Publicado 6 números por año
ISSN Imprimir: 1948-2590
ISSN En Línea: 1948-2604
ADVANCED HEAT ANALYSIS OF TURBINE ROTOR BLADES COUPLED WITH COMBUSTION CHAMBER SIMULATION
SINOPSIS
Development of an approach to estimating turbine cooled blade heat condition considering 3D and transitional effects is presented. The object of the investigation is the high-pressure turbine first stage blades of an advanced aircraft engine. To set boundary conditions at the investigated turbine stage inlet in detail, the combustor chamber was included in the computational domain. Therefore, combustion processes were also modeled. A fine-grid discretization of the computational domain was applied. This allowed modeling of all perforation holes on the stator vanes and the rotor blades, the cooling air film at end walls, and applying a conjugate heat transfer approach for the rotor blades. To model transition flow features in the boundary layer, the two-equation γ−θ model was chosen. This model allows describing the flow structure in the boundary layer and temperature distribution on the blade surface more accurately in comparison to the ordinary assumption of fully turbulent flow behavior. This numerical approach was applied to estimate the blades heat condition for two different engine operating modes. Experimental data of the investigated blade are presented as temperature values at the given locations on the blade surface. These data were obtained during full-scale testing of the engine core and were collected using two different types of temperature gauges: thermocouples and thermocrystals. The comparison of experimental data and numerical results shows that the developed numerical approach allows obtaining accurate estimation of temperature distribution on the investigated blades.