04. Propulsion

Introduction

There are currently a wide range of technologies for propulsion systems, however the miniaturization of these systems for small spacecraft has been a particular challenge. The purpose of this chapter is to identify and analyze the current developmental status of 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. Since the last edition of this report in 2015, there have been several flight demonstrations of different forms of smallsat propulsion, however given that propulsion systems are still not standard COTS components for CubeSat for factors, most of these systems are still in prototype version and the flight test serve as flight demonstration of operational design thus reaching a TRL of 7 and not 9.

While cold gas or pulsed plasma systems are targeted for small delta-v (ΔV) application, modules that can provide more demanding maneuvers still need 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.

Table 4-1: Propulsion Systems Types for Small Spacecraft
Product Thrust Specific Impulse (s) TRL Status
Hydrazine  0.5 – 30.7 N  200-235  7
Cold Gas 10 mN – 10 N 65 – 70 GN2/Butane 9
Non-toxic (Green) Propulsion 0.1 – 27 N  220 – 250  HAN 6, ADN 7
Pulsed Plasma and Vacuum Arc Thrusters  1 – 1300 μN 500 – 3000 Teflon 7, Titanium 7
 Electrospray Propulsion 10 – 120 μN  500 – 5000 7
Hall Effect Thrusters 10 – 50 mN  1000 – 2000 Xenon 7, Iodine 7
Ion Engines 1 – 10 mN 1000 – 3500 Xenon7, Iodine 4
 Solar Sails 0.25 – 0.6 mN N/A 6 (85 m2), 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.

While hazardous propellants introduce handling and safety challenges and increase total mission cost, several non-toxic (green) propellants are associated with a reduced amount of less safety and handling requirements. These non-toxic propellants also have a higher specific impulse and density which is beneficial for ΔV 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 ΔV, technologies such as Hall Effect and ion system, are still being developed. Finally, in regards to propellant-less systems, the 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 as this has not generally been a common smallsat (especially CubeSat) system. There are no COTS CubeSat propulsion systems yet as many of the flight demonstrated modules are still being updated for improved performance. 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. Development on propulsive modules used for deorbit maneuvers, like solar sails, will be addressed.

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.

Hydrazine propellant

There are a significant number of mature hydrazine propulsion systems used in large spacecraft that present a generally reliable option as 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. 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.

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. 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.

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.

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 MONARCH-1 and the MONARCH-22 series 4. All of these MONARCH thrusters have extensive flight heritage and are TRL 9.

Non-Toxic (Green) Propellant

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 exothermically 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 Ammonium DiNitramide (AND) based LMP-103S and AF-M215E, and AF-M315E, a hydroxylammonium nitrate (HAN) based monopropellant. Table 4.2 lists the current state of the art in green propellants

Table 4-2: Green Propulsion Systems
Product AND or HAN based Propellant Manufacturer Thrust (N) Specific Impulse (s) TRL Status
GR-1 HAN Aerojet Rocketdyne 0.26 – 1.42 231 6
GR-22 HAN Aerojet Rocketdyne 5.7 – 26.9 248 5
1 N HPGP AND Bradford Engineering 0.25 – 1.00 204 – 235 7
HYDROS-C Other Tethers Unlimited, Inc. 1.2 310 6
AMAC Other Busek 0.425 225 5
MiPS ADN VACCO 0.4 250 7
BGT-X5 HAN Busek 0.5 220 5
EPSS C1K ADN NanoAvionics 0.3 252 7
Green Hybrid Other Utah State 8 215 5
ECAPS
Figure 4.1: ECAPS HPGP thruster. Image Courtesy of SSC ECAPS

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 Bradford Engineering (formally ECAPS) to design a self-contained unit that can deliver up to 1808 N-s of total impulse and can be adapted for different sizes, 0.5U to 1U. The 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 N-s of total impulse for main ΔV applications and 69 N-s for RCS8. There are several upcoming opportunities for this module to be flight-proven: A hybrid MiPS system is being developed for ArgoMoon nanosatellite program that is planned to launch 2020 with EM-1 as well as four MiPS thrusters will be flown on Lunar Flashlight, a 3U. The thrusters for this mission underwent qualification testing June 2018.

