Editing Rocket engine (section)

==Types of rocket engines==

===Physically powered===
{| class="wikitable"
|-
! Type
! Description
! Advantages
! Disadvantages
|-
! [[Water rocket]]
| Partially filled pressurised carbonated drinks container with tail and nose weighting
| Very simple to build
| Altitude typically limited to a few hundred feet or so (world record is 830 meters, or 2,723 feet)
|-
! [[Cold gas thruster]]
| A non-combusting form, used for [[vernier thruster]]s
| Non-contaminating exhaust
| Extremely low performance
|}

===Chemically powered===
{{See also|Liquid rocket propellant}}

{| class="wikitable"
|-
! Type
! Description
! Advantages
! Disadvantages
|-
! [[Solid-propellant rocket]]
| Ignitable, self-sustaining solid fuel/oxidiser mixture ("grain") with central hole and nozzle
| Simple, often no [[moving parts]], reasonably good mass fraction, reasonable [[Specific Impulse|''I''<sub>sp</sub>]]. A thrust schedule can be designed into the grain.
| Throttling, burn termination, and reignition require special designs. Handling issues from ignitable mixture. Lower performance than liquid rockets. If grain cracks it can block nozzle with disastrous results. Grain cracks burn and widen during burn. Refueling harder than simply filling tanks. Cannot be turned off after ignition; will fire until all solid fuel is depleted.
|-
! [[Hybrid-propellant rocket]]
| Separate oxidiser/fuel; typically the oxidiser is liquid and kept in a tank and the fuel is solid.
| Quite simple, solid fuel is essentially inert without oxidiser, safer; cracks do not escalate, throttleable and easy to switch off.
| Some oxidisers are monopropellants, can explode in own right; mechanical failure of solid propellant can block nozzle (very rare with rubberised propellant), central hole widens over burn and negatively affects mixture ratio.
|-
! [[Monopropellant rocket]]
| Propellant (such as hydrazine, hydrogen peroxide or nitrous oxide) flows over a catalyst and exothermically decomposes; hot gases are emitted through nozzle.
| Simple in concept, throttleable, low temperatures in combustion chamber
| Catalysts can be easily contaminated, monopropellants can detonate if contaminated or provoked, [[Specific Impulse|''I''<sub>sp</sub>]] is perhaps 1/3 of best liquids
|-
! [[Liquid bipropellant rocket engine|Bipropellant rocket]]
| Two fluid (typically liquid) propellants are introduced through injectors into combustion chamber and burnt.
| Up to ≈99% efficient combustion with excellent mixture control, throttleable, can be used with turbopumps which permits incredibly lightweight tanks, can be safe with extreme care
| Pumps needed for high performance are expensive to design, huge thermal fluxes across combustion chamber wall can impact reuse, failure modes include major explosions, a lot of plumbing is needed.
|-
! [[Methane-oxygen gaseous thruster|Gas-gas rocket]]
| A bipropellant thruster using gas propellant for both the oxidiser and fuel
| Higher-performance than cold gas thrusters
| Lower performance than liquid-based engines
|-
! [[Dual mode propulsion rocket]]
| Rocket takes off as a bipropellant rocket, then turns to using just one propellant as a monopropellant.
| Simplicity and ease of control
| Lower performance than bipropellants
|-
! [[Tripropellant rocket]] 
| Three different propellants (usually hydrogen, hydrocarbon, and liquid oxygen) are introduced into a combustion chamber in variable mixture ratios, or multiple engines are used with fixed propellant mixture ratios and throttled or shut down
| Reduces take-off weight, since hydrogen is lighter; combines good thrust to weight with high average [[specific impulse|''I''<sub>sp</sub>]], improves payload for launching from Earth by a sizeable percentage
| Similar issues to bipropellant, but with more plumbing, more research and development
|-
! [[Air-augmented rocket]]
| Essentially a ramjet where intake air is compressed and burnt with the exhaust from a rocket
| Mach 0 to Mach 4.5+ (can also run exoatmospheric), good efficiency at Mach 2 to 4
| Similar efficiency to rockets at low speed or exoatmospheric, inlet difficulties, a relatively undeveloped and unexplored type, cooling difficulties, very noisy, thrust/weight ratio is similar to ramjets.
|-
! [[Turborocket]]
| A combined cycle turbojet/rocket where an additional oxidiser such as oxygen is added to the airstream to increase maximum altitude
| Very close to existing designs, operates in very high altitude, wide range of altitude and airspeed
| Atmospheric airspeed limited to same range as turbojet engine, carrying oxidiser like [[LOX]] can be dangerous. Much heavier than simple rockets.
|-
! [[Precooled jet engine]] / [[liquid air cycle engine|LACE]] (combined cycle with rocket)
| Intake air is chilled to very low temperatures at inlet before passing through a ramjet or turbojet engine. Can be combined with a rocket engine for orbital insertion.
| Easily tested on ground. High thrust/weight ratios are possible (≈14) together with good fuel efficiency over a wide range of airspeeds, mach 0–5.5+; this combination of efficiencies may permit launching to orbit, single stage, or very rapid intercontinental travel.
| Exists only at the lab prototyping stage. Examples include [[RB545]], [[Reaction Engines SABRE|SABRE]], [[ATREX]]
|}

