Paulo Lozano and Daniel Courtney

ON THE DEVELOPMENT OF HIGH SPECIFIC IMPULSE ELECTRIC PROPULSION THRUSTERS FOR SMALL SATELLITES Pestana Conference Centre Funchal, Madeira - Portugal 31 May 4 June 2010 ABSTRACT Paulo Lozano and Daniel Courtney Massachusetts Institute of Technology - Space Propulsion Laboratory 77 Mass. Ave. 37-401 Cambridge MA 02139 USA, (617) 258-0742, plozano@mit.edu The fundamental working principles and preliminary engineering design of electrospray-based electric propulsion systems for small satellites are presented. Electrospray propulsion is based on the electrostatic extraction and acceleration of ions from liquid surfaces. The exhaust is made up of ions falling across a high-voltage drop, thus reaching Isp s higher than 2500 s, making the propellant mass fraction negligible for most missions. The liquid is passively pumped by capillary forces inside a porous metal structure towards very sharp tips. Only when a voltage between these tips and a downstream aperture is applied, ions are emitted producing thrust. The thrust density of these devices is in the 0.1 to 1 N/m 2 range and can be scaled up or down with ease, making them ideal for small satellites. The applicability of this technology is discussed in the context of a series of examples of expected performance on CubeSat vehicles. 1. INTRODUCTION In recent years, miniaturization of space components has triggered a flurry of activity aimed at the design, construction, launching and operation of small satellites. At the same time, it has become increasingly clear that small satellites have a significant potential in the advancement of space technology for many different players, such as entrepreneurs and universities, research institutions and government agencies. This significance is for the most part due to the solid advantages introduced by low cost systems that are increasingly able to perform challenging missions. Advancements in micro-electro-mechanical systems (MEMS) make it possible to reduce the size of practically all spacecraft subsystems, a large number of payload hardware and ancillary components. On the other hand, the most demanding missions still require a level of complexity which is difficult to incorporate in small vehicles. A particular example is the propulsion subsystem, which provides the required maneuverability that enables a number of the most promising applications of small satellites, such as flight formation of satellite constellations, fine pointing capability, orbital maintenance and general orbit modification. The most common and convenient way to design and build subsystems with the required small dimensionality is direct scaling of components successfully used in larger platforms. For example, a macroscopic stiffening spring can be microfabricated in silicon to provide stiffness in a scaled version of hardware operating under the same physical principles. If this scaling is linear, then it should be possible to simply take a large space platform and miniaturize every component into a small volume to create a small satellite. However, traditional propulsion systems do not scale linearly. In fact, the smaller many of these systems are, the larger their relative size in comparison to the vehicle will be. Take for example a traditional chemical (hydrazine) monopropellant thruster. If the propulsive performance is to be maintained (same specific impulse, Isp) then the adiabatic flame temperature should remain constant, which at lower sizes means higher heat fluxes and the need of better dissipation structures

