Performance Of Hybrid Rocket Motor Research Articles

Transient numerical investigation on thermochemical erosion of C/C nozzles in hybrid rocket motors

Nozzle erosion is a vital important issue that impacts the performance of hybrid rocket motors. Serious nozzle erosion may significantly decrease the combustion chamber pressure and thrust, which increases the difficulty in designing flight control systems. This paper aims to reveal erosion mechanism systematically. In this paper, transient numerical simulations of thermochemical erosion in hybrid rocket motors are conducted, and combustion flow and thermochemical erosion are coupled calculated. The numerical computation of combustion flow is based on the chemical reactions of propellants and regression rate model, and that of thermochemical erosion is based on the surface reactions between oxidizing species and carbon. The movement of burning surface and nozzle inner surface is simulated through dynamic mesh method. The hybrid rocket motor adopts 90% hydrogen peroxide and hydroxyl-terminated polybutadiene. Distributions of flow field parameters and fuel regression rate are given. The spatial developing process of nozzle surface is presented, and it is found that the roughness of nozzle profile increases with time. The nozzle wall temperature and wall pressure decline with time. Erosion by different species is calculated. OH and H2O make a major contribution to the nozzle thermochemical erosion, while nozzle erosion contributed by CO2 and O2 are quite low. Time-varying characteristics of the erosion rate are unveiled. At the first 1.5 s, the total erosion rate remains almost constant, and then it reduces over time.

A comprehensive fluid–solid coupling dynamic simulation for spatiotemporal distribution of regression rate in hybrid rocket motors

The spatiotemporal distribution characteristics of the regression rate are crucial aspects of the research on Hybrid Rocket Motor (HRM). This study presents a pioneering effort in achieving a comprehensive numerical simulation of fluid dynamics and heat transfer in both the fluid and solid regions throughout the entire operation of an HRM. To accomplish this, a dynamic grid technique that incorporates fluid–solid coupling is utilized. To validate the precision of the numerical simulations, a firing test is conducted, with embedded thermocouple probes being used to measure the inner temperature of the fuel grain. The temperature variations in the solid fuel obtained from both experiment and simulations show good agreement. The maximum combustion temperature and average thrust obtained from the simulations are found to deviate from the experimental results by only 3.3% and 2.4%, respectively. Thus, it can be demonstrated that transient numerical simulations accurately capture the fluid–solid coupling characteristics and transient regression rate. The dynamic simulation results of inner flow field and solid region throughout the entire working stage reveal that the presence of vortices enhances the blending of combustion gases and improves the regression rate at both the front and rear ends of the fuel grain. In addition, oscillations of the regression rate obtained in the simulation can also be well corresponded with the corrugated surface observed in the experiment. Furthermore, the zero-dimension regression rate formula and the formula describing the axial location dependence of the regression rate are fitted from the simulation results, with the corresponding coefficients of determination (R2) of 0.9765 and 0.9298, respectively. This research serves as a reference for predicting the performance of HRM with gas oxygen and polyethylene, and presents a credible way for investigating the spatiotemporal distribution of the regression rate.

Numerical analysis on combustion characteristics of hybrid rocket motor with star-tube segmented grain

A method of star-tube combined segmented grain is proposed to improve the combustion performance of hybrid rocket motor. The star-tube combined segmented grain consists of a single-port star part and a single-port tube part. A mid-chamber forms between the fore-grain and the aft-grain for better mixing effect. The single-port feature gives hybrid rocket motor several advantages, such as simple structure, high reliability, and variable combinations. This paper is mainly aimed at studying the combustion characteristics of hybrid rocket motor with star-tube segmented grain through three-dimensional steady simulations. Combustion performance of the motors with different segmented grain combinations, including fore-tube/aft-tube, fore-tube/aft-star, fore-star/aft-star and fore-star/aft-tube, is contrastively analyzed. The motor in this paper adopts polyethylene and 90% hydrogen peroxide as the propellants. Simulations reveal that segmented grain with different-type grain combinations could greatly change the flow field in the second half of the combustion chamber. Transformation of the flow field is beneficial to the mixing between the fuel and the oxidizer, and it could increase the fuel regression rate and the combustion efficiency. The turbulence effect of tube aft-grain is better than that of star aft-grain. Among the four segmented grain combinations, the combination of star fore-grain and tube aft-grain is the preferred method with optimal overall performance. This grain configuration could increase the regression rate of tube aft-grain to surpass that of star aft-grain in other combinations. Besides, hybrid rocket motor with this grain configuration achieves the highest combustion efficiency.

