An analysis system for blade forced response
Copyright © 1992 by ASME. A frequent cause of turbomachinery blade failure is excessive resonant response. The most common excitation source is the nonuniform flow field generated by inlet distortion, wakes, and/or pressure disturbances from adjacent blade rows. The standard method for dealing with this problem is to avoid resonant conditions using a Campbell diagram. Unfortunately, it is impossible to avoid all resonant conditions. Therefore, judgements based on past experience are used to determine the acceptability of a blade design. A new analysis system has been developed to predict blade forced response. The system provides a design tool, over and above the standard Campbell diagram approach, for predicting potential forced response problems. The incoming excitation sources are modeled using a semi-empirical rotor wake/vortex model for wake excitation, measured data for inlet distortion, and a quasi-3D Euler code for pressure disturbances. Using these aerodynamic stimuli, and the blade's natural frequencies and mode shapes from a finite element model, the unsteady aerodynamic modal forces and the aerodynamic damping are calculated. A modal response solution is then performed. This system has been applied to current engine designs. A recent investigation involved fan blade response due to inlet distortion. An aeromechanical test had been run with two different distortion screens. The resulting distortion entering the fan was measured. With this as input data, the predicted response agreed almost exactly with the measured response. In another application, the response of the LPT blades of a counter-rotating supersonic turbine was determined. In this case the blades were excited by both a wake and a shock wave. The shock response was predicted to be three times larger than that of the wake. Thus, the system identified a new forcing function mechanism for supersonic turbines. This paper provides a basic description of the system, which includes: 1) models for the wake excitation, inlet distortion, and pressure disturbance; 2) a kernel function solution technique for unsteady aerodynamics; and 3) a modal aeroelastic solution using strip theory. Also, results of the two applications are presented.
Asme 1992 International Gas Turbine and Aeroengine Congress and Exposition, Gt 1992
Volume / Issue
International Standard Book Number 13 (ISBN-13)
Digital Object Identifier (DOI)