Thrusters are essential components of space systems—including satellites, spacecraft, and lunar modules—due to their ability to generate controlled thrust for precise attitude and trajectory adjustments. Among the factors influencing thruster efficiency, the propellant plays a critical role in determining key performance metrics. Ammonium dinitramide (ADN), a high-density oxidizer with a positive oxygen balance, low toxicity, non-carcinogenic characteristics, and chlorine-free composition, has emerged as a promising alternative to traditional oxidizers. This study simulates the influence of feed parameters on the performance of a single-propellant thruster utilizing an ADN-based liquid formulation. Two-dimensional, steady-state modeling was performed to evaluate the effects of input parameters on combustion and thrust generation. The propellant mixture consisted of 63% ADN, 11% methanol (CH₃OH), and 26% water (H₂O), delivered at a mass flow rate of 2/3 g/s. The preheat temperature was set at 473 K, and the catalyst bed porosity was 0.5. The governing equations for continuity, momentum, energy, and species transport, along with a detailed reaction mechanism involving 40 reactions and 18 chemical species in a porous medium, were solved using a pressure-based numerical approach. The effects of preheat temperature, catalyst bed porosity, and methanol mass fraction were investigated in relation to NO₂ and O₂ mass fractions, maximum pressure and temperature, specific impulse, and thrust force. Results indicated that porosity values of 0.4, 0.5, 0.6, and 0.7 yielded thrust forces of 13.5, 19.5, 18.5, and 16.5 N, respectively, with 0.5 producing the highest output. Increasing porosity initially raised the maximum pressure and temperature, enhancing both thrust and specific impulse, before declining beyond an optimal threshold. A direct correlation was observed between preheat temperature and performance enhancement; raising the temperature from 423 to 573 K increased thrust from 13.5 to 22.5 N. Variation of the methanol mass fraction from 0.07 to 0.2 resulted in an initial rise, followed by a drop in thrust, with a peak value of 19.5 N occurring at a mass fraction of 0.11. Higher methanol content increased the exhaust gas mass and, consequently, the propellant consumption rate. Therefore, optimizing the methanol fraction is essential for maximizing thrust efficiency.