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A hybrid cold gas microthruster system for spacecraft
Sensors and Actuators A 97±98 (2002) 587±598
A hybrid cold gas microthruster system for spacecraft Johan KoÈhler*, Johan Bejhed, Henrik Kratz, Fredrik Bruhn, Ulf Lindberg, Klas Hjort, Lars Stenmark The AÊngstroÈm Laboratory, The AÊngstroÈm Space Technology Centre, Uppsala University, Box 534, Uppsala, Sweden Received 25 June 2001; accepted 25 September 2001
Abstract A hybrid cold gas microthruster system suitable for low Dv applications on spacecraft have been developed. Microelectromechanical system (MEMS) components together with ®ne-mechanics form the microthruster units, intergrating four independent thrusters. These are designed to deliver maximum thrusts in the range of 0.1±10 mN. The system includes three different micromachined subsystems: a nozzle unit comprising four nozzles generating supersonic gas velocity, i.e. 455 m/s, four independent piezoelectric proportional valves with leak rates at 10 6 scc/s He, and two particle ®lters. The performances of all these MEMS subsystems have been evaluated. The total system performance has been estimated in two parameters, the system-speci®c impulse and the mass ratio of the propulsion system to the spacecraft mass. These ®gures provide input for spacecraft design and manufacture. # 2002 Elsevier Science B.V. All rights reserved. Keywords: Cold gas micropropulsion; Micronozzle; Piezoelectric valve
1. Introduction The development of microthruster systems, reaching for increasingly smaller thrust levels, strongly depends on the ef®cient utilisation of microelectromechanical system (MEMS) technology [1]. Several concepts have been discussed in [2±7] where main topics include the type of propulsion used and the level of integration aimed at. Here, we investigate a hybrid integration solution, i.e. connecting and mounting of different MEMS units by conventionally machined housing parts. Furthermore, the MEMS units are served by external electronics. Thus, the complete microthruster system is assembled as shown in Fig. 2. The system can be designed for maximum thrust levels from 0.1 to 10 mN. Such thrusts can be used for main propulsion of short-term nanoprobe missions, or attitude control and drag compensation on larger spacecraft. This cold gas microthruster system releases minute amounts of gas at supersonic speed through micronozzles, thus, providing precise momentum to the spacecraft. Piezoelectric valvesÐusing feedback from differential pressure sensors straddling each nozzleÐproportionally modulate
* Corresponding author. Tel.: 46-1847-17253; fax: 46-1855-5095. E-mail address: [email protected] (J. KoÈhler).
the gas ¯ow to the nozzles. Heat exchangers with thin ®lm heaters and temperature sensors in the nozzle unit increase the ef®ciency of the thruster. Micromachined ®lters protect these MEMS units from degrading particle contamination due to incoming gas and surrounding space. 2. Cold gas micropropulsion In cold gas micropropulsion, the force to the system is delivered by ejecting mass (i.e. gas) at high speeds. The force is determined by Newton's second law of motion under the assumption of constant exit velocity of the gas, thus, F Q m ve
(1)
Here, Qm is the mass ¯ow and ve is the exit velocity of the matter. Thus, the mass of the propellant is utilised to its maximum by optimising the exit velocity. In this context, a converging±diverging nozzle is suitable, named a Laval nozzle (Fig. 1). Subsonic gas entering a converging zone increases its velocity, while supersonic gas increases its velocity in a diverging zone. Thus, if the gas reach sonic speed at the nozzle throat, it will continue its velocity increase to supersonic values beyond the throat in the nozzle expansion area.
0924-4247/02/$ ± see front matter # 2002 Elsevier Science B.V. All rights reserved. PII: S 0 9 2 4 - 4 2 4 7 ( 0 1 ) 0 0 8 0 5 - 6
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Fig. 1. The basic concept of a Laval nozzle.
For an optimal converging±diverging nozzle, the exit velocity reaches maximum exactly at the nozzle outlet, and is given by v ( u k 1=k ) u2kRm T0 pout (2) 1 ve t k 1 pin
of propellant. This value is obtained by dividing the speci®c impulse by the standard gravitational acceleration g (9.81 m/s2).
where k is the ratio of speci®c heat capacities (cp /cv, for N2: k 1:4), Rm the gas constant with respect to gas mass (for N2: Rm 297 N m/kg K), T0 the inlet gas temperature, pout the pressure at the nozzle outlet and pin is the inlet pressure. In vacuum, the surrounding pressure does not determine the actual pressure ratio pout/pin, but rather the area expansion of the nozzle sets the ratio, according to s 1= k 1 r 2=k k1=k Athroat k1 k1 pout pout 2 k 1 Aexit pin pin
A hybrid microsystem solution comprises vital microsystem parts supported by innovative miniature ®xtures and connections made by conventional ®ne-mechanics methods. Hybrid solutions of this kind typically involves lower development risks, while the extreme miniaturisation possible in microsystems proper cannot be fully exploited. The possibilities of incorporating redundancy in the system are also more limited. The hybrid cold gas microthruster system presented here exhibits highly integrated micromechanical devices together with conventional components. These form a miniaturised high performance micropropulsion system intended for spacecraft attitude control, extreme stabilisation, noise reduction, and drag compensation. The successful miniaturisation of the system basically leads to the possibility to develop smaller fully capable spacecraft and providing better stability and precision to larger ones as well. Thus, the system actually enables several new concepts in space exploration, requiring features like formation ¯ying of spacecraft clusters, extreme stabilisation for ultra-sensitive measurements on earth phenomena, or excellent pointing accuracy for long distance observations. The micromachined subsystems are described in the Section 4. Before this, the mechanical and electronic designs are outlined.
(3) with de®nitions from Fig. 1. A numerical solution for different throat sections in the present design (see the following sections) is presented in Table 1, together with predicted exit velocities. Given these calculations, the ef®ciency of an experimentally characterised nozzle can be expressed as the ratio of the experimental and ideal exit velocities, respectively. A key ®gure used in determining the performance of a thruster is the speci®c impulse (Isp), which is numerically equivalent to the exit velocity, using the unit N s/kg. This is the impulse delivered from each mass unit of propellant. Often, speci®c impulse is quoted in the unit of seconds, corresponding to the impulse delivered per weight unit Table 1 Specific pressure ratios and ideal exit velocity for nozzles with different throat sections in the current design Nozzle (mm)
Pressure ratio
Predicted v (m/s)
10 30 40 15 35 45
2.31E 7.61E 1.12E 3.82E 9.59E 1.34E
716 685 671 704 677 664
100 100 100 94 100 100
03 03 02 03 03 02
3. Hybrid microsystem solution
3.1. Mechanical design The mechanical design of the microthruster system pod is quite conventionally made by ®ne-mechanics machining. An overview is given in the following, referring to Fig. 2. The pressurised housing consists of three major parts. First, a cylindrical bottom part contains the gas inlet and the ®lter stack. Second, the bottom surface with four threaded holes and two guide pins provides the mechanical interface to a mounting bracket. Third, the cylindrical centre part contains
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Fig. 2. The hybrid cold gas microthruster system, containing four independent thrusters. (A) Overview showing the gross dimensions of the system. (B) Cross-section showing the different parts, e.g. the microsystems: nozzle unit, valve assemblies, and particle filters. (C) Gas outlet in ground testing (left) and in operation (right).
