M.Sc. in SPACE STUDIES 2004/2005

A Microsystems Technology Implementation of an Optical Solar Reflector for Miniaturized Silicon Space Systems

Lise Bilhaut Individual Project Report submitted to the International Space University in partial fulfillment of the requirements of the M.Sc. Degree in Space Studies August, 2005

Internship Mentor:

Peter Nilsson Zandkarimi

Host Institution:

Ångström Space Technology Centre, Uppsala, Sweden

ISU Academic Advisor:

Isabelle Scholl

Lise Bilhaut

August, 2005

Acknowledgments I would like to thank my advisor Peter Nilsson Zandkarimi for his patience, guidance and confidence in my work, Henrik Kratz for his availability throughout my work, Arne Roos for his help with the spectrophotometer use, Johan Bejhed for his aid in drawing the photolithographic mask, my office mate Tobias Böhnke for the useful conversation and his kindness, Hugo Nguyen for his documents about the thermal simulation of NanoSpace-1 and Fredrik Bruhn, without whom I would not have had the chance to discover the fabulous work conducted in the Ǻngström Space Technology Center. Thanks also to Isabelle Scholl and Johan Köhler for their feedback on this report. I also would like to take this opportunity to thank the people who supported me during this internship, my parents and my grand-mother, and my friends Navtej, Erik, Florian, Emilie and Pierre.

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Acronyms °C

α ǺAC Al AM0

αS ǺSTC Cu DMC

ε εIR

Degree Celsius Absorptivity Ǻngström Aerospace Corporation Aluminum Air Mass Zero Solar absorptance Ǻngström Space Technology Center Copper Disaster Monitoring Constellation Emissivity Infrared emittance

GPS IR KOH LEO MEMS MFS MMNS MMS MTQ NASA PECVD

Global Positioning System Infrared Potassium Hydroxide Low Earth Orbit Micro-Electro-Mechanical Systems Multifunctional Structures Multifunctional Micro/Nano Systems Multifunctional Microsystems magnetic torquer National Aeronautics and Space Administration Plasma Enhanced Chemical Vapor Deposition

SEM

Scanning Electron Microscope

Si3N4

Silicon Nitride

SiO2

Silicon Dioxide

UV W

Ultra Violet Watt

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Abstract This report presents a work that has been carried out during an internship to complete the Master of Science degree at the International Space University. This internship was done at the Ǻngström Space Technology Center, in the University of Uppsala, Sweden. The ÅSTC focuses on the development of microtechnology for space applications, linking two technologies that serve the current trend of reducing the size of satellites. Indeed, ÅSTC is developing a multifunctional microsystems concept, liking both microtechnology and multifunctional structures into one concept. The Center is developing a nanosatellite, NanoSpace-1, built with such multifunctional systems and whose weigh shall not exceed 10 kg. This work was focusing on the thermal management of the satellite. While in space, a spacecraft should avoid to be overheated by the Sun or the Earth environment (albedo and Earth infrared) while emitting well in the infrared in order to red ride of excess heat generated internally by the electronics. This issue is prominent for miniaturized satellites such as NanoSpace-1, where the power density is higher than for regular satellites. This report considers some basic calculations that have been done to assess the thermal environment of NanoSpace-1, in the worse cases scenarios (hot and cold case). Several functions such as transmission and reception are accomplished by modules included in the satellite, called Multifunctional Micro/Nano Systems. These modules are square multilayer stack of wafers of 6.8 cm by side that also serve as structural elements. To their structural functions and their other purposes, these modules can also serve as a radiator to complete the active thermal control of the satellite. Indeed, a module facing the outer space can also act as a regular radiator provided than its surface has been tailored to answer to the specific constraint such a radiator demands. A low solar absorptance and a high infrared emittance are the key parameters to answer to the requirement of a thermal control coating, as well as a big emissive area. In a MMNS module with fixed dimensions, the area can be increased in a controlled way by creating some relief onto the surface with lithography. This work has investigated several solutions that could be applied to a module, combining relief manufacturing, metallic thin film deposition and dielectric deposition. It has been shown that a 25 µm grooves relief combined with a 100 nm layer of aluminum and a 3.5 µm layer of silicon dioxide is the best way to increase the infrared emittance while keeping the solar absorptance quite low and increasing the area of the module. Thus, it is recommended that the MMNS module uses such a solution as a thermal control coating, provided that this coating does not prevent the other functions of the module to be fulfilled. It is hence recommended to test this selected coating with for instance the receiver function of a S-band antenna.