Effect of High-Mass Fraction of Aluminum on Catalytic Hybrid Rocket Motor

Catalytic hybrid rocket motors using hydrogen peroxide (HP) are easy and efficient to achieve multiple starts and stops, and hydroxyl-terminated polybutadiene (HTPB) grains are commonly used due to their excellent mechanical properties. The low regression rate of HTPB grains limits the application of hybrid rocket motors. Furthermore, the addition of solid aluminum particles can effectively improve the regression rate and performance of hybrid rocket motors. However, the experimental results of a high mass fraction of aluminum and hydrogen peroxide available at present are not sufficient. In this research, the impact of a high mass fraction of aluminum on the motor performance and ablation rate of nozzles is studied experimentally. A solution of 95% hydrogen peroxide and HTPB with an aluminum additive are adopted as propellants. The variation in the axial regression rate of the grains is obtained by computed tomography (CT) scans and pre-test parameter measurements. The instantaneous regression rate method is adopted to obtain the real-time regression rate of the motor. The surface appearance and composition of the front and the end of the grains after the tests are analyzed by electron microscopy. Carbon ceramic and tungsten-bronze nozzles are used to explore the effect of a high mass fraction of aluminum on nozzle ablation. The experimental results show that the addition of aluminum raises the specific impulse and decreases the optimal oxygen-to-fuel ratio of the propellant combination. The high mass fraction of the aluminum particles has a severe ablative effect on carbon ceramic nozzles, while the effect on tungsten-bronze nozzles is minimal for a hot test lasting four to five seconds. Our results can provide experimental guidance for the application of a high mass fraction of aluminum and hydrogen peroxide hybrid rocket motor.

Dynamic Numerical Simulation of Hybrid Rocket Motor with HTPB-Based Fuel with 58% Aluminum Additives

The addition of aluminum (Al) to the fuel is an effective way to increase the regression rate of hybrid rocket motors (HRMs). Due to its high regression rate, the impact of the regression of combustion surface on the performance of HRMs cannot be ignored. Therefore, it is significant to establish a dynamic numerical simulation model to predict the performance of HRMs. In this study, the dynamic simulation model was established based on dynamic mesh technology and was verified by a firing test. The results show that the simulation results agree well with the experimental results, and the errors of the average thrust and combustion chamber pressure are 3.4% and 1.4%, respectively. The dynamic simulation shows that with the regression of the combustion surface, the vortex of the pre-combustion chamber is divided into two vortices. The vortex near the front of the grain will increase the regression rate downstream. The results show that the addition of Al can obviously improve the regression rate of HRMs. The fuel containing 58% Al can improve the regression rate by 88.8% compared with the fuel with pure hydroxyl-terminated polybutadiene (HTPB). Moreover, due to the higher combustion temperature and the scouring of metal particles, the ablation rate of the nozzle with carbon ceramic materials reaches 0.16 mm/s. This investigation provides a valuable reference for HRMs design and simulation.

Effects of aluminum and aluminum hydride additives on the performance of hybrid rocket motors based on 95% hydrogen peroxide

This paper aims to study experimentally and numerically the effects of aluminum (Al) and aluminum hydride (AlH3) 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% AlH3 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 AlH3 and Al are performed to investigate the effects on regression rate and combustion performance of the HRM. Due to the addition of H2, the thermal conductivity increases with the decrease of the mass fraction of AlH3, which leads to the increase of the regression rate. The regression rate of the fuel with 8% AlH3 and 50% Al increases by 35.5% than the fuel with 58% AlH3, demonstrating that the effect of Al on the regression rate enhancement is more than that of AlH3. The theoretical specific impulse increases with the mass fraction of AlH3. However, it is owing that the part of H2 does not participate in the reaction with O2 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.

Effect of metal and metalloid solid-fuel additives on performance and nozzle ablation in a hydroxy-terminated polybutadiene based hybrid rocket motor

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.

Numerical analysis of the hybrid rocket motor with axial injection based on oxidizer flow distribution

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.

Numerical and experimental investigation of throttleable hybrid rocket motor with aerospike nozzle

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.

Effect of multi-location swirl injection on the performance of hybrid rocket motor

An attempt has been made in the present study to realize the effect of multi-location swirl injection on the improvement of regression rate and combustion efficiency of hybrid rocket by using two different length to diameter (L/D) ratio of the motor. It is 12 and 24. The fuel used for the study was Polyvinyl chloride (PVC)-Dibutyl Phthalate (DBP) in the ratio of 50:50. The gaseous oxygen was used as the oxidizer. The significance of this method is that a single swirl injection at the head end of the hybrid rocket is not sufficient to provide swirling flow till the nozzle end, if the length of the motor is sufficiently long and the regression rate decreases beyond certain length. In the present study, it has been observed that the 220 mm length is sufficient to get the improvement with single swirling injection. For the 24 L/D motor, the use of multi-location swirl injector motor gave twice the improvement in regression rate compare to the conventional hybrid rocket motor where showerhead injector has been used. Combustion efficiency obtained with multi-location swirl injection gave lower improvement with 24 L/D motor (from 70% to 76%), but the improvement was significant with 12 L/D motor with single swirl injector (from 50% to 74%). Improvement in regression rate as well as the efficiency was observed to be significant, if the mid injector was used as the cavity instead of supplying oxygen through it. The value of mass flux exponent (n) was observed to be 0.51 with multi-location swirl injection and the relatively uniform regression rate from head end to the nozzle end was observed.

