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A new system design tool for the simulation of a full hybrid rocket engine is developed in this article. This tool enables to simulate the behaviour of the engine at different conditions/configurations in several minutes from a desktop computer by keeping a balance between accuracy and computation time. This makes its use especially attractive for the pre-design phases of the engine. The main components of the hybrid rocket engine from our laboratory-characterized by a catalytic injection of the oxidizer-, are modelled and implemented in the tool: the feed/injection sub-system through 0-D models of a mass flow rate regulator and a catalyst; the combustion chamber, with a 1.5-D model; and the nozzle through a 1-D model. An iterative method is employed to reach the convergence of pressure in the combustion chamber between these sub-systems. The corresponding set of equations is solved by a Newton–Raphson technique. Seven experiments performed on our lab-scale engine were used to validate the system design tool. In five of the simulations, the relative differences of the main physical quantities with the experiment were below 30%, being the largest errors found in the fuel regression rate and mixture ratio. The best agreements with the experiments were retrieved for the cases with the largest oxidizer mass fluxes (above 230 kg/ m 2 /s) and mixture ratios closest to stoichiometry, defining thus, the range of applicability of the system design tool. The results presented in this article were issued of a Ph.D. thesis. • A 1.5-D combustion chamber model of a hybrid rocket engine was developed and validated. • 0-D/1-D mass flow regulator, catalyst and nozzle models were developed and validated. • A system design tool of a hybrid rocket engine implementing these models was developed. • The tool is adapted to pre-design phases, with a few minutes of computation time. • Errors between 8%–40% in fuel regression rate and below 10% in thrust and chamber pressure. [ABSTRACT FROM AUTHOR]

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. • Results from single steps translate to multiple steps. • Regression rate profiles reproducible for multi-steps. • Multi-step grains increase regression rate by up to 81%. • Spatial resolution of Marxman law possible with genetic algorithm. • Complex fuel profiles can be approximated with steps. [ABSTRACT FROM AUTHOR]

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. [ABSTRACT FROM AUTHOR]

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. [ABSTRACT FROM AUTHOR]

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. [ABSTRACT FROM AUTHOR]

Hybrid rocket engines suffer from the restricted mechanical properties and low regression rates of current polymeric fuel grains. We propose a three-dimensional printed aluminium (Al) nested composite fuel grain with millimetre-scale lattice pores (referred to as Al-L). In this study, breathable Al blades with micrometer-scale interconnected pores (Al-B) and blades combining millimetre-scale and micrometer-scale pores (Al-B&L) are designed. The formation mechanisms, characteristics, and effects of the breathable blades are analysed in simulations, micro-computed tomography, and cyclic compression tests. The mechanical properties of the composite fuel grains are investigated numerically and in compression tests. Al-B has the highest Young’s modulus at more than 15 times that of a paraffin-based fuel grain and Al-B&L has the highest yield stress at 4 times that of the paraffin-based fuel grain. Referring to combustion properties, the regression rates of the Al-B and Al-B&L grains are respectively 63.3% and 58.2% greater than the regression rate of the paraffin-based fuel grain.

An experimental study is conducted to measure the ignition delay of ammonia borane doped paraffin wax with sprays of white fuming nitric acid as oxidizer. The injection pressure and injector diameter are varied to investigate the effects of spray characteristics on the ignition dynamics. A range of Weber (W e) number for the oxidizer jet covering 100 < W e <2700 and Reynolds (R e) number 1300 < R e <8900 is investigated, whereas the W e number at the droplet scale varies from 0.8 < W e d r o p <24.1. The sprays are characterized using phase-doppler anemometry to obtain the size and speed distribution of the oxidizer droplets across multiple positions. In addition to conventional high-speed imaging, the hypergolic ignition is also observed using Schlieren imaging and a high-speed mid-infrared camera. The results show that the ignition delay of a hypergolic hybrid fuel under spray conditions is unaffected by the mean droplet size of the oxidizer. The ignition delay decreases with increasing W e number at the droplet scale and increasing droplet velocity. The variability in the ignition delay also decreases with Weber number and droplet velocity. The values obtained for the ignition delay when using a spray are generally higher than those from simple oxidizer drop experiments, highlighting the need to investigate the hypergolic ignition in configurations representative of engine conditions. [ABSTRACT FROM AUTHOR]

