Chapter Contents

• Chapter Glossary
• 4.1 Introduction
  • 4.1.1 Document Organization
• 4.2 Public Data Sources and Disclaimers
• 4.3 Definitions
• 4.4 Technology Maturity
  • 4.4.1 Application of the TRL Scale to Small Spacecraft Propulsion Systems
  • 4.4.2 Progress Toward Mission Infusion (PMI)
• 4.5 Overview of In-Space Propulsion Technology Types
• 4.6 State-of-the-Art in Small Spacecraft Propulsion
  • 4.6.1 In-Space Chemical Propulsion
  • 4.6.2 In-Space Electric Propulsion
  • 4.6.3 In-Space Propellant-less Propulsion
• References

Chapter Glossary

(ABS)		Acrylonitrile Butadiene Styrene
(AC)		Alternating Current
(ACE)		Apollo Constellation Engine
(ACO)		Announcement for Collaborative Opportunity
(ACS)		Attitude Control System
(ADN)		Ammonium Dinitramide
(AFRL)		Air Force Research Laboratory
(AIS)		Applied Ion Systems
(AOCS)		Attitude and Orbit Control System
(AR)		Aerojet Rocketdyne
(ARC)		Ames Research Center
(BOL)		Beginning of Life
(CHIPS)		CubeSat High Impulse Propulsion System
(CMNT)		Colloid MicroNewton Thrusters
(CNAPS)		Canadian Nanosatellite Advanced Propulsion System
(CNES)		French National Center for Space Studies
(CPOD)		CubeSat Proximity Operations Demonstration
(CUA)		CU Aerospace LLC
(DFMR)		Design for Minimum Risk
(DRM)		Design Reference Mission
(DSSP)		Digital Solid State Propulsion LLC
(EOL)		End of Life
(EMC)		Electromagnetic Compatibility
(EMI)		Electromagnetic Interference
(EP)		Electric Propulsion
(EPL)		Electric Propulsion Laboratory, Inc.
(EPSS)		Enabling Propulsion System for Small Satellites
(ESA)		European Space Agency
(ESPs)		Electrically Controlled Solid Propellant
(FASTSAT)		Fast, Affordable, Science and Technology Satellite
(FEEP)		Field Emission Electric Propulsion
(FPPT)		Fiber-Fed Pulsed Plasma Thruster
(GEO)		Geostationary Equatorial Orbit
(GIT)		Gridded-ion Thrusters
(GOCE)		Gravity Field and Steady-State Ocean Circulation Explorer
(GOX)		Gaseous Oxygen
(GPIM)		Green Propellant Infusion Mission
(GPS)		Global Positioning System
(GRC)		Glenn Research Center
(GSFC)		Goddard Space Flight Center
(HAN)		Hydroxylammonium Nitrate
(HET)		Hall-effect Thruster
(HTP)		High Test Peroxide
(HTPB)		Hydroxyl-terminated Polybutadiene
(IPS)		Integrated Propulsion System
(ISS)		International Space Station
(JANNAF)		Joint Army Navy NASA Air Force
(JHU ERG)		Johns Hopkins University Energetics Research Group
(JPL)		Jet Propulsion Laboratory
(LFPS)		Lunar Flashlight Propulsion System
(LISA)		Laser Interferometer Space Antenna
(MAPS)		Modular Architecture Propulsion System
(MarCO)		Mars Cube One
(MCD)		Micro-cavity Discharge
(MEMS)		Microelectromechanical System
(MEO)		Medium Earth Orbit
(MMH)		Monomethyl Hydrazine
(MPUC)		Monopropellant Propulsion Unit for CubeSats
(MSFC)		Marshall Space Flight Center
(MVP)		Monofilament Vaporization Propulsion
(N2O)		Nitrous Oxide
(NEA)		Near-Earth Asteroid
(NODIS)		NASA Online Directives Information System
(NSTT)		Nanosat Terminator Tape
(OTS)		Orbital Transfer System
(OTV)		Orbital Transfer Vehicle
(PacSci EMC)		Pacific Scientific Energetic Materials Company
(PBM)		Plasma Brake Module
(PMD)		Propellant Management Device
(PMDs)		Propellant Management Devices
(PMI)		Progress toward Mission Infusion
(PMMA)		Polymethyl Methacrylate
(PPT)		Pulsed Plasma Thrusters
(PPU)		Power Processing Unit
(PTD)		Pathfinder Technology Demonstration
(PTFE)		Polytetrafluoroethylene
(PUC)		Propulsion Unit for CubeSats
(ROMBUS)		Rapid Orbital Mobility Bus
(SAA)		Space Act Agreement
(SBIR)		Small Business Innovative Research
(SCAPE)		Self Contained Atmospheric Protective Ensemble
(SEP)		Solar Electric Propulsion
(SMAP)		Soil Moisture Active Passive
(SMART-1)		Small Missions for Advanced Research in Technology
(SME)		Subject Matter Experts
(SSTL)		Surrey Satellite Technology Ltd.
(SSTP)		Small Spacecraft Technologies Program
(TBD)		To Be Determined
(TCMs)		Trajectory Correction Maneuvers
(TMA)		Technology Maturity Assessment
(TRL)		Technology Readiness Level
(UTIAS)		University of Toronto Institute for Aerospace Research
(VAT)		Vacuum arc thrusters
(VENuS)		Vegetation and Environment monitoring on a New Microsatellite
(WFF)		Wallops Flight Facility

4.1 Introduction

In-space propulsion devices for small spacecraft are rapidly increasing in number and variety. Although a mix of small spacecraft propulsion devices have established flight heritage, the market for new propulsion products continues to prove dynamic and evolving. In some instances, systems and components with past flight heritage are being reconsidered to meet the needs of smaller spacecraft. This approach minimizes new product development risk and time to market by creating devices similar to those with existing spaceflight heritage, although accounting for small spacecraft volume, mass, power, safety and cost considerations. Such incremental advancement benefits from existing spaceflight data, physics-based models, and customer acceptance of the heritage technologies, which eases mission infusion. In other instances, novel technologies are being conceived specifically for small spacecraft using innovative approaches to propulsion system design, manufacturing, and integration. While the development of novel technologies typically carries a higher risk and slower time to market, these new technologies strive to offer small spacecraft a level of propulsive capability not easily matched through the miniaturization of heritage technologies. Such novel devices are often highly integrated and optimized to minimize the use of a small spacecraft’s limited resources, lower the product cost, and simplify integration. Regardless of the development approach, the extensive investments by commercial industry, academia, and government to develop new propulsion products for small spacecraft suggests long-term growth in the availability of propulsion devices with increasingly diverse capabilities.

In the near-term, the surge in public and private investments in small spacecraft propulsion technologies, combined with the immaturity of the overall small spacecraft market, has resulted in an abundance of confusing, unverified, sometimes conflicting, and otherwise incomplete technical literature. Furthermore, the rush by many device developers to secure market share has resulted in some confusion surrounding the true readiness of these devices for mission infusion. As third parties independently verify device performance, and end-users demonstrate these new devices in their target environments, the true maturity, capability, and flight readiness of these devices will become evident. In the meantime, this report will attempt to reduce confusion by compiling a list of publicly described small spacecraft propulsion devices, identifying publicly available technical literature for further consideration, recognizing missions of potential significance, and organizing the data to improve comprehension for both neophytes and subject matter experts.

This chapter avoids a direct technology maturity assessment (TMA) based on the NASA Technology Readiness Level (TRL) scale, recognizing insufficient in-depth technical insight into current propulsion devices to perform such an assessment accurately and uniformly. An accurate TRL assessment requires a high degree of technical knowledge on a subject device as well as an understanding of the intended spacecraft bus and target environment. While the authors strongly encourage a TMA that is well-supported with technical data prior to infusing technologies into programs, the authors believe TRLs are most accurately assessed within the context of a program’s unique requirements. Rather than attempting to assess TRL in the absence of sufficient data, this chapter introduces a novel classification system that simply recognizes Progress toward Mission Infusion (PMI) as an indicator of the efficacy of the manufacturers’ approach to system maturation and mission infusion. PMI should not be confused with TRL as PMI does not directly assess technology maturity. However, PMI may prove insightful in early trade studies. The PMI classification system used herein is described in detail in Section 4.4.2.

4.1.1 Document Organization

This chapter organizes the state-of-the-art in small spacecraft propulsion into the following categories:

• In-Space Chemical Propulsion              (4.6.1)
• In-Space Electric Propulsion                  (4.6.2)
• In-Space Propellant-less Propulsion   (4.6.3)

Each of these categories is further subdivided by the prevailing technology types. The subsections organize data on each prevailing technology type as follows:

• Technology Description
• Key Integration and Operational Considerations
• Current & Planned Missions
• Summary Table of Devices
• Notable Advancements

The organizational approach introduces newcomers to each technology, presents technology-specific integration and operation concerns for the reader’s awareness, highlights recent or planned missions that may raise the TRL of specific devices, and tabulates procurable devices of each technology. Some sections also include an incomplete list of notable advancements. While the key integration and operational considerations are not comprehensive, they provide initial insights that may influence propulsion system selection. In the cases where a device has a long history of spaceflight, this chapter reviews only select missions.

4.2 Public Data Sources and Disclaimers

This chapter is a survey of small spacecraft propulsion technologies as discussed in open literature and does not endeavor to be an original source. As such, this chapter only considers literature found in the public domain to identify and classify devices. Commonly used sources for public data include manufacturer datasheets, press releases, conference papers, journal papers, public filings with government agencies, and news articles.

This chapter summarizes device performance, capabilities, and flight history, as presented in publicly available literature. Data not appropriate for public dissemination, such as proprietary, export controlled, or otherwise restricted data, are not considered. As such, actual device maturity and flight history may be more or less extensive than what is documented herein. Device manufacturers should be consulted for the most up-to-date and relevant data before performing a TMA.

This chapter’s primary data source is literature produced by device manufacturers. Unless otherwise published, do not assume independent verification of device performance and capabilities. Performance and capabilities described may be speculative or otherwise based on limited data.

The information presented is not intended to be exhaustive but to provide a general overview of current state-of-the-art technologies and their development status. It should be noted that technology maturity designations may vary with change to payload, mission requirements, reliability considerations, and/or the environment in which performance was demonstrated. Readers are highly encouraged to reach out to companies for further information regarding the performance and maturity of the described technology. There is no intention of mentioning certain companies and omitting others based on their technologies or relationship with NASA.

Suggestions or corrections for this document should be submitted to the NASA Small Spacecraft Virtual Institute Agency-SmallSat-Institute@mail.nasa.gov for consideration prior to the publication of future issues. When submitting comments, please cite appropriate publicly accessible references. Private correspondence is not considered an adequate reference.

4.3 Definitions

• Device refers to a component, subsystem, or system, depending on the context.
• Technology refers to a broad category of devices or intangible materials, such as processes.

4.4 Technology Maturity

4.4.1 Application of the TRL Scale to Small Spacecraft Propulsion Systems

NASA has a well-established guideline for performing TMAs, described in detail in the NASA Systems Engineering Handbook (1). A TMA determines a device’s technological maturity, which is usually communicated according to the NASA TRL scale. The TRL scale is defined in NASA Procedural Requirements (NPR) 7123 (2). The NASA Systems Engineering Handbook and NPR 7123 can be accessed through the NASA Online Directives Information System (NODIS) library. Assessment of TRLs for components, systems, or software allows for coherent communication between technologists, program managers, and other stakeholders regarding the maturity of a technology. Furthermore, TRL is a valuable tool to communicate the potential risk associated with the infusion of technologies into programs. For TRLs to be applied across all technology categories, the NASA TRL definitions are written broadly and rely on subject matter experts (SME) in each discipline to interpret appropriately.

Recently, U.S. Government propulsion SMEs suggested an interpretation of the TRL scale specifically for micro-propulsion. The Micro-Propulsion Panel of the JANNAF Spacecraft Propulsion Subcommittee in 2019 published the JANNAF Guidelines for the Application of Technology Readiness Levels (TRLs) to Micro-Propulsion Systems (3). The guideline was recently updated in 2022 to reflect the latest community input (4). This guideline suggests an interpretation of TRL for micro-propulsion and reflects both NASA and DOD definitions for TRL. The JANNAF panel consists of participants from the Air Force Research Laboratory (AFRL), Glenn Research Center (GRC), Jet Propulsion Laboratory (JPL), and Goddard Space Flight Center (GSFC). The panel further receives feedback from the non-Government propulsion community. While this JANNAF guideline focuses on micro-propulsion (e.g., CubeSats), the guideline still has relevance to rigorously assessing TRLs for the more general category of small spacecraft in-space propulsion. By establishing a common interpretation of TRL for small spacecraft propulsion, a more coherent and consistent communication of technology maturity can occur between small spacecraft propulsion providers and stakeholders. The JANNAF guideline is open to unlimited distribution and may be requested from the Johns Hopkins University Energetics Research Group (JHU ERG). Ensure the use of the latest JANNAF guideline, as the guideline may continue to evolve with further community input.

A fundamental limitation of the JANNAF guideline for TRL assessment, and TMA in general, is an assumption of in-depth technical knowledge of the subject device. In the absence of detailed technical knowledge, especially in a broad technology survey as presented herein, a TMA may be conducted inaccurately or inconsistently. Furthermore, assessment of TRL assumes an understanding of the end-user application. The same device may be concluded to be at different TRLs for infusion into different missions. For example, a device may be assessed at a high TRL for application to low-cost small spacecraft in low-Earth orbits, while assessed at a lower TRL for application to geosynchronous communication satellites or NASA interplanetary missions due to different mission requirements. Differences in TRL assessment based on the operating environment may result from considerations such as thermal environment, mechanical loads, mission duration, or radiation exposure. Propulsion-specific variances between missions might include propellant type, total propellant throughput, throttle set-points, burn durations, and the total number of on/off cycles. As such, an accurate TRL assessment not only requires an in-depth technical understanding of a device’s development history, including specifics on past flight-qualification activities, but also an understanding of mission-specific environments and interfaces. The challenge of assessing an accurate TRL in a broad technology survey poses a significant burden for data collection, organization, and presentation. Such activities are better suited for programs seeking to infuse new technologies into their missions.

Given the rapid evolution of small spacecraft propulsion technologies and the variety of mission environments, as well as generally limited device technical details in open literature, the propulsion chapter implements a novel system to classify technical maturity according to Progress toward Mission Infusion (PMI). This novel classification system is not intended to replace TRL but is a complementary tool to provide initial insight into device maturity when it is not yet feasible to accurately and consistently apply the TRL scale. This novel classification system is discussed in detail below.

Readers are strongly encouraged to perform more in-depth technical research on candidate devices based on the most up-to-date information available, as well as to assess risk within the context of their specific mission(s). A thoughtful TMA based on the examination of detailed technical data through consultation with device manufactures can reduce program risk and increase the likelihood of program success. This survey is not intended to replace the readers’ own due diligence. Rather, this survey and PMI seek to provide early insights that may assist in propulsion system down-select to a number of devices where an in-depth TMA becomes feasible.

4.4.2 Progress Toward Mission Infusion (PMI)

Rather than directly assessing a device’s technical maturity via TRL, propulsion devices described herein are classified according to evidence of progress toward mission infusion. This is a novel classification system first introduced in this survey. Assessing the PMI of devices in a broad survey, where minimal technical insight is available, may assist with down-selecting propulsion devices early in mission concept development. Once a handful of devices are identified as possible solutions for a specific mission concept, a detailed TMA and rigorous TRL assessment should be conducted. The PMI classification system sort devices into one of four broad technology development categories: Concept, In-Development, Engineering-to-Flight, and Flight-Demonstrated. The following sections describe the PMI classification system in-detail. Furthermore, figure 4.1 summarizes the PMI classifications.

Concept, ‘C’

The Concept classification reflects devices in an early stage of development, characterized by feasibility studies and the demonstration of fundamental physics. Concept devices typically align with the NASA TRL range of 1 to 3. At a minimum, these devices are established as scientifically feasible, perhaps through a review of relevant literature and/or analytical analysis. These devices may even include experimental verification that supports the validity of the underlying physics. These devices may even include notional designs. While Concept devices are generally not reviewed herein, particularly promising Concept devices will be classified in tables with a ‘C’.

In-Development, ‘D’

The In-Development classification reflects the bulk of devices being actively matured and covered in this survey, where only a modest number of devices may progress to regular spaceflight. In-Development devices typically align with the NASA TRL range of 4 to 5. While In-Development devices may have specific applications attributed by their developers, no selection for a specific mission has been publicly announced. In the absence of a specific mission, device development activities typically lack rigorous system requirements and a process for independent requirement validation. Furthermore, qualification activities conducted in the absence of a specific mission typically require a delta-qualification to address mission-specific requirements. At a minimum, In-Development devices are low-fidelity devices that have been operated in an appropriate environment to demonstrate basic functionality and support prediction of the device’s ultimate capabilities. They may even be medium- or high-fidelity devices operated in a simulated final environment, but lacking a specific mission pull to define requirements and a qualification program. They may even be medium- or high-fidelity devices operated in a spaceflight demonstration but lacking sufficient fidelity or demonstrated capability to reflect the anticipated final product. These devices are typically described as a technology push, rather than a mission pull. In-Development devices will be classified in tables with a ‘D’.

Engineering-to-Flight, ‘E’

The Engineering-to-Flight classification reflects devices with a publicly announced spaceflight opportunity. This classification does not necessarily imply greater technical maturity than the In-Development classification, but it does assume the propulsion device developer is receiving mission-specific requirements to guide final development and qualification activities. Furthermore, the Engineering-to-Flight classification assumes a mission team performed due diligence in the selection of a propulsion device, and the mission team is performing regular activities to validate that the propulsion system requirements are met. Thus, while the PMI classification system does not directly assess technical maturity, there is an underlying assumption of independent validation of mission-specific requirements, where a mission team does directly consider technical maturity in the process of device selection and mission infusion. Engineering-to-Flight devices typically align with the NASA TRL range of 5 to 6. At a minimum, these are medium-fidelity devices that have been operated in a simulated final environment and demonstrate key capabilities relative to the requirements of a specific mission. These devices may even be actively undergoing or have completed a flight qualification program. These devices may even include a spaceflight, but in which key capabilities failed to be demonstrated or further engineering is required. These devices may even include a previously successful spaceflight, but the devices are now being applied in new environments or platforms that necessitate design modifications and/or delta-qualification. These devices must have a specific mission pull documented in open literature. A design reference mission (DRM) may be considered in place of a specific mission pull, given detailed documentation in open literature, which includes a description of the DRM, well-defined propulsion system requirements, maturation consistent with the DRM requirements, and evidence of future mission need consistent with the DRM. Engineering-to-Flight devices will be classified in tables with an ‘E’.

Flight-Demonstrated, ‘F’

The Flight-Demonstrated classification reflects devices where a successful technology demonstration or genuine mission has been conducted and described in open literature. Flight-Demonstrated devices typically align with the NASA TRL range of 7 to 9. These devices are high-fidelity components or systems (in fit, form, and function) that have been operated in the target in-space environment (e.g., low-Earth orbit, GEO, deep space) on an appropriate platform, where all key capabilities were successfully demonstrated. These devices may even be final products, which have completed genuine missions (not simply flight demonstrations). These devices may even be in repeat production and routine use for several missions. The devices must be described in open literature as successfully demonstrating key capabilities in the target environment to be considered Flight-Demonstrated. If a device has flown, but the outcome is not publicly known, the classification will remain Engineering-to-Flight. Flight-Demonstrated devices will be classified in tables with an ‘F’.

4.5 Overview of In-Space Propulsion Technology Types

In-space small spacecraft propulsion technologies are generally categorized as (i) chemical, (ii) electric, or (iii) propellant-less. This chapter surveys propulsion devices within each technology category. Additionally, liquid-propellant acquisition and management devices are reviewed as an important component of in-space propulsion systems. Although other key subsystems have not yet been reviewed, such as small spacecraft propulsion power processing units, they may be included in future updates of this publication. Table 4-1 lists the in-space propulsion technologies reviewed. Figure 4.2 graphically illustrates the range of thrust and specific impulse for these small spacecraft propulsion devices. The thrust and specific impulse ranges provided in table 4-1 and figure 4.1 only summarize the performance of small spacecraft devices covered in this survey and may not reflect the broader capability of the technologies beyond small spacecraft or the limits of what is physically possible with further technology advancement. Furthermore, propulsion systems are often highly throttleable, and the devices surveyed herein may in many cases be capable of offering performance beyond the ranges presented in table 4-1.

