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Hybrid-propellant rocket

A hybrid-propellant rocket is a rocket with a rocket motor that uses rocket propellants in two different phases: one solid and the other either gas or liquid. The hybrid rocket concept can be traced back to the early 1930s.

Hybrid rockets avoid some of the disadvantages of solid rockets like the dangers of propellant handling, while also avoiding some disadvantages of liquid rockets like their mechanical complexity. Because it is difficult for the fuel and oxidizer to be mixed intimately (being different states of matter), hybrid rockets tend to fail more benignly than liquids or solids. Like liquid rocket engines, hybrid rocket motors can be shut down easily and the thrust is throttleable. The theoretical specific impulse ($$I_{sp}$$) performance of hybrids is generally higher than solid motors and lower than liquid engines. $$I_{sp}$$ as high as 400 s has been measured in a hybrid rocket using metalized fuels. Hybrid systems are more complex than solid ones, but they avoid significant hazards of manufacturing, shipping and handling solid rocket motors by storing the oxidizer and the fuel separately.

History
The first work on hybrid rockets was performed in the early 1930s at the Soviet Group for the Study of Reactive Motion. Mikhail Klavdievich Tikhonravov, who would later supervise the design of Sputnik I and the Luna programme, was responsible for the first hybrid propelled rocket launch, the GIRD-9, on 17 August 1933, which reached an altitude of 400 m. In the late 1930s at IG Farben in Germany and concurrently at the California Rocket Society in the United States. Leonid Andrussow, working in Germany, theorized hybrid propellant rockets. O. Lutz, W. Noeggerath, and Andrussow tested a 10 kN hybrid rocket motor using coal and gaseous N2O as the propellants. Oberth also worked on a hybrid rocket motor using LOX as the oxidizer and graphite as the fuel. The high heat of sublimation of carbon prevented these rocket motors from operating efficiently, as it resulted in a negligible burning rate.

In the 1940s, the California Pacific Rocket Society used LOX in combination with several different fuel types, including wood, wax, and rubber. The most successful of these tests was with the rubber fuel, which is still the dominant fuel in use today. In June 1951, a LOX / rubber rocket was flown to an altitude of 9 km.

Two major efforts occurred in the 1950s. One of these efforts was by G. Moore and K. Berman at General Electric. The duo used 90% high test peroxide (HTP, or H2O2) and polyethylene (PE) in a rod and tube grain design. They drew several significant conclusions from their work. The fuel grain had uniform burning. Grain cracks did not affect combustion, like it does with solid rocket motors. No hard starts were observed (a hard start is a pressure spike seen close to the time of ignition, typical of liquid rocket engines). The fuel surface acted as a flame holder, which encouraged stable combustion. The oxidizer could be throttled with one valve, and a high oxidizer to fuel ratio helped simplify combustion. The negative observations were low burning rates and that the thermal instability of peroxide was problematic for safety reasons. Another effort that occurred in the 1950s was the development of a reverse hybrid. In a standard hybrid rocket motor, the solid material is the fuel. In a reverse hybrid rocket motor, the oxidizer is solid. William Avery of the Applied Physics Laboratory used jet fuel and ammonium nitrate, selected for their low cost. His O/F ratio was 0.035, which was 200 times smaller than the ratio used by Moore and Berman.

In 1953 Pacific Rocket Society (est. 1943) was developing the XDF-23, a 4 × hybrid rocket, designed by Jim Nuding, using LOX and rubber polymer called "Thiokol". They had already tried other fuels in prior iterations including cotton, paraffin wax and wood. The XDF name itself comes from "experimental Douglas fir" from one of the first units. In the 1960s, European organizations also began work on hybrid rockets. ONERA, based in France, and Volvo Flygmotor, based in Sweden, developed sounding rockets using hybrid rocket motor technology. The ONERA group focused on a hypergolic rocket motor, using nitric acid and an amine fuel, developing the LEX sounding rocket. The company flew eight rockets: Once in April 1964, three times in June 1965, and four times in 1967. The maximum altitude the flights achieved was over 100 km. The Volvo Flygmotor group also used a hypergolic propellant combination. They also used nitric acid for their oxidizer, but used Tagaform (polybutadiene with an aromatic amine) as their fuel. Their flight was in 1969, lofting a 20 kg payload to 80 km.

Meanwhile, in the United States, United Technologies Center (Chemical Systems Division) and Beech Aircraft were working on a supersonic target drone, known as Sandpiper. It used MON-25 (mixed 25% NO, 75% N2O4) as the oxidizer and polymethyl methacrylate (PMM) and Mg for the fuel. The drone flew six times in 1968, for more than 300 seconds and to an altitude greater than 160 km. The second iteration of the rocket, known as the HAST, had IRFNA-PB/PMM for its propellants and was throttleable over a 10/1 range. HAST could carry a heavier payload than the Sandpiper. Another iteration, which used the same propellant combination as the HAST, was developed by Chemical Systems Division and Teledyne Aircraft. Development for this program ended in the mid-1980s. Chemical Systems Division also worked on a propellant combination of lithium and FLOx (mixed F2 and O2). This was an efficient hypergolic rocket that was throttleable. The vacuum specific impulse was 380 seconds at 93% combustion efficiency.

American Rocket Company (AMROC) developed the largest hybrid rockets ever created in the late 1980s and early 1990s. The first version of their engine, fired at the Air Force Phillips Laboratory, produced 312000 N of thrust for 70 seconds with a propellant combination of LOX and hydroxyl-terminated polybutadiene (HTPB) rubber. The second version of the motor, known as the H-250F, produced more than 1000000 N of thrust.

Korey Kline of Environmental Aeroscience Corporation (eAc) first fired a gaseous oxygen and rubber hybrid in 1982 at Lucerne Dry Lake, CA, after discussions on the technology with Bill Wood, formerly with Westinghouse. The first SpaceShipOne hybrid tests were successfully conducted by Kline and eAc at Mojave, CA.

In 1994, the U.S. Air Force Academy flew a hybrid sounding rocket to an altitude of 5 km. The 6.4 m rocket used HTPB and LOX for its propellant, and reached a peak thrust of 4400 N and had a thrust duration of 16 seconds.

Basic concepts

In its simplest form, a hybrid rocket consists of a pressure vessel (tank) containing the liquid oxidizer, the combustion chamber containing the solid propellant, and a mechanical device separating the two. When thrust is desired, a suitable ignition source is introduced in the combustion chamber and the valve is opened. The liquid oxidiser (or gas) flows into the combustion chamber where it is vaporized and then reacted with the solid propellant. Combustion occurs in a boundary layer diffusion flame adjacent to the surface of the solid propellant.

Generally, the liquid propellant is the oxidizer and the solid propellant is the fuel because solid oxidizers are extremely dangerous and lower performing than liquid oxidizers. Furthermore, using a solid fuel such as Hydroxyl-terminated polybutadiene (HTPB) or paraffin wax allows for the incorporation of high-energy fuel additives such as aluminium, lithium, or metal hydrides.

Combustion
The governing equation for hybrid rocket combustion shows that the regression rate is dependent on the oxidizer mass flux rate, which means the rate that the fuel will burn is proportional to the amount of oxidizer flowing through the port. This differs from a solid rocket motor, in which the regression rate is proportional to the chamber pressure of the motor.

 * $$\dot{r}=a_oG^n_o$$

 * where $$\dot{r}$$ is the regression rate, $a$$o$ is the regression rate coefficient (incorporating the grain length), $G$$o$ is the oxidizer mass flux rate, and $n$ is the regression rate exponent.

