Hybrid Rocket Motor Research Articles

Regression Rate Enhancement of Hybrid Rockets by Introducing Novel Distributed Tube Injector

Low fuel regression rate is one of the major drawbacks of hybrid rocket motors. In this study, a novel injector concept is proposed to provide a substantial enhancement in the fuel regression rate. A tubular injector, Distributed Tube Injector (DTI), is inserted in the center of the cylindrical fuel port in order to tailor the oxidizer flow introduced into the combustion chamber with the desired combination of radial, tangential, or axial components of velocity. This concept has been tested using a 500 N thrust class hybrid rocket motor, which uses a paraffin-based fuel and supercharged (L) as the oxidizer. As a result of 30 hot firings conducted in the test program, it is determined that the DTI configuration provides regression rates up to 3.9 times higher than the regression rates obtained using fore-end injector commonly employed in conventional hybrid rockets. Using the motor test data, a comprehensive nondimensional and scalable regression rate relation has been established. This nondimensional regression rate equation can be used to design the internal ballistic configuration of hybrid rocket motors using the novel injector.

Regression rate performance of paraffin wax and bees wax –Gox with varied grain configuration in a hybrid rocket motor

Purpose This paper aims to determine the regression rate using wax fuels for three different grain configurations and find a suitable grain port design for hybrid rocket application. Design/methodology/approach The design methodology of this work includes different grain port designs and subsequent selection of solid fuels for a suitable hybrid rocket application. A square, a cylindrical and a five-point star grained were designed and prepared using paraffin and beeswax fuels. They were tested in a laboratory-scale rocket with gaseous oxygen to study the effectiveness of solid fuels on these grain structures. The regression rate by static fire testing of these wax fuels was analyzed. Findings Beeswax performance is better than that of paraffin wax fuel for all three designs, and the five-slotted star fuel port grain attained the best performance. Beeswax fuel attained an average regression rate ≈of 1.35 mm/s as a function of oxidizer mass flux Gox ≈ 111.8 kg/m2 s and for paraffin wax 1.199 mm/s at Gox ≈ 121 kg/m2 s with gaseous oxygen. The local regression rates of fuels increased in the range of 0.93–1.194 mm/s at oxidizer mass flux range of 98–131 kg/m2 s for cylindrical grain, 0.99–1.21 mm/s at oxidizer mass flux range of 96–129 kg/m2s for square grain and 1.12–1.35 mm/s at oxidizer mass flux range of 91–126 kg/m2 s for a star grain. A complete set of the regression rate formulas is obtained for all three-grain designs as a function of oxidizer flux rate. Research limitations/implications The experiment has been performed for a lower chamber pressure up to 10 bar. Originality/value Different grain configurations were designed according to the required dimension of the combustion chamber, injector and exhaust nozzle of the design of a lab-scale hybrid rocket, and input parameters were selected and analyzed.

Computational algorithm for hybrid rocket motor preliminary design

Hybrid rockets present a compromise between solid and liquid rockets. It has a simpler design than liquid rockets and higher performance than solid rockets. It is constructed from an oxidizer tank followed by a solid fuel grain. Therefore, the combustion is controlled. This paper presents an algorithm to construct the design space of the hybrid rocket motor for a single cylindrical port. A simple regression rate formula is used to treat the combustion. The inner diameter of the grain, outer diameter of the grain, and length of the grain are used as parameters to construct the design space and a single design point is chosen. The performance of this design point is presented with the combustion time.