Chemical systems have enabled in-space maneuvering since the onset of the space age, proving highly capable and reliable. These include hydrazine-based systems, other mono- or bipropellant systems, hybrids, cold gas systems, and solid propellants. Typically, these systems are sought when high thrust or rapid maneuvers are required. As such, chemical systems continue to be the in-space propulsion technology of choice when their total impulse capability is sufficient to meet mission requirements.

On the other hand, the application of electric propulsion devices has been historically far more limited. While electric propulsion can provide an order of magnitude greater total impulse than chemical systems, research and development costs have typically eclipsed that of comparable chemical systems. Furthermore, electric propulsion generally provides thrust-to-power levels below 75 mN/kW. Thus, a small spacecraft capable of delivering 500 W to an electric propulsion system may generate no more than 38 mN of thrust. Therefore, while the total impulse capability of electric propulsion is generally considerable, these systems may need to operate for hundreds or thousands of hours, compared to the seconds or minutes that chemical systems necessitate for a similar impulse. That said, the high total impulse and low thrust requirements of specific applications, such as station keeping, have maintained steady investment in electric propulsion over the decades. Only in recent years has the mission pull for electric propulsion reached a tipping point where electric propulsion may overtake chemical for specific in-space applications. Electric propulsion system types considered herein include electrothermal, electrospray, gridded ion, Hall-effect, pulsed plasma and vacuum arc, and ambipolar.

Propellant-less propulsion technologies such as solar sails, tethers, electric sails (and plasma brakes), and aerodynamic drag devices have long been investigated, but they have yet to move beyond small-scale demonstrations. However, growing needs such as orbital debris removal may offer compelling future applications.

Some notable categories are not covered in this survey, such as nuclear in-space propulsion technologies. While substantial investment continues in such areas for deep space science and human exploration, such technologies are generally at lower TRL and typically aim to propel spacecraft substantially larger than the 180 kg wet mass limit covered by this report.

Whenever possible, this survey considers complete propulsion systems, which are composed of thrusters, feed systems, pressurization systems, propellant management and storage, and power processing units, but not the electrical power supply. However, for some categories, components (e.g., thruster heads) are mentioned without consideration of the remaining subsystems necessary for their implementation. Depending on the device’s intended platform (i.e., NanoSat, MicroSat, SmallSat), the propulsion system may be either highly integrated or distributed within the spacecraft. As such, it is logical to describe highly integrated propulsion units at the system level, whereas components of distributed propulsion systems may be logically treated at the sub-system level, where components from a multitude of manufacturers may be mixed-and-matched to create a unique mission-appropriate propulsion solution.

Table 4-1: Summary of Propulsion Technologies Surveyed
Technology	Thrust Range	Specific Impulse Range [sec]
4.6.1 CHEMICAL PROPULSION TECHNOLOGIES
Hydrazine Monopropellant	0.25 – 28 N	180 – 285
Alternative Mono- and Bipropellants	50 mN – 22 N	150 – 310
Hybrids	8 – 222 N	215 – 300
Cold Gas	10 μN – 3.6 N	40 – 110
Solid Motors	37 – 461 N	187 – 269
Propellant Management Devices	–	–
4.6.2 ELECTRIC PROPULSION TECHNOLOGIES
Electrothermal	0.1 mN – 1 N	20 – 350
Electrosprays	20 μN – 20 mN	225 – 3,000
Gridded Ion	0.1 – 20 mN	500 – 3,000
Hall-Effect	0.25 – 55 mN	200 – 1,920
Pulsed Plasma and Vacuum Arc Thrusters	4 – 500 μN	87 – 3,200
Ambipolar	0.5 – 17 mN	400 – 1,100
4.6.3 PROPELLANTLESS PROPULSION TECHNOLOGIES
Solar Sails	TBD	–
Tethers	TBD	–
Electric Sails	TBD	–
Aerodynamic Drag	TBD	–

4.6 State-of-the-Art in Small Spacecraft Propulsion

4.6.1 In-Space Chemical Propulsion

Chemical propulsion systems are designed to satisfy high-thrust impulsive maneuvers. They offer lower specific impulse compared to their electric propulsion counterparts but have significantly higher thrust to power ratios.

Hydrazine Monopropellant

1. Technology Description

Hydrazine monopropellant systems use catalyst structures (such as S-405 granular catalyst) to decompose hydrazine to produce hot gases. Hydrazine thrusters and systems have been in extensive use since the 1960’s. The low mass and volume of a significant number of heritage hydrazine propulsion systems makes them suitable for small spacecraft buses, and some hydrazine thrusters used on large spacecraft may be appropriate as the main propulsion system for small spacecraft. Hydrazine thrusters typically achieve a specific impulse between 200 – 235 seconds for 1-N class or larger thrusters.

2. Key Integration and Operational Considerations

• Extensive Flight Heritage: Since hydrazine has been used extensively in spaceflight applications, the technology’s traits are well understood (4).

• Extensive Component Ecosystem: A robust ecosystem of components and experience exists because hydrazine systems are widely used. As such, hydrazine propulsion systems are frequently customized for specific applications using the available components.

• Qualified for Multiple Cold Restarts: These systems have the advantage of typically being qualified for multiple cold starts.

• Extensive Safety and Handling Requirements: Hydrazine and its derivatives are corrosive, toxic, and potentially carcinogenic. Its vapor requires the use of Self Contained Atmospheric Protective Ensemble (SCAPE) suits. This overhead must be considered when planning ground processing workflow for spacecraft and may impose undesirable constraints on the spacecraft, the launch provider, or other spacecraft participating in the same launch opportunity. Hydrazine propulsion systems typically incorporate redundant serial valves to prevent spills or leaking vapor, which might harm ground personnel or hardware.

3. Current & Planned Missions

ArianeGroup has developed a 1-N class hydrazine thruster that has extensive flight heritage, including use on the ALSAT-2 small spacecraft (6) (7). The company also offers a 20-N hydrazine thruster with extensive flight history, although not with small spacecraft.

Aerojet-Rocketdyne (L3 Harris company) has several small hydrazine thrusters with extensive flight histories. The 0.09-N MR-401, 1-N MR-103 series, 4-N MR-111G, and the 22-N MR-106L hydrazine thrusters are flight proven (8) (9) (10). In addition to four MR-107S 222-N thrusters, the OSIRIS-Rex spacecraft propulsion system has six MR-106L, sixteen MR-111G, and two MR-401 thrusters (11).

Moog has the 1-N MONARC-1, 5-N MONARC-5, and 22-N MONARC-22 series hydrazine thrusters (12). These thrusters have decades-long flight histories, although primarily as attitude control thrusters in larger satellites. NASA JPL’s Soil Moisture Active/Passive (SMAP) spacecraft (a small satellite) used eight MONARC-5 thrusters (13).

Northrop Grumman has a line of hydrazine thrusters with long flight histories as attitude control thrusters for larger satellites. The thrust lines include the 1-N MRE-0.1, 5-N MRE-1.0, and 18-N MRE 4.0 (14).

Rafael offers 1-N, 5-N, and 25-N flight proven hydrazine thrusters (15).

IHI Aerospace offers the 1-N MT-9, 4-N MT-8A, and 20-N MT-2 flight proven hydrazine thrusters (17).

Stellar Exploration supplied their Monopropellant CubeSat System for an EchoStar spacecraft and the CAPSTONE spacecraft. The EchoStar nanosatellite built by Tyvak has been successfully commissioned and placed in the altitude prescribed in EchoStar’s license for its S-band frequency (18). The CAPSTONE spacecraft was similarly built by Tyvak and launched on a Rocket Lab Electron from New Zealand on 28 June 2022. The CAPSTONE spacecraft is a 12U, 25 kg CubeSat that will help reduce risk for future lunar spacecraft by validating navigation technologies and verifying the dynamics of a halo-shaped orbit as planned for Gateway, the Moon-orbiting outpost that is part of NASA’s Artemis program. Although CAPSTONE experienced communication and propulsion challenges along the way, the spacecraft achieved the expected lunar orbit (19) (20) (21).

4. Summary Table of Devices

See table 4-2 for current state-of-the-art hydrazine monopropellant devices applicable to small spacecraft.

5. Notable Advances

Aerojet-Rocketdyne, in work sponsored by NASA GSFC, is developing a hybridization of ionic liquid and conventional hydrazine constituents to form the green hydrazine propellant blend (GHPB). “Green hydrazine” would provide the low vapor toxicity and high-density specific impulse of ionic liquids while retaining the low combustion and preheat temperatures of conventional hydrazine. This approach would provide a direct drop-in replacement for hydrazine propulsion systems. In testing completed to date, green hydrazine blends have demonstrated long-term thermal stability/storability, low shock/impact sensitivity, and good operational stability. Furthermore, they have demonstrated a 100-fold reduction in vapor pressure/toxicity and a similar low-temperature start capability as compared to pure hydrazine (22).

Alternative Monopropellants and Bipropellants

1. Technology Description

For the past several decades, Earth storable propellants (i.e., storable at room temperature) have dominated in-space chemical propulsion systems, hydrazine for monopropellants and nitrogen tetroxide with either monomethylhydrazine or hydrazine for bipropellants. These propellants have the advantages of being ambient-temperature liquids (thus not requiring extensive thermal management) and easily ignited (catalyst for hydrazine, hypergolic for bipropellants). However, due to the extensive handling and toxicity concerns of the conventional chemical propellants, more propulsion systems using alternate propellants have been developed and adopted. Often described as “green” propellants, these alternative propellants are not necessarily benign, but are not vapor hazards like the conventional Earth storables and thus do not require specialized suits and breathing apparatus for handling.

Alternative monopropellants based on ionic liquids have received extensive technology development in recent years and have been matured to flight status. Ionic liquid monopropellants are ionic salts dissolved in water and blended with a fuel. Catalysts react the aqueous ionic salt and fuel blend for combustion. Thus, these formulations are not true monopropellants, since they have fuel and oxidizer components, but functionally they are no different than monopropellants.

The two matured ionic liquid monopropellant blends are LMP-103S, which is based on ammonium dinitramide (ADN) or ASCENT (Advanced Spacecraft Energetic Non-Toxic), formally referred to as AF-M315E, which is based on Hydroxylammonium Nitrate (HAN). These monopropellants do not present a vapor hazard and can be handled with conventional personal protection equipment (gloves, face shield). Depending on the formulations, they also can offer higher specific impulse and higher density-specific impulse than monopropellant hydrazine. The challenges with ionic liquid monopropellants are that their catalysts require more preheating than hydrazine and their higher combustion temperatures require higher-temperature, with more expensive catalyst and chamber materials. The ionic liquid monopropellants are not drop-in replacements for hydrazine, although their propulsion system components are similar.

Hydrogen peroxide is a high-density liquid that can be catalytically decomposed exothermically like hydrazine. Hydrogen peroxide does not present a vapor hazard, although high concentrations (> 90%) can rapidly react with impurities in storage containers. Hydrogen peroxide was used extensively in the early days of space propulsion before losing favor to the higher performing hydrazine. It is now being examined again as a nontoxic alternative to hydrazine. Development efforts have focused on high-test peroxide (HTP) which have concentrations from 85 to 98%.

Green bipropellant combinations have also been developed, particularly for small satellite applications. Hydrogen peroxide and nitrous oxide have been explored as the oxidizer options with an alcohol as the fuel. A water electrolysis system breaks water down to gaseous hydrogen and gaseous oxygen, which serve as propellants for the engine. An electrolysis system, using water as the source of propellant, could enable the use of in-situ resources.

2. Key Integration and Operational Considerations

• Improved Hazard Safety Classifications: Air Force Range Safety AFSPCMAN91-710 (23) requirements state that if a propellant is less prone to external leakage, which is often seen with the ionic liquid ‘green’ propellant systems due to higher viscosity of the propellant, then the hazardous classification is reduced. External hydrazine leakage is considered “catastrophic,” whereas using ionic liquid green propellants reduces the hazard severity classification to “critical” and possibly “marginal” per MIL-STD-882E (Standard Practice for System Safety) (24). A classification of “critical” or less only requires two-seals to inhibit external leakage, meaning no additional latch valves or other isolation devices are required in the feed system (24). While these propellants are not safe for consumption, they have been shown to be less toxic compared to hydrazine. This is primarily due to green propellants having lower vapor pressures, being less flammable, and producing more benign constituent product gases (such as water vapor, hydrogen, and carbon dioxide) when combusted.

• Simplified Safety and Handling Requirements: Fueling spacecraft with green propellants, generally permitted as a parallel operation, may require a smaller exclusionary zone, allowing for accelerated launch readiness operations (25). These green propellants are also generally less likely to exothermically decompose at room temperature due to higher ignition thresholds. Therefore, they require fewer inhibit requirements, fewer valve seats for power, and less stringent temperature storage requirements. The reduced hazard associated with some of these propellants may enable projects to take a Design for Minimum Risk (DFMR) approach to address some propulsion system safety concerns, but only with the support of associated range and payload safety entities.

• Immature Component Ecosystem: While there are thrusters that are relatively mature (PMI E/F), incorporating them into integrated propulsion systems is challenging, and the maturity of stand-alone propulsion systems has lagged the pace of component development. Historically, research and development efforts, like Small Business Innovative Research (SBIR) efforts, have focused on component development, and not the entire system. Efforts are now being made to focus on the development of system solutions. Most of these non-toxic propellants are still in some phase of development. Additionally, data on the propellants is widely restricted. Therefore, a comprehensive, public, peer-reviewed databased of compatible materials does not currently exist, and would-be system developers using these propellants may have difficultly accessing such data to guide their efforts.

• Other Considerations for Green Propellants: Other ‘green propellants’ such as Hydrogen Peroxide, High Test Peroxide (HTP), and HTP/Alcohol bipropellants also have their own unique handling considerations. For instance, HTP is a strong oxidizer and can exothermically decompose rapidly if improperly stored or handled. Hydrogen Peroxide, however, has been used as a rocket propellant for many decades, and there is a lot of information on safe handling, materials selection, and best practices. Electrolyzed water is another propellant option, wherein water is decomposed into hydrogen and oxygen and combusted as a traditional bi-propellant thruster. However, generating and managing the power required to electrolyze the water in a compact spacecraft presents its own unique challenges. Yet it does provide a safe-to-launch system with very benign constituents.

3. Current & Planned Missions

ECAPS offers a range of High Performance Green Propulsion (HPGP) thrusters, including the LMP-103S (figure 4.3), at 100-mN, 1-N, 5-N, and 22-N thrust levels. The 1-N HPGP thrusters were first demonstrated on orbit in the Prototype Research Instruments and Space Mission technology Advancement (PRISMA) mission completed in June 2011. PRISMA consisted of two small satellites, each carrying hydrazine and HPGP systems to show off a side-by-side comparison of their performance (26).

ECAPS developed an LMP-103S based propulsion system for a constellation of Earth observing satellites, called SkySat. SkySat has a propulsion system that uses four 1-N HPGP thrusters. Thirteen SkySat satellites with the ECAPS propulsion system have been launched and are fully operational (27).

ECAPS LMP-103S propulsion system is also used on the Autoscale demonstration of rendezvous technologies called ELSA-d, which launched in March 2021. ELSA-d has eight 1-N ECAPS HPGP thrusters to provide re-orbiting and de-orbiting capability. A system issue impacted three of eight ECAPS thrusters and an unresolved root cause resulted in the loss of a fourth thruster. Nevertheless, many mission goals were successfully accomplished, improving the providers readiness for offering a commercial deorbit service (28).

VACCO provided a hybrid Micro Propulsion System (MiPS) for the ArgoMoon 6U CubeSat built for the Italian Space Agency. The ArgoMoon MiPS, figure 4.4, has an ECAPS 100mN LMP-103S thruster with four VACCO 25 mN R134a cold gas thrusters. ArgoMoon was successfully deployed in the Artemis I mission in November 2022 (29) (30) (31).

Aerojet-Rocketdyne built the ASCENT-based propulsion system for the NASA Green Propellant Infusion Mission (GPIM). GPIM was an on-orbit demonstration of the ASCENT (then called AF-M315E) propulsion system, using five 1-N thrusters (figure 4.5) for small attitude control maneuvers (32). GPIM was integrated on a small spacecraft bus (Ball Aerospace’s BCP-100) and launched in June 2019 as a secondary payload on a Falcon Heavy. The five thrusters were successfully fired in space, across a range of operating modes, testing their ability to control the satellite’s attitude, and demonstrating their effectiveness at changing its orbital inclination (33).

An ASCENT propulsion system, the Lunar Flashlight Propulsion System (LFPS) was developed for the JPL Lunar Flashlight mission (34). Lunar Flashlight (figure 4.6) was launched as a secondary payload in a December 2022 Falcon 9 launch, to map the lunar south pole for volatiles. LFPS was a pump-fed monopropellant system, using four 0.1-N ASCENT thrusters (figure 4.7) built by Rubicon Space Systems (a division of Plasma Processes) and a micro-pump built by Flight Works Inc. The propellant management system was fabricated using additive manufacturing. During the first few days of flight, it was found that 3 of the 4 thrusters were underperforming. Based on ground testing it was thought the underperformance might have been caused by obstructions in the fuel lines that limited propellant flow to the thrusters. Improvements were seen by increasing fuel pump pressure to clear the suspected obstructions. However, the effort was not sufficient to keep the spacecraft in the vicinity of the Moon and the mission was terminated in May 2023 (35).

Benchmark Space Systems offers thrusters that can be used with High Test Peroxide (HTP) as a monopropellant or as an oxidizer in a bipropellant combination with a fuel. Benchmark provided a bipropellant Halcyon Avant propulsion system that uses four 22-N thrusters using HTP and isopropyl alcohol for the Sherpa-LTC2 orbital transfer vehicle (OTV) (36). The Sherpa-LTC2 was launched on a Falcon 9 in September 2022. The OTV was used to elevate the orbit of the Varuna Technology Demonstration Mission satellite (37). In June 2021, three monopropellant HTP Halcyon systems launched aboard the SpaceX Transporter 2 rideshare mission. One of the systems debuted Orbit Fab’s RAFTI refueling kit as part of their Tenzing mission. The other two missions supported an undisclosed mission partner (38).

CisLunar Explorer, part of a NASA Centennial Challenges program will use a water electrolysis propulsion system developed by Cornell University’s Space Systems Design Studio on a pair of 3U CubeSat. In this system, the water is electrolyzed in the propellant tank. The gaseous hydrogen/gaseous oxygen mixture is flowed through a flame arrestor into the combustion chamber where it is ignited by a glow plug (39). The CubeSat was originally intended to be launched on Artemis I but difficulties during integration bumped it from the mission (40) (41).

NASA’s Small Spacecraft Technology (SST) program at Ames Research Center (ARC) launched the first Pathfinder Technology Demonstration (PTD) mission in January 2021 (42)(43)(44). PTD-1 (figure 4.9) tested the HYDROS-C water electrolysis propulsion system, developed by Tethers Unlimited Inc. With a volume less than 2.4U, the HYDROS-C uses water as propellant. In-orbit, water was electrolyzed into oxygen and hydrogen, then combusted like a traditional bi-propellant thruster. Limited performance data has been evaluated and made public (45). The system requires 10 – 15 minutes of recharge time between pulses. A variant of the HYDROS-C system is the HYDROS-M system, which is intended to be sized for MicroSats. Tethers Unlimited became a subsidiary of ARKA Group LP in 2020.

NanoAvionics developed an ADN-based monopropellant propulsion system under the Enabling Propulsion System for Small Satellites (EPSS) program. The EPSS monopropellant system was demonstrated on LituanicaSAT-2, a 3U CubeSat, to correct orientation and attitude, avoid collisions, and extend orbital lifetime. LituanicaSAT-2 was launched in June 2017 and successfully separated from the primary payload (Cartosat-2) as part of the European QB50 initiative. According to product literature, multiple missions have since launched, with the latest being in April 2019 (46) (47).

4. Summary Table of Devices

See table 4-3 for the current state-of-the-art in other mono- and bipropellant devices applicable to small spacecraft.

