Abstract
An experimental investigation was carried out to study the effect of the boundary layer transition on the flow dynamics in the blade passage of a compressor cascade. A model of a turbine compressor passage was designed and assembled in a transonic wind tunnel for this purpose. Two different flow control methods were used in the experiment to induce the transition upstream of the shock wave, one concerning the microstep application and the other using distributed roughness strips. Two locations of spanwise microsteps for the transition trigger were chosen, one at the leading edge and the other closer to the shock wave position. The distributed roughness in the form of standard sandpaper strips with different heights was applied in three various locations on the blade upstream of the shock. The main objective of investigations is to present the influence of the laminar and turbulent shock wave-boundary layer interaction on the flow dynamics in a compressor fan passage, and the specific objective is to show how the parameters of particular transition control methods affect the flow dynamics in the investigated channel. The major challenge for this research was to minimize the disturbance caused by the microstep or roughness to the laminar boundary layer, while still ensuring a successful transition. Very interesting results were obtained in the flow control application for the boundary layer transition control, demonstrating a positive effect on the flow unsteadiness in changing the nature of the interaction.
Highlights
In the case of a civil turbofan engine operating at high altitudes, the Reynolds number can drop by a factor of 4 in comparison to the sea level values. e laminar boundary layer on the transonic compressor rotor blades will interact with shock waves, and a strong boundary layer separation will appear as a result. is can seriously affect the aero-engine performance and operation
When the transition is provoked too far upstream, a large part of the beneficial laminar flow region is lost. e main challenge is to gain knowledge of where the transition should be induced with respect to the mitigation of unsteadiness. is knowledge will enable the implementation of an effective laminar flow technology for engines in which the interaction of a laminar boundary layer with a shock wave takes place and causes severe problems
Different tripping devices may be used to induce the transition in the boundary layer; the presented experiments focused on the effects of roughness and microsteps on the unsteady aerodynamic phenomena in a compressor fan blade passage
Summary
Unsteady aerodynamic loads for the structure lead to the fatigue of the material [10, 11]. e sources of such unsteadiness come from the compressibility effects at high speed, and they are sources of flow-induced vibrations, crucial for the stability of components. E main objective of the paper is to present an experimental investigation concerning the influence of the laminar and turbulent shock wave-boundary layer interaction on the flow dynamics, i.e., on unsteadiness in a compressor fan passage. Different tripping devices may be used to induce the transition in the boundary layer; the presented experiments focused on the effects of roughness and microsteps on the unsteady aerodynamic phenomena in a compressor fan blade passage. For this purpose, a model of a compressor passage was designed and assembled in a wind tunnel [19] (at the Institute in Gdansk) in order to investigate the flow structure on the suction side of the blade. An experimental program was carried out in a transonic wind tunnel in the test sections shown in Figure 2. e flow structure was investigated in the model of a compressor blade passage. e lower and upper profiles were located in a similar configuration as in the reference linear cascade (delivered by the industrial partner—Rolls Royce) in order to keep a similar flow structure as in the reference one. ere was a nozzle with a uniform velocity region at its outlet upstream of the blade passage. e Mach number upstream of the blades was supersonic, M 1.22, and the Reynolds
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