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Cold gas microthruster
Abstract
This chapter covers the development of cold gas microthruster, which is widely regarded as the simplest way of generating thrust in space, for nanosatellites. A brief background and principles of operation were given, followed by the introduction of nozzle theory that could be used in the preliminary estimation of microthruster performance and considerations in selecting a suitable propellant. The current state of development in cold gas microthruster (at the time of writing in 2020) was provided, from its first use in SNAP-1 in 2000 to another 12 nanosatellites as well as one technology demonstration mission in Prototype Research Instruments and Space Mission technology Advancement small satellite. The chapter ends with a discussion on future nanosatellite missions that will feature a cold gas microthruster system and the challenges, such as fabrication of micronozzle and its design optimization, to improve overall efficiency.
... solid propellant [9,10] cold gas [11,12], liquid propellant [13][14][15], and electrospray [16,17], significantly to fit into the design envelop of cubesats. For example, MEMS-based cold gas microthruster has been space qualified in cubesat platform [18][19][20]. Nevertheless, a heavy tank for pressurized gas is still a limiting factor for further application into the smaller PocketQube. ...
... The micronozzle was designed using the quasi 1D nozzle flow theory [20]. The nozzle throat diameter was set as 500 μm, which is not arbitrary but a limitation due to the printing resolution of the 3D printer and the transparent resin, which cannot prevent some of the UV light from scattering into the adjacent layers. ...
Vaporizing liquid microthruster is a relatively new concept for evaporation based propulsive system. The research interest in vaporizing liquid microthruster mainly comes from its simplicity and the use of green propellant: water. This paper presents the study of a highly compact additively manufactured vaporizing liquid microthruster of 38 × 22 × 8 mm in dimension. The heat transfer enhancement effect of micro pin fins for vaporizing process has been investigated using modeling and computational fluid dynamics simulation, followed by experiments to validate the results. It has been shown that all micro pin fins (square, diamond, circular) with different array arrangements (staggered and inline) can enhance heat transfer. Square shaped pin fin exhibited the best heat enhancement performance with an increase in dimensionless total thermal resistance of 14.65% (as compared to blank vaporizing liquid microthruster) due to its large contacting area, while the circular shape was least effective. Meanwhile, the inline arrangement was found to induce a significant fluctuation in transverse flow velocity between the columns, which promotes heat convection. The vaporizing liquid microthruster prototype was produced using a specialty resin of good heat resistance up to 350 °C as confirmed by the thermogravimetric analysis. Thrust performance of the prototype was measured using a torsional thrust stand in a vacuum chamber. The vaporizing liquid microthruster with square micro pin fins and in inline configuration has produced the highest thrust of 740.2 μN at 2 μl/s flow rate, giving a specific impulse of 37.7s. The developed prototype could be implemented as micropropulsion system for nanosatellites, such as CubeSat and PocketQube.
The micro-nozzle is an important component of the micropropulsion system for accelerating the gas flow. However, the 3D ceramic micro-nozzle is difficult to machine due to its small complex structure and high hardness of material. In this study, micro-electrical discharge machining (EDM) segmented milling is proposed to fabricate the ZrB2–SiC–graphite ceramic micro-nozzle. To improve the machining efficiency and reduce the electrode wear, the micro-nozzle is machined using two tool electrodes with different diameters. The inlet and outlet segments of micro-nozzle are machined by the large diameter electrode, and the throat segments are fabricated with the small diameter electrode. The effects of layer thickness, tool electrode feedrate and electrode diameter on machining performances are investigated in preliminary experiments, and the proper machining parameters are determined. The processing tool-path is planned for producing the micro-nozzle with consideration of the dimension error induced by the machining gap. The micro-nozzle with the conical convergent section and parabolic divergent section is successfully machined, exhibiting high processing efficiency and machining accuracy. It is demonstrated that the proposed micro-EDM segmented milling method is feasible for fabricating the ZrB2–SiC–graphite ceramic micro-nozzle.
