microsats tc baturkin

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Acta Astronautica 56 (2005) 161–170 www.elsevier.com/locate/actaastro Mic ro-s atellites thermal control —con cept s and comp onents Volodymyr Baturkin  National Technical University of Ukraine, Kyiv Polytechnic Institute, Peremogy Pr., 37, Kyiv 03056, Ukraine Abstract The main idea of this paper is to present the survey of current tendencies in micro-satellites thermal control concepts that can be rational and useful for posterior missions due to intensive expansion of satellites of such type. For this purpose, the available references and lessons learned by the National Technical University of Ukraine during the elaboration of thermal control hardware for micro-satellites Magion 4, 5, BIRD and autonomous thermal control systems for interplanetary missions VEGA, PHOBOS have been used. The main parameters taken into consideration for analysis are the satellite sizes, mass, power consumption, orbit parameters, altitude control peculiarities and thermal control description. It was dened that passive thermal control concepts are widely used, excepting autonomous temperature regulation for sensitive components such as batteries, high-precision optics, and some types of sensors. The practical means for realization of passive thermal control design as multi-layer insulation, optical coatings, heat conductive elements, gaskets are briey described. © 2004 Elsevier Ltd. All rights reserved. 1. Intr oduct ion Last years conf eren ces on small satellites acti vi- ties [1,2], the information about currently elaborated proj ects and futu re small -satellite missions plann ed [3] emphasized the actuality of satellites and active int erest in the m. The intensive expan sio n of sma ll sat ell ite s cou ld be ex pla ine d by moder ate eno ugh cos t, short time of ela borat ion and exist ing possi- bility to include the complicated devices like multi- functional equipment and optical systems as payload. Accor ding to small -sate llite classi cati on [3], there are the following conditional groups: nano- and pico- satellites (< 10kg), mic ro-sa tel lit es (10–1 00kg), T el ./fax: +380 44 24175 97.  E-mail address: [email protected] (V. Baturkin). 0094-57 65/$ - see front matter © 2004 Elsev ier Ltd. All rights reserv ed. doi:10.1016/j.actaastro.2004.09.003 min i-s ate lli tes (100–500 kg), int erp lan eta ry sma ll missions (< 500 kg). T able 1 presents some tec h- nic al dat a for the abo ve con sid ere d sma ll- sat ell ite groups. Certain constraints originating in small-satellite de- signs due to limited mass and power denite available volume for payload and housekeeping systems pro- duce the set of requirements for each of the satellite systems and for the thermal control system as well. The survey of thermal means used in practice in small- satellite designs, visible perspectives and difculties can be useful to be more oriented in this topic. Main attention is devoted to small-satellite group with mass less and near 100 kg (micro-satellites) , where compro- mise in the power/mass/volume distribution between mandatory housekeeping needs and the payload re- quirements is actual.

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Acta Astronautica 56 (2005) 161–170

www.elsevier.com/locate/actaastro

Micro-satellites thermal control—concepts and components

Volodymyr Baturkin∗

  National Technical University of Ukraine, Kyiv Polytechnic Institute, Peremogy Pr., 37, Kyiv 03056, Ukraine

Abstract

The main idea of this paper is to present the survey of current tendencies in micro-satellites thermal control concepts that

can be rational and useful for posterior missions due to intensive expansion of satellites of such type. For this purpose, the

available references and lessons learned by the National Technical University of Ukraine during the elaboration of thermal

control hardware for micro-satellites Magion 4, 5, BIRD and autonomous thermal control systems for interplanetary missions

VEGA, PHOBOS have been used. The main parameters taken into consideration for analysis are the satellite sizes, mass,

power consumption, orbit parameters, altitude control peculiarities and thermal control description. It was defined that passive

thermal control concepts are widely used, excepting autonomous temperature regulation for sensitive components such as

batteries, high-precision optics, and some types of sensors. The practical means for realization of passive thermal control

design as multi-layer insulation, optical coatings, heat conductive elements, gaskets are briefly described.

© 2004 Elsevier Ltd. All rights reserved.

