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The hybrid rocket motor (HRM) emerges as a promising propulsion system offering thrust adjustment at a low cost. Prior research has indicated that the inclusion of micro Al additives enhances HRM performance, yet a comprehensive investigation remains absent. Thus, this study aims to explore the impact of Al particle additive content and size variation on the characteristics of hybrid rocket motors (HRMs) through experiments and simulations. Firing tests are conducted on HRMs employing unadulterated HTPB, 38 % additive Al (2 μm), and 58 % Al (2 μm). Experimental findings reveal enhancements of 12.6 % and 33.9 % in regression rate, and 2.5 % and 8.0 % in specific impulse for the 38 % Al and 58 % Al fuel configurations, respectively, compared to the base case. Subsequently, a numerical two-phase combustion model is built to analyze the inner flow field characteristics. The high consistency between experimental and simulation results validates the efficacy of the numerical approach employed. Following this, a series of simulations covering Al content ranging from 0 % to 58 % and Al particle sizes from 0.4 μm to 75 μm is conducted. Simulation outcomes indicate that specific impulse and combustion efficiency enhancements correlate with the increase in Al mass fraction for particles with diameters under 35 μm. However, high Al content proves detrimental to specific impulse and combustion efficiency when particle diameter exceeds 35 μm. Additionally, regression rate and thrust demonstrate an increase with Al content, albeit the rate of increment diminishes with larger Al particles. This study comprehensively examines the influence of Al particle addition on HRM performance and elucidates the underlying mechanisms, thereby offering insights into HRM design and manufacture. • Hybrid rocket firing tests are performed on HTPB, 38 %, and 58 % Al particle fuel. • Simulations with various Al particle contents and diameters are performed. • High Al content with diameter under 35 μm is beneficial for combustion efficiency. • High Al content with diameter over 35 μm is detrimental for combustion efficiency. • Large particle reduces the thrust enhancement brought by Al content rise. [ABSTRACT FROM AUTHOR]

This paper focuses on investigating ablation characteristics of different nozzle materials in hybrid rocket motors by firing tests and numerical simulations. Hybrid rocket nozzles with different ablation materials are designed and manufactured. Four firing tests with working time of 40 s are conducted, and 90% hydrogen peroxide and polyethylene are adopted. The test results indicate that erosion rates of T705 graphite, whole felt C/C composite, axial rod braided C/C composite, and carbon/ceramic composite are 0.045 mm/s, 0.022 mm/s, 0.03 mm/s, and 0.01 mm/s. Circumferentially uneven ablation is found in the nozzle throat of T705 graphite and C/C composite, and this is mainly due to asymmetric oxidizer injection, asymmetric combustion, and uneven material properties. Transient numerical simulations of 40 s on nozzle thermochemical erosion are carried out, and the regression process of fuel wall and nozzle wall is solved by a dynamic grid technique. The error of average throat erosion rate between simulation results and test results is only 1.82%. The evolution process of nozzle profile and flow field structure are exhibited, and coupling effects between the erosion rate and nozzle profile are obtained. Simulation results show that as the working time increases, protrusions and pits appear on the nozzle inner wall and gradually extend. The temperature of nozzle wall and nozzle erosion rate on the windward side is higher, while they are lower on the leeward side. Nozzle throat erosion rate history is divided into four stages, and time nonlinear characteristics of the throat erosion rate are revealed. • 40 s firing tests of different nozzle materials are conducted. • Ablation properties of graphite, C/C composite and carbon ceramic are obtained. • Transient simulations of 40 s on nozzle thermochemical erosion are performed. • Simulation results agree with test data and evolution of nozzle profile is shown. • Time nonlinear characteristics of nozzle throat erosion rate are revealed. [ABSTRACT FROM AUTHOR]

