04. Propulsion

Introduction

There is currently a wide range of technologies for propulsion systems, however the miniaturization of these systems for small spacecraft has been particularly challenging. 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 small satellite propulsion flight demonstrations. Due to the continuous redesign of smallsat propulsion systems post flight demonstration, their associated TRL will be reflected to match the NASA Standard guidelines (found on NASA Nodis website). A system is only TRL 9 when the actual system is flight proven through successful mission operations with documented mission operation results on a small spacecraft. A redesign or change in the component architecture and the environment drops the TRL to 5 until proven demonstration in high fidelity environment (NASA NPR 7123.1B). Additionally, it should be noted that flight proven systems on a small spacecraft that is larger than a nanosatellite may still require testing for a smaller (nanosatellite) platform.

Cold gas or pulsed plasma systems for small delta-V maneuvers are fairly well established. However, higher delta-V applications require propulsion systems that are still in 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.

Electric and chemical systems have experienced a significant maturation process with respect to the previous 2015 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. 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.

Alternative (green) propellants offer safety and handling advantages over traditional hazardous propellants, such as hydrazine. Finally, in regards to propellant-less systems, the launch of LightSail has advanced the state of the art of solar sails for small spacecraft.

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 in the State of the Art and On the Horizon sections: chemical, electric and propellant-less systems, which are divided into smaller subsections depending on the type of thrust generation. The State of the Art section is defined as technology assessed at TRL 5 and higher, while On the Horizon section describes technology assessed at TRL 4 and below. 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. Table 4-1 shows a summary of the current state of the art for different propulsion methods.

The author would like to highlight that the presented tables are not intended to be exhaustive but to provide an overview of current state-of-the-art technologies and their development status for this small spacecraft subsystem. There is no intention of mentioning certain companies and omitting others based on their technologies.

 

Table 4-1: Propulsion Systems Types for Small Spacecraft
Product Thrust Specific Impulse (s) TRL Status
Hydrazine  0.5 – 30.7 N  200-235  9
Cold Gas 10 mN – 10 N 65 – 70 GN2/Butane 9
Non-toxic (Green) Propulsion 0.1 – 27 N  220 – 250  HAN 6, ADN 9
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)

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 corrective 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. 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 20141. Currently, this system has some final development tasks remaining and depending on the level of required qualification a first system delivery could be accomplished in the next year. The TRL is assessed at 5

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 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 thrusters 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. While all of these MONARCH thrusters have extensive flight heritage on larger spacecraft, there is no evidence they have a flown on a small spacecraft, making the TRL for small spacecraft application 5.

Alternative (Green) Propellant

Alternative propellants are designated ‘green fuel’ as they have a reduced toxicity due to the lower danger of component chemicals and significantly reduced vapor pressure as compared to hydrazine. The ‘green’ affiliation results in the propellant being less flammable which in turn require less safety requirements for handling, potentially including the lack of required SCAPE suit (self-contained atmospheric protective ensemble). This reduces operational oversight by safety and emergency personnel.

Range Safety AFSPCMAN91-710 requirements state that if a propellant is less prone to external leakage, which is seen with the alternative “green” systems due to higher viscosity of the propellant, the hazardous classification is reduced. As hydrazine external leakage is ranked as “catastrophic”, the use of alternative “green” propellants have been indicated with a reduced hazard severity classification to “critical” and possibly “marginal” per MIL-STD-882E  (Standard  Practice  for  System  Safety) 3. A classification of “critical” or less only requires a two-seal inhibits to external leakage, meaning no  additional  latch valves  other  isolation  device  are  required  in  the  feed  system 4. While these propellants are not safe for consumption, they have been shown to be less toxic compared to hydrazine. This is primarily due to alternative propellants being less flammable; nontoxic gasses (such as water vapor, hydrogen and carbon peroxide) are released when combusted.

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. These alternative propellants are generally less likely to exothermically decompose at room temperature due to higher ignition thresholds. Therefore they require fewer inhibits resulting in fewer valve seats to power, including a less stringent temperature requirement with less system heater power.