As the motor burns, the increase in diameter of the fuel port results in an increased fuel mass flow rate. This phenomenon makes the oxidizer to fuel ratio (O/F) shift during the burn. The increased fuel mass flow rate can be compensated for by also increasing the oxidizer mass flow rate. In addition to the O/F varying as a function of time, it also varies based on the position down the fuel grain. The closer the position is to the top of the fuel grain, the higher the O/F ratio. Since the O/F varies down the port, a point called the stoichiometric point may exist at some point down the grain.

Properties
Hybrid rocket motors exhibit some obvious as well as some subtle advantages over liquid-fuel rockets and solid-fuel rockets. A brief summary of some of these is given below:

Advantages compared with liquid rockets

 * Mechanically simpler – requires only a single liquid propellant resulting in less plumbing, fewer valves, and simpler operations.
 * Denser fuel – fuels in the solid phase generally have higher density than those in the liquid phase, reducing overall system volume.
 * Metal additives – reactive metals such as aluminium, magnesium, lithium or beryllium can be easily included in the fuel grain increasing specific impulse ($$I_{sp}$$), density, or both.
 * Combustion instabilities – Hybrid rockets do not typically exhibit high frequency combustion instabilities that plague liquid rockets due to the solid fuel grain breaking up acoustic waves that would otherwise reflect in an open liquid engine combustion chamber.
 * Propellant pressurization – One of the most difficult to design portions of a liquid rocket system are the turbopumps. Turbopump design is complex as it has to precisely and efficiently pump and keep separated two fluids of different properties in precise ratios at very high volumetric flow rates, often cryogenic temperatures, and highly volatile chemicals while combusting those same fluids in order to power itself. Hybrids have far less fluid to move and can often be pressurized by a blow-down system (which would be prohibitively heavy in a liquid rocket) or self-pressurized oxidizers (such as N2O).
 * Cooling – Liquid rockets often depend on one of the propellants, typically the fuel, to cool the combustion chamber and nozzle due to the very high heat fluxes and vulnerability of the metal walls to oxidation and stress cracking. Hybrid rockets have combustion chambers that are lined with the solid propellant which shields it from the product gases. Their nozzles are often graphite or coated in ablative materials similarly to solid rocket motors. The design, construction, and testing of liquid cooling flows is complex, making the system more prone to failure.

Advantages compared with solid rockets

 * Higher theoretical $$I_{sp}$$ – Possible due to limits of known solid oxidizers compared to often used liquid oxidizers.
 * Less explosion hazard – Propellant grain is more tolerant of processing errors such as cracks since the burn rate is dependent on oxidizer mass flux rate. Propellant grain cannot be ignited by stray electrical charge and is very insensitive to auto-igniting due to heat. Hybrid rocket motors can be transported to the launch site with the oxidizer and fuel stored separately, improving safety.
 * Fewer handling and storage issues – Ingredients in solid rockets are often incompatible chemically and thermally. Repeated changes in temperature can cause distortion of the grain. Antioxidants and coatings are used to keep the grain from breaking down or decomposing.
 * More controllable – Stop/restart and throttling are all easily incorporated into most designs. Solid rockets rarely can be shut down easily and almost never have throttling or restart capabilities.

Disadvantages of hybrid rockets
Hybrid rockets also exhibit some disadvantages when compared with liquid and solid rockets. These include:

 * Oxidizer-to-fuel ratio shift ("O/F shift") – with a constant oxidizer flow-rate, the ratio of fuel production rate to oxidizer flow rate will change as a grain regresses. This leads to off-peak operation from a chemical performance point of view. However, for a well-designed hybrid, O/F shift has a very small impact on performance because $$I_{sp}$$ is insensitive to O/F shift near the peak.
 * Poor regression characteristics often drive multi-port fuel grains. Multi-port fuel grains have poor volumetric efficiency and, often, structural deficiencies. High regression rate liquefying fuels developed in the late 1990s offer a potential solution to this problem.
 * Compared with liquid-based propulsion, re-fueling a partially or totally depleted hybrid rocket would present significant challenges, as the solid propellant cannot simply be pumped into a fuel tank. This may or may not be an issue, depending upon how the rocket is planned to be used.

In general, much less development work has been completed with hybrids than liquids or solids and it is likely that some of these disadvantages could be rectified through further investment in research and development.

One problem in designing large hybrid orbital rockets is that turbopumps become necessary to achieve high flow rates and pressurization of the oxidizer. This turbopump must be powered by something. In a traditional liquid-propellant rocket, the turbopump uses the same fuel and oxidizer as the rocket, since they are both liquid and can be fed to the pre-burner. But in a hybrid, the fuel is solid and cannot be fed to a turbopump's engine. Some hybrids use an oxidizer that can also be used as a monopropellant, such as hydrogen peroxide, and so a turbopump can run on it alone. However, hydrogen peroxide is significantly less efficient than liquid oxygen, which cannot be used alone to run a turbopump. Another fuel would be needed, requiring its own tank and decreasing rocket performance.

Common fuel choices
A reverse-hybrid rocket, which is not very common, is one where the engine uses a solid oxidizer and a liquid fuel. Some liquid fuel options are kerosene, hydrazine, and LH2. Common fuels for a typical hybrid rocket engine include polymers such as acrylics, polyethylene (PE), cross-linked rubber, such as HTPB, or liquefying fuels such as paraffin wax. Plexiglass was a common fuel, since the combustion could be visible through the transparent combustion chamber. Hydroxyl-terminated polybutadiene (HTPB) synthetic rubber is currently the most popular fuel for hybrid rocket engines, due to its energy, and due to how safe it is to handle. Tests have been performed in which HTPB was soaked in liquid oxygen, and it still did not become explosive. These fuels are generally not as dense as solid rocket motors, so they are often doped with aluminum to increase the density and therefore the rocket performance.

Cast
Hybrid rocket fuel grains can be manufactured via casting techniques, since they are typically a plastic or a rubber. Complex geometries, which are driven by the need for higher fuel mass flow rates, makes casting fuel grains for hybrid rockets expensive and time-consuming due in part to equipment costs. On a larger scale, cast grains must be supported by internal webbing, so that large chunks of fuel do not impact or even potentially block the nozzle. Grain defects are also an issue in larger grains. Traditional fuels that are cast are hydroxyl-terminated polybutadiene (HTPB) and paraffin waxes.

Additive manufacturing
Additive manufacturing is currently being used to create grain structures that were otherwise not possible to manufacture. Helical ports have been shown to increase fuel regression rates while also increasing volumetric efficiency. An example of material used for a hybrid rocket fuel is acrylonitrile butadiene styrene (ABS). The printed material is also typically enhanced with additives to improve rocket performance. Recent work at the University of Tennessee Knoxville has shown that, due to the increased surface area, the use of powdered fuels (i.e. graphite, coal, aluminum) encased in a 3D printed, ABS matrix can significantly increase the fuel burn rate and thrust level as compared to traditional polymer grains.

Common oxidizer choices
Common oxidizers include gaseous or liquid oxygen, nitrous oxide, and hydrogen peroxide. For a reverse hybrid, oxidizers such as frozen oxygen and ammonium perchlorate are used.

Proper oxidizer vaporization is important for the rocket to perform efficiently. Improper vaporization can lead to very large regression rate differences at the head end of the motor when compared to the aft end. One method is to use a hot gas generator to heat the oxidizer in a pre-combustion chamber. Another method is to use an oxidizer that can also be used as a monopropellant. A good example is hydrogen peroxide, which can be catalytically decomposed over a silver bed into hot oxygen and steam. A third method is to inject a propellant that is hypergolic with the oxidizer into the flow. Some of the oxidizer will decompose, heating up the rest of the oxidizer in the flow.

