Abstract
The paper presents the results of an experimental study of the thermal modes of a supersonic combustion chamber at Mach numbers at the channel inlet 3-4, operating on hydrogen. Model is made in the form of a rectangular channel with a combined flame holder with a backward-facing step and wedge-shaped injectors. The experiments were carried out in the mode of a connected pipeline under conditions close to flight ones. It was found that the maximum level of heat flux is reached in the region where the ignition of the mixture is initiated, and this value is achieved during the non-stationary ignition of the combustion chamber. The position of this region shifts upstream with a decrease in the Mach number, which depends on the change in the position of the separation region of shock waves interaction with the boundary layer. With an increase in the Mach number to 4, intense combustion does not occur due to the displacement of this region downstream into the diverging channel and is accompanied by a significant increase in the flow rate. The data obtained show that when developing systems for thermal protection of the combustion chamber, it is necessary to take into account the uneven distribution of heat fluxes along the combustor length in order to prevent local destruction of the walls and possible saving of the coolant resource.
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