The combustion characteristics of a swirl-radial-injection composite fuel grain were experimentally and numerically investigated. This composite grain permits swirl-radial oxidizer injection based on three hollow helical blades, each having a constant hollow space allowing uniform oxidizer injection into the main chamber along the axial direction. The oxidizer enters from channel inlets located along a hollow outer wall. This wall, together with the three blades, is fabricated as one piece from acrylonitrile-butadiene-styrene using three-dimensional printing. Paraffin-based fuel is embedded in the spaces between adjacent blades. Firing tests were conducted with gaseous oxygen as the oxidizer, using oxidizer mass flow rates ranging from 7.45 to 30.68 g/s. Paraffin-based fuel grains using conventional fore-end injection were used for comparison. Regression rate boundaries were determined taking into account the erosion of the oxidizer channels. The data show that the regression rate was significantly increased even at the lower limit. Images of the combustion chamber flame and of the exhaust plume were also acquired. The flame was found to be concentrated in the main chamber and a smoky plume was observed, consistent with the high regression rate. A three-dimensional simulation was employed. The present design was found to improve fuel/oxidizer mixing and combustion efficiency compared with a fuel grain using fore-end injection. Both the experimental results and numerical simulations confirmed the potential of this swirl-radial-injection fuel grain. [ABSTRACT FROM AUTHOR]

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.

The paper is part of the research aimed at determining if the vortex flow pancake (VFP) hybrid rocket engine is feasible as green in-space chemical propulsion. The objective of this study is to test an N2O/HDPE VFP hybrid ignited with N2O/C3H8 torch igniter. The N2O is used in self-pressurizing mode, which results in two-phase flow and varying inlet conditions, thus better simulating real in-space behavior. The study begins with characterizing the torch igniter, followed by hot-fire ignition tests of the VFP. The results allow for the improved design of the torch igniter and VFP hybrid. The axial regression rate ballistic coefficients are reported for the N2O/HDPE propellants in the VFP configuration. [ABSTRACT FROM AUTHOR]

Hybrid rocket engines (HREs) offer a low-cost, reliable, and environmentally friendly solution for both launch and in-space applications. Hybrid propellants have been identified as green thanks to their use of non-toxic, non-carcinogenic oxidizers. Of particular relevance are storable oxidizers, namely high-concentration (≥90 wt.%) hydrogen peroxide (HP, H2O2) and nitrous oxide (N2O). This work provides a survey of experimental activities based on H2O2 and N2O for hybrid rocket propulsion applications. Open literature data are completed with original thermochemical calculations to support the discussion. [ABSTRACT FROM AUTHOR]

Firing test campaigns were carried out on a 200 N thrust-class hybrid rocket engine, using gaseous oxygen as an oxidizer and a paraffin-wax-based fuel. Different fuel grain lengths were adopted to extend the fuel characterization under different operating conditions, and to evaluate rocket performances and internal ballistics in the different configurations. In addition to data collected under a 220 mm propellant grain length, two further test campaigns were carried out considering 130 mm and 70 mm grain lengths. Two different injector types were adopted in the 130 mm configuration; in particular, a showerhead injection system was used with the aim to contain high-amplitude pressure oscillations observed during some firing tests in this engine configuration. Parameters such as the chamber pressure and temperature inside the graphite nozzle, space-averaged fuel regression rate and nozzle throat diameter were measured. The results allowed for the investigation of different issues related to hybrid rockets (e.g., fuel regression rate, engine performance, nozzle ablation under different conditions). The focus was mainly directed to the nozzle heat transfer, through the reconstruction of the convective heat transfer coefficient for different tests in the 70 mm grain length engine configuration. The reconstruction took advantage of the experimental data provided by the nozzle embedded thermocouple. Then, the experimental convective heat transfer coefficient was used to validate the results from some empirical correlations. The results showed significant differences between the experimental convective heat transfer coefficients when considering tests with different oxidizer mass flow rates. Furthermore, the predictions from the empirical correlations proved to be more reliable only in cases characterized by oxidizer-rich conditions. [ABSTRACT FROM AUTHOR]

