Ion thruster

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NASA's 2.3 kW NSTAR ion thruster for the Deep Space 1 spacecraft during a hot fire test at the Jet Propulsion Laboratory

An ion thruster is a form of electric propulsion used for spacecraft propulsion that creates thrust by accelerating ions. The term is strictly used to refer to gridded electrostatic ion thrusters, but may often more loosely be applied to all electric propulsion systems that accelerate plasma, since plasma consists of ions.

Ion thrusters are categorized by how they accelerate the ions, using either electrostatic or electromagnetic force. 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 either case, when an ion passes through an electrostatic grid engine, the potential difference of the electric field converts to the ion's kinetic energy.

Ion thrusters have an input power spanning 1–7 kilowatts, exhaust velocity 20–50 kilometers per second, thrust 20–250 millinewtons and efficiency 60–80%.[1][2]

The Deep Space 1 spacecraft, powered by an ion thruster, changed velocity by 4.3 km/s while consuming less than 74 kilograms of xenon. The Dawn spacecraft broke the record, reaching 10 km/s.[1][2]

Applications include control of the orientation and position of orbiting satellites (some satellites have dozens of low-power ion thrusters) and use as a main propulsion engine for low-mass robotic space vehicles (for example Deep Space 1 and Dawn).[1][2]

Ion thrusters are not the most promising type of electrically powered spacecraft propulsion (although in practice they have been more successful than others).[2] The ion drive is comparable to a car that takes two days to accelerate from zero to 60 miles per hour; a real ion engine's technical characteristics, and especially its thrust, are considerably inferior to its literary prototypes.[1][2] Technical capabilities of the ion engine are limited by the space charge created by ions. This limits the thrust density (force per cross-sectional area of the engine).[2] Ion thrusters create small thrust levels (for example the thrust of Deep Space 1's engine approximately equals the weight of one sheet of paper[2]) compared to conventional chemical rockets, but achieve very high specific impulse, or propellant mass efficiency, by accelerating their exhaust to high speed. However, ion thrusters carry a fundamental price: the power imparted to the exhaust increases with the square of its velocity while thrust increases linearly. Chemical rockets, on the other hand, can provide high thrust, but are limited in total impulse by the small amount of energy that can be stored chemically in the propellants.[3] Given the practical weight of suitable power sources, the accelerations given by ion thrusters are frequently less than one thousandth of standard gravity. However, since they operate as electric (or electrostatic) motors, a greater fraction of the input power is converted into kinetic exhaust power than in a chemical rocket. Chemical rockets operate as heat engines, hence Carnot's theorem bounds their possible exhaust velocity.

Due to their relatively high power needs, given the specific power of power supplies and the requirement of an environment void of other ionized particles, ion thrust propulsion is currently only practical on spacecraft that have already reached space, and is unable to take vehicles from Earth to space. Spacecraft rely on conventional chemical rockets to initially reach orbit.

Origins[edit]

The first person to publish mention of the idea was Konstantin Tsiolkovsky in 1911.[4] However, the first documented instance where the possibility of electric propulsion was considered is found in Robert H. Goddard's handwritten notebook in an entry dated September 6, 1906.[5] The first experiments with ion thrusters were carried out by Goddard at Clark University from 1916–1917.[6] 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” (Ways to Spaceflight), published in 1923, 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 gases.[4]

A working ion thruster was built by Harold R. Kaufman in 1959 at the NASA Glenn Research Center facilities. It was similar to the general design of a gridded electrostatic ion thruster with mercury as its fuel. Suborbital tests of the engine followed during the 1960s and in 1964 the engine was sent into a suborbital flight aboard the Space Electric Rocket Test 1 (SERT 1).[7][8] It successfully operated for the planned 31 minutes before falling back to Earth.[9] This test was followed by an orbital test, SERT-2, in 1970.[10][11]

An alternate form of electric propulsion, the Hall effect thruster was studied independently in the U.S. and the Soviet Union in the 1950s and 1960s. Hall effect thrusters had operated on Soviet satellites since 1972. Until the 1990s they were mainly used for satellite stabilization in North-South and in East-West directions. Some 100–200 engines completed their mission on Soviet and Russian satellites until the late 1990s.[12] 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 description[edit]

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 very high exhaust velocities. This reduces the amount of reaction mass or fuel required, but increases the amount of specific power required compared to chemical rockets. Ion thrusters are therefore able to achieve extremely high specific impulses. The drawback of the low thrust is low spacecraft acceleration, because the mass of current electric power units is directly correlated with the amount of power given. This low thrust makes ion thrusters unsuited for launching spacecraft into orbit, but they are ideal for in-space propulsion applications.

Various ion thrusters have been designed and they all generally fit under two categories. The thrusters are categorized as either electrostatic or electromagnetic. The main difference is how the ions are accelerated.