The electric pump in the propellant supply system, with a wide range of fast thrust adjustment capabilities, can significantly enhance rocket motor performance. Previous research has focused on the performance and feasibility of hydrogen peroxide electric pumps for hybrid rocket motors; however, their dynamic characteristics remain unclear. This study numerically and experimentally investigated the transition stages in the flow field and head for an electric pump during the start-up of a hybrid rocket motor, as well as analyzed the effect of pipeline length between the pneumatic valve and the adjustable Venturi on the flow field and head. The simulation and experimental results indicate that during the start-up of the hybrid rocket motor, six transient stages are recorded when the pipeline length between the pneumatic valve and adjustable Venturi is 3 m. The head and internal flow field of the electric pump during the transient process were determined by the throttling position as well as by the filling status of the pipeline. The simulation results suggest that the shorter the pipeline length, the steadier and safer is the supply system. Our investigation provides a practical optimization strategy for electric pump supply systems in rocket motors. • The transient process of the electric pump during start-up of the hybrid rocket motor was obtained. • Physical mechanisms for the variation of head and internal flow field were obtained. • The effect of pipeline length on the transient process of the electric pump was obtained. [ABSTRACT FROM AUTHOR]

Precise regulation of oxidizer mass flow rate is a crucial technology for hybrid rocket motors, and a flow control valve serves as a core component of an oxidizer delivery system. In this paper, we designed a variable area cavitating venturi (VACV) as a flow control valve to control oxidizer mass flow rate. Three types of experiments were carried out to investigate VACA-based flow feedback control and its application in hybrid rocket motors. Specifically, we measured flow characteristics at different pintle strokes and conducted five cold tests and one ground hot test based on flow feedback control. Test results are as follows: first, the VACA flow characteristic curve shows a wide regulation range with a flow regulation ratio of 21.8 and performs good linearity of 96.76%. Second, the mass flow rate curve measured with flow feedback control tracks the target curve. Flow feedback control effectively improves control accuracy by maintaining the steady-state error within 1.2%. Third, hot test results demonstrate that the VACA with flow feedback control can effectively maintain the accurate and consistent oxidizer mass flow rate during the firing stage of the hybrid rocket motor. The findings of this research indicate that VACA with flow feedback control can be used as an effective flow regulation component to maintain a constant and accurate oxidizer mass flow rate during the working stage of a hybrid rocket motor. • A firing test of a hybrid rocket motor based on feedback control was conducted. • VACV shows good linearity with a wide regulation range in flow characteristics curve. • Feedback control with VACV improves accuracy, keeping steady-state error within 1%. • Feedback control maintains an accurate and consistent oxidizer mass flow rate. [ABSTRACT FROM AUTHOR]

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]

To improve the ablation resistance under long-time working conditions, a multiple interface nozzle with a carbon ceramic throat cap and a copper infiltrated tungsten throat insert is designed. Firstly, a transient numerical model on thermochemical erosion (considering surface reactions of carbon and tungsten) of the multiple interface nozzle in a hybrid rocket motor is established, and the regression of nozzle inner profile is simulated by adopting a dynamic grid technique. Then a 200 s firing test of the full-scale hybrid rocket motor with the multiple interface nozzle is conducted, and 98% hydrogen peroxide and hydroxyl-terminated polybutadiene based fuel are adopted. Erosion steps are formed at the interface of the carbon ceramic and copper infiltrated tungsten, and its formation mechanism is discussed. The erosion step height on the windward side of the throat insert is larger than that on the leeward side of the throat insert. The nozzle inner profile after the test agrees well with the simulation results, and the error of the nozzle throat diameter between the simulation results and test data is about 3.45%, which verifies the accuracy of numerical models. Finally, effects of erosion steps on the ablation characteristics and thermal protection properties are studied by numerical simulations. The results indicate that the erosion rate on the leeward side is much lower than that on the windward side. On the leeward side, the obvious radial diffusion of the gas occurs and the erosion rate increases with the increase of the erosion step height. Local vortices at the position of erosion steps will improve the heat transfer coefficient. When the erosion step height increases, the temperature on the outside surface of the nozzle shell improves, and it is increased by 38.3 K on the shell surface in the nozzle throat at t = 200 s. Research findings in this paper can provide important support for the engineering application of hybrid rocket motors with multiple interface nozzles. • Transient simulations on thermochemical erosion of C and W are performed. • 200 s firing test of hybrid rocket motor with multiple interface nozzle is conducted. • Simulation results agree well with test data and erosion step is found. • Erosion rate on the leeward side is much lower than that on the windward side. • Nozzle shell temperature increases with the increase of erosion step height. [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 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 aims at the dynamic characteristics study of variable thrust HRMs, for which a two-dimensional transient numerical model of HRM is established. Complete transient simulation based on dynamic mesh has been performed sequentially. An HRM with a thrust of three stages is designed and tested. The pressure, regression rate and thrust of the simulation results are in good agreement with the experiment data, which verifies the feasibility of the dynamic mesh technique for fuel regression simulation in HRMs. Specifically, the maximum deviations of pressure and thrust are less than 3.6% and 3.2%, respectively at the large mass flow rate stage. The numerical simulations show that the regression rate is not only related to the oxidizer mass flux, but also affected by the dynamic regression process. Max deviation of regression rate is 35% at the same oxidizer mass flux. Moreover, it is also demonstrated that the presence of vortex near front edge of fuel grain has a great effect on motor performance. Regression rate at front part of solid fuel is significantly high at the initial stage of throttling process, and regression rate downstream of the vortex obviously increases with time. In addition, the gas temperature of the pre-combustion chamber continuously increases with regression of fuel surface. Gas temperature of the post-combustion chamber is positively correlated with the oxidizer mass flow rate. This research can predict performance of HRMs more accurately, and provide a valuable guidance for variable thrust HRMs design. • A numerical simulation model of variable thrust hybrid rocket motor is established. • Dynamic mesh technique is used to simulate the regression of solid fuel surface. • A firing test of variable thrust hybrid rocket motor are conducted. • Average regression rate is different at the same oxidizer mass flux. • Regression rate of downstream of the vortex obviously increases with time. [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]

