The Spherical Mobile Investigator for Planetary Surface (SMIPS) concept aims at making use of the latest developments within extreme miniaturization of space systems. The introduction of Microelectromechanical Systems (MEMSs) and higher level Multifunctional MicroSystems (MMSs) design solutions gives the robot high performance per weight unit. The untraditional spherical shape makes it easily maneuverable and thus provides a platform for scientific investigations of interplanetary bodies. Preliminary investigations of the SMIPS concept show several advantages over conventional robots and rovers in maneuverability, coverage, size, and mass. A locomotion proof-of-concept has been studied together with a new distributed on-board data system configuration. This paper discusses theoretical robot analysis, an overall concept, possible science, enabling technologies, and how to perform scientific investigations. A preliminary design of an inflatable multifunctional shell is proposed.

This thesis explores the potential of multifunctional silicon-based microsystems for advanced nanospacecraft (AIN). Especially, multifunctional microsystems with the coexistant functions of communications and thermal management implemented in multilayer silicon stacks are approached with systems study. Host vehicles, composed of microsystems, including micro/nano-spacecraft and spherical rovers are contemplated with respect to future performance and implications, system level design, and breadboard realizations. A module of great importance, named the integrated communications and thermal management system for advanced or ICTM, symbolizes the achievements within the field of self-contained microsystems and is a prioritized entity throughout the thesis. The ICTM is natively placable onboard all types of highly miniaturized craft.The single AIN spacecraft and future clusters of these are investigated with respect to future full scale implementation of space systems designed and implemented with the distributed reconfigurable nanospacecraft cluster (DRNC) concept. Here, a true entanglement of microsystems technology (MST) and miniaturized spacecraft technology can revolutionize the applications, cost, and span of conceivable space missions.An intended communications scenario supporting a data rate of 1 Mbps, for the transmitter, is achieved during 6 minutes with a maximum continuous power dissipation of 10 W. Thermal simulations support the expectation, of a thermally biased ICTM, that the module is capable of supporting this energy burst, by using the mechanisms of heat storage and heat switches, and still fulfilling the requirements imposed by AIN type of spacecraft. In addition, multiple functional surfaces for the ICTM are evaluated with respect to equilibrium temperature and process compatibility. The tailored surfaces provide temperature control using micromachining methods.A design of a micromachined Ka-band front end with several MST enabled features is presented including e.g. vias, phase-shifters, and antennas. Similar antennas have been manufactured resulting in an evaluation of ring- and slot-antennas on silicon substrate. Based on a primitive version of the ICTM, a S-band patch antenna has been successfully implemented and characterized. Included in the thesis is a microthruster, an enabling technology for DRNC.

Microsystem interfaces to the macroscopic surroundings and within the microsystems themselves are formidable challenges that this thesis makes an effort to overcome, specifically for enabling a spacecraft based entirely on microsystems. The NanoSpace-1 nanospacecraft is a full-fledged satellite design with mass below 10 kg. The high performance with respect to mass is enabled by a massive implementation of microsystem technology – the entire spacecraft structure is built from square silicon panels that allow for efficient microsystem integration. The panels comprise bonded silicon wafers, fitted with silicone rubber gaskets into aluminium frames. Each module of the spacecraft is added in a way that strengthens and stiffens the overall spacecraft structure. The structural integrity of the silicon module as a generic building block has been successfully proven. The basic design (silicon, silicone, aluminium) survived considerable mechanical loads, where the silicon material contributed significantly to the strength of the structural element. Structural modeling of the silicon building blocks enables rapid iterative design of e.g. spacecraft structures by the use of pertinent model simplifications. Other microsystem interfaces treats fluidic, thermal, and mechanical functions. First, solder sealing of microsystem cavities was demonstrated, using screen-printed solder and localized resistive heating in the microsystem interface. Second, a dismountable fluidic microsystem connector, using a ridged silicon membrane, intended for monopropellant thruster systems, was developed. Third, a thermally regulated microvalve for minute flows, made by a silicon ridge imprint in a stainless steel nipple, was investigated. Finally, particle filters for gas interfaces to microsystems, or between parts of fluidic microsystems, were made from sets of crossed v-grooves in the interface of a bonded silicon wafer stack. Filter manufacture, mass flow and pressure drop characterization, together with numeric modeling for filter design, was performed. All in all this reduces the weight and volume when microsystems are interfaced in their applications.