The increasingly successful use of nanosatellites has made CubeSat an attractive form factor for a number of missions that require both an attitude control system (ACS) and primary propulsion. The CubeSat High Impulse Propulsion System (CHIPS), under development by CU Aerospace with partner VACCO Industries, is designed to enable CubeSat mission operations beyond simple orbit maintenance, including significant altitude changes, formation flying, and proximity operations such as rendezvous and docking. CHIPS integrates primary propulsion, ACS and propellant storage into a single bolt-on module that is compatible with a variety of non-toxic, self-pressurizing liquid propellants. The primary propulsion system uses micro-resistojet technology developed by CU Aerospace to superheat the selected propellant before subsequent supersonic expansion through a micro-nozzle optimized for frozen-flow efficiency. The 1U+ baseline design is capable of providing an estimated 563 N-s total impulse at 30 mN thrust using R134a propellant, giving an impulse density of 552 N-s/liter. The ACS is a cold gas, 4-thruster array to provide roll, pitch, yaw, and reverse thrust with a minimum impulse bit of 0.4 mN-s. An engineering prototype, with integrated pressure control, power control, and data logging, was extensively tested in cold and warm gas modes on the University of Illinois thrust stand under a range of conditions. This paper presents thrust and specific impulse data for the resistojet thruster using two different propellants. Resistojet data include cold and warm gas performance as a function of mass flow rate, plenum pressure, geometry, and input power.
The CubeSat High Impulse Propulsion System (CHIPS), under development by CU Aerospace with partner VACCO Industries, is a system targeted at CubeSat and micro-satellite platforms designed to enable mission operations beyond simple orbit maintenance, including significant altitude changes, formation flying, and proximity operations such as rendezvous and docking. CHIPS integrates primary propulsion, ACS and propellant storage into a single bolt-on module that is compatible with a variety of non-toxic, self-pressurizing liquid propellants. The primary propulsion system uses micro-resistojet technology developed by CU Aerospace to superheat the selected propellant before subsequent supersonic expansion through a micro-nozzle optimized for frozen-flow efficiency. A 1U design is estimated to be capable of providing 471 N-s total impulse at ~20 mN thrust using R236fa propellant, resulting in a volumetric impulse of 525 N-s/liter. The ACS is a cold gas, 4-thruster array to provide roll, pitch, yaw, and reverse thrust with a minimum impulse bit of 0.4 mN-s. A high-fidelity, fully-integrated engineering unit was extensively tested in cold and warm gas modes on the University of Illinois thrust stand under a range of conditions. This paper presents thrust and specific impulse data for the resistojet thruster using two different propellants. Resistojet data include cold and warm gas performance as a function of mass flow rate, plenum pressure, geometry, and input power. Presently the CHIPS prototype is TRL 5 as a result of a Phase II SBIR effort funded by NASA.
Since combustion is an easy way to achieve large quantities of energy from a small volume, we developed a MEMS based solid propellant microthruster array for small spacecraft and micro-air-vehicle applications. A thruster is composed of a fuel chamber layer, a top-side igniter with a micromachined nozzle in the same silicon layer. Layers are assembled by adhesive bonding to give final MEMS array. The thrust force is generated by the combustion of propellant stored in a few millimeter cube chamber. The micro-igniter is a polysilicon resistor deposited on a low stress SiO 2 /SiN x thin membrane to ensure a good heat transfer to the propellant and thus a low electric power consumption. A large range of thrust force is obtained simply by varying chamber and nozzle geometry parameters in one step of Deep Reactive Ion Etching (DRIE). Experimental tests of ignition and combustion employing home made (DB+x%BP) propellant composed of a Double-Base and Black-Powder. A temperature of 250 °C, enough to propellant initiation, is reached for 40 mW of electric power. A combustion rate of about 3.4 mm/s is measured for DB+20%BP propellant and thrust ranges between 0.1 and 3,5 mN are obtained for BP ratio between 10% and 30% using a microthruster of 100 µm of throat wide.
In this study, we numerically examine thrust performance of the linear aerospike nozzle micro-thruster for various nozzle spike lengths and flow parameters in order to identify optimal geometry(s) and operating conditions. Decomposed hydrogen-peroxide is used as the monopro-pellant in the studies. Performance is characterized for different flow rates (Reynolds numbers) and aerospike lengths, and the impact of micro-scale viscous forces is assessed. It is found that 2-D full micro-aerospike efficiencies can exceed axisymmetric micro-nozzle efficiencies by as much as 10%; however, severe penalties are found to occur for truncated spikes at low Reynolds numbers.