Slow regression rate of the solid fuel is the main limitation for the use of hybrid rocket engines in high thrust applications. Paraffin-based fuels tackle this limitation thanks to the entrainment mass transfer. In this study, ballistic behaviors of conventional polymeric fuel (ABS) and paraffin-based blends are studied and compared with those of the armored grains. These latter are a new generation of fuels featuring 3D printed cellular structures embedded in the wax-based grain. The ballistic characterization focuses on the evaluation of the regression rate (r f) and its dependence on the oxidizer mass flux. Relative ballistic grading of the formulations is pursued via thickness over time methods and an optical technique for r f determination. The armored grains are reinforced by gyroid structures that are 3D printed using three different polymers (ABS, PLA, and Nylon 6) and two relative densities (10% and 15%). Despite the slow burning behavior of the printing polymers, the embedded reinforcement enhances the r f of the paraffin-based formulations, with percent increases ranging from +48% to +91%. This result could be explained by the uneven and irregular texture of the burning surface promoting turbulence (and therefore, propellant mixing) and convective heat transfer. For both the armored grains and the paraffin-based formulations, blending the pristine paraffin wax with polymeric additives results in more viscous formulations and in a r f reduction. Armored grain combustion performance makes this novel fuel an interesting candidate for high-thrust hybrid rockets. • Armored grains embed 3D-printed structures for paraffin-based fuel reinforcement. • Gyroid is taken as suitable 3D-printed open-cell reinforcing structure. • Armored grains enable fuel reinforcement with limited impact on melt layer viscosity. • Armored grains show faster regression rates (+48% to +91%) than conventional blends. • Embedded structure provides a combination of fast regression and strong mechanical properties. [ABSTRACT FROM AUTHOR]

In this work, we assess the impact of stepped hybrid rocket solid fuel grains on the performance of the engine. The influence of Forward Facing Steps (FFS) and Backward Facing Steps (BFS) on the regression rate is analysed on two different cylindrical motors and one slab burner with optical access. Accessible post-processing allows to quantify the local regression rate profiles induced by the steps for the cylindrical motors. For the slab burner, novel image processing approaches enable the authors to correlate the findings of the cylindrical motors precisely to the visual inspection of the slab burner. Using two different cylindrical motors with different fuel grain lengths stressed the importance of the total motor length on the effect of the steps. Shorter grains (118 mm) exhibit higher regression rates for FFS than for BFS (+41.3% versus +15.7%), whereas longer motors (500 mm) profit more from BFS than FFS (+26.3% versus ± 0%). Moreover, using three different oxidizers (nitrous oxide, hydrogen peroxide and gaseous oxygen) and two different fuels (high density polyethylene and paraffin) at different O/F ratios, we showed that the beneficial properties of steps are valid for numerous configurations. This would allow the employment of stepped geometries to a variety of existing motors. Finally, the authors propose the possibility to approximate fuel grain designs by a set of multiple steps. • Definition of characteristics for backward and forward facing steps in fuel grains. • Accessible analysis of axial regression rate profiles on three different motors. • Regression rate observable in intensity profile from slab burner with optical access. • Qualitative regression rate profile on cylindrical motors and slab burner is similar. • Total fuel grain length influences the performance of steps. • Multiple steps need to be distributed carefully according to their area of influence. [ABSTRACT FROM AUTHOR]

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.