5. Notable Advances

Aerojet Rocketdyne continues to develop its GR-M1 Advanced Green Monopropellant CubeSat Thruster. It employs the same advanced techniques, ultra-high-temperature catalyst, and refractory metal manufacture as the GPIM GR-1 thruster, but on a nanosat scale (48). To partially mitigate thermal management challenges exacerbated at the miniature scale, the GR-M1 is designed to operate on a reduced-flame-temperature variant of the ASCENT propellant containing 10% added water. The heat transfer to surrounding spacecraft structure both during heat up and operation are comparable to conventional hydrazine thrusters.

Rubicon Space Systems (a division of Plasma Processes) is maturing 1N and 5N ASCENT thrusters (49), intended for SmallSat applications. Both offerings are built using the same materials and processes as those used on the 0.1-N thrusters delivered for the Lunar Flashlight Mission. Additionally, Rubicon Space Systems intends to engineer a short-life, lower cost version of the 5N thruster. The prototype thruster accumulated > 1-kg throughput and over 500 seconds before the end of the NASA Phase I SBIR. The Phase II effort will continue to develop the 5N thruster.

CU Aerospace LLC (CUA) has developed the Monopropellant Propulsion Unit for CubeSats (MPUC) system, figure 4.10 (50). The monopropellant is an H₂O₂-ethanol blend denoted as CMP-X. Tests on a thrust stand with a bladder-fed propellant tank have demonstrated continuous constant thrust for >50 minutes at a thrust level of 250 mN at I_SP of 179 s with an average input power of ~6 W during catalyst warmup. 1.5U and 2U system designs have an estimated 1400 N-s and 2120 N-s total impulse, respectively. A ~900°C flame temperature allows the thrust chamber to use non-refractory construction materials. CMP-X has low toxicity and was subjected to UN Series 1, 2, 3, and 6 testing; CMP-X demonstrated no detonation propagation when confined under a charge of high explosive, it exhibited thermal stability with no explosion or detonation during bonfire testing, and was not sensitive to drop impact or friction. CMP-X passed the criteria for either a 1.4S or a “Not Class 1” determination and may be excluded from the explosive class. Long-term storage testing shows no degradation over > 1200 days with testing ongoing. This flight-like, additively-manufactured, thruster passed environmental (vibration and thermal vacuum) testing and post-environmental thrust data were within experimental error of the pre-environmental data (51).

Dawn Aerospace has developed the CubeDrive bipropellant (nitrous oxide and propylene) system that consists of a single B1 thruster, propellant tank, valves, and electronics. The thruster can also be operated in cold gas mode (for smaller impulse bits) by not engaging the spark igniter. The 0.8U (figure 4.11) to 4U configurations provide 425 to 3,500 N-s of total impulse, respectively (52).

VACCO has integrated four ECAPS 1-N LMP-103S thrusters into their Integrated Propulsion System (IPS) (figure 4.12), a bolt-on propulsion module for delta-V and attitude control applications (53) (54).

Hybrids

1. Technology Description

Hybrid propulsion is a mix of both solid and liquid/gas forms of propulsion. In a hybrid rocket, the fuel is typically a solid grain, and the oxidizer (often gaseous oxygen) is stored separately. The rocket is then ignited by injecting the oxidizer into the solid motor and igniting it with a spark or torch system. Since combustion can only occur while the oxidizer is flowing, these systems can readily be started or shut down by controlling the oxidizer flow.

2. Key Integration and Operational Considerations

• Improved Safety and Handling: Hybrid systems are inherently safer to handle than solid motor systems because there is no oxidizer pre-mixed into the solid motor, which reduces the risk of pre-mature ignition.

• Integrates Attributes of Solids and Liquids: Hybrids achieve many positive attributes of both solid motors (storability & handling) and liquid engines (restart & throttling).

• Combustion Efficiency: Combustion efficiency tends to be lower than either solid motors or liquid engines.

• Other Drawbacks: Regression rate control and fuel residuals tend to be more problematic in hybrid designs.

3. Current & Planned Missions

An arc-ignition ’green’ CubeSat hybrid thruster system prototype was developed at Utah State University and demonstrated in flight under the Undergraduate Student Instrument Project (USIP). The hybrid rocket design used a 3D printed acrylonitrile butadiene styrene (ABS) plastic as the fuel and high-pressure gaseous oxygen (GOX) as the oxidizer. For safety considerations the oxidizer was diluted to 60% nitrogen and 40% oxygen for the demonstration. On March 25, 2018, the system was successfully tested aboard a sounding rocket launched from NASA Wallops Flight Facility (WFF) into space and the motor was successfully re-fired 5 times. During the tests, 8 N of thrust and a specific impulse of 215 s were achieved as predicted (55) (56). The Space Dynamics Lab has miniaturized this technology to be better suited for CubeSat applications (0.25 – 0.5 N). A qualification unit is currently in development for the miniaturized system.

4. Summary Table of Devices

See table 4-4 for current state-of-the-art hybrid devices applicable to small spacecraft.

5. Notable Advances

Utah State University has an ongoing test series with Nytrox, a blend of nitrous oxide and oxygen, and ABS. This testing is focused on a 25-50 N system for a 12U sized vehicle. Investigating different nozzle materials for low erosion in long duration burns is a key concern (57) (58).

JPL is developing a hybrid propulsion system for a 12U CubeSat and a 100 kg SmallSat. Testing included regression rate characterization of clear and black Poly (Methyl MethAcrylate) fuels with GOX to be included in propulsion system sizing. Later vacuum testing included an improvement of the ignition system to a laser operated system that eliminates the need for a separate ignition fuel gas (59) (60) (61) (62).

NASA ARC developed a polymethyl methacrylate (PMMA) and nitrous oxide hybrid system that had ethylene and nitrous oxide thrusters. The ethylene and nitrous oxide also function as the hybrid ignition source. The hybrid system had a demonstrated efficiency of 91% and calculated I_SP of 247 sec, making it competitive with current small satellite propulsion systems (63) (64).

Aerospace Corporation and Penn State University developed an “Advanced Hybrid Rocket Motor Propulsion Unit for CubeSats (PUC)”. The design used additive manufacturing techniques for the carbon filled polyamide structure including the nitrous oxide tank and a paraffin grain within an acrylic shell, with acrylic diaphragms 3-D printed in-situ in the grain for enhanced performance. This design fits in a 1U space, for a 3 to 6U spacecraft (65).

Parabilis Space Technologies has developed two small satellite propulsion systems. Rapid Orbital Mobility Bus (ROMBUS) is a hybrid rocket-based system with nitrous oxide as the oxidizer and the attitude control system/reaction control system thruster propellant. It provides high-impulse thrust for satellite translational maneuvers which can be used for initial orbit insertion, rapid orbit rephasing, threat/collision avoidance, and targeted re-entry at the satellite’s mission end of life (66).Nano Orbital Transfer System (OTS) is a Hydroxyl-terminated polybutadiene (HTPB) and nitrous oxide (N₂O) hybrid system, with N₂O based ACS thrusters. Nano OTS leverages Parabilis’ proven hybrid engine and small satellite technologies for low-cost, high-performance maneuvers using non-toxic green propellants. The OTS has a modular design, enabling rapid and low-cost configuration of stages to accommodate 3U size NanoSats up to >50 kg MicroSat-size vehicles (67).

Cold Gas

1. Technology Description

Cold gas propulsion systems are simple, mature, and safe, although they provide relatively limited total impulse. There is no combustion in cold gas systems, with thrust produced by the expulsion of a gaseous propellant through a diverging nozzle. Propellants can be stored as compressed gases, saturated liquids, or solids. Gaseous nitrogen is a commonly used propellant for cold gas systems, although many other gases have been used in the long history of cold gas propulsion. For gases the tradeoff is between performance and storage; lower molecular weight gases offer higher specific impulse but require more voluminous storage. Saturated liquids are stored at low pressure and vaporized when flowed into a low-pressure chamber. The dense storage of saturated liquids and their property of self-pressurization has made them popular as a propellant for CubeSat missions. The liquid propellants may require a pressurant or are self-pressurized, if stored in a two-phase liquid-gas state. Solid propellants can be used in a cold gas system by subliming the solid propellant with adequate heat.

A derivative of cold gas systems is electrothermal or ‘warm gas’ systems, in which the propellant is somewhat heated without chemical reaction and accelerated through a nozzle. The additional heating results in a modest improvement in thrust and specific impulse compared to a pure cold gas system, but typically burdens the spacecraft with increased power consumption. Electrothermal systems are described in more detail in the Electric Propulsion section.

2. Key Integration and Operational Considerations

• Low Cost and Complexity: Cold gas thrusters are often attractive and suitable for small buses due to their relatively low cost and complexity.

• Safe: Most cold gas thrusters use inert, non-toxic propellants, which are an advantage for secondary payloads that must adopt “do no harm” approaches to primary payloads.

• Small Impulse Bit: Cold gas systems are often well suited to provide attitude control since they can provide very small minimum impulse bits for precise maneuvering.

• Small Total Impulse: The low specific impulse of these systems limits them from providing large orbital correction maneuvers.

• Integrated Systems Optimized for CubeSats: Designs optimized around a CubeSat’s limited resources have improved the capability of these systems for nanosatellite buses.

3. Missions

The Micro-Electromechanical-based PICOSAT Satellite Inspector, or MEPSI, built by the Aerospace Corporation flew aboard STS-113 in 2002 and STS-116 in 2006. The spacecraft included both target and imaging/inspector vehicles connected via a tether. The two vehicles were each 4 x 4 x 5 in³ in volume and had five cold-gas thrusters, producing approximately 20 mN. The MEPSI propulsion system was produced using stereo-lithography. It was suited as a propulsion research unit for PicoSats (68).

Marotta developed a cold gas micro-thruster, CGMT-000-9, for fine attitude adjustment maneuvers that flew on the NASA ST-5 mission in 2006. The thruster operated in blowdown mode with gaseous nitrogen, starting at 2.4 N and ending at 0.05 N (69) (70).

In June 2014, Space Flight Laboratory at University of Toronto Institute for Aerospace Research (UTIAS) launched two 15 kg small spacecraft (CanX-4 and CanX-5) to demonstrate formation flying. The Canadian Nanosatellite Advanced Propulsion System (CNAPS), shown in figure 4.13, consisted of four thrusters fueled with liquid sulfur hexafluoride. This propulsion module is a novel version of the previous NanoPS that flew on the CanX-2 mission in 2008. The non-toxic sulfur hexafluoride propellant was selected because it has a high vapor pressure and density, which are important properties for making a compact self-pressurizing system. The propulsion system was successfully used for drift recovery and autonomously maintaining a formation from 1-km range down to 50-m separation (71) (72) (73) (74).

Microspace Rapid Pte Ltd of Singapore developed a cold gas propulsion system for the POPSAT-HIP1 CubeSat demonstration mission that launched June 2014. It consists of eight micro-nozzles that provide control for three rotational axes with a single thrust axis for translational applications. The total delta-v was estimated from laboratory data to between 2.25 and 3.05 ms^-1. Each thruster has 1 mN of nominal thrust using argon propellant. An electromagnetic microvalve with a very short opening time of 1 m-s operates each thruster (75).

GomSpace (acquired NanoSpace in 2016) has developed two related propulsion systems called the NanoProp CGP3 and NanoProp 6U. Both use proportional thrust control of four nozzles to control spacecraft attitude and provide delta-v. The CGP3 was flown on the TW-1 3U CubeSat launched in 2015. The 6U configuration was flown on GOMX-4B in 2018 as a formation flight demonstration (76) (77) (78) (79).

An ACS cold gas propulsion system using R-236fa was produced and tested by Lightsey Space Research for the NASA ARC BioSentinel mission, a 6U CubeSat that launched on Artemis I in November 2022. The propulsion system enables detumbling and pointing for communication back to Earth. The propulsion system uses a 3D-printed propellant tank to reduce part count and make efficient use of the available volume. The system contains six RCS thrusters and one delta-v thruster. The delta-v thruster was included to allow for collision avoidance but was ultimately not needed. One of the RCS thruster valves failed closed during RCS checkout. Rather than further attempting to actuate the valve, and risk the valve failing open, a workaround was identified to perform momentum unloading with the remaining five RCS thrusters. As of May 2023, the propulsion system has accumulated 408 firings and continues to operate as expected. The initial phase of the mission was completed in April 2023. NASA has extended the mission by up to an additional 18 months, or as late as November 2024 (80) (81) (82) (83).

The ThrustMe I2T5 subliming iodine cold gas module, figure 4.14, was the first iodine propulsion system to be spaceflight tested, flown on the Xiaoxiang 1-08 satellite in 2019 (84) (85). Since then, the system was also launched on RTAF’s NAPA-2 in June 2021 and on Spire’s L3C in January 2023. Additional I2T5 are anticipated to launch on the Robusta-3A satellite, developed by CSUM, which will carry various scientific payloads related to meteorology and technology demonstration (86) (87).

The CubeSat Proximity Operations Demonstration (CPOD) is a mission led by Terra Orbital (88) to demonstrate autonomous on-orbit rendezvous and proximity operations using two identical 3U CubeSats. Each spacecraft incorporates a cold gas propulsion system built by VACCO Industries that provides up to 186 N-s of total impulse. This module uses the self-pressurizing refrigerant R236fa propellant, which is exhausted through a total of eight thrusters distributed in pairs at the four corners of the module. CPOD launched in May 2022 as part of SpaceX’s Transporter-5 mission. The CPOD mission demonstrated rendezvous of the CubeSats with intersatellite distances closing from 997 km to a minimum of 0.36 km. On-orbit limitations on experimentation were attributed to multiple factors as the launch date slipped due to many years of delay, including obsolete hardware, partial failure in the solar panels, and a propulsion system anomaly suspected to be a plenum leak (89) (90).

The Mars Cube One (MarCO) technology demonstration mission consisted of two identical CubeSats that followed the InSight spacecraft to Mars in loose formation in 2018. The MarCO spacecraft performed five trajectory correction maneuvers (TCMs) during the mission to Mars. The TCMs were conducted using an R236fa cold gas propulsion system developed by VACCO Industries, which contains four thrusters for attitude control and another four for the TCMs. MarCO-B developed a propulsion system leak which required constant maintenance. A tank to plenum leak was identified prior to launch, which the mission accepted. However, a second leak through a MarCO-B thruster valve developed during flight. The combination of both leaks resulted in a continuous moment on the MarCO-B spacecraft. The MarCO spacecraft succeeded in their mission to relay telemetry from the InSight lander during its descent to Mars (91) (92) (93) (94).

Near-Earth Asteroid Scout (NEA Scout) was a joint MSFC and JPL mission that had a VACCO cold gas MiPS (R236FA propellant), with six 25-mN thrusters, to assist the main propulsion system, a solar sail. However, after deployment from Artemis I in November 2022, the project team was not able to communicate with the spacecraft (95) (96) (97).

4. Summary Table of Devices

See table 4-5 for the current state-of-the-art cold gas devices applicable to small spacecraft.

Solid Motors

1. Technology Description

Solid rocket motors have an oxidizer and fuel mechanical mixture stored in solid form (propellant grain). For small satellites, solid rocket motors may be used for impulsive maneuvers such as orbit insertion or quick de-orbiting. They achieve moderate specific impulses and high thrust magnitudes. There are some electrically controlled solid thrusters that operate in the milli-newton (mN) range that are restartable and have steering capabilities. Solid rocket arrays can be compact and suitable for small buses. Composed of several miniature solid rockets, individual units can be fired, alone or together, as needed.

2. Key Integration and Operational Considerations

• Thrust Vector Control: Thrust vector control systems can be coupled with existing solid rocket motors to provide controllable high delta-v maneuvering.

• Usually Single-Burn: In general, solid motors are considered a single-burn event system. To achieve multiple burns, the system must be either electrically restartable (aka electric solid propellants), or several small units must be matrixed into an array configuration. Because electrically controlled solid propellant (ESPs) are electrically ignited, they are considered safer than traditional solid energetic propellants.

3. Current & Planned Missions

Northrop Grumman offers the STAR 3 motor which has a maximum thrust of 2050 N. The STAR 3 motor was used as the Transverse Impulse Rocket System for the Mars Exploration Rover (MER) program, and was used in 2004 to reduce the lateral velocity of the MER Spirit Lander. The STAR 4G motor was developed and tested by NASA GSFC as an orbit adjust motor for deploying nanosatellite constellations but was not flown, although it is still offered by Northrop Grumman along wth other, higher total impulse motors in the STAR line (98).

The Pacific Scientific Energetic Materials Company (PacSci EMC) Modular Architecture Propulsion System (MAPS) array (figure 4.15) has a 10-plus year in-orbit lifespan. The MAPS system provides three axes capability to control attitude control, deorbit, drag makeup, and plane and attitude changes with a delta-v greater than 50 m s^-1. The capability of MAPS “plug-and-play” bolt-on design and clean-burning propellant array is scalable and can be custom fit for a range of interfaces. MAPS was flown in 2017 aboard PacSciSat, which successfully completed all mission objectives (99) (100).

4. Summary Table of Devices

See table 4-6 for the current state-of-the-art solid motor devices applicable to small spacecraft.

Propellant Management Devices

1. Technology Description

While not directly a thrust producing device, propellant management devices (PMDs) are frequently used in liquid propulsion systems to reliably deliver propellant to thruster units. PMDs are commonly a critical part of in-space liquid propulsion systems that do not use bellows or membrane type tanks. As small spacecraft look toward more complex propulsion system requirements, PMDs will undoubtedly play an integral role. Historically, small spacecraft have used bellows or membrane tanks to ensure propellant delivery and expulsion. However, there is the potential to incorporate PMD structures into additively manufactured tanks and propulsion systems, permitting more conformal structures to be created and optimized for small spacecraft missions. As such, PMDs are briefly covered here for awareness. A more detailed treatment and explanation can be found in literature. A comprehensive, up-to-date list of the types of PMDs, as well as missions employing PMDs, is available in Hartwig (101).

2. Key Integration and Operational Considerations

The purpose of PMDs is to separate liquid and vapor phases within the propellant storage tank upstream of the thruster, and to transfer vapor-free propellant in any gravitational or thermal environment. PMDs have flight heritage with all classical storage systems, have been flown once with LMP-103S, have no flight heritage with cryogenic propellants, and have been implemented in electric propulsion systems. Multiple PMDs are often required to meet the demands of a particular mission, whether using storable or cryogenic propellants.

3. Current & Planned Missions

The Lunar Flashlight Propulsion System used a PMD sponge and ribbon vane. The sponge was additively manufactured, while the ribbon vane was cut from sheet metal and bent to conform to the required dimensions. Surface tension properties, a necessary parameter for PMD sizing, were determined for the ASCENT propellant by Kent State University, funded and managed by NASA. The design and modelling was a joint effort between MSFC and GRC.

4. Summary Table of Devices

No summary table is included for propellant management devices in this report edition.

5. Notable Advances

Northrop Grumman has made advances in the development of SmallSat and CubeSat scale diaphragm propellant tanks using materials that are compatible with hydrazine and some green monopropellant fuels (102), and demonstrated the utility of additive manufacturing in producing tank shells.

4.6.2      In-Space Electric Propulsion

In-space electric propulsion (EP) is any in-space propulsion technology wherein a propellant is accelerated through the conversion of electrical energy into kinetic energy. The electrical energy source powering in-space EP is historically solar, therefore these technologies are often referred to as solar electric propulsion (SEP), although other energy sources such as nuclear reactors or beamed energy are conceivable. The energy conversion occurs by one of three mechanisms: electrothermal, electrostatic, or electromagnetic acceleration (133) (134). Each of these technologies are covered herein.

This survey of the state-of-the-art in EP does not attempt to review all known devices but focuses on those devices that can be commercially procured or devices that appear on a path toward commercial availability. The intent is to aid mission design groups and other in-space propulsion end-users by improving their awareness of the full breadth of potentially procurable EP devices that may meet their mission requirements.

Instead of detailing the complete operating range for each propulsion device, the authors decided to provide only the metrics associated with the nominal operating condition to improve comprehension of the data and make initial device comparisons more straightforward. When a manufacturer does not specifically state a nominal operating condition in literature, the manufacturer may have been contacted to determine a preferred nominal operating condition, otherwise a nominal operating condition was assumed based on similarity to other devices. For those metrics not specifically found in published literature, approximations have been made when calculable from the available data. Readers are strongly encouraged to follow the references cited to the literature describing each device’s full performance range and capabilities.

