Prediction of flutter boundary in transonic cascades

The goal of this paper is to explore blade aeroelastic stability, significantly influencing the aero-engine safety which is required in airworthiness and certification standards of aero-engine, to guide the design of fan or compressor blade in aircraft engines. A fluid-structure interaction (FSI) numerical method is developed for turbomachinery flutter boundary in a transonic cascade to identify the key flutter parameters. The use of computational structural dynamics (CSD) for blade motion and computational fluid dynamics (CFD) for the flow with body-fitted moving blade boundary are calculated with 3D Finite Element interpolation algorithm using invariant distance vectors for data transfer in the overlapping region. Energy Method for aerodynamic modal damping ratio is proven to be accurate and efficient by computing a transonic cascade, the NASA Rotor 67 transonic fan. The validated computer codes are used in a series of parametric studies, and the influence of mode shapes is investigated to shown as an important contributor to determine the stability of blade design. The contour of aerodynamic modal damping ratio in rotor characteristic map, clearly indicating the flutter boundary and the region of aeroelastic stability and instability, is obtained in the different operating conditions with respect to different back pressures and different rotating speeds. The investigation is necessary for the detailed analysis stage of the typical blade design process. Utilization of the procedures will lead to blade designs free from aeroelastic instability for all design conditions, which should be paid close attention during the safety analysis of the airworthiness standards and the analysis of airworthiness compliance.