Hybrid safety
Generally, well designed and carefully constructed hybrids are very safe. The primary hazards associated with hybrids are:

 * Pressure vessel failures – Chamber insulation failure may allow hot combustion gases near the chamber walls leading to a "burn-through" in which the vessel ruptures.
 * Blow back – For oxidizers that decompose exothermically such as nitrous oxide or hydrogen peroxide, flame or hot gasses from the combustion chamber can propagate back through the injector, vaporising the oxidizer and mixing it with hot fuel rich gasses leading to a tank explosion. Blow-back requires gases to flow back through the injector due to insufficient pressure drop which can occur during periods of unstable combustion. Blow back is inherent to specific oxidizers and is not possible with oxidizers such as oxygen, or nitrogen tetroxide, unless fuel is present in the oxidizer tank.
 * Hard starts – An excess of oxidizer in the combustion chamber prior to ignition, particularly for monopropellants such as nitrous oxide, can result in a temporary over-pressure or "spike" at ignition.

Because the fuel in a hybrid does not contain an oxidizer, it will not combust explosively on its own. For this reason, hybrids are classified as having no TNT equivalent explosive power. In contrast, solid rockets often have TNT equivalencies similar in magnitude to the mass of the propellant grain. Liquid-fuel rockets typically have a TNT equivalence calculated based on the amount of fuel and oxidizer which could realistically intimately combine before igniting explosively; this is often taken to be 10–20% of the total propellant mass. For hybrids, even filling the combustion chamber with oxidizer prior to ignition will not generally create an explosion with the solid fuel, the explosive equivalence is often quoted as 0%.

Commercial companies
In 1998 SpaceDev acquired all of the intellectual property, designs, and test results generated by over 200 hybrid rocket motor firings by the American Rocket Company over its eight-year life. SpaceShipOne, the first private crewed spacecraft, was powered by SpaceDev's hybrid rocket motor burning HTPB with nitrous oxide. However, nitrous oxide was the prime substance responsible for the explosion that killed three in the development of the successor of SpaceShipOne at Scaled Composites in 2007. The Virgin Galactic SpaceShipTwo follow-on commercial suborbital spaceplane uses a scaled-up hybrid motor.

SpaceDev was developing the SpaceDev Streaker, an expendable small launch vehicle, and SpaceDev Dream Chaser, capable of both suborbital and orbital human space flight. Both Streaker and Dream Chaser use hybrid rocket motors that burn nitrous oxide and the synthetic HTPB rubber. SpaceDev was acquired by Sierra Nevada Corporation in 2009, becoming its Space Systems division, which continues to develop Dream Chaser for NASA's Commercial Crew Development contract. Sierra Nevada also developed RocketMotorTwo, the hybrid engine for SpaceShipTwo. On October 31, 2014, when SpaceShipTwo was lost, initial speculation had suggested that its hybrid engine had in fact exploded and killed one test pilot and seriously injured the other. However, investigation data now indicates an early deployment of the SpaceShip-Two feather system was the cause for aerodynamic breakup of the vehicle.

U.S. Rockets manufactured and deployed hybrids using self-pressurizing nitrous oxide (N2O) and hydroxyl-terminated polybutadiene (HTPB) as well as mixed High-test peroxide (HTP) and HTPB. The High-test peroxide (H2O2) 86% and (HTPB) and aluminum hybrids developed by U.S. Rockets produced a sea level delivered specific impulse (Isp) of 240, well above the typical 180 of N2O-HTPB hybrids. In addition to that, they were self-starting, restartable, had considerably lower combustion instability making them suitable for fragile or crewed missions such as Bloodhound SSC, SpaceShipTwo or SpaceShipThree. The company had successfully tested and deployed both pressure fed and pump fed versions of the latter HTP-HTPB style. Deliverables to date have ranged from 6 inch to 18 inch diameter, and developed units up to 54 inch diameter. The vendor claimed scalability to over 5 meters diameter with regression rates approaching solids, according to literature distributed at the November 2013 Defense Advanced Research Projects Agency (DARPA) meeting for XS-1. U.S. Rockets is no longer manufacturing large-scale rockets.

Gilmour Space Technologies began testing Hybrid rocket engines in 2015 with both N2O and HP with HDPE and HDPE+wax blends. For 2016 testing includes a 5000 lbf HP/PE engine. The company is planning to use hybrids for both sounding and orbital rockets.

Orbital Technologies Corporation (Orbitec) has been involved in some U.S. government-funded research on hybrid rockets including the "Vortex Hybrid" concept.

Environmental Aeroscience Corporation (eAc) was incorporated in 1994 to develop hybrid rocket propulsion systems. It was included in the design competition for the SpaceShipOne motor but lost the contract to SpaceDev. Environmental Aeroscience Corporation still supplied parts to SpaceDev for the oxidizer fill, vent, and dump system.

Rocket Lab formerly sold hybrid sounding rockets and related technology.

The Reaction Research Society (RRS), although known primarily for their work with liquid rocket propulsion, has a long history of research and development with hybrid rocket propulsion.

Copenhagen Suborbitals, a Danish rocket group, has designed and test-fired several hybrids using N2O at first and currently LOX. Their fuel is epoxy, paraffin wax, or polyurethane. The group eventually moved away from hybrids because of thrust instabilities, and now uses a motor similar to that of the V-2 rocket.

TiSPACE is a Taiwanese company which is developing a family of hybrid-propellant rockets.

bluShift Aerospace in Brunswick, Maine, won a NASA SBIR grant to develop a modular hybrid rocket engine for its proprietary bio-derived fuel in June 2019. Having completed the grant bluShift has launched its first sounding rocket using the technology.

Vaya Space based out of Cocoa, Florida, is expected to launch its hybrid fuel rocket Dauntless in 2023.

Reaction Dynamics based out Saint-Jean-sur-Richelieu, Quebec, began developing a hybrid rocket engine in 2017 capable of producing 21.6 kN of thrust. Their Aurora rocket will use nine engines on the first stage and one engine on the second stage and will be capable of delivering a payload of 50–150 kg to LEO. In May 2022, Reaction Dynamics announced they were partnering with Maritime Launch Services to launch the Aurora rocket from their launch site currently under construction in Canso, Nova Scotia, beginning with suborbital test flights in Summer, 2023 with a target of 2024 for the first orbital launch.

In 2017 DeltaV Uzay Teknolojileri A.Ş. was founded by Savunma Sanayi Teknolojileri A.Ş (SSTEK), a state company of Turkey, for hybrid-propellant-rocket research. The company CEO Arif Karabeyoglu is former Consulting Professor of Stanford University  in the area of rocket propulsion and combustion. According to company web site DeltaV achieved many firsts in hybrid-propellant-rocket technology including first paraffin/LOX dual fuel rocket launch, highest specific impulses for a hybrid-propellant-rocket, first sounding rocket to reach 100 km altittude, first orbital hybrid-propellant-rocket design, first orbital firing of hybrid-propellant-rocket.

