| In modern gas turbines,the gas temperature at the inlet of high-pressure turbines obviously exceeds the metal melting point of turbine components.In order to improve the cooling scheme of gas turbines and ensure the service life of hot gas path components,accurate prediction of the heat load distribution on the hub and end wall of turbine blades is becoming more and more important.In this harsh environment,gas film cooling and thermal barrier coating technologies are key technologies to ensure that turbine blades bear high-temperature gas.However,the temperature at the actual outlet of gas turbine has great inhomogeneity in circumferential and radial direction,that is,there are hot streaks.Secondly,there are strong residual swirling currents at the outlet.The simultaneous existence of hot streaks and swirling currents has a significant impact on the aerodynamic thermal performance of blade surface.Therefore,the effects of hot spots in different circumferential directions and hot spot coupled swirling flow on the surface cooling performance of metal blades and coated blades under rotor-stator interference were studied in this paper.The main research of this paper is as follows:(1)A geometric model of high temperature gas turbine cascade was established,and unsteady numerical simulation method was used to analyze the influence mechanism of different circumferential hot spots and hot spots coupled with positive and negative swirl on unsteady flow heat transfer characteristics in cascade channels within a stator period.The migration and decline of hot spots in different circumferential positions in the flow passage were analyzed.(2)The gas-thermal characteristics of the leading edge of the stator blade and the center of the flow channel of the hot spot in yeshan channel were compared and analyzed,and the migration law of the hot spot at different time points and circumferential positions was obtained.It is found that at 1/4t time point,the affected area of hot spots is mainly located at the suction side of the leading edge of the blade.At 2/4t,the affected area of the hot spot turns to the pressure side of the leading edge.As it moves towards the rotor blade,it is divided into hot fluid masses.(3)Under different circumferential hot spot inlet conditions,the comprehensive influence factors of hot spot and rotor-stator interference on the cooling performance of guide vane and rotor blade surface were studied,and the heat flux density and cooling efficiency of rotor blade surface were compared with that of thermal barrier coating,and the improvement situation of blade cooling performance was obtained.It is found that the influence of hot spots and unsteady wakes on the cooling performance of rotor blades is obviously changed by the swirl at the inlet of the rotor blades.Especially on the suction side near the gas film hole.Combined with the timing effect between rotor and stator blades,four kinds of inlet hot spots with relative rotor and stator positions and two kinds of circumferential positions were selected to study the comprehensive influence on the cooling performance of turbine blades with thermal barrier coatings.This study can provide guidance for the thermal design of turbine blades under the condition of hot spot inlet,so as to maximize the insulation characteristics of thermal barrier coatings. |
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