8 MICROPROPULSON

Andrew D. KetsdeverPropulsion Directorate, United States Air Force Research Laboratory, Edwards AFB, CA, USA

KEYWORDS: Micropropulsion, propulsion, microspacecraft, microscale, system specific impulse.

8.1 Introduction

In its broadest definition, a micropropulsion system is one capable of producing a change in momentum of a microspacecraft. (Ketsdever, 2000) The United States Air Force Research Laboratory (AFRL) defines a microspacecraft as having a total mass of less than 100 kg. This broad definition of micropropulsion allows the inclusion of a wide range of concepts ranging from microelectromechanical systems (MEMS) fabricated thrusters with micrometer characteristic sizes to simply scaled-down versions of macro-scale thrusters. Regardless of the characteristic size of a particular micropropulsion system, severe design constraints will be imposed by resource limited microspacecraft, including:

Mass Volume Power Maximum voltage

Micropropulsion concepts will not only have to address microspacecraft mission requirements, but will also have to fit within the systems constraints imposed. Although this is true of all spacecraft subsystems regardless of spacecraft mission or size, the constraints of some microspacecraft may pose such severe design restraints that new paradigms will need to be developed. A complication to the design process for micropropulsion systems will be the fact that, in many cases, microspacecraft missions will require exceedingly low thrust levels, accurate and very low minimum impulse bits, and low thrust noise. Also, many microspacecraft missions currently under investigation require relatively large V, indicating that high specific impulse micropropulsion systems are necessary to perform these missions under strict mass constraints.

A wide array of micropropulsion systems have been investigated, designed, built, and even flown. As with larger spacecraft propulsion systems, these micropropulsion concepts incorporate both chemical and electric means of producing high speed propellant species to create thrust. Both solid and liquid chemical thrusters have been developed for use on microspacecraft missions where relatively large thrust levels are required. Electric propulsion thrusters including micro-electrothermal, electrostatic, and electromagnetic systems have also been investigated for missions where higher specific impulse is desired. Electrothermal thrusters use electrical means to heat a propellant, which include arcjets [See 2.4.02 Arcjets] and resistojets. Electrostatic thrusters include a class of propulsion systems that use electrostatic means to accelerate ions formed in a plasma discharge [See 2.4.05 Ion Thrusters]. Electromagnetic thrusters achieve acceleration of plasma propellants through a combined interaction between electric and magnetic fields [See 2.4.03 Hall Thrusters and 2.4.04 Magnetoplasmadynamic

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Thrusters]. A comprehensive survey of micropropulsion systems under development has been compiled by Mueller (2000). Each system brings unique design challenges including:

Microscale combustion, heat transfer, and fluid dynamics Microscale plasma formation Material compatibility Micro-valve leakage Passage clogging System reliability and lifetime Manufacturing and integration complexities

Many of these design challenges will be addressed in the following sections.

Micropropulsion is considered an enabling technology for microspacecraft by making possible missions which otherwise could not be performed. However, micropropulsion could also find application on larger spacecraft to compensate for very small disturbing torques (e.g. solar pressure), providing extremely fine attitude control. Several propulsive requirements for potential missions are given in the following section. Reducing the physical size, operating power, and in some cases operating pressure and temperature, may lead to significant reductions in operational performance and efficiency. For example, a microchemical propulsion system will require high pressure operation for efficient combustion at small physical scales. However, a particular microspacecraft mission may require low thrust levels indicating that a micronozzle must be employed with a small throat diameter. High temperatures along with small nozzle throat diameters will lead to a reduction in the operational Reynolds number, leading to significant viscous losses. High temperature microchemical systems will also suffer from high heat transfer rates which will further reduce the efficiency of these systems, particularly in MEMS fabricated thrusters where the thermal conductivity of silicon-based materials is quite high. On a positive note, there is significant potential for improvement in all manner of micropropulsion systems, leaving future micropropulsion developers many fruitful areas of basic research and applied engineering practice.