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 1 N class (GR-1) thruster from Aerojet Rocketdyne will be flown as a technology demonstration on the NASA Green Propellant Infusion Mission (GPIM), scheduled to launch November 2018. This small spacecraft is designed to test the performance of this propulsion technology in space by using five GR-1 thrusters for small attitude control maneuver 10. Aerojet completed a hot-fire test of the GR-1 in 2016 and further qualification tests were performed in 2017. 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 GR-1 unit when launched in 2018 1. The TRL is currently 6 for the GR-1 (Figure 4‑1), and 5 for the larger GR-22 (Figure 4‑3).

Aerojet Rocketdyne also is developing MPS-130 green monopropellant propulsion system, which is derived from their MPS-120 hydrazine propulsion system. The current ΔV capability for a 3U is 340 m/s and 130 m/s for a 6U. In Autumn 2016, NASA contracted Aeroject Rocketdyne to mature this system 2.

GR1 thruster
Image 4.2: GR1 thruster. Image Courtesy of Masse et al. (2015).
GR22_v2 thruster
Figure 4.3: GR22 thruster. Image Courtesy of Masse et al. (2015).

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 13. Current status is unknown.

hydros engineering unit
Figure 4.4: HYDROS engineering unit. Image Courtesy of James et al. (2015).

Tethers Unlimited, Inc. is developing a water electrolysis propulsion system called HYDROS-C, 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 thruster provides around 200 m/s ΔV for a 6U CubeSat and has a base configuration mass of 1.02 kg. This system has been selected for NASA’s first Pathfinder Demonstration CubeSat Mission planned for launch early 2019 3. The current TRL for this unit is 6.

NanoAvionics has developed a non-toxic mono-propellant propulsion system called Enabling Propulsion System for Small Satellites (EPSS) and it was demonstrated on LituanicaSAT-2, a 3U CubeSat, to correct orientation and attitude, avoid collisions, and extend orbital lifetime 4. It uses ADN as propellant and gives 252 s of specific impulse that is designed to provide 0.3 N thrust and up to 200 m /s ΔV. LituanicaSAT-2 was launched June 2017 and successfully separated from the primary (Cartosat-2) as part of the European QB50 initiative. The current TRL is 7.

A novel arc-ignition “green” CubeSAT hybrid thruster system prototype is currently under development at Utah State University. This system is fueled by 3-D printed acrylonitrile butadiene styrene (ABS) plastic for its unique electrical breakdown properties. Initially, high-pressure gaseous oxygen (GOX) was to be used as the oxidizer, however after safety considerations by NASA Wallops High Pressure Safety Management Team, it was concluded the oxidizer needed to contain 60% nitrogen, only 40% oxygen. On March 25th 2018, the system was successfully tested aboard a sounding rocket launched from NASA WFF in a hard space vacuum and the motor was successfully refired 5 times in space. During the tests, 8 N thrust level and 215 specific impulse were achieved as predicted 5. For small spacecraft applications, the TRL is currently 5.

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 capability of these systems that allow their application into nanosatellite 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.

Table 4-3: Cold and Warm Gas Propulsion Systems
Product Manufacturer Thrust Specific Impulse (s) Propellant TRL Status
MicroThruster Marotta 0.05 – 2.36 N 65 Nitrogen 9
Butane Propulsion System SSTL 0.5 N 80 Butane 9
MEMS (Nanoprop 3U) GomSpace/NanoSpace 0.01 – 1 mN 60 – 110 Butane 9
Nanoprop 6U GomSpace/NanoSpace 4 – 40 mN 60 – 110 Butane 9
MiPS Cold Gas VACCO 53 mN 23 Butane 9
MarCO MiPS VACCO 50 mN 94 R236FA 9
POPSAT-HIP1 Micro Space 0.083 – 1.1 mN 32 – 43 Argon 8
CNAPS UTIAS/SFL 12.5 – 40 mN 40 Sulfur Hexafluoride 9
CPOD VACCO 25 mN 40 R134A/R236FA 6
marotta cold gas thruster
Figure 4.5: Marotta cold gas thruster. Image Courtesy of Marotta.

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 17.

Butane_propulsion system from sstl
Figure 4.6: SSTL butane propulsion system. Image Courtesy of Gibbon (2010).

