Full Annulus Simulations of a Transonic Axial Compressor Stage with Distorted Inflow at Transonic and Subsonic Blade Tip Speed

Abstract

It is reported on systematic numerical studies on an axial compressor stage with distorted inflow. Four operating points at two speedlines have been simulated with an inflow distortion generated by a 120• -sector segment with a wedge type cross section. With this setup the interaction between the distorted inflow and the rotor flow is studied. The focus is put on the differences of the interaction between the distorted inflow and rotor flow as well as on the compressor behaviour at subsonic and transonic blade tip speed, as the general mechanisms are analyzed in more detail in previous publications. The distorted flow itself is not influenced by the blade tip speed but the interaction phenomena depend strongly on the spool speed and operating point. Also, the blade tip speed influences the circumferential sector of the compressor stage exit which is affected by the distorted flow. The impact reaches from a small sector at 65% spool speed, peak efficiency operating point up to nearly the entire annulus at 100% spool speed, near stall operating point. KEYWORDS INTRODUCTIONThe design process of new aircraft engines heavily depends on reliable simulation environments, due to demanding and ambitious design goals as defined for instance by the ACARE group in Europe by the European Commission (2011). The requirements on aircraft concerning lower specific fuel consumption and less noise emissions are challenging. Today's modern compressors need to provide higher pressure ratios to achieve higher thermal efficiency while the number of turbomachinery stages and the weight is reduced by the usage of lightweight materials. To obtain this goal with less turbomachinery stages the remaining stages become higher loaded which makes the compressor stages prone to deviations from their designated operating range. Flow distortions induced by the integration and the intake of the jet engine also impact the performance and surge margin of the compressor as investigated by Schnell et al. (2015). As the deviations from the designated operating range are likely to occur in flight situations like high angle of attack or strong crosswinds, the compressor stages have to withstand them without significant loss in surge margin for safety reasons. As the surge margin of compressors decreases with inlet distortions, investigations on the behavior of compressors with distorted inflow are still necessary. Hah et al. (1998) have shown the loss in surge margin in their experimental and numerical studies on the effects of an inflow distortion on an axial compressor stage. While their approach with distortion screens and modified numerical boundary conditions gives insight into the reaction of the compressor on the inflow distortion, the upstream influence of the compressor on the distorted flow cannot be investigated. Requirements for the design process of low pressure compressors and fans are simulation tools capable of resolving strong secondary flow phenomena within turbomachines, the interaction between secondary flow phenomena and inf...