acting in the opposite direction to the desired scaling. In addition to this limitation, chemical thrusters require a number of ancillary components that are difficult to miniaturize, including high pressure valves, reservoir tanks, catalytic chambers, etc. Finally, the Isp of such systems is limited to about 200-250 sec, meaning that high ΔV maneuvers require a propellant mass fraction with respect to the initial spacecraft mass m0 is (g = 9.8 m/s 2 ), m p m 0 = 1 exp ΔV gisp (1) which for chemical rockets represents about 20% for a ΔV = 500 m/s. In consequence, there are incentives for a propulsion system with few ancillary components, little or no heat dissipation concerns and high Isp. In small satellites, an alternative is to use cold-gas (CG) systems. Many of their components can be miniaturized, however, their low Isp (at most 60-70 sec) limits their applicability to missions of small ΔV. In this paper we present an alternative for a propulsion system able to work at significantly higher Isp (> 3000 sec) in a system which is compact enough to be included in most small space platforms, including CubeSats. Such high Isp is only accessible through the use of electric propulsion, but in contrast with traditional plasma engines, which operate with degraded performance at small sizes [1], we make use of an ion-electrospray propulsion system which is typically too small even for the smallest of space platforms. The advantage in this concept is its high efficiency at the individual ion emitter level. The challenge is that in order for such system to be successful, a large number of ion emitting electrosprays need to be fired simultaneously to produce macroscopic thrust. Emitters are fabricated and integrated in a MEMS device using electrochemical etching techniques on porous metals. This concept is described in the following section. 2. ELECTROSPRAY PROPULSION Electrospray propulsion is based on the electrostatic extraction and acceleration of ions from ionic liquids, substances with exceptionally low vapor pressure and relatively high electrical conductivity [1,2]. Ionic liquids can be exposed to vacuum conditions with practically no thermal evaporation and can be electrically stressed to form conical tips (known as Taylor cones) where very strong electric fields develop, inducing ion emission when a potential difference of about Ve = 1-2 kv is applied between a relatively sharp emitter coated with an ionic liquid and a downstream aperture electrode. The thrust produced by each Taylor cone is about 1/10 of a micro-newton, which is too small except for very specific precision applications. On the other hand individual emitters are very small too and the thrust density would be more than 1000 N/m 2 for an emitter about 10 microns in diameter. This means that there is much to be gained in building densely packed arrays of individual emitters. Porous metal substrates are ideal materials for this application due to their ability to transport liquids via capillary forces. We have been able to produce emitter arrays using electrochemical microfabrication techniques with emitter-to-emitter spacing of about 300 microns, thus giving a current density of about 1-5 N/m 2, well below the theoretical maximum, but already higher than current ion engines, with the important benefit that electrospray thrusters are very compact since no gas-phase ionization is required and the propellant is passively fed from a nearby porous tank directly into the substrate containing the emitter tips. The basic configuration is schematically shown in Fig. 1.

Fig. 1. Conceptual view of an electrospray propulsion thruster Ions emitted from electrospray thrusters carry almost the full energy of acceleration provided by the power supply. Only about 5-8 ev of energy are lost during the emission process, while an amount on the same order corresponds to the width of the energy distribution [4]. This can be contrasted with the losses of existing plasma thrusters, typically in the 100 s of ev both for the energy deficit and distribution widths. This means that electrospray thrusters have substantial potential for operation at high efficiency, a characteristic of the utmost importance in electric propulsion technologies, all of which trade off propellant with power source mass. The specific impulse in this thruster is given by, Isp = 1 g 2 q m V acc (2) In here, Vacc is the acceleration potential and q m is the average specific charge of ions in the beam. For the nearly monochromatic distributions described above, we obtain an energy efficiency of η E = V acc V e > 98%. Additional losses make the efficiency drop to about 80-85%, mostly due to the polydispersive character of the emitted species. The ionic liquid propellant is fundamentally a room-temperature molten salt. As with any other ionic compound, its composition is given by cations [C+] and anions [A-] with no intervening solvent. Depending on the polarity of the power supply, emitted ions are of the form [CA]nC in the positive mode, and [CA]nA in the negative mode, where n is the degree of solvation of the ions in the vapor phase. In most operating conditions, ions are of the type n = 0 and 1. The effect on efficiency of this polydispersity is due to the fact that there are ions of dissimilar masses accelerated by the same potential field. This traduces to non-zero relative velocities in the beam with respect to the beam mean velocity, i.e., heating. Interestingly, this loss, the most severe in electrosprays occurs in the beam and not the device, which means that the net loss (~10%) does not traduce into heat dissipation in the engine [5]. Also, similar to other electric propulsion devices, electric neutralization is required as ions are ejected into space. In plasma thrusters, an electron emitting cathode is used for this purpose. In electrosprays this method would work. However, electrosprays are able to emit negative ions too, which are as effective as positives to produce thrust. In fact, an electrospray propulsion device would require two thruster heads firing in parallel with opposite charge and periodically alternate the polarity to avoid accumulation of counterions on the emitter surface that would otherwise produce electrochemical reactions, which might be detrimental to the thruster operation [2]. One of the main difficulties in the operation of electrospray thrusters is the efficient transport and distribution of propellant to the sharp structures from which emission occurs. This is particularly important in a system where compactness is important and there is little room for active pumping systems. Previous work has focused on silicon as the material of choice to construct the emitter structures [6]. The problem with silicon is that in its native or oxidized states has a strong