  • Electrostatic ion thrusters use the Coulomb force and are categorized as accelerating the ions in the direction of the electric field.
  • Electromagnetic ion thrusters use the Lorentz force to accelerate the ions.

Electric power supplies for ion thrusters are usually solar panels but, at sufficiently large distances from the Sun, nuclear power is used. In each case the power supply mass is essentially proportional to the peak power that can be supplied, and they both essentially give, for this application, no limit to the energy.

Electric thrusters tend to produce low thrust, which results in low acceleration. Using 1 g is 9.81 m/s2; F = m a ⇒ a = F/m. An NSTAR thruster producing a thrust (force) of 92 mN[13] will accelerate a satellite with a mass of 1,000 kg by 0.092 N / 1,000 kg = 0.000092 m/s2 (or 9.38×10−6 g).

thrust = 2*η*power/(g * Isp)
Where

thrust is the force in N
η is the efficiency, a dimensionless value between 0 and 1 (70% efficiency is 0.7)
power is the electrical energy going into the thruster in W
g is a constant, the acceleration due to gravity 9.81 m/s2
Isp is the Specific impulse in s

Electrostatic ion thrusters[edit]

Gridded electrostatic ion thrusters[edit]

Figure 2: A diagram of how a gridded electrostatic ion engine (multipole magnetic cusp type) works

Gridded electrostatic ion thrusters commonly utilize xenon gas. This gas has no charge and is ionized by bombarding it with energetic electrons. These electrons can be provided from a hot cathode filament and when accelerated in the electrical field of the cathode, fall to the anode. Alternatively, the electrons can be accelerated by the oscillating electric field induced by an alternating magnetic field of a coil, which results in a self-sustaining discharge and omits any cathode (radio frequency ion thruster).

The positively charged ions are extracted by an extraction system consisting of 2 or 3 multi-aperture grids. After entering the grid system via the plasma sheath the ions are accelerated due to the potential difference between the first and second grid (named screen and accelerator grid) to the final ion energy of typically 1–2 keV, thereby generating the thrust.

Ion thrusters emit a beam of positive charged xenon ions only. To avoid charging up the spacecraft, another cathode is placed near the engine, which emits electrons (basically the electron current is the same as the ion current) into the ion beam.[9] This also prevents the beam of ions from returning to the spacecraft and cancelling the thrust.[citation needed]

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
  • 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)[14][15]
Schematic of a Hall Thruster

Hall effect thrusters[edit]

Hall effect thrusters accelerate ions with the use of an electric potential maintained between a cylindrical anode and a negatively charged plasma that forms the cathode. The bulk of the propellant (typically xenon gas) is introduced near the anode, where it becomes ionized, and the ions are attracted towards the cathode; they 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, and 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 far more affected and 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 closing the circuit.[16]

Field-emission electric propulsion[edit]

Field-emission electric propulsion (FEEP) thrusters use a very simple system of accelerating ions to create thrust. Most designs use either caesium or indium as the propellant. 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 ("Taylor cones"). At a sufficiently high applied voltage, positive ions are extracted from the tips of the cones.[17][18][19] 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.

Electromagnetic thrusters[edit]

Pulsed inductive thrusters (PIT)[edit]

Pulsed inductive thrusters (PIT) use pulses of thrust instead of one continuous thrust, and have the ability to run on power levels in 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 commonly used in PIT engines. For each pulse of thrust the PIT gives, a large charge first 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 ammonia gas that has just been released in the opposite direction of the original current. This opposite current ionizes the ammonia and these positively charged ions are accelerated away from the PIT engine due to the electric field jθ crossing with the magnetic field Br, which is due to the Lorentz Force.[20]

Magnetoplasmadynamic (MPD) / lithium Lorentz force accelerator (LiLFA)[edit]

Magnetoplasmadynamic (MPD) thrusters and lithium Lorentz force accelerator (LiLFA) thrusters use roughly the same idea with the LiLFA thruster building off of the MPD thruster. Hydrogen, argon, ammonia, and nitrogen gas 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 first 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. 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, except for two main differences. The first difference is that the LiLFA uses lithium vapor, which has the advantage of being able to be stored as a solid. The other difference is that the cathode is replaced by multiple smaller cathode rods packed into a hollow cathode tube. The cathode in the MPD thruster is 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.[21][22][dead link][23][dead link]

In 2013 Russian company the Chemical Automatics Design Bureau successfully conducted a bench test of their magnetoplasmadynamic engine for long-distance space travel.[24]

Electrodeless plasma thrusters[edit]

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 takes away the factor of 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 stage give the engine the ability to throttle the speed of propellant flow, which then changes the thrust magnitude and specific impulse values.[25]

Helicon double layer thruster[edit]

A helicon double layer thruster is a type of plasma thruster, which ejects high velocity ionized gas to provide thrust to a spacecraft. In this thruster 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 the plasma. 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, there is a sharp boundary between 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. The 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.