1. Introduction

Last years conferences on small satellites activi-

ties [1,2], the information about currently elaborated

projects and future small-satellite missions planned

[3] emphasized the actuality of satellites and active

interest in them. The intensive expansion of small

satellites could be explained by moderate enough

cost, short time of elaboration and existing possi-

bility to include the complicated devices like multi-functional equipment and optical systems as payload.

According to small-satellite classification [3], there

are the following conditional groups: nano- and pico-

satellites (< 10kg), micro-satellites (10–100 kg),

∗ Tel./fax: +380 44 24175 97.

  E-mail address: [email protected] (V. Baturkin).

0094-5765/$ - see front matter © 2004 Elsevier Ltd. All rights reserved.

doi:10.1016/j.actaastro.2004.09.003

mini-satellites (100–500 kg), interplanetary small

missions (< 500 kg). Table 1 presents some tech-

nical data for the above considered small-satellite

groups.

Certain constraints originating in small-satellite de-

signs due to limited mass and power definite available

volume for payload and housekeeping systems pro-

duce the set of requirements for each of the satellite

systems and for the thermal control system as well.

The survey of thermal means used in practice in small-satellite designs, visible perspectives and difficulties

can be useful to be more oriented in this topic. Main

attention is devoted to small-satellite group with mass

less and near 100 kg (micro-satellites), where compro-

mise in the power/mass/volume distribution between

mandatory housekeeping needs and the payload re-

quirements is actual.

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162 V. Baturkin / Acta Astronautica 56 (2005) 161–170

Table 1

Attributes of small-satellite groups

Satellite class Mass (kg) Bus linear sizes (m) Power averaged (W)

Mini 100–500 More than 1 Up to 100

Micro 10–100 0.5–1 Tens

Nano Less than 10 Less than 0.2 Several

0

5

10

15

20

25

30

35

        1        9        8        0

        1        9        8        1

        1        9        8        2

        1        9        8        3

        1        9        8        4

        1        9        8        5

        1        9        8        6

        1        9        8        7

        1        9        8        8

        1        9        8        9

        1        9        9        0

        1        9        9        1

        1        9        9        2

        1        9        9        3

        1        9        9        4

        1        9        9        5

        1        9        9        6

        1        9        9        7

        1        9        9        8

        1        9        9        9

        2        0        0        0

        2        0        0        1

        2        0        0        2

        2        0        0        3

number of

micro satellites per year

(by SSHP 2002)

Fig. 1. Launches of micro-satellites in years [3]. Dash line is a

trend linear approximation of flying micro-satellites.

The paper does not cover the analysis of all thermal

control principles used and their technical embod-

iment, as according to [3] (Fig. 1) more than 250micro-satellite launches have been realized during

1980–2000, and each of them has something spe-

cific. An additional difficulty is the insufficiency of 

available descriptions of thermal control details.

2. Typical micro-satellite thermal concepts

Main principles of space thermal control design,

software and hardware used are collected in [4–11] and

many other works. Nevertheless, namely, the details of thermal control system (TCS) design are mostly im-

portant in the case when they are embodied in practi-

cally realized projects.

The diverse information about thermal designs is

collected by Surrey Satellite Technology Ltd., [2,

p. 417], Technical University of Berlin [2, p. 347],

by Design Bureau “Pivdenne” [1, p. 437] and other

leading small-satellite space hardware manufacturers

as well. Summarizing, it is possible to emphasize the

following regularities.

2.1. External heat exchange

Micro-satellites operate on low Earth circular orbits

(450–1200 km) with wide range of NASA angles

and on high elliptical orbits. This means that satellite

sides can be exposed to external disturbances as IR

Earth radiation, Albedo, Sun. IR and Albedo can benegligible for highly elliptical orbits. The typical max-

imal incident fluxes for 550 km orbit for flat surface

with normal to nadir are qIR ∼ 200W/m2, qALB,MAX

∼ 450W/m2 (averaged over orbit < 150W/m2). The

eclipse time can reach 0.5 h (circular) up to several

hours for elliptic orbits. Most often used attitude con-

trol systems (ACS) are spin attitude control, gravity

and 3-axes stabilization. This means that the satellites

external light disturbances can be predicted due to ex-

act knowledge of its attitude. Spin attitude presumes

the arrangement of rotation axes perpendicular or par-allel to Sun direction; ACS for 3-axes stabilized satel-

lites has to be designed to change, often enough, the

satellite orientation along orbit movement to decide

the intended tasks.