the valve assemblies and the hermetic interface connector. The connector size is the design driver regarding dimensions. The nozzle wafer is located between the centre part and the half-spherical top cap. Inside the top is a mounting block
where four differential pressure sensors are located. Miniaturised V-bands are used to join all pressurised parts. The connector I/F section serves double needs, providing an interface between the external connectors and the connector I/F board, while comprising the base connecting to
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the spacecraft power supply. This housing is not designed to be pressurised, but it will be dust tight as the connector backshells cover the only holes to the surroundings. The ®lter section is the ®rst section of the pressure-tight part in the thruster pod. It contains the gas input and a large micromachined silicon ®lter. This ®lter is suspended between the ®lter section and the intersection ring. Four O-rings are used to keep the silicon part ¯oating between the two aluminium sections for sealing and in order to avoid mechanical stress due to different thermal expansion coef®cients. The intersection ring provides space for electric components and wire connections to the valve assemblies. The valve section contains the four proportional valve assemblies and an interface board with preampli®ers for the strain gauge sensors. Each valve is accommodated in individual cavities with O-ring sealings in the bottom and a small hole in the centre for the gas transport to the nozzle unit. This, in turn, is mounted on the topside essentially in the same way as the ®lter stack. The valve is pushed against the O-ring in the cavity bottom by a small bracket. This bracket also pushes a second O-ring against the valve cap wafer. Thus, the silicon device is brought to ¯oat between Orings at the aluminium surface, as in the ®lter assembly, thereby avoiding thermal stresses. The pressure sensor section is mounted on top of the nozzle unit. Five miniature pressure sensors are mounted in this section. These are four differential sensors with a maximum pressure difference of 7 barsÐone for each nozzleÐand one absolute sensor in order to monitor the ambient pressure. A PC-board with ®ve preampli®ers is also mounted in this section. A thin-walled cover on the thruster pod protects the internal harness and keeps contaminating particles outside. A small protection ®lter of the same type as the large inlet ®lter permits the unit to ``breath'' and the differential
pressure sensors to refer to the true ambient pressure. The protection ®lter is mounted in the top by use of a locking ring and O-rings similar to the other silicon parts. The protective cap is used on ground to protect the nozzles form dust contamination. An O-ring at the bottom seals the cap when mounted in position. It is absolutely necessary to remove this cap before launch. For that reason it is clearly marked with both an inscription and an eye-catching colour. Finite element analyses have been performed on the pressurised parts of the housing at 32 bars to verify that the mechanical stress is acceptable. The outer coating to be used is a functional surface, exhibiting different optical properties for different wavelengths. The present most promising ®lm should maintain an equilibrium temperature in direct solar irradiation below 50 8C without signi®cant internal power dissipation [8]. Besides assisting in the thermal control, this coating is mechanically extremely durable, chemically inert, and electrically conductive which prevents electrical charging in space. The complete microthruster unit, without the external electronics and propellant storage, weighs approximately 0.15 kg. 3.2. Electronics The electronic system integration of the hybrid microthruster is divided in two conventional systems, a digital and an analogue part with a unique interconnection. The complete system begins with the design of the nozzle positions in perpendicular pairs, shown in Fig. 3 as X and Y. This approach has several advantages, the ®rst and most obvious is redundancy and the second is to enhance the steering and ®ne tuning of the nozzle functions. In this hybrid microthruster design, the electronics are contained outside of the
Fig. 3. Hybrid cold gas microthruster layout.
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Fig. 4. Electronics block diagram.
mechanical envelope presented in the earlier sections, with an estimated mass of 0.2 kg. However, all electronics will eventually be implemented in multichip modules (MCM) [9] and can, therefore, be made redundant in a three- or fourway system without much loss of space or gain of mass. The actuators and sensors in the pod are monitored and controlled by regular analogue circuitry allowing the thrusters to be run differentially for maximum performance. The analogue control electronics have an advanced control system in need of reference inputs. These inputs are set by an embedded microcontroller. Fig. 4 shows the system integration starting from the right with the thruster pairs. Analogue signals are carried in closed control loops to and from the thrusters. The reference points are used within these closed control loops, which are connected to the microcontroller through a number of digital±analogue and analogue±digital converters (ADC/DAC). The new approach is to create a system based on conventional control engineering while increasing its ¯exibility by introducing a microcontroller as a smart device sending input to the control loops. There are further advantages by introducing microcontrollers. First, they are very power saving. A normal 8 bit modern microcontroller, only a few square millimetres in die form, consumes less than 0.13 W in full operation. Second, they include advanced features like controller area network (CAN) and serial peripheral interface (SPI).
The microcontroller in this application connects each thruster to the CAN bus which provides an easy and modular structure of the microthruster system concept. Furthermore, several ADC/DAC are hooked on to the microcontroller through the Motorola SPI. By using widely accepted electronics solutions a large range of products are available for use within the system. Furthermore, most of the SPI chips are compatible, meaning that individual chips may be changed for different applications and requirements without having to redo the complete electronics layout. The CAN is also increasingly used within the space community and this provides a solid and simple base for including this thruster in a space project. The user only needs to specify the CAN address of the individual thrusters in order for them to be fully operational on the CAN bus. The use of both conventional analogue control electronics and new digital interfaces gives the possibility to maximise the abilities of electronics engineers with high analogue skills and modern computer programmers with high digital skills. 4. Microsystem manufacture The primary material for the different microsystem units of the hybrid cold gas microthruster system is single crystalline silicon, shaped by anisotropic wet etching and/or deep
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reactive ion etching (DRIE). Direct bonding joins structured silicon wafers in order to eventually forming the complete units. Evaporation and patterning of different metals, i.e. tin±lead alloy, tin±silver alloy, titanium, copper, gold, and chromiumÐprovides conductors and solder-pads for electronics, sensors, and actuators. Other thin ®lm techniques, e.g. low pressure chemical vapour deposition (LPCVD) give wear resistant coatings on particularly exposed details. 4.1. Nozzle unit The nozzle unit comprises four independent nozzles, divided by 908 angles on a 40 mm diameter disc. Heat exchangers are integrated directly upstream of each nozzle,
and platinum thin ®lm heater elements combined with temperature sensors are included on the outside surfaces of all heat exchanger structures. The nozzle shape is essentially a two-dimensional converging±diverging outlet with rectangular cross-section (Fig. 5). Bayt et al. [4] have demonstrated slightly larger nozzles of the same type, made from silicon and glass. In the present case, the inlet angle is 27±318 and the outlet 158, and the throat widths are 10, 30, or 40 mm. Each nozzle furthermore comprises two identical halves bonded together. The nominal height of two such halves is 100 mm. The actual shape of each manufactured nozzle is carefully monitored using white-light interferometric pro®lometry. The aligned bonding allows for smaller nozzle heights and homogenous
Fig. 5. The converging±diverging nozzle half. The throat width in this case is 10 mm. The close-up on the throat section readily shows the straight walls obtained.
