In a solid propellant rocket engine, combustion occurs in a confined space within the rocket motor, where a solid propellant undergoes combustion to produce hot gases that are expelled through a nozzle to generate thrust. The flame temperature in a solid propellant rocket engine is a crucial parameter influencing the engine’s performance. However, determining the flame temperature in a solid rocket motor can be more complex than in liquid or hybrid rocket engines due to the nature of the solid propellant combustion.
Solid propellants consist of a mixture of oxidizers, fuels, and binders in a solid form. The combustion of solid propellants involves the rapid decomposition of the propellant material, releasing hot gases that contribute to thrust.
Experimental testing and detailed computational simulations are often employed to characterize and understand the combustion behavior, temperature distribution, and performance of solid propellant rocket motors.
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)).