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Chapter Contents

Chapter Glossary.................................................................................................................ii

4.0 In-Space Propulsion............................................................................................... 61

4.1 Introduction.......................................................................................................... 61

Document Organization ............................................................................... 62

4.2 Public Data Sources and Disclaimers ................................................................ 62

4.3 Definitions............................................................................................................ 63

4.4 Technology Maturity............................................................................................ 63

Application of the TRL Scale to Small Spacecraft Propulsion Systems..... 63

Progress Toward Mission Infusion (PMI) .................................................... 64

4.5 Overview of In-Space Propulsion Technology Types........................................ 67

4.6 State-of-the-Art in Small Spacecraft Propulsion ................................................ 69

In-Space Chemical Propulsion .................................................................... 69

In-Space Electric Propulsion........................................................................ 82

In-Space Propellant-less Propulsion ......................................................... 103

References...................................................................................................................... 120
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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 
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(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 
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(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
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4.0 In-Space Propulsion

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.
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Document Organization

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

categories:

1. In-Space Chemical Propulsion (4.6.1)

2. In-Space Electric Propulsion (4.6.2)

3. 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: 

a. Technology Description

b. Key Integration and Operational Considerations

c. Current & Planned Missions

d. Summary Table of Devices

e. Notable Advancements

The organizational approach introduces newcomers to each technology, presents technologyspecific 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.
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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

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 
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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 flightqualification 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.

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 FlightDemonstrated. 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’.
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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. InDevelopment 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, InDevelopment 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 InDevelopment 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. FlightDemonstrated devices typically align with the NASA TRL range of 7 to 9. These devices are highfidelity components or systems (in fit, form, and function) that have been operated in the target 
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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’. 

Figure 4.1: Progress toward mission infusion (PMI) device classifications. Credit: NASA.
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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 
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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 subsystem 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 -
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4.6 State-of-the-Art in Small Spacecraft Propulsion 

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

a. 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.

b. Key Integration and Operational Considerations

• Extensive Flight Heritage: Since hydrazine has been used extensively in spaceflight 

applications, the technology’s traits are well understood (5).

• Extensive Component Ecosystem: A robust ecosystem of components and experience 

exists because hydrazine systems are widely used. As such, hydrazine propulsion 

Figure 4.2: Typical small spacecraft in-space propulsion trade space (thrust vs. specific impulse). 

Credit: NASA.
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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.

c. 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).

d. Summary Table of Devices

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

spacecraft.
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e. 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

a. 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 
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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.

b. 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.
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c. 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 AFM315E) 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).

Figure 4.3: ECAPS HPGP thruster. 

Credit: Bradford ECAPS.

Figure 4.5: GR1 thruster. Credit: Aerojet.

Figure 4.4: ArgoMoon hybrid Micro 

Propulsion System. Credit: VACCO.
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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 

Figure 4.7: Rubicon 0.1N ASCENT thruster. 

Credit: NASA MSFC.

Figure 4.6: Lunar Flashlight Propulsion 

System. Credit: NASA.

Figure 4.8: Halcyon Avant (Sherpa-LTC2

Configuration). Credit Benchmark Space 

Systems.
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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 HYDROSM 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).

d. 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.

e. 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.

Figure 4.9: PTD-1 HYDROS-C. Credit: 

NASA.
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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 ISP 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, additivelymanufactured, thruster passed environmental (vibration 

and thermal vacuum) testing and post-environmental 

thrust data were within experimental error of the preenvironmental 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

a. 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 

Figure 4.12: VACCO Industries IPS.

Credit: VACCO Industries.

Figure 4.11: CubeDrive. Credit: 

Dawn Aerospace.

Figure 4.10: MPUC System. Credit: 

CU Aerospace.
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the oxidizer is flowing, these systems can readily be started or shut down by controlling the 

oxidizer flow.

b. 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.

c. 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.

d. Summary Table of Devices

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

e. 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 

ISP 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).
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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

a. 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.

b. 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.

c. 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 in3in volume and had five cold-gas thrusters, producing approximately 20 mN. The 
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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 POPSATHIP1 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). 