2.2. Thermal concepts

There are three conditional approaches in the

thermal scheme design—autonomous concept, cen-

tralized concept and combined concept (Fig. 2).

Autonomous concept assumes the individual thermalcontrol for each device or group of devices. The

devices/equipment are thermally disconnected from

each other. The centralized concept guesses the tight

thermal coupling of all units and the use of one

centralized radiator for external heat exchange. The

most exploited concept is combined concept, where

some group of equipments (for example, housekeep-

ing equipment, satellite bus) has a common thermal

control, several devices (some payload) are thermally

disconnected from satellite bus and use autonomous

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V. Baturkin / Acta Astronautica 56 (2005) 161–170 163

Fig. 2. Conditional thermal concepts of micro-satellites.

thermal control. The thermal concept defines the way

length of heat roaming between the device and the

aimed radiator and effective thermal mass of satellite

and units. Each of the concepts has its own peculiari-

ties, depending on the satellite orientation and attitude

control. For spinning satellite with perpendicular po-

sition of rotation and Sun light axes (Fig. 2) the exter-

nal sides can be used as a thermal sink, as with a ratios/< 1 the wall temperature will trend to the level

of 10–20 ◦C. The solar cells have the similar optical

characteristics, and they can also be used as a thermalsink. For coinciding longitudinal and Sun light axes

(Fig. 2, spinning and 3-axes stabilized satellites) lat-

eral satellite surfaces can be used for heat rejecting.

The heat removing can be arranged through one cen-

tralized radiator (all thermal lines go to it) or several

radiators distributed. The tight thermal connection of 

all components enlarges the thermal mass of satellite,

reduces the temperature non-uniformity throughout

the satellite. Estimating that at power 40W the area

of the radiator is only 0.15–0.2 m2 and the rest of the

lateral area should be covered with multi-layer insu-

lation MLI (1), area ∼ 0.5–1m2. Satellite radiator

temperature is sensitive to unsteadiness of the power

rejected, and roughly 1W in energy variation causes

1–1.5 K temperature change. The coating of the front

satellite area with optical selective coating (OSC)

with s ∼ 0.2, ∼ 0.85 allows to reach the radiator

temperature level of 0–30 ◦C.

Predominant design assumes non-hermetic shell of 

satellites that requires conductive and radiative means

for a heat transfer.

2.3. Inner thermal tasks

Typical requirements of satellite inner components

are collected in Table 2 [8]. The most delicate devices

are batteries, optic instrumentation and individual pay-

load. Requirements of average heat generation inside

the satellite are in the range of 15–40W. This power

is produced mainly by housekeeping equipment, peak 

heat generation coincides with payload operation and

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164 V. Baturkin / Acta Astronautica 56 (2005) 161–170

Table 2

Typical thermal requirements

Component name Temperature (◦C) Peak power (W)

Electrical equipment −10 to +40 Max 10 per unit

Batteries −5 to +15 Up to 20

Consumable gas +9 to +50 —

Microprocessors −5 to +40 Up to 20

Microprocessors −5 to +40 Up to 20

Bearing mechanisms −5 to +40 —

Solar cells −60 to +55 —

Solid-state diodes −60 to +90 —

Orientation sensors −5 to +45 Max 5

Optics 21 Max 10

Payload Individually < 150

Satellite total 250, average 30–60

can reach 200 W; typical housekeeping heat generated

components are board computer, transmitter,ACS, and

batteries. If payload operates constantly, the power

generation does not change abruptly.

The most temperature sensitive units are battery

sets, which should be used in the range of −5– +15 ◦C.

More details about satellite component exploitation

temperatures are presented in [4–10]. Overwhelming

majority of the satellites are built as non-sealing ob-

 jects, and heat transfer inside is realized by conduc-tion of structure elements and radiation (black surface

coating ∼ 0.85). Heat pipes (HP) and high conduc-

tive carbon materials are effective means to improve

the temperature uniformity.

