Currently being updated – Available September 1st 2018
There are currently a wide range of technologies for propulsion systems, however the miniaturization of these systems for small spacecraft is a particular challenge. The purpose of this chapter is to identify and analyze the current status of the main propulsion technologies for small spacecraft and to present an overview of the available systems. Performance tests and technology demonstrations were considered in order to assess the maturity and robustness of each system. Some of the current systems are adaptable to a large variety of smaller buses.
While cold gas or pulsed plasma systems are targeted for small delta-v (dV) application, modules that can provide more demanding maneuvers need still development. Small spacecraft buses other than cubesats have more flexibility to accommodate systems with several thruster units to provide more attitude control and also large single axis maneuvers. Missions have demonstrated these technologies successfully and performance data gathered has paved the way for future modifications of the existing hardware in order to re-adapt the designs to satisfy demanding constraints. Table 4.1 shows a summary of the current state of the art for different propulsion methods.
|Product||Thrust||Specific Impulse (s)||Status|
|Hydrazine||0.5 – 4 N||150 – 250||TRL 6|
|Cold Gas||10 mN – 10 N||65 – 70||GN2/Butane TRL 9|
|Non-toxic Propulsion||0.1 – 27 N||220 – 250||HAN TRL 8, ADN TRL 6|
|Pulsed Plasma and Vacuum Arc Thrusters||1 – 1300 μN||500 – 3000||Teflon TRL 8, Titanium TRL 7|
|Electrospray Propulsion||10 – 120 μN||500 – 5000||TRL 6|
|Hall Effect Thrusters||10 – 50 mN||1000 – 2000||Xenon TRL 8, Iodine TRL 4|
|Ion Engines||1 – 10 mN||1000 – 3500||Xenon TRL 8, Iodine TRL 4|
|Solar Sails||0.25 – 0.6 mN||N/A||TRL 6 (85 m2), TRL 7 (35 m2)|
Electric and chemical systems have experienced a significant maturation process with respect to the previous report. Thrust stand measurements in vacuum and lifetime tests have been performed for an extensive variety of devices and a serious effort has been made by several companies, agencies and institutions to satisfy small spacecraft requirements. Fundamental components such as Power Processing Units (PPUs) and particular mass, power and volume constraints have been adjusted to smaller buses.
Hazardous propellants introduce handling and safety challenges and increase the total cost of the mission, while several non-toxic propellants also provide less safety and handling requirements and also higher specific impulse and density, which is beneficial for dV budgets. Electric propulsion devices have been miniaturized to successfully adapt to small buses and low thrust options for cubesats, such as electrosprays or Pulsed Plasma Thrusters (PPT), enable easy integration due to their low degree of complexity. For more ambitious mission concepts that require higher dV, technologies such as Hall Effect and ion system, are still being developed. Finally, in regards to propellant-less systems, the recent launch of LightSail has advanced the state of the art of solar sails for small spacecraft.
The TRL for small spacecraft propulsion is usually lower compared to other subsystems. This section considers systems that have been flown or are actively being developed in the last few years to account for the most recent advances in the technology. The chapter is divided in three main categories: chemical, electric and propellant-less systems, which are divided into smaller subsections depending on the type of thrust generation. Whenever pertinent, this report considers complete propulsion systems composed of thrusters, feed systems, propellant storage and Power Processing Units but not including electrical power supply. In addition, for some subsections, single thruster heads are also introduced.
State of the Art
Chemical Propulsion Systems
Chemical propulsion systems are designed to satisfy high thrust impulsive maneuvers. They are associated with lower specific impulse compared to their electric counterparts, but have significantly higher thrust to power ratios.
There are a significant number of mature hydrazine propulsion systems used in large spacecraft that present a generally reliable option as the characteristics in terms of mass and volume of these compact systems allow them to be a suitable fit for some small spacecraft buses. Thrusters that perform small correction maneuvers and attitude control in large spacecraft may be large enough to perform high thrust maneuvers for small spacecraft and can act as main propulsion system. Hydrazine propulsion systems typically incorporate a double stage flow control valve that regulates the propellant supply and a catalyst bed heater with thermal insulation. Typically, they have the advantage of being qualified for multiple cold starts which may be beneficial for power-limited buses if the lifespan of the mission is short. Typically, hydrazine achievable specific impulses are in the 150-250 s range. Because hydrazine systems are so widely used for large satellites, robust ecosystem components exist, and hydrazine propulsion systems are custom-designed for specific applications using available components. This section considers both commercial off-the-shelf (COTS) hydrazine thrusters and integrated complete propulsion systems.
Airbus Defense and Space has developed a 1 N class hydrazine thruster that has extensive flight heritage, including use on the small spacecraft, ALSAT-2. Aerojet Rocketdyne has leveraged existing designs with flight heritage from large spacecraft that may be applicable to small buses, such as MR-103 thruster used on New Horizons for attitude control application 1. Other Aerojet Rocketdyne thrusters potentially applicable to small spacecraft include the MR-111 and the MR-106 2.
The Cubesat High-Impulse Adaptable Modular Propulsion System (CHAMPS) project leverages the miniaturization effort performed for previous small hydrazine thrusters to develop cubesat monopropellant propulsion systems. These modules satisfy a wide range of maneuvers from station-keeping and orbit transfers to momentum management. There are various configurations, such as the MPS-120, that support up to four 1 N hydrazine thrusters configured to provide pitch, yaw, and roll control as well as single axis thrusting vector. Additional versions of the MPS series are under development that utilize various thruster technologies such as cold gas (MPS-110), non-toxic AF-M315E propellant (MPS-130) or electric propulsion devices (MPS-160) 2. The MPS-120 was selected and funded by NASA to go through extensive testing. The 3D printed titanium isolation and tank systems were demonstrated in mid-2014 and one engine performed a hot fire test in late 2014 3. Aerojet Rocketdyne is also developing integrated modular propulsion systems for larger small spacecraft. The MPS-220 consist of two 22 N primary engines and eight 1 N auxiliary thruster that use hydrazine as propellant 2.
