Film cooling
The cooling methods in which liquid propellant or gas forms a thin film on the heated wall to prevent the combustion gas from transferring heat to the wall are called liquid film cooling or air film cooling. The thrust chamber of a liquid rocket engine uses liquid film cooling near the nozzle throat where the heat flux density is the largest. More often, a circle of low-mixture ratio nozzles or fuel direct current perforations are set around the injector to form a low-temperature edge area or film near the inner wall of the combustion chamber to separate the high-temperature combustion gas in the center area from the wall. Air film cooling is often used for turbine blades and flame tubes of turbojet engines, and the blade temperature can be reduced by 400 to 600°C. Some liquid rocket engines introduce turbine exhaust into the thrust chamber nozzle to perform air film cooling on the inner wall of the nozzle.
Sweating cooling
Also known as divergent cooling. The heated parts are made of porous materials, and the coolant seeps out of the heated surface through the micropores to take away the heat, while forming a layer of cooling film on the wall, which has a good cooling effect. Since it is difficult to develop porous materials, carbon deposits after the combustion of hydrocarbon fuels easily clog the micropores. This cooling method has not been widely used and is only used for cooling the injector panel of liquid oxygen-liquid hydrogen engines.
Ablation cooling
It is widely used in the nozzle of solid rocket engines (see ablation heat protection, ablation materials). There is also a self-cooling method, which has a similar mechanism to ablation cooling. Porous tungsten is used as the matrix, and silver, copper, zinc and other materials with low melting points are infiltrated. When heated, the materials with low melting points sublimate and escape to play a cooling role. The matrix tungsten is refractory and corrosion-resistant, and can maintain its shape. It is suitable for making the throat lining of the nozzle of solid rocket engines.
Radiative cooling
Use the thermal radiation of hot objects to dissipate heat outward. Radiative cooling is generally used in the nozzle extension section with low heat flux density in rocket engines, the gas generator with low gas temperature, and the thrust chamber of unit propellant decomposition. The black heat sink of the piston engine cylinder head also plays a role in radiative heat dissipation. Improving the radiation heat dissipation effect mainly relies on the use of high-temperature resistant materials to manufacture heated parts; secondly, it relies on improving the surface blackness, that is, applying high-temperature paint with a blackness of more than 0.85 on the wall surface.
Heat insulation layer
Coating or pasting high-temperature resistant materials with low thermal conductivity on the heated wall surface to reduce the heat transfer of gas to the wall. Commonly used heat insulation materials such as zirconium oxide and aluminum oxide are suitable for liquid rocket engines; graphite, tungsten carbide, ceramics, etc. are suitable for solid rocket engines; ceramics, high-temperature insulation paint, asbestos, etc. are suitable for turbojet engines.