Influence of [formula omitted] oxidizer blends on soot formation and radiative heat flux in PMMA-air 2D slab burner for understanding hybrid rocket combustion

The focus of this study is to establish a correlation between the flame to fuel incident radiative heat flux (RHF) in 1D steady-state diffusion flames to non-intrusively measured RHF in 2D slab burner experiments for a range of O2/N2 oxidizer blends. Slab burner experiments are conducted for a range of increasing O2 concentrations from 17% to 100% in the oxidizer stream. Non-intrusive two-color pyrometry is used to determine flame temperature, soot volume fraction, and flame to fuel surface RHF. Measurements and modeling confirm soot volume fraction attains a local maximum for an O2 concentration of 45%. Low-strain diffusion flame modeling shows the reason for the local maximum is the relative location of the soot formation/nucleation fronts relative to oxidation which decreases with increasing O2 concentration. Overall good agreement between modeling and measurements of radiative heat flux to the fuel is observed for the fuel and range of O2 concentrations considered. These results demonstrate the viability of using flamelet modeling approaches for describing hybrid rocket motors.

Injector Effects on Hybrid Polymethylmethacrylate Combustion Assessed by Thermochemical Tomography

In this work, two representative hybrid rocket injector designs, a single-port and an axial showerhead design, were examined experimentally to determine the comparative influence on the development of the downstream combustion zone with polymethylmethacrylate (PMMA) fuel. One-dimensional laser absorption tomography was employed to obtain the thermochemical structure of the PMMA–gaseous oxygen reaction layer via quantitative in situ species and temperature measurements. Optical measurements were performed with tunable semiconductor lasers in the midwave infrared to target absorption features in the fundamental vibrational bands of CO, , and . Planar measurements were conducted at the exit plane of cylindrical fuel grains () at axial distances () ranging from 2 to 11 with a constant gaseous oxygen flow for each of the injectors. Assuming azimuthal symmetry, a Tikhonov-regularized Abel inversion technique was used to obtain the radially resolved absorption coefficient, from which temperature and mole fractions were extracted. One-dimensional measurements over all fuel-grain lengths were consolidated to obtain spatially resolved two-dimensional images of the reaction layer thermochemical structure in a hybrid rocket motor geometry. Significant variations in experimental results were observed between the two injectors, suggesting the importance of injection fluid dynamics in hybrid rocket combustion performance. Explanations of the variations are offered from both a theoretical and a computational analysis.

Transient investigation of nozzle erosion in a long-time working hybrid rocket motor

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.

Variability analysis of ABS solid fuel manufactured by fused deposition modeling for hybrid rocket motors

The process of fused deposition material (FDM) was used to manufacture propellant grains of Acrylonitrile Butadiene Styrene (ABS) as novel rocket fuel grain, with three types of geometry in the burning port. These solid fuel grains were used to measure the typical characteristics of combustion in rocket motors such as thrust and pressure inside the combustion chamber, seeking to obtain preliminary characteristics of operation and analyze the effect of combustion port geometry on pressure and thrust, using Multivariate Analysis of Variance (MANOVA) as statistical method. Two of the three geometries were manufactured with a helical-finocyl configuration, specially designed to be fabricated by Direct Digital Manufacturing (DDM), the other one was a straight-bore geometry also by DDM. This characterization experiment was performed on a static hybrid rocket engine, designed to inject 99.98% pure nitrous oxide into a combustion chamber with capacity to withstand 6.9 MPa of pressure, with an easy-to-exchange nozzle, avoiding erosive behavior in the throat. Statistical analyses made with the ABS fuel grains, suggest a significant effect on rocket motor pressure and thrust, due to helical geometric changes made to the combustion port of solid fuel grains made by FDM manufacture process.

Design optimization of velocity-controlled cruise vehicle propelled by throttleable hybrid rocket motor

Throttleable engines have greatly raised the technical performances of aircrafts and missiles, particularly for cruise vehicles. The hybrid rocket motor (HRM) combines the advantages of solid and liquid rocket engines, especially the ability of easily throttling and restart. This paper presents a design optimization of an HRM propelled cruise vehicle to explore the throttling ability of HRM on velocity control. First, a theoretical thrust model is presented and revised through several experimental tests. Based on it, a detailed design process of HRM is proposed. Second, the layout of the cruise vehicle, flight trajectory model and a closed-loop PID velocity control method are developed. The differential evolution algorithm is adopted to get the optimal design result, which is used to analyze velocity characteristics by real-time control of the oxidizer flow rate. The results show that the PID-control has good effect on velocity control performance and maintaining small deviation from the target. The influences of aerodynamic and thrust errors are also considered and analyzed. For different types of deviations, including the proportional and periodic perturbations, the PID-controlled scheme shows a good stability under the parameters' deviations, which indicates that the HRM with the real-time throttling thrust has a good capacity of high-precision velocity control when used in rocket propelled vehicles.