Microwave Electrothermal Thrusters[edit]

Thruster Components
Thruster Components
Discharge Chamber
Discharge Chamber
Sketch of the Microwave Electrothermal Thruster. 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.[26]

Under a research grant from the NASA Lewis Research Center during the 1980s and 1990s, professors Martin C. Hawley and Jes Assmussen led a team of the following engineers from Michigan State University in developing a Microwave Electrothermal Thruster (MET).[27]

  1. Terrence J. Morin focused his work solely upon a diatomic gas, primarily hydrogen, modeling collision induced heating in a non equilibrium environment for a weakly ionized plasma with less than 0.1% ionization. He combined the use of statistical mechanics and kinetic theory. Using the Boltzmann based kinetic theory of gases, he applied simple chemical reaction models of Plug Flow Reactor (PFR) and Continuous Stirred-Tank Reactor (CSTR) models. [28][29][30][31]
  2. Randall A. Chapman conducted many experiments using a microwave cavity system. His research focused on hydrogen gas at low pressures (0.5 - 10 torr) and at low microwave power (20 - 100 watts). He showed a 20% net power absorption to the exiting gas from the cavity system, calculated a vibrational temperature range of 4,000 - 17,000 Kelvin, and measured an ionization percentage of between 0.001 and 0.1 %, which demonstrated that the electron density increased with pressure and energy (temperature).[28][32][31]
  3. Stanley J. Whitehair obtained experimental measurements of microwave coupling efficiencies, thruster energy efficiencies, and specific impulses for nitrogen and helium propellants over discharge pressures ranging between 40-1,100 torr with varying propellent flow rates. With efficiencies in excess of 95%, he established that the MET is a very efficient thruster. However, he noted that melting and erosion of the nozzles limit the input power and specific impulses.[33][34][35][31]
  4. Scott S. Haraburda developed a simple equilibrium based theory of space-dependent parameters for transport design equations, using helium as the monatomic gas and nitrogen as the diatomic gas.[36][37][38][26][39][40] Applying the Finite Element Method (FEM) within the plasma in a Batch Reactor model of the TM011 and TM012 electromagnetic resonance modes of a microwave cavity, he predicted the residence time of the reaction to equilibrium. With his data, he conducted many simulations on NASA's Two-Dimensional Kinetic (TDK) computer program to determine the effects on engine performance from pressure and energy changes along with propellant contamination. He determined that one should use a minimum power of 1 kWatt to obtain specific impulses greater than 1,000 lbf*sec/lb.[41][42] He also determined that fouling of the walls in the discharge chamber significantly reduces the amount of energy transferred to the propellant, thus reducing the rocket efficiencies.

Comparisons[edit]

The following table compares actual test data of some ion thrusters:

Engine Propellant Required power
(kW)
Specific impulse
(s)
Thrust
(mN)
Thruster mass
(kg)
NSTAR Xenon 2.3 3,300 to 1,700[43] 92 max.[13]
NEXT[13] Xenon 6.9[44] 4,190[44][45][46] 236 max[13][46]
NEXIS[47] Xenon 20.5
HiPEP Xenon 20–50[48] 6,000–9,000[48] 460–670[48]
RIT 22[49] Xenon 5
Hall effect Bismuth 25[citation needed]
Hall effect Bismuth 140[citation needed]
Hall effect Xenon 25[citation needed] 3,250[citation needed] 950[citation needed]
Hall effect Xenon 75[citation needed]
FEEP Liquid Caesium 6×10−5–0.06 6,000–10,000[18] 0.001–1[18]
VASIMR Argon 200 3,000–12,000 ~5,000[50] 620[51]
CAT[52] Xenon, Iodine, water[53] 0.01 690[54] 1.1-2 <1 (73 mN/kW)[53]
DS4G Xenon 250 19,300 2,500 max. 5
KLIMT Krypton 0.5[55] 4[55]

The following thrusters are highly experimental and have been tested only in pulse mode.

Engine Propellant Required power
(kW)
Specific impulse
(s)
Thrust
(mN)
Thruster mass
(kg)
MPDT Hydrogen 1,500 4,900[citation needed] 26,300[citation needed]
MPDT Hydrogen 3,750 3,500[citation needed] 88,500[citation needed]
MPDT Hydrogen 7,500[citation needed] 6,000[citation needed] 60,000[citation needed]
LiLFA Lithium Vapor 500 4,077[citation needed] 12,000[citation needed]

Lifetime[edit]

A major limiting factor of ion thrusters is their small thrust; however, it is generated at a high propellant efficiency (mass utilisation, specific impulse). The efficiency comes from the high exhaust velocity, which in turn demands high energy, and the performance is ultimately limited by the available spacecraft power.