Electrothermal

1. Technology Description

Electrothermal technologies use electrical energy to increase the enthalpy of a propellant, whereas chemical technologies rely on exothermal chemical reactions. Once heated, the propellant is accelerated and expelled through a conventional converging-diverging nozzle to convert the acquired energy into kinetic energy, like chemical propulsion systems. The specific impulse achieved with electrothermal devices is typically of similar magnitude as chemical devices given that both electrothermal and chemical devices are fundamentally limited by the working temperature limits of materials. However, electrothermal technologies can achieve somewhat higher specific impulses than chemical systems since they are not subject to the limits of chemical energy storage.

Electrothermal devices are typically subclassified within one of the following three categories.

1. Resistojet devices employ an electrical heater to raise the temperature of a surface that in turn increases the temperature of a gaseous propellant.
2. Arcjet devices sustain an electrical arc through an ionized gaseous propellant, resulting in ohmic heating.
3. Electrodeless thrusters heat a gaseous propellant through an inductively or capacitively coupled discharge or by radiation.

2. Key Integration and Operational Considerations

• Propellant Selection: Electrothermal technologies offer some of the most lenient restrictions on propellant selection for in-space propulsion. Whereas chemical systems require propellants with both the right chemical and physical properties to achieve the desired performance, electrothermal systems primarily depend on acceptable physical properties. For example, electrothermal devices can often employ inert gases or even waste products such as water and carbon dioxide. They also allow use of novel propellants such as high storage density refrigerants or in-situ resources. That said, not all propellants can be electrothermally heated without negative consequences. Thermal decomposition of complex molecules may result in the formation of polymers and other inconvenient byproducts. These byproducts may result in clogging of the propulsion system and/or spacecraft contamination.

• Propellant Storage: Electrothermal devices may require that propellants be maintained at a high plenum pressure to operate efficiently. This may require a high-pressure propellant storage and delivery system.

• High Temperature Materials: The working temperature limit of propellant wetted surfaces in the thruster head is a key limitation on the performance of electrothermal devices. As such, very high temperature materials, such as tungsten and molybdenum alloys, are often employed to maximize performance. The total mass and shape of these high temperature materials are a safety consideration during spacecraft disposal. While most spacecraft materials burnup on re-entry, the re-entry behavior of devices using these high temperature materials will be scrutinized when assessing the danger of debris to life and property.

• Power Processing: While some simple resistojet devices may operate directly from spacecraft bus power, other electrothermal devices may require a relatively complex power processing unit (PPU). For example, a radio-frequency electrodeless thruster requires circuitry to convert the direct current bus power to a high-frequency alternating current. In some cases, the cost and integration challenges of the PPU can greatly exceed those of the thruster.

• Thermal Soak-back: Given the high operating temperatures of electrothermal devices, any reliance on the spacecraft for thermal management of the thruster head should be assessed. While the ideal propulsion system would apply no thermal load on the spacecraft, some thermal soak-back to the spacecraft is inevitable, whether through the mounting structure, propellant lines, cable harness, or radiation.

3. Missions

The Bradford (formerly Deep Space Industries) Comet water-based electrothermal propulsion system (figure 4.16) has been implemented by multiple customers operating in low-Earth orbit, including HawkEye 360, Capella Space, and BlackSky Global (135). All missions use the same Comet thruster head, while the BlackSky Global satellites use a larger tank to provide a greater total impulse capability. The HawkEye 360 Pathfinder mission is a constellation of three small microsatellites built by Space Flight Laboratory (SFL) based on its 15-kg NEMO platform; each spacecraft measures 20 x 20 x 44 cm³ with a mass of 13.4 kg (136) (137). The Comet provides each HawkEye 360 pathfinder a total delta-v capability of 96 ms^-1. The approximate dimensions of the BlackSky Global spacecraft are 55 x 67 x 86 cm³ with a mass of 56 kg (139).

The Propulsion Unit for CubeSats (PUC) system (140), figure 4.17, was designed and fabricated by CU Aerospace LLC (Champaign, IL) and VACCO Industries under contract with the U.S. Air Force to supply two government missions (141). The system was acquired for drag makeup capability to extend asset lifetime in low-Earth orbit. The system uses SO₂ as a self-pressurizing liquid propellant. The propulsion system electrothermally heats the propellant using a micro-cavity discharge (MCD) and expels the propellant through a single nozzle (142). It can alternatively use R134a or R236fa propellants, but only in a cold-gas mode with reduced performance. Eight (8) flight units were delivered to the Air Force in 2014, although it remains unknown if any of the units have flown.

In 2019, CU Aerospace was selected for a NASA STMD Tipping Point award to design, fabricate, integrate, and perform mission operations for the Dual Propulsion Experiment (DUPLEX) 6U CubeSat. DUPLEX has two of CU Aerospace’s micro-propulsion systems onboard, one Monofilament Vaporization Propulsion (MVP) system (143) (144) (145), figure 4.18, and one Fiber-Fed Pulsed Plasma Thruster (FPPT) system (149) (150) (151) (152) (153), figure 4.43. The MVP is an electrothermal device that vaporizes and heats an inert solid polymer propellant fiber to 725 K. The coiled solid filament approach for propellant storage and delivery addresses common propellant safety concerns, which often limit the application of propulsion on low-cost CubeSats. In-orbit operations will demonstrate multiple mission capabilities including inclination change, orbit raising and lowering, drag makeup, and deorbit burns. Launch is manifested in early-2024 (155).

AuroraSat-1 is a technology demonstration 1.5U CubeSat that is demonstrating multiple propulsion devices by Aurora Propulsion Technologies. AuroraSat-1 carries Aurora’s smallest version of Aurora Resistojet Module for Attitude control (ARM-A) (156), figure 4.19, and a demonstration unit of their Plasma Brake Module (PBM) (157). The ARM-A system integrated into AuroraSat-1 has six resistojet thrusters for full 3-axis attitude control and 70 grams of water propellant, providing a total impulse of 70 N-s. AuroraSat-1 is built by SatRevolution with Aurora providing the payloads. The satellite was launched by Rocket Lab in May 2022. (158) (159). See section 4.6.3 for discussion of the PBM module.

The OPTIMAL-1 technology demonstration 3U nanosatellite by ArkEdge Space Inc. launched in late 2022 and was deployed from the International Space Station in early 2023. Among other technology demonstration devices, OPTIMAL-1 contains a Pale Blue water resistojet thruster (160).

SPHERE-1 EYE, a 6U CubeSat developed by Sony Group Corporation, includes a Pale Blue Water Resistojet Thruster. The satellite was launched aboard a SpaceX Falcon 9 on January 3^rd, 2023, and has been orbiting Earth with an altitude between 500 km to 600 km. The water propulsion system is expected to prolong the satellite’s life by 2.5 years. The propulsion system operated for approximately 2 minutes on March 3^rd, 2023, and the company confirmed the successful generation of thrust from the obtained flight telemetry (161).

4. Summary Table of Devices

See table 4-7 for current state-of-the-art electrothermal devices applicable to small spacecraft.

Electrosprays

1. Technology Description

Electrospray propulsion systems generate thrust by electrostatically extracting and accelerating ions or droplets from a low-vapor-pressure, electrically conductive, liquid propellant (figure 4.20). This technology can be generally classified into the following types according to the propellant used:

Ionic-Liquid Electrosprays: These technologies use ionic liquids (i.e., salts in a liquid phase at room conditions) as the propellant. The propellant is stored as a liquid, and onboard heaters may be present to maintain propellant properties within the desired operational temperature range. Commonly used propellants include 1-ethyl-3-methylimidazolium tetrafluoroborate (EMI-BF4) and bis(trifluoromethylsulfonyl)imide (EMI-Im). Thrusters that principally emit droplets are also referred to as colloid thrusters.

Field Emission Electric Propulsion (FEEP): These technologies use low-melting-point metals as the propellant. The propellant is typically stored as a solid, and onboard heaters are used to liquefy the propellant prior to thruster operations. Common propellants include indium and gallium.

Feed systems for electrospray technologies can be actively fed via pressurant gas or passively fed via capillary forces. The ion (high-I_SP) or droplet (moderate-I_SP) emission can be controlled by modulation of the high-voltage (i.e., >1 kV) input in a closed-loop feedback system with current measurements. Stable operations in either emission mode can provide very precise impulse bits.

2. Key Integration and Operational Considerations

• Charge Balance: Ionic liquid propellants can support electrospray operations with just cation (i.e., positively charged) emission or bi-polar emission (i.e., both anions and cations). For cation-only emission, as with FEEP thrusters, a separate cathode neutralizer is needed to maintain overall charge balance; such neutralizers do not necessarily need to be tightly integrated with the thruster heads, with resultant mass, volume, and power impacts that must be understood for spacecraft integration. For electrospray technologies using bi-polar emission, reliable high-voltage switching in the power electronics becomes a critical consideration.

• Plume Contamination: Because propellants for electrospray propulsion systems are electrically conductive and condensable as liquids or solids, impingement of the thruster plume on spacecraft surfaces may lead to electrical shorting and surface contamination of solar panels and sensitive spacecraft components. Consideration of only the primary beam plume may not be sufficient, as droplet emission and molecular fragmentation within the plume can generate off-axis species. Plume shields are frequently employed on spacecraft to protect sensitive surfaces in the absence of high-fidelity electrospray plume models or characterization data.

• Propellant Handling and Thruster Contamination: Ionic liquids and metallic propellants can be sensitive to humidity and oxidation, so care is needed if extended storage is required prior to flight. Electrospray technologies can also be sensitive to contamination of the thruster head during propellant loading, ground testing (e.g., backsputter or outgassed materials from the test facility), and handling (i.e., foreign object debris). Precautions should be taken to minimize contamination risks from manufacturing, through test, and to launch. Post-launch, ionic liquids can outgas (e.g., water vapor) when exposed to the space environment, and such behavior should be accounted for in the mission operations. Due to environmental factors such as spacecraft outgassing and atomic oxygen levels, thruster operations following spacecraft deployment may need a conditioning or burn-in period before achieving full propulsive performance.

• Performance Stability and Lifetime: As an electrospray propulsion system operates over time, the propulsive performance can degrade as the plume impinges upon and deposits condensable propellant on thruster head surfaces; in time, sufficiently deposited propellant buildup can electrically short out the thruster electrodes and terminate thruster operation. Especially for missions with large total impulse requirements, lifetime testing or validated life models of the electrospray propulsion system in a relevant environment is important for understanding end-of-life behavior.

• Specific Impulse: Even for electrosprays that principally emit ions, operational thruster modes and instabilities can result in droplet emission that degrade the specific impulse and thrust efficiency. Caution is advised when considering claimed specific impulse or other propulsive properties (e.g., thrust vector and beam divergence) derived from plume characteristics; verification test data in a relevant environment is important for properly assessing these claims.

• Precision Thrust: Electrospray devices have the potential to provide very precise thrust during continuous operations. For devices that can operate in pulsed mode via pulsed modulation of the high-voltage input, fine impulse bits (i.e., <10 μN-s) may be achievable. Such operations can permit precise control over spacecraft attitude and maneuvering. Verification test data in a relevant environment should be used to properly assess the degree of thrust precision.

3. Missions

The ESA Laser Interferometer Space Antenna (LISA) Pathfinder spacecraft was launched in December 2015 on Vega flight VV06. Onboard were two integrated propulsion modules associated with the NASA Space Technology 7 Disturbance Reduction System (ST7 DRS). Each propulsion module contained four independent Busek Colloid MicroNewton Thrusters (CMNT), propellant-less cathode neutralizers, power processing units, digital control electronics, and low-pressure propellant tanks. The propulsion system was successfully commissioned in-orbit in January 2016 after having been fully fueled and stored for almost eight years. The electrospray modules (figure 4.21) were operated at the Earth-Sun Lagrange Point 1 for 90 days to counteract solar disturbance forces on the spacecraft; seven of the eight thrusters demonstrated performance consistent with ground test results, and the full propulsion system met mission-level performance requirements (163).

The Enpulsion Nano FEEP (figure 4.22), formerly the IFM Nano, was first integrated onboard a 3U Planet Labs Flock 3P CubeSat and launched via PSLV-C40 in January 2018. The indium-propellant propulsion system (with integrated thruster head, propellant storage, and power processing unit) was demonstrated in a 491-km by 510-km orbit. Two thruster firing sequences were reported, with the first a 15-minute firing in non-eclipse and the second a 30-minute firing in eclipse. Global Positioning System (GPS) telemetry data onboard the spacecraft indicated good agreement with the ~220 µN commanded thrust (164). Since this initial demonstration, the Enpulsion Nano has flown onboard other spacecraft, but limited on-orbit data is publicly available. These missions include the ICEYE-X2 (launched onboard Falcon-9 flight F9-64 in December 2018) to provide low-Earth orbit interferometric synthetic aperture radar observations (165) (166) and the DOD-funded Harbinger technology demonstrator (launched onboard Electron flight STP-27RD in May 2019) (167) (168). The Enpulsion Nano was also integrated onboard the Zentrum für Telematik (Würzburg) NetSat formation-flying demonstrator mission, which launched as a Soyuz-2 rideshare in September 2020 (169) (170). A summary of available on-orbit statistics, anomalies, and lessons learned for the Enpulsion Nano family (Nano, Nano R³, and Nano AR³) is available (171).

An Enpulsion Micro R³ (figure 4.23) was housed onboard the GMS-T mission, which launched in January 2021 onboard a Rocket Lab Electron. The telecommunications satellite uses an OHB Sweden Innosat platform with propulsive capability for collision avoidance. Inaugural on-orbit commissioning of the propulsion system was confirmed in March 2021 (173).

The University Würzburg Experimental Satellite 4 (UWE-4) was launched as a secondary payload onboard the Soyuz Kanopus-V 5 and 6 mission in December 2018. This 1U spacecraft housed two Morpheus Space NanoFEEP systems, with each system consisting of two gallium-propellant thrusters, a power processing unit board for the UNISEC Europe bus, and a propellant-less cathode neutralizer. An experiment using one thruster as an attitude control actuator was reported, with the increased spacecraft rotation rate corresponding to a derived thrust magnitude of ~5 µN; anomalous torque was attributed to unexpected impingement of the thruster plume upon the spacecraft antenna (174) (175). Orbit lowering capability was demonstrated in 2020; of the four individual thrusters, three experienced anomalous behavior during the UWE-4 mission (176). A 3U-Cubsat implementation of the same NanoFEEP technology is shown in figure 4.24.

4. Summary Table of Devices

See table 4-8 for current state-of-the-art electrospray devices applicable to small spacecraft.

Gridded-Ion

1. Technology Description

Gridded-ion propulsion systems ionize gaseous propellant via a plasma discharge, and the resultant ions are subsequently accelerated via electrostatic grids (i.e., ion optics). This technology can be generally classified according to the type of plasma discharge employed:

• Direct-Current (DC) Discharge: The propellant is ionized via electron bombardment from an internal discharge cathode (figure 4.25).
• Radio-Frequency (RF) Discharge: No internal discharge cathode is present. Instead, the propellant is ionized via RF or microwave excitation from an RF generator (figure 4.26).

Gridded-ion thrusters typically operate at high voltages and include an external neutralizer cathode to maintain plume charge neutrality. High specific impulses can be achieved, but the thrust density is fundamentally limited by space-charge effects. While the earliest thruster technologies used metallic propellants (i.e., mercury and cesium), modern gridded-ion thrusters use noble gases (e.g., xenon, krypton) or iodine.

2. Key Integration and Operational Considerations

• Performance Prediction: Due to the enclosed region of ion generation and acceleration, gridded ion thrusters tend to be less sensitive to test-facility backpressure effects than other devices such as Hall thrusters. This allows for more reliable prediction of in-flight performance based on ground measurements. Furthermore, the separation between ion generation and acceleration mechanisms within the device tend to make calculations of thrust and ion velocity (or I_SP) more straightforward.

• Grid Erosion: Charge-exchange ions formed in between and downstream of the ion optics can impinge upon and erode the grids. Over time, this erosion can lead to a variety of failure modes, including grid structural failure, an inability to prevent electrons from back streaming into the discharge chamber, or the generation of an inter-grid electrical short due to the deposition of electrically conductive grid material. Proper grid alignment must be maintained during thruster assembly, transport, launch, and operations to minimize grid erosion. Random vibration tests at the protoflight level should be conducted to verify the survivability of the ion optics against launch loads. Validated thermal modeling may be needed to assess the impact of grid thermal expansion during thruster operations.

• Foreign Object Debris: The grids are separated by a small gap, typically less than 1 mm, to maximize the electric field and thrust capability of the device. As a result, gridded-ion thrusters tend to be sensitive to foreign object debris, which can bridge the inter-grid gap and cause electrical shorting. Precautions should be taken to minimize such contamination risks from manufacturing, through test, and to launch.

• Cathode Lifetime: Cathodes for plasma discharge or plume neutralization may be sensitive to propellant purity and pre-launch environmental exposure. Feed system cleanliness, bake-out, and use of a high-purity propellant are key factors in maximizing cathode lifetime. The technology provider may recommend maximum cumulative atmospheric exposure and humidity to reduce risk.

• Roll Torque: Misalignments in the ion optics can lead to disturbances in the thrust vector, resulting in a torque around the roll axis that cannot be addressed by the mounting gimbal. For missions requiring extended thruster operations, a secondary propulsion system or reaction wheels may be needed to counter the torque buildup (177).

• Electromagnetic Interactions: For RF-discharge thrusters, electromagnetic interference and compatibility (EMI/EMC) testing may be critical to assess the impact of thruster operations on spacecraft communications and payload functionality.

• Power Electronics: Operation of gridded-ion thrusters requires multiple high-voltage power supplies for discharge operation (ion generation), ion acceleration, and neutralization, leading to potentially complex and expensive power electronics.

• Iodine Propellant: To address the volume constraints of small spacecraft, iodine is an attractive propellant. Compared to xenon, iodine’s storage density is three times greater. Furthermore, iodine stores as a solid with a low vapor pressure, which addresses spacecraft integration concerns associated with high-pressure propellant storage. However, iodine is a strong oxidizer, and long-duration impact on the thruster and potential spacecraft interactions and the long-term reliability of feed system and thruster components exposed to iodine.

3. Missions

The ESA Gravity Field and Steady-State Ocean Circulation Explorer (GOCE) was launched in March 2009 onboard a Rokot / Briz-KM to provide detailed mapping of Earth’s gravitational field and ocean dynamics from an altitude of ~220-260 km. Two QinetiQ T5 DC-discharge gridded-ion thrusters (figure 4.27), with one serving as a redundant backup, successfully provided drag-free control of the 1000-kg satellite until xenon propellant exhaustion in October 2013 (178) (179) (180) (181).

ThrustMe’s NPT30-I2 1U or 1.5U (figure 4.28), an integrated, RF-discharge gridded-ion propulsion system, has at least two units in space as of July 2023. The Beihangkongshi-1 satellite was launched in November 2020 onboard a Long March 6 rocket. As part of the first on-orbit demonstration of iodine-propellant electric propulsion, two 90-minute burns provided an orbit altitude change of 700 m (182) as described in a Nature publication (183). In April 2023, the 1.5U flew onboard a Space Flight Laboratory of the University of Toronto Institute for Aerospace Studies (UTIAS) 35-kg DEFIANT bus for the Norwegian Space Agency’s NorSat-TD mission (184), which included a demonstration of satellite collision avoidance maneuvers (185). With more than 100 systems ordered by clients (186) as of July 2023, upcoming launches include the NPT30-I2 onboard a GomSpace 12U CubeSat for the ESA GOMX-5 technology demonstration mission, NTU Singapore’s INSPIRESAT-4, and Turion Space’s DROID.002 (187).

The Lunar IceCube and LunarH-Map missions both included a Busek BIT-3 propulsion system (figure 4.29) with solid iodine propellant. The BIT-3 system was the primary propulsion for each spacecraft’s lunar transfer, orbit capture, orbit lowering, and spacecraft disposal. Each integrated BIT-3 system includes a low-pressure propellant tank with heated propellant-feed system components, a power processing unit to control the RF thruster and RF cathode, and a two-axis gimbal assembly.

The Lunar IceCube mission, led by Morehead State University, aimed to characterize the distribution of water and other volatiles on the Moon from a highly inclined lunar orbit with a perilune less than 100 km. The 6U spacecraft was launched as a secondary payload onboard Artemis I (188) (189). Communication issues with the CubeSat were reported shortly after launch (190).