2.4. Satellite structure concept 

Two tendencies are noticeable. The first: payload

and housekeeping are developed and arranged by one

institution. In this case it will provide with the thermal

solution for the whole satellite. The second tendency:the multimission satellite bus with changeable payload

is proposed to the consumer. That means that pay-

load should be accommodated into the satellite thermal

surroundings. The solar arrays may cover the satel-

lite outer surface or could be deployable. After open-

ing they can not be re-oriented. The modern micro-

satellite will carry the high-precision optics, which is

sensitive to satellite structure and geometrical stability,

and this requirement can essentially influence on the

structure design. Sometimes interesting details about

Fig. 3. Predicted rate of heat flux density for space electronic

components [16].

micro-satellite thermal arrangement can be found on

web sites of satellite developers [12–15].

3. Potential thermal control future tasks

The following tasks have been distinguished for

satellite thermal control: to cool the high power gen-

erating components. The future problem is associated

with growing heat dissipation of modern board com-

puter processors reaching 15–30W with density up to

100 kW/m2 [16], as seen in Fig. 3; to provide thermal

and geometrically stable mounting places for devices

(illustration in Fig. 4 [17]); to ensure the conditions for

operation of payloads with several temperature levels.

The other important direction is miniaturization of 

space equipment that leads to reduction of sizes of all

the components and increase in the heat flux densities

of thermal control components.

4. Software for thermal design

The wide set of the following products is available

in the market:

1. Lumped parameter methods: ESATAN (Ther-

mXL), SINDA/G/FLUINT, TRASSA,

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V. Baturkin / Acta Astronautica 56 (2005) 161–170 165

Fig. 4. Thermo- and geometry stable payload baseplate of mi-

cro-satellite BIRD [17]: provides the geometrical stability of 

mounting places and supports, parallel—precise of optical axes

(± arcmin); provides the comfortable temperature regime for op-

tics (+15–+20 ◦C); provides heat removal. Courtesy of DLR.

2. finite element and finite difference methods: NAS-

TRAN, COSMOS, ANSYS, FLOTHERM, TAS,

TAK2000,

3. radiation heat exchange (internal and external):

ESARAD, TERMICA, TRASYS, RadCAD,

SSPTA, OAZIS,

4. specific software is very useful for the thermal con-

tact conductance definition, thermoelectric cooler

design, and heat pipe design.

The following peculiarities and problems can be

discovered during the thermal model preparation:

1. Retrieval information and definition of exact ther-

mal properties: thermal conductivity, heat capac-

ity, density, uniformity of properties, and temper-

ature dependence of properties,

2. retrieval of information and definition of exact op-

tical properties and their variation during exploita-

tion• emittance, diffuse reflectivity, specular reflectiv-

ity, and transmissivity in IR band,

• absorptance, diffuse reflectivity, specular reflec-

tivity, and transmissivity in solar band,

3. definition of realistic values of thermal contacts in

 joints,

4. reasonable simplification of thermal structure of 

simulated object, preparation of the thermal loads

functions, and choice of heat transfer process func-

tionalities,

5. the trace ability of mechanical and structural de-

sign changes in the course of project and adequate

modification of thermal model.

5. Thermal hardware

Main thermal control components did not vary

abruptly. For micro-satellites they are mainly passive:

multi-layer insulation, optical coating and finishing,

thermal conductive lines, thermal isolators, heat stor-

age, heat pipes, and electrical heaters. Conditional

classification of hardware is presented in Table 3.

Typical multi-layer insulation consists of 20–30 in-

ner layers of less than 0.006 mm aluminized Mylar and

innermost and outermost layers 0.025 or 0.05 mm alu-

minized Kapton [7]. The effective emittance depends

on discontinuity of MLI blankets, averaged tempera-

ture, pressure, and pressing of blankets. The summary

on effective emittance is presented in Fig. 5 [7]. Rec-

ommended in [7] for draft estimation the value of ef-

fective emittance is 0.03 with tolerance ±0.02. This

value should be precised in the subsequent experi-

ments.

Optical coating and finishing, optical selecting coat-

ing can provide the value of ration emittance/solarabsorptance 0.1< /s < 10. Typical optical painting

for low-temperature radiator has s ∼ 0.2, ∼ 0.85.

Fig. 6 shows the optical properties for some finishing

and coatings [7].

