport large quantities of heat with a very

small difference in temperature between the hotter and colder interfaces.

Inside a heat pipe, at the hot interface a fluid turns to vapour and the gas naturally flows and

condenses on the cold interface. The liquid is moved by capillary action back to the hot interface

to evaporate again and the cycle is repeated.

A typical heat pipe consists of a sealed pipe or tube made of a material with high thermal

conductivity such as copper or aluminium. A vacuum pump is used to remove all air from the

empty heat pipe, and then the pipe is filled with a fraction of a percent by volume of working

fluid (or coolant) chosen to match the operating temperature. Some of the fluids that are used as

coolant are water, ethanol, acetone, sodium, or mercury.

Due to the partial vacuum that is near or below the vapour pressure of the fluid, some of the fluid

will be in the liquid phase and some will be in the gas phase. Having a vacuum eliminates the

need for the working gas to diffuse through another gas and so the bulk transfer of the vapour to

the cold end of the heat pipe is at the speed of the moving molecules. The only practical limit to

the rate of heat transfer is the speed with which the gas can be condensed to a liquid at the cold

end.

Inside the pipe's walls, an optional wick structure exerts a capillary pressure on the liquid phase

of the working fluid. This is typically a sintered metal powder or a series of grooves parallel to

the pipe axis, but it may be any material capable of exerting capillary pressure on the condensed

liquid to wick it back to the heated end. The heat pipe may not need a wick structure if gravity or

some other source of acceleration is sufficient to overcome surface tension and cause the

condensed liquid to flow back to the heated end.

Heat pipes use a closed two-phase liquid-flow cycle with an evaporator and a condenser to

transport relatively large quantities of heat from one location to another without electrical power.

A heat pipe can create isothermal surfaces; as a thermal "transformer," it can change the flux

density of a heat flow; and it can function in various ways as a thermal-control device. One-way

27

(diode) heat pipes have been tested and flown, as have variable-conductance heat pipes

(VCHPs), which maintain a constant-temperature evaporator surface under varying load

conditions. Because the driving mechanism in a heat pipe is capillary pumping, a relatively weak

force that is provided by a wick, the pipe may be susceptible to severe performance degradation

when operating in a gravitational field. Planning is therefore needed to facilitate the ground

testing of systems that include heat pipes.

A heat pipe typically consists of a sealed container lined with a wicking material. The container

is evacuated and back filled with just enough liquid to fully saturate the wick. Heat pipes operate

on a closed two phase cycle and only pure liquid and vapor are present within the container,

working fluid will remain at saturation conditions as long as the operating temperature is

between the triple point and the critical.

A heat pipe consists of three distinct regions: an evaporator or heat addition region, a condenser

or heat rejection region, and an adiabatic or isothermal region. When heat is added to the

evaporator region of the container, the working fuel present in the wicking structure is heated

until it vaporizes. The high temperature and corresponding high pressure in this region cause the

vapor to flow to the cooler condenser region, where the vapor condenses, giving up its latent heat

of vaporization. The capillary forces exciting in the wick structure then pump the liquid back to

the evaporator.

The wicking structure has two functions in the heat pipe operation: it is the vehicle through

which, and provides the mechanism by which, the working fluid is returned from the condenser

to the evaporator and also ensures that the working fluid is evenly distributed over the evaporator

surface.

28

Because the latent heat of vaporization of most heat-pipe working fluids is high, only small

amounts of fluid need to flow to transport significant quantities of heat. The driving mechanism,

the temperature difference between the evaporator wall and the condenser wall, is also small.

5.2 Limitations on the transport capacity

Heat pipe performance and operation are strongly dependent on shape, working fluid, and wick

structure. The fundamental phenomenon that governs the operation, arises from the difference in

the capillary pressure across the liquid-vapor interfaces in the evaporator and condenser. When

this capillary pressure is not sufficient to promote the flow of the liquid from the condenser to

the evaporator, the heat pipe is said to have reached its capillary limit and dry out of the

evaporator wick occurs.

