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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]

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]

Increasing the mixing efficiency of the fuel jet along the combustion chamber is a crucial step for the advancement of the current high-velocity vehicles. In this article, comprehensive computational investigations have been performed to disclose the role of an annular air jet on the fuel mixing of the single extruded nozzle located inside the combustion chamber. This study has tried to offer hydrodynamic insight about the jet flow feature when released from the extruded nozzle in supersonic cross flow. Computational fluid dynamic is used for the visualization of fuel jet interactions with air stream. Effects of nozzle height on the mechanism of fuel jets are extensively analyzed in this investigation. Our results show that the injection of the annular jet would increase the strength of the circulations, and consequently, fuel mixing improves inside the combustion chamber. Our findings display that fuel mixing increases about 100% by the injection of the annular air jet in the combustion chamber. [ABSTRACT FROM AUTHOR]

This paper provides a valuable method to perform transient numerical simulations for hybrid rocket motors (HRMs). A two-dimensional transient numerical simulation model based on Butterworth filter and dynamic mesh is established. The Butterworth filter adopted in this project is a second-order low-pass filter. Then a firing test is arranged to prove the accuracy of the numerical simulation model of the HRMs. This rocket motor adopts 98% hydrogen peroxide (HP) and hydroxyl-terminated polybutadiene (HTPB) fuel containing aluminum as the propellants. The results demonstrate that the Butterworth filter can significantly improve the smoothness of the combustion surface compared with non-filter. And the combustion chamber pressure and thrust of simulation and test results are in good agreement. The average deviation of combustion chamber pressure between the simulation and test is 0.9%, and the average deviation of thrust between the simulation and test is 3.3%, which verifies the feasibility of the numerical simulation model for this HRM. Comparing the regression rate and filtered regression rate along with axial distance over time, it is indicated that the high-frequency oscillations are filtered by the Butterworth filter. In order to compare the port of simulation and test, the combustion chamber is scanned by computed tomography (CT) after the test. The results demonstrate that the simulation and test results are in good agreement, and the average deviation of simulation and test results is 4.5% mainly due to the end-burning of grain and nitrogen blowing. Then, the long-time simulations that adopted the Butterworth filter and non-filter were conducted. The results demonstrate that the use of the Butterworth filter can significantly smooth the combustion surface and the model has the ability to perform long-time simulations. This project can make the simulation for HRMs closer to the test results and provide a valuable guide for HRMs simulation. • A simulation model of hybrid rocket motor based on Butterworth filter is established. • The Butterworth filter is adopted to smooth the combustion surface. • A firing test of hybrid rocket motor with fuel containing aluminum is carried out. • Simulated combustion surface is compared with the computed tomography scan results. • A long-time simulation is conducted to validate model accuracy. [ABSTRACT FROM AUTHOR]

This paper studies the effects of liquid oxidizer flow distribution on combustion characteristics and performance of hybrid rocket motor with axial injection through numerical simulations of two-phase flow. Axial injection is widely used in hybrid rocket motor, however, the regression of burning surface may change combustion efficiency of the motor at different moments. Consequently, it is necessary for hybrid rocket motor to design injector patterns that could increase the average combustion efficiency in the whole operating process. The orifice injector plate is divided into three areas, which are recorded as zone I, zone II and zone III, respectively. Oxidizer flow distribution is achieved through changing the diameter of injector holes in various zones, and three-dimensional steady numerical simulations of two-phase flow of hybrid rocket motor with axial injector patterns at different moments are conducted. The motor adopts 98% hydrogen peroxide and hydroxyl-terminated polybutadiene as the propellants. Numerical analysis reveals that oxidizer flow distribution mainly influences the flow field structure in the pre-combustion chamber and front part of fuel grain port. Increasing oxidizer mass flow rate in zone III is beneficial for the chamber head thermal environment. Relative position in the radial direction of fuel grain inner surface and oxidizer distributed in the injector plate affects fuel regression rates, and the influence mainly exists in the front part of solid fuel grain. When the oxidizer mass flow rate in zone I is equal, the average fuel regression rate rises with the increase of oxidizer flow rate in zone II at initial moment, and decreases as the oxidizer in zone II increases at middle moment. The average combustion efficiency increases with the increase of oxidizer mass flow rate in zone II when the oxidizer in zone I remains unchanged. Both the average combustion efficiency and the chamber head thermal protection need to be considered in the design of axial injector patterns. Research results of this paper provide an important reference for the design of high performance injector patterns of hybrid rocket motor. • Effects of oxidizer flow distribution on hybrid rocket motor performance are studied. • Steady two-phase flow simulations for various injector patterns are conducted. • Adding oxidizer in zone III is good for thermal environment of the chamber head. • Relative position of fuel inner surface and oxidizer affects fuel regression rates. • When zone I is fixed, combustion efficiency raises with oxidizer in zone II adding. [ABSTRACT FROM AUTHOR]