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.

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.

The mechanical and combustion properties of metal-nested composite hybrid rocket fuel grains composed of spiral aluminium (Al) frameworks fabricated using three-dimensional (3D) printing with an embedded paraffin-based fuel were investigated. The mechanical properties of the resulting grains were evaluated by compression tests. In addition, the combustion characteristics of the Al composite grains were examined in a lab-scale hybrid rocket engine with gaseous oxygen as the oxidizer at an initial mass flow rate of 17.9 g/s. Pure paraffin-based (PP) and acrylonitrile–butadiene-styrene (ABS) composite grains were also tested as baseline fuels for comparison. The Al-A composite grain exhibited superior mechanical and combustion properties, with Young modulus, yield stress, and regression rate increased by 757.1%, 381.3% and 52.5% compared with the PP grain. The Young modulus and combustion efficiency were also further improved, by 51.0% and 14.9%, respectively, by including perforations in the spiral blades. These improvements are discussed in detail herein based on experimental data together with numerical simulations. Emission spectra of the engine plumes were also acquired and used to qualitatively analyze the combustion characteristics of the Al blades.

The results of numerical simulation of a gas-droplet fuel mixture combustion in an oxidizer flow under the conditions typical of hybrid rocket engines are presented. The effect of the size and rate of supply of fuel droplets on the combustion completeness, temperature, and position of the diffusion flame in the boundary layer of the oxidizer flow has been studied. It is shown that the influence due to the liquid fuel droplets manifests itself as a local decrease of gas temperature and an increase of the concentration of the gaseous fuel in the immediate vicinity of droplets. The addition of droplets to the gaseous fuel flow results in a slight decrease of the flame thickness and flame temperature, which, however, experience no large-scale perturbations as a result of droplet movement and remain almost stationary. [ABSTRACT FROM AUTHOR]

• Paper title: hybrid rocket engine optimization by controlled oxidizer flow dynamics. • Mass flowrate equation was significantly simplified. • It is possible to improve the oxidizer-to-fuel ratio by varying the mass flowrate. • The thrust output of the hybrid rocket engine is stabilized. • Improvement in combustion efficiency is achieved. 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. [ABSTRACT FROM AUTHOR]

The mechanical and combustion properties of metal-nested composite hybrid rocket fuel grains composed of spiral aluminium (Al) frameworks fabricated using three-dimensional (3D) printing with an embedded paraffin-based fuel were investigated. The mechanical properties of the resulting grains were evaluated by compression tests. In addition, the combustion characteristics of the Al composite grains were examined in a lab-scale hybrid rocket engine with gaseous oxygen as the oxidizer at an initial mass flow rate of 17.9 g/s. Pure paraffin-based (PP) and acrylonitrile–butadiene-styrene (ABS) composite grains were also tested as baseline fuels for comparison. The Al-A composite grain exhibited superior mechanical and combustion properties, with Young modulus, yield stress, and regression rate increased by 757.1%, 381.3% and 52.5% compared with the PP grain. The Young modulus and combustion efficiency were also further improved, by 51.0% and 14.9%, respectively, by including perforations in the spiral blades. These improvements are discussed in detail herein based on experimental data together with numerical simulations. Emission spectra of the engine plumes were also acquired and used to qualitatively analyze the combustion characteristics of the Al blades. [ABSTRACT FROM AUTHOR]

A Hybrid Rocket Engine (HRE) in a conventional way refers to a combination of fuel as solid and oxidizer in the liquid or gaseous phase. This combination has received significant attention in the scientific community due to its intriguing characteristic of reducing chemical explosions and enhancing operational safety. Alongside, the impact on environmental sustainability can be escalated to a reasonable scale. In this context, the current paper emphasizes the major challenges involved with combustion, design of reaction mechanism, low regression rate, choosing the right fuel and poor mechanical properties. This review will help the scientific community to establish a state-of-art technology that would resolve the current challenges to provide a sustainable result. At the same stretch, most of the works haven't realized this technology to a commercial value. Hence the current review shall focus on the works cited in the global community and lacunae for the transfer of technology-to-technology readiness level. [ABSTRACT FROM AUTHOR]