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 electrical interface 18.

CNAPS spacecraft with UTIAS propulsion modules.
Figure 4.7: CNAPS spacecraft with UTIAS propulsion modules. Image Courtesy of UTIAS website.

In June 2014, Space Flight Laboratory at University of Toronto Institute for Aerospace Research (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 6. This propulsion module is a novel version of the previous NanoPS that flew in the CanX-2 mission in 2008 20.

Another flight-demonstrated propulsion system was flown in the POPSAT-HIP1 CubeSat mission (launched June 2014) 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 ΔV 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 m-s operates each thruster 21.

nanospace MEMS cold gas system
Figure 4.8: NanoSpace MEMS cold gas system. Image Courtesy of NanoSpace.

A complete Microelectromechanical systems (MEMS) cold gas propulsion system for CubeSats (Figure 4‑7) provides accurate thrust control by using four thrusters with butane propellant has been devleoped. 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 22. This system is based on a flight-proven technology flown on the larger spacecraft (PRISMA mission, launched in 2010). The MEMS cold gas system was included into the bus of the TW-1 CubeSat, launched in September 2015 23.

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 refrigerant 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 24.

The Space Flight Laboratory (SFL) at the University of Toronto Institute for Aerospace Studies (UTIAS) developed a nitrous oxide (N2O) fueled monopropulsion system that provided 100 mN thrust at 131 s Isp during environmental tests performed in 2016 7. N2O is a common oxidizer for hybrid systems, that can be safely stored and can readily decompose into breathing air. Current status is unknown.

As part of their MPS thruster series, Aerojet Rocketdyne is developing a cold gas MPS-110 propulsion system. This module has a scalable 0.5 – 1U volume and is optimized for CubeSats orbit maintenance and de-orbiting. Current status of this system is unknown.

Solid motors

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 ΔV 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~300 s developed by Industrial Solid Propulsion 26.

Table 4‑4: Solid Rocket Motors
Product Manufacturer Total Mass (kg) Average Thrust (N) Specific Impulse (s) TRL Status
ISP 30sec motor Industrial Solid Propulsion 0.95 37 187 TRL 7
STAR 4G Orbital ATK 1.5 258 277 TRL 6
CAPS-3 DSSP 2.33 0.3 <900 TRL 8
MAP PacSci EMC Customized N/A 210 TRL 9
Module of DSSP thrusters
Figure 4.10: Module of DSSP thrusters. Image Courtesy of Nicholas et al. (2013)

SPINSAT, a 57 kg spacecraft, was deployed from ISS in 2014 and incorporated a set of solid motors (Figure 4‑8 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.

SpinSat Mission deploying from International Space Station (SpinSat top center))
Module of DSSP thrusters. Image Courtesy of Nicholas et al. (2013)

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 27.

Electric Propulsion Systems

Electric Propulsion

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 ΔV 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 ΔV maneuvers for transfer trajectories. For smaller ΔV applications, solid state materials such as polytetrafluoroethylene (PTFE), also known as Teflon, are used in most Pulsed Plasma Thrusters (PPTs) while electrosprays use various forms of ionic liquid

Resistojets

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.

Table 4-5: Resistojet Propulsion Systems
Product Manufacturer Thrust Power (W) Specific Impulse (s) Status
Low Power Resistojet Busek 13 mN at 200 W 100- 300 1390 Xenon TRL 8, Iodine TRL 4
CHIPS VACCO/CU Aerospace/AFRL 30 mN 30 82 TRL 6
PUC CU Aerospace and VACCO 0.45 N 15 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.

PUC module
Figure 4.11: PUC module. Image Courtesy of CU Aerospace and VACCO.

CU Aerospace and VACCO have built a Propulsion Unit for CubeSats (PUC), see Figure 4‑10. 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 28.

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 N-s 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 N-s of total impulse and 19 mN of thrust. The TRL of the integrated system is 5 and a second phase is currently in development 29.

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 N-s for main ΔV applications and 23 N-s for ACS 30.

The University of Toronto Institute for Aerospace Studies has also developed a warm gas resistojet system that has been assessed as TRL 6. This propulsion system was to fly on the LEO 2 spacecraft to achieve flight heritage on November 28 2017 but failed due to a launch vehicle anomaly 8.

