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High Test Peroxide (HTP) is a well-known green and energetic material, widely used in propulsion systems in a variety of applications, including hypergolic systems. This experimental work mixed fumed silica with liquid HTP to synthesize Gelled HTP (GHTP). A high-speed camera and a time-resolved infrared camera were employed to analyze liquid/gel phase mixing, ignition, and flame spread through drop tests, adopting a promising low toxicity liquid fuel as a GHTP hypergolic pair. Liquid/gel impinging jets in ambient air were also performed, simulating a basic element of a hybrid liquid/gel propulsion system. Prior to ignition, GHTP revealed a lower mixing ratio compared to HTP, which ultimately led to an ignition delay time of 20.6 ±6.4 ms, while HTP reached 14.6 ±3.7 ms when 1 wt% of catalyst was used. The flame area, used as a parameter to quantify gas-phase reaction, also showed sensitivity to liquid/gel phase mixing. Using a less concentrated HTP solution, ignition of GHTP/liquid fuel seemed to occur around 6 s faster than liquid HTP/liquid fuel in impinging jets, probably due to differences in propellant residence time and the vapor formation process. A theoretical performance analysis considering GHTP as an oxidizer in a hybrid liquid/gel hypergolic system revealed that the gelled propellant, made with the addition of 6 wt% of fumed silica, can achieve 96% of liquid HTP's specific impulse with the selected fuel. Furthermore, the theoretical density specific impulse of the hybrid gel/liquid system would significantly surpass that of the liquid/liquid system if 10 wt% of aluminum were mixed with GHTP. To take advantage of the hybrid system, a new propulsion system composed of pistons coupled to propellant tanks is proposed to properly infuse the gelled and liquid propellants. The results from this experimental work, together with theoretical analysis, suggest that the hybrid liquid/gel system may be suitable for military applications. [Display omitted] • Use of gelled hydrogen peroxide as oxidizer in hypergolic propulsion systems. • Drop and impinging jets test reveals vapor and ignition delay times with green fuel. • Analysis before and after ignition using visible and infrared high-speed cameras. • Infrared data quantifies mixing prior to ignition and flame spread. • Gelled hydrogen peroxide may ignite faster than its liquid counterpart. [ABSTRACT FROM AUTHOR]

Hybrid rockets have many attractive properties but have not yet been used in any high performance applications, mainly due to the low regression rate of the solid fuels used. In this work, a number of crystalline organic fillers were assessed with the aim to increase the regression rate and density of HTPB‐based fuels. Hexamine, which was the most promising filler, catalyses the HTPB/isocyanate crosslinking reaction leading to unacceptable short pot life. However, it was found that the pot life could be substantially prolonged by the use of less reactive brands of HTPB, added plasticizer and reduced NCO/OH ratio. Fuel formulations containing 75 % hexamine were easy to cast and the cured slabs were of good quality, free from bubbles or voids, with a density 30 % higher than plain HTPB. The tensile strength of the fuels were similar to typical solid propellants but the elasticity might need to be improved. Hexamine decompose at substantially lower temperature than HTPB. This, and its high density, makes it promising for use in hybrid rocket fuels. [ABSTRACT FROM AUTHOR]

This paper describes a new active method of controlling the position of the burning surface of a solid fuel strand in an opposed-flow burner experiment. This method is designed to avoid biases to the combustion process introduced by retaining wires and to avoid experimental delays caused by wire failures. Regression rates are presented for hydroxyl-terminated polybutadiene fuel strands burning in various oxidizer streams. The active control method uses a control loop composed of a diode laser, a photodiode, and a stepper motor. For comparison, tests were also performed using the commonly used passive control method of spring-loading the fuel against a retaining wire. This active control method can only be used at high values of oxidizer mass flux; at low mass flux, soot accumulation obstructs the laser beam. In passive control tests, the retaining wire conducts heat from the flame into the fuel and increases regression rate. The presence of the wire also modifies the extinction limits of the flame. A model was developed for regression rate as a function of oxidizer mass flux, oxidizer composition, and retaining wire diameter. This model can be used to correct future regression rate data for the bias introduced by retaining wires. [ABSTRACT FROM AUTHOR]