The combustion characteristics of a swirl-radial-injection composite fuel grain were experimentally and numerically investigated. This composite grain permits swirl-radial oxidizer injection based on three hollow helical blades, each having a constant hollow space allowing uniform oxidizer injection into the main chamber along the axial direction. The oxidizer enters from channel inlets located along a hollow outer wall. This wall, together with the three blades, is fabricated as one piece from acrylonitrile-butadiene-styrene using three-dimensional printing. Paraffin-based fuel is embedded in the spaces between adjacent blades. Firing tests were conducted with gaseous oxygen as the oxidizer, using oxidizer mass flow rates ranging from 7.45 to 30.68 g/s. Paraffin-based fuel grains using conventional fore-end injection were used for comparison. Regression rate boundaries were determined taking into account the erosion of the oxidizer channels. The data show that the regression rate was significantly increased even at the lower limit. Images of the combustion chamber flame and of the exhaust plume were also acquired. The flame was found to be concentrated in the main chamber and a smoky plume was observed, consistent with the high regression rate. A three-dimensional simulation was employed. The present design was found to improve fuel/oxidizer mixing and combustion efficiency compared with a fuel grain using fore-end injection. Both the experimental results and numerical simulations confirmed the potential of this swirl-radial-injection fuel grain.

The paper is part of the research aimed at determining if the vortex flow pancake (VFP) hybrid rocket engine is feasible as green in-space chemical propulsion. The objective of this study is to test an N2O/HDPE VFP hybrid ignited with N2O/C3H8 torch igniter. The N2O is used in self-pressurizing mode, which results in two-phase flow and varying inlet conditions, thus better simulating real in-space behavior. The study begins with characterizing the torch igniter, followed by hot-fire ignition tests of the VFP. The results allow for the improved design of the torch igniter and VFP hybrid. The axial regression rate ballistic coefficients are reported for the N2O/HDPE propellants in the VFP configuration.

Hybrid rocket engines (HREs) offer a low-cost, reliable, and environmentally friendly solution for both launch and in-space applications. Hybrid propellants have been identified as green thanks to their use of non-toxic, non-carcinogenic oxidizers. Of particular relevance are storable oxidizers, namely high-concentration (≥90 wt.%) hydrogen peroxide (HP, H2O2) and nitrous oxide (N2O). This work provides a survey of experimental activities based on H2O2 and N2O for hybrid rocket propulsion applications. Open literature data are completed with original thermochemical calculations to support the discussion.

Firing test campaigns were carried out on a 200 N thrust-class hybrid rocket engine, using gaseous oxygen as an oxidizer and a paraffin-wax-based fuel. Different fuel grain lengths were adopted to extend the fuel characterization under different operating conditions, and to evaluate rocket performances and internal ballistics in the different configurations. In addition to data collected under a 220 mm propellant grain length, two further test campaigns were carried out considering 130 mm and 70 mm grain lengths. Two different injector types were adopted in the 130 mm configuration; in particular, a showerhead injection system was used with the aim to contain high-amplitude pressure oscillations observed during some firing tests in this engine configuration. Parameters such as the chamber pressure and temperature inside the graphite nozzle, space-averaged fuel regression rate and nozzle throat diameter were measured. The results allowed for the investigation of different issues related to hybrid rockets (e.g., fuel regression rate, engine performance, nozzle ablation under different conditions). The focus was mainly directed to the nozzle heat transfer, through the reconstruction of the convective heat transfer coefficient for different tests in the 70 mm grain length engine configuration. The reconstruction took advantage of the experimental data provided by the nozzle embedded thermocouple. Then, the experimental convective heat transfer coefficient was used to validate the results from some empirical correlations. The results showed significant differences between the experimental convective heat transfer coefficients when considering tests with different oxidizer mass flow rates. Furthermore, the predictions from the empirical correlations proved to be more reliable only in cases characterized by oxidizer-rich conditions.

