The effect of increasing the tip-gap size on the performance of a splittered transonic rotor is presented. Tip clearance has a large influence on the performance and efficiency of compressors and fans during operation. In a gas turbine engine the ratio of tip-gap to blade height or span usually increases in the direction of flow. The front stages usually have a smaller ratio of tip-gap to blade height than the aft core stages. In addition the front stages are usually operating in the transonic regime while the rear stages operate sub-sonically.

In order to be representative of these differing flow regimes the results of a range of tests at varying tip-gaps and speeds from subsonic to transonic are presented. A highly loaded transonic axial splittered rotor is used as the test article in this study. Three experiments with cold tip gaps of 0.53 [mm], 0.76 [mm] and 0.99 [mm] are presented. Each experiment was run at six tip-Mach numbers ranging from Mach 0.72 to Mach 1.2 each over a full speed-line from choked to stalling conditions. Exit temperature and pressure profiles at the rotor exit are presented along with performance maps of pressure ratio and efficiency. Significant differences in performance in terms of pressure ratio, efficiency and operating range due to the tip-gap increase were observed and are presented.

The full mechanical geometry is available upon request to provide an open test case to evaluate simulation codes. This includes the cold-shape, the cavity between the rotor and stationary exit and the mounting bolts and geometry. The manufacturing steps in the preparation of the material and methods used to form the abradable material used over the rotor are presented. In addition the inlet conditions taking into account the effect of relative humidity on gas properties are presented. The complete data set of experimental results is also available electronically.

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