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Fig. 6. The heat exchanger structure included directly upstream of each nozzle. This half consists of 3 mm thin curved fins, 40 mm high, and spaced 7 mm apart.
materials. The layout of the nozzle shape has not yet been optimised. The nozzles are free-hanging tubular structures in order to minimise heat loss to the supporting silicon. Into this tube, the heat exchanger structure is included, comprising a vast array of narrow curved ®ns, 3 mm thin, spaced 7 mm apart (Fig. 6). In order to address the evaluation of single nozzles correctly, units without the heat exchanger have been manufactured as well. Here, the throat widths are 15, 35, and 45 mm, respectively. 4.2. Valve unit This valve unit is a normally-closed proportional piezoelectrically actuated device. Five multilayered piezoelectric lead±zirconium±titanate (PZT) elements perform the valve
action, lifting a central silicon cap from the coated silicon valve seat (Fig. 7). The valve seat is 2.8 mm square with a 200 mm wide seal. The multilayered design limits the required drive voltage to 50 V for a maximum stroke of 4 mm. The total actuator height is 4 mm. The folded piezoactuator set-up has two major advantages. First, problems with the coef®cient of thermal expansion or ageing in the PZT elements are eliminated. Second, if both the inner and outer elements are operated while opening and closing, possible reaction disturbances can be avoided. The actuators are assembled using a thin ®lm soldering system. Here, the bottom platinum layer of each actuator is soldered to a chromium±copper±gold multilayer pad, using lead-tin solder. The top PZT layer of the actuator is deposited with titanium±copper±gold and soldered to chromium± copper±gold pads on the silicon yoke using tin±silver solder (Fig. 8). These techniques allow for precision mounting of
Fig. 7. The valve unit. Right: overview, showing the basic assembly of the complete valve. Left: IR image of an assembled valve seal, showing the seat and cap parts with electrodes and soldering pads after wafer bonding.
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Fig. 8. The actuator mounting and connection. Right: the five PZT elements soldered in place. Left: thin film conductors connect the electrodes of the multilayered actuator.
the actuators on the silicon parts, which is necessary for the successful operation of the valve. The internal electrodes of the multilayer structure are connected using evaporation of gold through shadow masks onto the sides of the actuators (Fig. 8). Wear resistant coatings cover the valve seat in order to avoid rapid degradation of the sealing ability. Parylene and titanium±aluminium nitride have been tested and compared to bare silicon valve seats. 4.3. Filter unit The ®lter unit is realised from combining two different types of ®lters, forming a three-wafer stack. The active ®lter parts are located in the bond interfaces (Fig. 9).
The ®rst ®lter part consists of approximately 50 mm wide remaining ridges on a slightly depressed surface (Fig. 9). These ridges are placed near the gas inlet. This ®lter type is single-sided, and bonded with a wafer containing outlet holes. The fenced slot is the active ®lter part. The second ®lter part, crossed v-grooves, consists of narrow v-grooves in the wafer, together with inlet holes (Fig. 9). By bonding this wafer to a similarly patterned wafer, the crossed v-groove ®lter is obtained, forming a 908 angle between v-grooves. The narrow v-grooves are the active ®lter part. To complete the ®lter stack, the slot ®lter is bonded to the inlet side of the v-groove ®lter. The ridge height in the slot ®lter was measured to 1.5± 4 mm by a stylus pro®lometer, and the v-grooves were fully etched from a mask width of 1.5±2 mm.
Fig. 9. The filter unit. Top left: 40 mm diameter filter disc after bonding (IR image). Bottom left: filter cross-section, the upper bond interface contains the vgroove part and the lower interface contains the slot part. Top right: the v-groove filter part (close-up). Bottom right: the slot filter part (close-up).
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Fig. 10. Mass flow and thrust force as a function of inlet pressure for nozzle with throat dimension 35 mm 100 mm, without the heat exchanger structure.
5. Results and discussion The performance of single nozzle elements, without the heat exchanger structure, have been evaluated by measuring the mass ¯ow and thrust force as a function of inlet pressure (Fig. 10) [10]. These data are derived from pressure drop measurements from ®xed volumes, releasing the gas into vacuum, while monitoring the impact of thrust by placing the thruster on top of a mg-resolution scale. By correlating the thrust force obtained to the current mass ¯ow, the exit velocity can be calculated (Eq. (1)), and thus, the speci®c impulse of the thruster. The nozzle with 35 mm 100 mm throat section reaches an exit velocity of 445 m/s, corresponding to a speci®c impulse of 45 s. This is 66% ef®ciency according to the present area expansion, considering the ideal exit velocity from Table 1. Obviously, there is much room for design improvements, changing, e.g. the area expansion, nozzle length, and rotational symmetry of these nozzles. Bayt and Breuer quoted 460±590 m/s for their nozzles at high pressures, while viscous losses degrade performance at lower pressures (i.e. lower thrusts) [11]. Such behaviour is also con®rmed by our experiments [10]. The mass ¯ows of the more elaborate nozzles in the nozzle unit are expected to be restricted, due to the heat exchanger, in direct comparison to the results from these single nozzles. Thus, the throat dimensions may need adjusting in this design. However, the operation of the heater is expected to increase the speci®c impulse. MEMS valves for space applications generally exhibit poor leakage characteristics. Micropropulsion missions are
inherently more sensitive to leakage than larger sized spacecraft. In order to maintain lifetime and performance when shrinking the system to micropropulsion, the valves need leakage rates below 10 5 scc/s of helium [12]. These valve parts have been tested for 30,000 work cycles with parylene or TiAlN coating on the valve seat. The leak rates as a function of sealing force have been measured using a leak detector (Fig. 11). For the coated samples, the leak rate remains suf®ciently low. The incoming gas needs ®ltration, so the mass ¯ow capacity of the ®lters must be tested in order to correctly dimension the ®lter necessary for the mission. The mass ¯ows of different gases through an 18 mm diameter ®lter unit to atmospheric pressure have been monitored at different inlet pressures (Table 2). The performance of a propulsion system can be described using the system-speci®c impulse [13], de®ned as the total Table 2 Mass flow of the filter for different inlet pressures of different gases Gas
Pressure (MPa)
Mass flow (10
N2
0.2 0.4
0.60 3.1
He
0.2 0.4
0.12 0.54
Ar
0.2 0.4
0.69 3.6
Xe
0.2 0.4
2.0 10.1
6
kg/s)
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Fig. 11. Leak rate as a function of sealing force for pristine and used valve seats with different coatings.