Moog ISP has extensive experience in the design and testing of propulsion systems and components for large spacecraft. These may also apply for smaller platforms as some of their flight-proven thrusters are light-weight and have moderate power requirements. The MONARCH-5 thruster flew in NASA JPL’s Soil Moisture Active Passive (SMAP) spacecraft in 2015 and provided 4.5 N of steady state thrust. Other thrusters potentially applicable into small spacecraft buses include the MONARC-1 and the MONARC-22 series 4.
Non-toxic propellants are designated green as they have a reduced toxicity due to the lower danger of component chemicals and reduced vapor pressure as compared to hydrazine. This results in less safety requirements for handling these propellants, potentially including the lack of required SCAPE suit (self contained atmospheric protective ensemble) and reduces operational oversight by safety and emergency personnel. Fueling for these may also be considered a parallel operation or have a less significant required exclusionary zone, allowing for the acceleration of launch readiness operations. Non-toxic propellants are less likely to exothermicallly decompose at room temperature even in the presence of a catalyst. Therefore they require less inhibits resulting in less valve seats to power, including a less stringent temperature requirement with less system heater power. Non-toxic propellants also provide higher performance than the current state of the art fuel and have higher density-specific impulse achieving improved mass fractions. As a majority of these non-toxic propellants are in development, systems using these propellants present technical challenges including increased power consumption and the selection of materials due to higher combustion temperatures. The primary ionic liquid propellants with flight heritage or upcoming spaceflight plans are LMP-103S and AF-M215E. Table 4.2 lists the current state of the art in green propellants.
|Product||Manufacturer||Thrust (N)||Specific Impulse (s)||Status|
|GR-1||Aerojet Rocketdyne||0.26 – 1.42||231||TRL 6|
|GR-22||Aerojet Rocketdyne||5.7 – 26.9||248||TRL 5|
|1 N HPGP||ECAPS||0.25 – 1.00||204 – 235||TRL 8|
|HYDROS||Tethers Unlimited, Inc.||0.2 – 0.6||258||TRL 5|
The Ecological Advanced Propulsion Systems, Inc. (ECAPS) High Performance Green Propulsion (HPGP) system, shown in Figure 4.1, uses ammonium dinitramide-based LMP-103S as propellant. Its density is slightly higher than hydrazine (1.24 gcm-3 vs 1.02 gcm-3). The PRISMA mission incorporated successfully the 1 N version of this system in 2010. Furthermore, Skybox Imaging conducted a trade study of various propulsion alternatives for application in spacecraft constellations and selected this system 5. HPGP systems are being implemented in SkySat missions such as SkySat-3, and SkySat block-I. HPGP systems are currently developed for three different thrust magnitudes: 1 N, 5 N and 22 N, with higher thrust systems in development 6 7. VACCO partnered with ECAPS to design a self-contained unit that can deliver up to 1808 Ns of total impulse and can be adapted for different sizes, 0.5U to 1U. The VACCO/ECAPS Micro Propulsion System (MiPS) is designed to meet the specific cubesat standards and has four 100 mN ADN-propellant thrusters. Each engine is throttlable in order to have vector control. This unit has also an alternative hybrid version that incorporates one 100 mN ADN thruster and four 10 mN cold gas thrusters for attitude control, providing up to 1036 Ns of total impulse for main dV applications and 69 Ns for RCS 8.
Another non-toxic propellant in development is the U.S. Air Force developed AF-M315E, a hydroxylammonium nitrate (HAN) based monopropellant. Aerojet Rocketdyne is currently developing propulsion systems utilizing this propellant. The AF-M315E has a density of 1.47 gcm-3 (about 45% more than hydrazine) and a specific impulse of 230 – 250 s can be achieved by using this propellant. While some components have heritage from previous hydrazine systems, others that are compatible with AF-M315E propellant, such as valves and filters, are at TRL 6 9. The propulsion system will be flown as a technology demonstration on the NASA Green Propellant Infusion Mission (GPIM), scheduled to launch in 2016. This small spacecraft is designed to test the performance of this propulsion technology in space by using five 1 N class thrusters (GR-1) for small attitude control maneuvers 10. Aerojet completed a hot-fire test of the GR-1 version in 2014 and further tests in 2015. Initial plans to incorporate the GR-22 thruster (22-N class) on the GPIM mission were deferred in mid-2015 in order to allow for more development and testing of the GR-22. As a result, the GPIM mission will only carry 1-N class GR-1 units when launched in 2016 11. The TRL is currently 6 for the GR-1 (Figure 4.2), and 5 for the larger GR-22 (Figure 4.3).
The AF-M315E propellant is used by a 0.5 N thruster that is being developed by Busek. This device was placed on an inverted-pendulum type thrust stand for a test campaign. Three performance profiles were demonstrated: steady state, long and short duration pulses. For operating the thruster, there is a catalyst pre-heat requirement of 12 W for about eight minutes. In addition, the thruster is combined with a piezo-actuated micro-valve that is suitable for long-term propellant compatibility. While integrated system testing of the thruster and microvalve have occurred, further development is required before raising the TRL of the integrated system. The integrated testing demonstrated minimum impulse bits of 36 mN. A full duty cycle test of the whole system is included in future activities 12.
Tethers Unlimited, Inc. is developing a water electrolysis propulsion system called HYDROS, illustrated in Figure 4.4 that fits into 1U volume and uses water as propellant. On-orbit, water is electrolyzed into oxygen and hydrogen and these propellants are combusted as in a traditional bi-propellant thruster. This system is designed to be integrated into any cubesat configuration due to a modular nozzle and its injector design. A ground test campaign between Tethers Unlimited and the Air Force Institute of Technology (AFIT) measured thrust and specific impulse for the ½ version 13.
Cold and Warm Gas
Cold gas systems are relatively simple systems that provide limited spacecraft propulsive capability and are one of the most mature technologies for small spacecraft. Thrust is produced by the expulsion of an inert, non-toxic propellant which can be stored in high pressure gas or saturated liquid forms. Warm gas systems have been used in several missions for pressurization and use the same basic principle yielding more specific impulse performance than cold gas.