Heat transfer inside the satellite is realized by means

of conduction through structural elements and special

conductive lines. Aluminum alloys are used as bus

and conductive lines and heat capacity storage ( =

2700 kg/m3, =150W/mK, cp=900J/kg K), beryl-

lium as heat transfer and heat storage (=1850 kg/m3,

= 180W/mK, cp = 1850 J/kg K). Conductive linescan be fabricated from flexible copper strings [18,19]

as geometrically adjustable thermal conductors (ther-

mal resistance less than 2 K/W). They allow to con-

  jugate the elements with unknown exact layout and

permit to reduce the torque on clamping points.

The other very effective heat conductance lines

are the heat pipes. For space application they are

mostly presented by ammonium axially grooved heat

pipes produced by extrusion, though other types of 

structures (as metal felts, screens) are used as well.

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166 V. Baturkin / Acta Astronautica 56 (2005) 161–170

Table 3

Classification of thermal hardware for micro-satellites

Name of hardware Frequency of application

Passive means to regulate the satellite structure temperature level

Radiating surfaces and finishing Often

Multilayer insulation Often

Heat transport means to manipulate with satellite integral heat capacity Not often

Semi-passive and active means to regulate the satellite structure temperature level

Heaters; Louvers Not often

Heat storages; Variable conductance heat pipes Very random

Changes of satellite orientation; Change of on-off of device Possible

  Hardware for individual devices installed on micro-satellite platform

Passive means to regulate the device/equipment temperature level

Radiating surfaces and finishing Very often

Multilayer insulation Often

Heat conductive lines (flexible lines, heat pipes) Often

Low conductive supports and standsoff Very oftenContact conductance means Very often

Semi-passive and active means to regulate the device temperature level

Heaters; Heat storages Often

Thermal switches, Thermoelectric coolers, Stirling coolers Random

Fig. 5. Effective emittance of MLI blankets as function of area

and discontinuity by Stimpson & Jaworski [11]. AMLI -sizes of 

MLI area for microsatellites, rec -recommended range of effective

emmitence.

Typical extruded profile has the diameter more than

8 mm, thermal resistance less than 0.1 K/W.

Selecting the type of capillary structure and liquid

heat transfer medium one can match these heat transfer

instruments over wide temperature range from −190

to +100 ◦C. Dimensional configurations of heat pipes

are presented in Fig. 7 [20].The illustrations of heat pipe application in micro-

satellite thermal design are small-satellites Magion 4,

Magion 5 [21] and BIRD [22]. The aim of heat pipes

in both cases is to transport heat between remote satel-

lite zones (front and back compartments for Magion

satellites, distance 0.5 m; payload and main radiator

for BIRD satellite, distance 0.3 m). Scheme of heat

pipe configuration for BIRD satellite thermal control

system is presented in Fig. 8.

For future micro-satellite missions one of the heat

pipe modifications will be useful, namely, micro-heatpipes by circumcircle diameters 1–6mm. They are

constant conductance heat pipes, have wire wick or

grooves as liquid transfer medium; material of shell-

copper and silver; typical circumcircle diameter from

1–6 mm; length up to 100 mm, heat carrier: alco-

hol, water; maximum transfer heat flux 2–10 W/cm2;

thermal resistance of HPs: less than 0.5–10 K/W.

Micro-heat pipe array dimensions are: 20–40 mm

width, 110 mm length with thickness of 1–3 mm. The

other type of new heat pipe technology, so-called

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V. Baturkin / Acta Astronautica 56 (2005) 161–170 167

Fig. 6. Values of emittance and solar absorptance [7]. Zone

1—selective blacks (solar absorbers); 2—sandblasted metals and

conversion coatings; 3—white paints and second surface mirrors;

4—bulk metals (unpolished); 5—dielectric films on polished met-

als. Ideal reflector = 0; s = 0—in left bottom corner; ideal

painting = 1; s = 0—in right bottom corner; ideal solar ab-

sorber = 0; s = 1—in left upper corner; ideal black painting

= 1; s = 1—is in right upper corner.

Fig. 7. Examples of heat pipe configurations with metal belt wick.

