Thrust-chamber cooling Because of high combustion temperatures and high heat-transfer rates from the hot gases to the chamber wall, thrust-chamber cooling will be a major design consideration. For non-reusable rocket engines uncooled chambers may sometimes be used; the heat will be absorbed by a heavy chamber wall acting as a heat sink, until the wall temperature approaches the failure level; but for a reusable engine a steady-state chamber cooling system must be employed. Here are some chamber-cooling techniques typically used in rocket industry:  Regenerative cooling. Regenerative cooling, the most widely applied method, utilizes one or possibly both of the propellants fed through passages in the thrust-chamber wall for cooling, before being injected into the combustion chamber.  Dump cooling. With this principle, a small percentage of the propellant, such as the hydrogen in a LO2/LH2 engine, is fed through passages in the thrust chamber wall for cooling and is subsequently dumped overboard through openings at the rear end of the nozzle skirt. Because of inherent problems,such as performance losses, this method has only limited application.  Film cooling. Here, exposed chamber-wall surfaces are protected from excessive heat by a thin film of coolant or propellant introduced through orifices around the injector periphery or through manifolded orifices in the chamber wall near the injector and sometimes in several more planes toward the throat. The method has been used, particularly for high heat fluxes, either alone or in combination with regenerative cooling.  Transpiration cooling. Transpiration cooling introduces a coolant (either gaseous or liquid propellant) through porous chamber walls at a rate sufficient to maintain the desired temperature of the combustion- gasside chamber wall. This method is essentially a special type of film cooling.  Ablative cooling. In this process, combustion-gas-side wall material is sacrificed by melting, vaporization, and chemical changes to dissipate heat. As a result, relatively cool gases flow over the wall surface, thus lowering the boundary-layer temperature and assisting the cooling process. In addition, the ablative material is usually a good thermal insulator, keeping to a minimum the heat transmitted to the outer structure.  Radiation cooling. With this method, heat is radiated away from the surface of the outer thrust-chamber wall. It has been successfully applied to very small, high-temperaturematerial combustion chambers and to low heat- flux regions, such as nozzle extensions. Selection of the best cooling method for a given thrust chamber depends on many considerations. There are no simple-and-fast rules. However, the main factors that influence the selected design approaches will be the following:  Propellants. The properties of the combustion products, such as temperature, specific heat, specific weight, viscosity, etc., have a direct bearing on the heat-transfer rate and thus affect chamber cooling requirements and methods. The properties and flowrates of the propellants determine whether they are suitable or sufficient for regenerative, transpiration, dump, or film cooling. Consequently, the propellants involved will be a primary consideration in the design of a chamber cooling system.  Chamber pressure. Higher chamber pressures are linked with higher combustion-gas mass flowrates per unit area of chamber cross section and therefore with higher heat transfer rates. Regenerative- and film- cooling methods are usually combined to meet the stringent requirements of high-chamber pressure applications.