Numerical simulation based on the discrete phase model and dynamic mesh model can more accurately simulate the working process for hybrid rockets, including the motion and evaporation process of particles and regression process of the fuel grain. The dynamic numerical simulation of combustion and flow in hybrid rocket motors based on a discrete phase model is carried out. Aluminum and aluminum hydride are added to the internal flow field as discrete droplets. Then, a hybrid rocket is designed with 95 % hydrogen peroxide as oxidizer and 38 % aluminum hydride+20 % aluminum+42 % hydroxyl-terminated polybutadiene as the fuel and used to verify the model by firing test. A high degree of agreement between experiment and simulation findings is found. The average deviations for thrust and pressure are 3.6 % and 5.9 %, respectively. It offers the required backing for using the simulation method to investigate the flow and combustion process for the hybrid rockets. The results also demonstrate the transient characteristics of the inner flow field, including the separation and merging of vortexes in the pre-chamber. It is found that a peak of the regression rate appears at the front of the grain and moves backward. This is mainly because the vortex appears here and enhances its heat transfer. Through the discrete phase model, it is worth noting that the aluminum droplet content gradually decreases at the nozzle over time, which is consistent with the firing test. Further research on the impact of aluminum hydride content on the performance of the hybrid rockets has been carried out. The results show that the higher aluminum hydride content can lead to better performance for the hybrid rockets. The thrust, pressure, specific impulse, and characteristic velocity enhance with the increase of aluminum hydride content. An interesting discovery is that the average regression rate is maximum at 48 % aluminum hydride content. This research innovatively couples dynamic mesh and discrete phase models, which can help better understand the working process for hybrid rockets and provide important guidance for the selection of fuel formula. • A simulation model of hybrid rocket motors based on a discrete phase model is established. • The motion and evaporation of particles and regression of the fuel grain are simulated. • A test of hybrid rockets with fuel containing aluminum and aluminum hydride is carried out. • A high degree of agreement between experiment and simulation findings is found. • The impact of aluminum hydride content on the performance of the hybrid rockets is 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]

This study is aimed to determine and better reveal the mixture enhancement and regression rate distribution of hybrid rocket motor with diaphragm by numerical approach. A numerical model based on the computational fluid dynamics software is built to simulate the flow and combustion inside the motor. Four firing tests of the motor, including one without diaphragm and three with diaphragm, are conducted on a standard experimental system and also used as a reference for numerical simulation, the consistency between the simulation and experiment demonstrates that the numerical approach is an effective method to study the diaphragm effect on the motor performance. The flow field characteristic and regression rate distribution inside the hybrid rocket motor are then calculated to analyze the effect of position and diameter of the diaphragm. The results indicate that the diaphragm almost have no effect on the regression rate before it. However, the regression rate after the diaphragm has a strong dependence on the position and diameter of the diaphragm. As the diameter decreases and the position moves backward, the regression rate increases larger and larger, this is mainly due to the augmentation of the eddy generated by the diaphragm, which enhances the heat feedback transferred to the grain surface. When the diameter of diaphragm located at middle of grain decreases from 50 mm to 20 mm, regression rate is increased from 0.30 mm/s to 0.57 mm/s. The use of the diaphragm does cause a combustion efficiency improvement; the maximum combustion efficiency is enhanced to 98.9% from lower than 90% of the motor with no diaphragm. The increasing amplitude displays a square relation with the diameter decrease, since the entrainment of the eddy make the reactants mix sufficiently to release more energy inside the motor. [ABSTRACT FROM AUTHOR]