During steady state operation, several other important mechanisms can limit maximum amount

of heat that a heat pipe can transfer. Among these are the viscous limit, sonic limit, entrainment

limit, and boiling limit. The capillary limit and the viscous limit deal with the pressure drops

occuring in the liquid and vapour phases, respectively.

The sonic limit results from the occurrence of choked flow in the vapor passage, while the

entrainment limit is due to the high viscous liquid vapour shear forces developed as the vapour

flows in the counterflow direction over the liquid saturated wick. All these limits are axial heat

flux limits, that is, they are function of axial heat transport capacity of the heat pipe.

The boiling limit, however, is a radial heat flux limit and is reached when the heat flux applied in

the evaporator is so high that nucleate boling accors in the evporator wick. This creates vapor

bubbles that partially block the return of fluid and may ultimately lead to premature dryout of the

evaporator.

For moderate temperature heat pipes, the most significant of these limits is usaully the capillary

wicking limit. However, the significance of this limit decreases some what for reduced gravity

applications. In low temperature applications such as those using cryogenic working fluids,

either the viscous limit or capillary limit occur first, while in high temperature heat pipes, such as

those using liquid metal working fluids, the sonic limit and entrainment limit are of increased

importance.

29

5.2.1 Capillary Pumping Limit

Return flow of liquid from the condenser to the evaporator is caused by differences in the

capillary pressure between the evaporator and condenser. The capillary pressure acting on the

liquid surface is inversely proportional to the radius of curvature of the fluid surface at the

liquid/vapour interface in the wick. For purposes of the analysis, the liquid surface in the

condenser is usually assumed to be flat, so that the radius of curvature (and hence the capillary

force) is zero. As liquid evaporates, the meniscus in the evaporator depresses, causing a

difference in capillary pressure between the evaporator and condenser surfaces. This difference

in pressure pulls liquid through the wick from the condenser to the evaporator in an attempt to

restore equilibrium.

A heat pipe "dries out" when the flow of working fluid through the wick caused by this pressure

difference is insufficient to supply liquid at the same rate at which working fluid is being

vaporized in the evaporator. This point is illustrated in the equation which balances the pressure

of the system:

ΔΔΔΔPcapillary - ΔΔΔΔPgravity = ΔΔΔΔPliquid + ΔΔΔΔPvapour

30

In the above equation, ∆Pcapillary i.e. capillary pressure rise is the maximum possible difference in

capillary pressure between the evaporator and the condenser. This term is a function of the

surface tension (which depends on the choice of working fluid and the temperature) and the wick

pore size (which depends upon the wick material and type of wick).

∆Pgravity i.e. gravity head loss is the "head loss" that must be overcome by capillary pressure to

sustain fluid in the evaporator. In addition to gravity, other accelerations, such as those on a

spinning spacecraft, affect the value of this term.

∆Pliquid i.e. liquid pressure drop is the pressure loss resulting from viscous flow through the wick.

This term is simple for an axial-groove wick, but it can become extremely complicated for a

composite-artery wick, where viscous pressure losses in liquid flowing through complicated

structures of layered screens, metal felt, or sintered powder must be modelled. Expressions for

these losses usually contain empirical constants, which is one of the reasons why performance

testing of each pipe is usually necessary.

∆Pvapor i.e. vapour pressure drop is the pressure loss resulting from vapour flow from the

evaporator to the condenser. This term is usually small unless the vapour density is very low or

the vapour velocity is high because of constricted vapour space.

5.2.2 Viscous Limitation

At very low temperatures, the vapour pressure difference between the evaporator and the

condenser regions of a heat pipe may be extremely small. In some cases, the viscous forces

within the vapour region may actually be larger than the pressure gradients caused by the

imposed temperature field. When this occurs, the pressure gradients within the vapour region

may not be sufficient to generate flow and vapour may stagnate. This no flow or low flow

condition in the vapour portion of a heat pipe is referred to as the viscous limitation. Because the

vapour pressure typically must be very low for this to occur, the viscous limit is most often

observed in cryogenic heat pipes, heat pipes with extremely long condenser regions, or heat

pipes undergoing startup from a frozen state.

