Rocket engine
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2008-07-16T02:44:28Z
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/* Combustion instabilities */
[[Image:RS-68 rocket engine test.jpg|thumb|right|[[RS-68]] being tested at NASA's Stennis Space Center. The nearly transparent exhaust is due to this engine's exhaust being mostly superheated steam (water vapor from its propellants, hydrogen and oxygen)]]
[[Image:Viking 5C rocketengine.jpg|thumb|right|[[Viking (rocket engine)|Viking 5C rocket engine]] ]]
A '''rocket engine''' is a [[jet engine]]<ref ="RPECH1">Rocket Propulsion Elements; 7th edition- chapter 1</ref> that takes all its reaction mass ("''[[propellant]]''") from within tankage or the body of a [[rocket vehicle|vehicle]] and forms it into a high speed [[Jet (fluid)|jet]], thereby obtaining thrust in accordance with [[Newton's third law]]. [[Rocket]] engines can be used for [[spacecraft propulsion]] as well as terrestrial uses, such as [[missile]]s. Most rocket engines are [[internal combustion engine]]s, although non combusting forms also exist.
==Principle of operation==
[[Image:Rückstoss1600.png|thumb|300pix|right|Rocket engines give part of their thrust due to unopposed pressure on the combustion chamber]]<!-- It's a rubbishy picture, since most of the thrust is from the nozzle, but best I could find... --->
Most rocket engines produce thrust by the expulsion of a high-temperature, high-speed gaseous exhaust. This is typically created by high pressure (10-200 bar) combustion of solid or liquid [[Rocket propellant|propellants]], consisting of [[fuel]] and [[oxidiser]] components, within a [[combustion chamber]].
===Introducing propellant into a combustion chamber===
[[Liquid rocket|Liquid-fueled rockets]] typically pump separate fuel and oxidiser components into the combustion chamber, where they mix and burn. [[Solid rocket]] propellants are prepared as a mixture of fuel and oxidizing components and the propellant storage chamber becomes the combustion chamber. [[Hybrid rocket]] engines use a combination of solid and liquid or gaseous propellants. Alternatively, a chemically inert [[reaction mass]] can be heated using a high-energy power source via a heat exchanger, and then no combustion chamber is used.
Rocket propellant is mass that is stored, usually in some form of propellant tank, prior to being used as the propulsive mass that is ejected from a rocket engine in the form of a fluid jet to produce thrust.
Chemical rocket propellants are most commonly used, which undergo exothermic chemical reactions which produce hot gas which is used by a rocket for propulsive purposes.
===Combustion chamber===
For chemical rockets the combustion chamber is typically just a cylinder. The dimensions of the cylinder is such that the propellant is able to combust thoroughly; different propellants require different combustion chamber sizes for this to occur. This leads to a number called <math>L*</math>:
:<math>L* = \frac {V_c} {A_t}</math>
where:
:<math>V_c</math> is the Volume of the chamber
:<math>A_t</math> is the area of the throat
L* is typically in the range of 25-60 inches (0.6 - 1.5 m)
The combination of temperatures and pressures typically reached in a combustion chamber are usually extreme by any standards. Unlike air-breathing jet engines no atmospheric nitrogen is present to dilute and cool the combustion and the temperature can reach true [[stochiometric]]. The high pressures mean that the rate of conduction of heat through the walls is very high indeed.
===Rocket nozzles===
{{main|Rocket engine nozzle}}
[[Image:Nozzle de Laval diagram.svg|thumb|right|Typical temperatures (T) and pressures (p) and speeds (v) in a De Laval Nozzle]]
The large bell or cone shaped expansion nozzle gives a rocket engine its characteristic shape.
In rockets the hot gas produced in the combustion chamber is permitted to escape from the combustion chamber through an opening (the "throat"), within a [[Rocket engine nozzles|high expansion-ratio 'de Laval nozzle']].
Provided sufficient pressure is provided to the nozzle (about 2.5-3x above ambient pressure) the nozzle ''[[choke]]s'' and a supersonic jet is formed, dramatically accelerating the gas, converting most of the thermal energy into kinetic energy. Exhaust speeds as high as ten times the [[speed of sound]] at sea level are not uncommon.
[[Image:rocketThrust.svg|thumb|right|Rocket thrust is caused by pressures acting in the combustion chamber and nozzle. From Newtons third law, equal and opposite pressures act on the exhaust, and this accelerates it to high speeds.]]
