Методика аеродинамічного вдосконалення осьових і відцентрових компресорних ступенів газотурбінних двигунів

The development of approaches to aerodynamic improvement of compressor stages of gas turbine engines based on numerical modeling of three-dimensional turbulent gas flows remains a pressing task, which is of great importance for the design and optimization of compressors of aircraft gas turbine engines (GTE). The subject of the study is the aerodynamic processes of flow around the blades of axial stages and the blades of centrifugal wheels of GTE compressors. This study aims to consider the main elements and examples of the application of the aerodynamic improvement technique of axial and centrifugal compressor stages of gas turbine engines. This study aims to test the performance of the proposed cost-effective technique for aerodynamic improvement of axial compressor blade rows and centrifugal compressor stage blades. Methods of solving tasks: methods of mathematical modeling and computational fluid dynamics. The following results were obtained. The main elements of the technique for aerodynamic improvement of axial and centrifugal compressor stages of gas turbine engines based on numerical modeling of spatial turbulent gas flows are presented, which has the following main features: the use of fairly "coarse" calculation grids that retain the sensitivity of the calculation results to changes in the shape of the blades; search for optimal geometric parameters of blades using points of uniformly distributed sequences in the space of variables; formulation of quality criteria as values of the energy characteristics of impellers (air compression ratio and adiabatic efficiency) averaged over air flow. The efficiency of the proposed technique was verified in the aerodynamic improvement of the highly loaded Rotor-37 impeller of the supersonic compressor stage, the guide vanes of the Stage 37 supersonic compressor stages, and the impellers and inlet guide vanes of the centrifugal compressor stages. Conclusions: the proposed technique reduced computational costs in the aerodynamic improvement of blade rows of axial compressors and blades of centrifugal compressor stages and increased the efficiency of gas turbine engine compressors. The obtained results can be used in the aerodynamic optimization of axial and centrifugal compressor stages of aircraft gas turbine engines.