This thesis explores the potential of multifunctional silicon-based microsystems for advanced nanospacecraft (AIN). Especially, multifunctional microsystems with the coexistant functions of communications and thermal management implemented in multilayer silicon stacks are approached with systems study. Host vehicles, composed of microsystems, including micro/nano-spacecraft and spherical rovers are contemplated with respect to future performance and implications, system level design, and breadboard realizations. A module of great importance, named the integrated communications and thermal management system for advanced or ICTM, symbolizes the achievements within the field of self-contained microsystems and is a prioritized entity throughout the thesis. The ICTM is natively placable onboard all types of highly miniaturized craft.The single AIN spacecraft and future clusters of these are investigated with respect to future full scale implementation of space systems designed and implemented with the distributed reconfigurable nanospacecraft cluster (DRNC) concept. Here, a true entanglement of microsystems technology (MST) and miniaturized spacecraft technology can revolutionize the applications, cost, and span of conceivable space missions.An intended communications scenario supporting a data rate of 1 Mbps, for the transmitter, is achieved during 6 minutes with a maximum continuous power dissipation of 10 W. Thermal simulations support the expectation, of a thermally biased ICTM, that the module is capable of supporting this energy burst, by using the mechanisms of heat storage and heat switches, and still fulfilling the requirements imposed by AIN type of spacecraft. In addition, multiple functional surfaces for the ICTM are evaluated with respect to equilibrium temperature and process compatibility. The tailored surfaces provide temperature control using micromachining methods.A design of a micromachined Ka-band front end with several MST enabled features is presented including e.g. vias, phase-shifters, and antennas. Similar antennas have been manufactured resulting in an evaluation of ring- and slot-antennas on silicon substrate. Based on a primitive version of the ICTM, a S-band patch antenna has been successfully implemented and characterized. Included in the thesis is a microthruster, an enabling technology for DRNC.

With the miniaturization of spacecraft the need for efficient, accurate and low-weight attitude control systems is becoming evident. To this end, the cold/hot gas microthruster system of this paper incorporates carbon microcoils—deposited via laser-induced chemical vapor deposition—for heating the propellant gas (nitrogen) before the nozzle inlet. By increasing the temperature of the propellant gas for such a system, the specific impulse (Isp) of the microthruster will increase. The benefits of a higher Isp are lower propellant mass, higher thrust and shorter burning times. Therefore, the feasibility of achieving this increase with the carbon microcoils is investigated. The carbon microcoils have been characterized experimentally with respect to their electrothermal performance, i.e. resistance, temperature, parasitic heat losses and degradation in ambient. The resulting heat losses from the heater and the heated gas have been estimated through a combination of experiments, numerical simulation and approximate analytical expressions. At high powers, degradation of the carbon material leads to coil failure in ambient where trace oxygen was present. Thus, the next generation of carbon microcoils to be tested will have a protective coating to extend their lifetime. Theoretical modeling showed that an increase in the propellant gas temperature from 300 to 1200 K and a corresponding two-fold increase in the Isp can be achieved if 1.0 W of power is supplied to each coil in a three-coil thruster. These simulation results show that if the coils are capable of dissipating 1 W of heat at 1700 K coil temperature, the doubling of the Isp may be achieved. Comparing to the electrothermal characterization results we find that the carbon coils can survive at 1700 K if protected, and that they can be expected to reach 1700 K at power below 1 W.

To use the high mass fraction of silicon material in a nanosatellite based on microelectromechanical systems, part of the structural function has been assigned to the flat silicon stacks embracing these systems. Three modules for destructive testing in bending, warping, and shearing cases were built with 68 x 68 x 1 mm silicon stacks bonded in aluminium frames by in situ casting of silicone rubber. A special module of the same size was built with strain gauges of Nichrome. Elastic deformation tests on this and simulations using finite element analysis were performed for bending, warping, and shearing loads of up to 80,40, and 99 N, respectively. The correlation between simulations and experimental measurements was good with deviation of about 30%. The results show that the rubber works well as a mechanical interface between the stiff and brittle silicon stacks and their weaker and ductile aluminium frames. Its thickness influences the stress in the silicon stack significantly. The silicon stack stiffens the module by a factor of 46 and lowers the stress in its frame 24 times in shearing mode, which is the most relevant loading case for the satellite framework. Thus, the concept of using flat silicon panels as structural elements is fully feasible.