A portion of the rocket engine's thrust comes from the unbalanced pressures inside the combustion chamber but the majority comes from the pressures against the inside of the nozzle (see diagram). As the gas expands ([[Adiabatic process|adiabatically]]) the pressure against the nozzle's walls forces the rocket engine in one direction while accelerating the gas in the other.
===Propellant efficiency===
For a rocket engine to be propellant efficient, it is important that the maximum pressures possible be created by a specific amount of propellant acting on the chamber and nozzle. This can be achieved by all of:
*heating the propellant to as high a temperature as possible (using a high energy fuel, containing hydrogen and carbon and sometimes metals such as aluminium, or even using nuclear energy)
*using a low specific density gas (as hydrogen rich as possible)
*using propellants which are, or decompose to, simple molecules with few degrees of freedom so as to maximise molecular speeds
Since all of these things minimise the mass of the propellant used, and since pressure is proportional to the amount of propellant present to be accelerated as it pushes on the engine, and since from Newton's third law the pressure that acts on the engine also reciprocally acts on the propellant, it turns out that the speed that the propellant leaves the chamber is unaffected by the chamber pressure (although the thrust is proportional). However, speed is significantly affected by all three of the above factors and the exhaust speed is an excellent measure of the engine propellant efficiency.
For aerodynamic reasons the flow goes sonic ("[[Choked flow|chokes]]") at the narrowest part of the nozzle, the 'throat'. Since the [[speed of sound]] in gases increases with the square root of temperature, the use of hot exhaust gas greatly improves performance. By comparison, at room temperature the speed of sound in air is about 340 m/s while the speed of sound in the hot gas of a rocket engine can be over 1700 m/s; much of this performance is due to the higher temperature, but additionally rocket propellants are chosen to be of low molecular mass, and this also gives a higher velocity compared to air.
Expansion in the rocket nozzle then further multiplies the speed, typically between 1.5 and 2 times, giving a highly [[collimated]] hypersonic exhaust jet. The speed increase 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 detailed properties of the gas are also important. Larger ratio nozzles are more massive but are able to extract more heat from the combustion gases, increasing the exhaust velocity.
Nozzle efficiency is affected by operation in the atmosphere because atmospheric pressure changes with altitude; but due to the supersonic speeds of the gas exiting from a rocket engine, the pressure of the jet may be either below or above ambient, and equilibrium between the two is not reached.
===Back pressure and optimal expansion===
For optimal performance the pressure of the gas at the end of the nozzle should just equal the ambient pressure: if the exhaust's pressure is lower than the ambient pressure, then the vehicle will be slowed by the difference in pressure between the top of the engine and the exit; on the other hand, if the exhaust's pressure is higher, then exhaust pressure that could have been converted into thrust is not converted, and energy is wasted.
To maintain this ideal of equality between the exhaust's exit pressure and the ambient pressure, the diameter of the nozzle would need to increase with altitude, giving the pressure a longer nozzle to act on (and reducing the exit pressure and temperature). This increase is difficult to arrange. A compromise nozzle is generally used and some 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 engine|aerospike]] have been proposed, each providing 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 altitudes.
==Overall rocket engine performance==
Rocket technology can combine very high thrust ([[meganewton]]s), very high exhaust speeds (around 10 times the speed of sound at sea level) and very high thrust/weight ratios (>100) ''simultaneously'' as well as being able to operate outside the atmosphere.
Rockets can be further optimised to even more extreme performance along one or more of these axes at the expense of the others.
===Specific impulse===
{{main|Specific impulse}}
The most important metric for the efficiency of a rocket engine is [[impulse]] per unit of [[propellant]], this is called [[specific impulse]] (usually written <math>I_{sp}</math>). This is either measured as a speed (<math>V_{e}</math> in metres/second or ft/s) or as a time (seconds). An engine that gives a large specific impulse is normally highly desirable.