8.2 Sample Microspacecraft Missions and Requirements

Constellations or platoons of microsatellites working in collaboration have been envisioned to perform the functions of larger spacecraft in a distributed way. Microspacecraft constellations will benefit from increased survivability, flexibility in mission performance, distributed functionality, and more graceful degradation of overall capability. Partitioning the functions of a single large spacecraft into a number of smaller satellites operating cooperatively in close proximity or in widely dispersed orbits could lead to significant improvement in mission performance. In some cases, this approach enables missions that cannot be performed by a single satellite of any size. AFRL has proposed a standard definition of spacecraft with masses below 100kg as detailed in Table 1.

By the AFRL definition, the first microspacecraft launched into orbit from Earth was also the first man-made object launched, Sputnik. Sputnik had a mass of approximately 84 kg. Figure 1 shows Sputnik I, which was launched on October 4, 1957. Since the late 1950’s, trends in electronics, material sciences,

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and general mechanics have led to advances in micro-scale components. This has led to a significant increase in the performance of micro-scale systems which have been used on microspacecraft, making today’s small spacecraft far more capable then their early predecessors. Today, universities are capable of launching very sophisticated microsatellites. (Sweeting, 2002) Sputnik was a fairly simple radio transmitter launched into orbit on a converted SS-6 nuclear missile, while today microsatellites with sophisticated communication, attitude control, and propulsion systems are being designed, developed, and launched.

Table 1: AFRL satellite classification

Total Spacecraft Mass Description10-100 kg Microspacecraft1-10 kg Nanospacecraft<1 kg Picospacecraft

Figure 2 shows FalconSat 3 designed and developed by the United States Air Force Academy (USAFA). FalconSat 3 was launched in March, 2007 as part of the secondary payload program for the Evolved Expendable Launch Vehicles (EELVs). Fifty years after the launch of Sputnik, a highly capable microsatellite with a mass of approximately 50 kg had payloads which demonstrated an advanced micropropulsion attitude control system and an ability to characterize the ionosphere and local spacecraft plasma environment.

Figure 1: Sputnik I launched October 4, 1957 by the Soviet Union (Photo courtesy of NASA).

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Figure 2: FalconSat 3 launched March 8, 2007 (50 years after Sputnik I) by the United States Air Force Academy (Photo courtesy of the Department of Astronautics, USAFA).

Several studies (Janson et al., 1999; Mueller, 2000; Schilling, Spores and Spanjers, 2000; Ziemer et al., 2008) have provided propulsive requirements for many microspacecraft missions including Earth observation, communications, space environmental sciences, satellite servicing, satellite inspection, and technology demonstration. For the proposed TechSat21 mission, which has since been canceled by AFRL but is still a good example of microspacecraft propulsive requirements, the proposed total V was 390 m/s of which 50 m/s was required for orbit raising, 20 m/s for drag makeup, 200 m/s for stationkeeping over a 10 year mission lifetime, and 120 m/s for a deorbit maneuver at the spacecraft’s end of life (EOL).(Schilling, Spores and Spanjers, 2000) A minimum impulse bit of 2mN-s was required for close proximity maneuvers. An early version of the TechSat 21 experiment called for formation flying maneuvers over a 30 day period with a V requirement of 30 m/s and a minimum thrust level of 2mN. This would have allowed for a series of formations to be flown with separation distances between 5 m and 5 km. Mueller (2000) also gives a range of propulsive requirements for microspacecraft slew maneuvers with minimum impulse bits ranging from 10-8 to 10-3 N-s.

Janson et al. (1999) gives a good overview of potential micropropulsion maneuvers and the V required. The V requirement for a microspacecraft to perform a minor altitude raising maneuver of 1 km at an initial altitude of 700 km was calculated to be about 0.5 m/s for a 10 minute firing time. For North-South stationkeeping in GEO for approximately 1 week duration the V requirement is over 300 m/s. A deorbit burn at an initial altitude of 700 km requires a V of over 330 m/s for a low thrust spiral maneuver over a 1 week period. On-orbit servicing or inspection missions may require up to 500 m/s of V for orbit changing and attitude control. Many microspacecraft missions currently under consideration will have relatively large V requirements, indicating that the specific impulse and efficiency of micropropulsion systems will need to be correspondingly large. However, a large number of microspacecraft missions will also require a rapid response to changing conditions to be effective.