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Content Acknowledgments................................................................................................................................. i Acronyms .............................................................................................................................................. ii Abstract.................................................................................................................................................iii Content .................................................................................................................................................iv Introduction ..........................................................................................................................................1 1.

Background ..................................................................................................................................2 1.1. Small Satellites ....................................................................................................................2 1.2. New Technologies for Small Satellites............................................................................4 1.3. ǺSTC and Nanospace-1 ...................................................................................................5 1.4. The MMNS Module ..........................................................................................................9

2.

Coatings for Satellite Thermal Control..................................................................................11 2.1. Basic Optics ......................................................................................................................11 2.2. Satellite Thermal Balance................................................................................................14

3.

Experiments...............................................................................................................................20 3.1. Process...............................................................................................................................20 3.1.1. Relief.........................................................................................................................20 3.1.2. Coatings....................................................................................................................27 3.1.3. Measurement ...........................................................................................................28 3.2. Results................................................................................................................................32 3.2.1. Type of Wafers........................................................................................................32 3.2.2. Relief.........................................................................................................................33 3.2.3. Metallic Thin Films.................................................................................................34 3.2.4. Dielectric Coatings..................................................................................................34 3.2.5. UV Treatment .........................................................................................................35 3.2.6. Thickness Influence................................................................................................36 3.3. Analysis..............................................................................................................................38 3.4. Comparison.......................................................................................................................41

Conclusion ..........................................................................................................................................42 References .............................................................................................................................................v Appendix – Results.............................................................................................................................vi

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Introduction The International Space University is an institution dedicated to the education of present and future space professionals. In the master program offered by the university, each student has to carry out a twelve-week internship in a space-related organization in order to test his recently acquired knowledge on the rough area of reality. Having a master in solid-state physics and some experience in microstructure processing, I was hoping to be able to link my background and my new understanding of the space sector. The Ǻngström Space Technology Center (ǺSTC), a department of the Uppsala University (Sweden), fully answered to this wish. Indeed, the ǺSTC is working on the application offered by microand nano-technologies to the space sector. This technology is particularly adapted to the new trend in spacecraft design, which is to quickly build small satellites, although it will certainly show applications for bigger spacecrafts. The subject of this internship was related to the Multifunctional Micro/Nano Systems (MMNS), an ongoing project sponsored by the European Space Agency. The goal of the project is to build a module that would integrate both functions of transceiver/receiver and thermal control as well as being able to serve as a structural element on a satellite. In space, where convection and conductive heat flow are absent in a microgravity and vacuum environment, the only way to transfer heat is to radiate it. It is thus indispensable for the MMNS module, with high power densities, to develop an efficient and effective thermal control management. A part of the solution is to tailor the surface of the module so it radiates effectively the internally generated power into space, while preventing the module to become overheated by its environment. This internship aimed to try different coatings and to assess their influences on the thermal balance of the satellite. These coatings had to meet different requirements such as a low solar absorptance and a high infrared emittance, as well as being available in the Ångström Microstructure Laboratory and showing a good adhesion to the silicon substrate. A microstructure processing of the surface was also tested to increase the area of the space-faced module, hence increasing its radiative capability. The first chapter of this report presents the renewed development of small spacecrafts and the related technologies that are answering and promoting it. It goes deeper into the philosophy of the ǺSTC, which is to link multifunctional systems and micro-technology, and then presents in detail the MMNS module. The second chapter deals with the theoretical background necessary to understand the conducted work. It presents basic optical laws, as well as thermal consideration of a satellite that orbits the Earth. The last chapter presents the effective work conducted during the internship. It explains the experimental steps that were followed, and give the related results. These are analyzed so a useful conclusion can be drawn.

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1. Background 1.1.

