This paper presents the results of a modal decomposition method applied to the time resolved data of two different test turbines. The analysis is carried out on the measurements performed by fast response aerodynamic pressure probes as well as on CFD simulations. As shown in the earlier aeroacoustic theory, a plurality of rotating patterns, also called spinning modes, are generated by the rotor-stator interactions. The modes may be computed from the flow quantities, such as total pressure, velocity and flow angles, through Fourier decompositions performed in time and space. The deterministic unsteadiness is then simplified to a limited number of Fourier coefficients. At a fixed radial position, circumferential lobes are identified for any multiple of the blade passing frequency. Therefore, the flow may be described as the superposition of rotating patterns, the spatial characteristics of which are correlated to the linear combinations of blade/vane number. This analysis has been applied to a one and a half stage low pressure turbine and to a two-stage counter-rotating transonic turbine. In the former test case there is a limited number of modes that characterize the flow field. Hence, the decomposition in modes simplifies considerably the evaluation of the sources of unsteadiness and deterministic stresses. The second test case presents more complex interactions. In fact, the presence of two rotors induces oscillations at frequencies that corresponds to the linear combinations of the two blade passing frequencies. Circumferential modes are identified for the most characteristic frequencies and their physical meaning is discussed.
Within previous EU projects, possible modifications to the engine architecture have been investigated, that would allow for an optimised aerodynamic or acoustic design of the exit guide vanes (EGV) of the turbine exit casing (TEC). However, the engine weight should not be increased and the aerodynamic performance must be at least the same. This paper compares a state-of-the art TEC (reference TEC) with typical EGVs with an acoustically optimised TEC configuration for the engine operating point approach. It is shown that a reduction in sound power level for the fundamental tone (1 blade passing frequency) for this acoustically important operating point can be achieved. It is also shown that the weight of the acoustically optimised EGVs (only bladings considered) is almost equal to the Reference TEC, but a reduction in engine length can be achieved. Measurements were conducted in the subsonic test turbine facility (STTF) at the Institute for Thermal Turbomachinery and Machine Dynamics, Graz University of Technology. The inlet guide vanes, the low pressure turbine (LPT) stage, and the EGVs have been designed by MTU Aero Engines.
This paper presents the effects of an unshrouded low pressure turbine (LPT) onto the following exit guide vane row (EGV). The measurement results were obtained in the subsonic test turbine facility at Graz University of Technology by means of a fast response pressure probe in planes downstream of the rotor as well as oil flow visualisation. The test rig was designed in cooperation with MTU Aero Engines and represents the last 1.5 stages of a commercial aero engine. Considerable efforts were put into the adjustment of all relevant model parameters to reproduce the full scale LPT situation. Different tip clearances were evaluated by means of CFD obtained using a commercial Navier-Stokes code and validated with experimental results. The goal is to evaluate the effect of the varying leakage flow on the flow in the low aspect ratio EGV. Special attention is given to the impact on the development of secondary flows as well as the flow structures downstream of the EGV. The effect of the leakage flow causes a change of the flow structure of the EGV, particularly losses. Considering the largest investigated tip-clearance, the losses increased by 71% when compared to a zero-leakage case.
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