Rocket engine
From Wikipedia, the free encyclopedia
A rocket engine is a reaction engine that can be used for spacecraft propulsion as well as terrestrial uses, such as missiles. Rocket engines take their reaction mass from within the vehicle and form it into a high speed jet, obtaining thrust in accordance with Newton's third law. Most rocket engines are internal combustion engines, although non combusting forms exist.
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[edit] Principle of operation
Classic rocket engines produce a high temperature, hypersonic gaseous exhaust. This is most often achieved by the combustion of solid, liquid or gaseous propellant, containing oxidiser and a fuel, within a combustion chamber at high pressure. Alternatively, a chemically inert reaction mass can be heated to high temperature using a high energy power source.
The hot gas produced is then allowed to escape through a narrow hole (the 'throat'), into a high-expansion ratio nozzle. The effect of the nozzle is to dramatically accelerate the mass, converting most of the thermal energy into kinetic energy. The large bell or cone shaped expansion nozzle gives a rocket engine its characteristic shape. Exhaust speeds as high as 10 times the speed of sound at sea level are not uncommon.
Part of the rocket engine's thrust comes from the gas pressure inside the combustion chamber but the majority comes from the pressure against the inside of the expansion nozzle. Inside the combustion chamber the gas produces a similar force against all the sides of the combustion chamber but the throat gives no force producing an unopposed resultant force from the diametrically opposite end of the chamber. As the gases (adiabatically) expand inside the nozzle they press against the bell's walls forcing the rocket engine in one direction, and accelerating the gases in the opposite direction.
For optimum performance hot gas is used because it maximises the speed of sound at the throat — for aerodynamic reasons the flow goes sonic ("chokes") at the throat, so the highest speed there is desirable. By comparison, at room temperature the speed of sound in air is about 340m/s, the speed of sound in the hot gas of a rocket engine can be over 1700m/s.
The expansion part of the rocket nozzle then multiplies the speed of the flow by a further factor, typically between 1.5 and 4 times, giving a highly collimated exhaust jet. The speed ratio of a rocket nozzle is mostly determined by its area expansion ratio — the ratio of the area of the throat to the area at the exit, but details of the gas properties are also important. Larger ratio nozzles are more massive and bulkier, but they are able to extract more heat from the combustion gases, which become lower in pressure and colder, but also faster.
A significant complication arises when launching a vehicle from the Earth's surface as the ambient atmospheric pressure changes with altitude. For maximum performance it turns out that the pressure of the gas leaving a rocket nozzle should be the same as ambient pressure; if lower the vehicle will be slowed by the difference in pressure between the top of the engine and the exit, if higher then this represents pressure that the bell has not turned into thrust. To achieve this ideal, the diameter of the nozzle would need to increase with altitude, which is difficult to arrange. A compromise nozzle is generally used and some percentage reduction in performance occurs. To improve on this, various exotic nozzle designs such as the plug nozzle, stepped nozzles, the expanding nozzle and the aerospike have been proposed, each having some way to adapt to changing ambient air pressure and each allowing the gas to expand further against the nozzle giving extra thrust at higher altitude.
[edit] Thermal issues
The reaction mass's combustion temperatures can fairly typically reach ~3500 K (~5800 F) which is often far higher than the melting point of the nozzle and combustion chamber materials (~1200K for copper). Indeed many construction materials can make perfectly acceptable propellants in their own right. It is important that these materials be prevented from combusting, melting or vapourising to the point of failure. Materials technology could potentially place an upper limit on the exhaust temperature of chemical rockets.
To avoid this problem rockets can use ablative materials that erode in a controlled fashion, or very high temperature materials, such as graphite, ceramics or certain exotic metals.
Alternatively, rockets may use more common construction materials such as aluminum, steel, nickel or copper alloys and employ cooling systems that prevent the construction material itself becoming too hot. Regenerative cooling, where the propellant is passed through tubes around the combustion chamber or nozzle, and other techniques such as curtain cooling or film cooling, may be employed to give essentially unlimited nozzle and chamber life. These techniques ensure that the gas boundary layer touching the material is kept below the point where the material would fail.