Small Satellites

The orbital mass of Sputnik 1, the first artificial satellite successfully placed into orbit around the Earth, was 83.6 kg. Since, mass of satellite have increased a lot, and it is not uncommon to see a telecommunication satellite weighing a few tons. The mass of a satellite is actually only limited by the mass that a launcher can place into orbit and the volume that can fit inside the fairing. Today’s launchers can lift several tons into orbit. For instance, Space Shuttle can place in Low Earth Orbit (LEO) up to 28.8 tons and Ariane 5G can put 6.8 tons in Geo-Transfer Orbit [Futron, 2002]. For heavier constructions such as a space station, one has to assemble them directly into orbit. It is generally accepted that the cost of the launch depends on the mass it can put into orbit. For instance, a Space Shuttle launch costs 300 million and an Ariane 5G launch 165 million. To place small satellites in LEO, smaller launcher such as Cosmos or Delta 2 are also available. In that case, the price of the launch drops below 100 million (55 million for Delta 2 and 13 million for Cosmos) [Futron, 2002]. The first satellites were small for the only reason that launchers could then not place huge payloads into orbit. Some of them, such as the VELA satellites (the first ones were launched in 1963) had very successful missions – in the VELA case, Earth observation; and they only weighed 152 kg. Other examples of these early small satellites include OSCAR I, a 5 kg spacecraft built for radio-amateurs, which was launched in 1961, and the first commercial communication satellite, Early Bird, launched in 1965, that weighed 39 kg [Wertz & Larson, 1999]. From 1965 to 1985, when the focus of the biggest space nations was on human spaceflight, small satellites, although always used and developed (mainly by the military sector), were not the main concern of space agencies or the commercial sector. However, in the mid-seventies, NASA began to promote small satellites with its program Get Away Special that provides diverse users with a low cost access to space, provided the satellite does not weigh above 68 kg and can be attached to the Space Shuttle. Europe has also promoted the launch of small satellites, first with Ariane 4 and now with Ariane 5 thanks to the program Ariane Structure for Auxiliary Payloads. Since 1985, together with the opportunities of low-cost access to space and improved digital capabilities, the use of small satellites is becoming wider and wider. But as we will see, the technology is still under development that should allow for building a 10kg satellite with the same abilities as a 500-kg satellite. Figure 1 shows a trend in the dry mass of NASA science satellites: after the dry masses have increased over more than 30 years, it has begun to decrease since the 90s. This reverse corresponds to the failure of several big and expensive spacecrafts, such as the loss of Mars Observer, and the new motto of the NASA, “faster, better, cheaper”, brought by its

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administrator Dan Goldin to answer the public demand for successful missions and the shrinkage of NASA budget.

Mass of the spacecraf (kg)

100000

10000

1000

100

10

1 1955

1960

1965

1970

1975

1980

1985

1990

1995

2000

2005

Launch year

Figure 1 – Dry mass trends of NASA science satellites [Jilla & Miller, 1997]

Indeed, faced by the reduction of financial resources, a trend has appeared to lower the cost of space missions. First of all, building smaller satellites allows using small, hence cheaper, launch vehicles. They can share the launch of bigger ones when their mission orbits are similar by being launched for instance in piggy-back. But small satellites have also opened the space arena to other organizations (universities, companies or developing countries) that could not otherwise afford to launch a satellite. The Disaster Monitoring Constellation, a project lead by Surrey Satellite Technology Ltd., is a typical example of the usefulness of small satellites: bringing together seven countries, the DMC is composed by satellites weighting between 90 and 130 kg with very good Earth observation capabilities. Building a small satellite requires less people, thus increase the level of personal involvement and the efficiency and quality of the team’s work. Another advantage is that they can be built quickly, a couple of years instead of ten years for conventional satellites. Being built quickly by a dynamic organization, newer generation of components can be used: the traditional space sector prefers to stick with old but reliable technology, but with small satellites, the low cost of the mission allows taking more risks and using state-of-the-art technologies. Furthermore, when building similar small satellites that can be used in a constellation, batch processing will help to put the cost down by dividing the development cost. Building cheap satellites also permits new mission concepts (formation flying and so on). But small satellites demand a trade-off in the system requirement. Indeed, the requirement will also change due to the miniaturization: the power requirement, the amount of propellant