impulse (Itot) delivered by the system divided by the system mass (mPS), thus, Issp
Itot mPS
(4)
Here, the mass of the propulsion system can be expressed in terms of the individual masses of thruster units (mt) control electronics (me), propellant (mp), and storage tank (mST). Using four microthruster systems for full three-axis stabilisation, the propulsion system mass become mPS 4mt 4me mP mST
(5)
The mass of propellant needed is determined from the mission requirement and the spacecraft total mass. Using conservation of momentum, the velocity change Dv obtained by the spacecraft ejecting propellant is described by mF Dv ve ln (6) m0 where m0 is the initial and mF the ®nal mass of the spacecraft, respectively. Thus, the total amount of propellant needed to deliver the total impulse in Eq. (4) for a certain mission-speci®c Dv requirement is described by mP m0 1
e
Dv=ve
(7)
Using this, the total impulse delivered by the system is readily obtained as Itot mP ve
(8)
Different propulsion assignments on speci®c spacecraft missions require certain Dv manoeuvres, or their equivalent. For example, a 5 years ®ne-pointing earth-observing cluster mission typically requires 5±50 m/s Dv equivalent for attitude control, while the deep space interferometer mission (0.5±1 years duration) needs 100±300 m/s [14]. Now, observing that the propellant mass and storage system mass depend on the total amount of propellant
needed for speci®c Dv requirements and, thus, on initial spacecraft mass, the system-speci®c impulse can be estimated. Furthermore, the required mass fraction of the propulsion system to the total spacecraft mass [13] can be expressed by combining Eqs. (4), (5), (7) and (8) into Isp mPS 1 e Dv=ve (9) ms=c Issp Here, the total spacecraft mass ms/c is identical to the initial spacecraft mass m0. Assuming a storage pressure of nitrogen at 24 MPa in a spherical titanium tank with wall thickness 4 mm, a speci®c impulse of 45 s, and mass estimates from the above descriptions, the system-speci®c impulse and mass fraction of the propulsion system have been estimated (Fig. 12). The amount of propellant have further been adjusted by a leakage assumption based on a He leak rate of 1 10 6 scc/s, which translates to 3:8 10 7 scc/s for N2 if the leak rate is considered to be inversely proportional to the square root of the molecular mass [14]. A 5 years mission, thus, entails a total amount of leaked nitrogen of 960 scc, i.e. 1:2 10 3 kg. These ®gures can be much improved by choosing a light-weight composite storage tank. Otherwise, the control electronics total weight of 0.8 kg may be eliminated by the successful development of MCMs. Furthermore, the expected increase of the speci®c impulse when operating the heater, and the possible gain of ef®ciency by better nozzle design also promises better performances. For cold gas propulsion systems, a mass fraction between 0.1 and 0.3 is recommended [1]. Now, the graphs in Fig. 12 can be used for comparing the merits of different propulsion systems, and estimate the total mass of the spacecraft. Inspecting Fig. 12, it is evident that the hybrid cold gas microthruster system is best suited for low Dv assignments, like attitude control, ®ne stabilisation, drag compensation, or short-term free-¯ying missions.
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Fig. 12. The system-specific impulse predicted for the hybrid cold gas microthruster system (solid lines), together with the estimated mass fraction occupied by the propulsion system (dashed lines). The estimates are presented as a function of mission Dv requirement and total spacecraft mass.
6. Conclusion The hybrid cold gas microthruster system has been detailed, suitable for spacecraft attitude control, stabilisation, and short-term light-weight missions. This system enables exacting space missions where extreme accuracy is needed. The performance of the microthruster system has been demonstrated. Data from which to design mission-speci®c thruster systems, capable of maximum thrusts ranging from 0.1 to 10 mN have been obtained. The speci®c impulse reaches 45 s for nitrogen propellant, which translate to an ef®ciency of 66% using this current nozzle design. Four piezoelectric proportional valves independently modulate the gas ¯ow to the nozzles, using an analogue closed-loop arrangement. Thus, the achieved thrust is continuously controlled. The leak rate of a sealed valve is below 10 5 scc/s He, even after extensive wearing. Ef®cient particle ®lters have also been included. Acknowledgements The European Space Agency (ESA) is gratefully acknowledged for ®nancing this project, together with The Advanced Microengineering research programme at Uppsala University, funded by the Swedish Foundation for Strategic Research.
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[10] J. KoÈhler, Bringing silicon microsystems to space. Manufacture, performance, and reliability, Comprehensive Summaries of Uppsala Dissertations from the Faculty of Science and Technology, Vol. 678, Acta Universitatis Upsaliensis, Uppsala, 2001, 66 pp. ISBN 91-5545196-9. [11] R.L. Bayt, K.S. Breuer, Viscous effects in supersonic MEMSfabricated micronozzles, in Proceedings of the 1998 ASME International Mechanical Engineering Congress and Expo on Microelectromechanical Systems (MEMS), ASME, New York, USA, 1998, pp. 117±123. [12] J. Mueller, Review and applicability assessment of MEMS-based microvalve technology for microspacecraft propulsion, in: M.M. Micci, A.D. Ketsdever (Eds.), Micropropulsion for Small Spacecraft: AIAA Progress in Astronautics and Aeronautics, Vol. 187, 2000, pp. 449±476. [13] P. Erichsen, Performance evaluation of spacecraft propulsion systems in relation to mission impulse requirements, in: Proceedings of the 2nd European Spacecraft Propulsion Conference, 27±29 May 1997 (ESA SP-398, August 1997). [14] J. Mueller, Thruster Options for Microspacecraft: A Review of Existing Hardware and Emerging Technologies, AIAA Paper 973058, 1997, pp. 1±29.
Biographies Johan KoÈhler currently investigates bonded silicon microsystems for Ê ngstroÈm spaceflight applications, primarily micropropulsion, at the A Space Technology Centre, Uppsala University. He was born in 1971 and received his MSc in Materials Physics from Uppsala University in 1997. Since then his research has been in silicon micromachining, mainly fusion bonding. Johan Bejhed was born in 1973. He received his MSc in Materials Physics from Uppsala University in 2001. Since 1999, he has been a technician at Ê ngstroÈm Laboratory, Uppsala University, and he has recently become the A Ê ngstroÈm Space Technology Centre. His main topics a PhD student at the A are fluid-handling microsystems in space applications. Henrik Kratz was born in 1974. He received his MSc in Engineering Physics from Uppsala University in 1999. He is a PhD student at the Ê ngstroÈm Space Technology Centre. His current research involves RF± A MEMS for space applications and piezoelectric actuators integration into silicon microsystems.
Fredrik Bruhn was born in 1976 and received his MSc in Physics from Ê ngstroÈm Uppsala University in 2000. He is currently a PhD student at the A Space Technology Centre where his research is into system integration for microsystems. Ulf Lindberg was born in 1958. He received his MSc in Engineering Physics in 1985 and in 1993 he received his PhD in Solid-State Electronics, both from Uppsala University. After working at the Electrum Foundation 1994±1997 and Biacore AB 1997±2000, he joined the Ê ngstroÈm Space Technology Centre and the Solid-State Electronics A department at Uppsala University in 2000. His main research interests is microstructure technology for space and biotechnical applications. Klas Hjort received his MSc in Engineering Physics in 1988 and his PhD in Materials Science in 1993, both from Uppsala University. He spent a post-doctoral year at the Institute of High-Frequency Electronics, TU Darmstadt, Germany, in 1994. He made his ability degree in 1997 (Docent, Ê ngstroÈm Material Science). Since 1997, he is an associate professor at the A Laboratory, Department of Materials Science, Uppsala University. He has 60 reviewed international papers (eight invited). Since 1988, he has worked on basic mechanical characterisation, surface and bulk microstructuring, solid-state wafer bonding, device fabrication of bulk acoustic wave microstructure devices (e.g. based on GaAs/Al±GaAs), and quartz BAW resonators. His present research areas are INP-based optoelectronic MEMS, diamond microstructures, accelerator-based deep ion projection lithography, and microsystems for space applications. Lars Stenmark was born in 1944 and received his MSc in 1969 from Stockholm University, Sweden. He participated in the first sounding rocket campaigns at ESRANGE 1968. He worked as a development engineer and later project manager at the Swedish Space Corporation 1968±1979. In 1979, he started the private company Micro Nova KB for developing new electro-optical instruments. In 1982, he started ACR Electronic AB, a hardware developer and supplier entirely dedicated to space applications. Since then, ACR has been a major supplier to the Swedish national space programme (e.g. ODIN and FREJA: primary and secondary structures, solar sensors, power conditioning units, cryogenic boxes for sub-mm receivers, etc.). Lars Stenmark has designed, tested, and flown numerous experiments and subsystems, e.g. micro gravity payload modules for the TEXUS programme, MASER, and GAS (Get Away Special on the Shuttle). Since 1998, he has left all conventional space technology in order to promote the introduction of microsystems technology in space. He is Ê ngstroÈm Laboratory, A Ê ngstroÈm currently founder and director of the A Space Technology Centre, Uppsala University.