hydrophobic tendency towards ionic liquids. Several alternatives have been proposed to increase wettability, some of which have been partially successful but still a method is required to bring the liquid passively to the wettable surface from a remote container. An interesting alternative we now pursue is to use porous metal materials for this purpose. This is advantageous for several reasons: (i) provides the required liquid transport through capillarity action alone, (ii) produces clean ion emission at increased current levels and (iii) provides a large liquid-metal contact area that prevents prompt appearance of electrochemical reactions and simplifies the electrical contact. Fig. 2. Electrochemical microfabrication on porous metals The disadvantage of using porous metals is that there is practically no current-practice on the shaping of microstructures on them. Nevertheless, the benefits outweigh the difficulties and we have developed masking techniques and electrochemical etching processes to fabricate emitter arrays on porous substrates [7,8]. Fig. 2 shows an example of our preliminary work on 2-D porous tungsten arrays with an emitter spacing of 300 µm. The main challenges are at the moment emitter uniformity and reproducibility, although there is a relatively large number of parameters controlling the process that could be varied to achieve optimal results. Once the emitters are fabricated on the porous substrate, this is aligned and bonded to a housing structure containing the extractor, and possibly accelerator, electrodes as shown in Fig. 3. This housing structure could be fabricated in a variety of materials. Of particular interest is silicon due to the large knowledge pool related to this material. The thruster head would then be a MEMS device with a specific volume (per unit power) of about 1 cm 3 /W. Such small devices would be quite advantageous for applications in small satellites. In fact, the thruster size and mass (including propellant due to the high Isp) would be negligible in relation to the power processing unit (PPU). Commercially available high-voltage modules have specific volumes of about 10 cm 3 /W and efficiencies near 70%. While not as compact and efficient as the thruster itself, these power modules provide initial building blocks to produce a system which overall is much smaller than the host satellite.

Fig. 3. Schematic of an electrospray thruster array based on porous metal emitters 3. THRUSTER DESIGN To increase the thrust magnitude and keep a low profile on the spacecraft surface, arrays of individual emitters are produced as described above. In this section we use an example of a CubeSat-like satellite for the purpose of preliminary thruster sizing, thus assuming an electrospray propulsion system with the following characteristics: Power, P = 1.5 W (post PPU) Efficiency, η = 0.8 Specific impulse, Isp = 3500 sec Thrust density, F/A = 0.5 N/m 2 Mean specific charge, q m = 5 105 C/kg Current per electrospray emitter, Ie = 1 µa The accelerating voltage is given by, V acc = ( gisp)2 2 q m = 1177 V (3) The thrust is given by, F = 2ηP = 70 µn (4) gisp and the propellant mass consumption, assuming uniform thrust during a time t is given by, m p = Ft gisp (5)

For a corresponding thruster active area of A = 1.4 cm 2. This area would be equally distributed in two thruster heads operating with opposite polarity to avoid spacecraft charging. The polarity needs to be alternated regularly (at frequencies < 1Hz) to eliminate electrochemical degradation of the ionic liquid propellant. Adding structural components to the thruster, the total area would be a rectangle of about 1 2 cm (2 cm 2 thruster) containing a total of, N = F I e q m 2V e = 1021 emitters (6) Which means that each array would contain about 23 23 emitters, spaced out by about 370 µm between each other, which is within current porous metal micromachining capabilities. Fig. 4 illustrates the dual thruster configuration. Fig. 4. A minimum of two thruster heads side-to-side need to be fired in parallel to avoid spacecraft charging. Voltage alternation is needed to maintain chemical neutrality. Adding support electronics to commercially available HV modules, propellant tanks (a porousbased tank attached at the back of the thruster) and propellant itself, the total mass and volume of the 1.5 W electrospray propulsion system would likely be less than 100 g and 100 cm 3, respectively, or about 10% of a standard 1U CubeSat bus. 4. APPLICATIONS In this section we present several in-space maneuvers that could be performed with this electrospray propulsion system. To be consistent with our previous discussion on thruster design we limit our analysis to CubeSat-like spacecraft. Slew and Spin Maneuvers For this application, the 2 cm 2 thruster could be distributed in 4 smaller arrays located at the edges of 2 opposite faces of a 1U CubeSat of side l = 10 cm as shown in Fig. 5.