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These missions will require relatively large thrust levels, perhaps as large as 100’s of mN. As with their larger counterparts, micropropulsion systems will continue to seek high thrust, high specific impulse solutions.

The Laser Interferometer Space Antenna (LISA) mission will require precision formation flying to detect gravitational waves produced by processes in the universe. To perform this mission and the precursor LISA Pathfinder mission, a series of propulsive requirements have been derived. Thrust levels from 5 to 30 N with precision better than 0.1N have been demonstrated with colloid and field emission thrusters. (Ziemer et al., 2008; Scharlemann et al., 2008) Thrust noise levels less than 0.1 N/Hz1/2 are also required, making the LISA mission one of the most stringent in terms of micropropulsion requirements. LISA Pathfinder will fly both the LISA Test Package (LTP) and the Space Technology 7 (ST7) payloads which contain FEEP and colloid propulsion respectively. Although this is not an exhaustive list of propulsive requirements for microspacecraft missions, it is intended to give a broad scope and range of requirements that will need to be met by future micropropulsion systems.

8.3 Micropropulsion System Design Considerations

As previously mentioned, there is a wide ranging definition for micropropulsion which includes concepts from scaled down versions of larger-scale thrusters to specifically designed, MEMS fabricated propulsion systems. This and following sections will investigate some of the design considerations required to successfully scale down existing thrusters and some of the techniques for fabricating unique MEMS designs.

Design issues for micropropulsion systems have arisen from scaling issues, the use of corrosive propellants with new materials, contaminant handling in propellants, limitations in MEMS fabrication techniques, thermal cycling, and environmental interactions. In general, micropropulsion systems will experience decreased performance over their larger-scale counterparts due to losses associated with small characteristic sizes, limitations on system mass and power, and the lagging development of materials and micromachining techniques for propulsion specific applications. Engineers must address these issues by making use of novel approaches that use small-scale properties to the overall system’s benefit. With this in mind, careful attention should be paid to the characteristics that scale favorable with reduced size. These characteristics hold the key to the design of efficient micropropulsion systems.

8.3.1 General Design Considerations

A complication for microsystems of all types is the fact that as a typical device is scaled down, the volume goes with the characteristic dimension, L, cubed whereas the surface area only goes with L2. Therefore, microsystems tend to have large surface area to volume ratios relative to their larger scale counterparts. This has consequences for heat transfer and gradients of temperature, velocity and species. In the heat transfer case, the heat loss in the system will increase as the characteristic length scale is reduced. For example in a 1-D heat transfer case, the energy transferred per unit area through conduction is given by Fourier’s Law as

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Q̇=−k dTdx (1)

where k is the thermal conductivity of the material. For microscale systems, the value of dx will be small while the device area to volume ratio remains relatively large. This indicates from Eq. (1) that the amount of heat transfer in microscale devices can be relatively large. Another way to look at this situation, is that large thermal gradients in small scale devices will not be possible unless very low thermal conductivity materials can be identified and used. For systems (such as propulsion) that rely on temperature gradients for efficiency, heat transfer issues will become increasingly important as the characteristic size of the device decreases.

Heat conduction through relatively thin material will generally produce uniform system temperatures and the thermal response time of a typical microstructure will be low. Heat conduction is complicated by MEMS fabrication in silicon substrates with thermal conductivities that are generally higher than many aluminum alloys. Alexeenko et al. (2005) showed in a coupled thermal-fluid simulation that a MEMS fabricated silicon thruster reached a uniform temperature, from the combustion chamber to the nozzle exit plane, almost instantaneously. In most cases, active cooling of the silicon microthruster was required for survivability. In micronozzle flows, the maximum heat transfer occurs at the nozzle throat (minimum area). The heat transfer rate at the throat varies as Re0.8/D where Re is the flow Reynolds number and D is the nozzle throat diameter. Therefore, in micronozzle flows, the amount of heat transfer between the gas and nozzle walls is a competition between the characteristic size, D, and the viscous losses in the nozzle governed by the Reynolds number.

