Editing Rocket engine (section)

==Cooling==
For efficiency reasons, higher temperatures are desirable, but materials lose their strength if the temperature becomes too high. Rockets run with combustion temperatures that can reach {{cvt|6,000|F|C K|-2}}.<ref name="HuzelAndHuang" />{{rp|98}}

Most other jet engines have gas turbines in the hot exhaust. Due to their larger surface area, they are harder to cool and hence there is a need to run the combustion processes at much lower temperatures, losing efficiency. In addition, [[wiktionary:duct engine|duct engines]] use air as an oxidant, which contains 78% largely unreactive nitrogen, which dilutes the reaction and lowers the temperatures.<ref name="Sutton" /> Rockets have none of these inherent combustion temperature limiters.

The temperatures reached by combustion in rocket engines often substantially exceed the melting points of the nozzle and combustion chamber materials (about 1,200 K for [[copper]]). Most construction materials will also combust if exposed to high temperature oxidiser, which leads to a number of design challenges. The nozzle and combustion chamber walls must not be allowed to combust, melt, or vaporize (sometimes facetiously termed an "engine-rich exhaust").

Rockets that use common construction materials such as aluminium, steel, nickel or copper alloys must employ cooling systems to limit the temperatures that engine structures experience. [[Regenerative cooling (rocket)|Regenerative cooling]], where the propellant is passed through tubes around the combustion chamber or nozzle, and other techniques, such as 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.

Material exceptions that can sustain rocket combustion temperatures to a certain degree are [[Reinforced carbon–carbon|carbon–carbon materials]] and [[rhenium]], although both are subject to oxidation under certain conditions. Other [[refractory]] alloys, such as alumina, [[molybdenum]], [[tantalum]] or [[tungsten]] have been tried, but were given up on due to various issues.<ref name="RocketProp8">{{cite book |author=George P. Sutton |url=https://archive.org/details/Rocket_Propulsion_Elements_8th_Edition_by_Oscar_Biblarz_George_P._Sutton/page/308/mode/2up |title=Rocket Propulsion Elements |author2=Oscar Biblarz |date=2010 |publisher=Wiley Interscience |isbn=9780470080245 |edition=8th |page=308 |name-list-style=amp}}</ref>

Materials technology, combined with the engine design, is a limiting factor in chemical rockets.

In rockets, the [[heat flux]]es that can pass through the wall are among the highest in engineering; fluxes are generally in the range of 0.8–80 MW/m{{sup|2}} (0.5-50 [[BTU]]/in{{sup|2}}-sec).<ref name="HuzelAndHuang" />{{rp|98}} 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 although lower than the chamber, the high temperatures seen there. (See {{section link||Nozzle}} above for temperatures in nozzle).

In rockets the coolant methods include:<ref name="HuzelAndHuang" />{{rp|98–99}}

#[[ablation|Ablative]]: The combustion chamber inside walls are lined with a material that traps heat and carries it away with the exhaust as it vaporizes.
#[[Radiative cooling]]: The engine is made of one or several [[refractory]] materials, which take heat flux until its outer thrust chamber wall glows red- or white-hot, radiating the heat away.
#Dump cooling: A cryogenic propellant, usually [[hydrogen]], is passed around the nozzle and dumped. This cooling method has various issues, such as wasting propellant. It is only used rarely.
#[[regenerative cooling (rocket)|Regenerative cooling]]: The fuel (and possibly, the oxidiser) of a [[liquid rocket engine]] is routed around the nozzle before being injected into the combustion chamber or preburner. This is the most widely applied method of rocket engine cooling.
#Film cooling: The engine is designed with rows of multiple orifices lining the inside wall through which additional propellant is injected, cooling the chamber wall as it evaporates. This method is often used in cases where the heat fluxes are especially high, likely in combination with [[regenerative cooling (rocket)|regenerative cooling]]. A more efficient subtype of film cooling is [[transpiration cooling]], in which propellant passes through a [[porous]] inner combustion chamber wall and transpirates. So far, this method has not seen usage due to various issues with this concept.

Rocket engines may also use several cooling methods. Examples:

* Regeneratively and film cooled combustion chamber and nozzle: [[V-2 rocket|V-2]] Rocket Engine<ref>{{cite web |title=Raketenmotor der A4 (V2)-Rakete |url=https://www.deutsches-museum.de/flugwerft-schleissheim/ausstellung/flugantriebe-und-raketen/raketenmotor-a-4 |access-date=19 September 2022 |language=de |quote=An additional coolant line takes alcohol to fine holes in the inner chamber wall. The alcohol flows alongside the wall, creating a thin, evaporating film for additional cooling.}}</ref>
* Regeneratively cooled combustion chamber with a film cooled nozzle extension: [[Rocketdyne F-1|Rocketdyne F-1 Engine]]<ref>{{cite web |author=McCutcheon, Kimble D. |date=3 August 2022 |title=U.S. Manned Rocket Propulsion Evolution Part 8.12: Rocketdyne F-1 Engine Description |url=https://www.enginehistory.org/Rockets/RPE08.11/RPE08.12.shtml |access-date=19 September 2022}}</ref>
* Regeneratively cooled combustion chamber with an ablatively cooled nozzle extension: The [[LR-91]] rocket engine<ref>{{cite web |author=McCutcheon, Kimble D. |date=3 August 2022 |title=U.S. Manned Rocket Propulsion Evolution Part 6: The Titan Missile |url=https://www.enginehistory.org/Rockets/RPE06/RPE06.shtml |access-date=19 September 2022}}</ref>
* Ablatively and film cooled combustion chamber with a radiatively cooled nozzle extension: [[Lunar module descent engine]] (LMDE), [[Apollo command and service module#Service propulsion system|Service propulsion system engine]] (SPS)<ref>{{cite book |last=Bartlett |first=W. |url=https://ntrs.nasa.gov/api/citations/19700026405/downloads/19700026405.pdf |title=Apollo spacecraft liquid primary propulsion systems |last2=Kirkland |first2=Z. D. |last3=Polifka |first3=R. W. |last4=Smithson |first4=J. C. |last5=Spencer |first5=G. L. |date=7 February 1966 |publisher=NASA, Lyndon B. Johnson Space Center |location=Houston, TX |pages=8 |archive-url=https://web.archive.org/web/20220823092501/https://ntrs.nasa.gov/api/citations/19700026405/downloads/19700026405.pdf |archive-date=23 August 2022 |access-date=10 September 2022 |url-status=bot: unknown }}</ref>
* Radiatively and film cooled combustion chamber with a radiatively cooled nozzle extension: [[R-4D]] storable propellant thrusters<ref name="RocketProp8" />

In all cases, another effect that aids in cooling the rocket engine chamber wall is a thin layer of combustion gases (a [[boundary layer]]) that is notably cooler than the combustion temperature. 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.<ref name="HuzelAndHuang" />{{rp|105–106}}

Liquid-fuelled engines are often run [[Air-fuel ratio|fuel-rich]], which lowers combustion temperatures.  This reduces heat loads on the engine and allows lower cost materials and a simplified cooling system. This can also ''increase'' performance by lowering the average molecular weight of the exhaust and increasing the efficiency with which combustion heat is converted to kinetic exhaust energy.