Editing Rocket engine (section)

==Principle of operation==
[[File:Liquid-Fuel Rocket Diagram.svg|thumb|upright=1.25|Simplified diagram of a liquid-fuel rocket: {{olist |
    |[[Liquid rocket propellant|Liquid fuel]] tank
    |[[Oxidizing agent|Liquid oxidiser]] tank
    |Pumps feed fuel and oxidiser under high pressure.
    |[[Combustion chamber]] mixes and burns the propellants.
    |[[Propelling nozzle|Exhaust nozzle]] expands and accelerates the gas jet to produce thrust.
    |Exhaust exits nozzle.
}}]]

[[File:Solid-Fuel Rocket Diagram.svg|thumb|upright=1.25|Simplified diagram of a solid-fuel rocket: {{olist
    |Solid [[Rocket propellant#Solid chemical propellants|fuel–oxidiser mixture]] (propellant) packed into casing 
    |[[Pyrotechnic initiator|Igniter]] initiates propellant combustion.
    |Central hole in propellant acts as the [[combustion chamber]].
    |[[Propelling nozzle|Exhaust nozzle]] expands and accelerates the gas jet to produce thrust.
    |Exhaust exits nozzle.
}}]]
Rocket engines produce thrust by the expulsion of an exhaust [[fluid]] that has been accelerated to high speed through a [[propelling nozzle]]. The fluid is usually a gas created by high pressure ({{convert|10|to|300|bar|psi|order=flip|adj=on}}) combustion of solid or liquid [[Rocket propellant|propellants]], consisting of [[fuel]] and [[oxidizing agent|oxidiser]] components, within a [[combustion chamber]]. As the gases expand through the nozzle, they are accelerated to very high ([[supersonic]]) speed, and the reaction to this pushes the vehicle ([[rocket]]) in the opposite direction. Combustion is most frequently used for practical rockets, as the laws of [[thermodynamics]] (more specifically [[Carnot's theorem (thermodynamics)|Carnot's theorem]]) dictate that high temperatures and pressures are desirable for the best [[thermal efficiency]]. [[Nuclear thermal rocket]]s are capable of higher efficiencies, but have low thrust, thanks to the low mass of the propellants used, and also have [[Nuclear thermal rocket#Risks|environmental problems]] which preclude their routine use in the [[Atmosphere of Earth|Earth's atmosphere]] and [[cislunar space]].

For [[model rocket]]ry, an available alternative to combustion is a [[water rocket]] pressurized by [[compressed air]], [[carbon dioxide]], [[nitrogen]], or any other readily available, inert gas.

===Propellant===
Rocket propellant is mass that is stored, usually in some form of tank, or within the combustion chamber itself, prior to being ejected from a rocket engine in the form of a fluid jet to produce thrust.

Chemical rocket propellants are the most commonly used. These undergo exothermic chemical reactions  producing a hot jet of gas for propulsion. Alternatively, a chemically inert [[reaction mass]] can be heated by a high-energy power source through a heat exchanger in lieu of a combustion chamber.

[[Solid rocket]] propellants are prepared in a mixture of fuel and oxidising components called ''grain'', and the propellant storage casing effectively becomes the combustion chamber.

===Injection===
[[Liquid-propellant rocket|Liquid-fueled rockets]] force separate fuel and oxidizer components into the combustion chamber, where they mix and burn. [[Hybrid rocket]] engines use a combination of solid and liquid or gaseous propellants. Both liquid and hybrid rockets use ''[[Liquid-fuel rocket#Injectors|injectors]]'' to introduce the propellant into the chamber. These are often an array of simple [[jet (nozzle)|jet]]s – holes through which the propellant escapes under pressure; but sometimes may be more complex spray nozzles. When two or more propellants are injected, the jets usually deliberately cause the propellants to collide as this breaks up the flow into smaller droplets that burn more easily.

===Combustion chamber===
For chemical rockets the combustion chamber is typically cylindrical, and [[flame holder]]s, used to hold a part of the combustion in a slower-flowing portion of the combustion chamber, are not needed. The dimensions of the cylinder are such that the propellant is able to combust thoroughly; different [[rocket propellant]]s require different combustion chamber sizes for this to occur.