Alternative 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 (ADN) 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 ADN Bradford Engineering 0.25 – 1.00 204 – 235 9
HYDROS-C Other Tethers Unlimited, Inc. 1.2 310 6
AMAC Other Busek 0.425 225 5
MiPS ADN VACCO 0.4 250 6
BGT-X5 HAN Busek 0.5 220 5
EPSS C1K ADN NanoAvionics 0.3 252 7
Green Hybrid Other Utah State 8 215 6
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). HPGP 1-N systems are being implemented in SkySat missions such as SkySat-3 (120 kg mass, launched June 2016), and SkySat block-I, and as of October 2017 13 SkySat small spacecraft were launched and are fully operational making the TRL for this system 9. The HGPG systems are currently developed for three different thrust magnitudes: 5-N and 22-N, with higher thrust systems in development 6 7.

VACCO partnered with Bradford Engineering (formerly ECAPS) to design a self-contained unit that can deliver up to 3320 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. This unit has also an ArgoMoon Prolusion thruster that incorporates one 100-mN ADN thruster and four 10-mN cold gas thrusters for attitude control, providing up to 783 N-s of total impulse for main ΔV applications and 72 N-s for RCS 8. 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 spacecraft. 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 (HAN-based). 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 2018-2019. 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 GR-1 unit when launched 5. The TRL is currently 6 for the GR-1 (Figure 4‑2), 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. 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 mNs. 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 and average thruster of 1.2-N with 310 s Isp. This system has been selected for NASA’s first Pathfinder Demonstration CubeSat Mission planned for launch early 2019 6. The current TRL for this unit is 6 as it has not flown yet.

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 7. 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 into space and the motor was successfully re-fired 5 times. During the tests, 8-N thrust level and a specific impulse of 215 s were achieved as predicted 8. For small spacecraft applications, the TRL is currently 6.

Cold 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.

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.

*Information was taken from brochure and may need to be updated by vendor

Table 4‑3: Cold 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
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 10 mN 40 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

A cold gas thruster developed by Marotta flew on the NASA ST-5 mission (launch mass 55 kg) 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 4-5: 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 4-5. 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 4-6: CNAPS spacecraft with UTIAS propulsion modules. Image Courtesy of UTIAS website.

IIn 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 4-6, 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 9. 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 4-7: 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 developed. 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 N-s 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.

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 6
STAR 4G Orbital ATK 1.5 258 277 TRL 6
CAPS-3 DSSP 2.33 0.3 <300 TRL 8
MAP PacSci EMC Customized N/A 210 TRL 9

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

The STAR motor was initially developed and tested for use of deploying constellations of small spacecraft in early 2000 under NASA Goddard Space Flight Center program, and the 4G motor was first tested in late 2000 10. The current status of this motor is unknown.

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 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), such 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) TRL Status
Micro Resistojet Busek 10 mN 15 150 5
CHIPS CU Aerospace and VACCO 30 mN 30 82 5
PUC CU Aerospace and VACCO 0.45 N 15 70 6
Resistojet Propulsion System SSTL 100 mN 30 – 50 48 – 99 9

The Micro Resistojet offer at Busek is still in development for small spacecraft, and current specs list a max of 10 mN thrust at 150 s Isp at 15 W of power. The delta-V capability for a 4 kg spacecraft was projected at 60 m/s and the total system mass is 1.25 kg using ammonia as the propellant.  The current status of this thruster is unknown.

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. There is heritage on small spacecraft, but is not scalable on a cubesat without redesign.

PUC module
Figure 4-09: 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 11. Additionally, UTIAS-SFL developed a nitrous oxide (N2O) fueled monopropulsion system that provided 100 mN thrust at 131 s Isp during environmental tests performed in 2016 12. N2O is a common oxidizer for hybrid systems that can be safely stored and can readily decompose into breathing air. Current status is unknown

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 may 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 Impulse (s) TRL

Status

S-iEPS MIT 74 μN 1.5 1160 6
TILE-5000 Accion Systems Inc. 1.5 8-25 1250 5
1 mN Electrospray Busek 0.7 mN 15 800 7
100μ Busek 0.1 mN 5 2300 5
Electrospray thruster
Figure 4-11: Electrospray thruster. Image Courtesy of MIT SPL.
MIT SiEPS
Figure 4-10: 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 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 S-iEPS thruster was planned to be integrated on CubeSat mission, Aerocube 8 that was launched November 2016 from Vandenberg on an Atlas V 13, 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.

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 14. 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 15. The BET-100 systems was selected in March 2016 to fly on a NASA Ames Pathfinder Technology Demonstration mission that is scheduled for launch in 2018, and underwent quality testing in late 2017. However this flight has been cancelled and there is no current information on this system.