Design Optimization and Parameter Analysis of a Hybrid Rocket Motor-Powered Small LEO Launch Vehicle

In this paper, the effects of different grain shapes of a hybrid rocket motor (HRM) and different payload mass/orbit heights on the design of small launch vehicles (SLVs) are systematically discussed. An integrated overall design model for the hybrid rocket motor-powered small launch vehicle (HPSLV) is established, and two groups of three-stage SLVs capable of sending small payloads to the low earth orbit (LEO) are designed and optimized. In the first group, the SLVs with different grain shapes and different numbers of chambers in HRMs at the 1st and the 2nd stages are optimized and analyzed. In the second group, the SLVs capable of sending different payload mass to different orbit heights are optimized and analyzed. Pareto graphs of the design results show that the design of HRM at the 1st stage has the greatest impact on the take-off mass, total velocity increment, and maximum axial overload of the SLV. Self-organizing maps show that the take-off mass, maximum diameter, overall length, and velocity increment of the SLVs have the same variation tendency. For the 1-chamber HRM at the 1st stage, the wheel-shaped grain is better than circle-shaped and star-shaped grains in terms of reducing the total mass and length of the SLV, and the 4-chamber parallel HRM has more advantages over all 1-chamber designs for the same reason. The theoretical velocity increments are calculated by the Tsiolkovsky formula, and the actual velocity increments are obtained based on the trajectory simulation data. The results indicate that the HPSLV has a regular distribution in terms of the ratio of theoretical (actual) velocity increments at three different stages, and the estimated distribution ratio is around 1 : 1.55 : 1.69 (1 : 1.9 : 2.39), which can provide some reference for future development of HPSLV.

Variations and Control of Thrust and Mixture Ratio in Hybrid Rocket Motors

Hybrid rocket motors have several attracting characteristics such as simplicity, low cost, safety, reliability, environmental friendliness. In particular, hybrid rockets can provide complex and flexible thrust profiles not possible with solid rockets in a simpler way than liquid rockets, controlling only a single fluid. Unfortunately, the drawback of this feature is that the mixture ratio cannot be directly controlled but depends on the specific regression rate law. Therefore, in the general case the mixture ratio changes with time and with throttling. Thrust could also change with time for a fixed oxidizer flow. Moreover, propellant residuals are generated by the mixture ratio shift if the throttling profile is not known in advance. The penalties incurred could be more or less significant depending on the mission profile and requirements. In this paper, some proposed ways to mitigate or eliminate these issues are recalled, quantitatively analysed and compared with the standard case. In particular, the addition of energetic additives to influence the regression rate law, the injection of oxidizer in the post-chamber and the altering-intensity swirling-oxidizer-flow injection are discussed. The first option exploits the pressure dependency of the fuel regression to mitigate the shift during throttling. The other two techniques can control both the mixture ratio and thrust, at least in a certain range, at the expense of an increase of the architecture complexity. Moreover, some other options like pulse width modulation or multi-chamber configuration are also presented. Finally, a review of the techniques to achieve high throttling ratios keeping motor stability and efficiency is also discussed.