This leads to a number called <math>L^*</math>, the [[characteristic length]]:
:<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 of the nozzle.
L* is typically in the range of {{convert|64|-|152|cm|in}}.

The temperatures and pressures typically reached in a rocket combustion chamber in order to achieve practical [[thermal efficiency]] are extreme compared to a [[afterburner|non-afterburning]] [[airbreathing jet engine]]. No atmospheric nitrogen is present to dilute and cool the combustion, so the propellant mixture can reach true [[stoichiometric]] ratios. This, in combination with the high pressures, means that the rate of heat conduction through the walls is very high.

In order for fuel and oxidiser to flow into the chamber, the pressure of the propellants entering the combustion chamber must exceed the pressure inside the combustion chamber itself.  This may be accomplished by a variety of design approaches including [[turbopump]]s or, in simpler engines, via [[Pressure-fed cycle (rocket)|sufficient tank pressure]] to advance fluid flow. Tank pressure may be maintained by several means, including a high-pressure [[helium]] pressurization system common to many large rocket engines or, in some newer rocket systems, by a bleed-off of high-pressure gas from the engine cycle to [[autogenous pressurization|autogenously pressurize]] the propellant tanks<ref name=nsf20160927>
{{cite news |last=Bergin|first=Chris |url=https://www.nasaspaceflight.com/2016/09/spacex-reveals-mars-game-changer-colonization-plan/ |title=SpaceX reveals ITS Mars game changer via colonization plan |work=[[NASASpaceFlight.com]] |date=2016-09-27 |access-date=2016-09-27 }}</ref><ref name=sfi20160927/>  For example, the self-pressurization gas system of the [[SpaceX Starship]] is a critical part of SpaceX strategy to reduce launch vehicle fluids from five in their legacy Falcon 9 vehicle family to just two in Starship, eliminating not only the helium tank pressurant but all [[hypergolic propellant]]s as well as [[nitrogen]] for cold-gas [[reaction control system|reaction-control thrusters]].<ref name=nsf20161003/>

===Nozzle===
{{Main|Rocket engine nozzle}}
[[File:Rocket thrust.svg|thumb|right|Rocket thrust is caused by pressures acting in the combustion chamber and nozzle. From Newton's third law, equal and opposite pressures act on the exhaust, and this accelerates it to high speeds.]]
The hot gas produced in the combustion chamber is permitted to escape through a narrow space, called as the throat, to increase the velocity until it reaches Mach 1, and then through a diverging expansion section. When sufficient pressure is provided to the nozzle (about 2.5–3 times ambient pressure), the nozzle ''[[choked flow|choke]]s'' and a supersonic jet is formed, dramatically accelerating the gas, converting most of the thermal energy into kinetic energy. Exhaust speeds vary, depending on the [[expansion ratio]] the nozzle is designed for, but exhaust speeds as high as ten times the [[speed of sound]] in air at sea level are not uncommon. About half of the rocket engine's thrust comes from the unbalanced pressures inside the combustion chamber, and the rest comes from the pressures acting 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.

{{Anchor|opt_expansion}} <!-- add anchor for diagram references --->
[[File:Rocket nozzle expansion.svg|thumb|right|upright|The four expansion regimes of a de Laval nozzle:
• under-expanded
• perfectly expanded
• over-expanded
• grossly over-expanded]]
The most commonly used nozzle is the [[de Laval nozzle]], a fixed geometry nozzle with a high expansion-ratio. The large bell- or cone-shaped nozzle extension beyond the throat gives the rocket engine its characteristic shape.

The exit [[static pressure#Static pressure in fluid dynamics|static pressure]] of the exhaust jet depends on the chamber pressure and the ratio of exit to throat area of the nozzle. As exit pressure varies from the ambient (atmospheric) pressure, a choked nozzle is said to be
* '''under-expanded''' (exit pressure greater than ambient),
* '''perfectly expanded''' (exit pressure equals ambient),
* '''over-expanded''' (exit pressure less than ambient; [[shock diamond]]s form outside the nozzle), or
* '''grossly over-expanded''' (a [[shock wave]] forms inside the nozzle extension).

