Hybrid Rocket Engine Research Articles

A Review of Recent Developments in Hybrid Rocket Propulsion and Its Applications

This paper extensively reviews hybrid rocket propulsion-related activities from combustion engine designs to launch tests. Starting with a brief review of rocket propulsion development history, a comparison among the three bi-propellant rocket propulsion approaches, and hybrid rocket engine design guidelines, a very thorough review related to hybrid rocket propulsion and its applications is presented in this paper. In addition to propellant choice, engine design also affects the hybrid rocket performance and, therefore, a variety of engine designs, considering, e.g., fuel geometry, swirl injection, ignition designs, and some innovative flow-channel designs are also explored. Furthermore, many fundamental studies on increasing hybrid rocket engine performances, such as regression rate enhancement, mixing enhancement, and combustion optimization, are also reviewed. Many problems that will be encountered for practical applications are also reviewed and discussed, including the O/F ratio shift, low-frequency instability, and scale-up methods. For hybrid rocket engine applications in the future, advanced capabilities and lightweight design of the hybrid rocket engine, such as throttling capability, thrust vectoring control concept, insulation materials, 3D-printing manufacturing technologies, and flight demonstrations, are also included. Finally, some active hybrid rocket research teams and their plans for flight activities are briefly introduced.

Investigation of the existing methods for designing pre-combustion chambers in hybrid rocket engines

Investigation of the existing methods for designing pre-combustion chambers in hybrid rocket engines

Development of a Lunar Vehicle with a Hybrid Rocket Engine Produced Using Lunar Resources

Development of a Lunar Vehicle with a Hybrid Rocket Engine Produced Using Lunar Resources

Hybrid rocket engine optimization by controlled oxidizer flow dynamics

Rocket engines are an essential component in delivering any payload such as satellites, rovers, and probes into space. One such kind of engine is the hybrid rocket engine. Classical hybrid rocket engines have first become popular in the early 1970s. These engines, while providing combined advantages of both solid rocket motors and liquid rocket engines, as well as being relatively cheap to manufacture when compared to liquid rocket engines, exhibit drawbacks in thrust throughout the burn of the engine caused by the oxidizer to fuel ratio shifts induced by fuel regression. Their loss in efficiency is what makes them undesirable in the construction of rockets and causes companies to use the more expensive liquid rocket engines which are higher in efficiency and overall thrust output. The negative side effects of the fuel regression can be reduced by eliminating the oxidizer to fuel ratio shifts during the operation of a hybrid rocket engine. In this paper, a new approach of controlled oxidizer flow is proposed to achieve a more favourable combustion. The differential equations of the fuel mass flow rate function are numerically solved in terms of oxidizer mass flow rate, the results of which are taken as a basis to define a control function. It was demonstrated that it is possible to eliminate the oxidizer-fuel ratio shifts for all fuel types.

Regenerative cooling in hybrid rocket engines based on Self-Pressurized liquid nitrous oxide

An experimental investigation of the reliability and feasibility of a regenerative cooling system in hybrid rocket engines based on saturated nitrous oxide is presented. The novelty of the work consists in the use of a cooling system based on saturated liquid nitrous oxide, wrapping the external surface of a COTS graphite nozzle. The detrimental heat fluxes developed at the throat are handled with a slight increase in assembly complexity. Ten firing tests are performed with an incremental logic of the mass flow rate from 9 to 45 g/s, which is controlled by cavitating the fluid through orifices with different diameters upstream of the cooling system. The tests have been performed at different ambient temperatures in different seasons. The multiphase state of the coolant is evaluated by measuring pressure and temperature upstream and downstream of the cavitation orifice and cooling channels. The test with the smallest mass flow rate developed the highest vapour fraction downstream of the orifice and cooling channels at 0.458 and 0.481, respectively. The cooling performance is indirectly evaluated by measuring the nozzle temperature at 3, 5, and 8.5 mm from the throat internal surface. The results show that steady temperatures are achieved inside the nozzle, with throat temperatures included between 700 and 1200 K in a chamber pressure range between 5 and 30 bar. Nozzle erosion never occurs in the entire experimental campaign, and the nozzles are totally reusable for more ignitions. The coolant heat transfer coefficient increased from 3912 to 21181 W/(m2∙K) by increasing the flow rate per channel from around 3 to 15 g/s. Compared to cryogenic oxygen, liquid nitrous oxide displays higher cooling performance. Finally, the cavitation orifice is moved downstream of the cooling system, showing worse performance of the overall propulsion system.

