Experimental investigations on a single stage centrifugal compressor showed that measured blade vibration amplitudes vary considerably along a constant speed line from choke to surge. The unsteady flow has been analyzed to obtain detailed insight into the excitation mechanism. Therefore, a turbocharger compressor stage impeller has been modeled and simulated by means of computational fluid dynamics (CFD). Two operating points at off-design conditions were analyzed. One was close to choke and the second one close to the surge line. Transient CFD was employed, since only then a meaningful prediction of the blade excitation, caused by the unsteady flow situation, can be expected. Actually, it was observed that close to surge a steady state solution could not be obtained; only transient CFD could deliver a converged solution. The CFD results show the effect of the interaction between the inducer casing bleed system and the main flow. Additionally, the effect of the nonaxisymmetric components, such as the suction elbow and the discharge volute, was analyzed. The volute geometry itself had not been modeled. It turned out to be sufficient to impose a circumferentially asymmetric pressure distribution at the exit of the vaned diffuser to simulate the volute. Volute and suction elbow impose a circumferentially asymmetric flow field, which induces blade excitation. To understand the excitation mechanism, which causes the measured vibration behavior of the impeller, the time dependent pressure distribution on the impeller blades was transformed into the frequency domain by Fourier decomposition. The complex modal pressure data were imposed on the structure that was modeled by finite element methods (FEM). Following state-of-the-art calculations to analyze the free vibration behavior of the impeller, forced response calculations were carried out. Comparisons with the experimental results demonstrate that this employed methodology is capable of predicting the impeller’s vibration behavior under real engine conditions. Integrating the procedure into the design of centrifugal compressors will enhance the quality of the design process.
Underplatform friction dampers are possible solutions for minimizing vibrations of rotating turbine blades. Solid dampers, characterized by their compact dimensions, are frequently used in real applications and often appear in patents in different forms. A different type of the friction damper is a thin-walled structure, which has larger dimensions and smaller contact stresses on a wider contact area in relation to the solid damper. The damping performance of a thin-walled damper, mounted under the platforms of two rotating, freestanding high pressure turbine blades, is investigated numerically and experimentally in this paper. The tangential and normal contact stiffness that are crucial parameters in optimal design of any friction damper are determined from three-dimensional finite element computations of the contact behavior of the damper on the platform including friction and centrifugal effects. The computed contact stiffness values are applied to nonlinear dynamic simulations of the analyzed blades coupled by the friction damper of a specified mass. These numerical analyses are performed in the modal frequency domain, which is based on the harmonic balance method for the complex linearization of friction forces. The numerical dynamic results are in good agreement with the measured data of the real mistuned system. In the analyzed excitation range, the numerical performance curve of the thin-walled damper is obtained within the scatter band of the experimental results. For the known friction coefficients and available finite element and harmonic balance tools, the described numerical process confirms its usability in the design process of turbine blades with underplatform dampers.
The increase of new gas turbine’s efficiency is connected with further rise of turbine inlet temperature and sometimes as well pressure. In these conditions, first cooled turbine stages of a gas turbine engine usually consist of freestanding airfoils, which do not use an integrated shroud, to avoid risk of shroud overheating. In order to better control the radial gap leakage flow between the rotating blade tip and turbine casing, special design features of the airfoil tip need to be considered in the design process to meet the best possible stage performance. In the general engineering practice, a blade tip squealer provides opportunities to control tip clearance loss. In this paper several simplified types of the tip squealer design are investigated to determine the most effective loss control. At this stage of the investigation, blade tip cooling was not taken into account, but aerodynamic effects were analysed in detail. Based on the most common designs of the blade tip in the literature, four geometry types were investigated: (i) a flat tip design as the reference baseline solution, (ii) full tip squealer, (iii) partial squealer along the pressure side (PS) wall with a cut-out at the pressure side near the trailing edge (TE) and (iv) partial squealer along the suction side (SS) wall with a cut-out at the suction side near TE. All these cases have been compared among each other for two relative radial gaps (gap to blade height) of 0.6% and 1.36%. The flow calculations were done with a full 3D Navier-Stokes CFD code. For the flat tip and for full squealer designs, numerical results were validated against well-known experimental data measured on the GE-E3 blade cascade test rig found in the open literature. By using the 3D numerical data, the special attention was considered to confirm reliability and predictive credibility of the blade tip flow obtained from the analytical model. The obtained loss values and flow details were compared for all studied cases, providing insight into turbine stage aerodynamics with respect to minimal and maximal radial clearance.
This paper describes an approach to unidirectional coupled CFD–FEM analysis developed at ABB Turbo Systems Ltd. Results of numerical investigations concerning the vibration behavior of an axial turbocharger turbine are presented. To predict the excitation forces acting on the rotating blades, the time-resolved two-dimensional coupled stator–rotor flow field of the turbine stage was calculated. The unsteady pressure, imposed on the airfoil contour, leads to circumferentially nonuniform and pulsating excitation forces acting on the rotating bladed disk. A harmonic transformation of the excitation forces into the rotating coordinate system of a single blade was elaborated and the complex Fourier amplitudes were determined. The bladed rotor was modeled by a single symmetric segment with complex circumferential boundary conditions. With respect to different nodal diameter numbers, free vibration analyses of the disk assembly were then effectively performed. For calculated resonance conditions, the steady-state responses of the turbocharger bladed disk were computed. By using this coupled CFD–FEM analysis, the dynamic loading of the turbine blades can be determined in the design process.
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