8.3.2 MEMS Fabrication Techniques

MEMS encompasses a wide range of systems that include the integration of mechanical components, sensors, actuators, and electronics onto a common silicon substrate.(Gad-el Hak, 2002) Generally, MEMS devices have characteristic dimensions of less than 1mm and more than 1m. MEMS fabrication techniques include silicon micromachining through chemical etching or electro-discharge, lithography, molding, and thin-film deposition. Micromachining selectively etches parts of a silicon wafer whereas thin-film deposition can add new layers all in an attempt to form mechanical and electronic devices. Today, nanoelectromechanical systems (NEMS) processes are being developed which could lead to further breakthroughs in propulsion system design. Although many of these fabrication techniques are currently limited, the development of these areas for terrestrial applications will continue to have a strong impact for propulsion system engineers. An excellent review of these technologies and their applications can be found Gad-el Hak (2002).

A MEMS fabricated micro-resistojet (Lee et al., 2008) is shown in Fig. 4. The process for fabrication of the device is shown in Fig. 5. First, a photoresistive material is spun on a silicon wafer. The wafer is exposed to Ultraviolet light through an intervening mask. Light exposed sections of the photoresistive material are removed in a developing process. Deep reactive ion etching (DRIE) (Gad-el-Hak, 2002) was used to micromachine the 100 m wide gas expansion slots (pseudo-nozzles) in areas where the photoresistive material was removed. The DRIE process does not affect the silicon substrate in areas

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Open Inlet

Vac

uum

B

oun

dary

Expansion Slots Etched with DRIE (Deep Reactive Ion Etcher)

Ultravioletlight

Photoresist

Silicon wafer

Mask

Metallization

Exposed photoresist Unexposed photoresist

Developer removes exposed photoresistexposed photoresist

Acetone removes unexposed photoresistphotoresist

Ink

- 300 Å Ti (bottom), 600 Å Pt (middle), 750 Å Au (top layer)

where the photoresistive material remains. Once the expansion slots are machined, a secondary photoresist process is performed to mask areas for the thin film heater deposition. The thin film heater is composed of a 0.03 m thick layer of titanium, a 0.06 m thick layer of platinum, and a 0.075 m thick layer of gold. The layers act as a bonding agent with the silicon, a diffusion barrier, and the current carrying layer respectively. Other examples of MEMS fabricated thrusters can be found in Mueller (2000).

Figure 4: MEMS fabricated heater chip for a micro-resistojet concept.

Figure 5: MEMS fabrication techniques employed in the fabrication of the micro-resistojet.

8.3.3 System Derived Specific Impulse

In an attempt to quantify the fundamental viability of a micropropulsion concept for a particular microspacecraft mission, the concept of an effective or system comparable specific impulse has been developed. The effective (or system) specific impulse modifies the thruster’s intrinsic specific impulse by taking into account system penalties such as:

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High-pressure propellant storage and handling Micro-valve leakage Unusable propellant at EOL of a thruster Massive power supplies

Although the idea of a system specific impulse is useful for all types of propulsion systems, it is particularly useful for micropropulsion systems where systems-level constraints are paramount.

The effective specific impulse of a thruster system in terms of the excess mass associated with high pressure propellant storage, micro-valve leakage, and left over propellant at EOL can be written as (Ketsdever, Wadsworth and Muntz, 2000)

Ispeff=

Isp [1−( po ,despi , stor )](1+δ )[1+( M t

M p ,s ) ] (2)

where Isp is the intrinsic specific impulse, po,des is the design stagnation pressure, pi,stor is the initial propellant tank storage pressure, is the fraction of propellant lost to valve leakage, M t is the propellant tank mass, and Mp,s is the mass of the stored propellant. Similar formulations that take into account electrical subsystem (power processing units) mass, propellant feed system mass, and other propulsion system dry mass can be envisioned. In this way, propulsion devices that may have relatively low intrinsic performance can find niches where their combined benefits to the overall microspacecraft system outweigh any single drawback.