Figure 4.13: CanX-4 and CanX-5 formation 

flying nanosatellites with CNAPS propulsion 

systems. Credit: UTIAS SFL.
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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 selfpressurizing 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).

d. Summary Table of Devices

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

Solid Motors

a. 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 

Figure 4.14: I2T5 Iodine Cold Gas 

Module. Credit: ThrustMe.
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and suitable for small buses. Composed of several miniature solid rockets, individual units can be 

fired, alone or together, as needed.

b. 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.

c. 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-andplay” 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).

d. 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

a. 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).

Figure 4.15: PacSci EMC MAPS 

sealed solid propellant rocket motor 

array. Credit: PacSci.
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b. 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.

c. 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.

d. Summary Table of Devices

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

e. 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. 

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.
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Electrothermal

a. 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.

b. 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.
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• 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.

c. 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 ms1

. 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 lowEarth 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 

Figure 4.16: Comet-1000. Credit: 

Bradford Space.

Figure 4.18: MVP module. Credit: CU 

Aerospace.

Figure 4.17: PUC module. Credit: CU 

Aerospace.
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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 3rd, 

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 3rd, 2023, and the company confirmed the 

successful generation of thrust from the obtained flight telemetry (161).

d. Summary Table of Devices

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

Electrosprays

a. 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.

Figure 4.19: Aurora Resistojet 

Module for Attitude Control. Credit: 

Aurora Propulsion Technologies.
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Feed systems for electrospray 

technologies can be actively fed 

via pressurant gas or passively fed 

via capillary forces. The ion (highISP) or droplet (moderate-ISP) 

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. 

b. 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 cationonly 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.

Figure 4.20: Schematic of typical electrospray emitter and 

electrode configurations. Credit: NASA.
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• 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.

c. 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 

Figure 4.21: Flight CMNT modules 

for LISA Pathfinder. Credit: Busek.

Figure 4.22: Nano. Credit: Enpulsion.
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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 STP27RD 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 onorbit commissioning of the propulsion system was 

confirmed in March 2021 (173).

The University Würzburg Experimental Satellite 4 (UWE4) 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.

d. Summary Table of Devices

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

Gridded-Ion

a. 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).

Figure 4.24: Eight NanoFEEP 

thrusters integrated on 3U-Cubesat 

bus. Credit: Morpheus Space.

Figure 4.23: Micro R3. Credit: 

Enpulsion.
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• 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.

b. 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 ISP) 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 
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spacecraft remain largely unknown. Recent and upcoming flights will provide insight into 

Figure 4.25: Schematic of typical DC-discharge gridded-ion thruster. Credit: NASA.

Figure 4.26: Schematic of typical RF-discharge gridded-ion thruster. Credit: NASA.
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potential spacecraft interactions and the long-term reliability of feed system and thruster 

components exposed to iodine.

c. 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).

Figure 4.29: BIT-3 thruster. Credit: 

Busek.

Figure 4.27: T5 gridded-ion thruster 

for GOCE mission. Credit: QinetiQ.

Figure 4.28: NPT30-I2-1U. Credit:

ThrustMe.
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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).

d. Summary Table of Devices

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

Hall-Effect

a. 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 welldemonstrated 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 2nd 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 
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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.

b. 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 

Figure 4.30: Hall-effect Thruster 

schematic. Credit: NASA.
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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 offnominal conditions may result in decreased specific impulse and/or electrical efficiency.

c. 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 

Figure 4.31: BHT-200 thruster. 

Credit: Busek.
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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 IHET300 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, 

Figure 4.32: IHET-300 thruster. 

Credit: Rafael.

Figure 4.33: HT100 thruster. Credit:

SITAEL.

Figure 4.34: Astra Spacecraft Engine 

(ASE). Credit: Astra.
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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 cluster2to 

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).

Figure 4.35: spaceware-nano

thruster. Credit: Exotrail.

Figure 4.36: spaceware-micro

thruster. Credit: Exotrail.
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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 opportunities. 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 plausible 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).

Figure 4.37: Halo thruster. Credit:

ExoTerra Resource.

Figure 4.38: Aurora HET, PPU and 

XFC. Credit: Orbion Space.
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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 NGHT1X, 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).

Figure 4.40: BHT-600 Installed in 

NASA GRC Vacuum Test Facility. 

Credit: Busek Co.