Regression Rate and Combustion Efficiency of Composite Hybrid Rocket Grains Based on Modular Fuel Units

This study investigated combustion characteristics of composite fuel grains designed based on a modular fuel unit strategy. The modular fuel unit comprised a periodical helical structure with nine acrylonitrile–butadiene–styrene helical blades. A paraffin-based fuel was embedded between adjacent blades. Two modifications of the helical structure framework were researched. One mirrored the helical blades, and the other periodically extended the helical blades by perforation. A laboratory-scale hybrid rocket engine was used to investigate combustion characteristics of the fuel grains at an oxygen mass flux of 2.1–6.0 g/(s·cm2). Compared with the composite fuel grain with periodically extended helical blades, the modified composite fuel grains exhibited higher regression rates and a faster rise of regression rates as the oxygen mass flux increased. At an oxygen mass flux of 6.0 g/(s·cm2), the regression rate of the composite fuel grains with perforation and mirrored helical blades increased by 8.0% and 14.1%, respectively. The oxygen-to-fuel distribution of the composite fuel grain with mirrored helical blades was more concentrated, and its combustion efficiency was stable. Flame structure characteristics in the combustion chamber were visualized using a radiation imaging technique. A rapid increase in flame thickness of the composite fuel grains based on the modular unit was observed, which was consistent with their high regression rates. A simplified numerical simulation was carried out to elucidate the mechanism of the modified modular units on performance enhancement of the composite hybrid rocket grains.

Fast Reconfiguration Maneuvers of a Micro-satellite Constellation Based on a Hybrid Rocket Engine

AbstractIn this work, the formation flight of the CubeSat cluster RODiO (Radar for Earth Observation by synthetic aperture DIstributed on a cluster of CubeSats equipped with high-technology micro-propellers for new Operative services) with respect to a small satellite in LEO (Low Earth Orbit) has been analyzed. RODiO is an innovative mission concept funded by the Italian Space Agency (ASI) in the context of the Alcor program. The small satellite is equipped with an antenna that allows it to function as a transmitter, whereas RODiO functions as a receiver. The extension of the virtual SAR (Synthetic Aperture Radar) antenna can be achieved by establishing an along-track baseline performing an orbital coplanar maneuver. Another interesting scenario is the possibility to create a cross-track baseline performing an inclination change maneuver. Such formation reconfiguration maneuvers can be achieved in relatively short times only by use of a high-thrust propulsion system, i.e., based on conventional chemical technologies. From the study of maneuvers, it is possible to identify the required ∆V (order of magnitude of 10 m/s), which represents an input parameter for the design of propulsion system. Among the different kinds of propulsion systems, a Hybrid Rocket Engine was chosen. Based on the previous experience acquired by Department of Industrial Engineering (University of Naples Federico II), the preliminary design of the thrust chamber for a Hybrid Rocket Engine based on Hydrogen Peroxide (91 wt%) of the 10 N-class could be carried out, whose dimensions meet the compactness requirements of the CubeSat (1.5 U, 2 kg).

