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Ion thruster

An ion thruster, ion drive, or ion engine is a form of electric propulsion used for spacecraft propulsion. An ion thruster creates a cloud of positive ions from a neutral gas by ionizing it to extract some electrons from its atoms. The ions are then accelerated using electricity to create thrust. Ion thrusters are categorized as either electrostatic or electromagnetic.

Electrostatic thruster ions are accelerated by the Coulomb force along the electric field direction. Temporarily stored electrons are reinjected by a neutralizer in the cloud of ions after it has passed through the electrostatic grid, so the gas becomes neutral again and can freely disperse in space without any further electrical interaction with the thruster.

By contrast, electromagnetic thruster ions are accelerated by the Lorentz force to accelerate all species (free electrons as well as positive and negative ions) in the same direction whatever their electric charge, and are specifically referred to as plasma propulsion engines, where the electric field is not in the direction of the acceleration.

Ion thrusters in operation typically consume 1–7 kW of power, have exhaust velocities around 20–50 km/s (Isp 2000–5000s), and possess thrusts of 25–250 mN and a propulsive efficiency 65–80% though experimental versions have achieved 100 kW, 5 N.

The Deep Space 1 spacecraft, powered by an ion thruster, changed velocity by 4.3 km/s while consuming less than 74 kg of xenon. The Dawn spacecraft broke the record, with a velocity change of 11.5 km/s, though it was only half as efficient, requiring 425 kg of xenon.

Applications include control of the orientation and position of orbiting satellites (some satellites have dozens of low-power ion thrusters), use as a main propulsion engine for low-mass robotic space vehicles (such as Deep Space 1 and Dawn), and serving as propulsion thrusters for crewed spacecraft and space stations (e.g. Tiangong).

Ion thrust engines are generally practical only in the vacuum of space as the engine's minuscule thrust cannot overcome any significant air resistance without radical design changes, as may be found in the 'Atmosphere Breathing Electric Propulsion' concept. The Massachusetts Institute of Technology (MIT) has created designs that are able to fly for short distances and at low speeds at ground level, using ultra-light materials and low drag aerofoils. An ion engine cannot usually generate sufficient thrust to achieve initial liftoff from any celestial body with significant surface gravity. For these reasons, spacecraft must rely on other methods such as conventional chemical rockets or non-rocket launch technologies to reach their initial orbit.

Origins

The first person who wrote a paper introducing the idea publicly was Konstantin Tsiolkovsky in 1911. The technique was recommended for near-vacuum conditions at high altitude, but thrust was demonstrated with ionized air streams at atmospheric pressure. The idea appeared again in Hermann Oberth's Wege zur Raumschiffahrt (1929; Ways to Spaceflight), where he explained his thoughts on the mass savings of electric propulsion, predicted its use in spacecraft propulsion and attitude control, and advocated electrostatic acceleration of charged gasses.

A working ion thruster was built by Harold R. Kaufman in 1959 at the NASA Glenn Research Center facilities. It was similar to a gridded electrostatic ion thruster and used mercury for propellant. Suborbital tests were conducted during the 1960s and in 1964, the engine was sent into a suborbital flight aboard the Space Electric Rocket Test-1 (SERT-1). It successfully operated for the planned 31 minutes before falling to Earth. This test was followed by an orbital test, SERT-2, in 1970.

On the 12 October 1964 Voskhod 1 carried out tests with ion thrusters that had been attached to the exterior of the spacecraft.

An alternate form of electric propulsion, the Hall-effect thruster, was studied independently in the United States and the Soviet Union in the 1950s and 1960s. Hall-effect thrusters operated on Soviet satellites from 1972 until the late 1990s, mainly used for satellite stabilization in north–south and in east–west directions. Some 100–200 engines completed missions on Soviet and Russian satellites. Soviet thruster design was introduced to the West in 1992 after a team of electric propulsion specialists, under the support of the Ballistic Missile Defense Organization, visited Soviet laboratories.

General working principle
Ion thrusters use beams of ions (electrically charged atoms or molecules) to create thrust in accordance with momentum conservation. The method of accelerating the ions varies, but all designs take advantage of the charge/mass ratio of the ions. This ratio means that relatively small potential differences can create high exhaust velocities. This reduces the amount of reaction mass or propellant required, but increases the amount of specific power required compared to chemical rockets. Ion thrusters are therefore able to achieve high specific impulses. The drawback of the low thrust is low acceleration because the mass of the electric power unit directly correlates with the amount of power. This low thrust makes ion thrusters unsuited for launching spacecraft into orbit, but effective for in-space propulsion over longer periods of time.

Ion thrusters are categorized as either electrostatic or electromagnetic. The main difference is the method for accelerating the ions.
 * Electrostatic ion thrusters use the Coulomb force and accelerate the ions in the direction of the electric field.
 * Electromagnetic ion thrusters use the Lorentz force to accelerate the ions in the direction perpendicular to the electric field.

Electric power for ion thrusters is usually provided by solar panels. However, for sufficiently large distances from the sun, nuclear power may be used. In each case, the power supply mass is proportional to the peak power that can be supplied, and both provide, for this application, almost no limit to the energy.

Electric thrusters tend to produce low thrust, which results in low acceleration. Defining $$1g = 9.81\; \mathrm{m/s^2}$$, the standard gravitational acceleration of Earth, and noting that $$F = ma \implies a = F/m$$, this can be analyzed. An NSTAR thruster producing a thrust force of 92 mN will accelerate a satellite with a mass of 1ton by 0.092N / 1000 kg = 9.2m/s$2$ (or 9.38g). However, this acceleration can be sustained for months or years at a time, in contrast to the very short burns of chemical rockets.

$$F = 2 \frac{\eta P}{g I_\text{sp}}$$ Where:
 * F is the thrust force in N,
 * η is the efficiency
 * P is the electrical power used by the thruster in W, and
 * Isp is the specific impulse in seconds.

The ion thruster is not the most promising type of electrically powered spacecraft propulsion, but it is the most successful in practice to date. An ion drive would require two days to accelerate a car to highway speed in vacuum. The technical characteristics, especially thrust, are considerably inferior to the prototypes described in literature, technical capabilities are limited by the space charge created by ions. This limits the thrust density (force per cross-sectional area of the engine). Ion thrusters create small thrust levels (the thrust of Deep Space 1 is approximately equal to the weight of one sheet of paper ) compared to conventional chemical rockets, but achieve high specific impulse, or propellant mass efficiency, by accelerating the exhaust to high speed. The power imparted to the exhaust increases with the square of exhaust velocity while thrust increase is linear. Conversely, chemical rockets provide high thrust, but are limited in total impulse by the small amount of energy that can be stored chemically in the propellants. Given the practical weight of suitable power sources, the acceleration from an ion thruster is frequently less than one-thousandth of standard gravity. However, since they operate as electric (or electrostatic) motors, they convert a greater fraction of input power into kinetic exhaust power. Chemical rockets operate as heat engines, and Carnot's theorem limits the exhaust velocity.

Gridded electrostatic ion thrusters

Gridded electrostatic ion thrusters development started in the 1960s and, since then, they have been used for commercial satellite propulsion  and scientific missions. Their main feature is that the propellant ionization process is physically separated from the ion acceleration process.

The ionization process takes place in the discharge chamber, where by bombarding the propellant with energetic electrons, as the energy transferred ejects valence electrons from the propellant gas's atoms. These electrons can be provided by a hot cathode filament and accelerated through the potential difference towards an anode. Alternatively, the electrons can be accelerated by an oscillating induced electric field created by an alternating electromagnet, which results in a self-sustaining discharge without a cathode (radio frequency ion thruster).

The positively charged ions are extracted by a system consisting of 2 or 3 multi-aperture grids. After entering the grid system near the plasma sheath, the ions are accelerated by the potential difference between the first grid and second grid (called the screen grid and the accelerator grid, respectively) to the final ion energy of (typically) 1–2 keV, which generates thrust.

Ion thrusters emit a beam of positively charged ions. To keep the spacecraft from accumulating a charge, another cathode is placed near the engine to emit electrons into the ion beam, leaving the propellant electrically neutral. This prevents the beam of ions from being attracted (and returning) to the spacecraft, which would cancel the thrust.

Gridded electrostatic ion thruster research (past/present):
 * NASA Solar Technology Application Readiness (NSTAR), 2.3 kW, used on two successful missions
 * NASA's Evolutionary Xenon Thruster (NEXT), 6.9 kW, flight qualification hardware built; used on DART mission
 * Nuclear Electric Xenon Ion System (NEXIS)
 * High Power Electric Propulsion (HiPEP), 25 kW, test example built and run briefly on the ground
 * EADS Radio-frequency Ion Thruster (RIT)
 * Dual-Stage 4-Grid (DS4G)

Hall-effect thrusters

Hall-effect thrusters accelerate ions by means of an electric potential between a cylindrical anode and a negatively charged plasma that forms the cathode. The bulk of the propellant (typically xenon) is introduced near the anode, where it ionizes and flows toward the cathode; ions accelerate towards and through it, picking up electrons as they leave to neutralize the beam and leave the thruster at high velocity.

The anode is at one end of a cylindrical tube. In the center is a spike that is wound to produce a radial magnetic field between it and the surrounding tube. The ions are largely unaffected by the magnetic field, since they are too massive. However, the electrons produced near the end of the spike to create the cathode are trapped by the magnetic field and held in place by their attraction to the anode. Some of the electrons spiral down towards the anode, circulating around the spike in a Hall current. When they reach the anode they impact the uncharged propellant and cause it to be ionized, before finally reaching the anode and completing the circuit.

Field-emission electric propulsion
Field-emission electric propulsion (FEEP) thrusters may use caesium or indium propellants. The design comprises a small propellant reservoir that stores the liquid metal, a narrow tube or a system of parallel plates that the liquid flows through and an accelerator (a ring or an elongated aperture in a metallic plate) about a millimeter past the tube end. Caesium and indium are used due to their high atomic weights, low ionization potentials and low melting points. Once the liquid metal reaches the end of the tube, an electric field applied between the emitter and the accelerator causes the liquid surface to deform into a series of protruding cusps, or Taylor cones. At a sufficiently high applied voltage, positive ions are extracted from the tips of the cones. The electric field created by the emitter and the accelerator then accelerates the ions. An external source of electrons neutralizes the positively charged ion stream to prevent charging of the spacecraft.

