Editing Liquid-propellant rocket (section)

==Engine cycles==
For liquid-propellant rockets, four different ways of powering the injection of the propellant into the chamber are in common use.<ref>{{cite web|url= http://www.aero.org/publications/crosslink/winter2004/03_sidebar3.html|title= Sometimes, Smaller is Better|access-date= 2010-06-01|archive-url= https://web.archive.org/web/20120414212704/http://www.aero.org/publications/crosslink/winter2004/03_sidebar3.html|archive-date= 2012-04-14|url-status= dead}}</ref>

Fuel and oxidizer must be pumped into the combustion chamber against the pressure of the hot gasses being burned, and engine power is limited by the rate at which propellant can be pumped into the combustion chamber. For atmospheric or launcher use, high pressure, and thus high power, engine cycles are desirable to minimize [[gravity drag]]. For orbital use, lower power cycles are usually fine.

;[[Pressure-fed cycle (rocket)|Pressure-fed cycle]]: The propellants are forced in from pressurised (relatively heavy) tanks. The heavy tanks mean that a relatively low pressure is optimal, limiting engine power, but all the fuel is burned, allowing high efficiency. The pressurant used is frequently helium due to its lack of reactivity and low density. Examples: [[AJ-10]], used in the Space Shuttle [[Orbital Maneuvering System|OMS]], Apollo [[Service Propulsion System|SPS]], and the second stage of the [[Delta II]].
;[[Electric pump-fed engine|Electric pump-fed]]: An [[electric motor]], generally a [[brushless DC electric motor]], drives the [[pump]]s. The electric motor is powered by a battery pack. It is relatively simple to implement and reduces the complexity of the [[turbomachinery]] design, but at the expense of the extra dry mass of the battery pack. Example engine is the [[Rutherford (rocket engine)|Rutherford]] designed and used by [[Rocket Lab]].
;[[Gas-generator cycle (rocket)|Gas-generator cycle]]: A small percentage of the propellants are burnt in a preburner to power a turbopump and then exhausted through a separate nozzle, or low down on the main one. This results in a reduction in efficiency since the exhaust contributes little or no thrust, but the pump turbines can be very large, allowing for high power engines. Examples: [[Saturn V]]'s [[F-1 engine|F-1]] and [[J-2 engine|J-2]], [[Delta IV]]'s [[RS-68]], [[Ariane 5]]'s [[HM7B]], [[Falcon 9 v1.1|Falcon 9]]'s [[Merlin 1D|Merlin]].
;[[Tap-off cycle]]: Takes hot gases from the main [[combustion chamber]] of the rocket engine and routes them through engine [[turbopump]] turbines to pump propellant, then is exhausted. Since not all propellant flows through the main combustion chamber, the tap-off cycle is considered an open-cycle engine. Examples include the [[J-2S]] and [[BE-3]].
;[[Expander cycle]]: Cryogenic fuel (hydrogen, or methane) is used to cool the walls of the combustion chamber and nozzle. Absorbed heat vaporizes and expands the fuel which is then used to drive the turbopumps before it enters the combustion chamber, allowing for high efficiency, or is bled overboard, allowing for higher power turbopumps. The limited heat available to vaporize the fuel constrains engine power. Examples: [[RL10]] for [[Atlas V]] and Delta IV second stages (closed cycle), [[H-II]]'s [[LE-5]] (bleed cycle).
;[[Staged combustion cycle (rocket)|Staged combustion cycle]]: A fuel- or oxidizer-rich mixture is burned in a preburner and then drives turbopumps, and this high-pressure exhaust is fed directly into the main chamber where the remainder of the fuel or oxidizer undergoes combustion, permitting very high pressures and efficiency. Examples: [[SSME]], [[RD-191]], [[LE-7]].
;[[Staged combustion cycle (rocket)#Full-flow staged combustion cycle|Full-flow staged combustion cycle]]: Fuel- and oxidizer-rich mixtures are burned in separate preburners and driving the turbopumps, then both high-pressure exhausts, one oxygen rich and the other fuel rich, are fed directly into the main chamber where they combine and combust, permitting very high pressures and high efficiency. Example: [[SpaceX Raptor]].

===Engine cycle tradeoffs===
Selecting an engine cycle is one of the earlier steps to rocket engine design.  A number of tradeoffs arise from this selection, some of which include:
{| class="wikitable"
|-
|+ Tradeoff comparison among popular engine cycles
|-
! rowspan=2 |
! colspan=4 | Cycle type
|-
! Gas generator
! Expander cycle
! Staged-combustion
! Pressure-fed
|-
| {{Yes|'''Advantages'''}}
| Simple; low dry mass; allows for high power turbopumps for high thrust
| High specific impulse; fairly low complexity
| High specific impulse; high combustion chamber pressures allowing for high thrust
| Simple; no turbopumps; low dry mass; high specific impulse
|-
| {{No|'''Disadvantages'''}}
| Lower specific impulse
| Must use cryogenic fuel; heat transfer to the fuel limits available power to the turbine and thus engine thrust
| Greatly increased complexity &, therefore, mass (more-so for full-flow)
| Tank pressure limits combustion chamber pressure and thrust; heavy tanks and associated pressurization hardware
|}