Electrosprays

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 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.

Table 4‑6: Electrospray Propulsion Systems
Product Manufacturer Thrust Power (W) Specific Impluse (s) TRL Status
S-iEPS MIT 74 μN 1.5 1160 6
IMPACT Accion Systems Inc. 60 μN per axis 0.75 per axis 1200 5
TILE-V1 Accion Systems Inc. 1.5 mN 25 1500 5
TILE-500 Accion Systems Inc. 1.5 8-25 1250 5
MAX-1 Accion Systems Inc. 120 μN 1.6 2000 5
1 mN Electrospray Busek 0.7 mN 15 800 7
100μ Busek 0.1 mN 5 2300 7
Electrospray thruster
Figure 4.12: Electrospray thruster. Image Courtesy of MIT SPL.
MIT SiEPS
Figure 4 11: S-iEPS propulsion system. Image Courtesy of MIT SPL

Electrospray technology has been advanced significantly at Massachusetts Institute of Technology (MIT) Space Propulsion Laboratory (SPL) and some companies have started to commercialize systems based on this effort. Figure 4‑11 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, where a magnetically levitated thrust balance was used to measure thrust at μN levels 32. 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 4‑12, 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 33.

The SiEPS thruster was planned to be integrated on CubeSat mission, Aerocube 8 that was launched November 2016 from Vandenberg on an Atlas V 9, however it has not been confirmed if this thruster was integrated and if it has operated successfully. Until confirmation, the unit is at TRL 6.

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 N-s per axis. MAX-1 provides single-axis thrust, has a wet mass of 0.3 kg and a total impulse of 86 N-s 35. Additionally, there are two TILE thrusters  under development at Accion Systems that are desired for cubesat – smallsat form factors.  The TILE-V1 and TILE-50 have underwent qualification testing in 2017, and are slated for a AFRL flight demo in late 2018. TILE-50 module is roughly 100 x 100 x 120 mm size thruster that is planned for a flight demonstration on cubesat Irvine-01, part of the  Irvine CubeSat STEM Program, and is awaiting launch on Electron rocket later 2018. The outward facing surface consists of an array of 36 thruster chips, and the top layer beneath the of the module are other components stacked in a particular order, including the salt-based propellant, which is stored in tiny tanks, and the system’s power electronics, which run off of the satellites’ solar panels and batteries 10.

Busek Inc. has developed fully integrated electrospray propulsion systems in the mN range, the 100   micro-Newton BET 100uN and the one milli-Newton BET-1mN. These modules include 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 that launched December 2015, where all eight electric propulsion systems successfully fired 11. The system uses 15 W of power and provides 675 N-s with 50 mL of propellant and has a mass of 1.15 kg, and the 100 μN class thruster that provides a specific impulse of 2300 s and consumes 5 W. The 100 μN 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 12. The BET-100 systems was selected in March 2016 for a $1.6 million NASA award that will fly on a NASA Ames Pathfinder Technology Demonstration mission that is scheduled for launch in 2018 (Busek 2017), and underwent quality testing in late 2017.

Indium MEP
Figure 4.14: Indium MEP. Image Courtesy of Jet Propulsion Laboratory.

The Micro Devices Laboratory (MDL) at the Jet Propulsion Laboratory (JPL) has developed a highly integrated and scalable indium MEP system (Figure 4‑13) 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 38.

Ion Engines

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 at lower thrust. 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.

Table 4‑7: Ion Propulsion Systems and Thrusters
Product Manufacturer Thrust Power (W) Specific Impulse (s) Propellant TRL Status
BIT-3 Busek 1.4 mN 60 3500 Xenon-Iodine 6
BIT-1 Busek 0.1 mN 10 2250 Xenon 5
1-COUPS University of Tokyo 0.3 mN N/A 1000 Xenon 7
RIT-μX Airbus 50 – 500 μN 50 300 – 3000 Xenon 5
IFM Nano Thruster Enpulsion 10 μN – 0.4 mN 40 2000- 6000 7

Busek is developing a RF ion thruster that can operate with both xenon and iodine propellants, achieving similar performances 39.