8.4 Chemical Combustion at Microscales

Liquid and solid propellant thrusters are specifically targeted at microspacecraft missions where high thrust is required or where minimal power might be available. Generally, missions requiring high thrust also require spacecraft rapid response. As is the case for all space chemical propulsion systems, propellants stored in either a liquid or solid state have several systems advantages over the use of gaseous propellants stored at high pressure. Typical liquid propellants (hydrazines, nitrogen tetroxide, methyl amine, chlorine trifluoride) are volatile and toxic which leads to material compatibility issues with MEMS materials like silicon. Combustion temperatures are also quite high leading to heat transfer issues discussed in previous sections.

Because of the large heat transfer rates expected in small scale devices, lower combustion chamber temperatures are expected in micropropulsion systems utilizing combustion where the heat loss will be proportional to L3/2. (Yetter et al., 2007) Lower combustion chamber temperatures will result in low specific impulses. At smaller scales, velocity and temperature gradients near the combustion chamber walls will tend to increase. This will produce larger mass, momentum and energy transport rates near the wall making it difficult to maintain large differences between wall and bulk flow properties critical for efficient operation.

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The mixing length of propellants (fuel and oxidizer) is often used as a measure of the combustion efficiency. Every chemical reaction process requires a finite time on the order of a few mean molecular collision times of gas phase species in the combustion chamber. Therefore the combustion chamber must have a minimum length of a few molecular mean free paths at the combustion chamber pressure to be effective. In this scenario, length can be reduced at the cost of higher combustion chamber pressure or shorter chemical timescales. The characteristic length of a combustion chamber is given by

L¿=V c

At (3)

where Vc is the combustion chamber volume, and At is the nozzle throat area. The characteristic length of a chemical propulsion system’s combustion chamber depends on several factors including the propellant type, injector type and configuration, chamber geometry, and whether a catalyst is used. As one example, evaporation time for a droplet from a spray will need to be reduced, which leads to equivalent reductions in droplet size in order to obtain efficient combustion. For micropropulsion systems, the design criteria is for L* to be as small as possible while maintaining acceptable performance.

Monopropellants appear attractive for micropropulsion applications since the mixing of a fuel and oxidizer is not required. (Hitt, Zakrzwski and Thomas, 2001; Wu et al., 2006) However, long-term survivability of catalyst material at elavated temperatures will be an issue for these systems. A variety of new propellants are being considered to produce efficient combustion. (Yetter et al., 2007)

8.5 Micronozzle Flows

Many micropropulsion systems will require the expansion of propellant gases through micronozzle geometries. As with large scale propulsion systems, nozzles convert thermal energy in a propellant into kinetic energy. For micropropulsion systems, the expansion of a propellant gas through a small scale nozzle poses several challenges. The low thrust levels needed for many microspacecraft missions will require either low pressure operation of a thruster or a decrease in nozzle throat dimensions. Both of these scenarios lead to low Reynolds number flows. Additionally, high temperature gas expansion from micro-chemical or micro-electrothermal thrusters will also continue the trend of lowering the operating Reynolds number of nozzle flows. The Reynolds number is a measure of nozzle efficiency in terms of the viscous losses inherent in the subsonic boundary layers near the nozzle surfaces. At low Reynolds number, the viscous losses in a micronozzle become large, degrading performance. At sufficiently low Reynolds number, the viscous boundary layer (i.e. the subsonic flow region) can encompass a significant fraction of the micronozzle geometry.

