Due to their desirable structural tailoring capability and weight saving potential, composite structures have been widely employed in aerospace applications. The damage evolution and the failure mechanism of composite laminates with internally terminated plies under compression load were investigated experimentally. Four types of tapered configurations were tested, including different taper angles and drop-off sequences. The results indicated that the geometry change and material discontinuities of the tapered configuration gave rise to high-level stress concentration, prompting the delamination occurrence in the vicinity of the terminated plies. Meanwhile, the eccentricity of mid-planes between thin and thick sections contributed to a bending tendency of the tapered structure. Different carrying capacities and failure process were observed in specimens with varying tapered configuration parameters. All tested articles exhibited a typical compressive failure pattern, characterised mainly by the fiber fracture and the micro-buckling of the longitudinal fibers in the thin section. Moreover, distinct delamination was observed in three of four tested types.