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and the moment of inertia will decrease but on another hand, the power density and heating will increase while decreasing the thermal mass, and the aperture size could not be as big as on regular satellites, which will have a direct influence on the payload capabilities. The same goes for the scaling laws: some of them, such as laminar and molecular flow, are positively improved when decreasing the size. At sufficient small scale, surface forces dominate over bulk forces and electrostatic forces over gravitational forces, which can be used to create new systems. On the other hand, decreasing the size of the satellite decreases the size of the payload. It is difficult to use a small satellite for telecommunication since it needs large antennas, large solar areas to supply the required energy and big tanks to assure a sufficient life-time into orbit. Small satellites can only be used for short space mission, since propellant is needed to keep the satellite into its operating orbit. The payload must also be adapted and the performances are limited. Moreover, when large satellites are over-redundant, subsystems in small satellites are at a first glance more vulnerable, multiplying the risk for single-point failure. But a redundancy of spacecraft, either in flight or on ground, can compensate for this apparent drawback. Nonetheless, small satellites are believed to be on the long run cheaper than bigger ones, according to the “faster, better, cheaper” motto of NASA. Although there is a trade-off between the three comparatives, small satellite are definitively faster and cheaper and current and future technology development may allow, in some cases, to make them if not better, at least as good as bigger ones. The term small satellite includes all the spacecrafts with a mass less than 500 kg. But there are more specific denominations for small satellites regarding their masses, presented in Table 1. Table 1 – Satellite classification regarding their masses Denomination Large satellite Medium satellite Mini satellite Micro satellite Nano satellite Pico satellite

1.2.

Mass (kg) > 1,000 500 – 1,000 100 – 500 10 – 100 1 – 10 10,000 Ω). For a plain wafer, the optical properties depend on these characteristics, but as they are quite expensive, it was interesting to do the experiment on cheaper wafer, providing that the results would be the same. Table 2 shows if there is a difference between both kinds of wafer, it is in the range of the experimental mistake (for 25 µm grooves, αS and εIR are smaller for high ohmic wafer whereas for 25 µm pyramids, they are higher). Table 2 – αS and εIR for simple wafers and high ohmic wafers with a 100 nm Al thin film and 2 µm SiO2 layer Relief

25 µm grooves

5 µm grooves

25 µm pyramids

5 µm pyramids

Metallic thin film

Dielectric coating

αS

εIR

Simple SiO2

0.262

0.933

High Ohmic SiO2

0.242

0.894

Simple SiO2

0.237

0.701

High Ohmic SiO2

0.306

0.671

Simple SiO2

0.329

0.965

High Ohmic SiO2

0.353

0.970

Simple SiO2

0.377

0.780

High Ohmic SiO2

0.385

0.769

Simple

0.607

0.669

High Ohmic

0.659

0.632

Al

Al

Al

Al

5 µm pyramids Nothing

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3.2.2. Relief Table 3 shows a comparison between grooves and no relief: in all the cases, αS and εIR are much lower when there is no relief than when there is some. Table 3 – αS and εIR for flat and rough wafers with a 100 nm Al thin film and 2 µm SiO2 layer Relief

Metallic thin film

Dielectric coating

Al

αS

εIR

SiO2

0.21

0.927

Nothing

Nothing

0.68

0.772

Al

SiO2

0.22

0.706

Nothing

Nothing

0.69

0.443

SiO2

0.11

0.389

Nothing

0.47

0.035

25 µm grooves

5 µm grooves

Nothing

Al

Table 4 shows αS and εIR for different relief, a 100 nm aluminum layer and a 2 µm SiO2 layer. 25 µm pyramids are better for our purpose regarding to the 5 µm pyramids, because αS is lower and εIR is higher and the same goes for the grooves. Both αS and εIR are higher with pyramids than with grooves, thus more analysis will be needed to determine which one is the best (see section 3.3). Table 4 – αS and εIR for different relief with a 100 nm Al thin film and 2 µm SiO2 layer Relief

Metallic thin film

Dielectric coatings

αS

εIR

25 µm grooves

Al

SiO2

0.242

0.894

5 µm grooves

Al

SiO2

0.306

0.671

25 µm pyramids

Al

SiO2

0.353

0.970

5 µm pyramids

Al

SiO2

0.385

0.769

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3.2.3. Metallic Thin Films Table 5 shows that αS increases a lot with Cu compared to Al, which is not good for our purpose, when εIR remains roughly the same. The decrease in the solar absorptance might be due to the oxidation of the metallic layer that decreases the reflectance. Table 5 – αS and εIR for 100 nm Cu thin film vs. 100 nm Al thin film as a metallic layer (with a 2 µm SiO2 layer) Relief

Metallic thin film

Dielectric coatings

αS

εIR

25 µm grooves

Al

SiO2

0.242

0.894

25 µm grooves

Cu

SiO2

0.791

0.874

5 µm grooves

Al

SiO2

0.306

0.671

5 µm grooves

Cu

SiO2

0.902

0.839

25 µm pyramids

Al

SiO2

0.353

0.970

25 µm pyramids

Cu

SiO2

0.816

0.981

5 µm pyramids

Al

SiO2

0.385

0.769

5 µm pyramids

Cu

SiO2

0.852

0.857

3.2.4. Dielectric Coatings Table 6 shows that both αS and εIR are higher with Si3N4 than with SiO2; more analysis will be needed to decide which one fits better to our purpose (see section 3.3).