Warm and cold gases are suitable for small buses due to their very low grade of complexity and are inexpensive and robust. They can be used when small total impulse is required. Primary advantages include small impulse bit for attitude control applications and the association of small volume and low weight. Recently, new designs have improved the relatively high power requirement of these systems and there are currently thrusters that can be implemented into small buses such as 3U cubesats. Table 4.3 shows current state of the art for cold and warm gas propulsion systems that are small spacecraft applicable.
|Product||Manufacturer||Thrust||Specific Impulse (s)||Propellant||Status|
|MicroThruster||Marotta||0.05 – 2.36 N||65||Nitrogen||TRL 9|
|Butane Propulsion System||SSTL||0.5 N||80||Butane||TRL 9|
|MEMS||NanoSpace||0.01 – 1 mN||50 – 75||Butane||TRL 9|
|POPSAT-HIP1||Micro Space||0.083 – 1.1 mN||32 – 43||Argon||TRL 8|
|CNAPS||UTIAS/SFL||12.5 – 40 mN||40||Sulfur hexafloride||TRL 9|
|CPOD||VACCO||25 mN||40||R134a||TRL 6|
A cold gas thruster developed by Marotta (Figure 4.5) flew on the NASA ST-5 mission for fine attitude adjustment maneuvers. It incorporates electronic drivers that can operate the thruster at a power of less than 1 W. It has less than 5 ms of response time and it uses gaseous nitrogen as propellant 14.
Surrey Satellite Technology Ltd. (SSTL) has a butane propulsion system included in several small spacecraft missions for a wide range of applications in Low Earth Orbit (LEO) and Medium Earth Orbit (MEO). In this system, propellant tanks are combined with a resistojet thruster and operation is controlled by a series of solenoid valves, Figure 4.6. It uses electrical power to heat the thruster and improve the specific impulse performance with respect to the cold gas mode. It has been through more than five years of design life and it uses a RS-422 interface 15.
In June 2014, the Institute for Aerospace Studies in Toronto (UTIAS) launched two small spacecraft of 15 kg each to demonstrate formation flying. The Canadian Nanosatellite Advanced Propulsion System (CNAPS), shown in Figure 4.7, consisted of four thrusters fueled with liquid sulfur hexafluoride. This non-toxic propellant was selected since it has high vapor pressure and density which is important for making a self-pressurizing system. This propulsion module is a novel version of the previous NanoPS that flew in the CanX-2 mission in 2008 16.
Another recently flight-demonstrated propulsion system was flown in the POPSAT-HIP1 cubesat mission and was developed by Microspace Rapid Pte Ltd in Singapore. It consisted of a total of eight micro-nozzles that provided three rotation axes control and single-axis thrust for translational applications. The total dV has been estimated from laboratory data to be between 2.25 and 3.05 ms-1. Each thruster has 1 mN of nominal thrust by using argon propellant. An electromagnetic microvalve with a very short opening time of 1 ms operates each thruster 17.
NanoSpace has developed a complete Microelectromechanical systems (MEMS) cold gas propulsion system for cubesats (Figure 4.8) that provides accurate thrust control by using four thrusters with butane propellant. While thrust is controlled in a closed loop system with magnitude readings, each thruster can provide a thrust magnitude from zero to full capacity (1 mN) with 5 μN resolution. The dry mass of the system is 0.220 kg and average power consumption is 2 W during operation 18. This system is based on a flight-proven technology flown on the PRISMA mission, launched in 2010. Here, two thruster pods with four thrusters each were tested using Nitrogen as the propellant and each thruster provided up to 8 mN 19. Additionally, the MEMS cold gas system was included into the bus of the TW-1 cubesat, launched in September 2015 20.
The CubeSat Proximity Operations Demonstration (CPOD) is a mission led by Tyvak Nano-Satellite Systems. It incorporates a cold gas propulsion system built by VACCO Industries that provides up to 186 Ns of total impulse. This module operates at a steady state power of 5 W and delivers 40 s of specific impulse while the nominal thrust is 25 mN 8. It uses self-pressurizing R134a propellant to fire a total of eight thrusters distributed in pairs at the four corners of the module. It has gone through extensive testing at the US Air Force Research Lab. Endurance tests consisted on more than 70000 firings 21.
Solid rocket technology is typically utilized for impulsive maneuvers such as orbit insertion or quick de-orbiting. Due to the solid propellant, they achieve moderate specific impulses and high thrust magnitudes that are compact and suitable for small buses. There are some electrically controlled solid thrusters that operate in the mN range. These are restartable, have steering capabilities and are suitable for small spacecraft applications, unlike larger spacecraft systems that provided too much acceleration. Table 4.4 shows current state of the art in solid motors for small spacecraft. These thrust vector control systems can be coupled with existing solid rocket motors to provide controllable high dV in relatively short time. A flight campaign tested the ability of these systems to effectively control the attitude of small rocket vehicles. Some of these tests were performed by using state of the art solid rocket motors such as the ISP~30 s developed by Industrial Solid Propulsion 22.
|Product||Manufacturer||Total Mass (kg)||Average Thrust (N)||Specific Impulse (s)||Status|
|ISP 30sec motor||Industrial Solid Propulsion||0.95||37||187||TRL 7|
|STAR 4G||Orbital ATK||1.5||258||277||TRL 6|
SPINSAT, a 57 kg spacecraft launched in 2014, incorporated a set of solid motors (Figure 4.9 and Figure 4.10) which were part of the attitude control system and were developed by Digital Solid State Propulsion LLC (DSSP). The system was based on a set of Electrically Controlled Solid Propellant (ESP) thrusters that consisted of two coaxial electrodes separated by a thin layer of electric solid propellant. This material is highly energetic but non-pyrotechnic and allow for better burn control. They are only ignited if an electric current is applied, the lack of moving parts and duration control make the system suitable for small spacecraft.
In total, 72 thrusters formed the propulsion system of the spacecraft. Six of them were included in each of the twelve plugs strategically located around the bus. Performance characterization is done by firing the thrusters in pairs and measuring the changes in the spin rate by both on-board and on-ground assets 23.