Characterizing combustion of a hybrid rocket using laser absorption spectroscopy

The combustion of an oxygen/paraffin hybrid rocket motor was experimentally characterized. Firing tests were conducted for different oxidizer mass fluxes ranging from 2.47 to 3.40 g/ (cm2·s). Variations in temperature and H2O partial pressure at the nozzle exit were diagnosed using mid-infrared tunable diode laser absorption spectroscopy (TDLAS) based on H2O absorption near 2.5 μm. Three H2O absorption lines were simultaneously covered by only one distributed feedback (DFB) laser using scanned-wavelength direct absorption (DA) mode with 2.0 kHz repetition rate. Measurement uncertainty was analyzed in detail considering line-strength uncertainty and Voigt fitting residuals. A two-dimensional (2D) model of the nozzle was constructed using the ANSYS FLUENT CFD software package. The combustion efficiency of the hybrid rocket motor was evaluated from the perspectives of chemical reaction and heat release, respectively, based on TDLAS results and CFD simulations. The effectiveness of the evaluation was validated by comparing its results with characteristic velocity (C*)-based combustion efficiency. Finally, comparisons of combustion efficiencies among different cases show that increasing the oxidizer mass flux or oxidizer-to-fuel ratio improves the combustion efficiency of the hybrid rocket motor under our experimental conditions.

Analytic target cascading with fuzzy uncertainties based on global sensitivity analysis for overall design of launch vehicle powered by hybrid rocket motor

An analytic target cascading (ATC) method with fuzzy uncertainties based on global sensitivity analysis is proposed to solve the optimization problem of typical multilevel multidisciplinary nonlinear system. The overall design of launch vehicle (LV) powered by hybrid rocket motors (HRMs) involves multiply disciplines, and is decomposed into two levels using an ATC framework. The fuzzy theory is applied to describe the uncertain design factors caused by decisions and cognition insufficiency. The rank correlation coefficient method (RCCM) based on feasible optimization solutions and the quadratic response surface method (QRSM) based on Latin hypercube sampling (LHS) are used for global sensitivity analyses of input uncertainty and model uncertainty, respectively. The multi-island genetic algorithm (MIGA) is adopted in all examples, and the known two-phase optimization method is used to verify the LV design results. The results show that global sensitivity analysis can significantly filter the fuzzy uncertain factors which have little influence on the responses. The ATC decomposition is applicable in solving the calculation burden caused by uncertainties. The fuzzy-based design optimization with ATC is more efficient than that with MDF, and gives more reliable and robust results than the deterministic ATC method.

Closed Loop Throttle Control of a Liquefying Fuel Hybrid Rocket Motor

The thrust produced by a hybrid rocket motor (HRM) can be controlled by varying the oxidizer flow rate to the combustion chamber. This feature is useful in shaping motor thrust profiles and optimizing a vehicle flight trajectory, but propellant throttling in solid-fuel hybrids is limited to the liquid component only, complicating the control scheme and potentially destabilizing combustion in the motor. While hybrid motor throttling ability remains a subject of considerable interest, there has been little investigation of throttling in motors that use high regression rate, liquefying fuels such as paraffin wax. This article describes the development and implementation of a closed loop thrust control scheme for a laboratory-scale paraffin wax/nitrous oxide HRM using a low-cost ball valve as the controlling hardware element. A model of motor performance is first developed from which proportional-integral-derivative (PID) controller constants are obtained through experimental testing. The control scheme is demonstrated through closed loop hot fire tests of a laboratory-scale motor in which thrust tracks a set-point value with feedback provided through a load cell. Upon reaching the setpoint, the motor remains throttled within ± 2.4% of the maximum thrust of the motor. Constant and ramping thrust profiles are demonstrated.

Explicit Modelling of the Ignition Transient Structural Response of a Paraffin Wax Hybrid Rocket Motor Fuel Grain

Paraffin wax has been identified as a hybrid rocket motor fuel, which offers enhanced regression rates and improved combustion performance. While various investigations into the performance of this class of fuels are being conducted around the world, the consideration of its structural performance is often overlooked. The research presented here establishes a simplified, yet accurate method of defining the structural performance of a paraffin wax hybrid fuel grain to be introduced early in the design phase of a motor. The use of the Johnson–Cook (J–C) material model has been verified to work within the “low speed” ignition range experienced in paraffin wax/N2O hybrid motors, and therefore is used to predict failure in a variety of motors. The resultant stress profiles within the grains indicate that the grain outer to inner diameter (OD/ID) ratio, as well as the outer diameter (OD) itself, play an important role in the grain ability to withstand the loading conditions applied. Additionally, the grain structural properties, and the stiffness of the combustion chamber affect the severity of the internal stresses in the grain. The feasibility of large-scale pure paraffin wax grains without structural enhancement additives is thus found to be poor. Fuel additives should be considered for structural enhancement.