Diamond is a highly desirable material for use in x-ray optics and instrumentation. However, due to its extreme hardness and resistance to chemical attack, diamond is difficult to form into a structure suitable for x-ray lenses. Refractive lenses are capable of delivering x-ray beams with nanoscale resolution. A moulding technique for the fabrication of diamond lenses is reported. High-quality silicon moulds were made using photolithography and deep reactive ion etching. The study of the etch process conducted to achieve silicon moulds with vertical sidewalls and minimal surface roughness is discussed. Issues experienced when attempting to deposit diamond into a high-aspect-ratio mould by chemical vapour deposition are highlighted. Two generations of lenses have been successfully fabricated using this transfer-moulding approach with significant improvement in the quality and performance of the optics observed in the second iteration. Testing of the diamond x-ray optics on the Diamond Light Source Ltd synchrotron B16 beamline has yielded a line focus of sub-micrometre width.
It is now feasible to construct highly capable ‘nanosatellites’ (i.e. sub–10 kg satellites) to provide cost–effective an rapid–response orbiting test vehicles for advanced space missions and technologies. The UK's first nanosatellite, SNAP–1—designe and built by Surrey Space Centre (SSC) and Surrey Satellite Technology Ltd (SSTL) staff — is an example of such a test vehicle in this case, built with the primary objective of demonstrating that a sophisticated, fully agile nanosatellite can be constructe rapidly, and at very low cost, using an extension of the modular–COTS–based design philosophy pioneered by SSC for its microsatellites.
SNAP–1 was successfully launched into orbit on 28 June 2000 from the Plesetsk Cosmodrome on board a Russian Cosmos rocket.
It flew alongside a Russian Cospas–Sarsat satellite called Nadezhda, and an SSTL–built Chinese microsatellite, called Tsinghua–1.
The first year of operations has been highly successful, with SNAP–1 becoming the first nanosatellite to have demonstrate full attitude and orbit control, via its miniature momentum–wheel–based attitude control system and its butane–propellant–base propulsion system. This paper reviews the initial results of the SNAP–1 mission.
To enable formation flying of microsatellites, small sized propulsion systems are required. Our research focuses on the miniaturization of a feeding and thruster system by means of microsystem technology (MST). Three fabrication methods have been investigated to make a conical converging–diverging nozzle. These methods are deep reactive ion etching (DRIE), femtosecond laser machining (FLM) and a combination of powderblasting and heat treatment. It is shown that the latter two methods are very promising.
Cold gas micropropulsion is a sound choice for space missions that require extreme stabilisation, pointing precision or contamination-free operation. The use of forces in the micronewton range for spacecraft operations has been identified as a mission-critical item in several demanding space systems currently under development. Cold gas micropropulsion systems share merits with traditional cold gas systems in being simple in design, clean, safe, and robust. They do not generate net charge to the spacecraft, and typically operate on low-power. The minute size is suitable not only for inclusion on high-performance nanosatellites but also for high-demanding future space missions of larger sizes. By using differently sized nozzles in parallel systems the dynamic range of a cold gas micropropulsion system can be quite wide (e.g. 0 – 10 mN), while the smallest nozzle pair can deliver thrust of zero to 0.5 or 1 mN using continuously proportional gas flow control systems. The leakage is turned into an advantage enabling the system for continuous drag compensation. In this manner, the propellant mass efficiency can be many times as higher than that in a conventional cold gas propulsion system using ON-OFF-control. The analysis in this work shows that cold gas micropropulsion has emerged as a high-performance propulsion principle for future state-of-the-art space missions. These systems enable spacecraft with extreme demands on stability, cleanliness and precision, without compromising the performance or scientific return of the mission.
The objective of the Canadian Advanced Nanospace eXperiment (CanX) program is to develop highly capable "nanospacecraft," or spacecraft under 10 kilograms, in short timeframes of 2-3 years. CanX missions offer low-cost and rapid access to space for scientists, technology developers, and operationally responsive missions. The Space Flight Laboratory (SFL) at the University of Toronto Institute for Aerospace Studies (UTIAS) has developed the Canadian Advanced Nanospace eXperiment 2 (CanX-2) nanosatellite that launched in April 2008. CanX-2, a 3.5-kg, 10 x 10 x 34 cm satellite, features a collection of scientific and engineering payloads that push the envelope of capability for this class of spacecraft. The primary mission of CanX-2 is to test and demonstrate several enabling technologies for precise formation flight. These technologies include a custom cold-gas propulsion system, a 30 mNms nanosatellite reaction wheel as part of a three-axis stabilized momentum-bias attitude control system, and a commercially available GPS receiver. The secondary objective of CanX-2 is to fly a number of university experiments including an atmospheric spectrometer. At the time of writing CanX-2 has been in orbit for three weeks and has performed very well during preliminary commissioning. The mission, the engineering and scientific payloads, and the preliminary on-orbit commissioning experiences of CanX-2 are presented in this paper.