Fig. 5. Thruster distribution on a CubeSat-like satellite face One of the benefits of using electrospray propulsion is scalability, and given that microfabrication techniques are used to produce the arrays in the first place, such distributions should not tax in a significant way the mass and volume fractions of the original monoblock thruster. Assuming uniform mass distribution, the time it would take to slew the satellite about its principal axis by an angle Δθ is, t slew 4 3 MlΔθ F (7) Fig. 6 shows plots of the slew time and propellant consumption requirements for a full rotation Δθ = 2π for different spacecraft mass as a function of available power. Fig. 6. Full rotation slew maneuver time as a function of input power for different spacecraft mass In the same category of maneuver, spinning the satellite to a given angular velocity might be required to provide spatial rigidity in a particular direction. In this case, the time to spin a CubeSatlike satellite to an angular velocity θ is, t spin Ml θ 3F (8)

Mission time and propellant consumption for spinning the satellite to 1 Hz are shown in Fig. 7. Notice that in this case the propellant mass is independent of input power. Spin time for!" = 1 Hz (hours) 8 6 4 2 M = 1 kg M = 2 kg M = 3 kg Propellant mass (mg) 20 15 10 M = 1 kg M = 2 kg M = 3 kg 0 1 2 3 4 5 Power (W) 5 1 2 3 4 5 Power (W) Fig. 7. Spin maneuver time to 1 Hz as a function of input power for different spacecraft mass Orbital Climb One of the most demanding maneuvers for a propulsion system is changing the orbital altitude. Since electric propulsion is inherently low thrust, the typical picture of an impulsive Hohmann transfer is substituted by a gentle continuous push along a spiral trajectory. Changing altitudes might be of interest in missions that require probing at varying distances from the earth, or could also be of interest to de-orbit satellites once they reach their end-of-life to avoid the accumulation of unwanted space debris. The time to complete a spiral trajectory from an initial circular orbital radius r0 to its final value r is given by, t climb v 0 M F 1 r 0 r (9) Altitude change time (days) 70 60 50 40 30 20 10 P = 1 W P = 2 W P = 3 W Propellant mass (g) 8 6 4 2 P = 1 W P = 2 W P = 3 W 0 0 100 200 300 400 500 Orbital altitude change (km) 0 0 100 200 300 400 500 Orbital altitude change (km) Fig. 8. Orbital change maneuver time from a 700 km circular orbit at different power levels In here, v0 is the circular velocity at the initial orbit. Orbital descent from a circular orbit is described by the same expression, except that F is negative. Fig. 8 shows the orbital change time for