The low thrust requires ion thrusters to provide continuous thrust for a long time to achieve the needed change in velocity (delta-v) for a particular mission. To cause enough change in momentum, ion thrusters are designed to last for periods of weeks to years.

In practice the lifetime of electrostatic ion thrusters is limited by several processes:

  • In electrostatic gridded ion thruster design, 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 a structural failure of the grid occurs or the holes in the accelerator grid become so large that the ion extraction is largely affected; e.g., by the occurrence of electron backstreaming. Grid erosion cannot be avoided and is the major lifetime-limiting factor. By a thorough grid design and material selection, lifetimes of 20,000 hours and far beyond are reached, which is sufficient to fulfil current space missions.

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. The test was concluded prior to any failure and examination indicated the engine was not approaching failure either.[56]

More recently, the NASA Evolutionary Xenon Thruster (NEXT) Project, conducted at NASA's Glenn Research Center in Cleveland, Ohio, operated continuously for more than 48,000 hours.[57] The test was conducted in a high vacuum test chamber at Glenn Research Center. Over the course of the 5 1/2 + year test, the engine consumed approximately 870 kilograms of xenon propellant. The total impulse provided by the engine would require over 10,000 kilograms of conventional rocket propellant for similar application. The engine was designed by Aerojet Rocketdyne of Sacramento, California.

  • Hall thrusters suffer from very strong erosion of the ceramic discharge chamber by impact of energetic ions: a test reported in 2010[58] 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.

NASA's Jet Propulsion Laboratory has created ion drives with a time of continuous operation of more than 3 years.[1][2]

Propellants[edit]

Ionization energy represents a very large percentage of the energy needed to run ion drives. The ideal propellant for ion drives is thus a propellant molecule or atom that is easy to ionize, that has a high mass/ionization energy ratio. In addition, the propellant should not cause erosion of the thruster to any great degree to permit long life; and should not contaminate the vehicle.[59]

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 very expensive.

Older designs used mercury, but this is toxic and expensive, tended to contaminate the vehicle with the metal and was difficult to feed accurately.

Other propellants, such as bismuth, show promise and are areas of research, particularly for gridless designs, such as Hall effect thrusters.

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 a relatively abundant and inexpensive gas.

The CubeSat Ambipolar Thruster (CAT) used on the Mars Array of Ionospheric Research Satellites Using the CubeSat Ambipolar Thruster (MARS-CAT) mission proposes to use solid Iodine as the propellant to minimize storage volume.[54]

Energy efficiency[edit]

Plot of instantaneous propulsive efficiency (blue) and overall efficiency for a vehicle accelerating from rest (red) as percentages of the engine efficiency- note that peak vehicle efficiency occurs at about 1.6 times exhaust velocity.

Ion thrusters are frequently quoted with an efficiency metric. This efficiency is the kinetic energy of the exhaust jet emitted per second divided by the electrical power into the device.

The actual overall system energy efficiency in use 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 Isps the overall efficiency drops, because the ionization takes up a larger percentage energy, and at the high end propulsive efficiency is reduced.

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

Applications[edit]

Ion thrusters have many applications for in-space propulsion. The best applications of the thrusters make use of the long lifetime when significant thrust is not needed. Examples of this include orbit transfers, attitude adjustments, drag compensation for low Earth orbits, transporting cargo such as chemical fuels between propellant depots and ultra-fine adjustments for scientific missions. Ion thrusters can also be used for interplanetary and deep-space missions where time is not crucial. Continuous thrust over a very long time can build up a larger velocity than traditional chemical rockets.

Missions[edit]

Of all the electric thrusters, ion thrusters have been the most seriously considered commercially and academically in the quest for interplanetary missions and orbit raising maneuvers. Ion thrusters are seen as the best solution for these missions, as they require very high change in velocity overall that can be built up over long periods of time.

Pure demonstration vehicles[edit]

SERT

Ion propulsion systems were first demonstrated in space by the NASA Lewis (now Glenn Research Center) missions "Space Electric Rocket Test" (SERT) I and II.[60] The first was SERT-1, launched July 20, 1964, successfully proved that the technology operated as predicted in space. These were electrostatic ion thrusters using mercury and cesium as the reaction mass. The second test, SERT-II, launched on February 3, 1970,[61][62] verified the operation of two mercury ion engines for thousands of running hours.[63]

Operational missions[edit]

Ion thrusters are routinely used for station-keeping on commercial and military communication satellites in geosynchronous orbit, including satellites manufactured by Boeing and by Hughes Aerospace. The pioneers in this field were the Soviet Union, who used SPT thrusters on a variety of satellites starting in the early 1970s.