Electric Propulsion Systems
Electric propulsion has experienced significant improvement in terms of available systems and maturity of components. For many small spacecraft concepts, high specific impulses are necessary to comply with dV budgets. Depending on thruster technology, specific impulse for electric propulsion can range between 700-3000 s. However, thrust is low meaning long maneuver times. Some thrusters are more suitable for small correction maneuvers and attitude control applications due to low impulse bits while others are designed to achieve high accelerations for interplanetary spiral trajectories. A wide spectrum in propellants is offered with electric propulsion. Iodine is proposed for some technologies due to its very high density that allow high dV maneuvers for transfer trajectories. For smaller dV applications, solid state materials such as polytetrafluoroethylene (PTFE) are used in most Pulsed Plasma Thrusters (PPTs) while electrosprays use various forms of ionic liquid.
Resistojets are the simplest form of electric propulsion. Thrust is produced by heating the propellant by electrical means so that the resulting gas can be expanded and expelled at large velocities out of the nozzle. Table 4.5 lists current state of the art Resistojets designs that are small spacecraft applicable.
|Product||Manufacturer||Thrust||Power (W)||Specific Impulse (s)||Status|
|Low Power Resisojet||Busek||13 mN at 200 W||100- 300||1390||Xenon TRL 8, Iodine TRL 4|
|PUC||CU Aerospace and VACCO||0.45 N||Unkn.||70||TRL 6|
Surrey Satellite Technology Ltd. (SSTL) has developed a resistojet propulsion system that has flown in several missions. It can work with different types of propellant such as xenon, butane or nitrogen. Thrust can be up to 100 mN and the specific impulse varies with the selected propellant ranging from 48 s for xenon to 99 s for nitrogen. The system uses power from 30 to 50 W and does not require a PPU since it works directly from the bus voltage input.
CU Aerospace and VACCO have built a Propulsion Unit for Cubesats (PUC), see Figure 4.11. It consist of a full integrated system that includes controller, PPU, valves, sensors and a Micro-Cavity Discharge (MCD) thruster. High density and self pressurizing liquids are used as propellants by using the MCD heating technology together with an optimized low mass flow nozzle 24.
CU Aerospace and VACCO Industries have also developed a Cubesat High Impulse Propulsion System (CHIPS). This module incorporates a main micro-resistojet plus four equally distributed cold gas thrusters acting as a 3-axis attitude control system. By leveraging VACCO’s compact friction-less valve technology and utilizing an inert and non-toxic R-134a propellant, this system achieves a high total impulse to volume ratio. It occupies a 1U+ space in order to target 2U and 6U spacecraft buses. A fully integrated system with flow and power control has been demonstrated at the Electric Propulsion Laboratory at the University of Urbana-Champaign, Illinois. Tests included thrust and specific impulse measurements that estimated 82 s for the warm fire mode and 47 s for the cold fire mode. It can provide up to 563 Ns of total impulse and a throttleable thrust of 30 mN in warm fire mode, which is used for primary propulsion. The cold gas mode is used for the three axis attitude control and provides 323 Ns of total impulse and 19 mN of thrust. The TRL of the integrated system is 5 and a second phase is currently in development 25.
Busek Co, Inc. has leveraged previous flight and design efforts to miniaturize fundamental components such as valves and PPUs for a micro-resistojet. This system uses non-toxic ammonia propellant and delivers a total impulse of 404 Ns for main dV applications and 23 Ns for ACS 26.
Electrospray propulsion systems use the principle of electrostatic extraction and acceleration of ions from a propellant consisting of a negligible vapor pressure conductive salt. One of the biggest advantages of this technology with respect to other traditional electric propulsion systems is that no gas-phase ionization is required. The propellant does not need to be pressurized for storage since it flows via capillary action due to the ion evaporation process. The emission can be controlled by modulation of the voltage input in a closed loop feedback with current measurements. In addition, in some cases, both species of negative and positive ions can be utilized, avoiding the need for a neutralizer which simplifies the design and operation of the system. Expelled ions achieve very high velocities which translates into high specific impulse. Typically, the most widely used propellant in electrosprays is the ionic liquid 1-Ethyl-3-Methyl-Imidazolium Tetrafluoroborate (EMI-BF4). NASA’s Advanced In-Space Propulsion (AISP) project has created a portfolio that includes the development of Microfluidic Electrospray Propulsion (MEP). Table 4.6 displays the current state of the art for small spacecraft applicable electrospray thrusters.
|Product||Manufacturer||Thrust||Power (W)||Specific Impluse (s)||Status|
|S-iEPS||MIT||74 μN||1.5||1160||TRL 6|
|IMPACT||Accion Systems Inc.||60 μN per axis||0.75 per axis||1200||TRL 5|
|MAX-1||Accion Systems Inc.||120 μN||1.6||2000||TRL 5|
|1 mN Electrospray||Busek||0.7 mN||15||800||TRl 5|
|100μ||Busek||0.1 mN||5||2300||TRl 5|
Electrospray technology has been advanced significantly at the The Massachusetts Institute of Technology (MIT) Space Propulsion Laboratory (SPL) and some companies have started to commercialize systems based on this effort. Figure 4.12 is the Electrospray thruster developed at MIT. Voltage versus current curves and time of flight spectroscopy among other tests have helped to understand the ionic and electrical characteristics of the thruster. MIT has demonstrated a total of 315 hours of continuous electrospray operation and a magnetically levitated thrust balance was used to measure
thrust at μN levels 27. Each thruster has a total of 480 emitters, a passive propellant management system that includes a 1.2 cm-3 tank and an acceleration chamber. At the system level, MIT has developed the Scalable ion Electrospray Propulsion System (S-iEPS), shown in Figure 4.13, that features a total of eight thrusters that fire along a single axis. This module is able to provide 74 μN and more than 1160 s of specific impulse at a power draw of less than 1.5 W. It is light weight, about 0.095 kg including PPU, and fits in a 0.2U volume 28.