===Net thrust===
Below is an approximate equation for calculating the net thrust of a rocket engine:
:<math>F_n = \dot{m}\;V_{e} = \dot{m}\;V_{e-act} + A_{e}(P_{e} - P_{amb})</math><ref>Rocket Propulsion Elements seventh edition eq-2-14</ref>
where:
:<math> \dot{m} = \,</math>exhaust gas mass flow
:<math>V_{e} =\,</math>effective exhaust velocity
:<math>V_{e-act} =\,</math>actual jet velocity at nozzle exit plane
:<math>A_{e} =\,</math>flow area at nozzle exit plane
:<math>P_{e} =\,</math>static pressure at nozzle exit plane
:<math>P_{amb} =\,</math>ambient (or atmospheric) pressure
Since, unlike a jet engine, a conventional rocket motor lacks an air intake, there is no 'ram drag' to deduct from the gross thrust. Consequently the net thrust of a rocket motor is equal to the gross thrust (apart from static back pressure).
The <math>\dot{m}\;V_{e-act}\,</math> term represents the momentum thrust, which remains constant at a given throttle setting, whereas the <math>A_{e}(P_{e} - P_{amb})\,</math> term represents the pressure thrust term. At full throttle, the net thrust of a rocket motor improves slightly with increasing altitude, because the reducing atmospheric pressure increases the pressure thrust term.
[[Image:NozzleExpansion.svg|thumb|right|275px|If the pressure of the exhaust jet varies from atmospheric pressure, nozzles can be said to be underexpanded, ambient or overexpanded. If under or overexpanded then loss of efficiency occurs, grossly overexpanded nozzles lose less efficiency, but the exhaust jet is usually unstable. Rockets become progressively more underexpanded as they gain altitude. Note that almost all rocket engines will be momentarily grossly overexpanded during startup in an atmosphere.<ref>{{cite book|author=Huzel, D. K. and Huang, D. H.|title=NASA SP-125, Design of Liquid Propellant Rocket Engines|edition=2nd Edition|publisher=NASA|year=1971|url=http://ntrs.nasa.gov/search.jsp?Ntt=sp-125&Ntk=all&Ntx=mode+matchall}}</ref>]]
===Vacuum Isp===
Due to the specific impulse varying with pressure, a quantity that is easy to compare and calculate with is useful. Because rockets [[choked flow|choke]] at the throat, and because the supersonic exhaust prevents external pressure influences travelling upstream, it turns out that the pressure at the exit is ideally exactly proportional to the propellant flow <math> \dot{m}</math>, provided the mixture ratios and combustion efficiencies are maintained. It is thus quite usual to rearrange the above equation slightly:
:<math> Fvac = C_f \dot{m} c^*</math><ref>Rocket Propulsion Elements seventh edition eq-3-33</ref>
and so define the ''vacuum Isp'' to be:
:<math>V_{evac} = C_f c^*</math>
Where:
:<math>c^* =\,</math> the speed of sound constant at the throat
:<math>C_f =\,</math> the thrust coefficient constant of the nozzle (typically between 0.8 and 1.9)
And hence:
:<math> F_n = \dot{m} V_{evac} - A_{e} P_{amb}</math>
===Throttling===
Rockets can be throttled by controlling the propellant rate <math> \dot{m}</math> (usually measured in kg/s or lb/s).
In principle rockets can be throttled down to an exit pressure of about one-third of ambient pressure (due to flow separation in nozzles) and up to a maximum limit determined only be the mechanical strength of the engine.
In practice, the degree to which rockets can be throttled varies greatly, but most rockets may be throttled by a factor of 2 without great difficulty; the typical limitation is combustion stability, as for example, injectors need a minimum pressure to avoid triggering damaging oscillations (chugging or combustion instabilities); but injectors can often be optimised and tested for wider ranges. Additionally, it is important that the exit pressure not be too far below ambient to avoid flow separation in the nozzle.
===Energy efficiency===
[[Image:PropulsiveEfficiency.GIF|left|thumb|Rocket energy efficiency as a function of vehicle speed divided by effective exhaust speed]]
Rocket engine nozzles are surprisingly efficient [[heat engines]] for generating a high speed jet, as a consequence of the high combustion temperature and high [[compression ratio]] in accordance with the [[carnot cycle]]. For a vehicle employing a rocket engine the energetic efficiency is very good if the vehicle speed approaches or somewhat exceeds the exhaust velocity (relative to launch); but at low speeds the energy efficiency goes to 0% at zero speed (as with all [[jet propulsion]].) <!--- it's very counterintuitive, a way to look at it is that energy= force x distance, but at zero speed you have no movement and the force is costing you energy in the jet --->See [[Rocket#Energy efficiency|Rocket energy efficiency]] for more details.