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%. In 2015, it was shown that the 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. In July 2017, the BIT-3 completed two Critical Design Reviews for upcoming small spacecraft missions, IceCube and LunaH-Map to be launched with EM-1 in 2020 (Busek 2017).

A smaller thruster version of just 1 cm grids, called the BIT-1, is also under development by Busek. This system has a mass o 0.053 kg, provides 100 μN thrust and 2150 s Iso with 10W of power; thrust can exceed 180 μN and 3200 s Isp when more power is available 39. Current status is unknown.

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. 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 40.

RIT-?X
Figure 4.15: RIT-μX . Image Courtesy of Airbus.

Airbus  offers a family of RF ion thrusters and their smallest is the RIT-μX (Figure 4‑14). This thruster is designed for small spacecraft buses and high precision maneuvers. Various thrust configurations were proposed and tested. It uses xenon as propellant and it has a dry mass of 0.440 kg. In 2013, a system in the 50-500 μN range was demonstrated and thrust resolution, linearity, response and noise met LISA Pathfinder mission requirements increasing the TRL to 5. The nominal power to open rate 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 and high thrust levels of 50-2500μN were established in 2015 41. Current status is unknown.

A type of ion thruster that uses liquid metal rather than gases like xenon as a propellant is the field-emission electric propulsion (FEEP) device. This is a relatively new and advanced area of electrostatic space propulsion, and currently Enpulsion is the only commercial manufacturer worldwide of an FEEP thruster. The IFM Nano Thruster fits in a 1U volume and can produce 220 mN of thrust with a specific impulse of 4,000 seconds, and has already been flown on a 3U nanosatellite, deployed in January 2018 13.

Pulsed Plasma and Vacuum Arc Thrusters

In Pulsed Plasma Thrusters (PPTs), thrust is produced by triggering a high voltage discharge between two electrodes that results in an electric arc that typically ablates a solid state material like PTFE (telfon). 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 short circuits 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.

Table 4‑8: Pulsed Plasma and Vacuum Arc Propulsion Systems
Product Manufacturer Thrust Power (W) Specific Impulse (s) Propellant TRL Status
PPTCUP Mars Space and Clyde Space 40 μN 2 655 PTFE 6
NanoSat PPT Mars Space and Clyde Space 90 μN 5 640 PTFE 5
μ-CAT GWU and USNA 1 – 50 μN 2 – 14 2500 – 3000 Titanium 7
BmP-220 Busek 20 μN-s Impulse bit 1.5 536 PTFE 5
MPACS Busek 80 μN-s Impulse bit 10 827 PTFE 7
PPTCUP propulsion system
Figure 4.16: PPTCUP propulsion system. Image Courtesy of Ciaralli et. al (2015).

Mars Space Ltd. and Clyde Space Ltd. have developed a compact propulsion module (Figure 4‑15) 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 43.

pptbusek
Figure 4.17: The BmP-220. Image Courtesy of Busek.

Busek has extensive experience in the development of PPT systems. 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 μN-s at moderate power of less than 10 W 44 by using PTFE propellant. The system had heritage from previous investigations conducted at the Air Force Research Laboratory (AFRL) 45 by using PTFE propellant. The system had heritage from previous investigations conducted at the Air Force Research Laboratory (AFRL) and has been evolving since this first approach. The BmP-220 is the latest version of the Busek PPT family, 0.7U volume, see Figure 4‑15. It can provide up to 220 N-s 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 mN-s. The system TRL is estimated to be 5 30.

Patrick Neumann, the chief scientist and director of Neumann Space in Australia, is developing a pulsed cathodic arc thruster, or Neumann Drive. The Neumann Drive has broken the record for specific impulse previously held by NASA’s HIPEP thruster. It boasts a specific impulse as high as 10,000 seconds. This thruster operates like an arc welder, where metal is heated as arcing current jumps between a cathode and an anode. As electrons jump, they carry some atoms with them in the form of plasma and these atoms are propelled into space, creating thrust. The company hopes to flight test the thruster in 2020 13.

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. Disadvantages include a lower specific impulse and thrust.

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 46. 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 47.

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 48.

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.

