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This paper aims to study experimentally and numerically the effects of aluminum (Al) and aluminum hydride (AlH 3) additives on the performance of HRMs. A steady-state numerical model of HRM is established, which adopts the 95% hydrogen peroxide (HP) as the oxidizer. A firing test is completed for the certification of the numerical simulation model, the results illustrate that the regression rate of addition of 38% AlH 3 and 20% Al increases by 55.6% compared with pure hydroxyl-terminated polybutadiene (HTPB). A good agreement exists between the simulation regression rate and the experimental data. The deviations of thrust and pressure are less than 3.3% and 3.7%, respectively. Specifically, the error of the regression rate is only 0.3%, indicating that the accuracy of adopting numerical simulation method to predict the characteristics of the HRM is acceptable. Several steady-state simulations based on the different content of AlH 3 and Al are performed to investigate the effects on regression rate and combustion performance of the HRM. Due to the addition of H 2 , the thermal conductivity increases with the decrease of the mass fraction of AlH 3 , which leads to the increase of the regression rate. The regression rate of the fuel with 8% AlH 3 and 50% Al increases by 35.5% than the fuel with 58% AlH 3 , demonstrating that the effect of Al on the regression rate enhancement is more than that of AlH 3. The theoretical specific impulse increases with the mass fraction of AlH 3. However, it is owing that the part of H 2 does not participate in the reaction with O 2 and the oxidizer to fuel ratio (O / F) is far from the optimal O / F , the specific impulse is lower than the theoretical specific impulse. [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]

Abstract: This paper proposes the use of an optimal design method, based on multi-island genetic algorithm (MIGA), for the design of a ballistic single-stage sub-orbit rocket vehicle employing a hybrid rocket motor. The mathematical model of sub-orbit hybrid rocket system and a 2 degree of freedom (DOF) trajectory simulation are developed as part of this investigation. The star port grain rocket vehicles with 98% and 90% hydrogen peroxide (H2O2, diluted with water, for short “HP”)/hydroxyl-terminated polybutadiene (HTPB) propellant combination, and the 90% HP/HTPB propellant combination rocket vehicles with four different gain types, including star port grain, single circle port grain, wheel port grain and multi-tube grain, are optimized respectively to carry out the parametric analysis on the relations of propellant combinations and design variables with the performances of the rocket vehicles. The final results show that the increase of oxidizer mass concentration has considerable contributions to reduce the lift-off mass of the rocket vehicles, while it has almost no effects on the initial grain shape parameters except the initial grain diameter. For the four initial design parameters, the value of initial chamber pressure, initial mixture ratio and nozzle expansion ratio have less effect on the rocket performances compared to the initial value of thrust, when the initial grain shape parameters are constant. Among the four type hybrid rockets, the one with a multi-tube grain performs best, because it has the largest burning perimeter length. [Copyright &y& Elsevier]