When the high-pressure turbine guide blade works in the aero-engine, it is in the complex gas-heat environment of the combustion chamber and the turbine. The temperature distribution and strength of the blade affect the reliability and life of the aero-engine. In order to more accurately analyze the complex flow field aero-thermal parameters between the aero-engine combustion chamber and the turbine, this paper considers the influence of the non-uniform aero-thermal parameters of the combustion chamber outlet on the turbine. Taking the first-stage guide vane of an aero-engine turbine as the research object, the high and low temperature cycle thermal shock test was carried out to obtain the crack location and strength of the turbine guide vane, which provided a certain reference for improving the design of the turbine guide vane. Through the test, it is found that after 750 thermal shock cycles, fatigue cracks appear in four concentrated areas of the blade, which are distributed in the leading edge and trailing edge of the blade respectively. This shows that the temperature at the leading edge and trailing edge of the blade is higher, the heat load is higher, and the risk of damage and ablation is also the highest. Based on the experiment, the cross-component model of the combustion chamber and the blade is established. On this basis, four kinds of swirler structures at the head of the combustion chamber are designed. ANSYS FLUENT software is used to study the unsteady parameters such as the surface temperature of the guide vane under the influence of strong swirling flow. The numerical results show that when the rotation direction of the cyclone is counterclockwise, the temperature of the guide vane wall is more uniform. Compared with the clockwise rotation of the cyclone, the temperature is reduced by about 5 %, and the installation angle of the 50° cyclone blade is better than the installation angle of the 45° blade.