Spacecraft on interplanetary missions or advanced satellites orbiting the Earth all require propulsion systems to complete their missions. Introducing microelectromechanical systems technology to the space industry will not only reduce size and weight of the propulsion system, but can also increase the performance of the mission.Fluid handling systems are used in chemical and electric propulsion. Some components incorporated in a fluidic handling system are presented and evaluated in this work.Microsystems are very sensitive to contamination. Reliable, robust, and easily integrated filters were modeled, manufactured, and experimentally verified.A fluid connector, designed to withstand large temperature variations and aggressive propellants was manufactured and characterized. Similar designs was also be used as a thermally activated minute valve.The feasibility of a cold gas system for precise attitude control has been demonstrated. Steps towards improving the performance (from specific im-pulse 45 s) have been taken, by the integration of suspended heater elements.For electric propulsion, two thermally regulated flow restrictors have been characterized. These devices can fine-tune the propellant flow to e.g. an ion engine.A single-use valve using a soldered seal has also been successfully dem-onstrated within a pressure range of 5 to 100 bar.The microsystem-based propulsion systems of tomorrow’s spacecraft need to be demonstrated in space, in order to gain necessary credibility.

Microsystem interfaces to the macroscopic surroundings and within the microsystems themselves are formidable challenges that this thesis makes an effort to overcome, specifically for enabling a spacecraft based entirely on microsystems. The NanoSpace-1 nanospacecraft is a full-fledged satellite design with mass below 10 kg. The high performance with respect to mass is enabled by a massive implementation of microsystem technology – the entire spacecraft structure is built from square silicon panels that allow for efficient microsystem integration. The panels comprise bonded silicon wafers, fitted with silicone rubber gaskets into aluminium frames. Each module of the spacecraft is added in a way that strengthens and stiffens the overall spacecraft structure. The structural integrity of the silicon module as a generic building block has been successfully proven. The basic design (silicon, silicone, aluminium) survived considerable mechanical loads, where the silicon material contributed significantly to the strength of the structural element. Structural modeling of the silicon building blocks enables rapid iterative design of e.g. spacecraft structures by the use of pertinent model simplifications. Other microsystem interfaces treats fluidic, thermal, and mechanical functions. First, solder sealing of microsystem cavities was demonstrated, using screen-printed solder and localized resistive heating in the microsystem interface. Second, a dismountable fluidic microsystem connector, using a ridged silicon membrane, intended for monopropellant thruster systems, was developed. Third, a thermally regulated microvalve for minute flows, made by a silicon ridge imprint in a stainless steel nipple, was investigated. Finally, particle filters for gas interfaces to microsystems, or between parts of fluidic microsystems, were made from sets of crossed v-grooves in the interface of a bonded silicon wafer stack. Filter manufacture, mass flow and pressure drop characterization, together with numeric modeling for filter design, was performed. All in all this reduces the weight and volume when microsystems are interfaced in their applications.

A silicon-based integrated communications and thermal management microsystem qualifying for use on Nanospace-1, a modularized microsystem-based advanced integrated nanospacecraft, is presented. The transmitter and receiver share the same module framework with essential differences only in the electronics implementation. A data rate of 1 Mbps for the transmitter and 114 kbps for the receiver is accomplished with a transmitter power for the spacecraft and ground station of 2 and 10 W, respectively. Concurrent triple usage of paraffin as low loss antenna substrate, actuator material, and heat sink is designed and analyzed for the first time. On low-power or short-time high-power dissipation of heat from the electronics, energy is stored as latent heat in this phase-change material acting as a heat sink. Thermal transport through the module is initiated,by actuation of thermal switches when 75% of the paraffin's latent heat is consumed. A static thermal analysis reveals a thermal modulation factor of 5.6 between the on and off states of the thermal switches. The size of the module is 6.6 x 68 x 68 mm, and its weight is 43 g.