A hybrid cold gas microthruster system for spacecraft Johan KoÈhler*, Johan Bejhed, Henrik Kratz, Fredrik Bruhn, Ulf Lindberg, Klas Hjort, Lars Stenmark The AÊngstroÈm Laboratory, The AÊngstroÈm Space Technology Centre, Uppsala University, Box 534, Uppsala, Sweden Received 25 June 2001; accepted 25 September 2001
Abstract A hybrid cold gas microthruster system suitable for low Dv applications on spacecraft have been developed. Microelectromechanical system (MEMS) components together with ®ne-mechanics form the microthruster units, intergrating four independent thrusters. These are designed to deliver maximum thrusts in the range of 0.1±10 mN. The system includes three different micromachined subsystems: a nozzle unit comprising four nozzles generating supersonic gas velocity, i.e. 455 m/s, four independent piezoelectric proportional valves with leak rates at 10 6 scc/s He, and two particle ®lters. The performances of all these MEMS subsystems have been evaluated. The total system performance has been estimated in two parameters, the system-speci®c impulse and the mass ratio of the propulsion system to the spacecraft mass. These ®gures provide input for spacecraft design and manufacture. # 2002 Elsevier Science B.V. All rights reserved. Keywords: Cold gas micropropulsion; Micronozzle; Piezoelectric valve
1. Introduction The development of microthruster systems, reaching for increasingly smaller thrust levels, strongly depends on the ef®cient utilisation of microelectromechanical system (MEMS) technology [1]. Several concepts have been discussed in [2±7] where main topics include the type of propulsion used and the level of integration aimed at. Here, we investigate a hybrid integration solution, i.e. connecting and mounting of different MEMS units by conventionally machined housing parts. Furthermore, the MEMS units are served by external electronics. Thus, the complete microthruster system is assembled as shown in Fig. 2. The system can be designed for maximum thrust levels from 0.1 to 10 mN. Such thrusts can be used for main propulsion of short-term nanoprobe missions, or attitude control and drag compensation on larger spacecraft. This cold gas microthruster system releases minute amounts of gas at supersonic speed through micronozzles, thus, providing precise momentum to the spacecraft. Piezoelectric valvesÐusing feedback from differential pressure sensors straddling each nozzleÐproportionally modulate
* Corresponding author. Tel.: 46-1847-17253; fax: 46-1855-5095. E-mail address: [email protected] (J. KoÈhler).
the gas ¯ow to the nozzles. Heat exchangers with thin ®lm heaters and temperature sensors in the nozzle unit increase the ef®ciency of the thruster. Micromachined ®lters protect these MEMS units from degrading particle contamination due to incoming gas and surrounding space. 2. Cold gas micropropulsion In cold gas micropropulsion, the force to the system is delivered by ejecting mass (i.e. gas) at high speeds. The force is determined by Newton's second law of motion under the assumption of constant exit velocity of the gas, thus, F Q m ve
(1)
Here, Qm is the mass ¯ow and ve is the exit velocity of the matter. Thus, the mass of the propellant is utilised to its maximum by optimising the exit velocity. In this context, a converging±diverging nozzle is suitable, named a Laval nozzle (Fig. 1). Subsonic gas entering a converging zone increases its velocity, while supersonic gas increases its velocity in a diverging zone. Thus, if the gas reach sonic speed at the nozzle throat, it will continue its velocity increase to supersonic values beyond the throat in the nozzle expansion area.
0924-4247/02/$ ± see front matter # 2002 Elsevier Science B.V. All rights reserved. PII: S 0 9 2 4 - 4 2 4 7 ( 0 1 ) 0 0 8 0 5 - 6
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Fig. 1. The basic concept of a Laval nozzle.
For an optimal converging±diverging nozzle, the exit velocity reaches maximum exactly at the nozzle outlet, and is given by v ( u k 1=k ) u2kRm T0 pout (2) 1 ve t k 1 pin
of propellant. This value is obtained by dividing the speci®c impulse by the standard gravitational acceleration g (9.81 m/s2).
where k is the ratio of speci®c heat capacities (cp /cv, for N2: k 1:4), Rm the gas constant with respect to gas mass (for N2: Rm 297 N m/kg K), T0 the inlet gas temperature, pout the pressure at the nozzle outlet and pin is the inlet pressure. In vacuum, the surrounding pressure does not determine the actual pressure ratio pout/pin, but rather the area expansion of the nozzle sets the ratio, according to s 1= k 1 r 2=k k1=k Athroat k1 k1 pout pout 2 k 1 Aexit pin pin
A hybrid microsystem solution comprises vital microsystem parts supported by innovative miniature ®xtures and connections made by conventional ®ne-mechanics methods. Hybrid solutions of this kind typically involves lower development risks, while the extreme miniaturisation possible in microsystems proper cannot be fully exploited. The possibilities of incorporating redundancy in the system are also more limited. The hybrid cold gas microthruster system presented here exhibits highly integrated micromechanical devices together with conventional components. These form a miniaturised high performance micropropulsion system intended for spacecraft attitude control, extreme stabilisation, noise reduction, and drag compensation. The successful miniaturisation of the system basically leads to the possibility to develop smaller fully capable spacecraft and providing better stability and precision to larger ones as well. Thus, the system actually enables several new concepts in space exploration, requiring features like formation ¯ying of spacecraft clusters, extreme stabilisation for ultra-sensitive measurements on earth phenomena, or excellent pointing accuracy for long distance observations. The micromachined subsystems are described in the Section 4. Before this, the mechanical and electronic designs are outlined.
