Analysis of the flow in a propulsion nozzle subjected to a fluid injection

The aero-thermodynamics propulsion systems is one of the fields of fluid mechanics where decisive progress is still needed to improve performance and meet the continued demand to orbit increasingly heavy payloads. It's the chemical propulsion that use the thermal energy during combustion of propellants, which remains the most reliable way to meet the requirements of high demands for the thrust, specifically to lift-off. For those of propulsion, the nozzle is the body to convert most effectively heat energy into kinetic energy of the mixture. The divergent portion of a nozzle is the seat of an exhaust gas flow at high enthalpy, where a fraction is transmitted to the walls, thus altering its lifetime. The long periods of operation and amplitude of the heat flux set in the issue, require cooling technique film for the wall of the diverging portion. For weight constraints, it is a fraction of the flow of fuel supplied to the combustion chamber, which is injected into a section of the diverging form of a gaseous film adjacent to the wall. The injection conditions should be adjusted in order to avoid problems associated with shock losses and local re-ignitions. This work represents a numerical investigation of the dynamic and thermal of a compressible flow in the nozzle BKE-DLR, representing a laboratory scale model of the Volvo nozzle of the European Vulcain cryogenic engine 2. The reactants H2/O2 are injected at 50 bar, with a mixture ratio of 6.2 in a cylindrical combustion chamber, wherein the thermodynamic conditions (pressure, temperature) are taken as reservoirs values for the flow relaxation. The axisymmetric nozzle with an exit area ratio of 57 is subjected to a GH2 film cooling in a section where the ratio is equal to 32. Energy & Euler's equations are solved on a non-uniform structured grid, using a finite volume method according to the Roe scheme for convective terms. A zero-dimensional calculation (0D) allows the use of thermodynamic relations to predict aero-thermochemical variables at each section of the divergent, especially in the downstream region of the injection section. The compressible flow in the absence of cooling is numerically reproduced on the axisymmetric model to predict the thrust level and the rate of loss by divergence. The results are compared with values predicted by thermodynamic calculations. The cooling is then carried out for the two cases, sonic and supersonic injection port, wherein a cooling index is defined as a reduced temperature. The distribution of the efficiency index for several downstream locations of the injection zone, is confronted with the measurements on the test bench (DLR). A parametric analysis is also conducted on the effects of the injection angle relative to the wall and the injected flow rate, on the level of thrust and on the cooling efficiency.