Numerical study of the hybrid rocket engine

Numerical study of the hybrid rocket engine

Superior Al90Mg10-Based Composite Fuel Grain for a Hybrid Rocket Engine

A paraffin-based fuel coupled with a nested helical matrix structure is a low-cost, high-performing alternative for hybrid rocket engine applications. The mechanical and combustion properties of the composite fuel grain can be enhanced by replacing the conventional polymer matrix with a metal skeleton. Three-dimensional printing was used to design an Al90Mg10 skeleton embedded with the paraffin-based fuel. The combustion characteristics of the composite fuel grain, including the ignition behavior, pressure oscillations, regression rate, and combustion efficiency, were comprehensively investigated. The properties of grains with and without a secondary perforated structure were compared. The flame structure and metal burning behavior of the hybrid rocket engine were monitored by endoscopic radioluminescence imaging, and the emission spectral characteristics of the plume were analyzed simultaneously. Good flammability makes the Al90Mg10 helical skeleton a promising candidate for enhancing the combustion performance of a paraffin-based fuel grain. The combustion process with rapid ignition was relatively stable, and no additional pressure oscillation frequency was observed. The metal-based composite fuel grains had a superior regression rate to that of a paraffin-based fuel grain (up to 100% higher using the perforated skeleton). Introducing a secondary structure into the fuel grain promoted the reaction and thereby enhanced the combustion performance.

Experimental Research into an Innovative Green Propellant Based on Paraffin–Stearic Acid and Coal for Hybrid Rocket Engines

This study focuses on an innovative green propellant based on paraffin, stearic acid, and coal, used in hybrid rocket engines. Additionally, lab-scale firing tests were conducted using a hybrid rocket motor with gaseous oxygen as the oxidizer, utilizing paraffin-based fuels containing stearic acid and coal. The mechanical performance results revealed that the addition of stearic acid and coal improved the mechanical properties of paraffin-based fuel, including tensile, compression, and flexural strength, under both ambient and sub-zero temperatures (−21 °C). Macrostructural and microstructural examinations, conducted through optical and scanning electron microscopy (SEM), highlighted its resilience, despite minimal imperfections such as impurities and micro-voids. These characteristics could be attributed to factors such as raw material composition and the manufacturing process. Following the mechanical tests, the second stage involved conducting a firing test on a hybrid rocket motor using the new propellant and gaseous oxygen. A numerical simulation was carried out using ProPEP software to identify the optimal oxidant-to-fuel ratio for the maximum specific impulse. Following simulations, it was observed that the specific impulse for the paraffin and for the new propellant differs very little at each oxidant-to-fuel (O/F) ratio. It is noticeable that the maximum specific impulse is achieved for both propellants around the O/F value of 2.2. It was observed that no hazardous substances were present, unlike in traditional solid propellants based on ammonium perchlorate or aluminum. Consequently, there are no traces of chlorine, ammonia, or aluminum-based compounds after combustion. The resulting components for the simulated motor include H2, H2O, O2, CO2, CO, and other combinations in insignificant percentages. It is worth noting that the CO concentration decreases with an increase in the O/F ratio for both propellants, and the differences between concentrations are negligible. Additionally, the CO2 concentration peaks at an O/F ratio of around 4.7. The test proceeded under normal conditions, without compromising the integrity of the test stand and the motor. These findings position the developed propellant as a promising candidate for applications in low-temperature hybrid rocket technology and pave the way for future advancements.

Effect of characteristic structure of nested composite hybrid rocket fuel grain on combustion properties

This work numerically and experimentally investigated the effect of the characteristic structure of nested composite fuel grains on their combustion properties. The characteristic structure comprised multiple helical grooves inside the fuel grain port, which formed during the combustion process. Three-dimensional steady-state simulations were employed to analyze the burning behavior of composite fuel grains with different scales of the characteristic structure (i.e., groove depth) of 0, 1, and 2 mm. An obvious swirl-gas flow field and large tangential velocity appeared for fuel grains with the characteristic structure, which was conductive to enhancing the propellant mixing. A larger flame region and more intense burning were correspondingly achieved, during which greater amounts of H2O and CO2 were generated than by fuel grain without the characteristic structure. Numerical results verified that the characteristic structure was theoretically beneficial to improving the combustion properties of the grains. Firing experiments were conducted using a laboratory-scale hybrid rocket engine with oxygen as the oxidizer. The flame structure and pulsation features were synchronously analyzed using a radiation imaging technique. Pure paraffin-based fuel grains with a circular port were tested for comparison. Both regression rates and combustion efficiencies of the composite fuel grains with characteristic structure scales of 1 mm and 2 mm showed significant improvement. The flame images exhibited an obvious helical shape and large area, which was consistent with the numerical analysis. Both the numerical and experimental results confirmed that the helical characteristic structure in the composite fuel grain played a key role in enhancing the combustion properties and demonstrate potential for further development.