===Electrically powered===
{{Main|Electrically powered spacecraft propulsion}}
{| class="wikitable"
|-
! Type
! Description
! Advantages
! Disadvantages
|-
! [[Resistojet rocket]] (electric heating)
| Energy is imparted to a usually inert fluid serving as reaction mass via [[Joule heating]] of a heating element. May also be used to impart extra energy to a monopropellant.
| Efficient where electrical power is at a lower premium than mass. Higher [[specific impulse|''I''<sub>sp</sub>]] than monopropellant alone, about 40% higher.
| Requires a lot of power, hence typically yields low thrust.
|-
! [[Arcjet rocket]] (chemical burning aided by electrical discharge)
| Identical to resistojet except the heating element is replaced with an electrical arc, eliminating the physical requirements of the heating element.
| 1,600 seconds [[specific impulse|''I''<sub>sp</sub>]]
| Very low thrust and high power, performance is similar to [[ion drive]].
|-
![[Variable specific impulse magnetoplasma rocket]]
| Microwave heated plasma with magnetic throat/nozzle
| Variable ''I''<sub>sp</sub> from 1,000 seconds to 10,000 seconds
| Similar thrust/weight ratio with ion drives (worse), thermal issues, as with ion drives very high power requirements for significant thrust, really needs advanced nuclear reactors, never flown, requires low temperatures for superconductors to work
|-
! [[Pulsed plasma thruster]] (electric arc heating; emits plasma)
| Plasma is used to erode a solid propellant
| High ''I''<sub>sp</sub>, can be pulsed on and off for attitude control
| Low energetic efficiency
|-
! [[Ion thruster|Ion propulsion system]]
| High voltages at ground and plus sides
| Powered by battery
| Low thrust, needs high voltage
|}

===Thermal===

====Preheated====
{| class="wikitable"
|-
! Type
! Description
! Advantages
! Disadvantages
|-
! [[Hot water rocket]]
| Hot water is stored in a tank at high temperature&thinsp;/&thinsp;pressure and turns to steam in nozzle
| Simple, fairly safe
| Low overall performance due to heavy tank; [[specific impulse|''I''<sub>sp</sub>]] under 200 seconds
|}

====Solar thermal====

The [[solar thermal rocket]] would make use of solar power to directly heat [[reaction mass]], and therefore does not require an electrical generator as most other forms of solar-powered propulsion do. A solar thermal rocket only has to carry the means of capturing solar energy, such as [[Concentrating solar power|concentrator]]s and [[mirror]]s. The heated propellant is fed through a conventional rocket nozzle to produce thrust. The engine thrust is directly related to the surface area of the solar collector and to the local intensity of the solar radiation and inversely proportional to the ''I''<sub>sp</sub>.

