Abstract: Combustion instability is a critical issue limiting the performance and reliability of hypersonic propulsion systems. In this study, Large Eddy Simulation (LES), Fast Fourier Transform (FFT), and Dynamic Mode Decomposition (DMD) were employed to investigate the combustion instability of a strut combustor under a Mach 2 inflow condition. By integrating frequency-domain analysis with modal decomposition techniques, unsteady pressure signals and flow field data were jointly analyzed, enhancing the interpretation of low-frequency instability components and their underlying mechanisms. Hydrogen was injected parallel to the freestream at different pressures to change the equivalence ratio (ER). Time-averaged results reveal three distinct flame stabilization modes: the turbulence–lift flame mode, the boundary-layer flame stabilization mode, and a transitional mode between the two. Although these modes appear statistically stable, significant instabilities are still observed in the instantaneous flow field. In the turbulence–lift flame stabilization mode, the pressure spectrum exhibits a dominant instability frequency of 416 Hz, characterized by flame oscillations among the upper boundary mode, lower boundary mode, and turbulence–lift mode. In the transitional mode, a dominant instability frequency of 651 Hz appears, with flame oscillating occurring only between the upper- and lower-boundary mode. In the boundary-layer flame stabilization mode, two instability frequencies of 1000 Hz and 916 Hz are observed, corresponding to flame flashback along the strut wall. The DMD further confirms that the instabilities in the turbulent-lift and the transitional modes are driven by oscillations of the shear layer near the strut base, whereas the instability in the boundary-layer flame stabilization mode is oscillations by boundary-layer separation induced by a high adverse pressure gradient.