Nitrous oxide is often used in the space industry, as an oxidiser or monopropellant, mostly in self-pressurised configurations. It has potential for growth in use due to the recent rising interest in green propellants. At the same time, modelling the behaviour of a self-pressurising nitrous oxide tank is a challenging task, and few accurate numerical models are currently available. Two-phase flow, heat transfer and rapid changes of mass and temperature in the investigated system all increase the difficulty of accurately predicting this process. To get a get better understanding of the emptying of a self-pressurised nitrous oxide tank, two models were developed: a phase equilibrium model (single node equilibrium), treating the control volume as a single node in equilibrium state, and a phase interface model, featuring a moving interface between parts of the investigated medium. The single node equilibrium model is a variation of equilibrium model previously described in the literature, while the phase interface model involves a novel approach. The results show that the models are able to capture general trends in the main parameters, such as pressure or temperature. The phase interface model predicts nitrous oxide as a liquid, a two-phase mixture, and vapour in the lower part of the tank, which is reflected in the dynamics of changes in pressure and mass flow rate. The models developed for self-pressurisation, while created for predicting nitrous oxide behaviour, could be adapted for other media in conditions near vapour-liquid equilibrium by adding appropriate state equations. [ABSTRACT FROM AUTHOR]

The purpose of this paper is to present method of the numerical modeling of the hybrid rocket engine's work as a tool for designing engines of this type. The model is intended as an accurate and simple to use development tool for use in preliminary design stages of hybrid rocket engines, improving the effectiveness and quality of this process. General assumptions underlying the use of a model are presented, together with an analysis of past work in this field. Results of an extensive experimental campaign are presented and compared with the results of numerical modelling in order to calibrate the proposed model and evaluate its accuracy. Parameter variation and optimization were conducted, proving functionality of the methodology. Presented numerical calculations show that the adopted approach to reduce the analysis time and complexity was correct. This method of numerical calculation of hybrid engine working parameters combines aspects of accuracy and simplicity at the early design stage to avoid time-consuming and costly changes in subsequent detailed stages of design.

For our earlier study, we added metal powder to microcrystalline wax, a solid fuel, to improve the fuel regression rate. Results of those earlier combustion experiments revealed that aluminum (Al) powder increases the combustion flame temperature in the combustion chamber of a hybrid rocket engine, and that Al powder addition improves the solid fuel performance. Nevertheless, details of Al powder ignition in the combustion chamber remain unconfirmed. Therefore, numerical analysis of combustion in a hybrid rocket engine was conducted using software (ANSYS 2020 R2; Fluent Inc.). In general, solid fuels for hybrid rocket engines are made of polymeric thermosets (e.g., hydroxyl-terminated polybutadiene (HTPB)) or hydrocarbon compounds (e.g., paraffin). To use materials and methods similar to those used for earlier studies, we decided to use gaseous oxygen as the oxidizer and kerosene as the melted solid fuel for analyses. Through this numerical analysis, we reproduced the actual combustion experiment conducted at our laboratory. Based on the temperature distribution in the combustion chamber, the residence time, and the temperature distribution of the Al powder obtained from the analysis, we confirmed locations in the combustion chamber at which the Al powder melting and ignition might occur. Results of numerical analyses demonstrated that boundary layer combustion occurs in the combustion chamber and that the temperature rises sufficiently for Al powder melting and ignition. Additionally, results confirmed that the Al powder residence time is the period during which the Al powder can reach its melting point and confirmed that ignition and combustion can occur in the combustion chamber. [ABSTRACT FROM AUTHOR]