The next-generation of small satellites ('nanosats') will feature masses 10 kg and require miniaturised propulsion systems capable of providing extremely low levels of thrust. The emergence of viscous, thermal and/or rarefaction effects on the micro-scale can significantly impact the flow behaviour in supersonic micronozzles resulting in performance characteristics which differ substantially from traditional macro-scale nozzle designs. In this paper, we provide an overview of key findings obtained from computational studies of supersonic micronozzle flow and discuss the implications for future micro-scale nozzle design and optimisation.
Micro spacecraft, which have gained huge popularity in the last decade, are termed as “CubeSats”, a well-known class of small satellites. The fast growing functionality and popularity of CubeSats have helped researchers push technology demonstration towards efficient performance and reliability needed for commercial and governmental applications. Keeping in mind the increasing space debris, dependence of satellite life on fuel, and the use of fossil fuels as propellants, recent efforts are being made to develop a cold gas propellant-based (CGP) micro propulsion system. The CGP system has various merits, namely, non-toxicity, easy to use, and low leakage concerns over other propulsion systems. Liquid propellant-based propulsion requires a liquid feed system, which uses pressurized gas (such as helium, argon and nitrogen) to pressure-feed liquid to the combustion or vaporization chamber so as to produce thrust. But once the liquid propellant is depleted, the pressurized gas is unusable while the system runs out of fuel. Thus, the proposed CGP system being studied, apart from being used as a normal cold gas propulsion system, also offers the advantage of being used as a propulsion system for a liquid fueled system, in the event the liquid fuel runs out. The present work reports on cold gas micro thruster development, utilization of feed gas in a liquid fueled thruster, and the experimental study on the CGP system in simulated vacuum conditions and under a range of pressure conditions. Pressure in the pressure feed gas system for a 1–50 kg category of satellites varies from 4 to 8 bar depending on the mission requirement. The pressure reduces as the liquid propellant gets exhausted. The experimental results are used to validate a computational fluid dynamics model of the system. Thrust values are obtained between the micro to milli-Newton range so as to fulfil the requirements of attitude and station keeping for CubeSats in the 1–50 kg dry mass range. Under vacuum conditions, we obtained thrust values of 0.8 mN at 1 bar feed pressure to 2.24 mN at 4 bar feed pressure. These thrust values are nearly twice those achieved for sea level tests for the corresponding feed pressure. Furthermore, the parameters such as Mach number, velocity vector, pressure, temperature, specific impulse, and nozzle efficiency are studied and reported for atmospheric and vacuum conditions.
At present, very few CubeSats have flown in space featuring propulsion systems. Of those that have, the literature is scattered, published in a variety of formats (conference proceedings, contractor websites, technical notes, and journal articles), and often not available for public release. This paper seeks to collect the relevant publically releasable information in one location. To date, only two missions have featured propulsion systems as part of the technology demonstration. The IMPACT mission from the Aerospace Corporation launched several electrospray thrusters from Massachusetts Institute of Technology, and BricSAT-P from the United States Naval Academy had four micro-Cathode Arc Thrusters from George Washington University. Other than these two missions, propulsion on CubeSats has been used only for attitude control and reaction wheel desaturation via cold gas propulsion systems. As the desired capability of CubeSats increases, and more complex missions are planned, propulsion is required to accomplish the science and engineering objectives. This survey includes propulsion systems that have been designed specifically for the CubeSat platform and systems that fit within CubeSat constraints but were developed for other platforms. Throughout the survey, discussion of flight heritage and results of the mission are included where publicly released information and data have been made available. Major categories of propulsion systems that are in this survey are solar sails, cold gas propulsion, electric propulsion, and chemical propulsion systems. Only systems that have been tested in a laboratory or with some flight history are included.