Performance comparison of oxidizer injectors in a 1-kN paraffin-fueled hybrid rocket motor

To investigate the effects of oxidizer injection on the performance of hybrid rocket motors (HRMs), we have designed, manufactured and tested four types of injector: showerhead (SH), hollow-cone (HC), pressure-swirl (PSW) and vortex (VOR). This study is motivated by the fact that the experimental measurements of N2O/paraffin firings are poorly presented in the open literature. Besides few publications are dedicated to the characterization of novel types of injectors in hybrid rocket propulsion application such as HC, PSW, and VOR. It is advantageous that the study was conducted in the same motor configuration, with the advantage that it allowed to compare the performance of different types of injectors.This paper analyzes the influence of the oxidizer injector design on the main performance parameters, such as fuel regression rate, specific impulse and combustion efficiency. First, in order to observe injector spray qualities, a series of cold tests using liquid water and liquid nitrous oxide are carried out, providing a good understanding of the spray profiles. Then, the motor performance data is obtained by a series of firing tests using N2O as oxidizer and paraffin as fuel. Comparison of the various injectors data is made with the same average oxidizer mass flux and feeding pressure. The showerhead is used as a benchmark in this study. During this experimental analysis, the VOR injector exhibits the highest regression rate, followed by HC and SH. Because the assessment of the regression rate was not enough to explain all the effects of the injectors on the motor performance, firing tests of small-scale hybrid motor have been carried out. In terms of spray properties, the PSW exhibits significant differences compared with the other injectors; it generates the smallest Sauter Mean Diameter (SMD) in the formed spray and achieves good atomization. In spite of the fact that the PSW injector leads to the lowest regression rate, it provides good specific impulse, increases the oxidizer-to-fuel ratio (O/F), as well as a uniform and smooth consumption of the paraffin fuel grain. VOR leads to the highest specific impulse. In terms of stability, VOR, HC and SH injectors exhibit lower oscillations in the chamber pressure. Some observations are made on exhaust plume intensity developed during combustion, and in firing tests with SH a blow-out phenomenon occur often. Similarly to liquid engines, it is possible in hybrid motors to increase the global propulsive performance using alternative designs of the injection system.

712 再使用型ハイブリッドロケットシステム基本設計と飛行性能予測(OS7-5 ハイブリッドロケット(システム,回収系))(OS7 北海道における航空宇宙工学研究の進展と未来)

Hybrid rocket motors, propelled by a combination of a solid fuel and a liquid oxidizer, are best suited to development of small launch systems in university laboratories, because of their advantageous characteristics such as safety, easy handling, and low costs. The performance of hybrid rocket motors of three classes of thrust - 10tonf, 1tonf, and 200kgf- is estimated. The feasibility of reusable winged launch systems based on the three motors is preliminarily analyzed for suborbital micro-gravity experiments. The results tell that a system with a 10tonf-class motor by a coolant bleed cycle with polystyrene and LOx fed by an LE-5B turbopump will be capable of exposing a payload of 360kg to a micro-gravity environment for three minutes. It is also predicted that a system with a 1tonf-class motor will be moderately capable and that with a 200kgf-class motor will not be feasible for micro gravity missions mainly because the weight of its helium pressurization system for feeding LOx will reduce the mass fraction.

Theoretical investigation of hybrid rocket combustion by numerical methods

A numerical method for calculating the flow field and the performance of hybrid rocket motors is presented. It is based on a model in which the flow has boundary layer character, with the entrance region, where this assumption does not hold, represented by appropriate initial conditions. The coundary layer equations are solved by a finite difference method, giving velocity, temperature, and concentration profiles in addition to the regression rate. A finite rate of reaction is assumed and a thin liquid sublayer is introduced at the fuel surface. The influence of various parameters (mass flow rate, pressure, viscosity, radiation, thickness of liquid layer, and initial conditions) on the regression rate, the flow velocity, the temperature, and the mass fractions has been investigated. The conclusions, which are in good agreement with experimental evidence, are that the flame is of diffusive type, as shown by the existence of different zones for the oxidizer and the fuel, divided by a reaction zone of appreciable thickness. The pressure influences the combustion process for low pressure and high mass flow rate conditions. The regression rate is a function of the mass flow rate raised to the n th power (with n = 0.3 for laminar flow, rising to 0.8 for fully turbulent flow) and of the axial distance, raised to the m th power (with m between −0.2 and −0.35). Good agreement was obtained between the computations and experimental results, obtained, in a two-dimensional motor.