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This paper presents the results of an experimental investigation of a high aspect ratio rotor design. Future axial compressor designs for jet engines require increased efficiency in conjunction with reduced weight and size. Therefore, the investigation determines the influence of a decreased chord length on the rotor behavior. The experiments were carried out at the transonic compressor test rig at TU Darmstadt. The test rig represents a modern front stage highpressure compressor of a jet engine. The analysis is based on several measurements, using extensive steady and unsteady instrumentation, to obtain data for global compressor characteristics and performance as well as transient behavior, focusing on aerodynamic and aeromechanical phenomena. This paper compares the aerodynamic behavior of different transonic compressor stage setups, including an inlet guide vane and stator as well as three varying rotor designs. The baseline geometry design represents a modern low aspect ratio rotor (Rotor A). Based on that, the second rotor (Rotor B) is designed with an increased aspect ratio, resulting from an average chord length reduction of more than 20 % across the blade height. The blade height and rotors tip clearance as well as the solidity is kept constant, resulting in an increased blade count. By retaining constant testing conditions, all measurement data and observations can be merged, in order to describe the effects with respect to varying rotor aspect ratios. Therefore, the IGV and stator vanes as well as the entire instrumentation stays untouched during a rotor BLISK change. Five-hole-probe measurements at the rotor inlet show equal flow conditions for all setups. Comparing the compressor characteristics at design and part speed operation, the investigation shows a significant loss in stall margin and total pressure ratio. The aim of the desired future rotor design is to recover the losses due to a higher aspect ratio. Therefore, the investigation of this paper also includes a third rotor design (Rotor C). The mechanical and geometrical design parameters are the same as for Rotor B. The design is based on a numerical 3D optimization process. The experimental study is in line with the numerical predictions. Rotor C shows an equivalent behavior in total pressure rise and stall margin compared to the baseline (Rotor A) geometry. As a conclusion, this paper indicates the possibility to design a compressor rotor with a considerable reduced chord length and 3D optimized blading, without suffering significant losses in stall margin or total pressure rise. INTRODUCTION Future aero engine design and optimization is driven by environmental awareness, combining higher efficiency with a reduction in exhaust and noise emissions. Besides performance, the engines’ thrust-to-weight-ratio is a key factor. The compressor, as a multi-stage setup, significantly affects the overall engine weight and length, hence, resulting in requirements for more lightweight and efficient designs along with constant power capabilities. In order to accomplish these goals, different measures are reasonable. Enabling the required overall pressure ratio with a minimum number of compressor stages, as well as combining a reduced blade count and thinner blades in a BLISK rotor, is a fundamental design challenge. An increased performance can also be enabled by reduced stage length, particularly requiring shorter blades and vanes, while maintaining the initial total pressure rise.

Experimental Investigation of a Transonic Compressor with High Aspect Ratio Rotor Design