This paper provides an update of the Delfi nanosatellite programme of the Delft University of Technology (TU Delft), with a focus on the recent in-orbit results of the second TU Delft satellite Delfi-n3Xt. In addition to the educational objective that has been reached with more than 80 students involved in the project, most of the technological objectives of Delfi-n3Xt have also been fulfilled with successful in-orbit demonstrations of payloads and platform. Among these demonstrations, four are highlighted in this paper, including a solid cool gas micropropulsion system, a new type of solar cell, a more robust Command and Data Handling Subsystem (CDHS), and a highly integrated Attitude Determination and Control Subsystem (ADCS) that performs three-axis active control using reaction wheels. Through the development of Delfi-n3Xt, significant experiences and lessons have been learned, which motivated a further step towards DelFFi, the third Delfi CubeSat mission, to demonstrate autonomous formation flying using two CubeSats named Delta and Phi. A brief update of the DelFFi mission is also provided.
A parametric, three-dimensional, computational study examining steady state performance of a planar plug micronozzle is conducted. Nozzle depth and throat Reynolds number are varied for a single plug contour designed using an approximation for the Method of Characteristics. The Reynolds number is varied from 80 − 820, for nozzle depths of 90μm-360μm. Thrust and specific impulse are used as metrics to assess nozzle performance. Results from 3D simulation are compared to 2D simulation for the same planar geometry and 3D results for a 30-degree linear-walled nozzle of the same throat size. For all Reynolds numbers examined, 3D nozzles were found to under-perform the 2D case. The margin between the 3D and 2D cases decreases with increasing nozzle depth. The 3D plug nozzle outperforms the 3D 30-degree linear-walled nozzle in both 2D and 3D simulation for the range of Reynolds numbers examined.
The Direct Simulation Monte Carlo (DSMC) method is used to numerically simulate and design a micronozzle with improved performance. Thrust calculations using the DSMC method demonstrate that the coaxial micronozzles can achieve milli-Newton thrust levels with specific impulses on the order of 45 s using argon in a cold gas expansion. Improved micronozzle designs of coaxial microthrusters are also proposed. Coaxial micronozzles utilizing center-body geometries to exploit pressure thrust show about 140% increase in specific impulse at low Reynolds numbers compared to a traditional converging nozzle.
A novel cold gas propulsion system for small satellites has been developed at the University of Texas at Austin’s Satellite Design Lab. The concept behind the propulsion system is to additively manufacture the main thruster module by means of the stereolithography process. In this manner, intricate features can be created and complex volumes can be used. This method is especially useful for small satellites. The propulsion system operates by releasing a saturated liquid propellant serially through three valves and a built-in converging-diverging nozzle. Through extensive tests, the propulsion system was measured with a specific impulse that ranged from 65 seconds at 24C to over 100 seconds at 85C. The measured thrust force provided by the propulsion system ranged from 110 mN at 24C to over 170 mN at 85C. A technology demonstration unit has been developed for flight onboard the University of Texas at Austin’s Bevo-2 satellite. This system will have a total mass under 400 grams, including 90 grams of Dupont 236-fa as propellant. The propulsion system is expected to provide at least 10 m/s of delta-v capability which has been verified through testing.
This article describes a new approach to fabricate various embedded ceramic microstructures. Design geometries of the microstructures are replicated onto a soft reusable polydimethylsiloxane (PDMS) mold. Rapid consolidation of ceramic suspension on the molds using gel-casting technique produces structured ceramic layers. 85% aqueous glycerol solution has been identified as suitable adhesive agent to laminate the ceramic layers. Capability of the approach in fabricating embedded microstructures of 3-dimensional and aspect ratio as high as 20 and 50 has been demonstrated. Ceramic microthruster and Y-shaped microfluidic reactor are fabricated using this new approach to show its potential to construct various ceramic microdevices.