The Lunar Polar Hydrogen Mapper (LunaH-Map) mission, led by Arizona State University, aimed to map hydrogen distributions at the lunar south pole from a lunar orbit with a perilune less than 20 km. The 6U spacecraft was launched as a secondary payload onboard Artemis I (191). The mission ended in late-2023 as the BIT-3 was inoperable and unable to perform the required mission maneuvers. The mission team concluded that the propulsion system’s tank valve was stuck closed (192) (193).

4. Summary Table of Devices

See table 4-9 for current state-of-the-art gridded-ion devices applicable to small spacecraft.

Hall-Effect

1. Technology Description

The Hall-effect thruster (HET) is arguably the most successful in-space EP technology by quantity of units flown. The Soviet Union first flew a pair of EDB Fakel SPT-60 HETs on the Meteor-1-10 spacecraft in 1971. Between 1971 and 2018, over 300 additional HETs flew internationally, although EDB Fakel produced the vast majority. The first flight of a non-Russian HET was on board the European Space Agency (ESA) Small Missions for Advanced Research in Technology (SMART-1) spacecraft in 2003. SMART-1 employed the French PPS-1350 HET, produced by Safran (194). The first flight of a U.S. manufactured HET, the Busek BHT-200, was onboard the TacSat-2 spacecraft (195), a U.S. Air Force Research Laboratory (AFRL) experimental satellite in 2006. In 2010, Aerojet, another U.S. entity, began commercially delivering their 4.5 kW XR5 HET (196), formerly BPT-4000. Launches of HETs greatly accelerated in 2019 with the launch of 120 SpaceX Starlink and 6 OneWeb spacecraft (197), each using an HET. As of November 2023, SpaceX has launched over 5,500 Starlink satellites, and OneWeb has launched over 600 satellites. Suffice to say that HETs have become a mainstream in-space propulsion technology.

The rapid growth in demand for HETs can be attributed to their simple design, historically well-demonstrated reliability, good efficiency, high specific impulse, and high thrust-to-power ratio. Although, the higher voltage gridded-ion thrusters (GIT) can achieve even higher specific impulse, HETs can achieve higher thrust-to-power ratios because the HET’s higher density quasi-neutral plasma is not subject to space-charge limitations. The HET’s higher thrust-to-power ratio will typically shorten spacecraft transit time. On the other end of the spectrum, arcjets provide significantly higher thrust than HETs, however material limitations prevent arcjets from matching the HET’s electrical efficiency and specific impulse. For many missions, HETs provide a good balance of specific impulse, thrust, cost, and reliability.

HETs are a form of ion propulsion, ionizing and electrostatically accelerating the propellant. Historically, all HETs flown in space have relied on xenon propellant, given its high molecular weight, low ionization energy, and ease of handling. The recent exception is the SpaceX Starlink spacecraft using krypton propellant. While HETs typically operate less efficiently with krypton propellant, and krypton has more challenging storage requirements, krypton gas is considerably lower cost than xenon gas. Lower cost is a compelling attribute when the potential number of spacecraft are projected in the thousands, as with constellations. While xenon is generally a superior propellant, krypton fed Hall thrusters will likely become more common in the future, especially for large constellations. SpaceX’s 2^nd Generation Starlink spacecraft made another first using argon as a Hall thruster propellant. While argon has been commonly tested with lab HETs due to its low cost and good availability, argon is not generally considered a good choice for spaceflight due to its challenging storage requirements. However, given SpaceX’s need for such large volumes of propellant for the thousands of satellites planned, supply chain challenges may provide good justification for their use of this generally non-ideal propellant. Many other propellants have been considered and ground tested for Hall-effect thrusters, but to date only Hall-effect thrusters using xenon, krypton, or argon have flown.

As schematically shown in figure 4.30, HETs apply a strong axial electric field and radial magnetic field near the discharge chamber exit plane. The E x B force greatly slows the mean axial velocity of electrons and results in an azimuthal electron current many times greater than the beam current. This azimuthal current provides the means by which the incoming neutral propellant is collisionally ionized. These ions are electrostatically accelerated and only weakly affected by the magnetic field. The electron source is a low work function material typically housed in a refractory metal structure (i.e., hollow cathode), historically located external to the HET body. Many recent thruster designs have begun centrally mounting the cathode in the HET body as shown in figure 4.30. The cathode feeds electrons to the HET plasma and neutralizes the plasma plume ejected from the thruster. The high voltage annular anode sits at the rear of the discharge chamber and typically functions as the propellant distribution manifold.

2. Key Integration and Operational Considerations

• Ground Facility Effects: Ground facility effects may result in inconsistencies between ground and flight performance. The significance of the inconsistencies depends on factors such as test facility scale, test facility pumping speed, intrusiveness of diagnostics, and thruster electrical configuration.

• Contamination: Plume ions of an HET can affect spacecraft surfaces by erosion or contamination, even at large plume angles. Ground facility measurement of ion density at large angles may under predict flight conditions.

• Thermal Soak-Back: HET core temperature may exceed 400°C with the cathode exceeding 1000°C. Most HET waste heat radiates directly from the HET surfaces. However, some thermal soak-back to the spacecraft will occur through the mounting structure, propellant feed lines, electrical harness, and radiation.

• Survival Heaters: Given the thermal isolation between the HET and spacecraft, the HET may require a survival heater depending on the qualification temperature and flight environments.

• Performance: HET performance may vary over the life of the device due to erosion and contamination of the plasma wetted HET surfaces. Magnetically shielded thrusters demonstrate less time dependency to their performance than classical HETs.

• Thruster Lifetime: Classical HETs are primarily life-limited by erosion of the discharge chamber wall. Magnetically shielded HETs are primarily life-limited by erosion of the front pole covers.

• Cathode Lifetime: Cathode lifetime may be sensitive to propellant purity and pre-launch environmental exposure. Feed system cleanliness, bake-out, and use of a high purity propellant are key factors in maximizing cathode lifetime. The HET manufacturer may recommend a maximum cumulative atmospheric exposure and humidity. Some cathode emitter formulations are less sensitive to propellant impurities and atmospheric exposure, but these formulations may require other trades such as a higher operating temperature.

• Roll Torque: The E x B force results in a slight swirl torque. For missions requiring extended thruster operations, a secondary propulsion system or reaction wheels may be needed to counter the torque buildup. The roll torque may largely be countered by periodically reversing the direction of the magnetic field. Field reversal requires switching the polarity of current to the magnet coils. Field reversal is only possible with HETs using electromagnets.

• Thrust Vector: Non-uniformity of the azimuthal plasma, magnetic field, or propellant flow may result in slight variations of the thrust vector relative to the HET physical centerline. Temperature variation of the HET, such as during startup, may result in a slight walking of the thrust vector.

• Heaterless Cathodes: Heaterless cathode technologies continue to mature. The benefit of a heaterless cathode is elimination of the cathode heater, typically an expensive component due to rigorous manufacturing and acceptance processes. However, the physics of heaterless cathode life-limiting processes require further understanding. Nevertheless, heaterless cathode demonstrations have empirically shown significant promise. Heaterless cathode requirements on the EP system differ from an HET with a cathode heater. Impacts on the power processing unit and feed system should be well understood when trading a heaterless versus heated cathode.

• Throttling Range: HETs typically throttle stably over a wide range of power and discharge voltage. This makes an HET attractive for missions requiring multiple throttle set-points. However, an HET operates most efficiently at specific throttle conditions. Operating at off-nominal conditions may result in decreased specific impulse and/or electrical efficiency.

3. Missions

Canopus-V (or Kanopus-V) is a Russian Space Agency spacecraft for Earth observation with a design life of 5 years. The 450 kg spacecraft launched in 2012 employed a pair of EDB Fakel SPT-50 thrusters. Similarly, the Canopus-V-IK (Kanopus-V-IK) launched in 2017 with a pair of SPT-50. The SPT-50 thrusters have a long history of spaceflight dating back to the late 1970s. Although the Canopus bus exceeds 450 kg, the power class and physical scale of the SPT-50 are appropriate for smaller spacecraft. The SPT-50 is nominally a 220 W thruster operated on xenon propellant (198) (199) (200). The KazSat-1 and KazSat-2 spacecraft produced by Khrunichev Space Center in cooperation with Thales Alenia Space launched in 2006 and 2011, respectively. The KazSat spacecraft are geosynchronous communication satellites. These spacecrafts employ the EDB Fakel SPT-70BR thruster. The SPT-70BR is Fakel’s latest version of the SPT-70 product line. EDB Fakel optimized the SPT-70 for operation between 600 and 700 W, but no more than 900 W. Experiments demonstrate a lifetime of 3,100 hours, equating to about 450 kNs. The SPT-70 thrusters have a long history of spaceflight dating back to the early 1980s. Control of KazSat-1 was lost in 2008 (201) (202).

The Busek BHT-200 (figure 4.31) has the distinction of being the first U.S.-made HET to operate in space. The BHT-200 has flight heritage from demonstrations on the TacSat-2 mission launched in 2006, FalconSat-5 mission launched in 2010, and FalconSat-6 mission launched in 2018. A Busek PPU powered the 200 W HET for each of the FalconSat missions (203) (204) (205). Ground testing of the BHT-200 includes multiple propellants, although all spaceflights of the BHT-200 have used xenon. Busek developed an iodine compatible derivative of the BHT-200 for the NASA iSat mission. It was determined during the course of the iSat project that additional development related to iodine compatible cathodes was required before conducting an in-space demonstration of the technology (206) (207).

The Israel Space Agency and the French National Center for Space Studies (CNES) jointly developed the Vegetation and Environment monitoring on a New Microsatellite (VENuS) spacecraft launched in 2017. The 268 kg VENuS spacecraft includes a pair of Rafael IHET-300 thrusters (figure 4.32) and 16 kg of xenon propellant. Inflight operations have demonstrated operation between 250 and 600 W. Rafael developed the IHET-300, nominally operating at 300 W, specifically for small spacecraft (208) (209) (210) (211) (212).

The European and Italian space agencies selected the SITAEL HT100 (figure 4.33) for an in-orbit validation program to evaluate the device’s capabilities for orbital maintenance and accelerated reentry of a small spacecraft. The uHETSat mission will be the first in-orbit demonstration of the HT100. The mission will use the thruster to raise and lower its orbit over a series of verification tests conducted during a planned two-year lifespan. The mission seeks to conduct at least one thousand ignition cycles. SITAEL recently completed ground qualification of the complete propulsion system. The HT100 is nominally a 175 W device operating on xenon propellant. The uHETSat will use the SITAEL S-75 microsatellite platform that is 75 kg with dimensions of 60 x 40 x 36 cm³. The spacecraft launched December 1, 2023 (213) (214) (215) (216). The HT100 is also baselined for the Italian Space Agency Platino-1 and Platino-2 spacecraft (217).

The Astra Spacecraft Engine (ASE), figure 4.34, successfully achieved orbital ignition onboard the Spaceflight Sherpa-LTE1 orbital transfer vehicle, which launched on SpaceX’s Transporter-2 mission on June 30, 2021 (218). This single-string system is sized to achieve a controlled de-orbit of Sherpa-LTE1 (219). On-orbit performance was demonstrated by operating the system for 5-minute durations. The first 54 maneuvers have been reported (220). After outgassing, performance metrics were nominal within one standard deviation of ground test data. On-orbit thrust averaged 22.4 mN, and specific impulse for each 5-minute thrust maneuver averaged 1108 seconds. Total propulsion system power processing efficiency averaged 94%, including feed system power, circuit efficiency, and housekeeping circuits (221). As of November 2023, the ASE aboard the Sherpa-LTE1 is continuing mission operations and has operated for more than 800 five-minute maneuvers (i.e., accumulated total duration of ~70 hours) (218). The ASE, (formerly the Apollo Constellation Engine) is a propulsion system that was acquired in Astra’s purchase of Apollo Fusion in 2021. The ASE is designed for operation with xenon and krypton propellants and sized to fit ESPA-class missions. The propulsion system includes several key technologies, including permanent magnets, a heatless instant start cathode, and a radiation hardened PPU. Astra has also reportedly sold ASE units to OneWeb (222) and LeoStella (223), Apex (224), Astroscale (225), and Maxar (226). As of March 2023, Astra has nine ACE in orbit (224).

Exotrail’s first in-orbit propulsion demonstration mission was in November 2020. NanoAvionics and Exotrail partnered to integrate Exotrail’s spaceware – nano (figure 4.35) into a NanoAvionics M6P nanosatellite 6U bus, which was built, integrated, and qualified in 10 months. Following this, Exotrail signed a contract with AAC Clyde Space to provide propulsion for the Eutelsat ELO 3 (2022) and ELO4 (2023) 6U CubeSats (227) (228) (229) (230) (231), and a contract with Aerospace Lab to provide its spaceware – nano for Aerospace Lab’s Risk Reduction Flight (RRF) mission. The Aerospace Lab spacecraft, known as “Arthur,” was launched in 2021. The propulsion system will be used to demonstrate the spacecraft maneuver capabilities (232) (233). Blue Canyon Technologies (BCT) selected the Exotrail’s spaceware system for the company’s Venus-class microsatellite platform, which will be used for the NASA’s INCUS mission, which is expected to launch in 2026 (234) (235).

A spaceware – micro cluster² (figure 4.36) will be integrated onboard York Space Systems S-Class platform for a satellite mission scheduled for launch in late 2023, aiming to orbit the Moon and deliver Earth-to-Moon telecommunication services in support of Intuitive Machines’ lunar south pole mission. With Exotrail’s spaceware – micro cluster², York Space Systems will be able to execute maneuvers such as a lunar transfer orbit (236). Exotrail also has agreements to provide the spaceware – micro for missions by OHB Luxspace (237), Satrec Initiative (238), Astro Digital (239), and Muon Space (240).

Exotrail will also launch its spacevan In-Orbit Demonstration (IOD) mission in October 2023. The spacevan will use Exotrail’s spaceware – micro cluster² to demonstrate its capabilities (e.g., plane change or altitude change) (241).

Airbus will integrate Exotrail’s spaceware – mini thruster, Exotrail’s 300W to 600W class electric propulsion system, onto Airbus’ Earth Observation satellite platform portfolio. Exotrail anticipates completion of the thruster’s qualification activities in 2024 (242).

Blue Canyon, a Raytheon subsidiary, is producing satellites for the DARPA Blackjack program. Blue Canyon selected Exoterra’s Halo thruster (figure 4.37), for its Phase 2 and Phase 3 satellites (243). Exoterra has demonstrated Halo on three Blackjack satellites that launched in June 2023 on the SpaceX Transporter-8 rideshare. These are the first flights of ExoTerra’s Halo electric propulsion system (244). Additionally, ExoTerra has received a NASA Tipping Point award to perform an in-orbit demonstration of their 12U Courier SEP spacecraft bus, tentatively planned for launch in 2024. The bus includes ExoTerra’s Halo thruster, xenon flow control system (XFC), power processing unit (PPU), and deployable solar arrays. The Courier spacecraft provides up to 1 km/s of delta-v, while hosting a 2U, 4 kg payload. The Tipping Point mission objective is to demonstrate the SEP system by spiraling to 800 km from a drop-off orbit of 400 km and then deorbiting. Primary mission objectives include demonstration of the solar array deployment and power generation, PPU efficiency, and 2 kg of thruster propellant throughput. For the Tipping Point mission, the 0.85 kg, 1/3U thruster will nominally operate at 135 W discharge power and produce ~8 mN of thrust (245) (246) (247) (248) (249).

Exoterra Resource is commercializing as Halo12 a version of JPL’s Magnetically Shielded Miniature (MaSMi) Hall thruster technology for multiple unnamed flight opportunties. JPL performed extensive development of the technology between 2011 and 2021, culminating in a 7,205-hour wear test of JPL’s engineering model thruster. The life test processed over 100 kg of xenon propellant and produced a total impulse of approximately 1.55 MN-s. Exoterra’s unnamed flight programs (250) require additional life time, beyond JPL’s demonstration, so the same JPL engineering model thruster was recently installed in JPL’s high-bay vacuum facility to extend the demonstration to 8187 hours and 1.73 MN-s. Other qualification tests conducted by JPL include a 25,000 ignition cycle heaterless cathode demonstration, a 3000-cycle coil TVAC test, and dynamic testing of the thruster (251). It is important to note that the thruster tested by JPL was not produced by Exoterra. However, given that the JPL produced thruster continues to be operated in support of Exoterra’s commercialization activities, it is plausable that Exoterra has retained the key life limiting features of the JPL design (250) (252).

Orbion’s Aurora Hall-effect thruster (figure 4.38) system was selected for a U.S. Space Force 400-kg prototype weather satellite, under contract with General Atomics Electromagnetic Systems (GA-EMS). The Aurora propulsion system consists of a thruster, cathode, power processing unit, propellant flow controller, and cable harness. The system will be used for orbit raising, orbit maintenance, and de-orbit over the 3-5 year mission (253) (254) (255).

Busek has supplied its BHT-350, figure 4.39, Hall-effect thruster to Airbus OneWeb Satellites (AOS) for a range of missions. Busek engineered and qualified the thrusters for orbit raising, orbit maintenance, and end-of-life de-orbit. Eighty OneWeb satellites with BHT-350s launched in December 2022 and January 2023. More than 100 units were operating on OneWeb satellites as of August 2023 with all thrusters reported as operational (256) (257) (258) (259).

Busek shipped its first flight BHT-600 Hall-effect thruster system to a U.S. Government customer in early 2021. The BHT-600 previously demonstrated a 7,000-hour ground test performed at NASA GRC as part of a NASA Announcement for Collaborative Opportunity (ACO) Space Act Agreement (SAA), figure 4.40. The thruster successfully demonstrated 70 kilograms of xenon propellant throughput before the test was terminated (260) (261).

Northrop Grumman’s (NG) Space Systems Sector has developed the NGHT-1X (figure 4.41) thruster for its next generation satellite servicing vehicle known as the Mission Extension Pod (MEP). MEP carries power and propulsion for self-propelled orbit raising as well as station keeping and momentum management when docked to a client vehicle. MEP is designed for a 6-year mission life but can carry a propellant load that permits even longer lifetimes. The NGHT-1X is designed to operate nominally in the 600 to 1,000 W range, but has demonstrated stable operation from 200 to 1,100 W. Within the nominal power range, the thruster is designed to generate a minimum total impulse of 2.1 MN-s, not including margin, to enable the MEP mission. NG partnered with the NASA Glenn Research Center (GRC) to develop and commercialize the NGHT-1X, licensing NASA’s technology for a high propellant throughput, sub-kilowatt hall-effect thruster. NG’s SpaceLogistics has sold all three MEPs that are on the manifest for the first launch on a Falcon 9 rocket planned for 2025 (262) (263) (264) (265) (266) (267) (268).

The EOS SAT-1 developed by Dragonfly Aerospace for the EOS Data Analytics space mission successfully tested its SETS SPS-25 propulsion system. The EOS SAT-1 launched on a Falcon 9 rocket on January 3, 2023 (269) (270).

The Aliena Pte Ltd MUSIC-SI Hall thruster (figure 4.42) was integrated on the NuX-1 3U nanosatellite in July 2021. In-orbit deployment was achieved in January 2022 during SpaceX Transport 3 Mission. Aliena’s Hall thruster became the first to operate onboard such a small satellite, featuring the lowest power consumption for a commercial Hall thruster at 20W (271) (272) (273) (274) (275).

The MUSIC-HM Hall thruster from Aliena and four ARM-A resistojets from Aurora Propulsion Technologies form the Multi-modal Electric Propulsion Engine (MEPE). MEPE was deployed on an Orbital Astronautics 12U satellite (ORB-12 Strider) in July 2023 by a PSLV rocket (PSLV C-56 DS-SAR Mission). The mission will demonstrate the capability of these propulsion systems to support a small satellite constellation. MEPE’s dual electric propulsion technologies allows both the high thrust, low specific impulse of resistojets and low thrust, high specific impulse of Hall thrusters. Both systems share a common propellant, electronics, and fluidics (271) (275) (276).

4. Summary Table of Devices

See table 4-10 for current state-of-the-art HET devices applicable to small spacecraft.

Pulsed Plasma and Vacuum Arc Thrusters

1. Technology Description

Pulsed Plasma Thrusters (PPT) produce thrust by triggering an electric arc between a pair of electrodes that typically ablates a solid-state propellant like polytetrafluoroethylene (PTFE) or ionizes a gaseous propellant. The plasma may be accelerated by either electrothermal or electromagnetic forces. Whether the mechanism of acceleration is electrothermal, electromagnetic, or often some combination thereof, is determined by the device topology (277).