Pulsed inductive thrusters
Pulsed inductive thrusters (PITs) use pulses instead of continuous thrust and have the ability to run on power levels on the order of megawatts (MW). PITs consist of a large coil encircling a cone shaped tube that emits the propellant gas. Ammonia is the gas most commonly used. For each pulse, a large charge builds up in a group of capacitors behind the coil and is then released. This creates a current that moves circularly in the direction of jθ. The current then creates a magnetic field in the outward radial direction (Br), which then creates a current in the gas that has just been released in the opposite direction of the original current. This opposite current ionizes the ammonia. The positively charged ions are accelerated away from the engine due to the electric field jθ crossing the magnetic field Br, due to the Lorentz force.

Magnetoplasmadynamic thruster
Magnetoplasmadynamic (MPD) thrusters and lithium Lorentz force accelerator (LiLFA) thrusters use roughly the same idea. The LiLFA thruster builds on the MPD thruster. Hydrogen, argon, ammonia and nitrogen can be used as propellant. In a certain configuration, the ambient gas in low Earth orbit (LEO) can be used as a propellant. The gas enters the main chamber where it is ionized into plasma by the electric field between the anode and the cathode. This plasma then conducts electricity between the anode and the cathode, closing the circuit. This new current creates a magnetic field around the cathode, which crosses with the electric field, thereby accelerating the plasma due to the Lorentz force.

The LiLFA thruster uses the same general idea as the MPD thruster, though with two main differences. First, the LiLFA uses lithium vapor, which can be stored as a solid. The other difference is that the single cathode is replaced by multiple, smaller cathode rods packed into a hollow cathode tube. MPD cathodes are easily corroded due to constant contact with the plasma. In the LiLFA thruster, the lithium vapor is injected into the hollow cathode and is not ionized to its plasma form/corrode the cathode rods until it exits the tube. The plasma is then accelerated using the same Lorentz force.

In 2013, Russian company the Chemical Automatics Design Bureau successfully conducted a bench test of their MPD engine for long-distance space travel.

Electrodeless plasma thrusters
Electrodeless plasma thrusters have two unique features: the removal of the anode and cathode electrodes and the ability to throttle the engine. The removal of the electrodes eliminates erosion, which limits lifetime on other ion engines. Neutral gas is first ionized by electromagnetic waves and then transferred to another chamber where it is accelerated by an oscillating electric and magnetic field, also known as the ponderomotive force. This separation of the ionization and acceleration stages allows throttling of propellant flow, which then changes the thrust magnitude and specific impulse values.

Helicon double layer thrusters
A helicon double layer thruster is a type of plasma thruster that ejects high velocity ionized gas to provide thrust. In this design, gas is injected into a tubular chamber (the source tube) with one open end. Radio frequency AC power (at 13.56 MHz in the prototype design) is coupled into a specially shaped antenna wrapped around the chamber. The electromagnetic wave emitted by the antenna causes the gas to break down and form a plasma. The antenna then excites a helicon wave in the plasma, which further heats it. The device has a roughly constant magnetic field in the source tube (supplied by solenoids in the prototype), but the magnetic field diverges and rapidly decreases in magnitude away from the source region and might be thought of as a kind of magnetic nozzle. In operation, a sharp boundary separates the high density plasma inside the source region and the low density plasma in the exhaust, which is associated with a sharp change in electrical potential. Plasma properties change rapidly across this boundary, which is known as a current-free electric double layer. The electrical potential is much higher inside the source region than in the exhaust and this serves both to confine most of the electrons and to accelerate the ions away from the source region. Enough electrons escape the source region to ensure that the plasma in the exhaust is neutral overall.

Variable Specific Impulse Magnetoplasma Rocket (VASIMR)
The proposed Variable Specific Impulse Magnetoplasma Rocket (VASIMR) functions by using radio waves to ionize a propellant into a plasma, and then using a magnetic field to accelerate the plasma out of the back of the rocket engine to generate thrust. The VASIMR is currently being developed by Ad Astra Rocket Company, headquartered in Houston, Texas, with help from Canada-based Nautel, producing the 200 kW RF generators for ionizing propellant. Some of the components and "plasma shoots" experiments are tested in a laboratory settled in Liberia, Costa Rica. This project is led by former NASA astronaut Franklin Chang-Díaz (CRC-USA). A 200 kW VASIMR test engine was in discussion to be fitted in the exterior of the International Space Station, as part of the plan to test the VASIMR in space; however, plans for this test onboard ISS were canceled in 2015 by NASA, with a free flying VASIMR test being discussed by Ad Astra instead. An envisioned 200 MW engine could reduce the duration of flight from Earth to Jupiter or Saturn from six years to fourteen months, and Mars from 7 months to 39 days.

Microwave electrothermal thrusters
Under a research grant from the NASA Lewis Research Center during the 1980s and 1990s, Martin C. Hawley and Jes Asmussen led a team of engineers in developing a microwave electrothermal thruster (MET).

In the discharge chamber, microwave (MW) energy flows into the center containing a high level of ions (I), causing neutral species in the gaseous propellant to ionize. Excited species flow out (FES) through the low ion region (II) to a neutral region (III) where the ions complete their recombination, replaced with the flow of neutral species (FNS) towards the center. Meanwhile, energy is lost to the chamber walls through heat conduction and convection (HCC), along with radiation (Rad). The remaining energy absorbed into the gaseous propellant is converted into thrust.

Radioisotope thruster
A theoretical propulsion system has been proposed, based on alpha particles ( or indicating a helium ion with a +2 charge) emitted from a radioisotope uni-directionally through a hole in its chamber. A neutralising electron gun would produce a tiny amount of thrust with high specific impulse in the order of millions of seconds due to the high relativistic speed of alpha particles.

A variant of this uses a graphite-based grid with a static DC high voltage to increase thrust as graphite has high transparency to alpha particles if it is also irradiated with short wave UV light at the correct wavelength from a solid-state emitter. It also permits lower energy and longer half-life sources which would be advantageous for a space application. Helium backfill has also been suggested as a way to increase electron mean free path.

Lifetime
Ion thrusters' low thrust requires continuous operation for a long time to achieve the necessary change in velocity (delta-v) for a particular mission. Ion thrusters are designed to provide continuous operation for intervals of weeks to years.

The lifetime of electrostatic ion thrusters is limited by several processes.

Gridded thruster life
In electrostatic gridded designs, charge-exchange ions produced by the beam ions with the neutral gas flow can be accelerated towards the negatively biased accelerator grid and cause grid erosion. End-of-life is reached when either the grid structure fails or the holes in the grid become large enough that ion extraction is substantially affected – e.g., by the occurrence of electron backstreaming. Grid erosion cannot be avoided and is the major lifetime-limiting factor. Thorough grid design and material selection enable lifetimes of 20,000 hours or more.

A test of the NASA Solar Technology Application Readiness (NSTAR) electrostatic ion thruster resulted in 30,472 hours (roughly 3.5 years) of continuous thrust at maximum power. Post-test examination indicated the engine was not approaching failure. NSTAR operated for years on Dawn.

The NASA Evolutionary Xenon Thruster (NEXT) project operated continuously for more than 48,000 hours. The test was conducted in a high-vacuum test chamber. Over the course of the test, which lasted more than five and a half years, the engine consumed approximately 870 kilograms of xenon propellant. The total impulse generated would require over 10,000 kilograms of conventional rocket propellant for a similar application.

Hall-effect thruster life
Hall-effect thrusters suffer from strong erosion of the ceramic discharge chamber by impact of energetic ions: a test reported in 2010 showed erosion of around 1 mm per hundred hours of operation, though this is inconsistent with observed on-orbit lifetimes of a few thousand hours.

The Advanced Electric Propulsion System (AEPS) is expected to accumulate about 5,000 hours and the design aims to achieve a flight model that offers a half-life of at least 23,000 hours and a full life of about 50,000 hours.

Propellants
Ionization energy represents a large percentage of the energy needed to run ion drives. The ideal propellant is thus easy to ionize and has a high mass/ionization energy ratio. In addition, the propellant should not erode the thruster to any great degree, so as to permit long life, and should not contaminate the vehicle.

Many current designs use xenon gas, as it is easy to ionize, has a reasonably high atomic number, is inert and causes low erosion. However, xenon is globally in short supply and expensive (approximately $3,000 per kg in 2021).

Some older ion thruster designs used mercury propellant. However, mercury is toxic, tended to contaminate spacecraft, and was difficult to feed accurately. A modern commercial prototype may be using mercury successfully however, mercury was formally banned as a propellant in 2022 by the Minamata Convention on Mercury.

From 2018–2023, krypton was used to fuel the Hall-effect thrusters aboard Starlink internet satellites, in part due to its lower cost than conventional xenon propellant. Starlink V2-mini satellites have since switched to argon Hall-effect thrusters, providing higher specific impulse.

Other propellants, such as bismuth and iodine, show promise both for gridless designs such as Hall-effect thrusters,  and gridded ion thrusters.

Iodine was used as a propellant for the first time in space, in the NPT30-I2 gridded ion thruster by ThrustMe, on board the Beihangkongshi-1 mission launched in November 2020,  with an extensive report published a year later in the journal Nature. The CubeSat Ambipolar Thruster (CAT) used on the Mars Array of Ionospheric Research Satellites Using the CubeSat Ambipolar Thruster (MARS-CAT) mission also proposes to use solid iodine as the propellant to minimize storage volume.

VASIMR design (and other plasma-based engines) are theoretically able to use practically any material for propellant. However, in current tests the most practical propellant is argon, which is relatively abundant and inexpensive.

Energy efficiency

Ion thruster efficiency is the kinetic energy of the exhaust jet emitted per second divided by the electrical power into the device.

Overall system energy efficiency is determined by the propulsive efficiency, which depends on vehicle speed and exhaust speed. Some thrusters can vary exhaust speed in operation, but all can be designed with different exhaust speeds. At the lower end of specific impulse, Isp, the overall efficiency drops because ionization takes up a larger percentage energy and at the high end propulsive efficiency is reduced.

