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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]

This paper investigates the effect of rotation angles on the fuel regression rate and the motor combustion efficiency in hybrid rocket motor with two-hole segmented grains. In this research, the propellant combination of 90% hydrogen peroxide (H 2 O 2) and polyethylene (PE) is adopted, and two two-hole grains with the same configuration are used to conduct both numerical and experimental tests. A 3D simulation model is established to obtain combustion efficiency, fuel regression rate, temperature distribution and species mass fraction distribution. Meanwhile, in order to achieve the motor combustion efficiency and the average regression rate, 20 firing tests were conducted on a lab-scale hybrid rocket motor. The numerical and experimental results agree well and demonstrate that the rotation of after-section grain has no influence on the fore-section grain regression rate, while the after-section grain regression rate arises, compared with the no rotation case. Regression rates are fitted with the following empirical equation, r ˙ = a G ox n (1 − e − D ‾ p m ) , in which the regression rate coefficients a and n of after-section grain changed with the rotation angle. The motor combustion efficiency is higher than the base operation condition. It increases gradually with the increase of rotation angle, and the increasing rate becomes slower. The combustion efficiency can reach to the maximum when the rotation angle is 45°, after which it drops slowly, and the rate of descent is slower than that of ascent. [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 concentrates on the fuel regression rate characteristics of hybrid rocket motor with helical grain and attempts to overcome the low regression rate shortcoming of conventional hybrid rocket motor. Both three-dimensional numerical simulations and lab-scale firing tests are conducted to analyze the regression behavior of solid helical fuel grain. Gaseous oxygen and polyethylene are utilized as the propellants. Head axial injection method and Laval nozzle are adopted in the motor. A steady numerical model based on pure-gas-phase reaction assumption is established. Flow field characteristics of the hybrid rocket motor with helical grain are identified through three-dimensional numerical simulations. Simulation results show that the helical structure can both increase turbulence kinetic energy and swirl number in the flow field. These two factors both contribute to the increase of the fuel regression rate. Geometric factors influencing the fuel regression rate are thread pitch, groove depth and groove width. Experimental results reveal that the fuel regression rate increases with the increase of the groove depth and the groove width. The thread pitch can also obviously influence the regression behavior of the helical grain. Overall, the helical grain can effectively increase the fuel regression rate of hybrid rocket motor. [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]

In this paper, to improve the combustion efficiency and the regression rate in hybrid rocket motor, we introduce a concept that two different grain configurations are deployed into one combustion chamber, which is called the segmented grain in following chapters. Both numerical and experimental investigations are conducted. In the simulation part, the combustion efficiency and the distributions of regression rate, temperature and mass fraction of species in segmented grain cases are obtained. The corresponding firing tests are performed by the lab-scale motor with 90% H 2 O 2 and PE propellant combination. The combustion efficiency and average regression rates are achieved in each test case. The numerical and experimental results agree well, and demonstrate that the segmented grain configuration proves its ability of enhancing the combustions efficiency and the regression rate of the hybrid rocket motor. Comparing with the contrast cases, the combustion efficiency of segmented grain cases are raised by 15% and around 10% in the simulations and tests, respectively. The regression rates of the after-section grain are also increased according to simulation and test results. [ABSTRACT FROM AUTHOR]

A transient ignition model is established based on the theoretical analysis of hybrid rocket motor ignition process. The ignition process can be divided into four stages: inert heating, ignition, flame propagation, and rapid pressure buildup. The inert heating up time takes up the main part of the ignition delay. An experiment system is designed for 90% hydrogen peroxide laboratory-scale hybrid rocket motor with catalytic ignition. Firing tests are carried out with polymethyl methacrylate (PMMA) and polyethylene (PE) fuels. The ignition criteria of 90% H 2 O 2 /PE hybrid rocket motor are established by comparing the theoretical calculations with experimental results. Comparison results show that the ignition process is governed by both the temperature criterion (TC) and the oxidizer-to-fuel ratio criterion (OFC). OFC determines ignition delay at a low oxidizer mass flux and TC determines ignition delay at a higher mass flux. In addition, the ignition delay time is more sensitive to the initial oxidizer temperature than the initial fuel temperature. [ABSTRACT FROM AUTHOR]

