Search

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 three-dimensional dynamic numerical simulation is rare in previous studies of hybrid rocket motors, therefore, a three-dimensional dynamic numerical model is established. A hybrid rocket motor is designed, which adopts 98 wt% hydrogen peroxide and polyethylene as the propellants. Then, a ground experiment is conducted to verify the model. The results show that the method of adopting dynamic mesh technology to simulate the three-dimensional regression of the combustion surface is feasible. The average errors of thrust and pressure in the simulation and experiment are 5.6% and 7.9%, respectively. Moreover, the inner surface of grain is scanned by computed tomography after the experiment. The average deviation of port diameter is 5% between the simulation and experiment, therefore, the accuracy of this model is acceptable. Then, the results also show that the thrust and pressure both gradually decrease with time in the simulation and experiment at 5-10 s. The main reason is the rapid reduction in the fuel mass flow rate. However, the pressure in simulation is stable but decreases with time in experiment at 10-25 s. The main reason is the expansion of the nozzle throat caused by the ablation. The results demonstrate that the numerical method can accurately predict motor performance. In addition, the flow field is approximately uniform along the circumference by the dynamic numerical method, which is difficult to obtain for the conventional steady-state and two-dimensional methods. It is the first time to realize the three-dimensional regression in hybrid rocket motors. Therefore, this method can be used to predict the three-dimensional regression of the complex shapes grain and also can guide the design of grain shape in hybrid rocket motors. [ABSTRACT FROM AUTHOR]

Electric pumps capable of operating at high rotational speeds have enormous potential for application in hybrid rocket motors because of their light weight and ability to made rapid adjustments in rotational speed. Previous studies have assessed the performance of electric pumps; however, the physical mechanisms underlying the operational instabilities of high-rotational-speed pumps remain obscure. Simulations were accordingly conducted in this study to explore the physical mechanism of experimentally observed operational instabilities. The average head deviation between the simulated and experimental results was within 2%. The results of the instability analysis at low mass flow rates indicated that the reflux at the impeller outlet intensified, resulting in increased velocity, turbulent Reynolds number, and vorticity. The flow structure at a mass flow rate of 4.2 kg/s caused significant rotor torque fluctuation (2.1 times that at 5.88 kg/s), which coupled with the electric motor rotation and gradually increased the instability of the electric pump until shutdown. The minimum mass flow rate at which the electric pump operated stably was between 4.2 and 4.47 kg/s. The results of the instability analysis also indicated that severe cavitation occurred at 7.4 kg/s that reduced the head and significantly disrupted the flow stability. Furthermore, this cavitation led to an increased torque fluctuation amplitude and more dispersed frequency distribution (compared with that at 4.2 kg/s), which affected the stability of the electric motor and ultimately led to shutdown. These findings provide theoretical guidance for the development of thrust adjustment schemes for electric-pump rocket motors, benefiting the aerospace community. [ABSTRACT FROM AUTHOR]

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]

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]

An analytic target cascading (ATC) method with fuzzy uncertainties based on global sensitivity analysis is proposed to solve the optimization problem of typical multilevel multidisciplinary nonlinear system. The overall design of launch vehicle (LV) powered by hybrid rocket motors (HRMs) involves multiply disciplines, and is decomposed into two levels using an ATC framework. The fuzzy theory is applied to describe the uncertain design factors caused by decisions and cognition insufficiency. The rank correlation coefficient method (RCCM) based on feasible optimization solutions and the quadratic response surface method (QRSM) based on Latin hypercube sampling (LHS) are used for global sensitivity analyses of input uncertainty and model uncertainty, respectively. The multi-island genetic algorithm (MIGA) is adopted in all examples, and the known two-phase optimization method is used to verify the LV design results. The results show that global sensitivity analysis can significantly filter the fuzzy uncertain factors which have little influence on the responses. The ATC decomposition is applicable in solving the calculation burden caused by uncertainties. The fuzzy-based design optimization with ATC is more efficient than that with MDF, and gives more reliable and robust results than the deterministic ATC method. [ABSTRACT FROM AUTHOR]

Throttleable engines have greatly raised the technical performances of aircrafts and missiles, particularly for cruise vehicles. The hybrid rocket motor (HRM) combines the advantages of solid and liquid rocket engines, especially the ability of easily throttling and restart. This paper presents a design optimization of an HRM propelled cruise vehicle to explore the throttling ability of HRM on velocity control. First, a theoretical thrust model is presented and revised through several experimental tests. Based on it, a detailed design process of HRM is proposed. Second, the layout of the cruise vehicle, flight trajectory model and a closed-loop PID velocity control method are developed. The differential evolution algorithm is adopted to get the optimal design result, which is used to analyze velocity characteristics by real-time control of the oxidizer flow rate. The results show that the PID-control has good effect on velocity control performance and maintaining small deviation from the target. The influences of aerodynamic and thrust errors are also considered and analyzed. For different types of deviations, including the proportional and periodic perturbations, the PID-controlled scheme shows a good stability under the parameters' deviations, which indicates that the HRM with the real-time throttling thrust has a good capacity of high-precision velocity control when used in rocket propelled vehicles. [ABSTRACT FROM AUTHOR]