==Cooling==
For efficiency reason, and because they physically can, rockets run with reaction mass's combustion temperatures that can reach ~3500 K (~5800 °F).
The temperatures employed are very often far higher than the melting point of the nozzle and combustion chamber materials, two exceptions are [[graphite]] and [[tungsten]] (~1200 K 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.
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, are employed to give longer nozzle and chamber life. These techniques ensure that a gaseous thermal [[boundary layer]] touching the material is kept below the temperature which would cause the material to catastrophically fail.
In rockets, the heat fluxes that can pass through the wall are among the highest in engineering, fluxes are generally in the range of 1-200 MW/m^2. The strongest heat fluxes are found at the throat, which often sees twice that found in the associated chamber and nozzle. This is due to the combination of high speeds (which gives a very thin boundary layer), and the high temperatures seen there.
Most other jet engines have gas turbines in the hot exhaust. Due to their larger surface area, they are harder to cool and hence need to run their combustion processes at much lower temperatures, losing efficiency.
In rockets the coolant methods include:
#uncooled (used for short runs mainly during testing)
#[[ablation|ablative]] walls (walls are lined with a material that is continuously vapourised and carried away).
#[[radiative cooling]] (the chamber becomes almost white hot and radiates the heat away)
#dump cooling (a propellant, usually [[hydrogen]], is passed around the chamber and dumped)
#[[regenerative cooling]] (uses the propellant to cool the chamber via a cooling jacket before being injected)
#curtain cooling (propellant injection is arranged so the temperature of the gases is cooler at the walls)
#film cooling (surfaces are wetted with liquid propellant, which cools as it evaporates)
In all cases the cooling effect that prevents the wall from being destroyed is caused by a thin layer of insulating fluid (a [[boundary layer]]) that is in contact with the walls that is far cooler than the combustion temperature. Provided this boundary layer is intact the wall will not be damaged.
Disruption of the boundary layer may occur during cooling failures or combustion instabilities, and wall failure typically occurs soon after.
With regenerative cooling a second boundary layer is found in the coolant channels around the chamber. This boundary layer thickness needs to be as small as possible, since the boundary layer acts as an insulator between the wall and the coolant. This may be achieved by making the coolant [[velocity]] in the channels as high as possible.
==Mechanical issues==
Rocket combustion chambers are normally operated at fairly high pressure, typically 10-200 bar (1 to 20 MPa), higher pressures often give better performance (by permitting a more efficient nozzle to be fitted). This causes the outermost part of the chamber to be under very large [[hoop stress]]es.
Worse, due to the high temperatures created in rocket engines the materials used tend to have a significantly lowered working tensile strength.
==Acoustic issues==
In addition, the extreme vibration and acoustic environment inside a rocket motor commonly results in peak stresses well above mean values, especially in the presence of [[organ pipe]]-like resonances and gas turbulence.
===Combustion instabilities===
Three different types of combustion instabilities occur
;Chugging
This is a low frequency oscillation at a few Hertz in chamber pressure usually caused by pressure variations in feed lines due to variations in acceleration of the vehicle. This can cause cyclic variation in thrust, and the effects can vary from merely annoying to actually damaging the payload or vehicle. Chugging can be minimised by using gas filled damping tubes on feed lines of high density propellant's.
;Buzzing
This can be caused due to insufficient pressure drop across the injectors. It generally is mostly annoying, rather than being damaging. However, in extreme cases combustion can end up being forced backwards through the injectors- this can cause explosions with monopropellants.