Universities
Space Propulsion Group was founded in 1999 by Arif Karabeyoglu, Brian Cantwell, and others from Stanford University to develop high regression-rate liquefying hybrid rocket fuels. They have successfully fired motors as large as 12.5 in. diameter which produce 13,000 lbf using the technology and are currently developing a 24 in diameter, 25,000 lbf motor to be initially fired in 2010. Stanford University is the institution where liquid-layer combustion theory for hybrid rockets was developed. The SPaSE group at Stanford is currently working with NASA Ames Research Center developing the Peregrine sounding rocket which will be capable of 100 km altitude. Engineering challenges include various types of combustion instabilities. Although the proposed motor was test fired in 2013, the Peregrine program eventually switched to a standard solid rocket for its 2016 debut. The University of Tennessee Knoxville has carried out hybrid rocket research since 1999, working in collaboration with NASA Marshall Space Flight Center and private industry. This work has included the integration of a water-cooled calorimeter nozzle, one of the first 3D-printed, hot section components successfully used in a rocket motor. Other work at the university has focused on the use of helical oxidizer injection, bio-derived fuels and powdered fuels encased in a 3D-printed, ABS matrix, including the successful launch of a coal-fired hybrid at the 2019 Spaceport America Cup.

At the Delft University of Technology, the student team Delft Aerospace Rocket Engineering (DARE) is very active in the design and building of hybrid rockets. In October 2015, DARE broke the European student altitude record with the Stratos II+ sounding rocket. Stratos II+ was propelled by the DHX-200 hybrid rocket engine, using a nitrous oxide oxidizer and fuel blend of paraffin, sorbitol and aluminium powder. On July 26, 2018, DARE attempted to launch the Stratos III hybrid rocket. This rocket used the same fuel/oxidizer combination as its predecessor, but with an increased impulse of around 360 kNs. At the time of development, this was the most powerful hybrid rocket engine ever developed by a student team in terms of total impulse. Unfortunately, the Stratos III vehicle was lost 20 seconds into the flight.

Florida Institute of Technology has successfully tested and evaluated hybrid technologies with their Panther Project. The WARR student-team at the Technical University of Munich has been developing hybrid engines and rockets since the early 1970s. Using acids, oxygen, or nitrous oxide in combination with polyethylene, or HTPB. The development includes test stand engines as well as airborne versions, like the first German hybrid rocket Barbarella. They are currently working on a hybrid rocket with Liquid oxygen as its oxidizer, to break the European height record of amateur rockets. They are also working with Rocket Crafters and testing their hybrid rockets.

Boston University's student-run "Rocket Propulsion Group", which in the past has launched only solid motor rockets, is attempting to design and build a single-stage hybrid sounding rocket to launch into sub-orbital space by July 2015.

Brigham Young University (BYU), the University of Utah, and Utah State University launched a student-designed rocket called Unity IV in 1995 which burned the solid fuel hydroxyl-terminated polybutadiene (HTPB) with an oxidizer of gaseous oxygen, and in 2003 launched a larger version which burned HTPB with nitrous oxide.

The University of Brasilia's (UnB) Hybrid Rocket Team initiated their endeavors in 1999 within the Faculty of Technology, marking the pioneering institution in the Southern Hemisphere to engage with hybrid rockets. Over time, the team has achieved notable milestones, encompassing the creation of various sounding rockets and hybrid rocket engines. Presently, the team is known as the Chemical Propulsion Laboratory (CPL) and is situated at Campus UnB Gama. CPL has made significant strides in the advancement of critical hybrid engine technologies. This includes the development of a modular 1 kN hybrid rocket engine for the SARA platform, an innovative methane-oxygen gas-torch ignition system, an efficient oxidizer feed system, precision flow control valves, and thrust vector control mechanisms tailored for hybrid engines. Additionally, they've achieved a breakthrough with a 3D-printed, actively cooled hybrid rocket engine. Furthermore, the Laboratory is actively engaged in diverse areas of research and development, with current projects spanning the formulation of hybrid engine fuels using paraffin wax and N2O, numerical simulations, optimization techniques, and rocket design. CPL collaborates extensively with governmental agencies, private investors, and other educational institutions, including FAPDF, FAPESP, CNPq, and AEB. A notable collaborative effort includes the Capital Rocket Team (CRT), a group of students from UnB, who are currently partnering with CPL to develop hybrid sounding rockets. In a remarkable achievement, CRT clinched the top spot in the 2022 Latin American Space Challenge (LASC).

University of California, Los Angeles's student-run "Rocket Project at UCLA" launches hybrid propulsion rockets utilizing nitrous oxide as an oxidizer and HTPB as the fuel. They are currently in the development process of their fifth student-built hybrid rocket engine.

University of Toronto's student-run "University of Toronto Aerospace Team", designs and builds hybrid engine powered rockets. They are currently constructing a new engine testing facility at the University of Toronto Institute for Aerospace Studies, and are working towards breaking the Canadian amateur rocketry altitude record with their new rocket, Defiance MKIII, currently under rigorous testing. Defiance MK III's engine, QUASAR, is a Nitrous-Paraffin hybrid engine, capable of producing 7 kN of thrust for a period of 9 seconds.

In 2016, Pakistan's DHA Suffa University successfully developed Raheel-1, hybrid rocket engines in 1 kN class, using paraffin wax and liquid oxygen, thereby becoming the first university run rocket research program in the country. In India, Birla Institute of Technology, Mesra Space engineering and rocketry department has been working on Hybrid Projects with various fuels and oxidizers.

Pars Rocketry Group from Istanbul Technical University has designed and built the first hybrid rocket engine of Turkey, the rocket engine extensively tested in May 2015.

A United Kingdom-based team (laffin-gas) is using four N2O hybrid rockets in a drag-racing style car. Each rocket has an outer diameter of 150 mm and is 1.4 m long. They use a fuel grain of high-density wound paper soaked in cooking oil. The N2O supply is provided by Nitrogen-pressurised piston accumulators which provide a higher rate of delivery than N2O gas alone and also provide damping of any reverse shock.

In Italy one of the leading centers for research in hybrid propellants rockets is CISAS (Center of Studies and Activities for Space) "G. Colombo", University of Padua. The activities cover all stages of the development: from theoretical analysis of the combustion process to numerical simulation using CFD codes, and then by conducting ground tests of small scale and large-scale rockets (up to 20 kN, N2O-Paraffin wax based motors). One of these engines flew successfully in 2009. Since 2014, the research group is focused on the use of high test peroxide as oxidizer, in partnership with "Technology for Propulsion and Innovation", a university of Padua spin-off company.

In Taiwan, hybrid rocket system developments began in 2009 through R&D projects of NSPO with two university teams. Both teams employed nitrous oxide / HTPB propellant system with different improvement schemes. Several hybrid rockets have been successfully launched by NCKU and NCTU teams so far, reaching altitudes of 10–20 km. Their plans include attempting 100–200 km altitude launch to test nanosatellites, and developing orbital launch capabilities for nanosatellites in the long run. A sub-scale N2O/PE dual-vortical-flow (DVF) hybrid engine hot-fire test in 2014 has delivered an averaged Isp of 280 sec, which indicates that the system has reached around 97% combustion efficiency.

In (Germany) the University of Stuttgart's Student team HyEnd is the current world record holder for the highest-flying student-built hybrid rocket with their HEROS rockets.

In Bangladesh, Amateur Experimental Rocketry Dhaka supported by the American International University Bangladesh has also tested the country's first hybrid rocket engine, and are now working towards larger paraffin/nitrous oxide based prototypes.

The Aerospace Team of the TU Graz, Austria, is also developing a hybrid-propellant rocket.

The Polish Student team PWr in Space at Wrocław University of Science and Technology has developed three hybrid rockets: R2 "Setka", R3 "Dziewięćdziesiątka dziewiątka" and the most powerful of all - R4 "Lynx" with a successful test at their test stand

Many other universities, such as Embry-Riddle Aeronautical University, the University of Washington, Purdue University, the University of Michigan at Ann Arbor, the University of Arkansas at Little Rock, Hendrix College, the University of Illinois, Portland State University, University of KwaZulu-Natal, Texas A&M University, Aarhus University, Rice University, and AGH University of Science and Technology have hybrid motor test stands that allow for student research with hybrid rockets.