an initial circular 700 km orbit for a 1 kg satellite at different power levels and the corresponding propellant consumption. Notice that eclipse time has been neglected in this calculation. Walking in Orbit In flight formation maneuvers it is important to provide and maintain separation. In this case we assume two spacecraft are delivered in a circular orbit with no further separation. The thruster is then used to increase (or decrease) the orbital radius so that the slower (or faster) orbit would move the selected spacecraft to its correct position on the same orbit. For instance, for a separation in the positive (aligned with the velocity vector) direction, the thruster needs to fire backwards for half of the time required by the maneuver, followed by forward thrust during the second half. The angular separation for this maneuver is illustrated in Fig 9. Fig. 9. On-orbit angular separation profile as a function of time Notice that continuous thrust is assumed to minimize the maneuver time, T. Assuming the orbit remains nearly circular during thrusting, the total maneuver time to achieve a separation s would be, T 4s 3F M (10) Fig. 10 shows plots for 1, 2 and 3 kg spacecraft using electrospray thrusters to achieve a s = 10 km separation. Notice that time and propellant usage are independent of orbit altitude. The tradeoff here is on eclipse time, which is if course much more of a common occurrence in low LEO, and therefore the time would be larger in non-sunsynchronous LEO orbits, but likely not more than by a factor of 2. Once the orbit separation has been established, on-orbit propulsion requirements depend on local perturbations (e.g., atmospheric drag, photon pressure) and uncertainty in the relative position of the satellites. Such perturbations and uncertainties are small, and the amount of ΔV (or propellant required) depends on mission time. Solar pressure on a CubeSat is small and would required a cancellation ΔV of less than about 3 m/s per year at both LEO and GEO altitudes for a 1 kg satellite. On the other hand, atmospheric drag is very different on such orbits. While negligible at GEO altitudes, a ΔV of about 10 m/s per year would be required at 500 km. This figure could increase to values as large as 100 m/s per year at 300 km. These are of course absolute perturbations, which would occur on all spacecraft on local orbits. Relative perturbations would be just a fraction of these values.

Fig. 10. Time to complete 10 km separation as a function of power level 5. CONCLUSIONS Current trends towards small satellite systems, including CubeSats, call for innovative solutions to the problem of miniaturization of propulsion systems. Chemical and cold gas thrusters are only viable when there is enough room left for payloads and other subsystems and ΔV requirements are not too demanding. Electrospray propulsion offers an alternative solution to the challenging problems of scaling down propulsion systems with high Isp. This type of devices are formed by microscopic elements that perform well at the smallest of scales. The main issue for these thrusters is scaling up to cover propulsion requirements of small, power-limited spacecraft. A particularly interesting application of this technology would be in CubeSats. Several thruster heads could be easily mounted on the surface of a CubeSat without interfering with other elements, including solar arrays. Using MEMS technologies, each thruster head could occupy volumes as small as 1 cm 3. Most of the mass and volume of the propulsion system would be that of the power processing unit. However, recent advancements in high-voltage electronics provide an avenue to scale the power subsystem to levels adequate for integration in CubeSats. Such a system would enable a number of maneuvers of potential interest for small satellite operators. 6. ACKNOWLEDGMENTS Support for this work has been provided by grants from the US Department of Defense. 7. REFERENCES [1] Khayms, V. Advanced Propulsion for Microsatellites, MIT PhD Thesis, Cambridge, MA, 2000. [2] Lozano, P., et al. Ionic Liquid Ion Sources: Characterization of Externally Wetted Emitters, Journal of Colloid and Interface Science, 282, 415-421, 2005. [3] Lozano, P., et al. Ionic Liquid Ion Sources: Suppression of electrochemical reactions using voltage alternation, Journal of Colloid and Interface Science, 280, 149-154, 2004. [4] Lozano, P.C. Energy properties of an EMI-Im ionic liquid ion source, Journal of Physics D: Applied Physics, Vol. 39 No. 1, 126-134, 2006.

[5] Lozano, P. Efficiency estimation of EMI-BF4 Ionic Liquid Electrospray Thrusters, AIAA-2005-4388, 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 2005. [6] Gassend, B., A Fully Microfabricated Two-Dimensional Electrospray Array with Applications to Space Propulsion, PhD Thesis, MIT, Cambridge, MA, 2007. [7] Legge, R., et al. Fabrication and Characterization of Porous Metal Emitters for Electrospray Thrusters, 30th International Electric Propulsion Conference IEPC-07-145, 2007. [8] Courtney, D., et al. Development of ionic liquid electrospray thrusters using porous emitter substrates, 27th International Symposium on Space Technology and Science, 2009.