This paper presents transient simulations and experimental studies of the ignition process of the hybrid rocket motors (HRMs) using 90% hydrogen peroxide (HP) as the oxidizer and polymethyl methacrylate (PMMA) and Polyethylene (PE) as fuels. A fluid-solid coupling numerically method is established based on the conserved form of the three-dimensional unsteady Navier-Stokes (N-S) equations, considering gas fluid with chemical reactions and heat transfer between the fluid and solid region. Experiments are subsequently conducted using high-speed camera to record the ignition process. The flame propagation, chamber pressurizing process and average fuel regression rate of the numerical simulation results show good agreement with the experimental ones, which demonstrates the validity of the simulations in this study. The results also indicate that the flame propagation time is mainly affected by fluid dynamics and it increases with an increasing grain port area. The chamber pressurizing process begins when the flame propagation completes in the grain port. Furthermore, the chamber pressurizing time is about 4 times longer than the time of flame propagation. [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]

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]

This paper is aimed to analyse the combustion characteristics of hybrid rocket motor with multi-section swirl injection by simulating the combustion flow field. Numerical combustion flow field and combustion performance parameters are obtained through three-dimensional numerical simulations based on a steady numerical model proposed in this paper. The hybrid rocket motor adopts 98% hydrogen peroxide and polyethylene as the propellants. Multiple injection sections are set along the axis of the solid fuel grain, and the oxidizer enters the combustion chamber by means of tangential injection via the injector ports in the injection sections. Simulation results indicate that the combustion flow field structure of the hybrid rocket motor could be improved by multi-section swirl injection method. The transformation of the combustion flow field can greatly increase the fuel regression rate and the combustion efficiency. The average fuel regression rate of the motor with multi-section swirl injection is improved by 8.37 times compared with that of the motor with conventional head-end irrotational injection. The combustion efficiency is increased to 95.73%. Besides, the simulation results also indicate that (1) the additional injection sections can increase the fuel regression rate and the combustion efficiency; (2) the upstream offset of the injection sections reduces the combustion efficiency; and (3) the fuel regression rate and the combustion efficiency decrease with the reduction of the number of injector ports in each injection section. [ABSTRACT FROM AUTHOR]

The objective of this effort is to study the combustion performance of a hybrid rocket motor with the help of 3D steady-state numerical simulation, which applies 90% hydrogen peroxide as the oxidizer and hydroxyl-terminated polybutadiene as the fuel. A method of secondary oxidizer injection in post-chamber is introduced to investigate the flow field characteristics and combustion efficiency. The secondary injection medium is the mixed gas coming from liquid hydrogen peroxide catalytic decomposition. The secondary injectors are uniformly set along the circumferential direction of the post-chamber. The simulation results obtained by above model are verified by experimental data. Three influencing parameters are considered: secondary injection diameter, secondary injection angle and secondary injection numbers. Simulation results reveals that this design could improve the combustion efficiency with respect to the same motor without secondary injection. Besides, the secondary injection almost has no effect on the regression rate and fuel sueface temperature distribution. It is also presented that the oxidizer is injected by 8 secondary injectors with a diameter of 7–8 mm in the direction of 120°in post-chamber is identified as the optimized secondary injection pattern, through which combustion efficiency, specific impulse efficiency as well as utilization of propellants are all improved obviously. [ABSTRACT FROM AUTHOR]