High power rocketry
There are a number of hybrid rocket motor systems available for amateur/hobbyist use in high-powered model rocketry. These include the popular HyperTek systems and a number of 'Urbanski-Colburn Valved' (U/C) systems such as RATTWorks, Contrail Rockets, and Propulsion Polymers. All of these systems use nitrous oxide as the oxidizer and a plastic fuel (such as Polyvinyl chloride (PVC), Polypropylene), or a polymer-based fuel such as HTPB. This reduces the cost per flight compared to solid rocket motors, although there is generally more ground support equipment required with hybrids.

In popular culture
An October 26, 2005 episode of the television show MythBusters entitled "Confederate Rocket" featured a hybrid rocket motor using liquid nitrous oxide and paraffin wax. The myth purported that during the American Civil War, the Confederate Army was able to construct a rocket of this type. The myth was revisited in a later episode entitled Salami Rocket, using hollowed out dry salami as the solid fuel.

In the February 18, 2007, episode of Top Gear, a Reliant Robin was used by Richard Hammond and James May in an attempt to modify a normal K-reg Robin into a reusable Space Shuttle. Steve Holland, a professional radio-controlled aircraft pilot, helped Hammond to work out how to land a Robin safely. The craft was built by senior members of the United Kingdom Rocketry Association (UKRA) and achieved a successful launch, flew for several seconds into the air and managed to successfully jettison the solid-fuel rocket boosters on time. This was the largest rocket launched by a non-government organisation in Europe. It used 6 × 40960 NS O motors by Contrail Rockets giving a maximum thrust of 8 tonnes. However, the car failed to separate from the large external fuel tank due to faulty explosive bolts between the Robin and the external tank, and the Robin subsequently crashed into the ground and seemed to have exploded soon after. This explosion was added for dramatic effect as neither Reliant Robins nor hybrid rocket motors explode in the way depicted.

title: Hybrid-propellant rocket


text:
Monopropellant rocket

A monopropellant rocket (or "monochemical rocket") is a rocket that uses a single chemical as its propellant. Monopropellant rockets are commonly used as small attitude and trajectory control rockets in satellites, rocket upper stages, manned spacecraft, and spaceplanes.

Chemical-reaction based monopropellant rockets
The simplest monopropellant rockets depend on the chemical decomposition of a storable propellant after passing it over a catalyst bed. The power for the thruster comes from the high pressure gas created during the decomposition reaction that allows a rocket nozzle to speed up the gas to create thrust.

The most commonly used monopropellant is hydrazine (N2H4, or H2N\sNH2), a compound unstable in the presence of a catalyst and which is also a strong reducing agent. The most common catalyst is granular alumina (aluminum oxide, Al2O3) coated with iridium. These coated granules are usually under the commercial labels Aerojet S-405 (previously made by Shell) or W.C. Heraeus H-KC 12 GA (previously made by Kali Chemie). There is no igniter with hydrazine. Aerojet S-405 is a spontaneous catalyst, that is, hydrazine decomposes on contact with the catalyst. The decomposition is highly exothermic and produces a 1000 C gas that is a mixture of nitrogen, hydrogen and ammonia. The main limiting factor of the monopropellant rocket is its life, which mainly depends on the life of the catalyst. The catalyst may be subject to catalytic poisoning and catalytic attrition which results in the catalyst failure. Another monopropellant is hydrogen peroxide, which, when purified to 90% or higher concentration, is self-decomposing at high temperatures or when a catalyst is present.

Most chemical-reaction monopropellant rocket systems consist of a fuel tank, usually a titanium or aluminium sphere, with an ethylene-propylene rubber container or a surface tension propellant management device filled with the fuel. The tank is then pressurized with helium or nitrogen, which pushes the fuel out to the motors. A pipe leads from the tank to a poppet valve, and then to the decomposition chamber of the rocket motor. Typically, a satellite will have not just one motor, but two to twelve, each with its own valve.

The attitude control rocket motors for satellites and space probes are often very small, 25 mm or so in diameter, and mounted in groups that point in four directions (within a plane).

The rocket is fired when the computer sends direct current through a small electromagnet that opens the poppet valve. The firing is often very brief, a few milliseconds, and — if operated in air — would sound like a pebble thrown against a metal trash can; if on for long, it would make a piercing hiss.

Chemical-reaction monopropellants are not as efficient as some other propulsion technologies. Engineers choose monopropellant systems when the need for simplicity and reliability outweigh the need for high delivered impulse. If the propulsion system must produce large amounts of thrust, or have a high specific impulse, as on the main motor of an interplanetary spacecraft, other technologies are used.

Solar-thermal based monopropellant thrusters
A concept to provide low Earth orbit (LEO) propellant depots that could be used as way-stations for other spacecraft to stop and refuel on the way to beyond-LEO missions has proposed that waste gaseous hydrogen—an inevitable byproduct of long-term liquid hydrogen storage in the radiative heat environment of space—would be usable as a monopropellant in a solar-thermal propulsion system. The waste hydrogen would be productively utilized for both orbital station-keeping and attitude control, as well as providing limited propellant and thrust to use for orbital maneuvers to better rendezvous with other spacecraft that would be inbound to receive fuel from the depot.

Solar-thermal monopropellant thrusters are also integral to the design of a next-generation cryogenic upper stage rocket proposed by U.S. company United Launch Alliance (ULA). The Advanced Common Evolved Stage (ACES) is intended as a lower-cost, more-capable and more-flexible upper stage that would supplement, and perhaps replace, the existing ULA Centaur and ULA Delta Cryogenic Second Stage (DCSS) upper stage vehicles. The ACES Integrated Vehicle Fluids option eliminates all hydrazine and helium from the space vehicle—normally used for attitude control and station keeping—and depends instead on solar-thermal monopropellant thrusters using waste hydrogen.

History
Soviet designers had begun experimenting with monopropellant rockets as early as 1933. They believed their monopropellant mixes of nitrogen tetroxide with gasoline, or toluene, and kerosene would lead to an overall simpler system; however, they ran into problems with violent explosions with pre-mixed fuel and oxidizer serving as a monopropellant that led the designers to abandon this approach.

Helmuth Walter was a German engineer an early pioneer of monopropellant rockets using hydrogen peroxide as fuel. Although his initial work was on submarine propulsion the same jets of oxygen produced by for combustion in gas turbines could be directed through a nozzle to generate thrust. The rocket Walter developed was used in the German ME-163 fighter aircraft in 1944, the first aircraft to break the 1000 km/h (635 mph).

After World War Two the British would continue to experiment with hydrogen peroxide monopropellants. They would develop the de Havilland Sprite a hydrogen peroxide rocket that could produce 5000lbf of thrust over 16 seconds. Not intended for space flight the rocket would provide hot and high takeoff capability to the de Havilland Comet 1 the first commercial jet airliner.

In the United States, when NASA began studying monopropellants at the Jet Propulsion Laboratory (JPL) the properties of the existing propellants demanded that the thrusters be impractically large. The addition of a catalyst and pre-heating propellant made them more efficient, but raised concerns over safety and handling of hazardous propellants like anhydrous hydrazine. However the simplicity of the thrusters designed around early monopropellants offered many simplicities and were first tested in 1959 on the Able-4 mission. This test allowed for the Ranger and Mariner missions to use a similar thruster for correction maneuvers and in the orbital insertion of Telstar, considered by the National Air and Space Museum to be the most significant communications satellite in the beginning of the space race. In 1964, NASA began use of the Lunar Landing Research Vehicle to train Apollo astronauts in piloting the Lunar Excursion Module (LEM) using an attitude control system consisting of 16 hydrogen peroxide monopropellant thrusters to steer the LEM to the lunar surface.