31

5.2.3 Entrainment Limitation

In heat pipes, the liquid and vapour flows in opposite directions. The interaction between the

counter current liquid and vapour flow and the viscous shear forces occurring at the liquid-

vapour interface may inhibit the return of liquid to the evaporator. When this occurs, the heat

pipe is said to have reached the flooding limit. Further increase in heat input results in increased

vapour velocities, which may cause the liquid flow to become unstable. In most cases, waves

may form and the interfacial shear forces may become greater than the liquid surface tension

forces, resulting in liquid droplets being picked up or entrained in the vapour flow and carried to

the condenser. This entrainment of liquid droplets can limit the axial heat flux and is referred to

as the entrainment limit.

5.2.4 Boiling Limitation

The "boiling limit" or "heat-flux limit" is concerned with the flux density of the thermal load on

the evaporator. Even if the heat-pipe wick could theoretically return the liquid from the

condenser required by the heat load, if the load is concentrated in too small an area, nucleate

boiling can occur in the evaporator wick. The creation of bubbles in an otherwise filled wick

reduces the area of the wick available for fluid flow, and hence reduces the capacity of the wick.

5.3 Thermodynamic Considerations

If operation near the freezing point is needed--as would be the case for water at typical room

temperatures, for almost any cryogenic liquid, or for liquid metals at start-up--high vapour

velocities and large vapour-pressure drops will be encountered, because in these situations the

vapour density and pressure are very low. These large pressure drops cause their own

temperature drops in the pipe (because saturation temperature is a function of pressure). In some

cases, the pressure drop in the vapour required to support the calculated heat-pipe capacity would

result in a negative vapour pressure in the condenser, an obvious impossibility. Under similar

low-density conditions, choked flow (the "sonic limit") has been observed in liquid-metal heat

pipes.

32

Although not a true limit, the operating temperature of the heat pipe rises, so that the

thermal equilibrium can be established, which may cause the temperature to rise beyond the

desired range? In short, a heat pipe that runs in a temperature regime where its working fluid has

a very low vapour pressure must not be designed. If the relative velocity of liquid and vapour is

high enough (as measured by the Weber number), liquid can be pulled out of the wick and

returned to the condenser in the form of droplets entrained in the vapour. It is an operating limit

in that, to support a given rate of heat transfer from the evaporator, an excess of liquid must be

pulled through the wick, because not all of the liquid will reach the evaporator.

5.4 Types of Heat Pipe

5.4.1 Constant-Conductance Heat Pipe

This is the most basic heat pipe and consists of a working fluid, a wick structure, and an

envelope. This pipe is used to move heat from one location to another (possibly by changing the

flux density in the process) or to isothermalize a surface. It need not be shaped like a

conventional cylindrical pipe--flat plates several feet across have been built and tested as heat

pipes for special applications. Constant-conductance heat pipes are often categorized according

to the type of wick structure they use.

Groove Wicks

The simplest heat-pipe wick design consists of axial grooves in the wall of extruded aluminium

tubing. Grooves can be formed in tubes of other materials, such as copper (by swaging) or even

refractory metals (by deposition), but they are formed most often in tubes of aluminium. This

class of wick is very susceptible to gravitational effects during ground testing, but it is relatively

inexpensive to produce and its performance is very consistent. Its moderate heat-transfer

capability is sufficient for many applications. Most grooves are rectangular or trapezoidal, but

some have more complex shapes, such as the "teardrop" or "keyhole," which can be extruded

with difficulty

33

"Monogroove" Design

The monogroove design is a high-capacity design consisting typically of a wick in one large

teardrop-shaped groove connected to a vapour space, and can be considered an extension of the

basic groove concept.

