An improved blade passage model for estimating off-design axial compressor performance

[摘要] Accurate estimates of multistage axial compressor performance at off-design operating conditions are essential to the determination of key performance metrics of aircraft gas turbine engines, such as fuel burn, thrust output, and stable operating range. However, conventional RANS based CFD calculations of multistage axial compressors diverge at off-design operating conditions where large separation occurs and the stages are mismatched. This thesis demonstrates the feasibility of a body force based approach to capturing the three-dimensional flow field through a turbomachinery blade row at off-design conditions. A first principles based blade passage model is introduced which addresses the limitations of previous approaches. The inputs to the improved blade passage model are determined from three-dimensional, steady, single-passage RANS CFD calculations. In a first step towards modeling multistage configurations, the improved blade passage model is validated using a fan rotor test case. At the design operating conditions, the stagnation pressure rise coefficient and the work coefficient are both estimated within 5%, and the adiabatic efficiency is estimated within 1 percentage point over most of the span relative to single-passage RANS CFD simulations. At low mass flow operating conditions, where the single-passage RANS CFD diverges, the blade passage model and related body force representation are capable of computing the three-dimensional throughflow with separation and reversed flow. These results pave the way for future unsteady calculations to assess compressor stability and for multistage compressor simulations at off-design conditions.

[发布机构] Massachusetts Institute of Technology