• 한국추진공학회:학술대회논문집
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• 한국추진공학회 2004년도 제22회 춘계학술대회논문집
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• pp.75-78
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• 2004

Various jet engines (Turbine engine family and RAM Jet engine) have been developed for high speed aircrafts. but their application to hypersonic flight is restricted by principle problems such as increase of total pressure loss and thermal stress. Therefore, the development of next generation propulsion system for hypersonic aircraft is a very important subject in the aerospace engineering field, SCRAM Jet engine based on a key technology, Supersonic Combustion. is supposed as the best choice for the hypersonic flight. Since Supersonic Combustion requires both rapid ignition and stable flame holding within supersonic air stream, much attention have to be given on the mixing state between air stream and fuel flow. However. the wider diffusion of fuel is expected with less total pressure loss in the supersonic air stream. So. in this study the direction of fuel injection is inclined 30 degree to downstream and the total pressure of jet is controlled for lower penetration height than thickness of boundary layer. Under these flow configuration both streams, fuel and supersonic air stream, would not mix enough. To spread fuel wider into supersonic air an aerodynamic force, baroclinic torque, is adopted. Baroclinic torque is generated by a spatial misalignment between pressure gradient (shock wave plane) and density gradient (mixing layer). A wedge is installed in downstream of injector orifice to induce an oblique shock. The schlieren optical visualization from side transparent wall and the total pressure measurement at exit cross section of combustor estimate how mixing is enhanced by the incidence of shock wave into supersonic boundary layer composed by fuel and air. In this study non-combustionable helium gas is injected with total pressure 0.66㎫ instead of flammable fuel to clarify mixing process. Mach number 1.8. total pressure O.5㎫, total temperature 288K are set up for supersonic air stream.

• 한국유체기계학회 논문집
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• 제18권2호
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• pp.60-66
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• 2015

Conjugate heat transfer analysis was performed to investigate the flow and cooling performance of the high pressure turbine nozzle of gas turbine engine. The CHT code was verified by comparison between CFD results and experimental results of C3X vane. The combination of k-${\omega}$ based SST turbulence model and transition model was used to solve the flow and thermal field of the fluid zone and the material property of CMSX-4 was applied to the solid zone. The turbine nozzle has two internal cooling channels and each channel has a complex cooling configurations, such as the film cooling, jet impingement, pedestal and rib turbulator. The parabolic temperature profile was given to the inlet condition of the nozzle to simulate the combustor exit condition. The flow characteristics were analyzed by comparing with uncooled nozzle vane. The Mach number around the vane increased due to the increase of coolant mass flow flowed in the main flow passage. The maximum cooling effectiveness (91 %) at the vane surface is located in the middle of pressure side which is effected by the film cooling and the rib turbulrator. The region of the minimum cooling effectiveness (44.8 %) was positioned at the leading edge. And the results show that the TBC layer increases the average cooling effectiveness up to 18 %.