;Screeching
This the most immediately damaging, and the hardest to control. It is due to acoustics within the combustion chamber that often couples to the chemical combustion processes that are the primary drivers of the energy release<ref>Acoustic oscillations can be excited by thermal processes, such as the flow of hot air through a pipe or combustion in a chamber. Specifically, standing acoustic waves inside a chamber can be intensified if combustion occurs more intensely in regions where the pressure of the acoustic wave is maximal. According to Lord Rayleigh’s criterion for thermoacoustic processes, "If heat be given to the air at the moment of greatest condensation, or be taken from it at the moment of greatest rarefaction, the vibration is encouraged. On the other hand, if heat be given at the moment of greatest rarefaction, or abstracted at the moment of greatest condensation, the vibration is discouraged." (Reference: John Wm. Strutt, Baron Rayleigh, ''The Theory of Sound'', 2nd ed. (London: Macmillan, 1896) (reprinted by Dover Publications (N.Y., N.Y.) in 1945), vol. 2, page 226.) See also: (1) Lord Rayleigh (1878) "The explanation of certain acoustical phenomena" (namely, the [[Rijke tube]]) ''Nature'', vol. 18, pages 319–321; and (2) E. C. Fernandes and M. V. Heitor, “Unsteady flames and the Rayleigh criterion” in F. Culick, M. V. Heitor, and J. H. Whitelaw, ed.s, ''Unsteady Combustion'' (Dordrecht, the Netherlands: Kluwer Academic Publishers, 1996), page 4. Available on-line at: http://books.google.com/books?id=Je_hG6UfnogC&printsec=copyright&dq=rayleigh+thermoacoustic+&ie=ISO-8859-1&source=gbs_toc_s&cad=1#PPA4,M1 </ref>, and can lead to unstable resonant "screeching" that commonly leads to catastrophic failure due to thinning of the insulating thermal boundary layer.<ref>G.P. Sutton & D.M. Ross, ''Rocket Propulsion Elements, 4th Ed.'' 1975 (WileyInterscience New York), Ch 8, Sec 6 & especially Sec. 7, re combustion instability.</ref>
Such effects are very difficult to predict analytically during the design process, and have usually been addressed by expensive and time consuming extensive testing, combined with trial and error remedial correction measures.
Screeching is often dealt with by detailed changes to injectors, or changes in the propellant chemistry, or vaporizing the propellant before injection of the use of [[Helmholtz damper]]s within the combustion chambers to change the resonant modes of the chamber.
Testing for the possibility of screeching is sometimes done by exploding small explosive charges in the combustion chamber to determine the engine's [[impulse response]] and then evaluating the time response of the chamber pressure- a fast recovery indicates a stable system.
===Exhaust noise===
For all but the very smallest sizes, rocket exhaust compared to other engines is generally very noisy indeed. As the [[hypersonic]] exhaust mixes with the ambient air, [[shock wave]]s are formed. The [[sound intensity]] from these shock waves depends on the size of the rocket, and on large rockets could potentially kill at close range.<ref name="CR566">[http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19660030602_1966030602.pdf NASA CR-566]</ref>
The [[Saturn V]] launch was detectable on [[seismometer]]s a considerable distance from the launch site. The [[sound intensity]] from the shock waves generated depends on the size of the rocket and on the exhaust velocity. Such shock waves seem to account for the characteristic crackling and popping sounds produced by large rocket engines when heard live. These noise peaks typically overload microphones and audio electronics, and so are generally weakened or entirely absent in recorded or broadcast audio reproductions. For large rockets the acoustic effects could 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 radiates up away from the plume, as well as reflecting off the ground. Also, when the vehicle is moving slowly, little of the chemical energy input to the engine can go into increasing the kinetic energy of the rocket (since useful power ''P'' transmitted to the vehicle is <math>P = F*V</math> for thrust ''F'' and speed ''V''). Then the largest portion of the energy is dissipated in the exhaust's interaction with the ambient air, producing noise. 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.
==Testing==
Rocket engines are usually statically tested at a [[rocket engine test facility|test facility]] before being put into production. For high altitude engines, either a shorter nozzle must be used, or the rocket must be tested in a large vacuum chamber.
==Safety==
[[Rocket]]s 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 all launch vehicles are test vehicles by normal aerospace standards ([[as of 2006]]).
The [[X-15]] rocket plane [[Albert Scott Crossfield|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 350 engine-flights.
==Chemistry==
[[Rocket propellant]]s require a high specific energy (energy per unit mass), because ideally all the reaction energy appears as kinetic energy of the exhaust gases, and exhaust velocity is the single most important performance parameter of an engine, on which vehicle performance depends.
