Abstract
This paper presents two optimized rotors. The first rotor is as part of a 3-blade row optimization (IGV-rotor-stator) of a high-pressure compressor. It is based on modifying blade angles and advanced control of curvature of the airfoil camber line. The effects of these advanced blade techniques on the performance of the transonic 1.5-stage compressor were calculated using a 3D Navier-Stokes solver combined with a vortex/vorticity dynamics diagnosis method. The first optimized rotor produces a 3-blade row efficiency improvement over the baseline of 1.45% while also improving stall margin. The throttling range of the compressor is expanded largely because the shock in the rotor tip area is further downstream than that in the baseline case at the operating point. Additionally, optimizing the 3-blade row block while only adjusting the rotor geometry ensures good matching of flow angles allowing the compressor to have more range. The flow diagnostics of the rotor blade based on vortex/vorticity dynamics indicate that the boundary-layer separation behind the shock is verified by on-wall signatures of vorticity and skin-friction vector lines. In addition, azimuthal vorticity and boundary vorticity flux (BVF) are shown to be two vital flow parameters of compressor aerodynamic performance that directly relate to the improved performance of the optimized transonic compressor blade. A second rotor-only optimization is also presented for a 2.9 pressure ratio transonic fan. The objective function is the axial moment based on the BVF. An 88.5% efficiency rotor is produced.
Highlights
The development of modern high thrust-weight ratio gas turbine engines for aircraft requires a compression system with high total pressure ratio and adiabatic efficiency, and needs sufficient stall margin for stable and robust operation
With the total pressure ratio being 2.896 and the adiabatic efficiency reaching 88.49% for the optimum case, the stall margin of the fan rotor is still higher than 10%, which is an inspiring result
An optimization process has been implemented for two different transonic fans
Summary
The development of modern high thrust-weight ratio gas turbine engines for aircraft requires a compression system with high total pressure ratio and adiabatic efficiency, and needs sufficient stall margin for stable and robust operation. An advanced compressor aerodynamic optimization design technique and flow diagnosis method that can modify geometry parametrically, and take into account complex flow physics such as viscous shear, shocks, and vortex structure will aid further development of these compression systems. Based on this motivation, many advanced aerodynamic design concepts and design system have been developed. In the early 1960s, Wang Zhongqi et al [1] [2] [3] [4] proposed the concept and design method of the bowed/lean blade based on the theory of boundary-layer low-energy fluid radial migration. The Air Force Research Labs at Wright-Patterson Air Force Base designed and tested a series of rotors to investigate the effects of swept blade [8] [9]
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