Fig. 8. Scheme of heat pipe arrangement in BIRD satellite. Heat

moves from evaporator (payload)—condensator (radiator) in mea-surement.

loop-heat pipes (LHPs), allows to operate indepen-

dently of gravity forces. Distance between heat input

and output zones can reach several meters at transfer

power up to 100 W. Also, the connecting lines in LHP

are flexible (thin wall tubes with diameter of 2–3 mm)

that allows to apply this device for thermal coupling of 

moveable parts. Alternative solution for heat transport

at a distance less than 0.3m and for heat spreading

is based on the use of variety of graphite materials,with effective conductivity more than copper [23].

Thermal contacts between satellite structural ele-

ments, thermal contact of high power generating com-

ponents with cooling means or substrates could be im-

proved by application of different types of interface

pads, greases, gap fillers and encapsulates. Typically,

these means are placed between contacting surfaces

that is presented in Fig. 9 [24].

Typical conductivity of pads is 1–6W/mK, they can-

not be applied with high pressure (typically 10–30 psi

or 0.7–2 bar), and allow to reduce the contact resis-tance in coupling joint [25]. Other variant to cool the

power-generated micro-elements is to use the con-

formable gap fillers, which allow to connect the heat

generated component with thermal shield and rear-

range the heat flux. Scheme of conformable pad ap-

plication is shown in Fig. 10.

Potential problems in space application of inter-

face pads, gap fillers, greases, encapsulates and ad-

hesives are associated with risk of mass losses and

subsequent spacecraft and payload contamination. The

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168 V. Baturkin / Acta Astronautica 56 (2005) 161–170

Fig. 9. Scheme of interface pad application to improve the contact

conductance [25].

Fig. 10. Cooling of electronic components of different

sizes: 1—printed card; 2—electromagnetic and thermal shield;

3—electronic components; 4—conformble pad; 5—ways of heat

movement to heat sink.

other important task is to verify the stability of the

component characteristics under space environment

factors and preliminary estimation of technical effi-

ciency of selected application.Important passive thermal control means are the low

conductive supports and standoffs. They are exploited

for thermal isolation of the device (part of the device)

from the satellite. Mainly such a measure is necessary

for devices with autonomous thermal control either for

devices or their parts, which have other temperature

levels, different from interior of the satellite. Typical

materials for these supports are a wide variety of low-

conductivity materials, including fiberglass, stainless

steel, titanium or plastics. The choice of material is dic-

Fig. 11. Scheme of low-conductance standoff  [19]. Insulating

material—glass fibers with epoxy, attachment elements made of 

stainless steel.

tated by the conductivity, temperature range, thermal

expansion and mechanical properties. Typical variant

of support design, used many times in space for the

temperature range of 170–350 K for thermal isolation

of 2 kg device is shown in Fig. 11. Four such supports

provide the thermal resistance higher than 400 K/W

and thermally disconnect the device from satellite bus

[19].

Such effective thermal control means as the louvers,

radiators with changeable optical properties, unfolded

radiators, radiators-variable conductance heat pipes,

heat storage, thermal switches, controlled electrical

heaters have not been considered in this paper.

The louvers and radiators with variable emitting

ability are more typical for larger satellite with massmore than 100kg. Heat storage, thermal switches, ther-

moelectric cooling and controlled electrical heaters are

the attributes of autonomous thermal control techno-

logy. More details concerning the above-mentioned

instruments can be found in special literature [7].

Among advanced thermal technologies one can note

the following means, which are very soon going to be

used in practice [7]:

1. Variable emittance coating technology (variation of 

emittance 0.2–0.8),2. micro-louvers (change of effective emittance in 2

times),

3. mini two-phase heat transfer devices

• mini heat pipes: diameters 1–2 mm, length

60–100 mm, heat power transferred 1–10 W,

• mini capillary pump loops: diameter 1–2mm,

length 300–1000mm, heat power transferred

3–100W,

4. mechanical thermal mini-switches (thermal resis-

tance change 100:1),

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V. Baturkin / Acta Astronautica 56 (2005) 161–170 169

5. high performance C–C composites with conductiv-

ity 400–1100W/mK, which can be used either for

the structure elements or heat transport elements.