Unlike a heat pipe with many smaller grooves of the same total area, the monogroove heat pipe

has a large single groove that provides relatively unrestricted longitudinal flow. Liquid is

distributed on the evaporator wall by means of a secondary wick consisting of small

circumferential grooves or screen. This design has shown very high capacity during ground

testing, but it encountered difficulties during early shuttle testing. Later experiments were more

successful.

34

Composite Wicks

Among composite wicks, the simplest consists of several layers of screen fastened to the inside

wall of a heat pipe. More capacity can be obtained by using more layers of screen, to increase the

wick flow area at the cost of increasing the heat-pipe temperature difference resulting from the

temperature drop needed to conduct heat through the thick saturated wick. To overcome this

penalty, some heat-pipe manufacturers separate the wick into two parts, the portion that spreads

the fluid circumferentially about the wall of the evaporator, and the portion that carries the fluid

down the length of the heat pipe.

The former, kept as thin as possible, can consist of circumferential grooves cut in the wall of the

heat pipe or of a single layer of screen or metal mesh bonded to the wall. The latter is held off the

wall by means of legs or straps, or makes contact with the wall in only a few places. This type of

wick has capacities similar to the axially grooved heat pipe, but has much more capability when

tilted. Because the wick must be assembled of relatively fragile materials, care is required in

building such a pipe, and no two supposedly identical pipes will perform in exactly the same

manner. Sample wick designs of this type are shown in figure.

Artery and Tunnel Wicks

This class of heat pipe wick is based on the composite wick, but provides one or more relatively

unrestricted liquid-flow paths in parallel with the longitudinal wick. These paths will fill with

fluid in space, because of minimum surface-energy considerations, and will greatly reduce the

viscous pressure drop in the heat pipe, thereby increasing capacity. When properly designed,

these arteries will fill as the heat pipes operate in a gravitational field.

35

Wicks in this class can be blocked by bubbles of non-condensable gas in the arteries but they are

attractive because of their large heat-transfer capability in a small envelope. If the liquid in the

artery remains subcooled when it reaches the evaporator, bubble formation can be avoided.

A number of mechanical schemes have been proposed and tested to prevent bubbles from

blocking the arteries of VCHPs. These pipes are particularly prone to bubble formation because

the liquid in the artery contains dissolved control gas, which tends to come out of solution as the

liquid warms during its transit of the pipe from condenser to evaporator. Cross sections of some

of these wick structures are shown in the figure.

5.4.2 Diode Heat Pipes

A constant-conductance heat pipe can be modified so that operation occurs normally in one

direction but ceases when an attempt is made to transfer heat in the other, "wrong" direction,

resulting in a diode action. Even when blocked, however, the pipe transfers some heat, if only by

conduction down the pipe wick and wall. This type of heat leak is particularly significant in

cryogenic systems. Common diode heat pipes are the liquid-trap, liquid-blockage, and gas-

blockage diodes.

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Liquid-Trap Diode

The most common type of heat-pipe diode, the liquid-trap diode has a wicked reservoir at the

evaporator end designed so that it is heated by the same environment that heats the evaporator.

Although the envelopes are connected, the reservoir wick is not connected to the rest of the heat

pipe. When, during normal operation, heat is applied to the evaporator and reservoir, heat is

transferred from the evaporator to the condenser as in the constant-conductance heat pipe, and

any fluid in the reservoir wick evaporates and joins the vapour flow to the condenser (the

reservoir wick should be dry during normal operation).

When ends of this pipe are reversed, and the evaporator and reservoir become cooler than the

condenser, some of the hot vapour coming from the condenser condenses in the reservoir and is

lost to the rest of the heat pipe. Sufficient liquid is tied up in the reservoir to cause the pipe to dry

out. "Shutoff' is neither instantaneous nor complete. A schematic of the operation of this type of

diode is shown in figure.

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Liquid-Blockage Diode

At its condenser end, the liquid-blockage diode has a wicked reservoir cooled by the same

environment that cools the condenser. The reservoir's wick is not in contact with that of the

remainder of the heat pipe, and it is normally full of working fluid- in effect it traps a large fluid

slug. When the ends of the pipe are reversed, the fluid slug travels to the normal evaporator end,

where it completely fills the evaporator vapour space (and that of a large portion of the transport

section), preventing condensation.