Table 4‑9: Hall Effect Propulsion Systems and Thrusters
Product Manufacturer Thrust (mN) Power (W) Specific Impulse (s) TRL 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 6
BHT-200 during operation
Figure 4.18: BHT-200 during operation. Image Courtesy of Busek Co Inc.

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 30. 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 and utilizes xenon as propellant to achieve a peak total efficiency of up to 35% and a maximum specific impulse of 1350 s. The HT100 will be validated in-orbit as part of the µHETsat microsatellite where it will be tested to both maintain the orbit and preform active reentry maneuvers. The µHETsat microsatellite will be launched into a 500-600 circular orbit by LauncherOne by end of 2018 14.

The MSHT100 (Magnetically Shielded HT100) is based on the HT100 thruster although uses magnetic topology that limits wall erosion and increases thruster lifetime performance than that of the traditional HT100. The effectiveness of the magnetic shielding on the thruster ceramic walls was studied using xenon as propellant and endurance tests showed no signs of erosion before 350 hr of continuous firing. This prototype also underwent vacuum chamber tests that MSHT100 performed at slightly lower thrust levels and total impulse. Preliminary tests using krypton as a propellant were performed and only point has been verified 15.

Cylindrical Hall Effect Thruster
Figure 4.19: Cylindrical Hall Effect Thruster. Image Courtesy of UTIAS SFL.

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 51. Current status is unknown.

Propellant-less Systems

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 52.

One of the most recent solar sail mission for small spacecraft was performed by The Planetary Society in 2015. The 3U LightSail-1 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 2018, a follow up mission called LightSail-2 that will be housed on 3U Prox-1, will demonstrate orbit raising maneuvers using the same 32 m2 of mylar sail at a circular 720 km orbit as part of the Space test Program (SPT-2). 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 53.

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

CAT
Figure 4.20: Cubesat Ambipolar Thruster (CAT). Image Courtesy of Phase Four LLC.

The CubeSat Ambipolar 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 4‑19) 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. However, the highly oxidizing nature of iodine presents other storage challenges in the propellant tank. These challenges include feeding the propellant tank; iodine must sublime into a gas before it can be fed and large amounts required present a greater challenge it will be to design a tank that will feed the iodine propellant in a constituent and reliable method 16.

This system can achieve an estimated specific impulse of 1010 s when using iodine. In 2015, the PPU was in its development phase and some of the components for iodine utilization are at TRL 3 55. 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 56. The company Phase Four LLC is developing an integrated flight unit of the CAT 57. Current status is unknown.

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 58.

Orbital Technologies Corporation (ORBITEC) is developing the Miniature Nontoxic Oxide-Propane (MINNOP) propulsion system which uses nitrous oxide as the oxidizer. 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. In 2014, efforts towards a demonstration of the bipropellant thrust chamber and ignition within suitable weight constraints in order to fit into a 1U form factor 58, although current development status is unknown.

Another high-performance propellant is a mixture of nitrous oxide (N2O) and hydrocarbons that were fist examined by FireStar Technologies called NOFBX (Nitrous Oxide Fuel Blend Experimental). Here, the N2O fuel blend is stored as a mono-propellant and offers ≥300 s specific impulse. Additional benefits are the nontoxicity and relatively inexpensive, however challenges in high combustion temperatures in the tank remain. Originally, Firestar technologies designed a self-pressurizing non-toxic propellant used nitrous oxide as an oxidizer where the specific impulse in a vacuum was 325 s, but this specific program has been cancelled 59.

Combustion tests using green monopropellant mixture of N2O and ethylene (C2H4) were conducted and analyzed by German Aerospace Center’s Institute of Space Propulsion in 2017. Here the combustion efficiency was evaluated to understand occurring losses (from heat fluxes, non-ideal combustion, etc.) as well as the specific impulse. A reduction in chamber length increased the chamber efficiency due to the reduced heat flux 17

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 61.

The Inductively Coupled Electromagnetic (ICE) thruster is a novel technology that is being developed by MSNW LLC. This system used a small integrated RF oscillator to generate plasma. One of the main advantages is that this system could virtually use any liquid propellant. The total volume of the thruster and the PPU was expected to be less than 0.125U, and anticipated operating power was 10-50 W. In 2015, the current goal was to achieve TRL 4, however the current status is unknown.