Hybrid rocket engines are studied internationally because they provide many benefits. However, hybrid rocket engines are rarely put into practical use because of their low regression rate. Earlier research efforts have been aimed at improving the regression rate using paraffin or microcrystalline wax (WAX) and adding metal powder to solid fuel. For hybrid rocket solid fuels, aluminum (Al) powder is used as a metallic fuel. Nevertheless, effects of adding Al powder have not been confirmed sufficiently because Al powder is reportedly difficult to ignite in the combustion chamber of a hybrid rocket engine. Magnesium (Mg) powder ignites more easily. This study specifically assessed addition of microscale magnalium (Mg-Al) powder particles because Mg-Al can overcome these and other shortcomings. Reportedly, Mg-Al powder has good combustion completeness in solid propellant. This study used combustion experiments to elucidate effects of Mg-Al powder addition to WAX. Combustion experiments using a small hybrid rocket engine demonstrated that 20 mass% Mg-Al powder in WAX produces an 18.7 % increased regression rate over that of WAX alone, which is greater than adding Mg or Al. These results suggest that high combustion heat of Mg-Al powder promotes WAX combustion. Adding Mg-Al powder raises the regression rate. Additionally, residue remaining in the aft chamber after combustion was analyzed using XRD. The results of these two analyses suggest that the Mg-Al powder was ignited and burned in the aft chamber. A small combustion chamber can be used because Mg-Al powder has high density and excellent completeness of combustion. Results demonstrate that Mg-Al powder is suitable for addition to solid fuel of small hybrid rocket engines. [ABSTRACT FROM AUTHOR]

The hybrid rocket engine (HRE) has attracted much attention as one of the promising engines for the propulsive subsystem of a VTVL. A recent study reported that the thrust response in HRE exhibits different delay curves depending on the direction of thrust modulation, which is a typical example of system nonlinearity called as plant dynamics change (PDC). A series of step response tests was conducted to investigate the PDC occurrence and its nonlinear characteristics in HRE combustion by measuring response delays of various components. Test results confirmed that thrust response delay varied depending on the thrust direction and thrust range, and thus PDC of different characteristics occurred. In addition, results showed that the hysteresis of the chamber response is the critical factor related to the development of PDC. To improve the thrust control accuracy of the VTVL system simply by reducing the PDC in HRE thrust control, advanced PID control method with gain scheduling was used. In the result, the size of the integrated absolute error (IAE) was successfully reduced. In addition, the results confirmed that PID with gain scheduling can be an effective solution to improve the thrust control accuracy of VTVL using a hybrid rocket engine. [ABSTRACT FROM AUTHOR]

Earlier studies revealed that addition of metal powder to microcrystalline􀀎wax improves the regression rate of hybrid rocket solid fuel, but ignition of the metal powder was not confirmed to have occurred in the hybrid rocket engine chamber. To confirm the ignition of metal powder in the hybrid rocket engine chamber, theoretical calculations and combustion experiments were conducted. Theoretical calculations were heat transfer calculations from the combustion flame of microcrystalline􀀎wax to metal powder. Combustion experiments were conducted in our laboratory using a hybrid rocket engine and 2D visualization engine. Theoretical calculations suggest that the metal powder ignited in the chamber. However, adding the aluminum powder did not improve the regression rate. Adding magnesium powder partially improved the regression rate. Both metal powders improved the characteristic velocity. According to the 2D visualization experiments, both metal powders ignited in the chamber. Results of these experiments suggest that metal powder raises the combustion flame temperature in the hybrid rocket engine chamber and that the addition of metal powder improves the solid fuel performance. [ABSTRACT FROM AUTHOR]

