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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]

The performance of hybrid rocket motor can be significantly boosted by metal and metalloid solid-fuel additives. However, the ablation rate of nozzle is dramatically increased by the additives at the same time. In this study, the effect of aluminum, boron and aluminum hydride additives on the performance and nozzle ablation in a lab-scale hybrid rocket motor is investigated. The propellant combination of 95% hydrogen peroxide and HTPB-based fuels is adopted. Nine tests are operated in various oxidizer mass flux to analyze the combustion performance. The regression rate for each test is measured by the fuel mass reducing. The ablation rate of nozzle is measured by three-jaw caliper and Computed Tomography scans. Electron microscope scanning and energy spectrum analysis are carried out on nozzle surface to investigate the microscopic morphology and composition of the nozzle surface after test. Results show that the difference between the specific impulse values of the two grains with the same aluminum element content is negligible. The specific impulse of aluminum-containing grains is generally higher than that of boron-containing grains. The regression rate of fuel adding 58% aluminum is higher than the fuel adding 20% aluminum and 38% aluminum hydride by 21.8%, suggesting that aluminum has a superior regression rate improvement than that of aluminum hydride. Among these three grains, the combustion efficiency and specific impulse of the fuel adding boron is the lowest, which indicates that the combustion performance of boron as an additive is inferior to aluminum and aluminum hydride. The addition of aluminum and aluminum hydride increases the ablation rate of nozzle and aggravates the deposition of alumina particles on the inner wall of the nozzle, whereas the addition of boron nearly has no effect on nozzle ablation. [ABSTRACT FROM AUTHOR]

Nozzle erosion is a common problem in a hybrid rocket motor with the passive thermal protection structure. Nozzle profile and erosion rate distribution are changing with time due to the chemical erosion. Serious erosion greatly influences the performance of the hybrid rocket motor. This paper is aimed to analyze the characteristics of nozzle erosion during the firing process. For this purpose, an unsteady numerical model of carbon-based nozzle erosion is established. The thermochemical erosion of nozzle is simulated coupled with the flow field. A dynamic grid technique is adopted to simulate the regression process of the nozzle surface. Based on this model, numerical simulations on a hydrogen peroxide(HP)/hydroxyl-terminated polybutadiene (HTPB) hybrid motor have been performed in a long-time working process. To validate the numerical results, a full-scale hybrid rocket motor is tested. The simulated nozzle erosion rate is about 0.069 mm/s in the steady state simulation, which agrees well with the test data (0.072 mm/s). The material properties of the carbon ceramic composite are affected by the high-temperature gas and suffering performance degradation in the firing process. Both of the combustion chamber pressure and oxidizer to fuel ratio have great influence on the erosion rate. [ABSTRACT FROM AUTHOR]

Hybrid rocket motors (HRMs) easily achieve variable thrust by changing oxidizer mass flow rate. This paper represents numerical and experimental studies on the throttleable HRM with aerospike nozzle, which can enhance the performances of the throttleable HRM theoretically. By comparing the performance of the throttleable HRM with the aerospike nozzle and that with the de Laval nozzle, the characteristics of the aerospike nozzle applied to the throttleable HRM are explored. Both numerical and experimental investigations are conducted. In the simulation part, characteristic velocity, combustion efficiency, thrust coefficient and specific impulse of the HRM with the two types of nozzles at different oxidizer mass flow rate are obtained. Comparing with the HRM with the de Laval nozzle, the HRM with the aerospike nozzle can improve specific impulse by about 8.9%∼23.4% when the working pressure ratio is lower than the design pressure ratio. The combustion efficiency of the throttleable HRM with the aerospike nozzle is 3.9%∼8.6% higher than that with the de Laval nozzle because of the core structure of the aerospike nozzle. When the oxidizer mass flow rate is 50 g/s, the thrust coefficient of the aerospike nozzle is increased by 12.3% relative to that of the de Laval nozzle which flow separation occurs in. The corresponding firing tests are performed by the lab-scale motor with 90% hydrogen peroxide (HP) and polyethylene (PE) grain. The numerical and experimental results agree well. The combination of the throttleable HRM and the aerospike nozzle has the advantages to improve the performance of HRMs. [ABSTRACT FROM AUTHOR]