Upper stage vehicles began using monopropellant thrusters as a convenient control device in the early 1960s when General Dynamics proposed the Centaur upper stage to the United States Airforce of which versions are still in use in United Launch Alliance's Atlas and Vulcan rockets.

New developments
NASA is developing a new monopropellant propulsion system for small, cost-driven spacecraft with delta-v requirements in the range of 10–150 m/s. This system is based on a hydroxylammonium nitrate (HAN)/water/fuel monopropellant blend which is extremely dense, environmentally benign, and promises good performance and simplicity.

The EURENCO Bofors company produced LMP-103S as a 1-to-1 substitute for hydrazine by dissolving 65% ammonium dinitramide, NH4N(NO2)2, in 35% water solution of methanol and ammonia. LMP-103S has 6% higher specific impulse and 30% higher impulse density than hydrazine monopropellant. Additionally, hydrazine is highly toxic and carcinogenic, while LMP-103S is only moderately toxic. LMP-103S is UN Class 1.4S allowing for transport on commercial aircraft, and was demonstrated on the Prisma satellite in 2010. Special handling is not required. LMP-103S could replace hydrazine as the most commonly used monopropellant.

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

title: Monopropellant rocket


text:
Dual mode propulsion rocket

Dual mode propulsion systems combine the high efficiency of bipropellant rockets with the reliability and simplicity of monopropellant rockets.

Dual mode systems are either hydrazine/nitrogen tetroxide, or monomethylhydrazine/hydrogen peroxide (the former is much more common). Typically, this system works as follows: During the initial high-impulse orbit-raising maneuvers, the system operates in a bipropellant fashion, providing high thrust at higher specific impulse; when it arrives on orbit, it closes off either the fuel or oxidizer, and conducts the remainder of its mission in a simple, predictable monopropellant fashion.

title: Dual mode propulsion rocket


text:
Resistojet rocket

A resistojet is a method of spacecraft propulsion (electric propulsion) that provides thrust by heating a typically non-reactive fluid. Heating is usually achieved by sending electricity through a resistor consisting of a hot incandescent filament, with the expanded gas expelled through a conventional nozzle.

Resistojets have been flown in space since 1965 on board military Vela satellites. However, they only became used in commercial applications in 1980 with the launch of the first satellites in the INTELSAT-V program. Many GEO spacecraft, and all 95 Iridium, used Aerojet MR-501/MR-502 series resistojet engines. Nowadays resistojet propulsion is used for orbit insertion, attitude control, and deorbit of LEO satellites, and do well in situations where energy is much more plentiful than mass, and where propulsion efficiency needs to be reasonably high but low thrust is acceptable.

Resistojets have also been proposed as means of using biowaste as reaction mass, particularly in conjunction with hydrazine. Studies focus on the characteristics of steam and carbon dioxide as major constituents of a biowaste stream, and typically use cubic zirconia as a heating element.

Advantages
Many satellite missions necessitate an ability for minor alterations in trajectory even after the craft has been inserted into orbit. Most satellites use monopropellant rocket motors or cold gas thrusters for such orbital adjustments. Both methods, however, suffer from some limiting drawbacks: Hydrazine, the most commonly used monopropellant, is highly expensive and due to its volatile nature unsuitable for smaller satellites that are sent to space as secondary cargo. Cold gas thrusters, while utilizing relatively cheap, inert and therefore "safe" gasses like nitrogen, suffer from low specific impulse in comparison to monopropellant motors. Resistojets are designed to bridge the gap between these two methods of propulsion, offering the safety of an inert propellant coupled with specific impulse nearing that of hydrazine.

Disadvantages
The main disadvantage of a resistojet design in comparison to simpler cold gas thrusters is the need for a power supply, which takes up space and is therefore sometimes a prohibitive factor for Microsat missions. In addition, the increased technical complexity of a resistojet relative to simpler solutions results in a greater risk of technical failure.

Since they do not take advantage of chemical combustion, resistojets (and similar designs) have a lower thrust that is orders of magnitude lower than those of more conventional solid fuel and liquid-propellant rockets. As a result, they are unsuitable for orbital maneuvers that require high delta V over shorter periods.

Physical Principles
Resistojets can be considered an evolution of traditional cold gas thrusters, which are the simplest form of rocket engine available. Their fuel tank holds the propellant, which is then led into the nozzle where it decompresses, propelling the craft forward. In a resistojet, a resistor is used to heat the fluid before it enters the nozzle, making it expand more forcefully, resulting in higher specific impulse.

A resistor is an electrical component which converts electrical energy into heat. Therefore, the thrust of a resistojet engine can be regulated by simply altering the wattage flowing through the resistor.

Heating a fluid by 300 °C in such a way results in a 41% increase in specific impulse. If heated by 900 °C, specific impulse could be doubled in comparison to a cold gas thruster using the same propellant.

title: Resistojet rocket


text:
Arcjet rocket

An arcjet rocket or arcjet thruster is a form of electrically powered spacecraft propulsion, in which an electrical discharge (arc) is created in a flow of propellant (typically hydrazine or ammonia). This imparts additional energy to the propellant, so that one can extract more work out of each kilogram of propellant, at the expense of increased power consumption and (usually) higher cost. Also, the thrust levels available from typically used arcjet engines are very low compared with chemical engines.

When the energy is available, arcjets are well suited to keeping stations in orbit and can replace monopropellant rockets.

Aerojet MR-510 series arcjet engines are currently used on Lockheed Martin A2100 satellites using hydrazine as a propellant, providing over 585 s average specific impulse at 2 kW.

In Germany, researchers at the University of Stuttgart's Institute of Space Aviation Systems have been looking into these challenges for years and have developed various hydrogen-powered arcjet engines capable of power outputs from 1 to 100 kW. The heated hydrogen reaches exit speeds of 16 km/s. An arcjet-propelled test satellite by the name of Baden-Württemberg 1 (BW1) was scheduled to go to the Moon by 2010. No such launch has yet occurred. Baden-Württemberg 1 would have used polytetrafluoroethylene (PTFE) propellant.

title: Arcjet rocket


text:
Hall-effect thruster

In spacecraft propulsion, a Hall-effect thruster (HET) is a type of ion thruster in which the propellant is accelerated by an electric field. Hall-effect thrusters (based on the discovery by Edwin Hall) are sometimes referred to as Hall thrusters or Hall-current thrusters. Hall-effect thrusters use a magnetic field to limit the electrons' axial motion and then use them to ionize propellant, efficiently accelerate the ions to produce thrust, and neutralize the ions in the plume. The Hall-effect thruster is classed as a moderate specific impulse (1,600s) space propulsion technology and has benefited from considerable theoretical and experimental research since the 1960s.

Hall thrusters operate on a variety of propellants, the most common being xenon and krypton. Other propellants of interest include argon, bismuth, iodine, magnesium, zinc and adamantane.

Hall thrusters are able to accelerate their exhaust to speeds between 10 and 80 km/s (1,000–8,000 s specific impulse), with most models operating between 15 and 30 km/s. The thrust produced depends on the power level. Devices operating at 1.35 kW produce about 83 mN of thrust. High-power models have demonstrated up to 5.4 N in the laboratory. Power levels up to 100 kW have been demonstrated for xenon Hall thrusters.