This paper presents the development of a miniaturized, thin film sun sensor for sun angle detection in space applications. The overall design, different thin film layer sequences, processes, materials used and system integration are described. The sun sensor has a field of view (FoV) of greater than 2π sr (±90°) and shall have a resolution of approximately 1° in elevation and azimuth angle. This sun sensor is therefore a coarse sun sensor with the advantage of having a large field of view. Key elements are the curved shape, the photosensitive layer consisting of copper indium gallium diselenide (CIGS), the transparent conductive layer consisting of thin molybdenum or aluminum-doped zinc oxide, and its integrated design. Different possible layer sequences are presented and discussed. The sun sensor will be one of the sensors in the attitude determination and control system (ADCS) of a nano-satellite.

Given the present, relatively limited deployment of low cost and mass space missions, there are clear opportunities for the application of small-scale propulsion systems in further enabling these small satellite missions. With this situation in mind, a team comprising ASTC, SSTL, TNO and QinetiQ – under funding from the European Space Agency – has undertaken the development of a MEMS-based micro-rocket engine concept with the intended capability of providing a specific impulse of greater than 100s. Both turbo-pump fed bi-propellant as well as mono propellant concepts were investigated. For demonstration a mono propellant design was selected of which an initial design concept was developed, based on hydrogen peroxide decomposition. Identified as the critical component in the mono-propellant system, several batches of the honeycomb wafers – upon which the decomposition occurs – have been manufactured. As a proof of concept, the wafers have been subjected to a set of structural and functionality tests. Given the results of the latest testing initiative, it is envisaged that, with adequate refinement and development of the current design, a reliable and deployable monopropellant micro-rocket engine solution may be realised. INTRODUCTION Given the present, relatively limited deployment of low-cost, low-mass space missions, there are clear opportunities for the application of small-scale propulsion systems in further enabling these small satellite missions. Requirements from Micro and nano satellites for constellation forming, LEO (low Earth orbit) drag compensation, manoeuvring for satellite inspection, formation flying, and end-of-life de-orbiting all indicate a need for significant ∆V (velocity change) capability from such a propulsion system, ideally coupled with high thrust levels, thus minimising energy losses. The requirement for such a technology is particularly evident when considering a problem commonly faced by low-cost missions; securing a precise orbit may often be compromised as a consequence of a shared or secondary launch. It is difficult to provide a suitable solution with conventionally engineered, miniature propulsion systems. Typically, these are expensive solutions, using toxic propellants that, in many cases, do not scale favourably to the sizes demanded by small satellite technology [1]. Micro System Technology based chemical propulsion is a candidate technology with the potential to fill the near term performance gap in the market for small satellites, defined by the lack of proposed systems providing an Isp in the range 100-300s. Such a system could conceivably be used for such tasks as orbit modification in microand nano-satellites, end-of-life de-orbit manoeuvring, and so on. Overview of MST propulsion research work Despite an increasing interest in miniaturised propulsion systems [2], very few MEMS-based chemical propulsion systems exist beyond the concept or breadboard stage. For example, the Ångström Space Technology Centre (ÅSTC) at Uppsala University has built and tested a MEMS hybrid system with a specific impulse (Isp) of 45s, with up to 100s projected for future developments [3]. Further, the Gas Turbine Laboratory at MIT has undertaken an in-depth development programme with the objective of demonstrating a high performance bipropellant system with a target Isp of 300s, a thrust of 15N, and with propellant feed achieved using a MEMSbased centrifugal turbo-pump [4]. These two concepts demonstrate the potential to be offered by such technology, and also serve to illustrate the wide gap in performance which exists between those systems currently being considered. NASA’s Goddard, together with Vermont University has been working on a HTP mono propellant engine. Problems with modelling were encountered and testing was performed on the engine, but no complete decomposition was reached [8]. In Austria, a team led by Mechatronics is working on a turbopump fed micro engine bi-propellant micro engine. This engine is small, but built out of standard materials [9]. Other work on the field of liquid micro propulsion is being performed in China and Japan. Different teams in Europe and the United States are also working on micro solid propellant engines. An example is the work performed by a team lead by CNRS [10]. REQUIREMENTS AND TECHNOLOGY SELECTION To establish a baseline for the micro-rocket engine concept, a top-level specification was derived from perceived market requirements (space mission level), and the current performance of state-of-the-art 