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Table 6 – αS and εIR for 2 µm Si3N4 layer vs. 2 µm SiO2 layer (with a 100 nm Al thin film) Metallic thin film

Relief

25 µm grooves

5 µm grooves

Dielectric coatings

αS

εIR

High Ohmic SiO2

0.242

0.894

Si3N4

0.365

0.939

High Ohmic SiO2

0.306

0.671

Si3N4

0.457

0.779

High Ohmic SiO2

0.353

0.970

Si3N4

0.447

0.979

High Ohmic SiO2

0.385

0.769

Si3N4

0.528

0.832

Al

Al

25 µm pyramids

5 µm pyramids

Al

Al

3.2.5. UV Treatment [Hass and al, 1973] and [Bradford and al, 1970] state that an ultraviolet exposure of the sample lowers αS while not having a noticeable influence on εIR. The experience has been made to expose samples to a 435 W UV light for 20 hours, at a distance about 5 cm. Results are presented in Table 7 but a close examination does not allow for any conclusion. Table 7 – αS and εIR for UV treated samples vs. non-UV treated samples Relief

Metallic thin film

Dielectric coatings

αS

εIR

Normal

UV treatment

Normal

UV treatment

Al

SiO2

0.325

0.32

0.854

0.845

Cu

SiO2

0.822

0.806

0.871

0.878

Al

SiO2

0.281

0.279

0.891

0.907

Cu

SiO2

0.642

0.679

0.919

0.892

5 µm pyramids

25 µm pyramids

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3.2.6. Thickness Influence The thickness of the SiO2 layers has a linear influence on the solar absorptance (see Table 8 and Figure 30) whereas the IR emittance will slightly increase up to a thickness between 3.5 and 5 µm and then remains quite stable. Table 8 - αS and εIR for different thickness of SiO2 (with a 100 nm Al thin film) Relief

25 µm grooves

25 µm pyramids

Metallic thin film

Dielectric coating

αS

εIR

SiO2 (1 µm)

0.235

0.844

SiO2 (2 µm)

0.242

0.890

SiO2 (3.5 µm)

0.262

0.934

SiO2 (5 µm)

0.281

0.943

SiO2 (7.5 µm)

0.313

0.950

SiO2 (10 µm)

0.353

0.927

SiO2 (1 µm)

0.268

0.943

SiO2 (2 µm)

0.283

0.963

SiO2 (3.5 µm)

0.292

0.975

SiO2 (5 µm)

0.295

0.971

SiO2 (7.5 µm)

0.306

0.972

SiO2 (10 µm)

0.327

0.972

Al

Al

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0

0,2

0,4

0,6

0,8

1

1,2

0

2

4

SiO2 thickness (um)

6

8

25 um 25 um 25 um 25 um

10

12

grooves IR emittance pyramids IR emittance grooves solar absorptance pyramids solar absorptance

Lise Bilhaut August, 2005

Figure 30 – αS and εIR vs. SiO2 thickness

37

Absorptance / emittance

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3.3. Analysis Using the equation A.(α S .I solar + α S .I solar .ρ albedo .Falbedo + ε .I EIR .FEIR ) + Qint = ε .σ . A.T 4 , the hot case and cold case have been computed for each sample. Values of parameters used for the calculation are shown in Table 9. The hot case (respectively cold case) corresponds to a sun synchronous orbit at 500 km of altitude (respectively 1,000 km), with a local time at ascending node at 12:00 (respectively 12:00), during the winter solstice (respectively summer solstice) and with the solar panels pointing to the Sun [Nguyen, 2005, unpublished]. These calculations are only a tool to determine the best coating, and the results do not correspond to the temperature of the module while the satellite is into orbit, since the module is in thermal equilibrium with the surrounding aluminum frame, that also acts as a radiator. Table 9 – Cold case and hot case values [Wertz & Larson, 1999 and personal communication with Fredrik Bruhn and Hugo Nguyen in July 2005] Parameter

