Search

With the aim of enhance the designability of burning surface, star and wagon wheel fuel types are applied extensively in hybrid rocket motors. However, mechanism of nozzle ablation in hybrid rocket motors with these complex fuel types is unclear, and this increases the difficulty of predicting ablation laws and thrust. This paper intends to investigate spatial distribution characteristics of nozzle ablation in the hybrid rocket motor with various fuel types. According to the principle that the grain port area and grain length are equal, tube, star, single-port wagon wheel, and multi-port wagon wheel grains are designed. Numerical models for combustion flow and thermochemical erosion are established, and three-dimensional numerical simulations are conducted. A firing test for nozzle erosion is conducted, and hydroxyl-terminated polybutadiene and 95 % hydrogen peroxide are employed. Errors of throat erosion rate and combustion chamber pressure among simulation results and experimental data are 4 % and 4.25 %, respectively. Calculation results demonstrate that the circumferential distributions of nozzle wall temperature and erosion rate show an obvious correlation with fuel types. The maximum throat erosion rates for star, single-port wagon wheel, and multi-port wagon wheel grains occur at the circumferential position of 0° (star slot, spoke slot, and midpoint of outer edge of wheel hole), while the minimum throat erosion rates locate at the circumferential position of 36° (star root, spoke root, and midline of adjacent wheel holes). The circumferential non-uniformity of throat erosion rate of multi-port wagon wheel grain is the highest. When the grain port area and oxidizer mass flow rate are the same, the throat erosion rate of multi-port wagon wheel grain is the highest, followed by the star grain, and the throat erosion rate of single-port wagon wheel grain is the lowest. • Effect of complex fuel types on spatial characteristics of nozzle erosion is studied. • Firing test of carbon-based nozzle erosion is performed to verify numerical models. • Three-dimensional simulations on erosion in hybrid rocket motors are conducted. • Simulation results agree well with test data. • Ablation mechanism of star, single-port, and multi-port wagon wheel grain is revealed. [ABSTRACT FROM AUTHOR]

The hybrid rocket motor (HRM) emerges as a promising propulsion system offering thrust adjustment at a low cost. Prior research has indicated that the inclusion of micro Al additives enhances HRM performance, yet a comprehensive investigation remains absent. Thus, this study aims to explore the impact of Al particle additive content and size variation on the characteristics of hybrid rocket motors (HRMs) through experiments and simulations. Firing tests are conducted on HRMs employing unadulterated HTPB, 38 % additive Al (2 μm), and 58 % Al (2 μm). Experimental findings reveal enhancements of 12.6 % and 33.9 % in regression rate, and 2.5 % and 8.0 % in specific impulse for the 38 % Al and 58 % Al fuel configurations, respectively, compared to the base case. Subsequently, a numerical two-phase combustion model is built to analyze the inner flow field characteristics. The high consistency between experimental and simulation results validates the efficacy of the numerical approach employed. Following this, a series of simulations covering Al content ranging from 0 % to 58 % and Al particle sizes from 0.4 μm to 75 μm is conducted. Simulation outcomes indicate that specific impulse and combustion efficiency enhancements correlate with the increase in Al mass fraction for particles with diameters under 35 μm. However, high Al content proves detrimental to specific impulse and combustion efficiency when particle diameter exceeds 35 μm. Additionally, regression rate and thrust demonstrate an increase with Al content, albeit the rate of increment diminishes with larger Al particles. This study comprehensively examines the influence of Al particle addition on HRM performance and elucidates the underlying mechanisms, thereby offering insights into HRM design and manufacture. • Hybrid rocket firing tests are performed on HTPB, 38 %, and 58 % Al particle fuel. • Simulations with various Al particle contents and diameters are performed. • High Al content with diameter under 35 μm is beneficial for combustion efficiency. • High Al content with diameter over 35 μm is detrimental for combustion efficiency. • Large particle reduces the thrust enhancement brought by Al content rise. [ABSTRACT FROM AUTHOR]

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. [ABSTRACT FROM AUTHOR]

In order to enhance energy characteristics of propellants in hybrid rocket motors, aluminum particles are added into the solid fuel. However, these particles may lead to nozzle mechanical erosion, which significantly affects the motor performance. This paper intends to study thermochemical and mechanical erosion mechanism of carbon-based nozzles in hybrid rocket motors with aluminum metallized fuel. A firing test is performed, and 95 % hydrogen peroxide and hydroxyl‑terminated polybutadiene based fuel with aluminum particles are utilized. Experimental results show that many metal particles are ejected from the nozzle, and aluminum oxide deposits heavily on the nozzle inner surface. A method of numerical calculation coupled with two-phase flow, combustion of aluminum metallized fuel, thermochemical ablation reactions, and mechanical erosion is established. The errors of combustion chamber pressure and throat erosion rate between numerical calculation and test results are 0.53 % and 7.76 %, respectively. Simulation results indicate that with the increase of aluminum content, nozzle wall pressure gradually increases, while the mass fraction of H 2 O, CO 2 , OH, and O declines. The nozzle wall temperature and thermochemical erosion rate raise first and then reduce as aluminum content increases. The mechanical erosion is mainly distributed in the convergent section, and it raises with the increase of aluminum content. When aluminum content is up to 38 %, the maximum mechanical and thermochemical erosion rates are 0.09185 mm/s and 0.0976 mm/s, respectively. This reveals that effects of mechanical erosion on hybrid rocket nozzles should be carefully considered and evaluated under conditions of high aluminum content. [ABSTRACT FROM AUTHOR]

Lunar missions are currently experiencing a significant surge in popularity, presenting expansive opportunities for further exploration and development. To thoroughly explore the design margins and potential of lunar landers, and to foster the development of overall designs driven by comprehensive performance objectives, it is crucial to conduct optimization design considering the coupling between key disciplines such as trajectory and propulsion. Considering the significant increase in computational complexity caused by conducting trajectory/propulsion integrated design optimization, the analytical target cascading method is employed to hierarchically decompose and coordinate optimization of the complex systems. This article presents a phased soft-landing strategy on the manned lunar lander propelled by hybrid rocket motors, utilizing powered explicit guidance and Apollo powered descent guidance, and proceeds with the trajectory/propulsion integrated design optimization involving diverse grain shapes and feed systems. This optimization process is separately undertaken utilizing multidisciplinary feasible method and analytical target cascading method. The analysis reveals that integrating trajectory and propulsion considerations into the optimization process facilitates a 5 % reduction in the overall mass relative to optimizations constrained solely by velocity increment and lack comprehensive trajectory design considerations. This highlights the profound impact of trajectory requirements on propulsion system design and the advantages of powered explicit guidance laws in minimizing fuel consumption. Crucially, the use of analytical target cascading achieves the better optimization results, and significantly reduces subsystem evaluation times, enhancing operational efficiency by 48 %, demonstrating the advantage in handling complex, large-scale systems. On another level, with different β values, the Mean Relative Error of the target values for the three schemes obtained by the analytical target cascading method is 0.0016, indicating good stability and strong robustness. The practical exploration in this article provides methods and frameworks for high-performance optimization design of complex aerospace mission profiles in the future. • The trajectory/propulsion integrated design optimization is conducted. • The two-phase guidance strategy for lunar lander descent is proposed. • Based on the strategy, the analytical target cascading-decomposed framework is set. • The influence of trajectory and propulsion coupling on overall design is revealed. • The optimization effectiveness, convergence rate, and robustness of the ATC is shown. [ABSTRACT FROM AUTHOR]