, Hall-effect thrusters ranged in input power levels from 1.35 to 10 kilowatts and had exhaust velocities of 10–50 kilometers per second, with thrust of 40–600 millinewtons and efficiency in the range of 45–60 percent. The applications of Hall-effect thrusters include control of the orientation and position of orbiting satellites and use as a main propulsion engine for medium-size robotic space vehicles.

History
Hall thrusters were studied independently in the United States and the Soviet Union. They were first described publicly in the US in the early 1960s. However, the Hall thruster was first developed into an efficient propulsion device in the Soviet Union. In the US, scientists focused on developing gridded ion thrusters.

Soviet
Two types of Hall thrusters were developed in the Soviet Union:
 * thrusters with wide acceleration zone, SPT (СПД, стационарный плазменный двигатель; SPT, Stationary Plasma Thruster) at Design Bureau Fakel
 * thrusters with narrow acceleration zone, DAS (ДАС, двигатель с анодным слоем; TAL, Thruster with Anode Layer), at the Central Research Institute for Machine Building (TsNIIMASH).

The SPT design was largely the work of A. I. Morozov. The first SPT to operate in space, an SPT-50 aboard a Soviet Meteor spacecraft, was launched December 1971. They were mainly used for satellite stabilization in north–south and in east–west directions. Since then until the late 1990s 118 SPT engines completed their mission and some 50 continued to be operated. Thrust of the first generation of SPT engines, SPT-50 and SPT-60 was 20 and 30 mN respectively. In 1982, the SPT-70 and SPT-100 were introduced, their thrusts being 40 and 83 mN, respectively. In the post-Soviet Russia high-power (a few kilowatts) SPT-140, SPT-160, SPT-200, T-160, and low-power (less than 500 W) SPT-35 were introduced.

Soviet and Russian TAL-type thrusters include the D-38, D-55, D-80, and D-100.

Over 200 Hall thrusters have been flown on Soviet/Russian satellites since the 1980s. No failures have ever occurred on orbit.

Non-Soviet
Soviet-built thrusters were introduced to the West in 1992 after a team of electric propulsion specialists from NASA's Jet Propulsion Laboratory, Glenn Research Center, and the Air Force Research Laboratory, under the support of the Ballistic Missile Defense Organization, visited Russian laboratories and experimentally evaluated the SPT-100 (i.e., a 100 mm diameter SPT thruster). Hall thrusters continue to be used on Russian spacecraft and have also flown on European and American spacecraft. Space Systems/Loral, an American commercial satellite manufacturer, now flies Fakel SPT-100's on their GEO communications spacecraft.

Since in the early 1990s, Hall thrusters have been the subject of a large number of research efforts throughout the United States, India, France, Italy, Japan, and Russia (with many smaller efforts scattered in various countries across the globe). Hall thruster research in the US is conducted at several government laboratories, universities and private companies. Government and government funded centers include NASA's Jet Propulsion Laboratory, NASA's Glenn Research Center, the Air Force Research Laboratory (Edwards AFB, California), and The Aerospace Corporation. Universities include the US Air Force Institute of Technology, University of Michigan, Stanford University, The Massachusetts Institute of Technology, Princeton University, Michigan Technological University, and Georgia Tech. A considerable amount of development is being conducted in industry, such as IHI Corporation in Japan, Aerojet and Busek in the US, SNECMA in France, LAJP in Ukraine, SITAEL in Italy, and Satrec Initiative in South Korea.

The first use of Hall thrusters on lunar orbit was the European Space Agency (ESA) lunar mission SMART-1 in 2003.

Hall thrusters were first demonstrated on a western satellite on the Naval Research Laboratory (NRL) STEX spacecraft, which flew the Russian D-55. The first American Hall thruster to fly in space was the Busek BHT-200 on TacSat-2 technology demonstration spacecraft. The first flight of an American Hall thruster on an operational mission, was the Aerojet BPT-4000, which launched August 2010 on the military Advanced Extremely High Frequency GEO communications satellite. At 4.5 kW, the BPT-4000 is also the highest power Hall thruster ever flown in space. Besides the usual stationkeeping tasks, the BPT-4000 is also providing orbit-raising capability to the spacecraft. The X-37B has been used as a testbed for the Hall thruster for the AEHF satellite series. Several countries worldwide continue efforts to qualify Hall thruster technology for commercial uses. The SpaceX Starlink constellation, the largest satellite constellation in the world, uses Hall-effect thrusters. Starlink initially used krypton gas, but with its V2 satellites swapped to argon due to its cheaper price and widespread availability.

The first deployment of Hall thrusters beyond Earth's sphere of influence was the Psyche spacecraft, launched in 2023 towards the Asteroid Belt to explore 16 Psyche.

Indian
Research in India is carried out by both public and private research institutes and companies.

In 2010, ISRO used Hall effect ion propulsion thrusters in GSAT-4 carried by GSLV Mk2 D3. It had four Xenon powered thrusters for North-South station keeping. Two of them were Russian and the other two were Indian. The Indian thrusters were rated at 13mN. However, GSLV D3 didn't make it to orbit.

In 2013, ISRO funded development of another class of electric thruster called Magnetoplasmadynamic Electric Propulsion Thruster. The project subsequently developed a technology demonstrator prototype Magneto Plasma Dynamic Thruster (MPD) using Argon propellant with a specific impulse of 2500s at a thrust of 25 mN.

The following year in 2014, ISRO was pursuing development of 75mN & 250 mN SPT thrusters to be used in its future high power communication satellites. The 75 mN thrusters were put to use for the GSAT-9 communication satellite.

By 2021 development of a 300 mN thruster was complete. Alongside it, RF powered 10kW plasma engine & Krypton based low power electric propulsion were being pursued.

With private firms entering the space domain, Bellatrix Aerospace became the first commercial firm to bring out commercial hall effect thrusters. The current model of the thruster uses Xenon as fuel. Tests were carried out at the spacecraft propulsion research laboratory in the Indian Institute of Science, Bengaluru. Heaterless Cathode Technology is the key innovation that sets them apart from the competition by increasing life and redundancy of the system. The spacetech startup had earlier developed world’s first commercial Microwave Plasma Thruster, which used water as fuel instantaneously heating it by microwave-induced plasma, and for which the company had bagged an order from ISRO.

Principle of operation
The essential working principle of the Hall thruster is that it uses an electrostatic potential to accelerate ions up to high speeds. In a Hall thruster, the attractive negative charge is provided by an electron plasma at the open end of the thruster instead of a grid. A radial magnetic field of about 100 – is used to confine the electrons, where the combination of the radial magnetic field and axial electric field cause the electrons to drift in azimuth thus forming the Hall current from which the device gets its name.

A schematic of a Hall thruster is shown in the adjacent image. An electric potential of between 150 and 800 volts is applied between the anode and cathode.

The central spike forms one pole of an electromagnet and is surrounded by an annular space, and around that is the other pole of the electromagnet, with a radial magnetic field in between.

The propellant, such as xenon gas, is fed through the anode, which has numerous small holes in it to act as a gas distributor. As the neutral xenon atoms diffuse into the channel of the thruster, they are ionized by collisions with circulating high-energy electrons (typically 10–40 eV, or about 10% of the discharge voltage). Most of the xenon atoms are ionized to a net charge of +1, but a noticeable fraction (c. 20%) have +2 net charge.

The xenon ions are then accelerated by the electric field between the anode and the cathode. For discharge voltages of 300 V, the ions reach speeds of around 15 km/s for a specific impulse of 1,500 s (15 kN·s/kg). Upon exiting, however, the ions pull an equal number of electrons with them, creating a plasma plume with no net charge.

