An experimental study was conducted in the present study to investigate the effects of upstream probe disturbance on the compressor cascade performance. The experiments were carried out using a plane cascade test facility where the aerodynamic coupling mechanisms between the inner probe and compressor were particularly evaluated in terms of the complex reversed pressure gradient. Moreover, the influence of probe install position near the cascade leading edge on the downstream flow characteristic was evaluated in detail. Results show that the presence of probe at cascade leading edge can reduce the total pressure recovery coefficient, raise the wake loss at cascade outlet and deteriorate the periodicity of conventional cascade flow field distribution. An optimum circumferential installing position at the upstream of the cascade is found with a minimum disturbance, which is not affected by the varying Mach numbers significantly. The probe installed at the intermediate of flow passage has a smaller disturbance on the downstream cascade performance than that installed in front of cascade leading edge.

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Involved in compressor performance testing, an inlet probe support causes flow blockage and produces shedding vortex, which affects aerodynamic performance of the compressor. In this paper a single-stage transonic compressor NASA Stage 35 is taken as an example, and a cylindrical probe support with 10mm diameter is located one chord away from the rotor leading edge. Steady and unsteady flow simulation for the compressor stage with and without the support is carried out by commercial CFX software tool. Results show that the support brings about substantial performance degradation of compressor. Meanwhile, the characteristics of the vortex shedding from the support and the flow field structure of compressor change by the interaction of the support and the compressor. The shedding vortex frequency of probe support is affected by the working condition of the compressor. Under the strong interaction of probe support and transonic rotor leading edge shock, the shedding vortex frequency of probe support is locked by the rotor blade passing. Wider distance between the support and leading edge of the rotor gives rise to homologous change of eddy phase in the rotor.

In the performance test of a compressor, the overstrain alarm of a rotor blade occurred and it was thought to be caused by a large-sized inlet probe. To explain and further avoid the occurrence of this phenomenon, the influences of probe strut configuration on the vibration strain of the compressor rotor blade and the corresponding flow mechanism are studied by using a one-way fluid-structure coupling calculation method. Firstly, the probe strut is simplified as a cylinder with a diameter of 10 mm and located upstream of the inlet stator with the strut wake impinging on the stator blade according to the compressor test. Then, the three scenarios are considered: moving the strut away from the stator blade in axial direction, shifting the strut half of the stator pitch circumferentially, and reducing the strut diameter. The analysis results show that the characteristics of blade vibration are determined by the excitation force on the rotor blade. Under the interaction between the large-sized strut and the compressor, in addition to the excitation force with the strut passing frequency, a force with a lower frequency, namely, the strut-wake-induced frequency, is also observed. The amplitude of the excitation force on the rotor blade depends on the probe configuration. When one of the excitation force frequencies is close to a natural frequency of the rotor blade, the blade resonates, and the amplitude of blade strain varies with the amplitude of the excitation force. In order to reduce the adverse effect of upstream strut wake on the compressor rotor blade vibration, the inlet probe strut should be designed with a smaller diameter and be placed further upstream of the stator in such a way that the strut wake vortex passes through the midpassage of the following stator.