Two geostationary satellites (ESA's Artemis in 2001–03[64] and the US military's AEHF-1 in 2010–12[65]) have used the ion thruster for orbit raising after the failure of the chemical-propellant engine. Boeing[66] have been using ion thrusters for station-keeping since 1997, and plan in 2013–14 to offer a variant on their 702 platform, which will have no chemical engine and use ion thrusters for orbit raising; this enables a significantly lower launch mass for a given satellite capability. AEHF-2 used a chemical engine to raise perigee to 10150 miles and is then proceeding to geosynchronous orbit using electric propulsion.[67]

In Earth orbit[edit]

GOCE

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

In deep space[edit]

Deep Space 1

NASA developed the NSTAR ion engine for use in their interplanetary science missions beginning in the late-1990s. This xenon-propelled ion thruster was first space-tested in the highly successful 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.[60] Based on the NASA design criteria, Hughes Research Labs, developed the XIPS (Xenon Ion Propulsion System) for performing station keeping on geosynchronous satellites.[citation needed]. Hughes (EDD) manufactured the NSTAR thruster used on the spacecraft.

Hayabusa

The Japanese space agency's Hayabusa, which was launched in 2003 and successfully rendezvoused with the asteroid 25143 Itokawa and remained in close proximity for many months to collect samples and information, was powered by four xenon ion engines. It used xenon ions generated by microwave electron cyclotron resonance, and a carbon / carbon-composite material (which is resistant to erosion) for its acceleration grid.[68] Although the ion engines on Hayabusa had some technical difficulties, in-flight reconfiguration allowed one of the four engines to be repaired, and allowed the mission to successfully return to Earth.[69]

Smart 1

The European Space Agency's satellite SMART-1, launched in 2003, used a Snecma PPS-1350-G Hall thruster to get from GTO to lunar orbit. This satellite completed its mission on September 3, 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 was launched on September 27, 2007, to explore the asteroid Vesta and the dwarf planet Ceres. To cruise from Earth to its targets it uses three Deep Space 1 heritage xenon ion thrusters (firing only one at a time) to take it in a long outward spiral. An extended mission in which Dawn explores other asteroids after Ceres is also possible. Dawn's ion drive is capable of accelerating from 0 to 60 mph (97 km/h) in 4 days, firing continuously.[70]

Planned missions[edit]

In addition, several missions are planned to use ion thrusters in the next few years.

BepiColombo

ESA will launch the BepiColombo mission to Mercury in 2016. It uses ion thrusters in combination with swing-bys to get to Mercury, where a chemical rocket will be fired for orbit insertion.

LISA Pathfinder

LISA Pathfinder is an ESA spacecraft to be launched in 2015. It will not use ion thrusters as its primary propulsion system, but will use both colloid thrusters and FEEP for very precise attitude control -— the low thrusts of these propulsion devices make it possible to move the spacecraft incremental distances very accurately. It is a test for the possible LISA mission.

International Space Station

As of March 2011, a future launch of an Ad Astra VF-200 200 kW VASIMR electromagnetic thruster was being considered for placement and testing on the International Space Station.[71][72] The VF-200 is a flight version of the VX-200.[73][74] Since the available power from the ISS is less than 200 kW, the ISS VASIMR will include a trickle-charged battery system allowing for 15 min pulses of thrust. Testing of the engine on ISS is valuable, because ISS orbits at a relatively low altitude and experiences fairly high levels of atmospheric drag, making periodic boosts of altitude necessary. Currently, altitude reboosting by chemical rockets fulfills this requirement. If the tests of VASIMR reboosting of the ISS goes according to plan, the increase in specific impulse could mean that the cost of fuel for altitude reboosting will be one-twentieth of the current $210 million annual cost.[71] Hydrogen is generated by the ISS as a by-product, which is currently vented into space.

NASA high-power SEP system demonstration mission

In June 2011, NASA launched a request-for-proposals[75] for a test mission (from context probably using the NEXT engine) capable of being extended to 300 kW electrical power; this was awarded to Northrop Grumman in February 2012.[76]

Mars Array of Ionospheric Research Satellites Using the CubeSat Ambipolar Thruster

The Mars Array of Ionospheric Research Satellites Using the CubeSat Ambipolar Thruster (MARS-CAT) mission is a two 6U CubeSat mission to study the ionosphere of Mars. The mission will investigate the plasma and magnetic structure of the Martian ionosphere, including transient plasma structures, magnetic field structure, magnetic activity and correlation with solar wind drivers. The Mars transit proposed is piggy back with Mars2020 using a CubeSat Ambipolar Thruster (CAT) burn for Mars orbit insertion and station keeping.[54]

Proposal[edit]

Geoffrey A. Landis proposed for interstellar travel future-technology project interstellar probe with supplying the energy from an external source (laser of base station) and ion thruster.[77][78]