{| class="wikitable"
|-
! Type
! Description
! Advantages
! Disadvantages
|-
! [[Solar thermal rocket]] 
| Propellant is heated by solar collector
| Simple design. Using hydrogen propellant, 900 seconds of [[specific impulse|''I''<sub>sp</sub>]] is comparable to nuclear thermal rocket, without the problems and complexity of controlling a fission reaction.{{citation needed|date=January 2011}} Ability to [[Solar thermal rocket#Proposed solar-thermal space systems|productively use]] waste gaseous [[hydrogen]]—an inevitable byproduct of long-term [[liquid hydrogen]] storage in the [[Radiative heat transfer|radiative heat]] environment of space—for both [[orbital stationkeeping]] and [[Spacecraft attitude control|attitude control]].<ref name=aiaa20100902>{{cite web|last=Zegler|first=Frank |title=Evolving to a Depot-Based Space Transportation Architecture |url=http://www.ulalaunch.com/site/docs/publications/DepotBasedTransportationArchitecture2010.pdf |archive-url=https://web.archive.org/web/20110717150155/http://www.ulalaunch.com/site/docs/publications/DepotBasedTransportationArchitecture2010.pdf |url-status=dead |archive-date=2011-07-17 |work=AIAA SPACE 2010 Conference & Exposition |publisher=AIAA |access-date=2011-01-25 |author2=Bernard Kutter  |date=2010-09-02 }} See page 3.</ref>
| Only useful in space, as thrust is fairly low, but hydrogen has not been traditionally thought to be easily stored in space,<ref name=aiaa20100902/> otherwise moderate/low [[Specific impulse|''I''<sub>sp</sub>]] if higher–molecular-mass propellants are used.
|}

====Beamed thermal====

{| class="wikitable"
|-
! Type
! Description
! Advantages
! Disadvantages
|-
! [[Laser propulsion|Light-beam-powered rocket]]
| Propellant is heated by light beam (often laser) aimed at vehicle from a distance, either directly or indirectly via heat exchanger
| Simple in principle, in principle very high exhaust speeds can be achieved
| ≈1 MW of power per kg of payload is needed to achieve orbit, relatively high accelerations, lasers are blocked by clouds, fog, reflected laser light may be dangerous, pretty much needs hydrogen monopropellant for good performance which needs heavy tankage, some designs are limited to ≈600 seconds due to reemission of light since propellant/heat exchanger gets white hot
|-
! [[Beam-powered propulsion|Microwave-beam-powered rocket]]
| Propellant is heated by microwave beam aimed at vehicle from a distance
| [[Specific impulse|''I''<sub>sp</sub>]] is comparable to Nuclear Thermal rocket combined with T/W comparable to conventional rocket.  While LH<sub>2</sub> propellant offers the highest I<sub>sp</sub> and rocket payload fraction, ammonia or methane are economically superior for earth-to-orbit rockets due to their particular combination of high density and I<sub>sp</sub>. [[Single-stage-to-orbit|SSTO]] operation is possible with these propellants even for small rockets, so there are no location, trajectory and shock constraints added by the rocket staging process.  Microwaves are 10-100× cheaper in $/watt than lasers and have all-weather operation at frequencies below 10&nbsp;GHz.
| 0.3–3{{nbsp}}MW of power per kg of payload is needed to achieve orbit depending on the propellant,<ref>{{cite web|url=http://parkinresearch.com/microwave-thermal-rockets/|title=Microwave Thermal Rockets|last=Parkin|first=Kevin|access-date=8 December 2016}}</ref> and this incurs infrastructure cost for the beam director plus related R&D costs. Concepts operating in the millimeter-wave region have to contend with weather availability and high altitude beam director sites as well as effective transmitter diameters measuring 30–300 meters to propel a vehicle to LEO.  Concepts operating in X-band or below must have effective transmitter diameters measured in kilometers to achieve a fine enough beam to follow a vehicle to LEO. The transmitters are too large to fit on mobile platforms and so microwave-powered rockets are constrained to launch near fixed beam director sites. 
|}