[edit] Mechanical issues
The combustion chamber is often under substantial pressure, typically 10-200 bar, higher pressures giving better performance. This causes the outermost part of the chamber to be under very large hoop stresses.
Worse, due to the high temperatures created in rocket engines the materials used tend to have a significantly lowered working tensile strength.
[edit] Safety
Rocket engines are tested at a test facility before being put into production.
Rockets have a reputation for unreliability and danger; especially catastrophic failures.
Contrary to this reputation, carefully designed rockets can be made arbitrarily reliable. In military use, rockets are not unreliable. However, one of the main non-military uses of rockets is for orbital launch. In this application, the premium is on minimum weight, and it is difficult to achieve high reliability and low weight simultaneously. In addition, if the number of flights launched is low, there is a very high chance of a design, operations or manufacturing error causing destruction of the vehicle. Essentially, as of 2006 all launch vehicles are test vehicles by normal aerospace standards.
The X-15 rocket plane achieved a 0.5% failure rate, with a single catastrophic failure during ground test, and the SSME has managed to avoid catastrophic failures in over 300 engine-flights.
[edit] Noise
The Saturn V launch was detectable on seismometers a considerable distance from the launch site. As the hypersonic exhaust mixes with the ambient air, shock waves are formed. The sound intensity from these shock waves depends on the size of the rocket, and on large rockets can actually kill. The Space Shuttle generates over 200 dB(A) of noise around its base.
Generally speaking noise is most intense when a rocket is close to the ground, since the noise from the engines radiate up away from the plume, as well as reflecting off the ground. This noise can be reduced somewhat by flame trenches with roofs, by water injection around the plume and by deflecting the plume at an angle.
[edit] Chemistry
Contrary to popular belief, while rocket propellants require reasonably high energy per kilogram, many common materials are more energetic; for example petrol/gasoline or paraffin has as much energy as a rocket fuel and far more than the fuel/oxidiser mix used for rocket fuels. This is due to the necessity of the propellant containing large amounts of oxidiser, normal propellants used on earth for say, Turbojet engines, are reacted with the atmosphere and hence can have several times higher energy density.
Good rocket propellants require large quantities of hydrogen in the propellant, as this gives the highest exhaust speeds primarily due to the low molecular mass; but this is not the whole story.[1]
Programs exist to predict the performance of propellants in rocket engines.(e.g.[2]).
[edit] Ignition
With liquid propellants immediate ignition of the propellants as they first enter the combustion chamber is essential.
Failure to ignite within milliseconds causes too much liquid propellant to be within the chamber, and if/when ignition occurs the amount of hot gas created will often exceed the maximum design pressure of the chamber. The pressure vessel will often fail catastrophically. This is sometimes called a hard start.
Ignition can be achieved by a number of different methods; a pyrotechnic charge can be used, the propellants can ignite spontaneously on contact (hypergolic), a plasma torch can be used, or electric spark plugs may be employed.
Gaseous propellants generally will not cause hardstarts, with rockets the total injector area is less than the throat thus the chamber pressure tends to ambient prior to ignition and high pressures cannot form even if the entire chamber is full of flammable gas at ignition.
Solid propellants are usually ignited with one-shot pyrotechnic devices.
Once ignited, rocket chambers are self sustaining and igniters are not needed, indeed chambers often spontaneously reignite if restarted after being shut down for a few seconds. However, when cooled, many rockets cannot be started more than once without minor maintenance, such as replacement of the pyrotechnic igniter.
[edit] Types of rocket engines
See also Liquid rocket propellants.