I solar (W.m-2 )

ρ albedo Falbedo

I EIR (W) FEIR

Cold case 1322 (summer solstice)

Hot case 1414 (winter solstice)

0.18 (min for orbit inclination of 0-30 deg en angle of Sun Out of orbit plane 0 deg) 0.0232 (min for altitude 500 km and theta 90 (etha 0)) 218

0.3 (max for orbit inclination of 30-90 deg en angle of Sun Out of orbit plane 0 deg) 0.8587 (max for altitude 500 km and theta 0 (etha 0)) 275

0.0006 (altitude of 1000 km, etha of 150)

0.8618 (altitude of 500 km, etha of 0) 10

0

Qint (W)

Results are shown in Table 10 to Table 13, where the best option has been highlighted. Table 10 – αS /εIR and hot and cold cases for flat and rough wafers Relief

Metallic thin film

Dielectric coating

αS/εεIR

Cold case (ºC)

Hot case (ºC)

Al

SiO2

0.227

-3

205

Nothing

Nothing

0.881

106

261

Al

SiO2

0.312

19

237

Nothing

Nothing

1.558

164

338

SiO2

0.283

12

304

Nothing

13.429

476

836

25 µm grooves

5 µm grooves

Nothing

Al

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Table 11 – αS /εIR and hot and cold cases for different relief with a 100 nm Al thin film and 2 µm SiO2 layer Relief

Metallic thin film

Dielectric coatings

αS/εεIR

Cold case (ºC)

Hot case (ºC)

25 µm grooves

Al

SiO2

0.271

9

212

5 µm grooves

Al

SiO2

0.456

48

251

25 µm pyramids

Al

SiO2

0.364

31

211

5 µm pyramids

Al

SiO2

0.501

56

240

Table 12 – αS /εIR and hot and cold cases for 100 nm Cu thin film vs. 100 nm Al thin film as a metallic layer (with a 2 µm SiO2 layer) Relief

Metallic thin film

Dielectric coatings

αS/εεIR

Cold case (ºC)

Hot case (ºC)

25 µm grooves

Al

SiO2

0.271

9

212

25 µm grooves

Cu

SiO2

0.905

108

252

5 µm grooves

Al

SiO2

0.456

48

251

5 µm grooves

Cu

SiO2

1.075

125

264

25 µm pyramids

Al

SiO2

0.364

31

211

25 µm pyramids

Cu

SiO2

0.832

100

240

5 µm pyramids

Al

SiO2

0.501

56

240

5 µm pyramids

Cu

SiO2

0.994

117

258

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Table 13 – αS /εIR and hot and cold cases for 2 µm Si3N4 layer vs. 2 µm SiO2 layer (with a 100 nm Al thin film) Relief

25 µm grooves

5 µm grooves

25 µm pyramids

5 µm pyramids

Metallic thin film

Dielectric coatings

αS/εεIR

Cold case (ºC)

Hot case (ºC)

High Ohmic SiO2

0.271

9

212

Si3N4

0.389

36

215

High Ohmic SiO2

0.456

48

251

Si3N4

0.587

69

244

High Ohmic SiO2

0.364

31

211

Si3N4

0.457

48

216

High Ohmic SiO2

0.501

56

240

Si3N4

0.635

76

241

Al

Al

Al

Al

Table 14 – αS /εIR and hot and cold cases for different thickness of SiO2 (with a 100 nm Al thin film) Relief

25 µm grooves

25 µm pyramids

Metallic thin film

Dielectric coating

αS/εεIR

Cold case (ºC)

Hot case (ºC)

SiO2 (1 µm)

0.278

11

218

SiO2 (2 µm)

0.272

9

212

SiO2 (3.5 µm)

0.281

12

208

SiO2 (5 µm)

0.298

16

209

SiO2 (7.5 µm)

0.329

23

210

SiO2 (10 µm)

0.381

34

216

SiO2 (1 µm)

0.284

12

208

SiO2 (2 µm)

0.294

15

207

SiO2 (3.5 µm)

0.299

16

206

Al

Al

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SiO2 (5 µm)

0.304

17

207

SiO2 (7.5 µm)

0.315

20

207

SiO2 (10 µm)

0.336

25

209

3.4. Comparison Table 10 to Table 14 show that the best candidates for the coatings is a 25 µm grooves relief, combined with an aluminum thin film of 100 nm and a 3.5 µm SiO2 layer (see Figure 31 and its spectra on Figure 28 and Figure 29). This coating gives for the cold case scenario a temperature of the MMNS module of 9°C and for the hot case scenario, a temperature of 212°C, if it is considered that the module is alone in space (which of course it is not, the surrounding structure lowering the hot case temperature).