Aside from inevitable losses and imperfections in the engine, incomplete combustion, etc., after specific reaction energy, the main theoretical limit reducing the exhaust velocity obtained is that, according to the laws of thermodynamics, a fraction of the chemical energy may go into rotation of the exhaust molecules, where it is unavailable for producing thrust. Monatomic gases like helium have only three degrees of freedom, corresponding to the three dimensions of space, {''x,y,z''}, and only such spherically symmetric molecules escape this kind of loss. A diatomic molecule like H<sub>2</sub> can rotate about either of the two axes perpendicular to the one joining the two atoms, and as the [[equipartition]] law of statistical mechanics demands that the available thermal energy be divided equally among the degrees of freedom, for such a gas in thermal equilibrium 3/5 of the energy can go into unidirectional motion, and 2/5 into rotation. A triatomic molecule like water has six degrees of freedom, so the energy is divided equally among rotational and translational degrees of freedom. For most chemical reactions the latter situation is the case. This issue is traditionally described in terms of the ratio, gamma, of the specific heat of the gas at constant volume to that at constant pressure. The rotational energy loss is largely recovered in practice if the expansion nozzle is large enough to allow the gases to expand and cool sufficiently, the function of the nozzle being to convert the random thermal motions of the molecules in the combustion chamber into the unidirectional translation that produces thrust. As long as the exhaust gas remains in equilibrium as it expands, the initial rotational energy will be largely returned to translation in the nozzle.
Although the specific reaction energy per unit mass of reactants is key, low mean molecular weight in the reaction products is also important in practice in determining exhaust velocity. This is because the high gas temperatures in rocket engines pose serious problems for the engineering of survivable motors. Because temperature is proportional to the mean ''energy per molecule'', a given amount of energy distributed among more molecules of lower mass permits a higher exhaust velocity at a given temperature. This means low atomic mass elements are favored. Liquid hydrogen (LH2) and oxygen (LOX, or LO2), are the most effective propellants in terms of exhaust velocity that have been widely used to date, though a few exotic combinations involving boron or liquid ozone are potentially somewhat better in theory if various practical problems could be solved.<ref>[http://yarchive.net/space/rocket/fuels/fuel_ratio.html Newsgroup correspondence], 1998-99</ref>
It is important to note in computing the specific reaction energy, that the ''entire mass of the propellants, including both fuel and oxidizer'', must be included. The fact that air-breathing engines are typically able to obtain oxygen "for free" without having to carry it along, accounts for one factor of why air-breathing engines are very much more propellant-mass efficient, and one reason that rocket engines are far less suitable for most ordinary terrestrial applications. Fuels for automobile or [[turbojet engine]]s, utilize atmospheric oxygen and so have a much better effective energy output per unit mass of fuel that must be carried.
Computer programs that predict the performance of propellants in rocket engines are available.<ref>[http://rocketworkbench.sourceforge.net/equil.phtml Complex chemical equilibrium and rocket performance calculations], Cpropep-Web</ref>
==Ignition==
With liquid and hybrid rockets, immediate ignition of the propellant(s) as they first enter the combustion chamber is essential.
With liquid propellants (but not gaseous), failure to ignite within milliseconds usually 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 hard starts, 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 they are restarted after being shut down for a few seconds. However, when cooled, many rockets cannot be restarted without at least minor maintenance, such as replacement of the pyrotechnic igniter.
==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 ([[Water rocket#Altitude Record|world record]] is 623 [[meter]]s/2044 [[Foot (unit of length)|feet]])
|-
![[cold gas thruster]]
|A non combusting form, used for [[attitude jet]]s
|Non contaminating exhaust
|Extremely low performance
|-
![[hot water rocket]]
|Hot water is stored in a tank at high temperature/pressure and turns to steam in exhaust
|Simple, fairly safe
|Low performance due to heavy tank
|}
===Chemically powered===
{{See also|Liquid rocket propellants}}
{| class="wikitable"
!'''Type'''
!'''Description'''
!'''Advantages'''
!'''Disadvantages'''
|-
! [[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 [[Specific Impulse|''I''<sub>sp</sub>]]. 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, Lower specific Impulse than Liquid Rockets.
|-
![[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.
|-
![[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, [[Specific Impulse|''I''<sub>sp</sub>]] 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 [[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 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 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]], [[SABRE]], [[ATREX]]
|}
=== Electrically powered ===
{| class="wikitable"
!'''Type'''
!'''Description'''
!'''Advantages'''
!'''Disadvantages'''
|-
! [[Resistojet rocket]] (electric heating)
|A monopropellant is electrically heated by a filament for extra performance
|Higher [[Specific Impulse|''I''<sub>sp</sub>]] 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 [[Specific Impulse|''I''<sub>sp</sub>]]
|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 [[Specific Impulse|''I''<sub>sp</sub>]] , 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 [[Specific Impulse|''I''<sub>sp</sub>]] 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
|}
{{See also|Ion thruster}}
=== Solar powered ===
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 [[solar concentrator|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.