Experimental investigation of regenerative cooling performance in hybrid rocket engines

An experimental investigation regarding the reliability and feasibility of a regenerative cooling system in hybrid rocket engines is presented. The novelty of the work is the introduction of a regeneratively cooled carbon-based nozzle throat using liquid oxidizer, for thermal management of the detrimental heat fluxes developed in the nozzle. The benefits and drawbacks of using regenerative cooling based on the observations of the current experimental campaign are critically discussed. Cryogenic oxygen in subcritical conditions is used as a coolant and oxidizer, for combustion with a fuel grain of high-density polyethylene stored within the combustion chamber. Pressure and temperature measurements are taken from multiple locations in the oxygen feed line and within the graphite nozzle throat. Eight firing tests were performed under various oxidizer mass flow rates and thus various thrusts. It was confirmed that cooling the graphite throat prevented the occurrence of nozzle erosion in all eight firing tests. The heat transfer rate and the physical state of the coolant are evaluated indirectly. Results show that increasing coolant gasification correlates with decreasing mass flow rate. The cooling performance improved more with higher vapour fractions than larger mass flow rates, however, dry-out occurred in tests with the highest gasification level. In summary, is it shown that while the current regenerative cooling system will supress nozzle erosion and limit nozzle temperatures in long duration firing, it reduces the predictability of flow rate and thus thrust during operations.

Multiphysics Modeling for Combustion Instability in Paraffin-Fueled Hybrid Rocket Engines

The use of paraffin-based fuels is a promising approach to a low regression rate in hybrid rocket engines, and the capability to describe and predict combustion instability in the presence of liquefying fuels becomes an enabling step towards the application of hybrid rockets in a wide range of space transportation systems. In this work, a multiphysics model having this purpose is presented and discussed. The model is based on a network of submodels in which chamber gas dynamics is described by a quasi-1D Euler model for reacting flows while thermal diffusion in the grain is described by the 1D heat equation in the radial direction. An artificial neural network is introduced to reduce the computational cost required by the chemical submodel. A sensitivity analysis is performed to identify the key parameters, which have the largest influence on combustion instability and to evaluate the predictive capability of the model despite the uncertainty introduced with the necessary modeling simplifications. Results are presented considering two test cases with different oxidizers: hydrogen peroxide and gaseous oxygen. The procedure shows good agreement with the experimental results available in the literature.

Computational Analysis and Regression Laws for Nozzle Erosion Prediction in Hybrid Rockets

The erosion of the nozzle throat can represent one of the major limitations against the future widespread use of hybrid rocket engines (HREs) in the space industry. In fact, nozzle erosion in HREs can be more severe and harder to predict than in solid rockets due to the higher concentration of oxidizing species in the combustion products and to mixture ratio shifts and/or throttling. Therefore, an accurate understanding of the erosion phenomenon is of fundamental importance for the technological advancement of HREs. This work is focused on the investigation of graphite nozzle erosion in HREs burning high-density polyethylene with two different oxidizers, oxygen and nitrous oxide. First, the results of a computational fluid dynamics parametric analysis are used to derive closed-form regression laws for the rapid estimation of nozzle throat erosion and wall temperature depending on chamber pressure and mixture ratio. Then, a one-dimensional transient heat conduction solver is loosely coupled with the aforementioned regression laws, allowing to reconstruct the transient heating process within the solid. The obtained numerical results are validated against experimental data. Finally, the effect of gas-phase reactions on the heterogeneous reactions occurring at the nozzle surface is highlighted when moving from fuel-rich to oxidizer-rich conditions.

Interaction of multiple steps in hybrid rocket engines: Experimental investigation

In this article, we investigate the interaction of multiple steps in hybrid rocket fuel grains. Previous studies have shown the potential of single steps to increase the regression rate. In this study, we evaluate if the single step results can be translated to a multi-step approach. Of special interest are the interactions between the different step configurations. Four grain profiles are assembled by using fuel grain segments with different inner diameters, therefore forming a multi-step approximation of the profiles that can easily be manufactured and scaled up. The multi-step grains enhance the regression rate because of increased mixing and convective heat transfer induced by the recirculation zones of the steps. The experimentally obtained regression rate profiles are similar to the single step experimental data. This signifies that they are reproducible and therefore predictable. Most importantly, multiple steps do not interfere negatively with each other, which proves that steps can be used to approximate different fuel grain profiles. We show experimentally that the regression rate can be increased up to 81% by accumulating the regression rate enhancing potential of single steps. Moreover, we developed a genetic algorithm to estimate the spatially resolved Marxman parameters with only one single burn.