Optimal efficiencies and exhaust velocities for any given mission can be calculated to give minimum overall cost.

Missions
Ion thrusters have many in-space propulsion applications. The best applications make use of the long mission interval when significant thrust is not needed. Examples of this include orbit transfers, attitude adjustments, drag compensation for low Earth orbits, fine adjustments for scientific missions and cargo transport between propellant depots, e.g., for chemical fuels. Ion thrusters can also be used for interplanetary and deep-space missions where acceleration rates are not crucial. Ion thrusters are seen as the best solution for these missions, as they require high change in velocity but do not require rapid acceleration. Continuous thrust over long durations can reach high velocities while consuming far less propellant than traditional chemical rockets.

SERT
Ion propulsion systems were first demonstrated in space by the NASA Lewis (now Glenn Research Center) missions Space Electric Rocket Test (SERT)-1 and SERT-2A. A SERT-1 suborbital flight was launched on 20 July 1964, and successfully proved that the technology operated as predicted in space. These were electrostatic ion thrusters using mercury and caesium as the reaction mass. SERT-2A, launched on 4 February 1970, verified the operation of two mercury ion engines for thousands of running hours.

Operational missions
Ion thrusters are routinely used for station-keeping on commercial and military communication satellites in geosynchronous orbit. The Soviet Union pioneered this field, using stationary plasma thrusters (SPTs) on satellites starting in the early 1970s.

Two geostationary satellites (ESA's Artemis in 2001–2003 and the United States military's AEHF-1 in 2010–2012 ) used the ion thruster to change orbit after the chemical-propellant engine failed. Boeing began using ion thrusters for station-keeping in 1997 and planned in 2013–2014 to offer a variant on their 702 platform, with no chemical engine and ion thrusters for orbit raising; this permits a significantly lower launch mass for a given satellite capability. AEHF-2 used a chemical engine to raise perigee to 16330 km and proceeded to geosynchronous orbit using electric propulsion.

Tiangong space station
China's Tiangong space station is fitted with ion thrusters. Its Tianhe core module is propelled by both chemical thrusters and four Hall-effect thrusters, which are used to adjust and maintain the station's orbit. The development of the Hall-effect thrusters is considered a sensitive topic in China, with scientists "working to improve the technology without attracting attention". Hall-effect thrusters are created with crewed mission safety in mind with effort to prevent erosion and damage caused by the accelerated ion particles. A magnetic field and specially designed ceramic shield was created to repel damaging particles and maintain integrity of the thrusters. According to the Chinese Academy of Sciences, the ion drive used on Tiangong has burned continuously for 8,240 hours without a glitch, indicating their suitability for the Chinese space station's designated 15-year lifespan. This is the world's first Hall thruster on a human-rated mission.

Starlink
SpaceX's Starlink satellite constellation uses Hall-effect thrusters powered by krypton or argon to raise orbit, perform maneuvers, and de-orbit at the end of their use.

GOCE
ESA's Gravity Field and Steady-State Ocean Circulation Explorer (GOCE) was launched on 16 March 2009. It used ion propulsion throughout its twenty-month mission to combat the air-drag it experienced in its low orbit (altitude of 255 kilometres) before intentionally deorbiting on 11 November 2013.

Deep Space 1
NASA developed the NSTAR ion engine for use in interplanetary science missions beginning in the late 1990s. It was space-tested in the space probe Deep Space 1, launched in 1998. This was the first use of electric propulsion as the interplanetary propulsion system on a science mission. Based on the NASA design criteria, Hughes Research Labs developed the Xenon Ion Propulsion System (XIPS) for performing station keeping on geosynchronous satellites. Hughes (EDD) manufactured the NSTAR thruster used on the spacecraft.

Hayabusa and Hayabusa2
The Japanese Aerospace Exploration Agency's Hayabusa space probe was launched in 2003 and rendezvoused with the asteroid 25143 Itokawa. It was powered by four xenon ion engines, which used microwave electron cyclotron resonance to ionize the propellant and an erosion-resistant carbon/carbon-composite material for its acceleration grid. Although the ion engines on Hayabusa experienced technical difficulties, in-flight reconfiguration allowed one of the four engines to be repaired and allowed the mission to successfully return to Earth.

Hayabusa2, launched in 2014, was based on Hayabusa. It also used ion thrusters.

Smart 1
The European Space Agency's satellite SMART-1 launched in 2003 using a Snecma PPS-1350-G Hall thruster to get from GTO to lunar orbit. This satellite completed its mission on 3 September 2006, in a controlled collision on the Moon's surface, after a trajectory deviation so scientists could see the 3-meter crater the impact created on the visible side of the Moon.

Dawn
Dawn launched on 27 September 2007, to explore the asteroid Vesta and the dwarf planet Ceres. It used three Deep Space 1 heritage xenon ion thrusters (firing one at a time). Dawn ion drive is capable of accelerating from 0 to 97 km/h in 4 days of continuous firing. The mission ended on 1 November 2018, when the spacecraft ran out of hydrazine chemical propellant for its attitude thrusters.

LISA Pathfinder
LISA Pathfinder is an ESA spacecraft launched in 2015 to orbit the Sun-Earth L1 point. It does not use ion thrusters as its primary propulsion system, but uses both colloid thrusters and FEEP for precise attitude control – the low thrusts of these propulsion devices make it possible to move the spacecraft incremental distances accurately. It is a test for the LISA mission. The mission ended on 30 December 2017.

BepiColombo
ESA's BepiColombo mission was launched to Mercury on 20 October 2018. It uses ion thrusters in combination with swing-bys to get to Mercury, where a chemical rocket will complete orbit insertion.

Double Asteroid Redirection Test
NASA's Double Asteroid Redirection Test (DART) was launched in 2021 and operated its NEXT-C xenon ion thruster for about 1,000 hours to reach the target asteroid on 28 September 2022.

Psyche
NASA's Psyche spacecraft was launched in 2023 and is operating its SPT-140 xenon ion thruster in order to reach asteroid 16 Psyche in August 2029.

International Space Station
, a future launch of an Ad Astra VF-200 200 kW VASIMR electromagnetic thruster was under consideration for testing on the International Space Station (ISS). However, in 2015, NASA ended plans for flying the VF-200 to the ISS. A NASA spokesperson stated that the ISS "was not an ideal demonstration platform for the desired performance level of the engines". Ad Astra stated that tests of a VASIMR thruster on the ISS would remain an option after a future in-space demonstration.

The VF-200 would have been a flight version of the VX-200. Since the available power from the ISS is less than 200 kW, the ISS VASIMR would have included a trickle-charged battery system allowing for 15 minutes pulses of thrust. The ISS orbits at a relatively low altitude and experiences fairly high levels of atmospheric drag, requiring periodic altitude boosts – a high-efficiency engine (high specific impulse) for station-keeping would be valuable; theoretically VASIMR reboosting could cut fuel cost from the current US$210 million annually to one-twentieth. VASIMR could in theory use as little as 300 kg of argon gas for ISS station-keeping instead of 7500 kg of chemical fuel – the high exhaust velocity (high specific impulse) would achieve the same acceleration with a smaller amount of propellant, compared to chemical propulsion with its lower exhaust velocity needing more fuel. Hydrogen is generated by the ISS as a by-product and is vented into space.

NASA previously worked on a 50 kW Hall-effect thruster for the ISS, but work was stopped in 2005.

Lunar Gateway
The Power and Propulsion Element (PPE) is a module on the Lunar Gateway that provides power generation and propulsion capabilities. It is targeting launch on a commercial vehicle in January 2024. It would probably use the 50 kW Advanced Electric Propulsion System (AEPS) under development at NASA Glenn Research Center and Aerojet Rocketdyne.

MARS-CAT
The MARS-CAT (Mars Array of ionospheric Research Satellites using the CubeSat Ambipolar Thruster) mission is a two 6U CubeSat concept mission to study Mars' ionosphere. The mission would investigate its plasma and magnetic structure, including transient plasma structures, magnetic field structure, magnetic activity and correlation with solar wind drivers. The CAT thruster is now called the RF thruster and manufactured by Phase Four.

Interstellar missions
Geoffrey A. Landis proposed using an ion thruster powered by a space-based laser, in conjunction with a lightsail, to propel an interstellar probe.

Popular culture

 * The idea of an ion engine first appeared in Donald W. Horner's By Aeroplane to the Sun: Being the Adventures of a Daring Aviator and his Friends (1910).
 * Ion propulsion is the main thrust source of the spaceship Kosmokrator in the East German/Polish science fiction film Der Schweigende Stern (1960). Minute 28:10.
 * In the 1968 Star Trek episode "Spock's Brain", Scotty is repeatedly impressed by a civilization's use of ion power.
 * The popular Imperial TIE Fighter spacecraft from the Star Wars franchise are propelled by twin ion engines, hence the name.
 * Ion propulsion is used by the Hermes spacecraft in the Andy Weir novel The Martian to transfer crew between Earth and Mars.

Articles

 * "NASA Trumps Star Trek: Ion Drive Live!" The Daily Galaxy 13 April 2009.
 * "The Ultimate Space Gadget: NASA's Ion Drive Live!" The Daily Galaxy, 7 July 2009.
 * An early experimental ion engine is on display at the Aerospace Discovery at the Florida Air Museum.

title: Ion thruster


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Magnetoplasmadynamic thruster

A magnetoplasmadynamic (MPD) thruster (MPDT) is a form of electrically powered spacecraft propulsion which uses the Lorentz force (the force on a charged particle by an electromagnetic field) to generate thrust. It is sometimes referred to as Lorentz Force Accelerator (LFA) or (mostly in Japan) MPD arcjet.

Generally, a gaseous material is ionized and fed into an acceleration chamber, where the magnetic and electric fields are created using a power source. The particles are then propelled by the Lorentz force resulting from the interaction between the current flowing through the plasma and the magnetic field (which is either externally applied or induced by the current) out through the exhaust chamber. Unlike chemical propulsion, there is no combustion of fuel. As with other electric propulsion variations, both specific impulse and thrust increase with power input, while thrust per watt drops.

