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The electric pump integrated into propellant supply system boasts the expansive capacity for swift thrust adjustments, making it an indispensable asset for optimizing the performance of a hybrid rocket motor. Although previous studies have explored the efficacy and practicability of hydrogen peroxide electric pumps in hybrid rocket motors, the implications of their filling process on start-up have remained elusive. In this study, the filling process and its effect on the start-up of the electric pump were numerically and experimentally investigated. The experimental and simulation results show that the electric pump cannot be completely filled by opening only the valve upstream of the pump and that there are still areas of air in the pump after the filling process. The results of the start-up process of the electric pump with different filling conditions indicate that the air, filling and acceleration process increase the pressure and axial force fluctuation, decrease the efficiency and compromise the operational stability of the pump after start-up. The results of the filling process of the electric pump at 2000 rpm demonstrate that filling the pump after starting it further reduces its efficiency and operational stability. This study provides practical guidance for filling electric pump supply systems and starting electric pumps of the rocket engines, which ensures the stable operation of rocket engines. • The filling process of electric pump in a hybrid rocket motor was studied by simulation and test. • The start-up process of the electirc pump with different filling conditions was investigated. • The filling process of the electric pump at 2000 rpm was investigated. [ABSTRACT FROM AUTHOR]

This paper attempts to research ignition delay time and ignition process of catalytic hybrid rocket motor (HRM) with 95% hydrogen peroxide (95HP)/HTPB-based aluminum additive fuel. Firing tests are conducted with 95HP and different fuels on lab-scale HRMs. The aluminum weight fractions of four HTPB-based fuels are 38%, 48%, 58% and 65% respectively. Corresponding two-dimensional axisymmetric transient simulations are established to verify the relationship of catalytic ignition delay to fuel grain aluminum weight fraction and to catalytic decomposition temperature. The mono-propellant catalysis experiments of 95HP are carried out and show that the maximum temperature of decomposition products can reach 1140 K. The catalytic ignition delay time of 95HP/HTPB-based fuels is affected by the aluminum additive. When the aluminum weight fraction of the fuels is increased from 38% to 65%, the effective heating time of the ignition is correspondingly increased from 0.250 s to 0.556 s. The flame pictures are taken by the high-speed camera in the firing tests. A large number of sparks appear in the flame when it reaches a stable state, but there is no spark in the flame at the initial stage of ignition. Therefore, it is indicated that the aluminum does not participate in the chemical reaction at the initial stage of ignition, but may change the physical properties of the fuel. Similarly, the simulation results prove that the increase of the aluminum weight fraction of the fuel will increase the ignition delay time while the increase of catalytic decomposition temperature will greatly reduce the ignition delay time. • Firing tests are conducted with HTPB-based aluminum additive fuel. • Increasing of the aluminum fraction of fuel will increase the ignition delay time. • Aluminum does not participate in the chemical reaction before the fuel is ignited. • Two-dimensional axisymmetric transient simulations are conducted. [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]

The purpose of this paper is to characterize the regression rate behavior of hybrid rocket motor propellant combinations, using hydrogen peroxide (HP), gaseous oxygen (GOX), nitrous oxide (N 2 O) as the oxidizer and hydroxyl-terminated poly-butadiene (HTPB) as the based fuel. In order to complete this research by experiment and simulation, a hybrid rocket motor test system and a numerical simulation model are established. Series of hybrid rocket motor firing tests are conducted burning different propellant combinations, and several of those are used as references for numerical simulations. The numerical simulation model is developed by combining the Navies–Stokes equations with the turbulence model, one-step global reaction model, and solid–gas coupling model. The distribution of regression rate along the axis is determined by applying simulation mode to predict the combustion process and heat transfer inside the hybrid rocket motor. The time-space averaged regression rate has a good agreement between the numerical value and experimental data. The results indicate that the N 2 O/HTPB and GOX/HTPB propellant combinations have a higher regression rate, since the enhancement effect of latter is significant due to its higher flame temperature. Furthermore, the containing of aluminum (Al) and/or ammonium perchlorate(AP) in the grain does enhance the regression rate, mainly due to the more energy released inside the chamber and heat feedback to the grain surface by the aluminum combustion. [ABSTRACT FROM AUTHOR]