1N hydrazine monopropellant thrusters. The key features of this baseline specification are: − Isp 100-150s, scalable in bipropellant mode to ~230s; − Propellant mass throughput of 200g (minimum) and 6000g (target), thus establishing the useful lifetime for the engine; − Power draw of 1W (nominal) and 3W (maximum), and with the possibility to extend this to 10W for small satellites in the >1kg category; − Minimum of 100 restarts / firings (a target of 1000 would be desirable to achieve competition with an industry standard hydrazine system); − Minimum continuous burn duration of 3 minutes; − Maximum mass of thruster and integrated propellant flow control system of 100g (system dry mass excluding tankage). A future target would be to achieve a 100g total dry mass including integrated propellant tank (i.e. total propulsion system mass); − Decomposition chamber pressure of 0.4MPa minimum, with 1.0MPa target for bipropellant mode. − An achievable thrust range of 20mN, through 200mN (nominal), and up to about 0.5-1.0N (preferred); − A minimum auto-ignition delay time of 50ms; − A peak decomposition temperature of around 900K assuming an uncooled silicon construction; The micro-rocket design being pursued should have the potential to be manufactured in silicon-based materials that are traditionally used in MEMS fabrication (i.e. good processing heritage exists). A technology selection roadmap, covering miniature pump technologies (for pressurised feed of propellant), micro-ancillary components (principally valves and environment sensors), and materials selection criteria, has been drawn up to aid in the monopropellant design concept development. Propellant selection The selection of a propellant well suited to meeting the defined requirement specification was accomplished through a semiFigure 1: A system block diagram for a pressure-fed monopropellant micro rocket engine, readily transferable to a bipropellant concept as part of a future development programme quantitative trade-off, in which potential propellant options were scored according to different, weighted criteria. The criteria considered in this exercise were: chemical and physical properties, storage and handling properties, performance properties, and technology and hardware. For a micro-rocket engine the process of selecting a suitable propellant differs substantially from that which would be used for a larger scale system. In a micro-engine the different criteria are much more interaction with each other Thus favouring systems that have an overall good trade-off performance instead of systems that are very strong in only one or two points. A number of possible monopropellants were examined in this trade-off, the three most promising candidates being concentrated (90%) H2O2, otherwise known as High Test Peroxide (HTP), N2O and high performance HNF (hydrazinium nitroformate). Out of this down-selected list, HTP was selected as offering most promise for the following reasons: − N2O: Techniques for achieving effective catalytic decomposition of this monopropellant are insufficiently mature; − HNF (and other propellants based on energetic salts): Considerable propellant development is required; − HTP: A well understood propellant from the point of view of achieving catalytic decomposition in a propulsion system. Other factors influencing this choice include prior experience, performance (relative to cold gas), low toxicity and applicability to a bipropellant system when used as an oxidizer. For the bi-propellant motor a hybrid motor using HTPB as a fuel and oxygen or HTP as oxidiser gained the highest score. This was mainly due to the ease of handling and storage and technology status of this combination. The second best was HTP in combination with kerosene or pentane, essentially for the same reasons. As the assignment was to develop a liquid motor and not a hybrid, the last combination was selected for the bi-propellant study. Kerosene was selected over pentane as more knowledge and experience was available on kerosene. Propellant feed strategies For the micro propulsion systems, the decomposition chamber pressure (Pc) and the resulting feed requirement were reviewed in detail. High Pc is desirable in micro-rockets principally to increase residence time of the propellant(s), compensating for small chamber sizes and high heat losses and maintaining c* (characteristic velocity) despite fixed chemical reaction times. However, high Pc in a pressure fed system poses challenges for MEMS fabrication, principally due to limited bond strengths and planar designs ill-suited to act as pressure vessels. Pumped propellant would be attractive if small high power density pumps can be fabricated. Analyses considered 1-100kg wet mass missions with 0.1-10kg of propellant respectively, with Pc values between 0.2 and 2.0MPa and a flow rate envelope of 0.01 to 0.5g/s (equivalent to a thrust of 20-500mN for an Isp of 100-150s). A pump fed propulsion system was considered to be competitive at low chamber pressures for the smallest (1kg) missions, a fully packaged pump mass of 15-50g being required. 10-100kg missions would benefit from 0.1-1kg mass pumps, which might also be able to meet the system requirements (head, flow rate) if based on miniaturized conventional materials. Several such pumps with 1-5kg mass are under test, for example [5], and provide a basis for future development. However MEMS micropump techno