Electrothermal PPTs characteristically include a chamber formed by a pair of electrodes and solid propellant, wherein propellant ablation and heating occurs. During and immediately following each electric discharge, pressure accumulates and accelerates the propellant through a single opening. Electromagnetic PPTs characteristically do not highly confine the propellant as plasma forms. The current pulse, which may exceed tens of thousands of amps, highly ionizes the ablated material or gas. The current pulse further establishes a magnetic field, where the j x B force accelerates the plasma. PPT devices that are predominantly electrothermal typically offer higher thrust, while devices that are predominantly electromagnetic offer higher specific impulse.

The simplest PPTs have no moving parts, which may provide a high degree of reliability. However, as the solid propellant is consumed, the profile of the propellant surfaces constantly changes. Thus, PPTs with static solid propellant demonstrate a change in performance over their life and inherently have a relatively limited lifetime. More complex solid propellant PPTs include a propellant feed mechanism. Typically, the propellant surface profile changes during an initial burn-in period, but then settles into a steady-state behavior where the propellant advancement is balanced by the propellant ablation.

PPT devices are suitable for attitude control and precision pointing applications. PPTs offer small and repeatable impulse bits, which allow for very high precision maneuvering. The complete propulsion system consists of a thruster, an ignitor, and a power processing unit (PPU). Energy to form the pulsed discharge is stored in a high voltage capacitor bank, which often accounts for a significant portion of the system mass. Once the capacitors are charged, resulting in a large differential voltage between the electrodes, the ignitor provides seed material that allows the discharge between the electrodes to form. Various materials and gases (including water vapor) have been tested with PPTs, however PTFE remains most common.

Vacuum arc thrusters (VAT) are another type of pulsed plasma propulsion (278). This technology consists of two metallic electrodes separated by a dielectric insulator. Unlike PPTs, one VAT electrode is sacrificial, providing the propellant source. The mechanism for propellant acceleration is predominantly electromagnetic, resulting in a characteristically high specific impulse and low thrust. One variant of the VAT is predominantly electrostatic, by the inclusion of a downstream electrostatic grid.

2. Key Integration and Operational Considerations

• Safety: PPT capacitor banks often store tens of joules of energy at potentially a couple thousand volts. Follow good electrical safety practices when operating and storing PPTs in a laboratory environment.

• Input Power Range: PPTs and VATs are pulsed devices, which operate by discharging energy stored in capacitors with each pulse. Thus, the propulsion system’s average power draw from the spacecraft bus can be quite low or high depending on the capacitor energy storage and pulse frequency. This flexibility allows PPTs to be applied to spacecraft with limited power budgets of just a few watts, or ample power budgets of hundreds of watts.

• Minimum Impulse Bit: A compelling capability of pulsed devices is the ability to generate small, precise, and well-timed impulse bits for precise spacecraft maneuvering. By controlling the discharge voltage, very small impulse bits on the order of micronewton-seconds are easily achieved.

• Compact and Simple Designs: PPTs and VATs are typically very simple and compact devices. While the total impulse capability is small compared to other forms of EP, these devices offer a particularly attractive solution for CubeSats, where low cost may be a more significant consideration than total impulse. The systems are also attractive for learning environments where propulsion expertise such as high-pressure feed systems and propellant management may be lacking.

• Late-Time Ablation: Although pulsed devices allow for operation over a wide range of pulse frequency, thruster efficiency typically improves with higher pulse rate. Late time ablation is a key inefficiency of solid propellant pulsed devices, where material continues to ablate from the propellant surface well after the discharge pulse. The amount of material accelerated may be maximized through higher frequency pulsing.

• Thrust-to-Power: Pulsed devices suffer from several inefficiencies including late time ablation, frozen flow, and wall heating. Propulsion system efficiency is typically below 20% and may be as low as a few percent. Thus, although pulsed devices may have high specific impulse, the thrust-to-power is low. Small spacecraft with limited power for propulsion may find that large propellant loads provide little benefit as there is inherently a limitation to the number of pulses achievable over the life of the power-limited spacecraft.

• Thermal Soak-back: The low thruster efficiencies may result in large thermal loads on the spacecraft due to thermal soak-back, especially at high rates of pulsing. The spacecraft’s ability to radiate this energy to limit heating may set an upper bound on pulse frequency.

• Ignitor: Pulsed devices usually require some form of ignitor to provide seed material to lower the impedance between the electrodes and initiate the discharge pulse. As such, the lifetime of the ignitor may drive the lifetime of the thruster. Ignitors may fail due to erosion or fouling that prevents sparking. Some devices may include multiple redundant ignitors to increase system lifetime.

• Shorting: The electrodes of pulsed devices are separated by isolating elements. Shadow shielding or other physical features are typically necessary to avoid shorting between electrodes as conductive material ejected by the thruster accumulates. While PTFE is an insulator, the PTFE is reduced to carbon and fluorine when ablated, where carbon accumulation provides a potentially conductive path. VATs employ metal propellants that can similarly result in unintended shorting.

• Spacecraft Contamination: As with any conductive propellant, spacecraft contamination is a concern. Plume interaction with the spacecraft must be understood to assess the impact of the plume on the operation of critical surfaces such as solar panels, antennas, and radiators.

3. Missions

In 2019, CU Aerospace was selected for a NASA STMD Tipping Point award to design, fabricate, integrate, and perform mission operations for the DUPLEX 6U CubeSat having two of CU Aerospace’s micro-propulsion systems onboard, one Monofilament Vaporization Propulsion (MVP) system (143) (144) (145), shown in figure 4.18, and one Fiber-Fed Pulsed Plasma Thruster (FPPT) system (149) (150) (151) (152) (153), shown in figure 4.43. The FPPT can provide a large total impulse primary propulsion for micro-satellites through implementation of a novel PTFE fiber propellant storage and delivery mechanism. A major enhancement of the FPPT technology over classical PPTs is the ability to control both the propellant feed rate and pulse energy, thereby providing control of both the specific impulse and thrust. The FPPT can also provide precision control capability for small spacecraft requiring capabilities such as precision pointing or formation flying. Thrust-vectoring capability of ±10° in the yaw and pitch axes (also with the potential for roll control authority) has been incorporated into the system allowing for limited ACS capability and wheel desaturation for deep space missions. In-orbit operations will demonstrate multiple mission capabilities including inclination change, orbit raising and lowering, drag makeup, and deorbit burns. Launch is manifested in early-2024 (155).

The Benchmark Space Systems’ Xantus thruster (figure 4.44) is a millinewton-class non-toxic metal plasma thruster that is compatible with multiple solid metal propellants. The baseline molybdenum propellant has a measured specific impulse of 1774 seconds. The Xantus thruster first launched in January 2023 onboard the USSF Rapid Revisit Optical Cloud Imager (RROCI) 12U CubeSat demonstration mission. However, the RROCI spacecraft failed to deploy after launch. Xantus is also manifested in 2024 onboard Orion Space Solutions’ EWS (Electro Optical Weather System) mission (279) (280).

Comat’s Plasma Jet Pack (PJP) vacuum arc thruster can be configured with up to four nozzles operating on a solid/inert propellant to provide vectorized thrust (281). The PJP has launched as part of the Vega VV23 mission on the ∑yndeo-2 6U CubeSat. At the time of this publication, no details have yet been reported on the commissioning of the propulsion system (282).

The Neumann Drive is a center-triggered pulsed cathodic arc propulsion technology developed by Neumann Space using molybdenum propellant. The Neumann Drive is offered in multiple configurations offering a range of propulsive performance depending on available spacecraft power (283). The ND-15 thruster, designed for nanosatellites, is currently in orbit undergoing testing on Australia’s Skykraft satellite (284) (285) (286), and another ND-15 will be flown in late 2023 on the University of Melbourne’s SpIRIT nanosatellite (287) (290). The ND-50, a more compact system, is planned for missions in 2024 on the 6U EDISON satellite (291) and 150 kg CarbSAR satellite (292).

4. Summary Table of Devices

See table 4-11 for current state-of-the-art pulsed plasma and vacuum arc devices applicable to small spacecraft.

Ambipolar

1. Technology Description

Ambipolar thrusters ionize gaseous propellant within a discharge cavity via various means, including DC breakdown or RF excitation. The escape of high-mobility electrons from the discharge cavity creates a charge imbalance in the plasma discharge, and the subsequent ambipolar diffusion accelerates ions out of the cavity to generate thrust.

Because the thruster plume is charge neutral, no neutralizer assembly is necessary. A variety of propellants are theoretically usable due to the absence of exposed electrodes (and their associated material compatibility concerns).

2. Key Integration and Operational Considerations

• Propellant Agnostic: While ambipolar thrusters may be operable on a variety of propellants thanks to the devices’ lack of exposed electrodes, different propellants will have different ionization costs (i.e., impact on thruster efficiency), plume behavior, and propellant storage requirements that should be considered during propellant selection.

• Electromagnetic Interactions: For RF-discharge thrusters, electromagnetic interference and compatibility (EMI/EMC) testing may be critical to assess the impact of thruster operations on spacecraft communications and payload functionality.

• Thermal Soakback: Low thruster efficiencies may result in large thermal loads on the spacecraft due to thermal soakback. Validated thermal modeling should be considered to assess impacts to the host spacecraft.

3. Missions

The SpaceX Falcon 9 Transporter-1 launch in January 2021 included two small satellites with the Phase Four Maxwell Block 1 onboard. This integrated propulsion system (figure 4.45) includes the RF thruster and power electronics along with a xenon propellant tank and feed system (293). Phase Four has further reported in 2023 that its Maxwell Block 2, with improved modularity and performance compared to Block 1, has been demonstrated on orbit (294).

The UniSat-7 mission, led by GAUSS, is a 36-kg microsatellite that launched via Soyuz-2-1a Fregat in March 2021. This technology demonstration mission included a T4i iodine-propellant REGULUS (50-I2-Small) module (figure 4.46), an integrated propulsion system that includes thruster, power processing unit, and heated propellant-feed components. The propulsion demonstration is expected to include orbit raising and lowering around the 600-km orbit (296) (297) (298).

A 6U CubeSat from Team Miles was awarded a rideshare slot onboard Artemis I, as one of the winning teams in NASA’s Cube Quest Challenge. The objective of the mission was to demonstrate deep space communications from beyond a 2.5-million mile range. Twelve ConstantQ water-propellant thrusters (figure 4.47), an earlier version of Team Miles’ current M1.4 system, were integrated onboard the CubeSat to provide primary propulsion as well as 3-axis control (299) (300). While the spacecraft was deployed in November 2022, communications contact was not established (405).

4. Summary Table of Devices

See table 4-12 for current state-of-the-art ambipolar devices applicable to small spacecraft.

4.6.3      In-Space Propellant-less Propulsion

Propellant-less propulsion systems generate thrust via interaction with the surrounding environment (e.g., solar photon pressure, planetary magnetic fields, solar wind and ionospheric plasma pressures, and planetary atmospheres). By contrast, chemical and electric propulsion systems generate thrust by expulsion of reaction mass (i.e., propellant). Four propellant-less propulsion technologies have undergone in-space demonstrations to date, including solar sails, tethers, electric sails (and plasma brakes), and aerodynamic drag devices.

Solar Sails

1. Technology Description

Solar sails use solar radiation pressure to generate thrust by reflecting photons via lightweight, highly reflective membranes. While no commercial products are presently available, a handful of missions have sought to demonstrate the technology using small spacecraft.

2. Missions

NASA’s NanoSail-D2, figure 4.48, launched as a 3U CubeSat secondary payload onboard the Fast, Affordable, Science and Technology Satellite (FASTSAT) bus in November 2010. The 10-m² sail made of CP-1 deployed from a 650 km circular orbit and de-orbited the spacecraft after 240 days in orbit (301).

The Planetary Society’s LightSail 2 mission launched as a 3U CubeSat secondary payload on the Department of Defense’s Space Test Program (STP-2) in June 2019. The 32-m² mylar solar sail was deployed at 720-km altitude and demonstrated apogee raising of ~10 km. Its mission was still ongoing as of September 2022 (302).

NASA’s Near-Earth Asteroid (NEA) Scout (figure 4.49) mission launched as a secondary payload onboard Artemis I in November 2022. Unfortunately, contact was never made with the spacecraft, and the sail was not able to be tested. The 6U CubeSat was to deploy an 85-m² solar sail and conduct a flyby of an asteroid within two years of launch (303) (304).

In Space Propulsion SOA chapter

NASA’s Advanced Composite Solar Sail System (ACS3) technology demonstration uses composite materials for solar sail propulsion. The unfurled sail, approximately 9-m2, will be deployed from a 12U CubeSat that is planned for launch in 2024 (305).

Tethers

1. Technology Description

NASA has developed tether technology for space applications since the 1960’s. A tether is nothing more than a long wire, conducting or nonconducting, deployed from a spacecraft in orbit. Two types of tethers and systems can be used for space transportation, with both having had their fundamental physical principles demonstrated in space:

• Electrodynamic Tethers (EDT): A propulsive force of F = IL x B (Lorentz force) is generated on a spacecraft-tether system when a current I from an onboard power supply is fed into a tether of length L against the electromotive force (emf) induced in it by the geomagnetic field B. This concept will work near any planet with a magnetosphere (Earth, Jupiter, etc.). The resulting force may be used to accelerate the deorbit of a spacecraft, boost a spacecraft’s orbital altitude (by reversing the direction of current flow using an onboard power supply), or change its orbital inclination.

• Momentum Exchange Tethers: Using completely different physical principles, long nonconducting tethers can exchange momentum between two masses in orbit to place one body into a higher orbit or a transfer orbit for lunar and planetary missions. Recently completed system studies of this concept indicate that it would be a relatively low-cost in-space asset with long-term, multi-mission capability.

2. Missions

Important milestones include retrieval of a tether in space on the Tethered Satellite System (TSS) mission (TSS-1, 1992), the accidental momentum-exchange boost of the TSS-1R satellite when the tether broke during flight (TSS-1R, 1996), successful deployment of a 20-km-long tether in space (SEDS-1, 1993), closed loop control of tether deployment (SEDS-2, 1994), and operation of an electrodynamic tether with tether current driven in both directions (i.e., power and thrust modes) (PMG, 1993).

More recently, Georgia Institute of Technology’s Prox-1 mission was launched as a secondary payload on the Department of Defense’s Space Test Program (STP-2) in June 2019. The 70 kg spacecraft served as the host and deployer for the LightSail 2 mission. The Prox-1 spacecraft housed a Tethers Unlimited Nanosat Terminator Tape (NSTT), shown in figure 4.50, which deployed a 70-m tether in September 2019 to lower the orbit from 717 km. Data from the Space Surveillance Network indicate that the NSTT caused Prox-1 to deorbit more than 24 times faster than otherwise expected. This rate of orbital decay will enable Prox-1 to meet its 25-year deorbit requirement (306) (307) (308).

The Naval Postgraduate School’s NPSat-1 was launched as a secondary payload on STP-2 and deployed its NSTT in late 2020 (308). TriSept’s DragRacer technology demonstration mission, launched as a rideshare onboard an Electron rocket in November 2020, sought to conduct a direct comparison of the deorbiting rates of two Millennium Space Systems satellites, one of which used a 250-m NSTT (308) (309). A comparison of flight data for operation of the NSTT from each of these three missions has been publicly released (310).

Electric Sails

1. Technology Description

An electric sail (also known as an electric solar wind sail or an E-sail) uses the dynamic pressure of the solar wind as a source of thrust. By using positively charged wires deployed from a spacecraft to create electric fields around each wire, a large ‘virtual’ sail is created that deflects solar wind protons and extracts their momentum. Unlike EDTs, they do not derive thrust using the Lorentz force. Electric sails must be used outside the Earth’s magnetosphere.

A closely related cousin to the electric sail is the plasma brake, an electrostatic deorbit propulsion system that, like the electric sail, uses Coulomb collisions between the charged wire and the ions in the surrounding environment (e.g., the Earth’s ionosphere) to generate an electrostatic force orthogonal to the tether direction, which lowers the spacecraft’s orbital altitude.

These technologies are relatively immature, but researchers in Europe and the USA have been making incremental progress on their development.

2. Missions

The AuroraSat-1 satellite was launched on an Electron rocket on May 5, 2022 (158) (159). The spacecraft was built by SatRevolution with Aurora Propulsion Technologies providing the payloads. The mission serves as a technology demonstration for a Plasma Brake module (157) (figure 4.51) and an Aurora Resistojet Module for Attitude control (ARM-A) (156) (figure 4.19), both produced by Aurora. The Plasma Brake module on AuroraSat-1 is a dual-redundant system for demonstration purposes. A 50-m tether will be deployed to demonstrate its deorbiting capability.

Aerodynamic Drag

1. Technology Description

Satellites have historically deorbited from low-Earth orbits with the aid of thrusters or passive atmospheric drag. Given the increasing rate of new spacecraft launched and the associated potential for new orbital debris following completion of missions, orbital debris management has gained increasing attention. Space debris poses a growing threat to active satellites and human activity in space. Allowing decades for defunct spacecraft to decay naturally from low-Earth orbit is no longer sufficient, as the Federal Communications Commission in late 2022 adopted a draft rule that requires <2000-km altitude spacecraft to deorbit within 5 years of mission completion (312). Aerodynamic drag devices may provide one method to rapidly remove spacecraft from low-Earth orbits upon mission completion.

Below about 1,000 km altitude, the atmosphere exerts a measurable drag force opposite the relative motion of any spacecraft, which results in a slow orbital decay. The intensity of the drag force exerted on the spacecraft depends on numerous factors such as local atmospheric density, the spacecraft forward facing area, the spacecraft velocity, and a drag coefficient. The drag coefficient accounts for the drag force’s dependency on an object’s unique geometric profile. While the spacecraft velocity and local atmospheric density are largely mission dependent, a spacecraft’s forward-facing area and drag coefficient can be altered by introducing aerodynamic drag devices such as exo-brakes and ballutes. These deployable or inflatable parachutes and balloons can greatly increase the drag force exerted on spacecraft by an order of magnitude or more and significantly increase the rate of orbital decay.

Furthermore, aerodynamic drag devices may be useful to reduce spacecraft propellant mass required for orbit capture and disposal at other planetary bodies, given sufficient atmospheric density exists.

For further details on these devices, see chapter on Deorbit Systems.