There are two main types of MPD thrusters, applied-field and self-field. Applied-field thrusters have magnetic rings surrounding the exhaust chamber to produce the magnetic field, while self-field thrusters have a cathode extending through the middle of the chamber. Applied fields are necessary at lower power levels, where self-field configurations are too weak. Various propellants such as xenon, neon, argon, hydrogen, hydrazine, and lithium have been used, with lithium generally being the best performer.

According to Edgar Choueiri magnetoplasmadynamic thrusters have input power 100–500 kilowatts, exhaust velocity 15–60 kilometers per second, thrust 2.5–25 newtons and efficiency 40–60 percent. However, additional research has shown that exhaust velocities can exceed 100 kilometers per second.

One potential application of magnetoplasmadynamic thrusters is the main propulsion engine for heavy cargo and piloted space vehicles (example engine $$a^2$$ for human mission to Mars).

Advantages
In theory, MPD thrusters could produce extremely high specific impulses (Isp) with an exhaust velocity of up to and beyond $110,000 m/s$, triple the value of current xenon-based ion thrusters, and about 25 times better than liquid rockets. MPD technology also has the potential for thrust levels of up to 200 newtons (N) ($45 lbf$), by far the highest for any form of electric propulsion, and nearly as high as many interplanetary chemical rockets. This would allow use of electric propulsion on missions which require quick delta-v maneuvers (such as capturing into orbit around another planet), but with many times greater fuel efficiency.

Development
MPD thruster technology has been explored academically, but commercial interest has been low due to several remaining problems. One small problem is that power requirements on the order of hundreds of kilowatts are required for optimum performance. Current interplanetary spacecraft power systems (such as radioisotope thermoelectric generators and solar arrays) are incapable of producing that much power. NASA's Project Prometheus reactor was expected to generate power in the hundreds of kilowatts range but was discontinued in 2005.

A project to produce a space-going nuclear reactor designed to generate 600 kilowatts of electrical power began in 1963 and ran for most of the 1960s in the USSR. It was to power a communication satellite which was in the end not approved. Nuclear reactors supplying kilowatts of electrical power (of the order of ten times more than current RTG power supplies) have been orbited by the USSR: RORSAT; and TOPAZ.

Plans to develop a megawatt-scale nuclear reactor for the use aboard a crewed spaceship were announced in 2009 by Russian nuclear Kurchatov Institute, national space agency Roskosmos, and confirmed by Russian President Dmitry Medvedev in his November 2009 address to the Federal Assembly.

Another plan, proposed by Bradley C. Edwards, is to beam power from the ground. This plan utilizes 5 200 kW free electron lasers at 0.84 micrometres with adaptive optics on the ground to beam power to the MPD-powered spacecraft, where it is converted to electricity by GaAs photovoltaic panels. The tuning of the laser wavelength of 0.840 micrometres ($1.48 eV$ per photon) and the photovoltaic panel bandgap of $1.43 eV$ to each other produces an estimated conversion efficiency of 59% and a predicted power density of up to $540 kW/m^{2}$. This would be sufficient to power a MPD upper stage, perhaps to lift satellites from LEO to GEO.

Another problem with MPD technology has been the degradation of cathodes due to evaporation driven by high current densities (in excess of $100 A/cm^{2}$). The use of lithium and barium propellant mixtures and multi-channel hollow cathodes has been shown in the laboratory to be a promising solution for the cathode erosion problem.

Research
Research on MPD thrusters has been carried out in the US, the former Soviet Union, Japan, Germany, and Italy. Experimental prototypes were first flown on Soviet spacecraft and, most recently, in 1996, on the Japanese Space Flyer Unit, which demonstrated the successful operation of a quasi-steady pulsed MPD thruster in space. Research at Moscow Aviation Institute, RKK Energiya, National Aerospace University, Kharkiv Aviation Institute, Institute of Space Systems of the University of Stuttgart, ISAS, Centrospazio, Alta S.p.A., Osaka University, University of Southern California, Princeton University's Electric Propulsion and Plasma Dynamics Lab (EPPDyL) (where MPD thruster research has continued uninterrupted since 1967), and NASA centers (Jet Propulsion Laboratory and Glenn Research Center), has resolved many problems related to the performance, stability and lifetime of MPD thrusters.

An MPD thruster was tested on board the Japanese Space Flyer Unit as part of EPEX (Electric Propulsion Experiment) that was launched March 18, 1995 and retrieved by space shuttle mission STS-72 January 20, 1996. To date, it is the only operational MPD thruster to have flown in space as a propulsion system. Experimental prototypes were first flown on Soviet spacecraft.

The applied-field MPD thruster in development at the Institute of Space Systems of the University of Stuttgart reached a thruster efficiency of 61.99% in 2019, corresponding to a specific impulse of Isp = 4665 s and 2.75 N of thrust.

license: Creative Commons - Attribution Share-Alike - https://creativecommons.org/licenses/by-sa/4.0/

title: Magnetoplasmadynamic thruster


text:
Pulsed plasma thruster

A pulsed plasma thruster (PPT), also known as a Pulsed Plasma Rocket (PPR), or as a plasma jet engine (PJE), is a form of electric spacecraft propulsion. PPTs are generally considered the simplest form of electric spacecraft propulsion and were the first form of electric propulsion to be flown in space, having flown on two Soviet probes (Zond 2 and Zond 3) starting in 1964. PPTs are generally flown on spacecraft with a surplus of electricity from abundantly available solar energy.

Operation
Most PPTs use a solid material (normally PTFE, more commonly known as Teflon) for propellant, although very few use liquid or gaseous propellants. The first stage in PPT operation involves an arc of electricity passing through the fuel, causing ablation and sublimation of the fuel. The heat generated by this arc causes the resultant gas to turn into plasma, thereby creating a charged gas cloud. Due to the force of the ablation, the plasma is propelled at low speed between two charged plates (an anode and cathode). Since the plasma is charged, the fuel effectively completes the circuit between the two plates, allowing a current to flow through the plasma. This flow of electrons generates a strong electromagnetic field which then exerts a Lorentz force on the plasma, accelerating the plasma out of the PPT exhaust at high velocity. Its mode of operation is similar to a railgun. The pulsing occurs due to the time needed to recharge the plates following each burst of fuel, and the time between each arc. The frequency of pulsing is normally very high and so it generates an almost continuous and smooth thrust. While the thrust is very low, a PPT can operate continuously for extended periods of time, yielding a large final speed.

The energy used in each pulse is stored in a capacitor. By varying the time between each capacitor discharge, the thrust and power draw of the PPT can be varied allowing versatile use of the system.

Comparison to chemical propulsion
The equation for the change in velocity of a spacecraft is given by the rocket equation as follows:
 * $$\Delta v = v_\text{e} \ln \frac{m_0}{m_1}$$

where:
 * $$\Delta v\ $$ is delta-v - the maximum change of speed of the vehicle (with no external forces acting),
 * $$v_\text{e}$$ is the effective exhaust velocity ($$v_\text{e} = I_\text{sp} \cdot g_0$$ where $$I_\text{sp}$$ is the specific impulse expressed as a time period and $$g_0$$ is standard gravity),
 * $$\ln$$ refers to the natural logarithm function,
 * $$m_0$$ is the initial total mass, including propellant,
 * $$m_1$$ is the final total mass.

PPTs have much higher exhaust velocities than chemical propulsion engines, but have a much smaller fuel flow rate. From the Tsiolkovsky equation stated above, this results in a proportionally higher final velocity of the propelled craft. The exhaust velocity of a PPT is of the order of tens of km/s while conventional chemical propulsion generates thermal velocities in the range of 2–4.5 km/s. Due to this lower thermal velocity, chemical propulsion units become exponentially less effective at higher vehicle velocities, necessitating the use of electric spacecraft propulsion such as PPTs. It is therefore advantageous to use an electric propulsion system such as a PPT to generate high interplanetary speeds in the range 20–70 km/s.

NASA's research PPT (flown in 2000) achieved an exhaust velocity of 13,700 m/s, generated a thrust of 860 μN, and consumed 70W of electrical power.

Advantages and disadvantages
PPTs are very robust due to their inherently simple design (relative to other electric spacecraft propulsion techniques). As an electric propulsion system, PPTs benefit from reduced fuel consumption compared to traditional chemical rockets, reducing launch mass and therefore launch costs, as well as high specific impulse improving performance.

However, due to energy losses caused by late time ablation and rapid conductive heat transfer from the propellant to the rest of the spacecraft, propulsive efficiency (kinetic energy of exhaust / total energy used) is very low compared to other forms of electric propulsion, at around just 10%.

Uses
PPTs are well-suited to uses on relatively small spacecraft with a mass of less than 100 kg (particularly CubeSats) for roles such as attitude control, station keeping, de-orbiting manoeuvres and deep space exploration. Using PPTs could double the life-span of these small satellite missions without significantly increasing complexity or cost due to the inherent simplicity and relatively low cost nature of PPTs.

The first use of PPTs was on the Soviet Zond 2 space probe which carried six PPTs that served as actuators of the attitude control system. The PPT propulsion system was tested for 70 minutes on the 14 December 1964 when the spacecraft was 4.2 million kilometers from Earth.

A PPT was flown by NASA in November, 2000, as a flight experiment on the Earth Observing-1 spacecraft. The thrusters successfully demonstrated the ability to perform roll control on the spacecraft and demonstrated that the electromagnetic interference from the pulsed plasma did not affect other spacecraft systems. Pulsed plasma thrusters are also an avenue of research used by universities for starting experiments with electric propulsion due to the relative simplicity and lower costs involved with PPTs as opposed to other forms of electric propulsion such as Hall-effect ion thrusters.

Ongoing NASA research
On July 11, 2024 Howe Industries announced that it had contracted with NASA via a $725,000.00 grant to continue research on PPT/PPR propulsion technology. Howe Industries claimed that should PPT/PPR propulsion technology succeed in becoming a fully functional means of propelling space ships to Mars, then PPT/PPR technology should be capable of shortening travel time to Mars, down from the current requirement of approximately 1 year, to a much shorter travel time of only 2 months. Howe Industries further stated that at the current rate of their PPT/PPR research and development program, the technology may not be fully ready to propel a manned space ship to Mars for approximately another 20 years (as of 2024).

title: Pulsed plasma thruster


text:
Nuclear thermal rocket

A nuclear thermal rocket (NTR) is a type of thermal rocket where the heat from a nuclear reaction replaces the chemical energy of the propellants in a chemical rocket. In an NTR, a working fluid, usually liquid hydrogen, is heated to a high temperature in a nuclear reactor and then expands through a rocket nozzle to create thrust. The external nuclear heat source theoretically allows a higher effective exhaust velocity and is expected to double or triple payload capacity compared to chemical propellants that store energy internally.