Fully integrated electrospray systems, designed mainly for cubesat applications, are being developed by Accion Systems. IMPACT and MAX-1 are two different complete electrospray modules that have been through thrust measurements and lifetime and efficiency tests. IMPACT offers thrust in one direction and also 2-axis attitude control, has a wet mass of 0.5 kg and provides a total impulse of 45 Ns per axis. MAX-1 provides single-axis thrust, has a wet mass of 0.3 kg and a total impulse of 86 Ns 29.
Busek Inc. is developing a fully integrated electrospray propulsion system in the mN range. This module includes a propellant-less cathode neutralizer and a low pressure customizable tank that were leveraged from the module incorporated into the NASA ST-7/ESA LISA Pathfinder spacecraft. The system uses 15 W of power and provides 675 Ns with 50 mL of propellant and has a mass of 1.15 kg. Tests in relevant conditions are being performed to raise the TRL from 5 to 6 26. The system features a 100 μN class thruster that provides a specific impulse of 2300 s and consumes 5 W. It can deliver 85 ms-1 to a 4 kg cubesat by having a wet mass of 0.320 kg and 10 mL of an ionic liquid propellant that has been fully characterized during the ST-7 flight program 26.
The Micro Devices Laboratory (MDL) at the Jet Propulsion Laboratory (JPL) has developed a highly integrated and scalable indium MEP system (Figure 4.14) that has a dry mass of less than 0.010 kg and provides thrust in the 20-100 μN range. Indium metal is stored in solid form and heated afterwards to be used as propellant. Over 10 hours of continuous operation tested an initial prototype assembly 30.
In ion thrusters, propellant is ionized by using various plasma generation techniques. Radio Frequency (RF) engines achieve thrust by producing ions with electrode-less inductive discharges that are typically achieved by using a helical coil at frequencies in the range of 1 MHz. The particles are then accelerated at very high exhaust velocities by electrostatic grids. These devices have a high efficiency when compared to other electric propulsion systems. In addition, the absence of electrodes avoids potential threats to thruster lifetime which is only limited by grid erosion. Table 4.7 displays the current state of the art ion engines for small spacecraft.
|Product||Manufacturer||Thrust||Power (W)||Specific Impulse (s)||Propellant||Status|
|BIT-3||Busek||1.4 mN||60||3500||Xenon-Iodine||TRL 5|
|BIT-1||Busek||0.1 mN||10||2250||Xenon||TRL 5|
|1-COUPS||University of Tokyo||0.3 mN||N/A||1000||Xenon||TRL 9|
|RIT-μX||Airbus||50 – 500 μN||50||300 – 3000||Xenon||TRL 5|
Busek is developing a RF ion thruster that can operate with both xenon and iodine propellants, achieving similar performances 31. The BIT-3 engine has 3 cm diameter grids and is capable of providing variable specific impulse and thrust. At 60 W of operating power, it can achieve an efficiency of 35%. Recent test performance results on the iodine version have shown that thrust-to-power ratios are similar to the ones achieved with xenon as propellant. Complementary technology associated with the thruster such as propellant tanks and feed system have been demonstrated as well for this propellant. The compatibility with iodine is made possible since the plasma-generation chambers in RF engines are generally built with ceramic materials that are resistant to corrosion. A lower TRL, smaller thruster version of just 1 cm grids, the BIT-1, is also under development by Busek 31.
Recently, the Japanese Proximate Object Close flyby with Optical Navigation (PROCYON) mission has shown successful operation of a propulsion system in Space. The Ion thruster and Cold-gas thruster Unified Propulsion System (I-COUPS) was designed at the University of Tokyo and is an integrated system comprised of two sets of ion and cold gas thrusters. Both technologies share the same gas feed system that provides xenon to be used as propellant. This combines high thrust and large dV capabilities. Cold gas thrusters are used for reaction wheel de-saturation and small correction burns, while ion engines are kept for deep space maneuvers. In total, the mass of the propulsion system is less than 10 kg, including propellant. The ion engines in the I-COUPS unit are an evolution of the Miniature Ion Propulsion System (MIPS), which was previously launched on board the Hodoyoshi-3/4 mission in October 2014. This spacecraft was placed on a Sun Synchronous Orbit and had 65 kg of mass. The MIPS had a wet mass of 8.1 kg with 1 kg of propellant mass. Ion thruster operation was proven by providing continuous acceleration 32.
Airbus is developing a family of RF ion thrusters over the last few years and has designed the RIT-μX (Figure 4.15) for small spacecraft buses and for high precision maneuvers. Various thrust configurations were proposed and tested. In 2013, a system in the 50-500 μN range was demonstrated and thrust resolution, linearity, response and noise met LISA Pathfinder mission requirements, which increased the TRL to 5. The nominal power to operate is less than 50 W and the specific impulse is between 300 and 3000 s, depending on the configuration. The maximum demonstrated specific impulse was 3500 s. It uses xenon as propellant and it has a dry mass of 0.440 kg 33.
Pulsed Plasma and Vacuum Arc Thrusters
In Pulsed Plasma Thrusters (PPTs), thrust is produced by triggering a high voltage discharge between two electrodes that result in an electric arc that typically ablates a solid state material. A self-generated magnetic field is produced and then accelerates and expels particles from the thruster head. Typically the propellant is pushed forward by a spring as it is being consumed. This technology has significant heritage from larger spacecraft versions and due to its simplicity, miniaturization was more achievable compared to other electric propulsion systems. Major problems such as electrode shortcuts or non-uniform propellant ablation are under active research. These systems are suitable for attitude control and fine pointing applications since the trigger pulse of the discharge can be adjusted, small impulse bits can be achieved that allow for high precision. Typically the propulsion system consists of just a PPU that controls the required discharge to operate the thrusters by storing energy in a capacitor bank, which accounts for a significant portion of the system mass. Various materials have been tested for PPT utilization, however, PTFE is the industry standard. Table 4.8 accounts for current small spacecraft applicable state of the art PPT thrusters.