Table 4-2: Hydrazine Chemical Propulsion

Manufacturer

Product

Propellant

Thrust per Thruster (Quantity)

Specific Impulse

Total Impulse

Mass

Envelope

Power

ACS

PMI Status

Missions

References

—

—

—

[N]

[s]

[kN-s]

[kg]

[cm3 or U]

[W]

Y/N

C,D,E,F

—

—

Integrated Propulsion Systems

Aerojet Rocketdyne USA

MPS-120

Hydrazine

0.25 – 1.0 (4)

2 (2U) 0.8 (1U)

2.5† (2U) 1.6† (1U)

2U 1U

–

Y

D

–

(104)

Aerojet Rocketdyne USA

MPS-125

Hydrazine

0.25 – 1.0 (4)

–

14 (8U) 9.9 (6U) 5.2 (4U)

12.1† (8U) 9.3† (6U) 6.2† (4U)

8U 6U 4U

–

Y

D

–

(104)

Stellar Exploration USA

Monopropellant CubeSat System

Hydrazine

–

200s

–

–

–

–

Y

F

Echostar Global 3, NASA Capstone

(18) (19) (20) (21) (105)

Stellar Exploration USA

Bipropellant CubeSat system

Hydrazine/ NTO

–

285

–

–

–

–

Y

D

–

(105)

Thruster

Aerojet Rocketdyne USA

MR-103

Hydrazine

1

202 – 224

183

0.33 – 0.37

–

< 16

–

F

numerous

(8)

Aerojet Rocketdyne USA

MR-106

Hydrazine

22

228 – 235

561

0.59

–

< 36

–

F

numerous

(8)

Aerojet Rocketdyne USA

MR-111

Hydrazine

4

219 – 229

262

0.37

–

< 16

–

F

numerous

(8)

Aerojet Rocketdyne USA

MR-401

Hydrazine

0.08

182

200

0.60

–

–

–

F

Numerous

(8)

ArianeGroup France

1 N

Hydrazine

1

200 – 223

135

0.29

–

–

–

F

ALSAT-2, numerous

(6) (7)

ArianeGroup France

20 N

Hydrazine

20

222 – 230

517

0.65

–

–

–

F

Numerous

(7)

IHI Aerospace Japan

MT-9

Hydrazine

1

208 – 215

(100 kg)*

–

–

–

–

F

Numerous

(17)

IHI Aerospace Japan

MT-8A

Hydrazine

5

212 – 225

(190 kg)*

–

–

–

–

F

Numerous

(17)

IHI Aerospace Japan

MT-2

Hydrazine

20

210 – 226

(115 kg)*

–

–

–

–

F

Numerous

(17)

Moog USA

MONARC-1

Hydrazine

1

227

111

0.38

113×50 mm

18 (valve)

–

F

numerous

(12)

Moog USA

MONARC-5

Hydrazine

4.5

226

613

0.49

203×380 mm

18 (valve)

–

F

NASA SMAP, numerous

(12) (13)

Moog USA

MONARC-22-6

Hydrazine

22

228

533

0.72

203×380 mm

30 (valve)

–

F

numerous

(12)

Moog USA

MONARC-22-12

Hydrazine

22

228

1,173

0.69

229×530 mm

30 (valve)

–

F

numerous

(12)

Northrop Grumman USA

MRE-0.1

Hydrazine

1

216

(34 kg)*

0.5

114×175 mm

15

–

F

numerous

(14)

Northrop Grumman USA

MRE-1.0

Hydrazine

5

218

(544 kg)*

0.5

114×188 mm

15

–

F

numerous

(14)

Northrop Grumman USA

MRE-4.0

Hydrazine

18

217

(249 kg)*

0.5

61×206 mm

30

–

F

numerous

(14)

Rafael Israel

1N

Hydrazine

1

205 – 214

100

0.31

–

9 (valve)

F

numerous

(15) (16)

Rafael Israel

5N

Hydrazine

6

210 – 220

74

0.31

–

9 (valve)

F

numerous

(15) (16)

Rafael Israel

25N

Hydrazine

28

205 – 220

100

0.53

–

15 (valve)

F

numerous

(15) (16)

Note that all data is documented as provided in the references. Unless otherwise published, do not assume the data has been independently verified. † denotes a wet mass, ‡ denotes a dry mass, “-“ = denotes data not available or applicable, * denotes mass of propellant throughput, rather than total impulse

Table 4-3: Alternative Monopropellant and Bipropellant Propulsion

Manufacturer

Product

Propellant

Thrust per Thruster (Quantity)

Specific Impulse

Total Impulse

Mass

Envelope

Power

ACS

PMI Status

Missions

References

—

—

—

[N]

[s]

[kN-s]

[kg]

[cm3 or U]

[W]

Y/N

C,D,E,F

—

—

Integrated Propulsion Systems

Aerojet Rocketdyne USA

MPS-130

ASCENT

0.25 – 1.0 (4)

–

2.7 (2U) 1.1 (1U)

2.8† (2U) 1.7† (1U)

2U 1U

–

Y

D

–

(103) (104)

Aerojet Rocketdyne USA

MPS-135

ASCENT

0.25 – 1.0 (4)

–

19.4 (8U) 13.7 (6U) 7.3 (4U)

14.7† (8U) 11.2† (6U) 7.2† (4U)

8U 6U 4U

–

Y

D

–

(104)

Aerospace Corp. USA

HyPer

H₂O₂

–

–

–

–

0.25U

–

–

D

–

(106)

Benchmark Space Systems USA

Halcyon

HTP

0.25 1 5 10

150 – 170

(1.56 kg)* (6.6 kg)* (32.5 kg)* (65 kg)*

–

–

< 4

Y

E

Transporter 2 (Undisclosed missions)

(38) (107) (108) (109)

Benchmark Space Systems USA

Halcyon Avant

HTP + Fuel

2 10 22

290 – 310

(6.8 kg)* (34 kg)* (74.8 kg)*

–

–

< 15

Y

F

Sherpa-TLC2

(36) (37) (109)

Benchmark Space Systems USA

Peregrine

HTP & Alcohol

0.1 – 22 (1-8)

270 – 300

5 – 150

–

–

–

–

D

–

(109)

ECAPS Sweden

SkySat 1N HPGP Propulsion System

LMP-103S

1.0 (4)

> 200

21

22†

55 x 55 x 15

10

Y

F

Skysat, PRISMA, Astroscale ELSA-d

(26) (27) (28) (29) (110) (111) (112) (113)

Busek USA

BGT-X5 System

ASCENT

0.5

220 – 225

0.57

1.5†

1U

20

N

D

–

(114) (115) (116)

Cornell University USA

Cislunar Explorer

Water (Electrolysis)

–

–

–

–

6U total (2-units)

–

–

E

CubeQuest Challenge (removed from Artemis 1)

(39) (40) (41)

CU Aerospace USA

MPUC

(CMP-X) Peroxide/ Ethanol blend

0.23

178

1.4 (1.5U) 2.1 (2U)

2.5† (1.5U) 3.3† (2U)

1.5U 2U

3

N

D

–

(50) (51)

Dawn Aerospace New Zealand

CubeDrive

Nitrous Oxide & Propylene

0.5

–

0.425 1.450

1.17† 2.70†

0.8U 2U

–

N

D

–

(52)

Dawn Aerospace New Zealand

SatDrive

Nitrous Oxide & Propylene

0.5 – 16.7

250 – 280

5 – 500

–

–

–

N

D

–

(52)

Moog USA

Monopropellant Propulsion Module

Green or ‘Traditional’

0.5

224

0.5

1.0†

1U

2 x 22.5 W/Thruster

N

D

–

(117)

Rubicon Space Systems USA

Sprite

ASCENT

0.05 – 0.15

215

> 1.2

< 2†

1.5U

2 – 16

N

D

–

(118)

Rubicon Space Systems USA

Phantom

ASCENT

0.3 – 0.5

–

> 9

< 10†

8U

15 – 50

Y

D

–

(119)

NASA MSFC USA

LFPS

ASCENT

0.1 (4)

> 200s

> 3.5

< 5.5†

~2.4U

15 – 47

Y

E

Lunar Flashlight

(34) (35)

NanoAvionics Lithuania

EPSS C1K

IADN-blend

1.0 BOL 0.22 EOL

213

> 0.4

1.2†

1.3U

9.6 (preheat) 1.7 (firing)

N

F

Lituanica-2

(46) (47)

NanoAvionics Lithuania

EPSS C2

IADN-blend

1.0 BOL 0.25 EOL

220

> 1.7

2.6†

2U

–

N

D

–

(46)

Table 4-3 (cont.): Other Monopropellant and Bipropellant Propulsion

Manufacturer

Product

Propellant

Thrust per Thruster (Quantity)

Specific Impulse

Total Impulse

Mass

Envelope

Power

ACS

PMI Status

Missions

References

—

—

—

[N]

[s]

[kN-s]

[kg]

[cm3 or U]

[W]

Y/N

C,D,E,F

—

—

Integrated Propulsion Systems (cont.)

Tethers Unlimited USA (Subsidiary of ARKA Group LP)

HYDROS-C

Water (Electrolysis)

> 1.2

> 310

> 2.1

2.7†

19 x 13 x 9.2

5 – 25

N

F

Pathfinder Technology Demonstration

(44) (45) (120) (121)

Tethers Unlimited USA (Subsidiary of ARKA Group LP)

HYDROS-M

Water (Electrolysis)

> 1.2 (1)

> 310

> 18

13.7†

38.1 dia. x 19.1

7 – 40

N

D

–

(122)

VACCO USA

ArgoMoon Hybrid MiPS

LMP-103S/ R134a

0.1 (1)

190

0.78

2.07†

1.3U

4.3 20 (max)

Y

F

ArgoMoon (Artemis I)

(29) (30) (31)

VACCO USA

Green Propulsion System (MiPS)

LMP-103S

0.1 (4)

190

3.3

5†

3U

15 (max)

Y

D

–

(123)

VACCO USA

Integrated Propulsion System

LMP-103S

1 (4)

200

12

14.7†

11U

50 (max)

Y

D

–

(53) (54)

Thruster Heads

Aerojet Rocketdyne USA

GR-M1

ASCENT

0.25 – 0.5

206

(3.4 kg)*

—

—

14 (max)

–

D

–

(48)

Aerojet Rocketdyne USA

GR-1

ASCENT

0.3 – 1.4

231

(12 kg)*

–

–

18 (max)

–

F

GPIM

(8) (32) (33) (48)

Aerojet Rocketdyne USA

GR-22

ASCENT

8 – 25

250

74

–

–

28

–

F

GPIM

(8) (32) (33)

Benchmark Space Systems USA

Felicette

HTP

0.4 – 1.2

161

> 10

–

–

–

–

E

Transporter 2 (Undisclosed missions)

(38) (109)

Benchmark Space Systems USA

Lynx

HTP

2 – 2.5

295

> 20

–

–

–

–

D

–

(109)

Benchmark Space Systems USA

Ocelot

HTP

18 – 22

302

> 240

–

–

–

–

F

Sherpa-TLC2

(36) (37) (109)

ECAPS Sweden

0.1 N HPGP

LMP-103S

0.03 – 0.1

196 – 209

(1 kg)*

0.04 excl. FCV

–

6.3 – 8

–

F

ArgoMoon

(26) (29) (112)

ECAPS Sweden

1 N HPGP

LMP-103S

0.25 – 1

204 – 235

(24 kg)*

0.38

–

8 – 10

–

F

PRISMA, SkySat, ELSA-d

(26) (27) (28) (29) (112) (113)

ECAPS Sweden

1 N GP

LMP-103S/LT

0.25 – 1

194 – 227

(5 – 8 kg)*

0.38

–

8 – 10

–

D

(113)

ECAPS Sweden

5 N HPGP

LMP-103S

1.5 – 5.5

239 – 253

–

0.48

–

15 – 25

–

D

–

(112)

ECAPS Sweden

22 N HPGP

LMP-103S

5.5 – 22

243 – 255

(150 kg)*

1.1

–

25 – 50

–

D

–

(112)

Dawn Aerospace New Zealand

B20 Thruster

Nitrous Oxide & Propylene

6.1 – 16.7

–

–

0.6

17.6 x 8.0 x 7.9

–

–

F

numerous

(52)

Dawn Aerospace New Zealand

B1 Thruster

Nitrous Oxide & Propylene

0.46 – 1.28

–

–

0.26

10.8 x 7.9 x 4.0

–

–

F

numerous

(52)

Rubicon Space Systems USA

0.1N

ASCENT

0.03 – 0.28

215 – 235

(3.1 kg)*

0.06

–

7 – 9

–

E

Lunar Flashlight

(34) (35) (49)

Rubicon Space Systems USA

1N

ASCENT

0.2 – 1.1

236 – 250

(> 7 kg)*

0.18

–

< 15

–

D

–

(49)

Rubicon Space Systems USA

5N LT

ASCENT

1 – 6

221 – 261

(3 kg)*

0.25

–

< 25

–

D

–

(49)

Rubicon Space Systems USA

5N HT

ASCENT

1 – 6

221 – 261

(110 kg)*

0.30

–

< 25

–

D

–

(49)

Note that all data is documented as provided in the references. Unless otherwise published, do not assume the data has been independently verified. † denotes a wet mass, ‡ denotes a dry mass, “-“ = denotes data not available or applicable, * denotes mass of propellant throughput, rather than total impulse

Table 4-4: Hybrid Chemical Propulsion

Manufacturer

Product

Propellant

Thrust (Quantity)

Specific Impulse

Total Impulse

Mass

Envelope

Power

ACS

PMI Status

Missions

References

—

—

—

[N]

[s]

[kN-s]

[kg]

[cm3 or U]

[W]

Y/N

C,D,E,F

—

—

Aerospace Corp. USA

Propulsion Unit for CubeSats

Paraffin/Nitrous Oxide

–

–

–

–

1U

–

–

D

(65)

JPL USA

Hybrid Rocket

PMMA/GOX

–

> 300

–

–

–

–

–

D

–

(59) (60) (61) (62)

NASA Ames USA

Hybrid Rocket

PMMA/ Nitrous Oxide

25

247

–

–

–

–

–

D

(63) (64)

Parabilis Space Technologies USA

ROMBUS

Various/N2O

222

260

44.6

49

55 dia. x 46

–

Y

D

(66)

Parabilis Space Technologies USA

NanoSat Obrital Transfer System

HTPB/N2O

9.4

245

–

–

–

–

Y

C

(67)

Utah State Univ. USA

Green Hybrid Rocket

ABS/Nytrox

25 – 50

220 – 300

–

–

–

< 30 for 1-2 seconds

–

D

(57) (58)

Utah State Univ. USA

Green Hybrid Rocket

ABS/GOX

8

215

–

–

–

–

–

D

–

(55) (56)

Note that all data is documented as provided in the references. Unless otherwise published, do not assume the data has been independently verified. † denotes a wet mass, ‡ denotes a dry mass, “-“ = denotes data not available or applicable

Table 4-5: Cold Gas Propulsion

Manufacturer

Product

Propellant

Thrust (Quantity)

Specific Impulse

Total Impulse

Mass

Envelope

Power

ACS

PMI Status

Missions

References

—

—

—

[mN]

[s]

[N-s]

[kg]

[cm3 or U]

[W]

Y/N

C,D,E,F

—

—

Integrated Propulsion Systems

Aerospace Corp. USA

MEPSI

R236fa

20

–

–

0.188†

–

–

Y

E

STS-113 and STS-116

(68)

Benchmark Space Systems USA

Starling

Nitrogen

10 – 1000

70

–

0.75†

0.5U

< 4

Y

D

–

(109)

GomSpace Sweden

NanoProp CGP3

Butane

0.01 – 1 (4)

60 – 110

40

0.35†

5 x 10 x 10

< 2

Y

E

TW-1

(76) (78) (79) (126)

GomSpace Sweden

NanoProp 6U

Butane

0.1 – 10 (4)

60 – 110

80

0.900†

2 x 10 x 20

< 2

Y

F

GOMX-4B

(76) (77) (127)

Lightsey Space Research USA

BioSentinel Propulsion System

R236fa

20

47

79.8

1.28†

4 x 10 x 20

< 1 idle, < 4 operating

Y

E

BioSentinel

(80) (81) (82) (83)

Microspace Rapid Pte Ltd Singapore

POPSAT-HIP1

Argon

0.083 – 1.1 (8)

43

–

–

–

–

–

E

POPSAT-HIP1

(75)

ThrustMe France

I2T5

Iodine

0.35

–

75

0.9†

0.5U

5

N

F

Xiaoxiang 1-08, NAPA-2, Spire L3C Demo, Robusta-3A (2024**)

(84) (85) (86) (87)

UTIAS/SFL Canada

CNAPS

Sulfur Hexafluoride

12.5 – 50

40

100

–

–

–

N

F

CanX-4/CanX-5

(71) (72) (73) (74)

VACCO USA

NEA Scout

R236fa

25 (6)

–

500

2.54†

2U

< 55 warmup < 9 operating

Y

E

NEA Scout

(95) (96) (97)

VACCO USA

MiPS Standard

R236fa

25 (4)

–

82 – 515

0.85 – 2.46†

0.4 – 1.5U

< 12

Y

D

–

(128)

VACCO USA

MarCO-A and -B MiPS

R236fa

25 (8)

–

755

3.49†

8 x 15 x 20

–

Y

E

MarCO-A & -B

(91) (92) (93) (94)

VACCO USA

C-POD

R236fa

10 (8)

40

186

1.25†

1U

5

Y

E

CPOD

(88) (89) (90)

Thruster Heads

Marotta USA

CGMT

Nitrogen

105 – 2360

–

–

< 0.06

–

< 7

–

F

NMP ST5

(69) (70)

Moog USA

058E143-146

Nitrogen

10 – 40

60

–

0.04

1.4 x 5.7

< 10

–

F

CHAMP, GRACE

(129)

Moog USA

058E142A

Nitrogen

120

57

–

0.016

1.4 x 2.0

< 35

–

F

Spitzer Space Telescope

(129)

Moog USA

058E151

Nitrogen

120

65

–

0.07

1.9 x 4.1

< 10.5

–

F

Spitzer Space Telescope

(129)

Moog USA

058-118

Nitrogen

3600

57

–

0.023

0.7 x 2.5

< 30

–

F

SAFER, Pluto Fast Flyby

(129)

Moog USA

58E163A

Nitrogen, Xenon, Argon

1300

70 N2, 21 Xe, 54 Ar

–

0.115

2.4 x 5.3

< 10.5

–

F

GEO applications

(129)

Note that all data is documented as provided in the references. Unless otherwise published, do not assume the data has been independently verified. † denotes a wet mass, ‡ denotes a dry mass, “-“ = denotes data not available or applicable, ** anticipated launch date

Table 4-6: Solid Motor Chemical Propulsion

Manufacturer

Product

Propellant

Thrust (Quantity)

Specific Impulse

Total Impulse

Mass

Envelope

Power

ACS

PMI Status

Missions

References

—

—

—

[N]

[s]

[N-s]

[kg]

[cm3 or U]

[W]

Y/N

C,D,E,F

—

—

Integrated Propulsion Systems

PacSci EMC USA

MAPS

–

–

–

–

–

–

–

–

E

PacSciSat

(99) (100)

Thruster Heads

Industrial Solid Propulsion USA

ISP 30 sec. Motor

80% Solids HTPB/AP

37

187

996

0.95†

5.7

–

–

D

Optical target at Kirtland AFB

(131)

Northrop Grumman USA

STAR 3

TP-H-3498

461

266

1,250

1.16†

8 dia. x 29

–

–

E

Mars Exploration Rover Spirit lander

(132)

Northrop Grumman USA

STAR 4G

TP-H-3399

258

269

2,650

1.5†

11.3 dia. x 13.8

–

–

D

–

(132)

Note that all data is documented as provided in the references. Unless otherwise published, do not assume the data has been independently verified. † denotes a wet mass, ‡ denotes a dry mass, “-“ = denotes data not available or applicable

Table 4-7: Electrothermal Electric Propulsion

Manufacturer

Product

Propellant

Thrust*

Specific Impulse*

Total Impulse*

Mass

Envelope

Power

ACS

PMI Status

Missions

References

—

—

—

[mN]

[s]

[N-s]

[kg]

[cm3 or U]

[W]

Y/N

C,D,E,F

—

—

Integrated Propulsion Systems

AIS USA

AIS-SWAG1-PQ

Adamantane

0.12

20

3

0.224†

4.2 x 4.2 x 5.4

5

N

D

–

(311)

AIS USA

AIS-SWAG1-DUO

Adamantane

0.24

20

6

0.449†

8.6 x 4.2 x 5.4

10

N

D

–

(311)

AIS USA

AIS-SWAG1-QUAD

Adamantane

0.48

20

12

0.898†

8.6 x 8.6 x 5.4

20

N

D

–

(311)

Aurora Finland

ARM-A

Water

0.5

100

70

0.280†

0.3U

10£

Y

E

AuroraSat-1 (2022), ORB-12 STRIDER

(156) (158) (159) (276)

Aurora Finland

ARM-C

Water

1

–

–

0.050†

45

12 (max)

N

D

–

(162)

Benchmark Space Systems USA

Starling Ardent

Nitrogen

10 – 1000

70 – 140

–

1.15

1U

10 – 100

Y

D

–

(109)

Bradford Space Netherlands

Comet-1000

H2O

17

175

1,150

1.55†

2,300

50 (max)

N

F

HawkEye 360, Capella Space, Transporter 7

(135) (136) (137) (138)

Bradford Space Netherlands

Comet-8000

H2O

17

175

8,348

6.68†

23,760

55 (max)

N

F

BlackSky

(135) (138) (139)

CU Aerospace USA

CHIPS-180

R134a

15

67

176

1.03†

540

20

Y

D

–

(313) (314) (315) (316)

CU Aerospace USA

CHIPS-500

R134a

25

69

505

1.84†

1300

25

Y

D

–

(313) (314) (315) (316)

CU Aerospace USA

CHIPS-1000

R134a

31

70

1,030

3.13†

2500

30

Y

D

–

(313) (314) (315) (316)

CU Aerospace and VACCO USA

PUC

SO2

4.5

70

184

0.72†

0.35U

15

N

E

8 flight units delivered to AFRL

(140) (141) (142)