See also[edit]

References[edit]

  1. ^ a b c d e Choueiri, Edgar Y. (2009). New dawn of electric rocket. The Ion Drive
  2. ^ a b c d e f g h Choueiri, Edgar Y (2009). "New dawn of electric rocket". Scientific American 300: 58–65. doi:10.1038/scientificamerican0209-58. (subscription required (help)). 
  3. ^ Electric Spacecraft Propulsion, Electric versus Chemical Propulsion, ESA Science & Technology
  4. ^ a b Choueiri, E. Y. "A Critical History of Electric Propulsion: The First 50 Years (1906–1956)" (PDF). Retrieved 2007-11-07. 
  5. ^ Mark Wright, April 6, 1999, science.nasa.gov, Ion Propulsion 50 years in the making
  6. ^ "Robert H. Goddard: American Rocket Pioneer". Smithsonian Scrapbook. Smithsonian Institution Archives. Retrieved March 28, 2012. 
  7. ^ NASA Glenn Contributions to Deep Space 1
  8. ^ Ronald J. Cybulski, Daniel M. Shellhammer, Robert R. LoveII, Edward J. Domino, and Joseph T. Kotnik, RESULTS FROM SERT I ION ROCKET FLIGHT TEST, NASA Technical Note D2718 (1965).
  9. ^ a b "Innovative Engines - Glenn Ion Propulsion Research Tames the Challenges of 21st Century Space Travel". Retrieved 2007-11-19. 
  10. ^ NASA Glenn, "SPACE ELECTRIC ROCKET TEST II (SERT II) (Accessed July 1, 2010)
  11. ^ SERT page at Astronautix (Accessed July 1, 2010)
  12. ^ "Native Electric Propulsion Engines Today" (in Russian) (7). Novosti Kosmonavtiki. 1999. Archived from the original on 6 June 2011. 
  13. ^ a b c d Shiga, David (2007-09-28). "Next-generation ion engine sets new thrust record". NewScientist. Retrieved 2011-02-02. 
  14. ^ "ESA and ANU make space propulsion breakthrough" (Press release). ESA. 2006-01-11. Retrieved 2007-06-29. 
  15. ^ ANU Space Plasma, Power & Propulsion Group (SP3) (2006-12-06). "ANU and ESA make space propulsion breakthrough". DS4G Web Story. The Australian National University. Archived from the original on 2007-06-27. Retrieved 2007-06-30. 
  16. ^ Oleson, S. R.; Sankovic, J. M. "Advanced Hall Electric Propulsion for Future In-Space Transportation" (PDF). Retrieved 2007-11-21. 
  17. ^ "FEEP - Field-Emission Electric Propulsion". Retrieved 2012-04-27. 
  18. ^ a b c Marcuccio, S. et al. "Experimental Performance of Field Emission Microthrusters" (PDF). Retrieved 2012-04-27. 
  19. ^ Marrese-Reading, Colleen; Polk, Jay; Mueller, Juergen; Owens, Al. "In-FEEP Thruster Ion Beam Neutralization with Thermionic and Field Emission Cathodes" (PDF). Retrieved 2007-11-21. liquid state and wicked up the needle shank to the tip where high electric fields deform the liquid and extract ions and accelerate them up to 130 km/s through 10 kV. 
  20. ^ Mikellides, Pavlos G. "Pulsed Inductive Thruster (PIT): Modeling and Validation Using the MACH2 Code" (PDF). Retrieved 2007-11-21. 
  21. ^ Sankaran, K.; Cassady, L.; Kodys, A.D.; Choueiri, E.Y. "A Survey of Propulsion Options for Cargo and Piloted Missions to Mars". Retrieved 2007-11-21. 
  22. ^ LaPointe, Michael R.; Mikellides, Pavlos G. "High Power MPD Thruster Development at the NASA Glenn Research Center" (PDF). Retrieved 2007-11-21. 
  23. ^ Conley, Buford Ray (May 22, 1999). "Utilization of Ambient Gas as a Propellant for Low Earth Orbit Electric Propulsion" (PDF). 
  24. ^ http://sdelanounas.ru/blogs/44948
  25. ^ Emsellem, Gregory D. "Development of a High Power Electrodeless Thruster" (PDF). Retrieved 2007-11-21. 
  26. ^ a b Haraburda, Scott (June 1992). "Developmental Research for Designing a Microwave Electrothermal Thruster". 18th Army Science Conference 2. Orlando, Florida. pp. 15–29. Retrieved 24 July 2015. 
  27. ^ "Less Fuel, More Thrust: New Engines are Being Designed for Deep Space". The Arugus-Press 128 (48) (Owosso, Michigan). 26 February 1982. p. 10. 