This paper focuses on the research on regression rate of hybrid rocket motor with single-port wagon wheel fuel grain. Both experiment and numerical investigations are carried out to study the three-dimension regression rate of hybrid rocket motor on several characteristic points. A laboratory 500 N-class firing test of hybrid rocket motor with single-port wagon wheel fuel grain has been conducted on the standard ground experimental system. The 98% hydrogen peroxide oxidizer and hydroxyl terminated polybutadiene based fuel are adopted as propellants. The digital image processing of three-dimension regression rate is carried out, and a post-processing software has been developed to calculate the time-averaged three-dimension distribution of regression rate after the firing test. Three-dimensional numerical simulations with gaseous combustion considered are carried out to study the axial distribution of regression rate. The results of digital image processing manifest that the average regression rates of the middle position of wagon wheel spokes is higher than those in the slot of wagon wheel. A good agreement exists between the calculated regression rate from numerical simulation and the experimental results conducted by computed tomography scanning. The comparison results show that the digital image processing is potentially an effective tool for investigating and predicting the three-dimension regression rate of hybrid rocket motor with single-port wagon wheel fuel grain. • Firing test of HRM with single-port wagon wheel fuel grain has been carried out. • The method of CT scanning digital image processing and software are developed. • Three-dimension regression rates of HRM are characterized by simulation and test. • A good agreement exists between regression rates from simulation and CT scanning. [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]

Abstract: This paper presents three-dimensional numerical simulations and experimental studies of the hybrid rocket motor with multi-port fuel grain. The numerical model is established based on the Navier–Stokes equations with turbulence, chemical reactions, fuel pyrolysis, and solid–gas boundary interactions. The simulation is performed based on the 98% hydrogen peroxide (HP) and hydroxyl terminated polybutadiene (HTPB) propellant combination. The results indicate that the flow field and fuel regression rate distributions present apparent three-dimensional characteristics. The fuel regression rates decrease first and then gradually increase with the axial location increasing. At a certain cross section, the fuel regression rates are lower in the points on arcs with smaller radius of curvature when the fuel port is a derivable convex figure. Two experiments are carried out on a full scale motor with the simulation one. The working process of the motor is steady and no evident oscillatory combustion is observed. The fuel port profiles before and after tests indicate that the fuel regression rate distributions at the cross section match well with the numerical simulation results. [Copyright &y& Elsevier]

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]

Chemical ablation and mechanical erosion of hybrid rocket nozzles can cause loss of nozzle thermal protection material, resulting thrust reduction. In this paper, the chemical ablation and mechanical erosion mechanisms of nozzle thermal protection materials are studied through numerical simulation. For chemical ablation, a coupled thermal-mechanical-chemical model is established, considering three mechanisms: wall regression caused by non-homogeneous chemical reactions at the wall, convective heat transfer and fluid-structure interaction. For mechanical erosion, a Fluent DPM model and an Abaqus DEM model are established, which are investigated from the perspectives of empirical formula and solid mechanics mechanisms respectively, and validated against each other. The comprehensive effect of chemical ablation and mechanical erosion on the morphology of the nozzle is considered. The results show that both chemical ablation and mechanical erosion dominate the wall regression rate at the downstream of the converging section and upstream of the throat while the other location is only dominate by chemical ablation. After 5s of operation, thrust of rocket engine decreases by 5.8%. The Abaqus DEM model presented by this paper breaks the limitations of the empirical formulae due to the introduction of the material constitutive model and the failure criterion. • A coupled thermal-mechanical-chemical model is established. • Mechanical erosion is studied from perspective of empirical formula and the solid mechanics mechanism. • Comprehensive effect of chemical ablation and mechanical erosion on nozzle inner wall morphology is studied. [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]