A typical two-dimensional, MEMS fabricated micronozzle is shown schematically in Fig. 5. These 2-D nozzles have square or rectangular throat and exit plane sections which leads to inefficiencies due to sidewall boundary layer formation; however, this micronozzle geometry is ideal for MEMS fabrication techniques. Results from experiments with 34 micron throat diameter, 2-D nozzles indicated that thrust efficiencies from 0.1 to approximately 0.8 could be expected in the Reynolds number range of 500-3500

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for a nozzle expansion ratio of 7.1:1. A larger expansion ratio (16.9:1) micronozzle showed improved performance for the same range of operating Reynolds number. A recent study by Louisos and Hitt (2009) showed the numerical performance of a traditionally MEMS fabricated 2-D micronozzle geometry computed in a 3-D domain. Figure 5 shows numerical results for a simulated 2-D MEMS fabricated micronozzle.

Figure 5: An illustration of the subsonic layer as denoted by Mach contours less than unity for selected cross sections in a 200μm deep nozzle operating at Re=60 (top) and Re =30 (bottom).

Figure 6 shows Mach number contours for a helium flow through a 3-D, conical micronozzle geometry at two different Reynolds numbers. These results were obtained using the Direct Simulation Monte Carlo (DSMC) numerical method. Note that the subsonic boundary layers have grown significantly at the low Reynolds number in both Fig. 5 and Fig. 6 indicating degradation of micronozzle performance at the lower pressure operation. Figure 7 shows the intrinsic specific impulse of the micronozzle flow at a Reynolds number of 60 as a function of axial position through the micronozzle (X=0 corresponds to the micronozzle throat). Figure 7 indicates that the intrinsic specific impulse decrease from a maximum just past the throat to a minimum at the exit plane regardless of the micronozzle length. In the case of the full length nozzle, the specific impulse at the micronozzle exit plane is actually lower than the specific impulse at the micronozzle throat indicating that the nozzle at this Reynolds number is actually degrading performance. In fact, Fig. 8 shows that at Reynolds numbers below approximately 100, a thin-walled, sonic orifice tends to outperform a typical micronozzle geometry.

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Figure 6: Predicted Mach number contours for helium flow inside a micronozzle geometry at two different throat Reynolds numbers. Subsonic, viscous boundary layer is indicated by series of contours with M<1.

Figure 7: Impact of micronozzle length on performance for a helium flow with a throat Reynolds number of 60. The full length nozzle is 1.07x10-2 m in length.

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0 50 100 1500.600000000000001

0.700000000000001

0.800000000000001

0.900000000000001

1

1.1

Re*

Fn /

Fo

Figure 8: Ratio of nozzle to orifice thrust for helium propellant as a function of throat Reynolds number.

8.6 Plasma Formation at Microscales

Because of the large surface area to volume ratios in micropropulsion systems, electron losses to chamber walls will be the limiting factor in the efficiency of microscale plasma devices. In large scale ion and Hall-effect thrusters, magnetic fields are used to contain the electron motion and prevent them from being lost to the discharge chamber walls. The power lost to the walls can be estimated by the ratio of the current to the wall to ionization potential of the propellant. In this case, the power lost at the walls scales directly with the characteristic scale of the thruster, L. These magnetic fields also increase the electron path length in the discharge chamber, increasing the probability of ionization. As these types of systems scale down, proportionately larger magnetic fields are required to provide efficient ionization. Essentially, the magnitude of the magnetic field required scales as

B=me v pq Rg

(4)

where me is the electron mass, vp is the electron velocity perpendicular to the magnetic field, q is the elemental charge, and Rg is the electron radius of gyration. To effectively contain the electrons, R g has to be a reasonably small fraction of the discharge chamber radius. As the scale size of the discharge chamber decreases, the required magnetic field increases. For example, a 1mm diameter ion engine would require on the order of a 10 Tesla magnetic field for electron containment. High power solenoids or heavy permanent magnets would be required, clearly indicating that these types of thrusters will encounter serious issues at small scales.

8.7 Micropropulsion Systems: Current State of the Art

A complete review of micropropulsion systems cannot be performed within the context of this section. Instead, this section is intended to give an update on system development since the last comprehensive review of Mueller (2000).