NTRs have been proposed as a spacecraft propulsion technology, with the earliest ground tests occurring in 1955. The United States maintained an NTR development program through 1973 when it was shut down for various reasons, including to focus on Space Shuttle development. Although more than ten reactors of varying power output have been built and tested,, no nuclear thermal rocket has flown.

Whereas all early applications for nuclear thermal rocket propulsion used fission processes, research in the 2010s has moved to fusion approaches. The Direct Fusion Drive project at the Princeton Plasma Physics Laboratory is one such example, although "energy-positive fusion has remained elusive". In 2019, the U.S. Congress approved US$125 million in development funding for nuclear thermal propulsion rockets.

In May 2022 DARPA issued an RFP for the next phase of their Demonstration Rocket for Agile Cislunar Operations (DRACO) nuclear thermal engine program. This follows on their selection, in 2021, of an early engine design by General Atomics and two spacecraft concepts from Blue Origin and Lockheed Martin. The next phases of the program will focus on the design, development, fabrication, and assembly of a nuclear thermal rocket engine. In July 2023, Lockheed Martin was awarded the contract to build the spacecraft and BWX Technologies (BWXT) will develop the nuclear reactor. A launch is expected in 2027.

Principle of operation
Nuclear-powered thermal rockets are more effective than chemical thermal rockets, primarily because they can use low-molecular-mass propellants such as hydrogen.

As thermal rockets, nuclear thermal rockets work almost exactly like chemical rockets: a heat source releases thermal energy into a gaseous propellant inside the body of the engine, and a nozzle at one end acts as a very simple heat engine: it allows the propellant to expand away from the vehicle, carrying momentum with it and converting thermal energy to coherent kinetic energy. The specific impulse (Isp) of the engine is set by the speed of the exhaust stream. That, in turn, varies as the square root of the kinetic energy loaded on each unit mass of propellant. The kinetic energy per molecule of propellant is determined by the temperature of the heat source (whether it be a nuclear reactor or a chemical reaction). At any particular temperature, lightweight propellant molecules carry just as much kinetic energy as heavier propellant molecules and therefore have more kinetic energy per unit mass. This makes low-molecular-mass propellants more effective than high-molecular-mass propellants.

Because chemical rockets and nuclear rockets are made from refractory solid materials, they are both limited to operate below 3000 C, by the strength characteristics of high-temperature metals. Chemical rockets use the most readily available propellant, which is waste products from the chemical reactions producing their heat energy. Most liquid-fueled chemical rockets use either hydrogen or hydrocarbon combustion, and the propellant is therefore mainly water (molecular mass 18) and carbon dioxide (molecular mass 44). Nuclear thermal rockets using gaseous hydrogen propellant (molecular mass 2) therefore have a theoretical maximum specific impulse that is 3 to 4.5 times greater than those of chemical rockets.

Early history
In 1944, Stanisław Ulam and Frederic de Hoffmann contemplated the idea of controlling the power of nuclear explosions to launch space vehicles. After World War II, the U.S. military started the development of intercontinental ballistic missiles (ICBM) based on the German V-2 rocket designs. Some large rockets were designed to carry nuclear warheads with nuclear-powered propulsion engines. As early as 1946, secret reports were prepared for the U.S. Air Force, as part of the NEPA project, by North American Aviation and Douglas Aircraft Company's Project Rand. These groundbreaking reports identified a reactor engine in which a working fluid of low molecular weight is heated using a nuclear reactor as the most promising form of nuclear propulsion but identified many technical issues that needed to be resolved.

In January 1947, not aware of this classified research, engineers of the Applied Physics Laboratory published their research on nuclear power propulsion and their report was eventually classified. In May 1947, American-educated Chinese scientist Qian Xuesen presented his research on "thermal jets" powered by a porous graphite-moderated nuclear reactor at the Nuclear Science and Engineering Seminars LIV organized by the Massachusetts Institute of Technology.

In 1948 and 1949, physicist Leslie Shepherd and rocket scientist Val Cleaver produced a series of groundbreaking scientific papers that considered how nuclear technology might be applied to interplanetary travel. The papers examined both nuclear-thermal and nuclear-electric propulsion.

Early NASA engine development
Through Project Rover, Los Alamos National Laboratory began developing nuclear thermal engines as soon as 1955 and tested the worlds first experimental nuclear rocket engine, KIWI-A, in 1959. This work at Los Alamos was then continued through the NASA's NERVA program (1961–1973). NERVA achieved many successes and improved upon the early prototypes to create powerful engines that were several times more efficient than chemical counterparts. However, the program was cancelled in 1973 due to budget constraints. To date no nuclear thermal propulsion system has ever been implemented in space.

Nuclear fuel types
A nuclear thermal rocket can be categorized by the type of reactor, ranging from a relatively simple solid reactor up to the much more difficult to construct but theoretically more efficient gas core reactor. As with all thermal rocket designs, the specific impulse produced is proportional to the square root of the temperature to which the working fluid (reaction mass) is heated. To extract maximum efficiency, the temperature must be as high as possible. For a given design, the temperature that can be attained is typically determined by the materials chosen for reactor structures, the nuclear fuel, and the fuel cladding. Erosion is also a concern, especially the loss of fuel and associated releases of radioactivity.

Solid core

Solid core nuclear reactors have been fueled by compounds of uranium that exist in solid phase under the conditions encountered and undergo nuclear fission to release energy. Flight reactors must be lightweight and capable of tolerating extremely high temperatures, as the only coolant available is the working fluid/propellant. A nuclear solid core engine is the simplest design to construct and is the concept used on all tested NTRs.

Using hydrogen as a propellant, a solid core design would typically deliver specific impulses (Isp) on the order of 850 to 1000 seconds, which is about twice that of liquid hydrogen-oxygen designs such as the Space Shuttle main engine. Other propellants have also been proposed, such as ammonia, water, or LOX, but these propellants would provide reduced exhaust velocity and performance at a marginally reduced fuel cost. Yet another mark in favor of hydrogen is that at low pressures it begins to dissociate at about 1500 K, and at high pressures around 3000 K. This lowers the mass of the exhaust species, increasing Isp.

Early publications were doubtful of space applications for nuclear engines. In 1947, a complete nuclear reactor was so heavy that solid core nuclear thermal engines would be entirely unable to achieve a thrust-to-weight ratio of 1:1, which is needed to overcome the gravity of the Earth at launch. Over the next twenty-five years, U.S. nuclear thermal rocket designs eventually reached thrust-to-weight ratios of approximately 7:1. This is still a much lower thrust-to-weight ratio than what is achievable with chemical rockets, which have thrust-to-weight ratios on the order of 70:1. Combined with the large tanks necessary for liquid hydrogen storage, this means that solid core nuclear thermal engines are best suited for use in orbit outside Earth's gravity well, not to mention avoiding the radioactive contamination that would result from atmospheric use (if an "open-cycle" design was used, as opposed to a lower-performance "closed cycle" design where no radioactive material was allowed to escape with the rocket propellant. )

One way to increase the working temperature of the reactor is to change the nuclear fuel elements. This is the basis of the particle-bed reactor, which is fueled by several (typically spherical) elements that "float" inside the hydrogen working fluid. Spinning the entire engine could prevent the fuel element from being ejected out the nozzle. This design is thought to be capable of increasing the specific impulse to about 1000 seconds (9.8 kN·s/kg) at the cost of increased complexity. Such a design could share design elements with a pebble-bed reactor, several of which are currently generating electricity. From 1987 through 1991, the Strategic Defense Initiative (SDI) Office funded Project Timberwind, a non-rotating nuclear thermal rocket based on particle bed technology. The project was canceled before testing.

Pulsed nuclear thermal rocket

In a conventional solid core design, the maximum exhaust temperature of the working mass is that of the reactor, and in practice, lower than that. That temperature represents an energy far below that of the individual neutrons released by the fission reactions. Their energy is spread out through the reactor mass, causing it to thermalize. In power plant designs, the core is then cooled, typically using water. In the case of a nuclear engine, the water is replaced by hydrogen, but the concept is otherwise similar.

Pulsed reactors attempt to transfer the energy directly from the neutrons to the working mass, allowing the exhaust to reach temperatures far beyond the melting point of the reactor core. As specific impulse varies directly with temperature, capturing the energy of the relativistic neutrons allows for a dramatic increase in performance.

To do this, pulsed reactors operate in a series of brief pulses rather than the continual chain reaction of a conventional reactor. The reactor is normally off, allowing it to cool. It is then turned on, along with the cooling system or fuel flow, operating at a very high power level. At this level the core rapidly begins to heat up, so once a set temperature is reached, the reactor is quickly turned off again. During these pulses, the power being produced is far greater than the same sized reactor could produce continually. The key to this approach is that while the total amount of fuel that can be pumped through the reactor during these brief pulses is small, the resulting efficiency of these pulses is much higher.

Generally, the designs would not be operated solely in the pulsed mode but could vary their duty cycle depending on the need. For instance, during a high-thrust phase of flight, like exiting a low earth orbit, the engine could operate continually and provide an Isp similar to that of traditional solid-core design. But during a long-duration cruise, the engine would switch to pulsed mode to make better use of its fuel.

Liquid core
Liquid core nuclear engines are fueled by compounds of fissionable elements in liquid phase. A liquid-core engine is proposed to operate at temperatures above the melting point of solid nuclear fuel and cladding, with the maximum operating temperature of the engine instead of being determined by the reactor pressure vessel and neutron reflector material. The higher operating temperatures would be expected to deliver specific impulse performance on the order of 1300 to 1500 seconds (12.8-14.8 kN·s/kg).