  28. ^ a b Chapman, Randall A.; Filpus, John W.; Morin, Terrence J.; Snellenberger, Reed; Asmussen, Jes; Hawley, Martin C.; Kerber, Ronald L. (1982). "Microwave Plasma Generation of Hydrogen Atoms for Rocket Propulsion". Journal of Spacecraft and Rockets 19 (6): 579–585. Retrieved 24 July 2015. 
  29. ^ Morin, Terrence J. (1985). Collision Induced Heating of a Weakly Ionized Dilute Gas in Steady Flow (Ph.D.). Michigan State University. 
  30. ^ Morin, Terrence J.; Hawley, Martin C. (1987). "The Efficacy of Heating Low-Pressure H2 in a Microwave Discharge". Plasma Chemistry and Plasma Processinq 7 (2): 181–199. Retrieved 24 July 2015. 
  31. ^ a b c Hawley, Martin C.; Asmussen, Jes; Filpus, John W.; Whitehair, Stanley J.; Hoekstra, Craig; Morin, Terrence J.; Chapman, Randall A. (1989). "Review of research and development on the microwave electrothermal thruster". Journal of Propulsion and Power 5 (6): 703–712. Retrieved 24 July 2015. 
  32. ^ Chapman, Randall A. (1986). Energy Distribution and Transfer in Flowing Hydrogen Microwave Plasmas (Ph.D.). Michigan State University. Retrieved 24 July 2015. 
  33. ^ Whitehair, Stanley J. (1986). Experimental Development of a Microwave Electrothermal Thruster (Ph.D.). Michigan State University. 
  34. ^ Whitehair, Stanley J.; Frasch, Lydell L.; Asmussen, Jes (May 1987). "Experimental Performance of a Microwave Electrothermal Thruster with High Temperature Nozzle Materials". 19th International Electric Propulsion Conference. Colorado Springs, Colorado: AIAA/SAE/ASME/ASEE. Retrieved 24 July 2015. 
  35. ^ Whitehair, Stanley J.; Asmussen, Jes; Nakanshi, Shigeo (1987). "Microwave Electrothermal Thruster Performance in Helium Gas". Journal of Propulsion and Power 3 (2): 136–144. Retrieved 24 July 2015. 
  36. ^ Haraburda, Scott S.; Hawley, Martin C. (July 1989). "Investigations of Microwave Plasmas (Applications in Electrothermal Thruster Systems)". 25th Joint Propulsion Conference. Monterey, California: AIAA/ASME/SAE/ASEE. doi:10.2514/3.62304. Retrieved 24 July 2015. 
  37. ^ Haraburda, Scott S.; Hawley, Martin C. (July 1990). "Diagnostic Evaluations of Microwave Generated Helium and Nitrogen Plasma Mixtures". 21st International Electric Propulsion Conference. Orlando, Florida: AIAA/DGLR/JSASS. Retrieved 24 July 2015. 
  38. ^ Haraburda, Scott S.; Hawley, Martin C.; Dinkel, Duane W. (October 1991). "Diagnostic Evaluations of Microwave Generated Helium and Nitrogen Plasma Mixtures". 21nd International Electric Propulsion Conference. Viareggio, Italy: AIDAA/AIAA/DGLR/JSASS. 
  39. ^ Haraburda, Scott S.; Hawley, Martin C. (July 1992). "Theoretical Nozzle Performance of a Microwave Electrothermal Thruster Using Experimental Data". 28th Joint Propulsion Conference. Nashville, Tennessee: AIAA/SAE/ASME/ASEE. Retrieved 24 July 2015. 
  40. ^ Haraburda, Scott S.; Hawley, Martin C.; Asmussen, Jes (September 1992). "Review of Experimental and Theoretical Research on the Microwave Electrothermal Thruster". 43rd Congress (World Space Congress). Washington, DC: International Astronautical Federation. 
  41. ^ Nickerson, G.R.; Dang, L.D.; Coats, D.E., Engineering and programming manual: Two-dimensional kinetic reference computer program (TDK) (PDF), NASA, retrieved 24 July 2015 
  42. ^ Haraburda, Scott S. (2001). Transport phenomena of flow through helium and nitrogen plasmas in microwave electrothermal thrusters (Ph.D.). Michigan State University. Retrieved 23 July 2015. 
  43. ^ ION PROPULSION
  44. ^ a b Szondy, David. "NASA's NEXT ion thruster runs five and a half years nonstop to set new record". Retrieved June 26, 2013. 
  45. ^ Schmidt, George R.; Patterson, Michael J.; Benson, Scott W. "The NASA Evolutionary Xenon Thruster (NEXT): the next step for US deep space propulsion" (PDF). 