Development of pre-molded internal thread on composite tubes

Premolded internal threads on composite tubes were developed. The composite tubes with threads on both ends were made using a mandrel with a male thread. The threads can be applied to struts with adjustable end fits and composite pressure vessels with threaded caps that enable disassembly for inspection and repair. Carbon fiber-reinforced plastic (CFRP) prepregs were laid up on a mandrel, wrapped with shrink tape, and cured in an oven. The threads were built-in, without using machine cutting, and the fibers on the thread were continuous through the thread and tubes for high strength. The thread was alternately rounded, convex, and concave in shape to enable CFRP prepregs to be laid up. Two types of specimen were made and tested. The layup sequence of specimen A was [0/h/90/h/0/h(1/2)]s, and that of specimen B was [0/h/90/h/90/h/0/h/90/h/90/h/90/h/90/h/0/h/90/90/0], where “h” denotes a helical layer along the concave part of the threads. The relation between load and strain is nonlinear because of the rounded shape of the threads; however, a simple and closed form analytical model was able to predict the strength of the threads and design of the threads. The model was compared with the experimental results. In addition, an application of threads for the pressure vessel of the hybrid rocket motor is also reported. The combustion test proceeded without failure. Visual inspection after the test indicated that the threads and tubes were not damaged, and thus, they can be applied to high-pressure and high-temperature rocket motors.

Catalytic Decomposition of Nitrous Oxide for Use in Hybrid Rocket Motors

In recent years, there has been significantly increased interest in using nitrous oxide as an oxidizer in hybrid rocket motor systems. The attraction of nitrous oxide is its low relative toxicity and high vapor pressure, potentially enabling self-pressurizing systems. Its decomposition is highly exothermic with a large energy barrier, typically requiring high temperature (700°C) to achieve appreciable decomposition rates. One solution is to use a catalyst to enable decomposition at significantly lower temperatures. This Paper uses a commercially available rhodium-based catalyst in a custom test cell to decompose nitrous oxide at temperatures well below the uncatalyzed self-sustaining decomposition temperature. This Paper also demonstrates that decomposition using the tested catalyst is direct to and at all temperature ranges observed and is incomplete with residual present at lower temperatures. Further, this Paper couples our test cell with a counterflow burner to ignite and burn hydroxyl-terminated polybutadiene fuel with the decomposed nitrous oxide and evaluate performance via fuel regression rate at a variety of conditions. The fuel autoignites when the gas exit temperature reaches 750°C, and regression rate is shown to increase with both mass flux and temperature.

Thermal reaction characteristics of paraffin in the presence of combustion catalysts

Paraffin fuels are regarded to be the best solid fuel of hybrid rocket motor due to the high regression rate, low cost, and environment friendly products. However, the regression rate of the fuels still cannot completely meet the requirement of the motor. Catocene, copper chromite, and cobalt stearate have been used as the combustion catalysts in this study, where the simultaneous TG-FTIR-MS and TG experiments at different heating rates were carried to investigate the thermal behavior of paraffin and paraffin/combustion catalysts mixtures. The results showed that both paraffin and the mixtures could decompose or be oxidized into gases products almost completely, with the main decomposition product of C2H6 and oxidation products of C2H4O, CO, CO2, and H2O. There was only one stage for both decomposition and oxidation of paraffin, and the kinetic model of decomposition process are close to two-dimensional diffusion (D2 model); cobalt stearate favored the decomposition of paraffin owing to the higher pre-exponential factor and the lower activation energy, whereas only copper chromite played an obvious positive role in both oxidation and regression rate of paraffin.