A liquid-core reactor would be extremely difficult to build with current technology. One major issue is that the reaction time of the nuclear fuel is much longer than the heating time of the working fluid. If the nuclear fuel and working fluid are not physically separated, this means that the fuel must be trapped inside the engine while the working fluid is allowed to easily exit through the nozzle. One possible solution is to rotate the fuel/fluid mixture at very high speeds to force the higher-density fuel to the outside, but this would expose the reactor pressure vessel to the maximum operating temperature while adding mass, complexity, and moving parts.

An alternative liquid-core design is the nuclear salt-water rocket. In this design, water is the working fluid and also serves as the neutron moderator. Nuclear fuel is not retained, which drastically simplifies the design. However, the rocket would discharge massive quantities of extremely radioactive waste and could only be safely operated well outside the Earth's atmosphere and perhaps even magnetosphere.

Gas core

The final fission classification is the gas-core engine. This is a modification to the liquid-core design which uses rapid circulation of the fluid to create a toroidal pocket of gaseous uranium fuel in the middle of the reactor, surrounded by hydrogen. In this case, the fuel does not touch the reactor wall at all, so temperatures could reach several tens of thousands of degrees, which would allow specific impulses of 3000 to 5000 seconds (30 to 50 kN·s/kg). In this basic design, the "open cycle", the losses of nuclear fuel would be difficult to control, which has led to studies of the "closed cycle" or nuclear lightbulb engine, where the gaseous nuclear fuel is contained in a super-high-temperature quartz container, over which the hydrogen flows. The closed-cycle engine has much more in common with the solid-core design, but this time is limited by the critical temperature of quartz instead of the fuel and cladding. Although less efficient than the open-cycle design, the closed-cycle design is expected to deliver a specific impulse of about 1500 to 2000 seconds (15 to 20 kN·s/kg).

Soviet Union and Russia
The Soviet RD-0410 went through a series of tests at the nuclear test site near Semipalatinsk Test Site.

In October 2018, Russia's Keldysh Research Center confirmed a successful ground test of waste heat radiators for a nuclear space engine, as well as previous tests of fuel rods and ion engines.

United States

Development of solid core NTRs started in 1955 under the Atomic Energy Commission (AEC) as Project Rover and ran to 1973. Work on a suitable reactor was conducted at Los Alamos National Laboratory and Area 25 (Nevada National Security Site) in the Nevada Test Site. Four basic designs came from this project: KIWI, Phoebus, Pewee, and the Nuclear Furnace. Twenty individual engines were tested, with a total of over 17 hours of engine run time.

When NASA was formed in 1958, it was given authority over all non-nuclear aspects of the Rover program. To enable cooperation with the AEC and keep classified information compartmentalized, the Space Nuclear Propulsion Office (SNPO) was formed at the same time. The 1961 NERVA program was intended to lead to the entry of nuclear thermal rocket engines into space exploration. Unlike the AEC work, which was intended to study the reactor design itself, NERVA's goal was to produce a real engine that could be deployed on space missions. The 334 kN thrust baseline NERVA design was based on the KIWI B4 series.

Tested engines included Kiwi, Phoebus, NRX/EST, NRX/XE, Pewee, Pewee 2, and the Nuclear Furnace. Progressively higher power densities culminated in the Pewee. Tests of the improved Pewee 2 design were canceled in 1970 in favor of the lower-cost Nuclear Furnace (NF-1), and the U.S. nuclear rocket program officially ended in the spring of 1973. During this program, the NERVA accumulated over 2 hours of run time, including 28 minutes at full power. The SNPO considered NERVA to be the last technology development reactor required to proceed to flight prototypes.

Several other solid-core engines have also been studied to some degree. The Small Nuclear Rocket Engine, or SNRE, was designed at the Los Alamos National Laboratory (LANL) for upper stage use, both on uncrewed launchers and the Space Shuttle. It featured a split-nozzle that could be rotated to the side, allowing it to take up less room in the Shuttle cargo bay. The design provided 73 kN of thrust and operated at a specific impulse of 875 seconds (8.58 kN·s/kg), and it was planned to increase this to 975 seconds, achieving a mass fraction of about 0.74, compared with 0.86 for the Space Shuttle main engine (SSME), one of the best conventional engines.

A related design that saw some work, but never made it to the prototype stage, was Dumbo. Dumbo was similar to KIWI/NERVA in concept, but used more advanced construction techniques to lower the weight of the reactor. The Dumbo reactor consisted of several large barrel-like tubes, which were in turn constructed of stacked plates of corrugated material. The corrugations were lined up so that the resulting stack had channels running from the inside to the outside. Some of these channels were filled with uranium fuel, others with a moderator, and some were left open as a gas channel. Hydrogen was pumped into the middle of the tube and would be heated by the fuel as it traveled through the channels as it worked its way to the outside. The resulting system was lighter than a conventional design for any particular amount of fuel.

Between 1987 and 1991, an advanced engine design was studied under Project Timberwind, under the Strategic Defense Initiative, which was later expanded into a larger design in the Space Thermal Nuclear Propulsion (STNP) program. Advances in high-temperature metals, computer modeling, and nuclear engineering, in general, resulted in dramatically improved performance. While the NERVA engine was projected to weigh about 6803 kg, the final STNP offered just over 1/3 the thrust from an engine of only 1650 kg by improving the Isp to between 930 and 1000 seconds.

Test firings

KIWI was the first to be fired, starting in July 1959 with KIWI 1. The reactor was not intended for flight and was named after the flightless bird, Kiwi. The core was simply a stack of uncoated uranium oxide plates onto which the hydrogen was dumped. The thermal output of 70 MW at an exhaust temperature of 2683 K was generated. Two additional tests of the basic concept, A1 and A3, added coatings to the plates to test fuel rod concepts.

The KIWI B series was fueled by tiny uranium dioxide (UO2) spheres embedded in a low-boron graphite matrix and coated with niobium carbide. Nineteen holes ran the length of the bundles, through which the liquid hydrogen flowed. On the initial firings, immense heat and vibration cracked the fuel bundles. The graphite materials used in the reactor's construction were resistant to high temperatures but eroded under the stream of superheated hydrogen, a reducing agent. The fuel species was later switched to uranium carbide, with the last engine run in 1964. The fuel bundle erosion and cracking problems were improved but never completely solved, despite promising materials work at the Argonne National Laboratory.

NERVA NRX (Nuclear Rocket Experimental), started testing in September 1964. The final engine in this series was the XE, designed with flight representative hardware and fired into a low-pressure chamber to simulate a vacuum. SNPO fired NERVA NRX/XE twenty-eight times in March 1968. The series all generated 1100 MW, and many of the tests concluded only when the test-stand ran out of hydrogen propellant. NERVA NRX/XE produced the baseline 334 kN thrust that Marshall Space Flight Center required in Mars mission plans. The last NRX firing lost 38 lb of nuclear fuel in 2 hours of testing, which was judged sufficient for space missions by SNPO.

Building on the KIWI series, the Phoebus series were much larger reactors. The first 1A test in June 1965 ran for over 10 minutes at 1090 MW and an exhaust temperature of 2370 K. The B run in February 1967 improved this to 1500 MW for 30 minutes. The final 2A test in June 1968 ran for over 12 minutes at 4000 MW, at the time the most powerful nuclear reactor ever built.

A smaller version of KIWI, the Pewee was also built. It was fired several times at 500 MW to test coatings made of zirconium carbide (instead of niobium carbide) but Pewee also increased the power density of the system. A water-cooled system is known as NF-1 (for Nuclear Furnace) used Pewee 2's fuel elements for future materials testing, showing a factor of 3 reductions in fuel corrosion still further. Pewee 2 was never tested on the stand and became the basis for current NTR designs being researched at NASA's Glenn Research Center and Marshall Space flight Center.

The NERVA/Rover project was eventually canceled in 1972 with the general wind-down of NASA in the post-Apollo era. Without a human mission to Mars, the need for a nuclear thermal rocket is unclear. Another problem would be public concerns about safety and radioactive contamination.

Kiwi-TNT destructive test
In January 1965, the U.S. Rover program intentionally modified a Kiwi reactor (KIWI-TNT) to go prompt critical, resulting in immediate destruction of the reactor pressure vessel, nozzle, and fuel assemblies. Intended to simulate a worst-case scenario of a fall from altitude into the ocean, such as might occur in a booster failure after launch, the resulting release of radiation would have caused fatalities out to 600 ft and injuries out to 2000 ft. The reactor was positioned on a railroad car in the Jackass Flats area of the Nevada Test Site.

United Kingdom
As of January 2012, the propulsion group for Project Icarus was studying an NTR propulsion system, but has seen little activity since 2019.

Israel
In 1987, Ronen & Leibson published a study on applications of 242mAm (one of the isotopes of americium) as nuclear fuel to space nuclear reactors, noting its extremely high thermal cross section and energy density. Nuclear systems powered by 242mAm require less fuel by a factor of 2 to 100 compared to conventional nuclear fuels.

Fission-fragment rocket using 242mAm was proposed by George Chapline at Lawrence Livermore National Laboratory (LLNL) in 1988, who suggested propulsion based on the direct heating of a propellant gas by fission fragments generated by a fissile material. Ronen et al. demonstrate that 242mAm can maintain sustained nuclear fission as an extremely thin metallic film, less than 1/1000 of a millimeter thick. 242mAm requires only 1% of the mass of 235U or 239Pu to reach its critical state. Ronen's group at the Ben-Gurion University of the Negev further showed that nuclear fuel based on 242mAm could speed space vehicles from Earth to Mars in as little as two weeks.

The 242mAm as a nuclear fuel is derived from the fact that it has the highest thermal fission cross section (thousands of barns), about 10x the next highest cross section across all known isotopes. The 242mAm is fissile (because it has an odd number of neutrons) and has a low critical mass, comparable to that of 239Pu.

It has a very high cross section for fission, and if in a nuclear reactor is destroyed relatively quickly. Another report claims that 242mAm can sustain a chain reaction even as a thin film, and could be used for a novel type of nuclear rocket.

Since the thermal absorption cross section of 242mAm is very high, the best way to obtain 242mAm is by the capture of fast or epithermal neutrons in Americium-241 irradiated in a fast reactor. However, fast spectrum reactors are not readily available. Detailed analysis of 242mAm breeding in existing pressurized water reactors (PWRs) was provided. Proliferation resistance of 242mAm was reported by the Karlsruhe Institute of Technology 2008 study.

