An investigation has been carried out to examine one aspect of supersonic jet-engine inlet flow, namely shock-wave and boundary-layer interaction in the presence of an expansion corner. The experiment was undertaken at Mach numbers of 1.8 and 2.5. The local Reynolds number per metre varied from 2.79 to 3.13 × 107 for M = 1.8 and from 4.20 to 6.4-8 × 107 for M = 2.5. Adiabatic turbulent boundary layers of about 6 mm thick on an expansion-corner plate were disturbed at different streamwise positions by oblique shocks with deflection angles of 4°, 6° and 8°. The pressure distributions, wave patterns and the extent of the separation regions were strongly influenced by the presence of the expansion corner, especially when the shock impingement positions were upstream of the corner. The pressure rise induced by the incident shock began to decrease in level when the shock impingement position was about 3-4 boundary-layer thicknesses upstream of the corner. As the incident shock moved towards the corner, the pressure rise and the extent of the separation region decreased. When the shock impingement position was at the corner, the reflected shock was effectively neutralised by the Prandtl-Meyer expansion; especially when the deflection angle of incident shock matched that of the expansion corner. The three dimensionality of the incident shock and separation region is due to the interactions between the glancing oblique incident shock and the tunnel’s sidewall boundary layers.