(3) with de®nitions from Fig. 1. A numerical solution for different throat sections in the present design (see the following sections) is presented in Table 1, together with predicted exit velocities. Given these calculations, the ef®ciency of an experimentally characterised nozzle can be expressed as the ratio of the experimental and ideal exit velocities, respectively. A key ®gure used in determining the performance of a thruster is the speci®c impulse (Isp), which is numerically equivalent to the exit velocity, using the unit N s/kg. This is the impulse delivered from each mass unit of propellant. Often, speci®c impulse is quoted in the unit of seconds, corresponding to the impulse delivered per weight unit Table 1 Specific pressure ratios and ideal exit velocity for nozzles with different throat sections in the current design Nozzle (mm)
Pressure ratio
Predicted v (m/s)
10 30 40 15 35 45
2.31E 7.61E 1.12E 3.82E 9.59E 1.34E
716 685 671 704 677 664
100 100 100 94 100 100
03 03 02 03 03 02
3. Hybrid microsystem solution
3.1. Mechanical design The mechanical design of the microthruster system pod is quite conventionally made by ®ne-mechanics machining. An overview is given in the following, referring to Fig. 2. The pressurised housing consists of three major parts. First, a cylindrical bottom part contains the gas inlet and the ®lter stack. Second, the bottom surface with four threaded holes and two guide pins provides the mechanical interface to a mounting bracket. Third, the cylindrical centre part contains
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Fig. 2. The hybrid cold gas microthruster system, containing four independent thrusters. (A) Overview showing the gross dimensions of the system. (B) Cross-section showing the different parts, e.g. the microsystems: nozzle unit, valve assemblies, and particle filters. (C) Gas outlet in ground testing (left) and in operation (right).
the valve assemblies and the hermetic interface connector. The connector size is the design driver regarding dimensions. The nozzle wafer is located between the centre part and the half-spherical top cap. Inside the top is a mounting block
where four differential pressure sensors are located. Miniaturised V-bands are used to join all pressurised parts. The connector I/F section serves double needs, providing an interface between the external connectors and the connector I/F board, while comprising the base connecting to
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the spacecraft power supply. This housing is not designed to be pressurised, but it will be dust tight as the connector backshells cover the only holes to the surroundings. The ®lter section is the ®rst section of the pressure-tight part in the thruster pod. It contains the gas input and a large micromachined silicon ®lter. This ®lter is suspended between the ®lter section and the intersection ring. Four O-rings are used to keep the silicon part ¯oating between the two aluminium sections for sealing and in order to avoid mechanical stress due to different thermal expansion coef®cients. The intersection ring provides space for electric components and wire connections to the valve assemblies. The valve section contains the four proportional valve assemblies and an interface board with preampli®ers for the strain gauge sensors. Each valve is accommodated in individual cavities with O-ring sealings in the bottom and a small hole in the centre for the gas transport to the nozzle unit. This, in turn, is mounted on the topside essentially in the same way as the ®lter stack. The valve is pushed against the O-ring in the cavity bottom by a small bracket. This bracket also pushes a second O-ring against the valve cap wafer. Thus, the silicon device is brought to ¯oat between Orings at the aluminium surface, as in the ®lter assembly, thereby avoiding thermal stresses. The pressure sensor section is mounted on top of the nozzle unit. Five miniature pressure sensors are mounted in this section. These are four differential sensors with a maximum pressure difference of 7 barsÐone for each nozzleÐand one absolute sensor in order to monitor the ambient pressure. A PC-board with ®ve preampli®ers is also mounted in this section. A thin-walled cover on the thruster pod protects the internal harness and keeps contaminating particles outside. A small protection ®lter of the same type as the large inlet ®lter permits the unit to ``breath'' and the differential
pressure sensors to refer to the true ambient pressure. The protection ®lter is mounted in the top by use of a locking ring and O-rings similar to the other silicon parts. The protective cap is used on ground to protect the nozzles form dust contamination. An O-ring at the bottom seals the cap when mounted in position. It is absolutely necessary to remove this cap before launch. For that reason it is clearly marked with both an inscription and an eye-catching colour. Finite element analyses have been performed on the pressurised parts of the housing at 32 bars to verify that the mechanical stress is acceptable. The outer coating to be used is a functional surface, exhibiting different optical properties for different wavelengths. The present most promising ®lm should maintain an equilibrium temperature in direct solar irradiation below 50 8C without signi®cant internal power dissipation [8]. Besides assisting in the thermal control, this coating is mechanically extremely durable, chemically inert, and electrically conductive which prevents electrical charging in space. The complete microthruster unit, without the external electronics and propellant storage, weighs approximately 0.15 kg. 3.2. Electronics The electronic system integration of the hybrid microthruster is divided in two conventional systems, a digital and an analogue part with a unique interconnection. The complete system begins with the design of the nozzle positions in perpendicular pairs, shown in Fig. 3 as X and Y. This approach has several advantages, the ®rst and most obvious is redundancy and the second is to enhance the steering and ®ne tuning of the nozzle functions. In this hybrid microthruster design, the electronics are contained outside of the
Fig. 3. Hybrid cold gas microthruster layout.
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Fig. 4. Electronics block diagram.
mechanical envelope presented in the earlier sections, with an estimated mass of 0.2 kg. However, all electronics will eventually be implemented in multichip modules (MCM) [9] and can, therefore, be made redundant in a three- or fourway system without much loss of space or gain of mass. The actuators and sensors in the pod are monitored and controlled by regular analogue circuitry allowing the thrusters to be run differentially for maximum performance. The analogue control electronics have an advanced control system in need of reference inputs. These inputs are set by an embedded microcontroller. Fig. 4 shows the system integration starting from the right with the thruster pairs. Analogue signals are carried in closed control loops to and from the thrusters. The reference points are used within these closed control loops, which are connected to the microcontroller through a number of digital±analogue and analogue±digital converters (ADC/DAC). The new approach is to create a system based on conventional control engineering while increasing its ¯exibility by introducing a microcontroller as a smart device sending input to the control loops. There are further advantages by introducing microcontrollers. First, they are very power saving. A normal 8 bit modern microcontroller, only a few square millimetres in die form, consumes less than 0.13 W in full operation. Second, they include advanced features like controller area network (CAN) and serial peripheral interface (SPI).
The microcontroller in this application connects each thruster to the CAN bus which provides an easy and modular structure of the microthruster system concept. Furthermore, several ADC/DAC are hooked on to the microcontroller through the Motorola SPI. By using widely accepted electronics solutions a large range of products are available for use within the system. Furthermore, most of the SPI chips are compatible, meaning that individual chips may be changed for different applications and requirements without having to redo the complete electronics layout. The CAN is also increasingly used within the space community and this provides a solid and simple base for including this thruster in a space project. The user only needs to specify the CAN address of the individual thrusters in order for them to be fully operational on the CAN bus. The use of both conventional analogue control electronics and new digital interfaces gives the possibility to maximise the abilities of electronics engineers with high analogue skills and modern computer programmers with high digital skills. 4. Microsystem manufacture The primary material for the different microsystem units of the hybrid cold gas microthruster system is single crystalline silicon, shaped by anisotropic wet etching and/or deep
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reactive ion etching (DRIE). Direct bonding joins structured silicon wafers in order to eventually forming the complete units. Evaporation and patterning of different metals, i.e. tin±lead alloy, tin±silver alloy, titanium, copper, gold, and chromiumÐprovides conductors and solder-pads for electronics, sensors, and actuators. Other thin ®lm techniques, e.g. low pressure chemical vapour deposition (LPCVD) give wear resistant coatings on particularly exposed details. 4.1. Nozzle unit The nozzle unit comprises four independent nozzles, divided by 908 angles on a 40 mm diameter disc. Heat exchangers are integrated directly upstream of each nozzle,
and platinum thin ®lm heater elements combined with temperature sensors are included on the outside surfaces of all heat exchanger structures. The nozzle shape is essentially a two-dimensional converging±diverging outlet with rectangular cross-section (Fig. 5). Bayt et al. [4] have demonstrated slightly larger nozzles of the same type, made from silicon and glass. In the present case, the inlet angle is 27±318 and the outlet 158, and the throat widths are 10, 30, or 40 mm. Each nozzle furthermore comprises two identical halves bonded together. The nominal height of two such halves is 100 mm. The actual shape of each manufactured nozzle is carefully monitored using white-light interferometric pro®lometry. The aligned bonding allows for smaller nozzle heights and homogenous
Fig. 5. The converging±diverging nozzle half. The throat width in this case is 10 mm. The close-up on the throat section readily shows the straight walls obtained.
