Supersonic H-2/air combustion experiments of a fixed entry Mach number 2.5 were conducted using eight different model combustors, at various stagnation conditions and global equivalence ratios. Specifically, stagnation temperature varied from 1200 to 2000 K, stagnation pressure ranged from 1 to 1.4 MPa, and the global equivalence ratio covered the range from lean to rich. In addition, the static pressure distribution in the axial direction and total pressure at the combustor exit were measured. Effects of wall injection, strut injection, and cavity flameholder were systematically investigated and compared. A one-dimensional model was further applied for data reduction and analysis. The calculated results were found to be fairly consistent with the experimental measurements. Performances of various model combustors, as well as the factors affecting combustion efficiency and total pressure recovery, were discussed.