A non-intrusive method for determining the time-resolved fuel mass flow rate in a hybrid rocket engine is proposed based on tunable diode laser absorption spectroscopy (TDLAS). The present work performed experimental tests using oxygen/paraffin as the propellant in a laboratory-scale hybrid rocket engine, applying oxidizer-to-fuel ratios in the range of 2.5–2.9 corresponding to combustion chamber pressures of 1.68–2.46 MPa. Variations in temperature and H 2 O partial pressure at the nozzle outlet were determined adopting TDLAS based on H 2 O absorption near 2.5 μm. Three H 2 O absorption lines at 4029.5, 4030.6 and 4030.7 cm−1 were simultaneously measured using only one diode laser at 2.0 kHz repetition rate and adopting a scanned-wavelength direct absorption strategy. The airflow velocity at the nozzle outlet was ascertained in two-dimensional numerical simulations. In this manner, the time-resolved fuel mass flow rate participating in combustion can be derived from the ideal state equation. A detailed discussion of the measurement uncertainty was provided herein considering the influence from determination of temperature, H 2 O partial pressure and velocity. The effectiveness of the novel method was validated by comparing its results with that acquired with the traditional weight-loss technique. Finally, it is initially analyzed that the combustion efficiency of pure paraffin fuel gradually increases during the combustion process according to the measurement results. • An optical fuel mass flow measurement method for hybrid rocket engines is proposed. • The quantitative fuel mass flow rate participating in combustion is obtained. • A compact tunable diode laser absorption spectroscopy system was designed. • Variations of temperature and H 2 O partial pressure at the nozzle exit were diagnosed. • Combustion characteristics of the oxygen/paraffin hybrid rocket were discussed. [ABSTRACT FROM AUTHOR]

To determine the suitability of the control performance of the VTVL system with hybrid rocket engine, simulations were done for the vertical and horizontal motion using various engine data. In the simulation, the thrust modulation characteristics of each engine were modeled with the same second-order transfer function in all cases. The results showed that the VTVL system with HRE showed better performance in terms of vertical motion error and touch down speed. In addition, this confirms the possibility for successful VTVL flight missions using HRE. In the next step, a better control algorithm and an efficient control technique will be attempted for developing the VTVL demonstrator. [ABSTRACT FROM AUTHOR]

Burning experiments were conducted to better understand the fuel regression behavior in a swirling-injection end-burning hybrid rocket engine using paraffin wax/gaseous oxygen propellant. The oxidizer mass flow rate, grain diameter, and the distance between the oxidizer injector and the grain end-surface were the variable parameters taken as influencing the regression rate. The engine attained an overall axial regression rate as high as approximately 5 mm/s, whereas unstable combustion occurred with increasing burning time owing to low melting temperature of paraffin wax. The fuel grain with a diameter of 90 mm also resulted in unstable combustion caused by the initial shallow crack of the cast grain. The radial distribution of the local regression rate exhibited dependency on the radial position and had two peaks: close to the periphery and the middle of the chamber. From the analogy of the heat transfer at the end surface in a vortex flow chamber, the controlling parameter of the overall axial regression rate was derived. Since this parameter depends on the distance between the oxidizer injector outlet and the fuel grain end-surface, to hold the grain end-surface by using some kind of actuator is necessary to attain constant O/F combustion.

To understand the initiating mechanism of low frequency instability (LFI) in hybrid rocket combustion, numerical calculation was performed using a direct numerical simulation (DNS). Special attention was paid to the effect of the equivalence ratio on the shift in the phase difference between pressure (p') and heat release oscillations (q'). A change in the equivalence ratio was considered by varying the temperature and composition of the combustion gas that entered into the post combustion chamber. In the obtained results, additional combustion was successfully simulated along with a vortex generation over the backward step, and combustion pressure and heat release oscillations were clearly reproduced as small vortices that move downward. In addition, the obtained results confirmed that the phase difference between pressure and heat release oscillations shifted because of the changes in the propagation speed of the pressure wave when the temperature of the combustion gas changes. [ABSTRACT FROM AUTHOR]