Optimum design of the wick structure and vapour space must be compromised to control the

liquid slug during shutoff, and such control requires maintaining close tolerances during the

manufacturing process. Proper control of the fluid (and therefore operation of the diode) in a

gravitational field requires maintaining the gap between the evaporator wall and the blocking

plug at a size that enables the gap to fill with liquid if it is available.

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Gas-Blockage Diode

The gas-blockage diode is similar in design to the liquid-blockage diode, except the reservoir,

which can be unwicked and contains a non-condensable gas. When the ends of the pipe are

reversed, the gas flows to the evaporator and, as above, completely fills the vapour space,

preventing condensation. However, as the temperature rises, the gas slug can be compressed to

the point where the heat pipe will start working again. Furthermore, convection within the gas

slug may be a significant heat-leak component. A schematic of the operation of this type of diode

is shown in the figure.

5.5 Working Fluids

The choice of working fluid is usually governed by the temperatures of the desired operating

range. A heat-pipe working fluid can be used effectively between a temperature somewhat above

its triple point and another that is below its critical temperature. If the triple point is approached

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too closely, temperature drop in the vapour flow increases. As the critical point is approached,

the distinction between liquid and vapour blurs, and the surface tension drops to zero. The

pressure that must be contained by the envelope also increases significantly. The triple points

and critical temperatures of several heat-pipe working fluids are given in Table

Two parameters have been developed to aid in comparing the relative performance of heat-pipe

working fluids.

The "zero-g figure of merit," is given by σρλ/µ, where σ is the surface tension, ρ is the liquid

density, λ is the latent heat of vaporization, and µ is the dynamic viscosity. This parameter

neglects vapour flow entirely, but for most applications, vapour flow is not the limiting factor.

The group of fluid properties included in the parameter definition appears in the heat-pipe

capacity equation.

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A second parameter, the "one-g figure of merit" or "wicking height factor," compares the relative

sensitivity to gravity effects of working fluids: σ/ρ, where the properties are as defined above. It

is a relative measure of how high a given wick structure will be able to pump a working fluid in a

gravitational field (or as a result of inertia effects, as in a spinning spacecraft).

5.6 Material Compatibility

Because a heat pipe is a completely sealed container, any chemical reactions between the

working fluid and the wall or wick material can be disastrous. None of the reaction products can

escape, and any material that is consumed cannot be replaced. Certain combinations of materials,

such as ammonia and copper, are known to react quickly with one another, and hence are not

likely to be chosen.

Many combinations of materials that are traditional and acceptable in the chemical-process

industry (such as water and stainless steel, or water and nickel) have been demonstrated to react

with one another, generating non-condensable gas. In general, the cryogenic working fluids up

through ammonia can be used with either stainless steel or aluminum (although some evidence

indicates that ammonia reacts slowly with aluminum, and the combination of ammonia,

aluminum [as is found in a wall material], and stainless steel [such as would be found in a typical

wick material] can react more quickly with one another).

Methanol works well with stainless steel but reacts with aluminum. Water seems to work well

with copper, and possibly monel, but not with 304 or 316 stainless steel or nickel. Some short-

term success has been achieved with carbon steel, but pipes using it appear to be generating

hydrogen gas, which diffuses through the pipe wall; this observation indicates an internal

reaction is taking place.

Materials available for higher-temperature (liquid-metal) heat pipes must hold together at those

higher temperatures and be inert to some very corrosive working fluids. This area is still under

investigation.

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5.7 Testing of Heat pipes

5.7.1 During Fabrication

The heat-pipe envelope will be checked for leaks during the fabrication process, usually with a

helium mass-spectrometer leak detector. However, once the pipe is sealed at the fill tube, the

integrity of this seal is open to question. Although some chemical tests have been used, the most

thorough seems to be checking for the presence of working fluid outside the heat pipe when it is

placed in an evacuated chamber.