Italy
In 2000, Carlo Rubbia at CERN further extended the work by Ronen and Chapline on a Fission-fragment rocket using 242mAm as a fuel. Project 242 based on Rubbia design studied a concept of 242mAm based Thin-Film Fission Fragment Heated NTR by using a direct conversion of the kinetic energy of fission fragments into increasing of enthalpy of a propellant gas. Project 242 studied the application of this propulsion system to a crewed mission to Mars. Preliminary results were very satisfactory, and it has been observed that a propulsion system with these characteristics could make the mission feasible. Another study focused on the production of 242mAm in conventional thermal nuclear reactors.

Current research in the US since 2000

Current solid-core nuclear thermal rocket designs are intended to greatly limit the dispersion and break-up of radioactive fuel elements in the event of a catastrophic failure.

As of 2013, an NTR for interplanetary travel from Earth orbit to Mars orbit is being studied at Marshall Space Flight Center with Glenn Research Center. In historical ground testing, NTRs proved to be at least twice as efficient as the most advanced chemical engines, which would allow for quicker transfer time and increased cargo capacity. The shorter flight duration, estimated at 3–4 months with NTR engines, compared to 6–9 months using chemical engines, would reduce crew exposure to potentially harmful and difficult to shield cosmic rays. NTR engines, such as the Pewee of Project Rover, were selected in the Mars Design Reference Architecture (DRA).

In 2017, NASA continued research and development on NTRs, designing for space applications with civilian approved materials, with a three-year, US$18.8 million contract.

In 2019, an appropriation bill passed by the U.S. Congress included US$125 million in funding for nuclear thermal propulsion research, including planning for a flight demonstration mission by 2024.

As of 2021, there has been much interest in nuclear thermal rockets by the United States Space Force and DARPA for orbital and cis-lunar uses. In addition to the U.S. military, NASA administrator Jim Bridenstine has also expressed interest in the project and its potential applications for a future mission to Mars. DARPA has awarded 2 contracts for their Demonstration Rocket for Agile Cislunar Operations (DRACO) program, which aims to demonstrate a nuclear thermal propulsion system in orbit: one award in September 2020 to Gryphon Technologies for US$14 million, and another award in April 2021 to General Atomics for US$22 million, both for preliminary designs for the reactor. Two conceptual spacecraft designs by Blue Origin and Lockheed Martin were selected. Proposals for a flight demonstration of nuclear thermal propulsion in FY2026 were due on August 5, 2022.

In January 2023, NASA and DARPA announced a partnership on DRACO to demonstrate an NTR engine in space, an enabling capability for NASA crewed missions to Mars. In July 2023, U.S. agencies announced that Lockheed Martin had been awarded a $499 million contract to assemble the experimental nuclear thermal reactor vehicle (X-NTRV) and its engine.

Risks
An atmospheric or orbital rocket failure could result in the dispersal of radioactive material into the environment. A collision with orbital debris, material failure due to uncontrolled fission, material imperfections or fatigue, or human design flaws could cause a containment breach of the fissile material. Such a catastrophic failure while in flight could release radioactive material over the Earth in a wide and unpredictable area. The amount of contamination would depend on the size of the nuclear thermal rocket engine, while the zone of contamination and its concentration would be dependent on prevailing weather and orbital parameters at the time of re-entry.

It is considered unlikely that a reactor's fuel elements would be spread over a wide area, as they are composed of materials such as carbon composites or carbides and are normally coated with zirconium hydride. Before criticality occurs, solid core NTR fuel is not particularly hazardous. Once the reactor has been started for the first time, extremely radioactive short-life fission products are produced, as well as less radioactive but extremely long-lived fission products. The amount of fission products is zero at fresh-fueled startup, and roughly proportional to (actually: limited by) the total amount of fission heat produced since fresh-fueled startup. Additionally, all engine structures are exposed to direct neutron bombardment, resulting in their radioactive activation.

title: Nuclear thermal rocket


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Nuclear electric rocket

A nuclear electric rocket (more properly nuclear electric propulsion) is a type of spacecraft propulsion system where thermal energy from a nuclear reactor is converted to electrical energy, which is used to drive an ion thruster or other electrical spacecraft propulsion technology. The nuclear electric rocket terminology is slightly inconsistent, as technically the "rocket" part of the propulsion system is non-nuclear and could also be driven by solar panels. This is in contrast with a nuclear thermal rocket, which directly uses reactor heat to add energy to a working fluid, which is then expelled out of a rocket nozzle.

Conceptual overview
The key elements to NEP are:
 * 1) A compact reactor core
 * 2) An electric generator
 * 3) A compact waste heat rejection system such as heat pipes
 * 4) An electric power conditioning and distribution system
 * 5) Electrically powered spacecraft propulsion

United States
A 1963 paper by Myron Levoy proposed a hybrid nuclear-electric engine design, which would have been able to work both in open-cycle mode as a nuclear thermal engine during mission phases requiring high thrust, as well as in closed-cycle mode as a nuclear-electric engine with low thrust, but high efficiency during remaining mission phases. The proposed application of this engine design was for a fast human-crewed round-trip mission to Mars.

In 2001, the Safe affordable fission engine was under development, with a tested 30 kW nuclear heat source intended to lead to the development of a 400 kW thermal reactor with Brayton cycle gas turbines to produce electric power. Waste heat rejection was intended to be accomplished using low-mass heat pipe technology. Safety was intended to be assured by a rugged design.

Project Prometheus was an early 2000s NASA study on nuclear electric spacecraft.

Kilopower is the latest NASA reactor development program, but is intended for surface use only.

Russia
The TEM project started in 2009 with the goal of powering a Mars engine.

March 2016 - First batch of nuclear fuel received

Pebble bed reactor combined with gas turbine
A pebble bed reactor using high mass-flow gaseous nitrogen coolant near normal atmospheric pressures is a possible heat source. Power generation could be accomplished with gas turbine technology, which is well developed. Nuclear fuel would be highly enriched uranium encapsulated in low-boron graphite balls probably 5–10 cm in diameter. The graphite would also moderate the neutrons of the nuclear reaction.

This style of reactor can be designed to be inherently safe. As it heats, the graphite expands, separating the fuel and reducing the reactor's criticality. This property can simplify the operating controls to a single valve throttling the turbine. When closed, the reactor heats, but produces less power. When open, the reactor cools, but becomes more critical and produces more power.

The graphite encapsulation simplifies refueling and waste handling. Graphite is mechanically strong, and resists high temperatures. This reduces the risk of an unplanned release of radioactive elements, including fission products. Since this style of reactor produces high power without heavy castings to contain high pressures, it is well suited to power spacecraft.

Novel electric propulsion concepts
A variety of electric propulsion technologies have been proposed for use with high power nuclear electrical generation systems, including VASIMR, DS4G, and pulsed inductive thruster (PIT). PIT and VASIMR are unique in their ability to trade between power usage, specific impulse (a measure of efficiency, see specific impulse) and thrust in-flight. PIT has the additional advantage of not needing conditioned power.

Electrical generation
A number of heat-to-electricity schemes have been proposed. In the near term, Rankine cycle, Brayton cycle, and Stirling cycle generators go through an intermediate mechanical phase, with attendant energy losses. More exotic technologies have also been proposed: thermoelectric (including graphene-based thermal power conversion  ), pyroelectric, thermophotovoltaic, thermionic and magnetohydrodynamic type thermoelectric materials.

Other types of nuclear power concepts in space
Radioisotope thermoelectric generators, radioisotope heater units, radioisotope piezoelectric generators, and the radioisotope rocket all use the heat from a static radioactive source (usually Plutonium-238) for a low level of electric or direct propulsion power. Other concepts include the nuclear thermal rocket, the fission fragment rocket, nuclear pulse propulsion, and the possibility of a fusion rocket, assuming that nuclear fusion technology is developed at some point in the near future.

title: Nuclear electric rocket


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Mass driver

A mass driver or electromagnetic catapult is a proposed method of non-rocket spacelaunch which would use a linear motor to accelerate and catapult payloads up to high speeds. Existing and proposed mass drivers use coils of wire energized by electricity to make electromagnets, though a rotary mass driver has also been proposed. Sequential firing of a row of electromagnets accelerates the payload along a path. After leaving the path, the payload continues to move due to momentum.

Although any device used to propel a ballistic payload is technically a mass driver, in this context a mass driver is essentially a coilgun that magnetically accelerates a package consisting of a magnetizable holder containing a payload. Once the payload has been accelerated, the two separate, and the holder is slowed and recycled for another payload.

Mass drivers can be used to propel spacecraft in three different ways: A large, ground-based mass driver could launch spacecraft away from Earth, the Moon, or another body. A small mass driver could act as a rocket engine on board a spacecraft, flinging pieces of material into space to propel itself. Another variation would have a massive facility on a moon or asteroid send projectiles to assist a distant craft.

Miniaturized mass drivers can also be used as weapons in a similar manner as classic firearms or cannon using chemical combustion. Hybrids between coilguns and railguns such as helical railguns are also possible.

Fixed mass drivers
Mass drivers need no physical contact between moving parts because they guide their projectiles by dynamic magnetic levitation, allowing extreme reusability in the case of solid-state power switching, and a functional life of – theoretically – up to millions of launches. While marginal costs tend to be accordingly low, initial development and construction costs are highly dependent on performance, especially the intended mass, acceleration, and velocity of projectiles. For instance, while Gerard O'Neill built his first mass driver in 1976–1977 with a $2000 budget, a short test model firing a projectile at 40 m/s and 33 g, his next model had an order-of-magnitude greater acceleration after a comparable increase in funding, and, a few years later, researchers at the University of Texas estimated that a mass driver firing a 10 kilogram projectile at 6000 m/s would cost $47 million.

For a given amount of energy involved, heavier objects go proportionally slower. Lightweight objects may be projected at 20 km/s or more. The limits are generally the expense of energy storage able to be discharged quickly enough and the cost of power switching, which may be by semiconductors or by gas-phase switches (which still often have a niche in extreme pulse power applications). However, energy can be stored inductively in superconducting coils. A 1 km long mass driver made of superconducting coils can accelerate a 20 kg vehicle to 10.5 km/s at a conversion efficiency of 80%, and average acceleration of 5,600 g.