|Product||Manufacturer||Thrust||Power (W)||Specific Impulse (s)||Propellant||Status|
|PPTCUP||Mars Space and Clyde Space||40 μN||2||655||PTFE||TRL 6|
|NanoSat PPT||Mars Space and Clyde Space||90 μN||5||640||PTFE||TRL 5|
|μ-CAT||GWU and USNA||1 – 50 μN||2 – 14||2500 – 3000||Titanium||TRL 7|
|BmP-220||Busek||20 μN-s Impulse bit||1.5||536||PTFE||TRL 5|
|MPACS||Busek||80 μN-s Impulse bit||10||827||PTFE||TRL 8|
Mars Space Ltd. and Clyde Space Ltd. have developed a compact propulsion module (Figure 4.16) specifically designed to provide maneuvering capabilities to cubesats. At the University of Southampton, thermal cycling, vibration, Electro Magnetic Compatibility (EMC) and lifetime tests were performed. Vibration test results showed that the module sustains the mechanical vibrations during launch and Electro-Magnetic (EM) noise levels during discharge were mostly compliant with guidelines. The system has a total mass of 0.270 kg and is characterized by an average specific impulse of 655 s and a total impulse of 48.2 Ns. It has a single thruster that uses PTFE propellant and is side-fed to maximize discharge length, with an electrode design that minimizes carbonization 34.
Busek has extensive experience in the development of pulsed plasma propulsion. Their Micro Pulsed Plasma Altitude Control System (MPCAS) flew on the FalconSat-3 mission in 2007. This module consisted of eight thrusters and provided attitude control with precise impulse bits of 80 μNs at moderate power of less than 10 W 35 by using PTFE propellant. The system had heritage from previous investigations conducted at the Air Force Research Laboratory (AFRL) 36 and has been evolving since this first approach. The BmP-220 is the latest version of the Busek PPT family, consumes less than 7.5 W, weighs less than 0.5 kg and all required components fit in a 10x10x7 cm volume, see Figure 4.17. It can provide up to 220 Ns of total impulse with 0.040 kg of propellant. An innovative solid state switching technology enables the implementation of several emitters in a single unit. The specific impulse is 536 s and the minimum impulse bit is 0.02 mNs. The system TRL is estimated to be 5 26.
Vacuum arc thrusters are another type of plasma-based propulsion device that produces thrust by propellant ionization. This technology consist of two metallic electrodes separated by a dielectric insulator. One of them is used as solid metallic propellant and it is consumed as the thruster operates. Advantages of using a metallic solid propellant over the more traditional option of PTFE are a lower energy consumption per ionized mass, high pulse stability and higher repetition rates due to the thermal properties of metals.
The Micro-Cathode Arc Thruster (μCAT) developed by The George Washington University (GWU), uses vacuum discharges to ablate the cathode material. It consists of a 5 mm thruster head that contains concentrically aligned and cylindrically shaped anode, cathode and insulator. By sending a pulse created by the PPU to the electrode interface, a high voltage arc is produced across it 37. The μCAT offers a quasi-perfect ionization degree of the plasma particles in the exhaust plume, giving a near zero back flux. This propulsion technology generates thrust by consuming cathode material made of titanium with a high voltage vacuum arc, producing highly ionized plasma jets with high exhaust velocities. In addition, the incorporation of an external magnetic coil improves significantly the capabilities of the thruster 38.
An autonomous and modular micro electric propulsion system based on this technology has been designed and built at NASA Ames Research Center in partnership with GWU. This module fits into a 0.2U volume and consists of one Printed Circuit Boards (PCB) that command and operate up to four vacuum arc thrusters. Two PPUs, implemented in the main PCB, create the necessary discharges to operate the thruster that have an average thrust in the μN range which is controlled by selecting different thrusting frequencies. This system was tested and measured in relevant conditions of vacuum at NASA Glenn Research Center with a high accuracy torsional thrust stand.
Furthermore, a partnership between GWU and The United States Naval Academy resulted in the integration of a μCAT propulsion system into the Ballistically Reinforced Communication Satellite (BRICSAT). This mission was launched in May of 2015 and consisted of four PPUs to operate four thrusters in total. Preliminary retrieved data has shown that the system successfully accomplished the objective of detumbling the spacecraft. After two days, the propulsion system was able to reduce the initial tumbling from 30°s-1 to nearly 1.5°s-1, increasing the TRL of this system from 6 to 7 39.
Hall Effect Thrusters
Hall Effect propulsion is a mature technology for large spacecraft systems. Miniaturization of some of the components, such as neutralizers, is complicated to achieve and power consumption is relatively high compared to other electric propulsion technologies. However, an improvement has been made to integrate complete Hall Effect propulsion systems that can potentially support large transfers for interplanetary missions. See Table 4.9 for current state of the art technology in Hall Effect Thrusters for small spacecraft.
|Product||Manufacturer||Thrust (mN)||Power (W)||Specific Impulse (s)||Status|
|BHT-200||Busek||13||200||1390||Xenon TRL 8, Iodine TRL 4|
|HT100||SITAEL||5 – 15||175||<1350||Xenon TRL 6|
|CHT||UTIAS SFL||6.2||200||1139||Xenon TRL 5|
Busek has developed a complete Hall Effect thruster propulsion system for small spacecraft. The BHT-200, shown in Figure 4.18, is suitable for small spacecraft buses of relatively high mass and power supply since it needs 100-300 W to operate. This system has flight heritage from the 2006 TacSat-2 mission, and was part of the payload in the FalconSat-5 mission in 2010. In addition, it will be flown in the FalconSat-6 mission, scheduled for 2016. This model can operate with multiple propellants 26. The utilization of iodine will advance the technology due to its increased density over xenon and its lower operating pressure, which reduces cost and risk implications. More details can be found in the On the Horizon section.
The HT100, developed by Sitael Aerospace, has been extensively tested through campaigns that include characterization under thermal-vacuum conditions and structural analysis under heavy loads. Erosion has been observed in an endurance test that lasted for 1650 hours where no thermal problems or important performance reduction was observed. The nominal operation power at 175 W gives a thrust range of 5-15 mN. The thruster mass is 0.440 kg, it utilizes xenon as propellant and it can achieve a peak total efficiency of up to 35% and a maximum specific impulse of 1350 s. The HT100 has been selected for an in-orbit validation program by the European and Italian space agencies. A larger version, the HT400, operates at a nominal power of 400 W and it is at TRL 5 40.