Development of MEMS-based (Micro Electro Mechanical System) components
and subsystems for space applications has been pursued by various
research groups and organizations around the world for at least a
decade. The main driver for developing MEMS-based components for space
is the achievable miniaturization. MEMS technology can not only save
orders of magnitude in mass and volume of individual components, it can
also allow increased redundancy, and enable novel spacecraft designs and
mission scenarios. This paper presents a number of miniaturized
components, their development status and their planned maiden
spaceflight onboard the PRISMA satellite. One of the two PRISMA
satellites will have a cold gas propulsion system onboard including a
number of miniaturized MEMS-based components. NanoSpace has developed
and manufactured several of the critical components using MEMS
technology, i.e. the isolation valve, the pressure relief valve, the
thrust chamber/nozzle assemblies, the proportional valves, and the
filters. Presens has developed the MEMS-based pressure sensor
technology.
A comprehensive numerical investigation of a steady viscous How through a two-dimensional supersonic linear micronozzle has been performed. The baseline model for the study is derived from the NASA Goddard Space Flight Center microelectromechanical systems-based hydrogen peroxide prototype microthruster. On the microscale, substantial viscous subsonic layers may form on the nozzle expander walls, which reduce thrust and efficiency. One approach to compensate for the presence of these layers has been to design micronozzles with expansion angles larger than traditional macroscale nozzles. Numerical simulations have been conducted for a range of Reynolds numbers (Re = 15-800) and for expander half-angles of 10-50 deg. Two different monopropellant fuels have been considered: decomposed 85% pure hydrogen peroxide and decomposed hydrazine. It is found that an inherent tradeoff exists between the viscous losses and the losses resulting from the nonaxial exit How at larger expansion angles. Our simulations indicate that the maximum nozzle efficiency occurs for both monopropellants at a nozzle expansion half-angle of approximately 30 deg, which is significantly larger than that of traditional conical nozzle designs.
The need for low thrust propulsion systems for maneuvers on micro-and nano-spacecraft is growing. Low thrust characteristics generally lead to low Reynolds number flows from propulsive devices that utilize nozzle expansions. Low Reynolds number flows of helium and nitrogen through a small conical nozzle and a thin-walled orifice have been investigated both numerically, using the Direct Simulation Monte Carlo technique, and experimentally, using a nano-Newton thrust stand. For throat Reynolds number less than 100, the nozzle to orifice thrust ratio is less than unity; however, the corresponding ratio of specific impulse remains greater than one for the Reynolds number range from 0.02 to 200.
It has frequently been proposed to use very small nanosatellites for missions requiring orbital agility. Whether it be swarms of satellites for scientific or remote-sensing measurements, constellations for communications or single satellites for remote inspection, all require some way of modifying their respective orbits. Novel, high-technology solutions to this requirement have been proposed from MEMS to solar sails. Notwithstanding the eventual availability of such advanced nanosatellite propulsion technologies, the Surrey Space Centre has developed a miniature propulsion subsystem using technology readily available today. On 28 th June 2000 Surrey launched SNAP-1, the first in a series of Surrey Nanosatellite Application Platform missions. Amongst other features of this new 6.5 kg nanosatellite is a butane liquefied gas propulsion subsystem to meet the spacecraft's mission requirement of 1 m/s delta V. With a total mass budget of 450 grams, including propellant, dry mass, structural support and drive electronics, this propulsion system will be one of the smallest ever to have flown on a spacecraft. This paper describes some of the interesting challenges in producing such a small system, especially in a seven month "concept to launch site" program. The flight propulsion system will be described, including novel techniques such as using a coiled tube in the place of a conventional propellant tank. The choice of butane as a propellant will be discussed.
The second Canadian Advanced Nanospace eXperiment (CanX) satellite, CanX-2, aims to support Canadian researchers while expanding the capabilities of nanosatellites. Designed and built at the University of Toronto Institute for Aerospace Studies’ Space Flight Laboratory (UTIAS/SFL), CanX-2 will include experiments in GPS technologies, earth observation, advanced materials, and space communications protocols. In addition to the science payloads, CanX-2 will also fly engineering payloads such as a momentum-bias attitude control system, an experimental S-band communications system, a custom on-board computer, and a miniature propulsion system. With such an ambitious science platform, CanX-2 hopes to demonstrate the use of a nanosatellite as a valuable scientific tool that is cost- and schedule-effective for today's researchers. With a target launch in late 2005 into a highly inclined orbit, the experiments and satellite subsystems described in this paper will help pave the way for future nanosatellite science missions both at UTIAS/SFL and other institutions.