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8.7.1 Microchemical thruster development

Louisos and Hitt (2009) discussed the general performance of a high purity, hydrogen peroxide (H2O2) monopropellant microthruster. One benefit of hydrogen peroxide is that the products of the combustion are non-toxic. Decomposition of the monopropellant occurs in a catalyst bed producing a flame temperature of 886 K with 85% pure H2O2. Simulations for stagnation pressures from 5 kPa to 250 kPa yielded thrusts in the range of 1-20 mN after gas expansion through a micronozzle. Yetter et al. (2007) described the development of a liquid monopropellant thruster that uses electrolytic ignition. This thruster was fabricated from alumina-based ceramic tape with maximum operating temperatures of 1200 K. Thrust levels of approximately 100 to 200 mN were measured using both XM46 and RK315A propellants.

8.7.2 Micro-Pulsed Plasma Thrusters (PPT)

Pulsed plasma thrusters use an arc struck across a non-conducting material to ablate propellant and generate thrust. [See 2.4.06 Pulsed Plasma Thrusters] The PPT uses a surface discharge across a solid Teflon™ propellant to provide impulses in the range of 10 N-sec. (Spanjers et al., 2002) In recent designs, a high voltage (~3 kV) capacitor is directly coupled with a coaxial propellant module. The module consists of three electrodes each surrounded by an annulus of propellant as shown in Fig. 9. The PPT is operated by initiating breakdown between the innermost electrodes which produces enough plasma to initiate breakdown between the outermost electrodes. Pulses on the order of several Joules of energy have been produced at repetition rates of approximately 1 Hz. Thrust-to-power ratios of approximately 10 N/W have been demonstrated. Initial lifetime tests have been conducted that have run the thruster to the point where all of the propellant has been consumed. A benefit of the PPT scaling from its larger counterparts has been the ability to remove the initiating spark plug and igniter circuitry. The PPT has been flown on FalconSat 3, a microspacecraft designed and built by USAFA. Two of the PPT thruster modules are visible at the top of FalconSat 3 in Fig. 2 (large cubes with copper rods).

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Figure 9: Three-electrode PPT schematic (after Spanjers et al., 2002).

8.7.3 Vacuum Arc Thruster (VAT)

In the VAT design, a self-triggered vacuum arc is utilized to form a fully ionized metal plasma from a cold cathode. (Schein et al., 2002) Plasma velocities upwards of 30 km/s have been observed for aluminum cathodes with a corresponding specific impulse of about 3000 sec. The specific impulse varies with cathode material. Discharges with arc currents between 10A and several kA have been produced with pulse lengths on the order of several microseconds. One benefit of the VAT over a typical PPT is its ability to operate at much lower voltages (<12 V). The VAT is operated in a pulsed mode up to 100 Hz with typical average power levels from 1 to 100 W. Initial results have shown that the VAT can obtain individual impulses on the order of 1 N-sec with thrust-to-power ratios of approximately 20 N/W. The overall VAT mass, including the power processing unit (PPU), is about 0.1 kg with an overall efficiency estimated to be about 13%.

8.7.4 Field Emission Electric Propulsion (FEEP)

The FEEP concept uses electrostatic acceleration of metal ions formed through field ionization of liquid metal propellants, typically cesium or indium, as shown in Fig. 10. Both cesium and indium FEEP thrusters are being developed for the LISA Pathfinder mission (as part of the LTP payload) which could launch as early as 2010. Single FEEP modules can operate at continuous thrust levels from 0.1 to 15 N with peaks near 35 N. They are capable of very small impulse bits (<5 nN-sec) with a very large specific impulse (4000-8000 sec). The great benefit of FEEP thrusters for extremely accurate missions like LISA is their low thrust noise values and high thrust resolution and repeatability. Thrust-to-power ratios of nearly 15 N/W have been produced with relatively high electrical efficiencies. (Scharlemann et al., 2008) Tajmar (2004) has investigated scaling FEEP thrusters using MEMS and other micro-fabrication techniques to produce large arrays of emitters. Because indium does not wet silicon, traditional MEMS techniques proved unsuitable. Laser machining was incorporated to produce an array of 441 emitters in a 5x5 mm area. Although the thrusters were able to be fired, currents of only 5A were able to be produced. Future design optimization could significantly improve the micro-FEEP performance.