Earth-based mass drivers for propelling vehicles to orbit, such as the StarTram concept, would require considerable capital investment. The Earth's relatively strong gravity and relatively thick atmosphere make the implementation of a practical solution difficult. Also, most if not all plausible launch sites would propel spacecraft through heavily-traversed air routes. Due to the massive turbulence such launches would cause, significant air traffic control measures would be needed to ensure the safety of other aircraft operating in the area.

With the proliferation of reusable rockets to launch from Earth (especially first stages) whatever potential might have once existed for any economic advantage in using mass drivers as an alternative to chemical rockets to launch from Earth is becoming increasingly doubtful. For these reasons many proposals feature installing mass drivers on the Moon where the lower gravity and lack of atmosphere greatly reduce the required velocity to reach lunar orbit; also, lunar launches from a fixed position are much less likely to generate issues with respect to matters such as traffic control.

Most serious mass-driver designs use superconducting coils to achieve reasonable energetic efficiency (often 50% to 90+%, depending on design). Equipment may include a superconducting bucket or aluminum coil as the payload. The coils of a mass driver can induce eddy currents in a payload's aluminum coil, and then act on the resulting magnetic field. There are two sections of a mass driver. The maximum acceleration part spaces the coils at constant distances, and synchronizes the coil currents to the bucket. In this section, the acceleration increases as the velocity increases, up to the maximum that the bucket can take. After that, the constant acceleration region begins. This region spaces the coils at increasing distances to give a fixed amount of velocity increase per unit of time.

Based on this mode, a major proposal for the use of mass drivers involved transporting lunar-surface material to space habitats for processing using solar energy. The Space Studies Institute showed that this application was reasonably practical.

In some designs, the payload would be held in a bucket and then released, so that the bucket can be decelerated and reused. A disposable bucket, on the other hand, would avail acceleration along the whole track. Alternatively, if a track were constructed along the entire circumference of the Moon (or any other celestial body without a significant atmosphere) then a reusable bucket's acceleration would not be limited by the length of the track – however, such a system would need to be engineered to withstand substantial centrifugal forces if it were intended to accelerate passengers and/or cargo to very high velocities.

On Earth
In contrast to cargo-only chemical space-gun concepts, a mass driver could be any length, affordable, and with relatively smooth acceleration throughout, optionally even lengthy enough to reach target velocity without excessive g forces for passengers. It can be constructed as a very long and mainly horizontally aligned launch track for spacelaunch, targeted upwards at the end, partly by bending of the track upwards and partly by Earth's curvature in the other direction.

Natural elevations, such as mountains, may facilitate the construction of the distant, upwardly targeted part. The higher up the track terminates, the less resistance from the atmosphere the launched object will encounter.

The 40 megajoules per kilogram or less kinetic energy of projectiles launched at up to 9000 m/s velocity (if including extra for drag losses) towards low Earth orbit is a few kilowatt-hours per kilogram if efficiencies are relatively high, which accordingly has been hypothesized to be under $1 of electrical energy cost per kilogram shipped to LEO, though total costs would be far more than electricity alone. By being mainly located slightly above, on or beneath the ground, a mass driver may be easier to maintain compared with many other structures of non-rocket spacelaunch. Whether or not underground, it needs to be housed in a pipe that is vacuum pumped in order to prevent internal air drag, such as with a mechanical shutter kept closed most of the time but a plasma window used during the moments of firing to prevent loss of vacuum.

A mass driver on Earth would usually be a compromise system. A mass driver would accelerate a payload up to some high speed which would not be enough for orbit. It would then release the payload, which would complete the launch with rockets. This would drastically reduce the amount of velocity needed to be provided by rockets to reach orbit. Well under a tenth of orbital velocity from a small rocket thruster is enough to raise perigee if a design prioritizes minimizing such, but hybrid proposals optionally reduce requirements for the mass driver itself by having a greater portion of delta-v by a rocket burn (or orbital momentum exchange tether). On Earth, a mass-driver design could possibly use well-tested maglev components.

To launch a space vehicle with humans on board, a mass driver's track would need to be almost 1000 kilometres long if providing almost all the velocity to Low Earth Orbit, though a lesser length could provide major launch assist. Required length, if accelerating mainly at near a constant maximum acceptable g-force for passengers, is proportional to velocity squared. For instance, half of the velocity goal could correspond to a tunnel a quarter as long needing to be constructed, for the same acceleration. For rugged objects, much higher accelerations may suffice, allowing a far shorter track, potentially circular or helical (spiral). Another concept involves a large ring design whereby a space vehicle would circle the ring numerous times, gradually gaining speed, before being released into a launch corridor leading skyward.

Mass drivers have been proposed for the disposal of nuclear waste in space: a projectile launched at much above Earth's escape velocity would escape the Solar System, with atmospheric passage at such speed calculated as survivable through an elongated projectile and a very substantial heatshield.

Spacecraft-based mass drivers
A spacecraft could carry a mass driver as its primary engine. With a suitable source of electrical power (probably a nuclear reactor) the spaceship could then use the mass driver to accelerate pieces of matter of almost any sort, boosting itself in the opposite direction. At the smallest scale of reaction mass, this type of drive is called an ion drive.

No absolute theoretical limit is known for the size, acceleration or muzzle energy of linear motors. However, practical engineering constraints apply for such as the power-to-mass ratio, waste heat dissipation, and the energy intake able to be supplied and handled. Exhaust velocity is best neither too low nor too high.

There is a mission-dependent limited optimal exhaust velocity and specific impulse for any thruster constrained by a limited amount of onboard spacecraft power. Thrust and momentum from exhaust, per unit mass expelled, scales up linearly with its velocity (momentum = mv), yet kinetic energy and energy input requirements scale up faster with velocity squared (kinetic energy = $undefined 1/2$ mv2). Too low an exhaust velocity would excessively increase propellant mass needed under the rocket equation, with too high a fraction of energy going into accelerating propellant not used yet. Higher exhaust velocity has both benefit and tradeoff, increasing propellant usage efficiency (more momentum per unit mass of propellant expelled) but decreasing thrust and the current rate of spacecraft acceleration if available input power is constant (less momentum per unit of energy given to propellant).

Electric propulsion methods like mass drivers are systems where energy does not come from the propellant itself. (Such contrasts to chemical rockets where propulsive efficiency varies with the ratio of exhaust velocity to vehicle velocity at the time, but near maximum obtainable specific impulse tends to be a design goal when corresponding to the most energy released from reacting propellants). Although the specific impulse of an electric thruster itself optionally could range up to where mass drivers merge into particle accelerators with fractional-lightspeed exhaust velocity for tiny particles, trying to use extreme exhaust velocity to accelerate a far slower spacecraft could be suboptimally low thrust when the energy available from a spacecraft's reactor or power source is limited (a lesser analogue of feeding onboard power to a row of spotlights, photons being an example of an extremely low momentum to energy ratio).

For instance, if limited onboard power fed to its engine was the dominant limitation on how much payload a hypothetical spacecraft could shuttle (such as if intrinsic propellant economic cost was minor from usage of extraterrestrial soil or ice), ideal exhaust velocity would rather be around 62.75% of total mission delta v if operating at constant specific impulse, except greater optimization could come from varying exhaust velocity during the mission profile (as possible with some thruster types, including mass drivers and variable specific impulse magnetoplasma rockets).

Since a mass driver could use any type of mass for reaction mass to move the spacecraft, a mass driver or some variation seems ideal for deep-space vehicles that scavenge reaction mass from found resources.

One possible drawback of the mass driver is that it has the potential to send solid reaction mass travelling at dangerously high relative speeds into useful orbits and traffic lanes. To overcome this problem, most schemes plan to throw finely-divided dust. Alternatively, liquid oxygen could be used as reaction mass, which upon release would boil down to its molecular state. Propelling the reaction mass to solar escape velocity is another way to ensure that it will not remain a hazard.

Hybrid mass drivers
A mass driver on a spacecraft could be used to "reflect" masses from a stationary mass driver. Each deceleration and acceleration of the mass contributes to the momentum of the spacecraft. The lightweight, fast spacecraft need not carry reaction mass, and does not need much electricity beyond the amount needed to replace losses in the electronics, while the immobile support facility can run off power plants able to be much larger than the spacecraft if needed. This could be considered a form of beam-powered propulsion (a macroscopic-scale analogue of a particle beam propelled magsail). A similar system could also deliver pellets of fuel to a spacecraft to power another propulsion system.

Another theoretical use for this concept of propulsion can be found in space fountains, a system in which a continuous stream of pellets in a circular track holds up a tall structure.

Mass drivers as weapons
Small to moderate size high-acceleration electromagnetic projectile launchers are currently undergoing active research by the US Navy for use as ground-based or ship-based weapons (most often railguns but coilguns in some cases). On larger scale than weapons currently near deployment but sometimes suggested in long-range future projections, a sufficiently high velocity linear motor, a mass driver, could in theory be used as intercontinental artillery (or, if built on the Moon or in orbit, used to attack a location on Earth's surface). As the mass driver would be located further up the gravity well than the theoretical targets, it would enjoy a significant energy imbalance in terms of counter-attack.

Practical attempts
One of the first engineering descriptions of an "Electric Gun" appears in the technical supplement of the 1937 science fiction novel "Zero to Eighty" by "Akkad Pseudoman", a pen name for the Princeton physicist and electrical entrepreneur Edwin Fitch Northrup. Dr. Northrup built prototype coil guns powered by kHz-frequency three-phase electrical generators, and the book contains photographs of some of these prototypes. The book describes a fictional circumnavigation of the moon by a two-person vehicle launched by a Northrup electric gun.

Later prototype mass drivers have been built since 1976 (Mass Driver 1), some constructed by the U.S. Space Studies Institute in order to prove their properties and practicality. Military R&D on coilguns is related, as are maglev trains.

SpinLaunch, a company founded in 2014, conducted the initial test of their test accelerator in October 2021.

People

 * Eric Laithwaite and the Maglifter project
 * Gerard K. O'Neill
 * Henry Kolm

title: Mass driver