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Fig. 6. The heat exchanger structure included directly upstream of each nozzle. This half consists of 3 mm thin curved fins, 40 mm high, and spaced 7 mm apart.
materials. The layout of the nozzle shape has not yet been optimised. The nozzles are free-hanging tubular structures in order to minimise heat loss to the supporting silicon. Into this tube, the heat exchanger structure is included, comprising a vast array of narrow curved ®ns, 3 mm thin, spaced 7 mm apart (Fig. 6). In order to address the evaluation of single nozzles correctly, units without the heat exchanger have been manufactured as well. Here, the throat widths are 15, 35, and 45 mm, respectively. 4.2. Valve unit This valve unit is a normally-closed proportional piezoelectrically actuated device. Five multilayered piezoelectric lead±zirconium±titanate (PZT) elements perform the valve
action, lifting a central silicon cap from the coated silicon valve seat (Fig. 7). The valve seat is 2.8 mm square with a 200 mm wide seal. The multilayered design limits the required drive voltage to 50 V for a maximum stroke of 4 mm. The total actuator height is 4 mm. The folded piezoactuator set-up has two major advantages. First, problems with the coef®cient of thermal expansion or ageing in the PZT elements are eliminated. Second, if both the inner and outer elements are operated while opening and closing, possible reaction disturbances can be avoided. The actuators are assembled using a thin ®lm soldering system. Here, the bottom platinum layer of each actuator is soldered to a chromium±copper±gold multilayer pad, using lead-tin solder. The top PZT layer of the actuator is deposited with titanium±copper±gold and soldered to chromium± copper±gold pads on the silicon yoke using tin±silver solder (Fig. 8). These techniques allow for precision mounting of
Fig. 7. The valve unit. Right: overview, showing the basic assembly of the complete valve. Left: IR image of an assembled valve seal, showing the seat and cap parts with electrodes and soldering pads after wafer bonding.
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Fig. 8. The actuator mounting and connection. Right: the five PZT elements soldered in place. Left: thin film conductors connect the electrodes of the multilayered actuator.
the actuators on the silicon parts, which is necessary for the successful operation of the valve. The internal electrodes of the multilayer structure are connected using evaporation of gold through shadow masks onto the sides of the actuators (Fig. 8). Wear resistant coatings cover the valve seat in order to avoid rapid degradation of the sealing ability. Parylene and titanium±aluminium nitride have been tested and compared to bare silicon valve seats. 4.3. Filter unit The ®lter unit is realised from combining two different types of ®lters, forming a three-wafer stack. The active ®lter parts are located in the bond interfaces (Fig. 9).
The ®rst ®lter part consists of approximately 50 mm wide remaining ridges on a slightly depressed surface (Fig. 9). These ridges are placed near the gas inlet. This ®lter type is single-sided, and bonded with a wafer containing outlet holes. The fenced slot is the active ®lter part. The second ®lter part, crossed v-grooves, consists of narrow v-grooves in the wafer, together with inlet holes (Fig. 9). By bonding this wafer to a similarly patterned wafer, the crossed v-groove ®lter is obtained, forming a 908 angle between v-grooves. The narrow v-grooves are the active ®lter part. To complete the ®lter stack, the slot ®lter is bonded to the inlet side of the v-groove ®lter. The ridge height in the slot ®lter was measured to 1.5± 4 mm by a stylus pro®lometer, and the v-grooves were fully etched from a mask width of 1.5±2 mm.
Fig. 9. The filter unit. Top left: 40 mm diameter filter disc after bonding (IR image). Bottom left: filter cross-section, the upper bond interface contains the vgroove part and the lower interface contains the slot part. Top right: the v-groove filter part (close-up). Bottom right: the slot filter part (close-up).
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Fig. 10. Mass flow and thrust force as a function of inlet pressure for nozzle with throat dimension 35 mm 100 mm, without the heat exchanger structure.
5. Results and discussion The performance of single nozzle elements, without the heat exchanger structure, have been evaluated by measuring the mass ¯ow and thrust force as a function of inlet pressure (Fig. 10) [10]. These data are derived from pressure drop measurements from ®xed volumes, releasing the gas into vacuum, while monitoring the impact of thrust by placing the thruster on top of a mg-resolution scale. By correlating the thrust force obtained to the current mass ¯ow, the exit velocity can be calculated (Eq. (1)), and thus, the speci®c impulse of the thruster. The nozzle with 35 mm 100 mm throat section reaches an exit velocity of 445 m/s, corresponding to a speci®c impulse of 45 s. This is 66% ef®ciency according to the present area expansion, considering the ideal exit velocity from Table 1. Obviously, there is much room for design improvements, changing, e.g. the area expansion, nozzle length, and rotational symmetry of these nozzles. Bayt and Breuer quoted 460±590 m/s for their nozzles at high pressures, while viscous losses degrade performance at lower pressures (i.e. lower thrusts) [11]. Such behaviour is also con®rmed by our experiments [10]. The mass ¯ows of the more elaborate nozzles in the nozzle unit are expected to be restricted, due to the heat exchanger, in direct comparison to the results from these single nozzles. Thus, the throat dimensions may need adjusting in this design. However, the operation of the heater is expected to increase the speci®c impulse. MEMS valves for space applications generally exhibit poor leakage characteristics. Micropropulsion missions are
inherently more sensitive to leakage than larger sized spacecraft. In order to maintain lifetime and performance when shrinking the system to micropropulsion, the valves need leakage rates below 10 5 scc/s of helium [12]. These valve parts have been tested for 30,000 work cycles with parylene or TiAlN coating on the valve seat. The leak rates as a function of sealing force have been measured using a leak detector (Fig. 11). For the coated samples, the leak rate remains suf®ciently low. The incoming gas needs ®ltration, so the mass ¯ow capacity of the ®lters must be tested in order to correctly dimension the ®lter necessary for the mission. The mass ¯ows of different gases through an 18 mm diameter ®lter unit to atmospheric pressure have been monitored at different inlet pressures (Table 2). The performance of a propulsion system can be described using the system-speci®c impulse [13], de®ned as the total Table 2 Mass flow of the filter for different inlet pressures of different gases Gas
Pressure (MPa)
Mass flow (10
N2
0.2 0.4
0.60 3.1
He
0.2 0.4
0.12 0.54
Ar
0.2 0.4
0.69 3.6
Xe
0.2 0.4
2.0 10.1
6
kg/s)
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Fig. 11. Leak rate as a function of sealing force for pristine and used valve seats with different coatings.
