Jingfei Gao

Hybrid rocket motors have great development potential due to their outstanding thrust adjustment flexibility and long-term operation ability. However, nozzle erosion during the motor operation can cause an increase in the throat area of the nozzle, which leads to a decrease in combustion chamber pressure and nozzle efficiency. Therefore, a performance prediction model for hybrid rocket motors considering nozzle erosion has become a key technology that must be addressed when developing hybrid rocket motors. This study uses dynamic grid technology to simulate the regression of the combustion surface and nozzle erosion, which fits well with experimental values. The behavior of high-energy particles in the combustion chamber is simulated through a discrete phase model. Notably, distinctive behavior is observed in Al and Mg droplets, with Al droplets exhibiting incomplete vaporization in the combustion chamber while Mg droplets completely vaporize. A ground firing test using the Φ336 mm hybrid rocket motor lasting 200 s is conducted. The results show that the dynamic numerical simulation, accounting for nozzle erosion, substantially enhances performance prediction accuracy. The average deviation in motor thrust remains below 1.8%, and the combustion chamber pressure deviation stays under 2.6%, confirming the precision of the model. Ultimately, both simulation and experimental outcomes indicate a gradual decrease in specific impulse and characteristic velocity over the long-term operation, attributed to the gradual deviation of the oxygen-fuel ratio. This research provides valuable insights for guiding hybrid rocket motor design and optimizing design parameters to improve overall performance. This model can achieve long-duration and high-precision performance predictions for hybrid rocket motors.

Space exploration greatly facilitates the development of advanced propulsion systems. Extensive research has shown that hybrid rocket motors have bright prospects for use in variable-thrust propulsion systems. However, the variable-thrust precision control of a hybrid rocket motor with a high-mass fraction of aluminum has not been adequately explored. In this paper, we propose a closed-loop control system for a high-performance laboratory-scale hybrid rocket motor, and verify its performance through tests on a hybrid rocket motor containing 98% hydrogen peroxide and hydroxyl-terminated polybutadiene with 58% of an aluminum additive. The results show that, first, the average value of thrust in the stable sections in the three stages were 400.7 N, 599.1 N, and 400.1 N when the target values were 400 N, 600 N, and 400 N, respectively. Second, the average thrust was stable, and the control error of the average value was better than 0.5%. Third, the real-time error in thrust was controlled to within ± 20 N with a steady-state error smaller than 5%. These results indicate that the proposed closed-loop control strategy for hybrid rocket motors with a high-mass fraction of aluminum can maintain a constant thrust and smooth transitions in case of variable thrust.

Catalytic hybrid rocket motors using hydrogen peroxide (HP) are easy and efficient to achieve multiple starts and stops, and hydroxyl-terminated polybutadiene (HTPB) grains are commonly used due to their excellent mechanical properties. The low regression rate of HTPB grains limits the application of hybrid rocket motors. Furthermore, the addition of solid aluminum particles can effectively improve the regression rate and performance of hybrid rocket motors. However, the experimental results of a high mass fraction of aluminum and hydrogen peroxide available at present are not sufficient. In this research, the impact of a high mass fraction of aluminum on the motor performance and ablation rate of nozzles is studied experimentally. A solution of 95% hydrogen peroxide and HTPB with an aluminum additive are adopted as propellants. The variation in the axial regression rate of the grains is obtained by computed tomography (CT) scans and pre-test parameter measurements. The instantaneous regression rate method is adopted to obtain the real-time regression rate of the motor. The surface appearance and composition of the front and the end of the grains after the tests are analyzed by electron microscopy. Carbon ceramic and tungsten-bronze nozzles are used to explore the effect of a high mass fraction of aluminum on nozzle ablation. The experimental results show that the addition of aluminum raises the specific impulse and decreases the optimal oxygen-to-fuel ratio of the propellant combination. The high mass fraction of the aluminum particles has a severe ablative effect on carbon ceramic nozzles, while the effect on tungsten-bronze nozzles is minimal for a hot test lasting four to five seconds. Our results can provide experimental guidance for the application of a high mass fraction of aluminum and hydrogen peroxide hybrid rocket motor.