The fuel regression rate is an important parameter in the design process of the hybrid rocket motor. Additives in the solid fuel may have influences on the fuel regression rate, which will affect the internal ballistics of the motor. A series of firing experiments have been conducted on lab-scale hybrid rocket motors with 98% hydrogen peroxide (H 2 O 2 ) oxidizer and hydroxyl terminated polybutadiene (HTPB) based fuels in this paper. An innovative fuel regression rate analysis method is established to diminish the errors caused by start and tailing stages in a short time firing test. The effects of the metal Mg, Al, aromatic hydrocarbon anthracene (C 14 H 10 ), and carbon black (C) on the fuel regression rate are investigated. The fuel regression rate formulas of different fuel components are fitted according to the experiment data. The results indicate that the influence of C 14 H 10 on the fuel regression rate of HTPB is not evident. However, the metal additives in the HTPB fuel can increase the fuel regression rate significantly. [ABSTRACT FROM AUTHOR]

Orifice injector plates are widely used in hybrid rocket motors (HRM), but little work has been done on the influences of their design parameters on combustion performances of HRM. The effects of orifice diameter, injection pressure drop and orifices distribution of the orifice injector plate on combustion efficiency of HRM are investigated by experimental and numerical methods in this paper. The hydrogen peroxide (H 2 O 2) and hydroxyl terminated polybutadiene (HTPB) with tube grain are selected as propellant. Firstly, 5 kinds of orifice injector plates with different orifice diameter and orifices distribution are designed, and 7 fire experiments are carried out. The experimental results show that the combustion efficiency of HRM increases with the decrease of orifice diameter and injection pressure drop. Moreover, the combustion efficiency of motor with plat using single-ring distribution is the highest, while the one using three-ring distribution is the lowest. Secondly, 36 orifice injector plates with different design parameters are calculated by three-dimensional numerical simulation of two-phase flow. The liquid oxidizer evaporation, solid-gas interaction, gaseous phase combustion, and turbulence are considered in the simulation model. The conclusions draw from simulation results agree with the experimental data, which illustrates the validity of numerical simulation model. The conclusions based on experimental and numerical results can provide guidance for designing high-performance orifice injector plate in HRMs. • Numerical and experimental studies on orifice injector plates are conducted. • The effects of design parameters of orifice injector plate on combustion efficiency are investigated. • The combustion efficiency increases with the decrease of orifice diameter. • The combustion efficiency rises with the decrease of injection pressure drop. • The orifices distribution should be choose based on mission requirements of HRMs. [ABSTRACT FROM AUTHOR]

This paper is aimed to analyze heat transfer in injector plate of hydrogen peroxide hybrid rocket motor by two-dimensional axisymmetric numerical simulations and full-scale firing tests. Long-time working, which is an advantage of hybrid rocket motor over conventional solid rocket motor, puts forward new challenges for thermal protection. Thermal environments of full-scale hybrid rocket motors designed for long-time firing tests are studied through steady-state coupled numerical simulations of flow field and heat transfer in chamber head. The motor adopts 98% hydrogen peroxide (98HP) oxidizer and hydroxyl-terminated poly-butadiene (HTPB) based fuel as the propellants. Simulation results reveal that flowing liquid 98HP in head oxidizer chamber could cool the injector plate of the motor. The cooling of 98HP is similar to the regenerative cooling in liquid rocket engines. However, the temperature of the 98HP in periphery portion of the head oxidizer chamber is higher than its boiling point. In order to prevent the liquid 98HP from unexpected decomposition, a thermal protection method for chamber head utilizing silica-phenolics annular insulating board is proposed. The simulation results show that the annular insulating board could effectively decrease the temperature of the 98HP in head oxidizer chamber. Besides, the thermal protection method for long-time working hydrogen peroxide hybrid rocket motor is verified through full-scale firing tests. The ablation of the insulating board in oxygen-rich environment is also analyzed. [ABSTRACT FROM AUTHOR]

The objectives of this study are to develop a more accurate regression rate considering the oxidizer mass flow and the fuel grain geometry configuration with numerical and experimental investigations in polyethylene (PE)/90% hydrogen peroxide (HP) hybrid rocket. Firstly, a 2-D axisymmetric CFD model with turbulence, chemistry reaction, solid–gas coupling is built to investigate the combustion chamber internal flow structure. Then a more accurate regression formula is proposed and the combustion efficiency changing with the length–diameter ratio is studied. A series experiments are conducted in various oxidizer mass flow to analyze combustion performance including the regression rate and combustion efficiency. The regression rates are measured by the fuel mass reducing and diameter changing. A new regression rate formula considering the fuel grain configuration is proposed in this paper. The combustion efficiency increases with the length–diameter ratio changing. To improve the performance of a hybrid rocket motor, the port length–diameter ratio is suggested 10–12 in the paper. [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]