Indium MEP
Figure 4-12: 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 4-12) 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. The current TRL for this system is unknown.

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 Iodine 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 Indium 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 16.

RRecently, 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 ΔV 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 40.

RIT-?X
Figure 4-13: RIT-μX . Image Courtesy of Airbus.

Airbus  offers a family of RF ion thrusters and their smallest is the RIT-μX (Figure 4‑13). 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 50-2500 μN 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. 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 .

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 (Teflon). A self-generated magnetic field is produced and then accelerates and expels particles from the thruster head, and the propellant is typically 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
Metal Plasma Thruster Applied Sciences Corp. 15 μN/W 100 826 s (Pt) up to 2400 s (Al) Any solid metal, Mo/Nb 5
PPTCUP propulsion system
Figure 4-14: PPTCUP propulsion system. Image Courtesy of Ciaralli et. al (2015).

Mars Space Ltd. and Clyde Space Ltd. have developed a compact propulsion module (Figure 4-14) 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 4-15: 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) and has been evolving since this first approach making its TRL 7. 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 30. The system TRL is estimated to be 5.

Dr. Patrick Neumann is developing a Pulsed Cathodic Arc Thruster (PCAT), or Neumann Drive that has broken the record for specific impulse previously held by NASA’s HiPEP thruster. It boasts a specific impulse as high as 14,000 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. This ion drive will be installed on the Airbus Defense and Space Bartolomeo platform part of the FAST (Facility for Australian Space Testing) mission in mid-2019 17 .

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 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 4-17: 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 4-17, 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. Additionally, it was launched with the FalconSat-6 (150 kg) mission on a Falcon Heavy in 2018. 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 of this chapter.

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. Cathode 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 where it will be tested to both maintain the orbit and accelerated reentry. A larger version, the HT400, operates at a nominal power of 400 W and is TRL 5 18.

Cylindrical Hall Effect Thruster
Figure 4-16: 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 4-16) 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 the 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 were done in order to raise the TRL from 5 to 6 in early 2016 51. Current status is unknown.

Radio Frequency (RF) Thrusters

The Phase Four RF Thrusters (RFT) that leverages the developed ambipolar technology from the CAT was tested at The Aerospace Corporation and Phase Four laboratories. Similarly to its predecessor, the RFT has a ceramic plasma liner which is wrapped in an inductive RF antenna or coil that is itself located inside a magnetic field generated by a permanent magnet. Inside the liner, xenon is ionized and the subsequent plasma is heated by induced oscillating magnetic fields. Electrons get accelerated at very high energies and this quick flux produces charge imbalance in the system. Then, propellant ions are expelled out of the nozzle due to the momentary imbalance, becoming the main source of thrust.

There are several notable advantages of the RFT: The size reduction and power density improvements in the RF switching electronics have allowed the PPU to be less than 500 g for LEO CubeSat applications; second, the ambipolar nature of the technology obviates the need for a cathode neutralizer, which implies that no high voltage electronics are required; finally, since the thruster does not have electrodes, more propellants can be used, since they no longer can be corrosive to cathodes or anodes in their plasma state.

There have been proof of concept Phase Four RF thrusters, RFT-0 RFT-2 and RFT-X, that showed comparable performance on a direct thrust stand to other RF thrusters that operate at much higher powers or have higher dimensions and mass 19. Despite the differences in the electrical, mechanical and magnetic characteristics, the specific impulse performance results scaled to the same linear trend, and are in the same order of magnitude than equivalent low power Hall Effect thrusters, within 50% of the thrust output at similar power levels, and with the advantage of being electrodeless 20. Based on these technologies, the Maxwell RF Thruster propulsion system corresponds to a 400 W class engine, and it is operated at a power range of 342-480 W, achieving 4.3-9 mN at specific impulses of 1463-918 s 21. The TRL on the Maxwell thruster is currently 5.

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.

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 22.

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.

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 0.125\U, and anticipated operating power was 10-50 W. In 2015, the current goal was to achieve TRL 4 61, 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 23.

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
Figure 4-18: 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 4‑18. 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 24.

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 25. 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 26. 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, had consisted 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 this mission was cancelled.

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 27. 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 28.

NEA 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 29. 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 29

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.

For Feedback solicitation, please email: arc-sst-soa@mail.nasa.gov. Please include a business email so someone may contact you further.

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