In practice, perfect expansion is only achievable with a variable–exit-area nozzle (since ambient pressure decreases as altitude increases), and is not possible above a certain altitude as ambient pressure approaches zero. If the nozzle is not perfectly expanded, then loss of efficiency occurs. Grossly over-expanded nozzles lose less efficiency, but can cause mechanical problems with the nozzle. Fixed-area nozzles become progressively more under-expanded as they gain altitude. Almost all de Laval nozzles will be momentarily grossly over-expanded during startup in an atmosphere.<ref name="HuzelAndHuang">{{cite book
 |last = Huzel
 |first = Dexter K.
 |last2 = Huang
 |first2 = David H.
 |date = 1 January 1971
 |title = NASA SP-125, Design of Liquid Propellant Rocket Engines, Second Edition
 |url = https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19710019929_1971019929.pdf
 |publisher = NASA
 |page = <!-- or pages= -->
 |archive-url = https://web.archive.org/web/20170324150551/https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19710019929_1971019929.pdf/
 |archive-date = 24 March 2017
 |url-status = dead
 |access-date = 7 July 2017
 }}</ref>

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 at all altitudes (see diagram).

====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 in a lightweight fashion, although is routinely done with other forms of jet engines. In rocketry a lightweight compromise nozzle is generally used and some reduction in atmospheric performance occurs when used at other than the 'design altitude' or when throttled. 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.

When exhausting into a sufficiently low ambient pressure (vacuum) several issues arise. One is the sheer weight of the nozzle—beyond a certain point, for a particular vehicle, the extra weight of the nozzle outweighs any performance gained. Secondly, as the exhaust gases adiabatically expand within the nozzle they cool, and eventually some of the chemicals can freeze, producing 'snow' within the jet. This causes instabilities in the jet and must be avoided.

On a [[De Laval nozzle]], exhaust gas flow detachment will occur in a grossly over-expanded nozzle.  As the detachment point will not be uniform around the axis of the engine, a side force may be imparted to the engine. This side force may change over time and result in control problems with the launch vehicle.

Advanced [[altitude compensating nozzle|altitude-compensating]] designs, such as the [[aerospike engine|aerospike]] or [[plug nozzle]], attempt to minimize performance losses by adjusting to varying expansion ratio caused by changing altitude.

===Propellant efficiency===
{{See also|Specific impulse}}
[[Image:Nozzle de Laval diagram.svg|thumb|right|upright|Typical temperature (T), pressure (p), and velocity (v) profiles in a de Laval Nozzle]]
For a rocket engine to be propellant efficient, it is important that the maximum pressures possible be created on the walls of the chamber and nozzle by a specific amount of propellant; as this is the source of the thrust. 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 to maximise translational velocity

Since all of these things minimise the mass of the propellant used, and since pressure is proportional to the mass 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 for any given engine, 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. This is termed ''exhaust velocity'', and after allowance is made for factors that can reduce it, the '''[[effective exhaust velocity]]''' is one of the most important parameters of a rocket engine (although weight, cost, ease of manufacture etc. are usually also very important).

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&nbsp;m/s while the speed of sound in the hot gas of a rocket engine can be over 1700&nbsp;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 exit to the area of the throat, 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.

===Thrust vectoring===
{{Main|Thrust vectoring}}
Vehicles typically require the overall thrust to change direction over the length of the burn. A number of different ways to achieve this have been flown:

* The entire engine is mounted on a [[hinge]] or [[gimbal]] and any propellant feeds reach the engine via low pressure flexible pipes or rotary couplings.
* Just the combustion chamber and nozzle is gimballed, the pumps are fixed, and high pressure feeds attach to the engine.
* Multiple engines (often canted at slight angles) are deployed but throttled to give the overall vector that is required, giving only a very small penalty.
* High-temperature vanes protrude into the exhaust and can be tilted to deflect the jet.