6. Ground testing and verification of thermal

control design

Verification of the thermal design should be con-

ducted prior to the launch. For micro-satellites the vac-

uum chamber with inner sizes of about 2 m length and

1.5 m in diameter is often enough for arbitrary satellite

layout inside the chamber. The vacuum chamber with

pressure of 0.0013 Pa (10−5 Torr) or less, cooled by

liquid nitrogen screen simulates the thermal conditionof the outer space (Figs. 12 and 13).

The Earth, Sun heat fluxes are simulated by aiding

heaters, installed on the back side of radiating sur-

faces, on the base of the absorbed fluxes knowledge.

Sometimes the Earth radiation is simulated by radia-

tive heat sinks with regulated temperature. The ther-

mal balance test of the satellite system, for cold and

hot cases of exploitation, should confirm or precise

the accepted parameters of most of the thermal control

components and approve used thermal concept.

Exploitation temperature boundary for satellitecomponents should be enlarged by ±11 ◦C [7] to

compensate the uncertainty connecting with calcula-

tions, design and environmental conditions. Among

possible recommendations, based on satellite design

and flight performance experience, the following ones

could be useful:

• Keep in mind the possibility of unpredictable sit-

uation with satellite orientation with respect to

sun light and with variation of power generation

value inside the satellite. They can shift satellitetemperature level essentially.

• Maintain the battery in comfortable temperature

range during charge–recharge mode during the

satellite’s shadow period.

• Degradation of optical surfaces is the real pro-

cess, which is important especially for the front

direction toward the Sun surfaces.

• Not all the factors can be simulated and estimated

in the ground tests and by calculations. Design

your equipment for wider temperature range.

Fig. 12. Space simulation chamber of DLR’s Space Center in

Berlin with volume of 3.2 m3 [26].

Fig. 13. Installation of micro-satellite BIRD into vacuum chamber

for thermo-vacuum testing. Courtesy of DLR [27,28].

7. Conclusion

Summary of the information concerning present-

day thermal control concepts conformably to micro-

satellite group (mass less than 100 kg) is surveyed andanalyzed. The description of the conventional thermal

hardware used in micro-satellite design is briefly pre-

sented. The base for the survey was the lessons learned

by the National Technical University of Ukraine “Kyiv

Polytechnic Institute” during the elaboration of ther-

mal control hardware for micro-satellites Magion 4,

5 (launched in 1995 and 1996), BIRD (launched in

2001), autonomous thermal control systems for inter-

planetary missions VEGA (launched in 1984), PHO-

BOS (launched in 1986) and other public data.

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170 V. Baturkin / Acta Astronautica 56 (2005) 161–170

Acknowledgements

The author thanks the BIRD team of the Institute

of Space Sensor Technology and Planet Exploration(DLR, Berlin, Germany) for the permission to use

illustrations and technical data.

Exchange of information in related fields can be

realized via email: [email protected] and fax:

(38044) 241-75-97.

References

[1] Small satellites for Earth observation, Digest of the First

International Symposium of the IAA, Berlin, November 4–8,

Copyright 1996 by Walter de Gruyter & Co., 1996.[2] H.P. Roser, R. Sandau, A. Valenzuela (Eds.), Small satellites

for Earth observation, Digest of the Second International

Symposium of the International Academy of Astronautics

(IAA), Berlin, April 12–16, Wissenschaft und Technic Verl.,

Berlin, ISBN 3-89685-561-1, 1999.

[3] Website of Surrey Space Center, Small Satellite Home Page,

http://www.ee.surrey.ac.uk/SSC/SSHP .

[4] O. Favorsky, Ya. Kadaner, Questions of Heat Transfer in

Space, Vyschya shkola, Moscow, 1972.

[5] V. Andreanov, V. Artamonov, et al., Autonomous Planet

Stations, Nauka, Moscow, 1973.

[6] V. Zaletaev, Yu. Karpinos, O. Surguchev, Spacecraft Heat

Transfer Calculation, Mashinostroenie, Moscow, 1979.

[7] D. Gilmore (Ed.), Satellite Thermal Control Handbook, TheAerospace Corporation Press, El Segundo, CA, 2002.

[8] P. Fortescue, J. Stark (Eds.), Spacecraft System Engineering,

third ed., ISBN 0 471 95220 6, Copyright John Wiley,

1995.

