A Rocket Engine is a type of jet engine that uses only stored rocket propellant mass for forming its high-speed propulsive jet. Rocket engines are reaction engines obtaining thrust in accordance with Newton’s third law of motion. Most rocket engines are internal combustion engines although non-combusting forms (such as cold gas thrusters) also exist.
In Rocket Engine, the heat created in the combustion chamber during combustion process is about 2800K-3500K, which contains exhaust gases.Most of this heat is expelled along with the gas that contains it; however, heat is transferred to the thrust chamber walls. Due to this high temperature, there will be an existence of damaging the walls of the thrust chamber and nozzle. cooling system cooling system cooling system cooling system cooling system
To hold such high temperature effectively, the Regenerative Cooling method is used. Regenerative cooling is the most widely used method of cooling a thrust chamber and is accomplished by flowing high-velocity coolant over the back side of the chamber wall to cool the hot gas liner. In this method, the fuel itself acts as a coolant because in liquids huge amount of heat transfer takes place quickly as compared to air or other gases. The coolant with the heat input from cooling the liner is then discharged into the injector and utilized as a propellant.
Earlier thrust chamber designs had low chamber pressure, low heat flux and low coolant pressure requirements, which could be satisfied by a simplified “double wall chamber” design with regenerative and film cooling. Later, chamber pressures were increased and the cooling requirements became more difficult to satisfy.
This led to the design of “tubular wall” thrust chambers, by far the most widely used design approach for the vast majority of large rocket engine applications even in Air Force and NASA. The primary advantage is its light weight and the large experience base. But as chamber pressures and hot gas wall heat fluxes continued to increase resulted in more effective methods.
One solution has been “channel wall” thrust chambers, the hot gas wall cooling is accomplished by flowing coolant through rectangular channels, which are machined into a hot gas liner fabricated from a high-conductivity material, such as copper or a copper alloy. Heat transfer and structural characteristics are excellent in this case.
Basically, there are three domains in a regeneratively cooled rocket engine-
1. Gas Domain (Combusted Gases) – Convection and Radiation heat transfer
2. Liquid Domain (Coolant) – Convection heat transfer
3. Solid Domain (Thrust chamber wall) – Conduction heat transfer
Heat transfer from the outer surface of thrust chamber to the environment can be neglected and the outer surface wall is assumed as adiabatic.
In addition to the regeneratively cooled designs mentioned above, other thrust chamber designs have been fabricated for rocket engines using dump cooling, film cooling, transpiration cooling, ablative liners and radiation cooling. Although regeneratively cooled combustion chambers have proven to be the best approach for cooling large liquid rocket engines, other methods of cooling have also been successfully used for cooling thrust chamber assemblies.
Dump cooling, which is similar to regenerative cooling because the coolant flows through small passages over the back side of the thrust chamber wall. The difference, however, is that after cooling the thrust chamber, the coolant is discharged overboard through openings at the aft end of the divergent nozzle. This method has limited application because of the performance loss resulting from dumping the coolant overboard.
Film cooling provides protection from excessive heat by introducing a thin film of coolant or propellant through orifices around the injector periphery or through manifold orifices in the chamber wall near the injector or chamber throat region. This method is typically used in high heat flux regions and in combination with regenerative cooling.
Transpiration cooling provides coolant (either gaseous or liquid propellant) through a porous chamber wall at a rate sufficient to maintain the chamber hot gas wall to the desired temperature. The technique is really a special case of film cooling.
With radiation cooling, heat is radiated from the outer surface of the combustion chamber or nozzle extension wall. Radiation cooling is typically used for small thrust chambers with a high-temperature wall material (refractory) and in low-heat flux regions, such as a nozzle extension.