  46. ^ a b Herman, Daniel A. (May 3–7, 2010), "NASA’s Evolutionary Xenon Thruster (NEXT) Project Qualifi cation Propellant Throughput Milestone: Performance, Erosion, and Thruster Service Life Prediction After 450 kg" (PDF), 57th Joint Army-Navy-NASA-Air Force (JANNAF) Propulsion Meeting, Colorado Springs, Colorado, USA: NASA - Glenn Research Center, retrieved 2014-03-08 
  47. ^ An overview of the Nuclear Electric Xenon Ion System (NEXIS) program (2006), 2006-02-10 (Polk, Jay E., Goebel, Don, Brophy, John R., Beatty, John, Monheiser, J., Giles, D.), Scientific Commons
  48. ^ a b c "HIGH POWER ELECTRIC PROPULSION PROGRAM (HiPEP)". NASA. 2008-12-22. 
  49. ^ Astrium Radiofrequency Ion Thruster, Model RIT-22., EADS Astrium[dead link]Archive copy at the Wayback Machine
  50. ^ VASIMR VX-200 Performance and Near-term SEP Capability for Unmanned Mars Flight, Tim Glover, Future in Space Operations (FISO) Colloquium, 2011-01-19, accessed 2011-01-31.
  51. ^ [1]
  52. ^ Mike Wall (July 8, 2013). "New Space Engine Could Turn Tiny CubeSats into Interplanetary Explorers". Space.com (Purch). Retrieved June 25, 2015. 
  53. ^ a b "PEPL Thrusters: CubeSat Ambipolar Thruster". pepl.engin.umich.edu. University of Michigan. Retrieved June 25, 2015. 
  54. ^ a b c "MARS-CAT Mission Implementation". www.marscat.space. University of Houston College of Natural Sciences and Mathematics. Retrieved June 25, 2015. 
  55. ^ a b Krypton Hall effect thruster for space propulsion, IFPiLM, accessed 2014-01-29.
  56. ^ "Destructive Physical Analysis of Hollow Cathodes from the Deep Space 1 Flight Spare Ion Engine 30,000 Hr Life Test" (PDF). Retrieved 2007-11-21. 
  57. ^ "NASA Thruster Achieves World-Record 5+ Years of Operation". Retrieved 2012-06-27. 
  58. ^ "A closer look at a stationary plasma thruster" (PDF). 
  59. ^ Rocket Propulsion Elements — Sutton & Biblarz 7th edition
  60. ^ a b Sovey, J. S.; Rawlin, V. K.; Patterson, M. J. (May–June 2001). "Ion Propulsion Development Projects in U. S.: Space Electric Rocket Test 1 to Deep Space 1". Journal of Propulsion and Power 17 (3): 517–526. doi:10.2514/2.5806. 
  61. ^ SPACE ELECTRIC ROCKET TEST II (SERT II), NASA Glenn Research Center. Accessed July 1, 2010.
  62. ^ SERT page at Astronautix (Accessed July 1, 2010)
  63. ^ Space Electric Rocket Test at NASA.gov
  64. ^ ESA. "Artemis team receives award for space rescue". Retrieved 2006-11-16. 
  65. ^ "Rescue in Space". 
  66. ^ Spaceflight Now. "Electric propulsion could launch new commercial trend". 
  67. ^ "AEHF 2 communications satellite keeps on climbing". 
  68. ^ ISAS. "小惑星探査機はやぶさ搭載イオンエンジン (Ion Engines used on Asteroid Probe Hayabusa)" (in Japanese). Retrieved 2006-10-13. 
  69. ^ Tabuchi, Hiroko (1 July 2010). "Faulty Space Probe Seen as Test of Japan’s Expertise". The New York Times. 
  70. ^ The Prius of Space, September 13, 2007, NASA Jet Propulsion Laboratory
  71. ^ a b "Executive summary" (PDF). Ad Astra Rocket Company. January 24, 2010. Retrieved 2010-02-27. 
  72. ^ Klotz, Irene (7 August 2008). "Plasma Rocket May Be Tested at Space Station". Discovery News. Retrieved 2010-02-27. 
  73. ^ Whittington, Mark (March 10, 2011). "NASA to Test VF-200 VASIMR Plasma Rocket at the ISS". Yahoo. Retrieved 2012-01-27. 
  74. ^ Mick, Jason (August 11, 2008). "Commercially Developed Plasma Engine Soon to be Tested in Space". DailyTech. Retrieved 2010-02-27. 
  75. ^ "Solar Electric Propulsion System Demonstration Mission Concept Studies". 
  76. ^ "NASA Awards Solar Electric Propulsion Flight System Contract". 
  77. ^ Laser-Powered Interstellar Probe G Landis - APS Bulletin, 1991
  78. ^ Geoffrey A. Landis. Laser-powered Interstellar Probe on the Geoffrey A. Landis: Science. papers available on the web
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