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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 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]