Abstract
A key problem in transonic compressor and fan design is that although a 3D description of the flow is necessary to correctly capture the shock; accounting for it during the sectional design is difficult because the key driving design parameters are unknown. In this paper, it is shown that for transonic inlet relative Mach numbers, the pressure rise across the shock is a function of the 3D streamtube area ratio At/A1. This key finding is based on three key transonic flow features, discussed within this paper, being present together across a wide range of 10,000 representative transonic compressor and fan designs published online. This unique wide-ranging dataset reveals that the changes in the shock's behaviour can be explained simply by keeping track of the changes in At/A1. Surprisingly, the pre-shock Mach number at a given At/A1 is shown to be insensitive to the detailed blade surface geometry. Only geometric design choices made in the preliminary design phase are shown to have a second-order effect. These findings suggest, that the purpose of the design should be to make the desired changes in the real spanwise 3D At/A1. The second half of the paper concerns itself with the level of fidelity necessary when calculating the spanwise 3D At/A1 for it to positively influence design. A key conclusion is that not resolving the subtle changes in the 3D radial flow at the appropriate level of fidelity could potentially mislead the transonic design process.