Figure 4.41: NGHT-1X Engineering 

Model Hall-Effect Thruster. Credit: 

Northrop Grumman.

Figure 4.39: BHT-350 Flight Units. 

Credit: Busek Co.
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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 C56 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).

d. 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

a. 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 burnin 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 

Figure 4.42: NuX-1 satellite with 

MUSIC-SI Hall Thruster. Credit: 

Aliena Pte Ltd
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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.

b. 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 micronewtonseconds 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 
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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.

c. 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).

Figure 4.43: FPPT 1.7U module. 

Credit: CU Aerospace.

Figure 4.44: Xantus Thruster. Credit: 

Benchmark Space Systems.
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National Aeronautics and Space Administration

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).

d. 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

a. 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).

b. 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.

c. 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 Figure 4.45: Maxwell Block 1. Credit 

Phase Four.
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National Aeronautics and Space Administration

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).

d. Summary Table of Devices

See table 4-12 for current state-of-the-art ambipolar

devices applicable to small spacecraft.

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

a. 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. 

Figure 4.46: REGULUS propulsion 

module. Credit: T4i.

Figure 4.47: ConstantQ thruster 

head. Credit: Miles Space.
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National Aeronautics and Space Administration

b. 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).

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

a. 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 

Figure 4.48: Deployment test of the NanoSail

solar sail. Credit: NASA.

Figure 4.49: Deployment test of the NearEarth Asteroid Scout solar sail. Credit: 

NASA.
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National Aeronautics and Space Administration

completed system studies of this concept indicate that it would be a relatively low-cost inspace asset with long-term, multi-mission capability.

b. 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

a. 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.

Figure 4.50: Nanosat Terminator 

Tape (NSTT). Credit: Tethers 

Unlimited.
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National Aeronautics and Space Administration

b. 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

a. 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 lowEarth 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.

Figure 4.51: Plasma Brake Module 

(PBM) demo unit. Credit: Aurora 

Propulsion Technologies.
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National Aeronautics and Space Administration

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 113x50 mm 18 (valve) - F numerous (12)

Moog USA MONARC-5 Hydrazine 4.5 226 613 0.49 203x380 mm 18 (valve) - F NASA SMAP, numerous (12) (13)

Moog USA MONARC-22-6 Hydrazine 22 228 533 0.72 203x380 mm 30 (valve) - F numerous (12)

Moog USA MONARC-22-12 Hydrazine 22 228 1,173 0.69 229x530 mm 30 (valve) - F numerous (12)

Northrop Grumman USA MRE-0.1 Hydrazine 1 216 (34 kg)* 0.5 114x175 mm 15 - F numerous (14)

Northrop Grumman USA MRE-1.0 Hydrazine 5 218 (544 kg)* 0.5 114x188 mm 15 - F numerous (14)

Northrop Grumman USA MRE-4.0 Hydrazine 18 217 (249 kg)* 0.5 61x206 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
108

National Aeronautics and Space Administration

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)

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, BOL = Beginning of Life, EOL = End of Life
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National Aeronautics and Space Administration

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 4.320 (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
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National Aeronautics and Space Administration

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/Nitro

us 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
111

National Aeronautics and Space Administration

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
112

National Aeronautics and Space Administration

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
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National Aeronautics and Space Administration

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
114

National Aeronautics and Space Administration

Table 4-8: Electrospray Electric Propulsion

Manufacturer Product Propellant Thrust*

Specific 

Impulse*

Total 

Impulse*

Mass Envelope Power Neutralizer PMI 

Status Missions References

--- --- --- [N] [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 R

3

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 R3Indium (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

Propellantless 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 Propellantless 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
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National Aeronautics and Space Administration

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; LunaHMap

(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 2823 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
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National Aeronautics and Space Administration

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 E

1 Unnamed Flights (250) (251) (252) (366)

ExoTerra USA Halo12 Krypton 55 1,575 > 2,000 3.4 1,700 1,000‡ CM-HL D

1

- (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
117

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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
118

National Aeronautics and Space Administration

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

--- --- --- [N] [Ns] [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
119

National Aeronautics and Space Administration

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

(306) (307) (308) (309)

(310) (407)

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.
120

National Aeronautics and Space Administration

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