====Nuclear thermal====

{| class="wikitable"
|-
! Type
! Description
! Advantages
! Disadvantages
|-
! [[Radioisotope rocket|Radioisotope rocket/"Poodle thruster"]] (radioactive decay energy)
| Heat from radioactive decay is used to heat hydrogen
| About 700–800 seconds, almost no moving parts
| Low thrust/weight ratio.
|-
! [[Nuclear thermal rocket]] (nuclear fission energy)
| Propellant (typically, hydrogen) is passed through a nuclear reactor to heat to high temperature
| [[Specific impulse|''I''<sub>sp</sub>]] can be high, perhaps 900 seconds or more, above unity thrust/weight ratio with some designs
| Maximum temperature is limited by materials technology, some radioactive particles can be present in exhaust in some designs, nuclear reactor shielding is heavy, unlikely to be permitted from surface of the Earth, thrust/weight ratio is not high.
|}

===Nuclear===
[[Nuclear propulsion]] includes a wide variety of [[spacecraft propulsion|propulsion]] methods that use some form of [[nuclear reaction]] as their primary power source. Various types of nuclear propulsion have been proposed, and some of them tested, for spacecraft applications:

{| class="wikitable"
|-
! Type
! Description
! Advantages
! Disadvantages
|-
! [[Gas core reactor rocket]] (nuclear fission energy)
| Nuclear reaction using a gaseous state fission reactor in intimate contact with propellant
| Very hot propellant, not limited by keeping reactor solid, [[Specific Impulse|''I''<sub>sp</sub>]] between 1,500 and 3,000 seconds but with very high thrust
| Difficulties in heating propellant without losing fissionables in exhaust, massive thermal issues particularly for nozzle/throat region, exhaust almost inherently highly radioactive. Nuclear lightbulb variants can contain fissionables, but cut [[Specific Impulse|''I''<sub>sp</sub>]] in half.
|-
! [[Fission-fragment rocket]] (nuclear fission energy)
| Fission products are directly exhausted to give thrust.
|
| Theoretical only at this point.
|-
! [[Fission sail]] (nuclear fission energy)
| A sail material is coated with fissionable material on one side.
| No moving parts, works in deep space
| Theoretical only at this point.
|-
! [[Nuclear salt-water rocket]] (nuclear fission energy)
| Nuclear salts are held in solution, caused to react at nozzle
| Very high [[Specific Impulse|''I''<sub>sp</sub>]], very high thrust
| Thermal issues in nozzle, propellant could be unstable, highly radioactive exhaust. Theoretical only at this point.
|-
! [[Nuclear pulse propulsion]] (exploding fission/fusion bombs)
| Shaped nuclear bombs are detonated behind vehicle and blast is caught by a 'pusher plate'
| Very high [[Specific Impulse|''I''<sub>sp</sub>]], very high thrust/weight ratio, no show stoppers are known for this technology.
| Never been tested, pusher plate may [[spall|throw off fragments]] due to shock, minimum size for nuclear bombs is still pretty big, expensive at small scales, nuclear treaty issues, fallout when used below Earth's magnetosphere.
|-
! [[Antimatter catalyzed nuclear pulse propulsion]] (fission and/or fusion energy)
| Nuclear pulse propulsion with antimatter assist for smaller bombs
| Smaller sized vehicle might be possible
| Containment of antimatter, production of antimatter in macroscopic quantities is not currently feasible. Theoretical only at this point.
|-
! [[Fusion rocket]] (nuclear fusion energy)
| Fusion is used to heat propellant
| Very high exhaust velocity
| Largely beyond current state of the art.
|-
! [[Antimatter rocket]] (annihilation energy)
| Antimatter annihilation heats propellant
| Extremely energetic, very high theoretical exhaust velocity
| Problems with antimatter production and handling; energy losses in [[neutrino]]s, [[gamma ray]]s, [[muon]]s; thermal issues. Theoretical only at this point.
|}