{| class="wikitable"
!'''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 [[Specific Impulse|''I''<sub>sp</sub>]] if higher molecular mass propellants are used
|}
{{See also|Ion thruster}}
===Beam powered===
{| class="wikitable"
!'''Type'''
!'''Description'''
!'''Advantages'''
!'''Disadvantages'''
|-
![[Beam-powered 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
|microwaves avoid reemission of energy, so ~900 seconds exhaust speeds might be achieveable
|~1 MW of power per kg of payload is needed to achieve orbit, relatively high accelerations, microwaves are absorbed to a degree by rain, reflected microwaves may be dangerous, pretty much needs hydrogen monopropellant for good performance which needs heavy tankage, transmitter diameter is measured in kilometres to achieve a fine enough beam to hit a vehicle at up to 100 km.
|}
=== Nuclear powered ===
[[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'''
|-
! [[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 (typ. 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.
|-
! [[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 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
|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 isn't 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
|}
==History of rocket engines==
According to the writings of the Roman [[Aulus Gellius]], in c. [[400 BC]], a [[Greek people|Greek]] [[Pythagorean]] named [[Archytas]], propelled a wooden bird along wires using steam.<ref>Leofranc Holford-Strevens, ''Aulus Gellius: An Antonine Author and his Achievement'' (Oxford University Press; revised paperback edn. 2005)
*{{1911}}</ref> However, it would not appear to have been powerful enough to take off under its own thrust.
The ''[[aeolipile]]'' invented in the [[1st century|first century]] (known as ''[[Hero's engine]]'') essentially consists of a [[hot water rocket]] on a [[bearing]]. It was created almost two millennia before the industrial revolution. Apparently Hero's steam engine was taken to be little more than a toy, the principles behind it were not well understood, and its full potential not realized for a millennium.
The availability of [[black powder]] to propel projectiles was a precursor to the development of the first solid rocket. [[9th century|Ninth Century]] [[Chinese people|Chinese]] [[Taoist]] [[Alchemy|alchemists]] discovered [[black powder]] in a search for the [[Elixir of life]]; this accidental discovery led to [[fire arrow]]s which were the first rocket engines to leave the ground.
Rocket engines were also brought in use by [[Tippu Sultan]], The king of Mysore. These rockets could be of various sizes, but usually consisted of a tube of soft hammered iron about 8" long and 1½ - 3" diameter, closed at one end and strapped to a shaft of bamboo about 4ft. long. The iron tube acted as a combustion chamber and contained well packed black powder propellant. A rocket carrying about one pound of powder could travel almost 1,000 yards. These 'rockets', fitted with swords used to travel long distance, several meters above in air before coming down with swords edges facing the enemy. These rockets were used against British empire very effectively.
Slow development of this technology continued up to the later 20th Century, when the writings of [[Konstantin Tsiolkovsky]] first talked about [[liquid rocket|liquid fuelled rocket engines]].
These independently became a reality thanks to [[Robert Goddard (scientist)|Robert Goddard]].
Liquid hydrogen engines were first successfully developed in America, the [[RL-10]] engine first flew in 1962. Hydrogen engines were used as part of the [[Project Apollo]]; the liquid hydrogen fuel giving a rather lower stage mass and thus reducing the overall size and cost of the vehicle.
The Space Shuttle's [[SSME]] is the highest [[specific impulse]] rocket engine to fly.
==References==
{{reflist}}
{{Citations missing|date=September 2007}}
== 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
*[[Jet damping]] an effect of the exhaust jet of a rocket that tends to slow a vehicle's rotation speed
{{Machine configurations|state=uncollapsed}}
==External links==
*[http://www.pwrengineering.com/articles/longterm.htm Designing for rocket engine life expectancy]
*[http://www.pwrengineering.com/articles/plume.htm Rocket Engine performance analysis with Plume Spectrometry]
*[http://www.pwrengineering.com/articles/heart.htm Rocket Engine Thrust Chamber technical article]
[[Category:Spacecraft propulsion]]
[[Category:Spacecraft components]]
[[Category:Aerospace engineering]]
[[Category:Rocket engines|*]]
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