The Space Flight Laboratory (SFL) at the University of Toronto is developing a low power cylindrical Hall thruster (Figure 4.19) that operates below 200 W and has a diameter of 26 mm for the ionization chamber. The cylindrical geometry of the ionization chamber was chosen in order to overcome the challenges of the annular chamber of traditional Hall thrusters. With this configuration, better efficiencies can be achieved while maintaining a sufficient thrust magnitude between 2.5-12 mN. Annular ionization chambers are mechanically simpler and produce high thrust to power ratios that are beneficial for small spacecraft applications. However, the efficiency still gets reduced when this chamber gets redesigned to optimize low power operation.
Excluding cathode, the weight of the first prototype was 1.6 kg. This device went under magnetic characterization and performance tests in vacuum. It uses xenon as a baseline propellant due to its improved performance over other gases such as argon. Further testing and design modifications will be done in order to raise the TRL from 5 to 6 in early 2016 41.
Systems that do not carry propellant for thrust generation are an ideal candidate for small spacecraft. They avoid complexity and reduce mass limitations. They can achieve high accelerations that can potentially propel an object for interplanetary travel.
Solar sails are the most popular method of propellant-less propulsion. They take advantage of solar radiation pressure by reflecting photons on a large sail made of a highly reflective material. Several missions have been conducted to demonstrate this technology for large buses such as the Japanese IKAROS, launched in 2010. Regarding small spacecraft, NASA has been conducting extensive research that resulted in the launch in 2010 of NanoSail-D2, a technology demonstration mission managed and designed by NASA Ames Research Center and NASA Marshall Space Flight Center. The sail had a deployed surface area of 10 m2, was made of a thin highly reflective material called CP-1 and weighted 4.2 kg 42.
One of the most recent solar sail mission for small spacecraft was performed by The Planetary Society in 2015. The 3U LightSail-A spacecraft completed its technology demonstration test in Space by fully deploying a solar sail in LEO. The dimensions were 5.6 m on a side and 32 m2 of total area once it was deployed. In 2016, a follow up mission, LightSail-B will complete the project by leveraging the experience acquired in the first flight and perform additional maneuvers. This spacecraft will fly in a Falcon heavy rocket to an approximately 720 km LEO orbit, where an orbital change in altitude or inclination will be performed 43.
On the Horizon
As propulsion technology matures, more small spacecraft missions will incorporate propulsion systems on board allowing for more complex mission architectures. This section will cover near-term spacecraft with propulsion as well as promising technologies that will become an important propulsion asset for future missions.
The Cubesat Ampibolar Thruster (CAT) is a novel device developed by the University of Michigan that utilizes a magnetic helicon discharge to ionize the propellant. The thruster (Figure 4.20) does not require a separate electron source and no resultant magnetic dipole is produced. High plasma density is created through a high efficiency helicon RF source and a large accelerating electric field is achieved. A large variety of propellants in solid or liquefied state can be used thanks to the electrode-less design of the thruster. Iodine has been presented as the most promising propellant due to its low cost and high storage density. This system can achieve an estimated specific impulse of 1010 s when using iodine. Currently, the PPU is still in development phase and some of the components for iodine utilization are at TRL 3 44. Initial tests were performed by using both xenon and argon as propellant. For xenon, CAT was designed to operate on 10-50 W in order to address some of the power limitations that small spacecraft face. In this configuration, the TRL is 4 and thrust and specific impulse are in the 0.5-4 mN range and the 400-800 s range respectively 45. The company Phase Four LLC is developing an integrated flight unit of the CAT 46.
There are several other propulsion technologies currently being developed: Ventions LLC is working on an integrated 3U cubesat propulsion system using non-toxic propellant; hybrid non-toxic/cold gas propulsion system for 6U and 12U spacecraft by Planetary Resources Development Corporation; and a non-toxic solid rocket for cubesats that allows for second ignition and utilizes an aluminized version of an Electric Solid Propellant (ESP) from Digital Solid State Propulsion (DSSP). ESPs provide more safety for handling compared to traditional solid energetic propellants and are electrically ignited 47.
Orbital Technologies Corporation (ORBITEC) is developing the Miniature Nontoxic Oxide-Propane (MINNOP) propulsion system. It consists of a bipropellant system for small spacecraft that can provide a significant increment in specific impulse performance with respect to hydrazine systems when used in bi-propellant mode and small levels of minimum impulse bit when used in cold gas mode. Current efforts are pointed towards the demonstration of the bipropellant thrust chamber and ignition within suitable weight constraints in order to fit into a 1U form factor 47.
Another high-performance propellant is a nitrous oxide fuel blended mono-propellant known as NOFBX developed by Firestar technologies. This self pressurizing non-toxic propellant can offer more than 320 s of specific impulse and it provides 3.5 to 3.9 times higher specific energy density than hydrazine 48.
The Inductively Coupled Electromagnetic (ICE) thruster is a novel technology that is being developed by MSNW LLC. This system uses a small integrated RF oscillator to generate plasma. The total volume of the thruster and the PPU is expected to be less than 0.125\,U. One of the main advantages is that this system can virtually use any liquid propellant. Anticipated operating power is between 10-50 W. The current goal is to achieve TRL 4 49.
An experimental characterization of a low power helicon thruster has been performed at the Stanford University’s Plasma Physics Laboratory. Tests were conducted by operating on water and argon propellants. Thrust was observed at various performance levels, achieving magnitudes of 2-5 μN. Future work in on-going and include optimization for greater performance and thrust stand measurements 50.
NanoAvionics JSC is developing a non-toxic mono-propellant propulsion system. It uses ADN as propellant and gives 252 s of specific impulse. Current efforts are focused on the miniaturization of a catalyst bed heater system and development of fuel feeding equipment. This module will be ready for flight in the LituanicaSAT-2 3U cubesat as part of the European QB50 initiative and is currently TRL 4.