Morphology and anti-ablation properties of composites nozzles under the Ф100mm H2O2- polyethylene hybrid rocket motor test

Environmentally friendly commercial applications spurred us to screen a suitable ablative throat material for the hybrid rocket, while preserving its cost-effective advantages as a special chemical motor. In this research, two types of carbon fiber reinforced composites, i.e. carbon/carbon (C/C) composite and Cf/C–SiC–ZrC composite utilized in high temperature environment, were employed to make the hybrid rocket nozzle. By comparison with the high-density graphite, the anti-ablation properties under the firing environment of Ф100mm H2O2-polyethylene hybrid rocket motor were characterized. We used whole felt preform to make C/C composite, whose matrix carbon was coming from chemical vapor infiltration of propylene; and the Cf/C–SiC–ZrC composite, which employs the same whole felt preform to make the low-density C/C billet, by infiltrated with Si and Zr organic precursors and pyrolysis at elevated temperatures repeatedly to make the advanced ceramic matrix composite. The firing test lasted 40s for all the candidate materials and the result indicated that the Cf/C–SiC–ZrC composite, whose average linear ablation rate was only 0.003 mm/s, was the most stable one in the firing environment. The SEM images gave detailed morphologies of those nozzle throat materials and proved that the fiber architecture, together with the glassy ceramic oxide, helped the nozzle to withstand the hybrid motor firing environment.

Regression rate of paraffin-based fuels in hybrid rocket motor

Paraffin-based fuels are regarded as a promising solid fuel for hybrid rocket motor (HRM) owing to high regression rate, excellent mechanical performance, and high dropping point. The regression rate of paraffin-based fuels was characterized by static fire test with nitrous oxide (N2O) and gaseous oxygen (GOX) as oxidizer, and the effect of combustion catalyst, injector and chamber pressure on the regression rate was evaluated in this study. The results show that the average regression rates of paraffin-based fuel with GOX and N2O are similar under the optimal oxidizer/fuel (O/F) ratio, and the exponent n is larger with GOX as the oxidizer, indicating high sensibility to oxidizer mass flux. The average regression rate increased with the increase in swirling number Sg, and the tangential injector with Sg of 6.83 improved the regression rate significantly. The local regression rate unchanged with axial position when a jet injector was used, but a complex local regression rate appeared at the first half of fuel grain for tangential injector. Both copper chromite and cobalt stearate obviously increased the regression rate; however, only copper chromite could be used as a preferable regression rate improver owing to significant curing acceleration of fuels by cobalt stearate. The average regression rate slightly increased with the increase in chamber pressure with GOX as the oxidizer, but a noticeable positive relationship could be found for the cases of N2O.

Experimental Studies of Protrusion-Inserted Hybrid Rocket Motor with Varying Ratio

This study investigates the effects of single, double, and multi-protrusion on the performance of a hybrid rocket with two different ratio motors, mainly 12.5 and 18. Wax and gaseous oxygen were used as fuel and oxidizer, respectively. It has been observed from studies that the performance of a hybrid rocket is sensitive to the number of protrusions and their locations within the motor. Between the two motors, single protrusion was outperformed in the motor, whereas double and multi-protrusion were more effective in the motor. The highest improvement in regression rate was observed from the use of three protrusions (0.2, 0.5, and ) in both motors. The use of four protrusions (0.2, 0.4, 0.6, and ) was more significant in improving the regression rate with the motor than with the motor. Upon insertion of protrusion, combustion efficiency was enhanced in the motor and decreased in the motor. The motor was more stable in terms of the shift in oxidizer-to-fuel ratio than the motor. Studies on exhaust plume reveal that a large amount of unburned fuel comes out of the motor and remains as white smoke for all the protrusion cases tested with the motor.