Hybrid Rocket Engine Burnback Simulations Using Implicit Geometry Descriptions

The performance of hybrid rocket engines is significantly influenced by the fuel geometry. Burnback simulations, to determine the fuel surface and fluid volume, are therefore an important tool for preliminary design. This work presents a method for the simulation of spatially constant burn-ups on arbitrary geometries. An implicit surface definition by means of a signed distance function is used to represent the fluid volume and the fuel block on tetrahedral meshes. Two methods each are used to determine the fluid volume and the burning surface. The first method is based on a direct integration of the signed distance function with the Heaviside function or the Dirac delta distribution, respectively. The second method linearly interpolates the position of an isosurface and thus reconstructs the fuel surface. Both methods are compared and validated with analytical results of four example geometries. Both calculations of the fluid volume and the calculation of the surface content with the interpolation method are characterized as first-order methods. With practicable mesh resolutions of one million computational cells, errors below two percent can be achieved. With the interpolation method, numerical meshes can also be exported for any time points of the burn. Finally, the application of the program to the fuel geometry of the Viserion hybrid rocket engine is demonstrated.

Experimental test analysis of a 300 N hybrid rocket engine

Experimental test analysis of a 300 N hybrid rocket engine

CFD Analysis of De Laval Nozzle of a Hybrid Rocket Engine

: The study gives a general overview of the CFD-simulated gas flow via the nozzle of a De Laval rocket engine. In CFD software, the mesh is generated using the nozzle’s geometry. In this work, CFD methods are used to compare a number of factors with those derived from classical theory, such as pressure, velocity, temperature, and arear ratios. The results of computational modelling of CFD simulation are comparable to those found from theoretical calculations.

Design, Manufacturing, and Testing Process of a Lab Scale Test Bench Hybrid Rocket Engine

The current paper presents the architecture of a test bench for small (laboratory) scale hybrid rocket motors destined for teaching purposes. The sustainability of the proposed methodology is emphasized, as it addresses the development of small-scale hybrid rocket motor test benches, to be used either as didactic means for students or for research laboratories, by using low-cost materials, while preserving efficiency. In this regard, the entire development process is approached, from designing to manufacturing (including the casting of the fuel rod) and testing of the product. The developed product uses a mixture of 88% paraffin, 10% stearic acid, and 2% coal for the solid phase and liquid oxygen for the liquid phase. The testing of the hybrid rocket motor demonstrator was performed outdoors, in controlled conditions. The results showed a good correlation between the theoretical testing parameters with the obtained ones.

Effect of grain length on GOx-paraffin hybrid rocket engines performance and regression rate

An experimental campaign was carried out on a hybrid rocket engine, using gaseous oxygen and paraffin wax-based fuel, to investigate the effect of fuel grain length on motor performance and internal ballistics. The results of the current work offer the reader a comprehensive overview of the pros and cons of reducing the fuel grain length and the difficulties in optimizing combustion efficiency. Steady-state numerical analyses were also performed to support the experimental findings. The motivation for the current work originates from a previous explorative test campaign performed with a 220 mm fuel grain length where only fuel-rich conditions were experienced. A trade-off analysis was carried out by introducing two additional configurations with grain lengths of 130 mm and 70 mm, since the grain length directly affects the mixture ratio and engine performance. Propellant mixing, combustion instability, entrainment, and regression rate are the main topics discussed in the paper. The results show that the reduction of the grain length increases the mixture ratio close to the optimum value, but mixing and combustion instability issues are promoted. No significant effects of the grain length on the regression rate have been found, leading to a discussion on the physics of the mass transfer mechanism of paraffin wax.