The PRISMA project for autonomous formation flying and rendezvous has passed its critical design review in February–March 2007. The project comprises two satellites which are an in-orbit testbed for Guidance, Navigation and Control (GNC) algorithms and sensors for advanced formation flying and rendezvous. Several experiments involving GNC algorithms, sensors and thrusters will be performed during a 10 month mission with launch planned for the second half of 2009.The project is run by the Swedish Space Corporation (SSC) in close cooperation with the German Aerospace Center (DLR), the French Space Agency (CNES) and the Technical University of Denmark (DTU). Additionally, the project also will demonstrate flight worthiness of two novel motor technologies: one that uses environmentally clean and non-hazardous propellant, and one that consists of a microthruster system based on MEMS technology.The project will demonstrate autonomous formation flying and rendezvous based on several sensors—GPS, RF-based and vision based—with different objectives and in different combinations. The GPS-based onboard navigation system, contributed by DLR, offers relative orbit information in real-time in decimetre range. The RF-based navigation instrument intended for DARWIN, under CNES development, will be tested for the first time on PRISMA, both for instrument performance, but also in closed loop as main sensor for formation flying. Several rendezvous and proximity manoeuvre experiments will be demonstrated using only vision based sensor information coming from the modified star camera provided by DTU. Semi-autonomous operations ranging from 200 km to 1 m separation between the satellites will be demonstrated.With the project now in the verification phase particular attention is given to the specific formation flying and rendezvous functionality on instrument, GNC-software and system level.
The Space Flight Laboratory (SFL) at the University of Toronto Institute for Aerospace Studies (UTIAS) is developing its second nanosatellite, the Canadian Advanced Nanospace eXperiment 2 (CanX-2) as a part of the CanX program. The objective of the CanX program is to produce highly capable nanospacecraft, each within a 2-year period, i.e. the time it takes to complete a graduate degree. CanX missions offer extremely low cost and rapid access to space for scientists and commercial exploitation. CanX-2, which is compatible with the Stanford University and California Polytechnic State University “CubeSat” standard, is a “triple CubeSat” with dimensions measuring and a mass of 3.5 kg. This nanosatellite features an ambitious suite of scientific and engineering payloads, and enlarges the envelope of capability for this class of spacecraft. The primary mission of CanX-2 is to test and demonstrate several enabling technologies for precise formation flying. The secondary purpose of CanX-2 is to fly a number of university experiments. The spacecraft mission, the engineering and scientific payloads, and their implication to planned SFL nanosatellite formation flight missions form the basis for this paper. CanX-2 is currently well within the 2-year design period and fast approaching flight-ready status with launch scheduled for mid 2006.
Development of MEMS-based (micro electro mechanical system) components and subsystems for space applications has been pursued by various research groups and organizations around the world for at least two decades. The main driver for developing MEMS-based components for space is the miniaturization that can be achieved. Miniaturization can not only save orders of magnitude in mass and volume of individual components, but it can also allow increased redundancy, and enable novel spacecraft designs and mission scenarios. However, the commercial breakthrough of MEMS has not occurred within the space business as it has within other branches such as the IT/telecom or automotive industries, or as it has in biotech or life science applications.
A novel micronozzle geometry with different depths at converging and diverging section was designed with the aim in mitigating viscous loss due to formation of subsonic viscous boundary layers at nozzle sidewalls. Feasible fabrication schemes based on available fabrication techniques, deep reactive ion etching (DRIE) and low temperature co-fired ceramic (LTCC) tapes, were proposed to prove practicality of the design. Numerical simulations were performed in evaluating the parameters of micronozzle performance, thrust, mass flow rate and specific impulse efficiency, for both 3D linear and two-depth micronozzles with 15° and 30° expander half-angles over a range of throat Reynolds numbers (Rethroat≈60–1068). Pressurized nitrogen gas was selected as propellant in this study. Comparison of numerical results shows two-depth micronozzles outperform linear micronozzles for the entire range of Rethroat and all performance parameters. The new micronozzle design has successfully reduced viscous loss through shorter expander length and mitigation of subsonic layers merging at nozzle expander section. Although the asymmetric geometry was found inducing a Z-axis thrust component, it can be offset through proper arrangement of the micronozzles. In conclusion, the design demonstrates improved performance over conventional micropropulsion system utilizing linear micronozzles by approximately 5%.
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