[9] Vincent L. Pisacane, Robert C. More (Eds.), Fundamental of 

Space Systems, ISBN 0-10-507497-1, Copyright Oxford

University Press Inc., Oxford, 1994.

[10] P. Zemlianoy, C. Combes, Thermal control of space

electronics, http://www.electronic-cooling.com/Resources/ 

EC-articles/SEPT96/Sep96_03.htm, 2000.

[11] C. Jilla, D. Miller, Satellite design: past, present and

future, International Journal of Small Satellite Engineering,

http://www.ee.surrey.ac.uk/EE/CSER/OUSAT/IJSSE/IJSSC/ 

issue1/cjilla/cjilla.html .

[12] Website of Institute of Atmospheric Physics of Czech

Academy of Science, http://www.ufa.cz, 2000.

[13] Website of Institute for Space Research Russian Academy of 

Science, http://www.iki.rssi.ru , 2000.

[14] Website of OHB System, http://www.fuchs-gruppe.com/ 

ohb-system, 2000.

[15] Website of Institute of Space Sensor Technology and Planet

Exploration, http://www.ba.dlr.de/ne/ws .

[16] Kaveh Azar, History of power dissipation, http://electronics-

cooling.com/html/2000_jan_a2.html.

[17] I. Walter, B. Heym, W. Stadthaus, Microsatellite structure

solutions: highly integrated, modular design of BIRD,

Proceedings of the Fourth International Symposium, SmallSatellites Systems and Services, 14–18 September, Antibes-

Juan les PINS/France, CNES, 1998.

[18] G. Avanesov, Ya. Ziman, V. Tarnopolsky, et al., TV Survey of 

Comet Halley, Nauka, Moscow, p. 295, ISBN 5-02-000093-0,

1989.

[19] V. Kostenko, V. Baturkin, N. Grechina, et al., Cryogenic

System for the Videospectrometric Complex (VSC) Cooling

in Project “PHOBOS”, preprint, Space Research Institute,

Academy of Sciences of the USSR, Pr-1409, 1988, p. 39.

[20] V. Baturkin, S. Zhuk et al., Autonomous heat pipe systems

for electronic components thermostating at near-earth orbit

exploitation, Proceedings of the 24th International Conference

on Environmental Systems, Friedrichshafen, Germany, June20–23, Rep. No. 941302, 1994.

[21] J. Vojta, S. Zhuk, V. Baturkin, Thermocontrol system concept

of Magion small subsatellites of Intreball Mission, Digest

of the First International Symposium of the International

Academy of Astronautics (IAA), Berlin, November 4–8, 1996,

pp. 380–383.

[22] F. Lura, B. Biering, V. Baturkin, et al., Heat pipe application

for thermal stable bench arrangement in small satellite

design, Proceedings of the 30th International Conference

on Environmental Systems (ICES) and Seventh European

Symposium on Space Environmental Control Systems,

Toulouse, France, July 10–13, Rep. No. 2000-01-2460, 2000.

[23] Material TC1050, Website http://www.Advceramic.com .

[24] Journal Electronics Cooling 5 (1) (1999) 38.[25] C. Latham, Technical brief, thermal resistance of interface

materials as a function of pressure, Journal Electronics

Cooling 2 (3) (1996) 35.

[26] F. Lura, D. Hagelschuer, System conditioning—our ways

and testing tools for the development of reliability for

spaceborne components and small satellites, Digest of the

First International Symposium of the International Academy

of Astronautics (IAA), Berlin, November 4–8, pp. 365–368,

1996.

[27] B. Biering, F. Lura, H.G. Lotzke, I. Walter, et al.,

TCS design of the microsatellite BIRD for infrared Earth

observation, Proceedings of the 28th International Conference

on Environmental Systems, Dancers, Massachusetts, USA,

July 13–16, Rep. No. 981639, 1998.

[28] K. Brieß, W. Bärwald, F. Lura, et al., The BIRD mission

is completed for launch with the PSLV-C3 in 2001, in: H.

Röser, R. Sandau, A. Valenzuela (Eds.), Digest of the Third

International Symposium of IAA, Small Satellites for Earth

Observation, Berlin, April 2–6, Wissenschaft und Technic

Verl., Berlin, 2001, ISBN 3-89685-566-2, 2001, pp. 323–326.