The radial magnetic field is designed to be strong enough to substantially deflect the low-mass electrons, but not the high-mass ions, which have a much larger gyroradius and are hardly impeded. The majority of electrons are thus stuck orbiting in the region of high radial magnetic field near the thruster exit plane, trapped in E×B (axial electric field and radial magnetic field). This orbital rotation of the electrons is a circulating Hall current, and it is from this that the Hall thruster gets its name. Collisions with other particles and walls, as well as plasma instabilities, allow some of the electrons to be freed from the magnetic field, and they drift towards the anode.

About 20–30% of the discharge current is an electron current, which does not produce thrust, thus limiting the energetic efficiency of the thruster; the other 70–80% of the current is in the ions. Because the majority of electrons are trapped in the Hall current, they have a long residence time inside the thruster and are able to ionize almost all of the xenon propellant, allowing mass use of 90–99%. The mass use efficiency of the thruster is thus around 90%, while the discharge current efficiency is around 70%, for a combined thruster efficiency of around 63% (= 90% × 70%). Modern Hall thrusters have achieved efficiencies as high as 75% through advanced designs.

Compared to chemical rockets, the thrust is very small, on the order of 83 mN for a typical thruster operating at 300 V and 1.5 kW. For comparison, the weight of a coin like the U.S. quarter or a 20-cent euro coin is approximately 60 mN. As with all forms of electrically powered spacecraft propulsion, thrust is limited by available power, efficiency, and specific impulse.

However, Hall thrusters operate at the high specific impulses that are typical for electric propulsion. One particular advantage of Hall thrusters, as compared to a gridded ion thruster, is that the generation and acceleration of the ions takes place in a quasi-neutral plasma, so there is no Child-Langmuir charge (space charge) saturated current limitation on the thrust density. This allows much smaller thrusters compared to gridded ion thrusters.

Another advantage is that these thrusters can use a wider variety of propellants supplied to the anode, even oxygen, although something easily ionized is needed at the cathode.

Xenon
Xenon has been the typical choice of propellant for many electric propulsion systems, including Hall thrusters. Xenon propellant is used because of its high atomic weight and low ionization potential. Xenon is relatively easy to store, and as a gas at spacecraft operating temperatures does not need to be vaporized before usage, unlike metallic propellants such as bismuth. Xenon's high atomic weight means that the ratio of energy expended for ionization per mass unit is low, leading to a more efficient thruster.

Krypton
Krypton is another choice of propellant for Hall thrusters. Xenon has an ionization potential of 12.1298 eV, while krypton has an ionization potential of 13.996 eV. This means that thrusters utilizing krypton need to expend a slightly higher energy per mole to ionize, which reduces efficiency. Additionally, krypton is a lighter ion, so the unit mass per ionization energy is further reduced compared to xenon. However, xenon can be more than ten times as expensive as krypton per kilogram, making krypton a more economical choice for building out satellite constellations like that of SpaceX's Starlink V1, whose original Hall thrusters were fueled with krypton.

Argon
SpaceX developed a new thruster that used argon as propellant for their Starlink V2 mini. The new thruster had 2.4 times the thrust and 1.5 times the specific impulse as SpaceX's previous thruster that used krypton. Argon is approximately 100 times less expensive than Krypton and 1000 times less expensive than Xenon.

Variants
As well as the Soviet SPT and TAL types mentioned above, there are :

Cylindrical Hall thrusters
Although conventional (annular) Hall thrusters are efficient in the kilowatt power regime, they become inefficient when scaled to small sizes. This is due to the difficulties associated with holding the performance scaling parameters constant while decreasing the channel size and increasing the applied magnetic field strength. This led to the design of the cylindrical Hall thruster. The cylindrical Hall thruster can be more readily scaled to smaller sizes due to its nonconventional discharge-chamber geometry and associated magnetic field profile. The cylindrical Hall thruster more readily lends itself to miniaturization and low-power operation than a conventional (annular) Hall thruster. The primary reason for cylindrical Hall thrusters is that it is difficult to achieve a regular Hall thruster that operates over a broad envelope from c.1 kW down to c. 100 W while maintaining an efficiency of 45–55%.

External discharge Hall thruster
Sputtering erosion of discharge channel walls and pole pieces that protect the magnetic circuit causes failure of thruster operation. Therefore, annular and cylindrical Hall thrusters have limited lifetime. Although magnetic shielding has been shown to dramatically reduce discharge channel wall erosion, pole piece erosion is still a concern. As an alternative, an unconventional Hall thruster design called external discharge Hall thruster or external discharge plasma thruster (XPT) has been introduced. The external discharge Hall thruster does not possess any discharge channel walls or pole pieces. Plasma discharge is produced and sustained completely in the open space outside the thruster structure, and thus erosion-free operation is achieved.

Applications

Hall thrusters have been flying in space since December 1971, when the Soviet Union launched an SPT-50 on a Meteor satellite. Over 240 thrusters have flown in space since that time, with a 100% success rate. Hall thrusters are now routinely flown on commercial LEO and GEO communications satellites, where they are used for orbital insertion and stationkeeping.

The first Hall thruster to fly on a western satellite was a Russian D-55 built by TsNIIMASH, on the NRO's STEX spacecraft, launched on October 3, 1998.

The solar electric propulsion system of the European Space Agency's SMART-1 spacecraft used a Snecma PPS-1350-G Hall thruster. SMART-1 was a technology demonstration mission that orbited the Moon. This use of the PPS-1350-G, starting on September 28, 2003, was the first use of a Hall thruster outside geosynchronous Earth orbit (GEO). Like most Hall thruster propulsion systems used in commercial applications, the Hall thruster on SMART-1 could be throttled over a range of power, specific impulse, and thrust. It has a discharge power range of 0.46–1.19 kW, a specific impulse of 1,100–1,600 s and thrust of 30–70 mN.

Early small satellites of the SpaceX Starlink constellation used krypton-fueled Hall thrusters for position-keeping and deorbiting, while later Starlink satellites used argon-fueled Hall thrusters.

Tiangong space station is fitted with Hall-effect thrusters. Tianhe core module is propelled by both chemical thrusters and four ion thrusters, which are used to adjust and maintain the station's orbit. 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.

The Jet Propulsion Laboratory (JPL) granted exclusive commercial licensing to Apollo Fusion, led by Mike Cassidy, for its Magnetically Shielded Miniature (MaSMi) Hall thruster technology. In January 2021, Apollo Fusion announced they had secured a contract with York Space Systems for an order of its latest iteration named the "Apollo Constellation Engine".

The NASA mission to the asteroid Psyche utilizes xenon gas Hall thrusters. The electricity comes from the craft's 75 square meter solar panels.

NASA's first Hall thrusters on a human-rated mission will be a combination of 6 kW Hall thrusters provided by Busek and NASA Advanced Electric Propulsion System (AEPS) Hall thrusters. They will serve as the primary propulsion on Maxar's Power and Propulsion Element (PPE) for the Lunar Gateway under NASA's Artemis program. The high specific impulse of Hall thrusters will allow for efficient orbit raising and station keep for the Lunar Gateway's polar near-rectilinear halo orbit.

In development
The highest power Hall-effect thruster in development (as of 2021) is the University of Michigan's 100 kW X3 Nested Channel Hall Thruster. The thruster is approximately 80 cm in diameter and weighs 230 kg, and has demonstrated a thrust of 5.4 N.

Other high power thrusters include NASA's 40 kW Advanced Electric Propulsion System (AEPS), meant to propel large-scale science missions and cargo transportation in deep space.

title: Hall-effect thruster