impulse (Itot) delivered by the system divided by the system mass (mPS), thus, Issp
Itot mPS
(4)
Here, the mass of the propulsion system can be expressed in terms of the individual masses of thruster units (mt) control electronics (me), propellant (mp), and storage tank (mST). Using four microthruster systems for full three-axis stabilisation, the propulsion system mass become mPS 4mt 4me mP mST
(5)
The mass of propellant needed is determined from the mission requirement and the spacecraft total mass. Using conservation of momentum, the velocity change Dv obtained by the spacecraft ejecting propellant is described by mF Dv ve ln (6) m0 where m0 is the initial and mF the ®nal mass of the spacecraft, respectively. Thus, the total amount of propellant needed to deliver the total impulse in Eq. (4) for a certain mission-speci®c Dv requirement is described by mP m0 1
e
Dv=ve
(7)
Using this, the total impulse delivered by the system is readily obtained as Itot mP ve
(8)
Different propulsion assignments on speci®c spacecraft missions require certain Dv manoeuvres, or their equivalent. For example, a 5 years ®ne-pointing earth-observing cluster mission typically requires 5±50 m/s Dv equivalent for attitude control, while the deep space interferometer mission (0.5±1 years duration) needs 100±300 m/s [14]. Now, observing that the propellant mass and storage system mass depend on the total amount of propellant
needed for speci®c Dv requirements and, thus, on initial spacecraft mass, the system-speci®c impulse can be estimated. Furthermore, the required mass fraction of the propulsion system to the total spacecraft mass [13] can be expressed by combining Eqs. (4), (5), (7) and (8) into Isp mPS 1 e Dv=ve (9) ms=c Issp Here, the total spacecraft mass ms/c is identical to the initial spacecraft mass m0. Assuming a storage pressure of nitrogen at 24 MPa in a spherical titanium tank with wall thickness 4 mm, a speci®c impulse of 45 s, and mass estimates from the above descriptions, the system-speci®c impulse and mass fraction of the propulsion system have been estimated (Fig. 12). The amount of propellant have further been adjusted by a leakage assumption based on a He leak rate of 1 10 6 scc/s, which translates to 3:8 10 7 scc/s for N2 if the leak rate is considered to be inversely proportional to the square root of the molecular mass [14]. A 5 years mission, thus, entails a total amount of leaked nitrogen of 960 scc, i.e. 1:2 10 3 kg. These ®gures can be much improved by choosing a light-weight composite storage tank. Otherwise, the control electronics total weight of 0.8 kg may be eliminated by the successful development of MCMs. Furthermore, the expected increase of the speci®c impulse when operating the heater, and the possible gain of ef®ciency by better nozzle design also promises better performances. For cold gas propulsion systems, a mass fraction between 0.1 and 0.3 is recommended [1]. Now, the graphs in Fig. 12 can be used for comparing the merits of different propulsion systems, and estimate the total mass of the spacecraft. Inspecting Fig. 12, it is evident that the hybrid cold gas microthruster system is best suited for low Dv assignments, like attitude control, ®ne stabilisation, drag compensation, or short-term free-¯ying missions.
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Fig. 12. The system-specific impulse predicted for the hybrid cold gas microthruster system (solid lines), together with the estimated mass fraction occupied by the propulsion system (dashed lines). The estimates are presented as a function of mission Dv requirement and total spacecraft mass.
6. Conclusion The hybrid cold gas microthruster system has been detailed, suitable for spacecraft attitude control, stabilisation, and short-term light-weight missions. This system enables exacting space missions where extreme accuracy is needed. The performance of the microthruster system has been demonstrated. Data from which to design mission-speci®c thruster systems, capable of maximum thrusts ranging from 0.1 to 10 mN have been obtained. The speci®c impulse reaches 45 s for nitrogen propellant, which translate to an ef®ciency of 66% using this current nozzle design. Four piezoelectric proportional valves independently modulate the gas ¯ow to the nozzles, using an analogue closed-loop arrangement. Thus, the achieved thrust is continuously controlled. The leak rate of a sealed valve is below 10 5 scc/s He, even after extensive wearing. Ef®cient particle ®lters have also been included. Acknowledgements The European Space Agency (ESA) is gratefully acknowledged for ®nancing this project, together with The Advanced Microengineering research programme at Uppsala University, funded by the Swedish Foundation for Strategic Research.
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Biographies Johan KoÈhler currently investigates bonded silicon microsystems for Ê ngstroÈm spaceflight applications, primarily micropropulsion, at the A Space Technology Centre, Uppsala University. He was born in 1971 and received his MSc in Materials Physics from Uppsala University in 1997. Since then his research has been in silicon micromachining, mainly fusion bonding. Johan Bejhed was born in 1973. He received his MSc in Materials Physics from Uppsala University in 2001. Since 1999, he has been a technician at Ê ngstroÈm Laboratory, Uppsala University, and he has recently become the A Ê ngstroÈm Space Technology Centre. His main topics a PhD student at the A are fluid-handling microsystems in space applications. Henrik Kratz was born in 1974. He received his MSc in Engineering Physics from Uppsala University in 1999. He is a PhD student at the Ê ngstroÈm Space Technology Centre. His current research involves RF± A MEMS for space applications and piezoelectric actuators integration into silicon microsystems.
Fredrik Bruhn was born in 1976 and received his MSc in Physics from Ê ngstroÈm Uppsala University in 2000. He is currently a PhD student at the A Space Technology Centre where his research is into system integration for microsystems. Ulf Lindberg was born in 1958. He received his MSc in Engineering Physics in 1985 and in 1993 he received his PhD in Solid-State Electronics, both from Uppsala University. After working at the Electrum Foundation 1994±1997 and Biacore AB 1997±2000, he joined the Ê ngstroÈm Space Technology Centre and the Solid-State Electronics A department at Uppsala University in 2000. His main research interests is microstructure technology for space and biotechnical applications. Klas Hjort received his MSc in Engineering Physics in 1988 and his PhD in Materials Science in 1993, both from Uppsala University. He spent a post-doctoral year at the Institute of High-Frequency Electronics, TU Darmstadt, Germany, in 1994. He made his ability degree in 1997 (Docent, Ê ngstroÈm Material Science). Since 1997, he is an associate professor at the A Laboratory, Department of Materials Science, Uppsala University. He has 60 reviewed international papers (eight invited). Since 1988, he has worked on basic mechanical characterisation, surface and bulk microstructuring, solid-state wafer bonding, device fabrication of bulk acoustic wave microstructure devices (e.g. based on GaAs/Al±GaAs), and quartz BAW resonators. His present research areas are INP-based optoelectronic MEMS, diamond microstructures, accelerator-based deep ion projection lithography, and microsystems for space applications. Lars Stenmark was born in 1944 and received his MSc in 1969 from Stockholm University, Sweden. He participated in the first sounding rocket campaigns at ESRANGE 1968. He worked as a development engineer and later project manager at the Swedish Space Corporation 1968±1979. In 1979, he started the private company Micro Nova KB for developing new electro-optical instruments. In 1982, he started ACR Electronic AB, a hardware developer and supplier entirely dedicated to space applications. Since then, ACR has been a major supplier to the Swedish national space programme (e.g. ODIN and FREJA: primary and secondary structures, solar sensors, power conditioning units, cryogenic boxes for sub-mm receivers, etc.). Lars Stenmark has designed, tested, and flown numerous experiments and subsystems, e.g. micro gravity payload modules for the TEXUS programme, MASER, and GAS (Get Away Special on the Shuttle). Since 1998, he has left all conventional space technology in order to promote the introduction of microsystems technology in space. He is Ê ngstroÈm Laboratory, A Ê ngstroÈm currently founder and director of the A Space Technology Centre, Uppsala University.