CU Aerospace USA

MVP

Delrin Fiber

4.5

66

280

1.06†

0.93U

45

N

E

DUPLEX (launch 2024**)

(143) (144) (145) (146)

EPL USA

APS 100

Ammonia

28

280

2,200

2.63†

2.5U

100

N

D

–

(317)

EPL USA

APS 500

Ammonia

150

350

25,200

13.4†

27,300

450

N

E

Atomos Space (2024**)

(317) (318)

Pale Blue Japan

PBR-9

Water

1.0

45

35

0.6†

0.5U

9

N

D

–

(319)

Pale Blue Japan

PBR-20

Water

1.0

70

200

1.5†

1U

20

N

E

ArkEdge OPTIMAL-1

(160) (319)

Pale Blue Japan

Water Resistojet Propulsion System

Water

2.7

60

170

1.4†

1.25U

< 30

N

E

SPHERE-1 EYE

(161)

SteamJet Space Systems UK

Steam TunaCan Thruster

Water

6

172

219

0.54†

402

< 19.9

N

D

–

(147)

SteamJet Space Systems UK

Steam Thruster One

Water

6

172

–

–

–

19.9

N

D

–

(148)

Thruster Heads

SITAEL Italy

XR-50

Ar, Xe, N2

100

55-85

≤ 72,000

0.22‡

45.8

≤ 50

–

D

–

(320) (321)

SITAEL Italy

XR-100

Ar, Xe, N2

125

63-105

≤ 90,000

0.22‡

45.8

≤ 80

–

D

–

(320) (321)

SITAEL Italy

XR-150

Ar, Xe, N2

250 (Ar) 100 (Xe)

58-110

≤ 180,000

0.22‡

45.8

≤ 95

–

D

–

(320) (321)

Note that all data is documented as provided in the references. Unless otherwise published, do not assume the data has been independently verified. *nominal values (see references for full performance ranges), ** anticipated launch date, † denotes a wet mass, ‡ denotes a dry mass, £ per active thruster, “-“ = denotes data not available or applicable

Table 4-8: Electrospray Electric Propulsion

Manufacturer

Product

Propellant

Thrust*

Specific Impulse*

Total Impulse*

Mass

Envelope

Power

Neutralizer

PMI Status

Missions

References

—

—

—

[mN]

[s]

[N-s]

[kg]

[cm3 or U]

[W]

—

C,D,E,F

—

—

Integrated Propulsion Systems

Busek USA

CMNT (4x heads)

EMI-Im (ionic)

4 x 20

225

980

14.8†

29U

16.5

Carbon Nanotube

F

LISA Pathfinder

(163)

Busek USA

BET-MAX (Config. A)

EMI-Im (ionic)

4 x 55

850

4 x 92§

0.8†

1250

12

Carbon Nanotube

E

US Government

(322) (323) (324) (325) (326) (327) (328)

Busek USA

BET-MAX (Config. B)

EMI-Im (ionic)

4 x 40

2300

4 x 250

0.8†

1250

14

Carbon Nanotube

D

–

(328)

Enpulsion Austria

Nano

Indium (FEEP)

330

1,500 (alpha); 3,000 (gamma)

> 3,000 (alpha); > 5,000 (gamma)

0.90†

10 x 10 x 8.3

40

Thermionic

F

Flock 3p’, ICEYE X2, Harbinger, NetSat, and others

(164) (165) (166) (167) (168) (169) (170) (171) (329) (330) (331)

Enpulsion Austria

Nano R3

Indium (FEEP)

350

1,500 (alpha); 2,700 (gamma)

> 3,000 (alpha); > 5,000 (gamma)

1.4†

9.8 x 9.9 x 9.5

45

Thermionic

E

(Evolution of Nano design)

(171) (172) (332)

Enpulsion Austria

Micro R3

Indium (FEEP)

1,000

2,100 (alpha); 2,750 (gamma)

> 25,000 (alpha); > 35,000 (gamma)

3.9†

14 x 12 x 13.3

105

Thermionic

F

GMS-T

(172) (173) (333) (334)

Enpulsion Austria

Neo

Indium (FEEP)

20,000

1,500

> 550,000

30†

34 x 34 x 15

800

Thermionic

D

–

(335)

Morpheus Space Germany

NanoFEEP (2x heads)

Gallium (FEEP)

< 40

–

–

0.16‡

9 x 2.5 x 4.3

< 3

Propellant-less

E

UWE-4

(174) (175) (336) (337)

Morpheus Space Germany

MultiFEEP (2x heads)

Gallium (FEEP)

< 140

–

–

0.28‡

9 x 4.5 x 4.5

< 19

Propellant-less

D

–

(336)

Note that all data is documented as provided in the references. Unless otherwise published, do not assume the data has been independently verified. *nominal values (see references for full performance ranges), ** anticipated launch date, † denotes a wet mass, ‡ denotes a dry mass, § demonstrated, “-“ = denotes data not available or applicable

Table 4-9: Gridded-Ion Electric Propulsion

Manufacturer

Product

Type

Propellant

Thrust*

Specific Impulse*

Total Impulse*

Mass

Envelope

Power

Cathode Type

PMI Status

Missions

References

—

—

—

[mN]

[s]

[kN-s]

[kg]

[cm3 or U]

[W]

—

C,D,E,F

—

—

Integrated Propulsion Systems

Avant Space Russia

GT-50

RF Ion

Xenon

< 7

–

–

< 8†

< 4U

< 240

Hollow

D

–

(338) (339)

Busek USA

BIT-3

RF Ion

Iodine

1.0

1,960

31.7

2.9†

18 x 8.8 x 10.2

70

RF

E

Lunar IceCube; LunaH-Map

(188) (189) (190) (191) (192) (193) (340) (341) (342) (343) (344) (345)

Pale Blue Japan

PBI-40 Hybrid

RF Ion (Resistojet)

Water (Water)

0.15 (0.9)

> 500 (40)

–

< 2.5†

9 x 12 x 12

28 23

RF

E

RAISE-3 (DDL); RAISE-4 (2024**)

(346) (347) (348) (349)

ThrustMe France

NPT30-I2

RF Ion

Iodine

< 1.3

2,400

5.5 (1U) 14 (1.5U)

1.2† (1U)  1.85† (1.5U)

1U 1.5U

< 81

Thermionic

F

Beihangkongshi-1; NORSAT-TD; INSPIRESAT-4 (2023**); GOMX-5 (2024**); DROID.002 (2024**)

(182) (183) (184) (185) (186) (187) (350) (351) (352)

Thruster Heads

Ariane Group Germany

RIT µX

RF Ion

Xenon

< 0.5

–

–

0.44‡

7.8 x 7.8 x 7.6

< 50

RF

D

–

(353) (354) (355) (356) (357) (358)

Ariane Group Germany

RIT 10 EVO

RF Ion

Xenon

< 15

–

–

1.8‡

18.6 x 18.6 x 13.4

< 435

Hollow

E

(Identical to flight-heritage RIT-10 with contemporary grid design)

(353) (355) (359)

QinetiQ UK

T5

DC Ion

Xenon

< 20

< 3,000

–

2‡

19 x 19 x 24.2

< 600

Hollow

F

GOCE

(178) (179) (180) (181)

Note that all data is documented as provided in the references. Unless otherwise published, do not assume the data has been independently verified. *nominal values (see references for full performance ranges), ** anticipated launch date, † denotes a wet mass, ‡ denotes a dry mass, “-“ = denotes data not available or applicable, RF = Radio Frequency, DC = Direct Current, DDL = Destroyed During Launch

Table 4-10: Hall-Effect Electric Propulsion Thrusters

Manufacturer

Product

Propellant

Thrust*

Specific Impulse*

Total Impulse*

Mass

Envelope

Thruster Power*

Cathode Type

PMI Status

Missions

References

—

—

—

[mN]

[s]

[kN-s]

[kg]

[cm3 or U]

[W]

Notes

C,D,E,F

—

—

Aliena Pte Ltd Singapore

MUSIC-SI

Xenon

< 0.25

< 200

1

2†

1.5U

20

None

E

NuX-1

(271) (272) (273) (274) (275)

Aliena Pte Ltd Singapore

MUSIC-HM

Xenon

3

1,000

15

5†

4U

100

EM-HL

E

ORB-12 STRIDER

(271) (275) (276)

Astra USA

ASE

Xenon

25

1,400

300

1.0

–

400‡

CM-HL

F

Sherpa-LTE, Aries (2024**), ELSA-M (2024**)

(218) (220) (219) (220) (221) (222) (223) (224)

Astra USA

ASE

Krypton

18

1,300

300

1.0

–

400‡

CM-HL

D

–

(218)

Busek USA

BHT-100

Xenon

6.3

1,086

150

1.2

275 wo cath.

105

EM-SH

D

–

(203) (360)

Busek USA

BHT-200

Xenon

13

1,390

84§

1.2

675 wo cath.

250‡

EM-SH

F

TacSat-2, FalconSat-5, -6

(203) (204) (205) (206)

Busek USA

BHT-200-I

Iodine

14

1390

–

1.2

675 wo cath.

250

EM-SH

E

NASA iSat (Cancelled)

(204) (206) (207)

Busek USA

BHT-350

Xenon

17

1,244

212§

1.9

–

350

EM-SH

F

OneWeb Satellites

(256) (257) (258) (259)

Busek USA

BHT-600

Xenon

39

1,500

1000§

2.8

1,470 wo cath.

680‡

EM-SH

E

US Government

(203) (260) (261) (361) (362)

Busek USA

BHT-600-I

Iodine

39

–

–

2.8

1,470 wo cath.

600

EM-SH

D

–

(206) (361) (362) (363)

EDB Fakel Russia

SPT-50

Xenon

14

860

126§

1.2

1,092

220

EM-SH

F

Canopus-V

(198) (199) (200) (201) (364)

EDB Fakel Russia

SPT-50M

Xenon

14.8

930

266

1.3

—

220

EM-SH

D

–

(364)

EDB Fakel Russia

SPT-70BR

Xenon

39

1,470

435§

2.0

1,453

660

EM-SH

F

KazSat-1, KazSat-2

(201) (202)

EDB Fakel Russia

SPT-70M

Xenon

41.3

1,580

–

–

–

660

EM-SH

D

–

(202)

EDB Fakel Russia

SPT-70M

Krypton

31.3

1,460

–

–

–

660

EM-SH

D

–

(202)

ExoTerra USA

Halo

Xenon

19.6

1,294

>375

0.83

375

400‡

CM-HL

F

Tipping Point (2024**), Blackjack

(243) (244) (245) (246) (247) (248) (249)

ExoTerra USA

Halo

Krypton

12

900

>175

0.83

375

300‡

CM-HL

D

–

(365)

ExoTerra USA

Halo12

Xenon

50

1,900

> 2,000

3.4

1,700

1,000‡

CM-HL

E1

Unnamed Flights

(250) (251) (252) (366)

ExoTerra USA

Halo12

Krypton

55

1,575

> 2,000

3.4

1,700

1,000‡

CM-HL

D1

–

(367)

Exotrail France

spaceware nano

Xenon

2.5

800

6

–

2.5U

60

EM-SH

F

M6P Demo, Arthur, ELO3 and ELO4, Otter Pup, INCUS (2026)

(227) (228) (229) (230) (231) (232) (233) (234) (235)

Exotrail France

spaceware micro

Xenon

7

1,000

60

–

960

150

EM-SH

E

York Space (2023**), SpaceVan (2023**), Astro Digital (2024), Satrec Initiative (2025), Muon Space (2026)

(227) (230) (236) (237) (238) (239) (240) (241)

Exotrail France

spaceware mini

Xenon

23

1,300

300

–

–

400

EM-SH

E

Airbus’ Earth Observation satellite platform portfolio

(227) (242)

Note that all data is documented as provided in the references. Unless otherwise published, do not assume the data has been independently verified. *nominal values (see references for full performance ranges), ** anticipated launch date, ‡ PPU input power, § demonstrated, CM = Center Mounted, EM = Externally Mounted, SH = Swaged Heater, HL = Heater-less, JPL = Jet Propulsion Laboratory, SETS = Space Electric Thruster Systems, EDB = Experimental Design Bureau, 1ExoTerra is commercializing the JPL developed MaSMi thruster, “-“ = denotes data not available or applicable

Table 4-10 (cont.): Hall-Effect Electric Propulsion Thrusters

Manufacturer

Product

Propellant

Thrust*

Specific Impulse*

Total Impulse*

Mass

Envelope

Thruster Power*

Cathode Type

PMI Status

Missions

References

—

—

—

[mN]

[s]

[kN-s]

[kg]

[cm3]

[W]

Notes

C,D,E,F

—

—

JPL USA

MaSMi

Xenon

55

1,920

> 1,550§

3.4

1,700

1,000

CM-HL

D

–

(250) (251) (368) (369) (370) (371) (372) (373) (374) (375) (376) (377) (378) (379)

Northrop Grumman USA

NGHT-1X

Xenon

55

1,700

> 2,000

3.1

–

900

CM-SH

E

MEP (2025**)

(262) (263) (264) (265) (266) (267) (268)

Orbion USA

Aurora

Xenon

15

1,320

200

1.5

1,500

250

EM-SH

E

GA-EMS (**)

(253) (254) (255)

Rafael Israel

R-200

Xenon

13

1,160

200

–

–

250

EM-HL

D

–

(208) (380) (381)

Rafael Israel

IHET-300

Xenon

> 14.3

> 1,210

> 135

1.5

1,836

300

EM-SH

F

VENuS

(208) (209) (210) (211) (212)

Rafael Israel

R-800

Xenon

–

–

> 600

–

–

800

EM-HL

D

–

(208) (382)

Safran France

PPS-X00

Xenon

43

1,530

1,000

< 3.2

–

650

EM-SH

D

–

(383) (384) (385)

SITAEL Italy

HT100

Xenon

9

1,300

73

–

407 wo cath.

175

EM-SH

E

uHETSat (2023**)

(213) (214) (215) (216)

SITAEL Italy

HT400

Xenon

27.5

1230

1,000

2.77

1,330

615

EM-SH

D

(386) (387) (388)

SETS Ukraine

ST25

Xenon

7.6

1,000

82

0.95 (with 2 cathodes)

1,003

140

EM-SH

E

EOS SAT-1

(269) (270) (389) (390)

SETS Ukraine

ST40

Xenon

25

1,450

400

1.2 (with 2 cathodes)

1,170

450

EM-HL

D

–

(391)

Note that all data is documented as provided in the references. Unless otherwise published, do not assume the data has been independently verified. *nominal values (see references for full performance ranges), ** anticipated launch date, ‡ PPU input power, § demonstrated, CM = Center Mounted, EM = Externally Mounted, SH = Swaged Heater, HL = Heater-less, JPL = Jet Propulsion Laboratory, SETS = Space Electric Thruster Systems, EDB = Experimental Design Bureau, 1ExoTerra is commercializing the JPL developed MaSMi thruster, “-“ = denotes data not available or applicable

Table 4-11: Pulsed Plasma and Vacuum Arc Electric Propulsion

Manufacturer

Product

Propellant

Thrust*

Impulse Bit

Specific Impulse*

Total Impulse*

Mass

Envelope

Power*

ACS

PMI Status

Missions

References

—

—

—

[mN]

[mNs]

[s]

[N-s]

[kg]

[cm3 or U]

[W]

Y/N

C,D,E,F

—

—

Integrated Propulsion Systems

AIS USA

AIS-VAT1-PQ

Bismuth

26

–

87

0.13

0.056†

37

5

N

D

—

(311)

AIS USA

AIS-VAT1-DUO

Bismuth

52

–

87

0.26

0.084†

74

10

N

D

—

(311)

AIS USA

AIS-VAT1-QUAD

Bismuth

104

–

87

0.52

0.177†

148

20

N

D

—

(311)

Benchmark Space Systems USA

Xantus Metal Plasma Thruster

Molybdenum

400

1

1,774

5,000

1.2†

0.53U

40

N

E

USSF RROCI (not deployed), USSF EWS (2024**)

(392) (279) (280)

Comat France

Plasma Jet Pack

Metal

150

20

2,000

120£

1.2†£

0.66U£

30

N

E

∑yndeo-2

(281) (282) (393)

CU Aerospace USA

FPPT-1.7

PTFE Fiber

500

165

3,200

24,000

3.0†

1.7U

96

Y

E

DUPLEX (2024**)

(149) (150) (151) (152) (153) (154)

Hypernova Space Technologies South Africa

NanoThruster A, Size XS

Solid-state fuel

80

–

≥500

–

0.8†

0.5

10

N

D

—

(394)

Mars Space Ltd UK Clyde Space Sweden

PPTCUP

PTFE

40

40

600

44

0.28†

0.33U

2

N

D

—

(395) (396)

Neumann Space Australia

ND-15

Molybdenum

3.75

45

2,000

880

1.9†

1.5U

15

N

E

Skyride, Apogee

(283) (284) (285) (286) (287) (288) (289) (290)

Neumann Space Australia

ND-50

Molybdenum

100

150

2,000

1,800

1.3†

1U

50

N

E

Edison (2024**), CarbSAR (2024**)

(283) (291) (292)

Note that all data is documented as provided in the references. Unless otherwise published, do not assume the data has been independently verified. *nominal values (see references for full performance ranges), ** anticipated launch date, † denotes a wet mass, ‡ denotes a dry mass, “-“ = denotes data not available or applicable, ESV = Ejector Spring Volume, £ assumes power unit and 2 nozzles

Table 4-12: Ambipolar Electric Propulsion

Manufacturer

Product

Propellant

Thrust*

Specific Impulse*

Total Impulse*

Mass

Envelope

Power

ACS

PMI Status

Missions

References

—

—

—

[mN]

[s]

[kN-s]

[kg]

[cm3]

[W]

Y/N

C,D,E,F

—

—

Integrated Propulsion Systems

Phase Four USA

Maxwell (Block 1)RF

Xenon

7

400

5

5.9‡

19 x 13.5 x 19

450

N

F

Capella

(293) (295) (397) (398) (399) (400)

Phase Four USA

Maxwell (Block 2)RF

Xenon

13

700

–

5.0 (without tank)

22 x 12 x 24 (without tank)

450

N

E

(Mission Not Specified)

(400) (401) (294)

Phase Four USA

Maxwell (Kr)RF

Krypton

13.6

1,110

–

–

–

550

N

D

–

(294) (402)

T4i Italy

REGULUS-50-I2RF

Iodine

0.55

550

3 (Small); 7 (Medium); 11 (Large)

2.5† (Small); 4† (Medium); 5.1† (Large)

9.4 x 9.5 x {15, 18, 20} {Small, Medium, Large}

50

N

E

UniSat-7

(296) (297) (403) (298) (404)

T4i Italy

REGULUS-150-I2RF

Iodine

2

1000

11

6.1†

12 x 12 x 27

150

N

D

–

(404)

Miles Space USA

M1.4

Water

17.2

–

–

1.0†

9.5 x 9.5 x 9.5

6

N

E

Team Miles

(299) (300) (405) (405)

Note that all data is documented as provided in the references. Unless otherwise published, do not assume the data has been independently verified. *nominal values (see references for full performance ranges), ** anticipated launch date, † denotes a wet mass, ‡ denotes a dry mass, “-“ = denotes data not available or applicable, RF = Radio Frequency

Table 4-13: Propellant-less Propulsion

Manufacturer

Product

Propellant

Thrust*

Specific Impulse*

Total Impulse*

Mass

Envelope

Power

ACS

PMI Status

Missions

References

—

—

—

[mN]

[s]

[kN-s]

[kg]

[cm3]

[W]

Y/N

C,D,E,F

—

—

Aurora Propulsion Technologies Finland

Plasma Brake

–

< 100 mN/m

–

–

< 2

1U

< 4

N

E

AuroraSat-1

(157) (158) (159)

Tethers Unlimited USA

NSTT

–

–

–

–

0.81

18 x 18 x 1.8

–

N

F

Prox-1, NPSat-1, DragRacer

(302) (303) (304) (305) (306) (404)(306) (307) (308) (309) (310) (407) (305) (306) (404)

Note that all data is documented as provided in the references. Unless otherwise published, do not assume the data has been independently verified. *nominal values (see references for full performance ranges), ** anticipated launch date, † denotes a wet mass, ‡ denotes a dry mass, “-“ = denotes data not available or applicable See Chapter on Passive Deorbit Systems for review of aerodynamic drag devices.

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