In 2015, an experimental characterization of a low power helicon thruster was performed at the Stanford University’s Plasma Physics Laboratory. Tests were conducted by operating on water and argon propellants and thrust was observed at various performance levels, achieving magnitudes of 2-5 μN. Planned future work included include optimization for greater performance and thrust stand measurements 62. As power is regarded as a significant barrier towards advancing this technology, current efforts have been focused on developing a dc-RF power supply that can achieve substantial improvements in weight power density 18.

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 64.

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 65. Current status is unknown.

FENIX
FENIX. Image Courtesy of D-Orbit

A modular mirco-propulsion system called FENIX is being designed to raise or lower CubeSats into a different orbit at D-Orbit, see Figure 4‑20. This system consists of four small solid rocket motors that can be configured to any size CubeSat. The capabilities of this system can boost CubeSats into a higher orbit after deployment or be used for decommissioning maneuvers. The assessed TRL of this system is currently 4 19.

The design of a prototyped propulsion system called B125 Propulsion System is being studied at Benchmark Space Systems. The bipropellant is hydrogen peroxide (H2O2) as the oxidizer and is fueled by 2-propanol (alcohol blend). Studies published in 2018, have identified a benefit when using a homogenous catalysis process that provides the ability to operate in two modes: pseudo-monopropellant and bipropellant, which are achieved by varying the flow rates of the catalyst solution and hydrogen peroxide, however a technology challenge here is the development of an effective and reliable catalytic bed 20. This system provides 1.25 N of thrust at 260 s specific impulse, and with a total mass of 1.5 kg it can provide an 8 kg nanosatellite 145 m/s of ΔV 21. The current status is unknown.

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. The spacecraft will include Busek’s BHT-200-I propulsion system using iodine as the propellant instead of xenon. It was expected to be delivered for launch in the second quarter of 2017, however in May 2017 the decision was taken delay the launch even further to allow the maturation of the BHT-200-I 22.

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, PDT-1, the HYDROS-C water-based propellant thruster will be demonstrated to change the spacecraft’s velocity and altitude 23. PDT-1 is expected to launch in 2019.

JPL is supporting the InSight mission which launched in March 2018 that incorporated two identical CubeSats as part of the Mars Cube One (MarCO) technology demonstration. These spacecraft performed 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 refrigerant R-236FA as propellant, produces 755 N-s of total impulse and weighs 3.49 kg 71.

A team at Purdue University and NASA Goddard Space Flight Center is developing the Film Evaporation MEMS Tunable Array (FEMTA). This Microelectromechanical systems (MEMS) thruster uses deionized liquid water as propellant and consists of nozzles that produce thrust by applying local heat to a propellant capillary interface. The main advantages are the absence of any power required mechanism thus operating at low power consumption, order of mW. This technology plans to achieve TRL 6 by of fiscal year 2019 by targeting technology maturation activities to achieve payload requirements for a Pathfinder Technology Demonstration 6U mission 24.

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 on several of the CubeSat MEPSI-3 mission from the Aerospace Corporation between 2002 and 2006 73,74. 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 75. A version of this propulsion unit is intended to be used on the INSPIRE CubeSats for attitude control maneuvers and nominal flight operations 76.

NNEA Scout and Lunar Flashlight are two NASA MSFC missions that are going to be launched as part of EM-1, scheduled for 2020. For its main propulsion system, NEA Scout will deploy a sail of 80 m2 of area with 0.0601 mms-2 of characteristic acceleration, and will be steered by active mass translation that will have a VACCO cold gas MiPS (R236FA propellant). This module is approximately 2U in volume and will use six 23 mN thrusters to provide 30 m/s of ΔV 25. The propulsion system on Lunar Flashlight is a VACCO green mono propellant MiPS (AND propellant), that will be used for station keeping and attitude control. VACCO Lunar Flashlight MiPS is approximately 3U in volume and uses four Bradford/ECAPS 100 mN thrusters to develop 3,320 N-sec of total impulse that provides 237 m/s ΔV 25

Summary

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 ΔV 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, despite high cost per kg. 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 ΔV 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 important subsystem.

 

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