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(a) (b)

Figure 10: FEEP thruster a) operating principle and b) operation of an 8cm-diameter module on indium (Photo courtesy of ARC, Austria)

8.7.5 Colloid thruster development

Colloid propulsion utilizes the charging of small liquid drops produced in Taylor cones which are accelerated through a high electrostatic potential [See 2.4.07 Colloid Thrusters] Colloid propulsion is currently being developed for use on the ST7 payload for the LISA Pathfinder mission. Thrust in the range of 5 to 30 N with specific impulse greater than 150 sec has been demonstrated. (Ziemer et al., 2008) As with FEEP thrusters, high thrust precision and low thrust noise also makes colloid thrusters attractive for high precision maneuvers on microspacecraft. The LISA mission architecture and flight unit for the ST7 payload are shown in Fig. 11. Arrays of micro-fabricated colloid emitters have been developed by Valasquez-Garcia, Martinez-Sanchez and Akinwande (2004). These colloid thrusters can have between 4 and 1025 emitters in a 0.64 cm2 area. Recent advances have also been made in the development of propellants for colloid thrusters where ionic liquids are being investigated.

(a) (b)

Figure 11: a) LISA mission architecture, b) Colloid thruster module for the ST7 payload. (Photo courtesy of the Jet Propulsion Laboratory)

8.7.6 Small-scale Hall thruster development

Low power Hall thrusters [See 2.4.03 Hall Thrusters] have been designed and tested. Recent developments in small scale Hall thruster development has been significant. (Diamant et al., 2008; Hargus and Charles, 2008; Ito et al., 2006; Biagioni et al., 2005) Diamant et al. (2008) investigated a low power cylindrical Hall thruster over a range of input powers from 70 to 220 W. Over this range, they

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observed anode efficiencies on the order of 15 to 35% indicating the issues that Hall thrusters will have for low power microspacecraft mission applications. Hargus and Charles (2008) investigated a 200W-class Hall thruster system operating with a 250 V anode potential. They observed maximum ion velocities corresponding to 192 eV with xenon propellant which yields an approximate specific impulse of 1700 sec. A system similar to the one tested by Hargus and Charles (2008) will fly on FalconSat 5, a follow-on microspacecraft designed and built by USAFA. A small Hall effect thruster is described by Bianioni et al. (2005). Stable operation was observed in the power range between 10 and 235 W with anode efficiencies in the range of 15 to 32%. This system utilizes permanent magnets to contain the electrons formed in the discharge. This study attempted to address issues in small scale Hall thruster development in which limited operational efficiency was caused by electron losses to discharge chamber walls and high heat transfer rates.

Ito et al. (2006) designed a much smaller Hall thruster system with a 0.5mm channel width and a 4mm outer diameter. A permanent magnet with a field strength of 1 T in the ionization channel was used. Average powers of 10-40 W were demonstrated with efficiencies in the range of 10-15%. The thrust ranged from 0.6 to 1.6 mN with specific impulses in the range of 300-850 sec. All of these small scale, low power Hall thrusters show the limitations on efficiency as the physical size and operational power levels are reduced. These limitations are obviously not unique nor restricted to Hall thrusters.

Acknowledgements

The DSMC micronozzle calculations in Figs. 6 and 7 were performed by Dr. Sergey Gimelshein of the University of Southern California. The micronozzle calculation in Fig. 5 was performed by Will Louisos and Prof. Darren Hitt of the University of Vermont. The photo in Fig. 2 is courtesy of Sgt. Michael Wickersheim of the United States Air Force Academy. The photo in Fig. 10(b) is courtesy of Dr. Martin Tajmar of ARC, Austria. Figure 11(a) is courtesy of NASA/JPL. The photo in Fig. 11(b) is courtesy of Dr. John Ziemer of the Jet Propulsion Laboratory. The author wishes to thank Dr. David Scharfe and Mr. Shawn Laabs for their help in preparing this article.

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