A generic dielectric constant reduction method for silicon antenna substrates is presented in detail along with a process description to produce functional dielectric layers for planar antennas. Virtually any dielectric constant below 11.9 down to 3.8 aimed for in this paper can be produced. Very small honeycomb cells with wall thickness of 16 /spl mu/m and inner wall length of 86.6 /spl mu/m is etched down using deep reactive ion etch (DRIE) to 475 /spl mu/m depth in each of two 525 /spl mu/m 4-inch high ohmic wafers. These two wafers are bonded together with the etched side of both wafers facing each other. A volumetric averaging yields an average dielectric constant of 3.8 for the two bonded wafers. By adjusting the etch depth, different dielectric constants for bonded pairs of silicon wafers are attainable. A demonstration of the concept has been physically realized showing an increase in the resonance frequency of a simple coaxial-fed disk-patch antenna with a simulated resonance frequency of 2.5 GHz.

This thesis describes and evaluates the design of nanospacecraft based on advanced multifunctional microsystems building blocks. These systems bring substantial improvements of the performance of nanosatellites and enable new space exploration, e.g. interplanetary science missions using minute space probes. Microsystems, or microelectromechanical systems, allows for extreme miniaturization using heritage from IC industry. Reducing mass and volume of spacecraft gives large savings in terms of launch costs. Definition and categorization of system and module level features in multifunctional microsystems are used to derive a spacecraft optimization algorithm which is compatible with commonly used concurrent engineering methods. The miniaturization of modules enables modular spacecraft architectures comprising powerful multifunctional microsystems, which are applicable to satellites between 10 and 1000’s of kg. This kind of complete spacecraft architecture has been developed for the NanoSpace-1 technology demonstrator satellite. The spacecraft bus uses multifunctional design to enable distributed intelligence and autonomy, graceful degradation, functional surfaces, and distributed power systems. The increase in performance of the new spacecraft architecture as compared with conventional nanosatellites is orders of magnitudes in terms of power storage, scientific payload mass ratio, pointing stabilization, and long time space operation. This high-performance system-of-microsystems architecture has been successfully employed on two space robotic concepts: a miniaturized submersible vehicle for Jupiter’s Moon Europa and a miniaturized spherical robot. The submersible is enabled by miniaturization of electronics into 3-dimensional, vertically integrated multi-chip-modules together with new interconnection methods. These technologies enabled the submersible vehicle tube-shaped design within 20 cm length and 5 cm diameter. The spherical rover was developed for long range and networked science investigations of interplanetary bodies. The rover weighs 3.5 kg and is shown to endure direct reentry on Mars, which increases the ratio between the landed mobile payload mass and the initial mass in Mars orbit by a factor of 18.

In this paper, a detailed design of a monopropellant microrocket engine demonstrator model is described, together with the thermal modelling and directions for manufacture and assembly. In order to reach the goal of an extremely miniaturized solution, the design draws heavily on Micro Electro Mechanical Systems (MEMS) manufacturing of silicon. The thermal modelling of the monopropellant decomposition chamber (using Hydrogen Peroxide, HTP, decomposed in a catalyst pack) shows that silicon as the engine bulk material can endure the thermal loads applied. Furthermore, the use of a suspended catalyst bed and a thermal standoff layer below the catalyst should enable rapid ignition times. The design is made by stacking structured silicon wafer parts together in order to form the main motor package. The thermal standoff bridges this part to the bottom interface part (electric and fuel) with a Pyrex wafer. The motor package is completed with a precision-machined boron nitride nozzle. The whole assembly is clamped and bonded together and reinforced by pin bolts. The model engine sits on four pipe legs, two for the fuel inlet and two for electric lines and pressure feedback to the fuel distribution unit. The pipe legs also assist in isolating the hot engine parts from the rest of the spacecraft.

A Xenon feed system, manufactured by the means of micro machining, is under development. The system comprises several multifunctional modules described in the paper. A large part of the effort has been directed to the flow control unit, in which the flow is modulated using a thermal flow regulator. The regulator includes a heater and a temperature sensor, and has been manufactured and evaluated. A flow sensor is also included in the flow control unit. The mass for the xenon feed system, including housing and electronics is estimated to 150g.