Performance of each heat pipe as a function of tilt should be measured at some typical operating

temperature(s) to determine whether the wick functions properly. Testing at a low temperature

will show whether non-condensable gas is present. (At high temperatures, the non-condensable

gas can be compressed into a thin plug so that it isn't detectable using thermocouples mounted on

the heat pipe).

If the heat pipe is to be installed in a spacecraft in a position where it will be tested vertically

(with gravity assist) during system-level testing, such as a thermal-vacuum or thermal-balance

test, then it must be tested in the same orientation with a similar heat load before installation. In

this way, the performance of the heat pipe that will be seen in the vacuum chamber will be

known before the test is performed.

This data will be helpful at the time when the heat pipe can't be reached without breaking

vacuum and tearing open the spacecraft. In the case of a heat pipe that is to be curved in three

dimensions and can't be tested in a single plane, some manufacturers build a test pipe with the

same number of curves in the wick, but with all of the curves in a single plane. In this way, the

wick performance to be expected in space can be characterized.

5.7.2 After Integration into the System

After integration of a heat pipe into a system, the heat pipe should be verified to determine

whether any deterioration took place during the integration procedure, and also to verify the

performance of the integrated thermal control system.

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5.8 Heat-Pipe Applications and Performance

The most obvious application of a heat pipe is one requiting physical separation of the heat

source and sink. If a heat pipe is used, all hardware to be cooled need not be mounted directly on

radiator panels, and relatively inefficient conductive couplings need not be used. By the same

token, heaters need not be mounted directly on hardware to be heated if a heat pipe is employed.

A closely related class of applications is that of the thermal transformer.

In this scenario, a small high-powered box is mounted on one side of a radiator with integral heat

pipes; the heat generated is spread and dissipated at a much lower flux density over the entire

surface of the radiator. Heat pipes have been used to reduce temperature gradients in structures to

minimize thermal distortion.

The diode heat pipe was first proposed as a means of connecting a device to two radiator panels

on opposite sides of a spacecraft, with the understanding that at least one of the radiators would

be free of any direct solar load at all times during the orbit. The diodes would couple the device

to the cold radiator, while preventing heat from leaking back into the system from the radiator in

the sun. This type of thermal-design problem in which, heat from a temporarily warm radiator or

from a failed refrigerator must be kept from leaking back into the system.

The VCHP can control the amount of active radiator area, providing reasonably good

temperature control without the use of heaters. This capability is particularly attractive if

electrical power is limited, and this type of design has been flown on a number of satellite

experiments. However, if the application requires maintaining a box or baseplate at a virtually

constant temperature, feedback control may be employed.

A sensor on the baseplate of the device to be controlled is routed to an onboard computer, and

whenever the temperature drops below the desirable range, heaters on the VCHP reservoirs are

activated, causing the control gas to expand and block off more of the radiator area. If the

temperature rises above the range desired, power to the reservoir heaters is reduced, increasing

the active radiator area. This concept usually requires less power than the direct use of heaters on

the box or system to be controlled.

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The use of flexible heat pipes or rotatable joints in heat pipes to cool devices on rotating

or gimballed platforms has been proposed, but flexible heat pipes tend to have too much

resistance to motion, and rotating joints in heat-pipe walls leak under extreme conditions. These

areas are still under active investigation.

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6. THERMAL CONTROL PAINTS

6.1 Introduction

The external surfaces of a spacecraft radiatively couple the spacecraft to space. Because these

surfaces are also exposed to external sources of energy, such as sunlight and Earth-emitted IR,

their radiative properties must be selected to achieve an energy balance at the desired

temperature between spacecraft internal dissipation, external sources of heat, and radiations to

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

absorptance.

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.

While space-qualified paints are available in a variety of colours, black and white are by far the

most commonly used. Almost all paints have a high emittance, so the choice is really between

solar absorptance (and its degradation in the space environment), ease of application, and