A cascade geometry derived from a research program on high-throughflow, transonic, axial-flow compressors was tested with similar supersonic entrance flow conditions in three linear cascade test facilities. The airfoil section used was representative of advanced rotor-blade, tip-region profiles designed to operate at inlet relative Mach numbers of 1.4 to 1.8. Objectives in the experiments were to study the reproducibility of test conditions and measured performance in facilities that are considered to be “state-of-the-art,” and to generate data sets that could be used as test cases or “benchmark” results to validate computational methods for turbomachine application. It was recognized from the beginning of the project that the aerodynamic regime involved represents a very difficult combination of problems in both experimentation and computation. This difficulty was certainly encountered; the experimental problems are fully discussed in this paper and the companion papers originating in two of the test groups. An excellent series of data sets has been obtained, and our confidence in the results is supported by the exchange of information and personnel that occurred during all phases of the experiments. The results presented here and in the forthcoming AGARD Propulsion and Energetics Panel “test case” compendium should serve as a standard for evaluating current and future computational efforts.

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