The Mechanical and Aerospace Engineering Department at Utah State University has built and tested a non-toxic 22 N thruster for small spacecraft. This unit uses innovative propellants: compressed gaseous oxygen and ABS plastic. Additive manufacturing is used to build various system components such as the nozzle or the fuel grain. The system is restartable and can be throttled from 1 N to 22 N while maintaining performance and robustness. The achieved laboratory specific impulse was above 230 s 51.
Princeton Plasma Physics Laboratory, with The Aerospace Corporation, have tested the performance of a small Cylindrical Hall Thruster with permanent magnets. The measured thrust was in the 3-6.5 mN range with a specific impulse of 1000-1900 s. Efficiency studies were also conducted at a discharge voltage of 300 V achieving a maximum thruster efficiency over 20%. This version demonstrated even superior performance in comparison to another version that utilizes electromagnets coils 52.
Future Small Spacecraft Missions with Propulsion
Due to the significant improvement in propulsion technologies, mission concepts that were previously limited to large spacecraft are now possible with small buses. Interplanetary missions are becoming less costly, and therefore several institutions are assuming more risks to perform science missions with higher payoffs. As an example, NASA’s Exploration Mission (EM-1) is going to be used to provide secondary payload opportunity for up to eleven 6U cubesat. The mission trajectory would provide access to deep space or a Moon orbit.
The iodine satellite (iSAT) mission, a partnership project between NASA Marshall Space Flight Center, Busek Co. Inc. and NASA Glenn Research Center, consists of a 12U cubesat in a high performance integrated bus configuration that will perform propulsive inclination and altitude plane changes. This spacecraft will include a Busek’s BHT-200-I propulsion system with iodine propellant that offers a similar performance than the xenon version when operating at the same power level. It is expected to be delivered for launch in the second quarter of 2017 53 and an 80 hour endurance test of the engineering model has been performed at NASA Glenn Research Center. The objective was to characterize the performance of the thruster over the throttling range, to demonstrate feed system components and to study the plume and thermal models during operation 54.
NASA Ames and Glenn Research Centers are working on the Pathfinder Technology Demonstration (PTD) project which consists of a series of 6U cubesats that will be launched to test the performance of new subsystem technologies on orbit. For the first flight version, various state of the art electrospray systems, previously discussed in the Electric Propulsion section, are being considered.
JPL is developing the InSight mission which is going to be launched in March 2016 that will incorporate two identical cubesats as part of the Mars Cube One (MarCO) technology demonstration. These spacecraft will need to perform up to five Trajectory Correction Maneuvers (TCMs) during the mission to Mars. These cubesats include an integrated propulsion system, developed by VACCO Industries, that contains four thrusters for attitude control and other four for the TCMs. The module uses cold gas R-236FA as propellant, produces 755 Ns of total impulse and weighs 3.49 kg 55.
A team at Purdue University and NASA Goddard Space Flight Center is developing the Film Evaporation MEMS Tunable Array (FEMTA). This technology consists of a group of nozzles made of high aspect ratio slots. Each nozzle produces thrust by applying local heat to a propellant capillary interface and the main advantages are the absence of any mechanisms and a low power consumption, in the order of mW. Up to three generations of these devices have been built and improved over time. Vanadium and platinum heaters were used for the most updated version and thrust, propellant and mass flow rate response have been characterized. Thrust levels from 15 to 600 μN were observed at less than 100 mW of input power while specific impulse ranged from 5 to 40 s. Repeatable thrust pulses were consistent in magnitude and could be controlled. This system is a promising option for attitude control and small maneuver applications in cubesats 56.
Two separate 3U cubesats are part of the Interplanetary NanoSpacecraft Pathfinder In a Relevant Environment (INSPIRE) mission. These spacecraft will be placed in an Earth escape trajectory in order to test the performance of the communication, navigation and operations segments in deep space. A cold gas system developed by the University of Texas, Austin, has been included that utilizes the additive manufacturing techniques that were previously used for the MiPS cold gas module. The MiPS flew in the MEPSI-3 mission from the Aerospace Corporation 57,58. Further research was conducted by UT Austin to redefine the 3D printing process to adapt the system to the Bevo-2 and ARMADILLO mission concepts. The additive manufacturing process allows the fabrication of complex features in small volumes and a saturated liquid propellant is released through a converging-diverging nozzle in order to produce thrust. Tests have measured specific impulse in the range of 65-89 s and thrust in the range of 110-150 mN across different temperatures 59. A version of this propulsion unit will be used for attitude control maneuvers and nominal flight operations for the INSPIRE mission 60.
NEA Scout and Lunar Flashlight are two NASA JPL missions that are going to be launched as part of EM-1, scheduled for 2018. Both of the interplanetary 6U cubesats will deploy an identical sail of 80 m2 of area with 0.0601 mms-2 of characteristic acceleration. NEA Scout solar sail will be used as a main propulsion system whereas Lunar Flashlight sail will be used mainly for station keeping and to reflect light into selected lunar craters once in orbit. The duty cycle for NEA Scout is 90% while for Lunar Flashlight is 85% 61.
A significant variety of propulsion technologies are currently available for small spacecraft. While cold gas and pulsed plasma thrusters present an ideal option for attitude control applications, they have limitations for more ambitious maneuvers such as large orbital transfers. Other alternatives such as hydrazine, non-toxic propellants and solid motors provide a high capability and are suitable for medium size buses and missions that require higher dV budgets. Some spacecraft have already flown with these systems or are being scheduled to fly in the next year. For the near future, the focus is placed on non-toxic propellants that avoid safety and operational complications and provide sufficient density and specific impulse. The application of this technology in cubesats is still in development as some of the components need to be scaled down to comply with volume, power and mass constraints.
Electrosprays, Hall Effect thrusters and ion engines are in an active phase of development and active testing and technology demonstrations are expected for different bus sizes. These propulsion technologies will allow spacecraft to achieve very high dV and, therefore, to perform interplanetary transfers with low thrust.
Several other technologies, as well as new versions of existing systems with improved capabilities, are being proposed and a wide range of mature options in the following years are forecasted. As the industry progresses and more launches are scheduled, more propulsion systems will be included on board small spacecraft , increasing the average TRL for this subsystem.
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