In a liquid propellant rocket engine, combustion occurs in a combustion chamber where liquid oxidizer and liquid fuel are injected, mixed, and then ignited. The combustion of liquid propellants results in the generation of hot gases that are expelled through a nozzle to produce thrust. The flame temperature in a liquid propellant rocket engine is a critical parameter influencing engine performance.
The actual flame temperature in a liquid propellant rocket engine may differ from the adiabatic flame temperature due to heat losses, mixing effects, and other factors.
Accurate determination of flame temperature often involves detailed computational simulations and experimental testing tailored to the specific propellant composition and combustion conditions.
The adiabatic flame temperature Tadiabatic is an estimate of the maximum temperature reached during combustion when no heat is lost to the surroundings. The formula for adiabatic flame temperature is:
Where:
- Tadiabatic: Adiabatic flame temperature (in Kelvin).
- : Inlet air temperature (in Kelvin).
- : Fuel temperature (in Kelvin).
- : Specific heat at constant pressure (in J/(kg·K)).
- : Ratio of specific heats ().
- : Specific gas constant (in J/(kg·K)).