3.5 µm of SiO2 100 nm of Al 25 µm grooves

Figure 31 – Best candidate for the thermal coating of the MMNS module

This coating should be tested with the MMNS module to ensure that it does not endanger the other function of the module. To minimize the charge buildup issues that can occur into orbit, a metallic grid can be evaporated onto the dielectric to permit a better repartition of charged particles. This coating is the best that has been tested, but other might worth at least to conduct similar survey: a metallic layer of silver may lower further the solar absorptance [Hass and al, 1973] and a dielectric layer of silicon monoxide might prove itself more effective than SiO2 for the infrared emittance [Bradford and al, 1970]. The influence of a larger relief (50 and 100 µm grooves) should also be tested.

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Conclusion Different solutions have been tested to find the best way to implement an optical solar reflector on NanoSpace-1 modules. The solution found is a mix of microstructure processing and thin film deposition. The radiator area is increased thank to a micro-relief created directly in the upper wafer of the module by photolithography. A 100 nm aluminum layer deposited onto this relief prevents the module to absorb too much solar radiation and a coating of 3.5 µm of silicon dioxide increases the infra-red emittance of the module without having a high influence on its solar absorptance. This internship just touched on the complex research needed to develop the Multifunctional Microsystems. The required interdisciplinary knowledge and the working environment in a foreign country well corresponded to the ISU 3I’s, Interdisciplinary, Intercultural and International. The fourth I was the Interest to witness the development of a new philosophy of satellite design. Things are definitively changing in satellite design, and small organizations may take the lead over big and heavy companies. The dynamism of research linked to small satellites development is well illustrated by the creation in July 2005 of the Ǻngström Aerospace Corporation (ǺAC). The ǺAC will transform the research conducted in the Ǻngström Space Technology Center into commercial products. Already, spin-off of these initially designed-for-space products are appearing, since ǺAC’s products are suitable for use in a variety of vehicles other than miniaturized satellites. For instance, the S-band antenna module can be used in airplanes or cars. Multifunctional Microsystems allow extreme miniaturization and the construction of satellites of a few kilograms equivalent to a 300-kg traditional satellite. However, the development cost for MMS is high, but when a product has been developed, its production cost becomes very low thanks to batch manufacturing. The ǺAC, beneficing of the technology developed with the ǺSTC, is a very good example of successfully commercialized university research.

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References •

Futron Corporation, 2002, Space Transportation Costs: Trends in Price Per Pound to Orbit 1990-2000

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Werts, JR & Larson, WJ, 1999, Space Mission Analysis and Design, Space Technology Library, fifth printing, Microcosm Press, Torrance, California.

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Jilla, CD & Miller DW, 1997, Satellite Design: Past, Present and Future, viewed August 2005,

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Renewable Resource Data Center, 2000, Solar Spectra: Air Mass Zero, last viewed August 2005,

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Karam RD, 1998, Satellite Thermal Control for Systems Engineers, American Institute of Aeronautics and Astronautics

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Harada X & Mell RJ, 1983, Inorganic thermal control coatings - A review, American Institute of Aeronautics and Astronautics, Aerospace Sciences Meeting, 21st, Reno, NV; United States; 10-13 Jan. 1983. 9 pp. 1983

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Hass G, Heaney JB & Triolo JJ, 1973, Evaporated Ag coated with double layers of Al2O3 and SiO2 to Produce Surface Films with Low Solar Absorptivity and High Thermal Emissivity, Optics Communications, Vol 8, No 3, pp. 183-185

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Bradford AP, Hass G, Heaney JB & Triolo JJ, 1970, Solar Absorptivity and Thermal Emissivity of Aluminum Coated with Silicon Oxide Films Prepared by Evaporation of Silicon Monoxide, Applied Optics, Vol 9, No 2, pp. 339-344

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Appendix – Results

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