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A detailed experimental investigation to understand and quantify the development of blockage in the flow field of a transonic, axial flow compressor rotor (NASA Rotor 37) has been undertaken. Detailed laser anemometer measurements were acquired upstream, within, and downstream of a transonic, axial compressor rotor operating at 100%, 85%, 80%, and 60 % of design speed which provided inlet relative Mach numbers at the blade tip of 1.48, 1.26, 1.18, and 0.89 respectively. The impact of the shock on the blockage development, pertaining to both the shock / boundary layer interactions and the shock / tip clearance flow interactions, is discussed. The results indicate that for this rotor the blockage in the endwall region is 2–3 times that of the core flow region, and the blockage in the core flow region more than doubles when the shock strength is sufficient to separate the suction surface boundary layer