Abstract: To mitigate micro-vibrations during on-orbit spacecraft operation and enhance structural stability and service life, this study investigates a cantilever beam structure with an end-stop clearance. The force integration method was employed to incorporate both solar heat flux and contact reaction forces into the system equations, establishing a thermally induced vibration model that accounts for the end-stop effects. The model was solved numerically using finite element-fitted mode shapes. Results demonstrate that the proposed model effectively predicts the nonlinear dynamic responses induced by both thermal excitation and clearance collisions. When the stop stiffness is below 100 N/m, increasing the stiffness significantly suppresses the vibration peaks, with the effect saturating beyond this threshold. Sensitivity analysis reveals that the primary influencing factors, in descending order of significance, are: the solar heat flux incidence angle, the damping ratio, the clearance size, and the stop stiffness. This study elucidates the relationships between key parameters and the system response, providing theoretical and numerical support for the modeling of thermally induced nonlinear vibrations, structural vibration control, and parameter optimization design.