| Type | Description | Advantages | Disadvantages |
|---|---|---|---|
| water rocket | Partially filled pressurised carbonated drinks container with tail and nose weighting | Very simple to build | altitude limited to a few hundred feet or so |
| cold gas thruster | A non combusting form, used for attitude jets | Non contaminating exhaust | Low performance |
| Solid rocket | Ignitable, self sustaining solid fuel/oxidiser mixture ("grain") with central hole and nozzle | Simple, often no moving parts, reasonably good mass fraction, reasonable Isp. A thrust schedule can be designed into the grain. | Once lit, extinguishing it is difficult although often possible, cannot be throttled in real time; handling issues from ignitable mixture, lower performance than liquid rockets, if grain cracks it can block nozzle with disastrous results, cracks burn and widen during burn. Refuelling grain harder than simply filling tanks. |
| Hybrid rocket | Separate oxidiser/fuel, typically oxidiser is liquid and kept in a tank, the other solid with central hole | 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, central hole widens over burn and negatively affects mixture ratio. Replacing grain harder than simply refuelling tanks. |
| Monopropellant rocket | Propellant such as Hydrazine, Hydrogen Peroxide or Nitrous Oxide, flows over catalyst and exothermically decomposes and 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, Isp is perhaps 1/3 of best liquids |
| 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. |
| 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 Isp, improves payload for launching from Earth by a sizeable percentage | Similar issues to bipropellant, but with more plumbing, more R&D |
| 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 oxidizer 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 oxidizer like LOX can be dangerous. Much heavier than simple rockets. |
| Precooled jets / 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, SABRE, ATREX |
[edit] Electric heating
| Type | Description | Advantages | Disadvantages |
|---|---|---|---|
| Resistojet rocket (electric heating) | A monopropellant is electrically heated by a filament for extra performance | Higher Isp than monopropellant alone, about 40% higher. | Uses a lot of power and hence gives typically low thrust |
| Arcjet rocket (chemical burning aided by electrical discharge) | Similar to resistojet in concept but with inert propellant, except an arc is used which allows higher temperatures | 1600 seconds Isp | Very low thrust and high power, performance is similar to Ion drive. |
| Pulsed plasma thruster (electric arc heating; emits plasma) | Plasma is used to erode a solid propellant | High Isp, can be pulsed on and off for attitude control | Low energetic efficiency |
| Variable specific impulse magnetoplasma rocket | Microwave heated plasma with magnetic throat/nozzle | Variable Isp from 1000 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 |
[edit] Solar heating
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 concentrators and mirrors. 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.
| Type | Description | Advantages | Disadvantages |
|---|---|---|---|
| Solar thermal rocket | Propellant is heated by solar collector | Reasonably simple, good performance with liquid hydrogen propellant, adequate performance with in-situ water for short-range interplanetary flight | only useful once in space, as thrust is fairly low, but hydrogen is not easily stored in space, otherwise moderate/low Isp if higher molecular mass propellants are used |
[edit] Nuclear heating
Nuclear propulsion includes a wide variety of 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:
| Type | Description | Advantages | Disadvantages |
|---|---|---|---|
| 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 (typ. hydrogen) is passed through a nuclear reactor to heat to high temperature | Isp 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 |
| 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, Isp between 1500 and 3000 seconds but with very high thrust | difficulties in heating propellant without losing fissionables in exhaust, exhaust inherently highly radioactive, massive thermal issues particularly for nozzle/throat region |
| 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 | |
| Nuclear salt-water rocket (nuclear fission energy) | Nuclear salts are held in solution, caused to react at nozzle | Very high Isp, very high thrust | Thermal issues in nozzle, propellant could be unstable, highly radioactive exhaust |
| Nuclear pulse propulsion (exploding fission/fusion bombs) | Shaped nuclear bombs are detonated behind vehicle and blast is caught by a 'pusher plate' | Very high Isp, very high thrust/weight ratio, no show stoppers are known for this technology | Never been tested, pusher plate may throw off fragments due to shock, minimum size for nuclear bombs is still pretty big, expensive at small scales, nuclear treaty issues |
| 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 isn't currently feasible |
| 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 reaction is used to heat propellant | Extremely energetic, very high exhaust velocity is possible on paper | Antimatter containment issues, thermal issues, beyond current state of the art. |
[edit] See also
- NERVA - NASA's Nuclear Energy for Rocket Vehicle Applications, a US nuclear thermal rocket programme
- Project Prometheus, NASA development of nuclear propulsion for long-duration spaceflight, begun in 2003
[edit] References
- ^ Newsgroup correspondence, 1998-99
- ^ Complex chemical equilibrium and rocket performance calculations, Cpropep-Web