This paper attempts to research and predict the effect of the convention aluminum particle additives on the performance of the HTPB/98HP hybrid rocket motor with the numerical approach. The aluminum particle is injected to the flow field in form of droplet from the grain surface, and then simulated by the discrete phase model. A CFD code has been developed to solve the coupling mass, momentum and energy equations of gas and solid at the grain surface, as well as the initial parameters of the aluminum droplet. A good agreement exists between the calculated regression rate from numerical model and the experimental data, indicating that the accuracy of applying numerical approach to predict the characteristics of aluminum metallized HRM is acceptable. Series of static numerical simulation have been conducted to determine the effect of aluminum particle property on the combustion and specific impulse performance of the HRM. The regression rate increases with the weight fraction of convention aluminum particle, but shows no distinguished correlation with the particle diameter. The characteristic velocity and vacuum specific impulse display a strong dependence on the particle diameter. The results illustrate that the proposed numerical approach is potentially an effective tool for predicting and investigating the enhancement of the regression rate and performance of the HRM with the addition of aluminum particle. [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]

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]

The objective of this study is to investigate the mechanism of mixing and atomization of pintle injectors numerically and experimentally using water and nitrogen. During the experiments, three cases with identical water injection condition and different nitrogen mass flow rate were performed. The nitrogen injection velocity was adjusted by changing the cross section of the exit gap. In other words, three cases with different gas/liquid momentum ratio and Weber number were analyzed. Meanwhile, simulations consistent with test conditions were conducted to compare with the experimental data. High-speed imaging was applied to capture of the movement of the flow field. In order to obtain the atomization cone angles, a series of image processing algorithms named principal component analysis (PCA) and region growing were adopted. Conclusions can be drawn that both the experimental and numerical results show a decrease tendency of atomization cone angle with the increase of gas/liquid momentum ratio. The PCA and region growing method was verified to be feasible and accurate for calculating atomization cone angle. The distribution characteristics of Sauter mean diameter (SMD) are determined together by Weber number and gas/liquid momentum ratio. Increasing the Weber number will help to reduce SMD and make its distribution more uniform. However, the SMD near the axis line is relatively large in the case of large momentum ratio. [ABSTRACT FROM AUTHOR]

This paper is aimed to analyze the transient process and ignition delay time of catalytic hybrid rocket motor through two-dimensional axisymmetric transient simulations and lab-scale firing tests. The overall catalytic ignition process is divided into five stages by analyzing typical firing traces. The total ignition delay time includes valve response time, pipeline filling time, catalytic bed startup time, monopropellant pressure buildup time and fuel grain heating time. The numerical simulations mainly focus on the fuel grain heating stage of catalytic hybrid rocket motor. The ignition transient flow fields of the motor with 90% hydrogen peroxide (90HP) and polyethylene (PE) propellants are obtained. Simulation results also reveal that the fuel grain heating time exponentially decreases with the increase of the oxidizer mass flow flux. Several firing tests of 90HP-PE lab-scale catalytic hybrid rocket motor are conducted to analyze the delay time of different ignition stages. The fuel grain heating time trend summarized by the experimental data accords well with that by the simulation results. Besides, the experimental data also show that the pipeline filling time, the catalytic bed startup time, the monopropellant pressure buildup time and the fuel grain heating time all decrease when the oxidizer mass flow rate increases. Therefore, the total ignition delay time almost linearly decreases with the increase of the oxidizer mass flow rate. [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]

Abstract: The scale effect of solid fuel regression rate in hybrid rocket motors is investigated theoretically based on pipe turbulent boundary combustion theory firstly, and a practicable scaling criterion of solid fuel regression rate has been obtained. The criterion indicates that solid fuel regression rate decreases with fuel port diameter and is proportional to when convective heat transfer is dominant in hybrid rocket motor. Based on the criterion, a numerical model combined with turbulence, combustion, solid fuel pyrolysis and solid–gas coupling model was established to simulate three different size motors, and an analysis was also performed on experimental data of different motors. Both the numerical results and experimental data verified the scaling criterion of solid fuel regression rate under different fuel port sizes. [Copyright &y& Elsevier]

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]