Fukuoka University offers opportunities for students to participate in manufacturing based on rockets. The objectives are to provide an environment in which students can practice the knowledge and skills learned in classroom lectures, support students manufacturing activities, develop space transportation engineers, and provide an environment where they can feel connected to society. The educational effects were evaluated based on the questionnaire results. This initiative is helping students on their path to becoming engineers. It also contributes to solving problems faced by the space industry. The effectiveness of education in terms of technology was confirmed, and 70% of students were able to utilize what they learned in their work even after graduation. Higher educational effects can be obtained by providing financial and management support from students and teachers. [ABSTRACT FROM AUTHOR]

The present study explores the influence of protrusions on the combustion stability of a hybrid rocket motor. The primary aim is to analyze single, double, and triple-protrusion configurations taking into account variables such as their location, material, and spacing between them. Protrusions used were made of Nylon (regress during combustion) and graphite (non-regress during combustion). For the base-case (without protrusion), significant pressure oscillations were observed, arising from the interaction between vortex shedding and fundamental longitudinal acoustic modes of the chamber at the fuel grain exit. A protrusion at 0.8X/L (X-protrusion insertion point from head-end, L-fuel grain length) disrupted the interaction between acoustic modes and vortex shedding, substantially dampened its amplitude (around 96% reduction). In case of double (0.2–0.8X/L, 0.5–0.8X/L) and triple-protrusions (0.2–0.5-0.8X/L), the distance between protrusions is playing a key role in inducing the pressure oscillations. When protrusions are kept close (0.5–0.8X/L and 0.2–0.5-0.8X/L), the excitation of third mode acoustics were observed. This was attributed to the continuous increase in protrusion height. If protrusion height exceeds 6–7 mm, the flow over the front protrusion separated and impinges on the subsequent protrusion. This causes pressure oscillations in double and triple-protrusion cases. These pressure oscillations were mitigated using regressable protrusions made of nylon, as they keep the height within the range of 6–7 mm. The results of this research provide valuable perspectives on the significance of protrusions in controlling combustion instability in hybrid rockets. • Influence of Single and multi-protrusion on combustion stability of hybrid rocket was investigated. • Hybrid rocket combustion instabilities associated with acoustic modes-vortex shedding were detected. • A single protrusion near the fuel grain exit suppressed the excited acoustic mode instabilities. • For multi-protrusions, distance between the protrusions plays a key role in stability of the hybrid rocket. • Significance of protrusion material in controlling instability was demonstrated. [ABSTRACT FROM AUTHOR]

The propulsion research lab at Utah State University has developed a minimally-intrusive optical sensing system for high-temperature/high-velocity gas-generator exhaust plumes. For this application glass fiber-optic cables, acting as radiation conduits, are inserted through the combustion chamber or nozzle wall and look directly into the flow core. The cable transmits data from the flame zone to externally-mounted spectrometers. In order to capture the full-optical spectrum, a blended dual-spectrum system was employed, with one spectrometer system tuned for best-response across the visible-light and near-infrared spectrum, and one spectrometer tuned for best-response in the near- and mid-infrared spectrum. The dual-band sensors are radiometrically-calibrated and the sensed-spectra are spliced together using an optimal Wiener filtering algorithm to perform the deconvolution. The merged spectrum is subsequently curve-fit to Planck's black-body radiation law, and flame temperature is calculated from associated curve maxima (Wien's law). The presented fiber-optic sensing systems performs a function that is analogous to Raman spectroscopy. The system non-contact, high-temperature measurement, and does not interfere with the heat transfer processes. In this report data collected from a lab-scale (200 N) hybrid rocket system are analyzed using the described method. Optically-sensed flame-temperatures are correlated to analytical predictions, and shown to generally agree within a few degrees. Additionally, local maxima in the optical spectra are shown to correspond to emission frequencies of atomic and molecular oxygen, water vapor, and molecular nitrogen; all species known to exist in the hybrid combustion plume. Presented data demonstrate that selected fiber-optics can survive temperature greater than 3000 ℃, for durations of up to 25 seconds.

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 H 2 O and CO 2 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. • The characteristic structure improved the combustion properties of composite grain. • Combustion properties improving mechanisms were studied numerically and experimentally. • Radiation imaging was employed to assess the flame pulsation features. • Experimental results were consistent well with the numerical results. [ABSTRACT FROM AUTHOR]

High-speed submerged marine vehicles, such as torpedoes, traveling at velocities of an order of 100 m/s and above, require powerful propulsion to overcome the tremendous hydrodynamic drag. This paper aims to investigate a marine hybrid-propellant water-breathing ram rocket (marine ramjet or ducted rocket) under various underwater cruise conditions. At high underwater cruise speeds, the ram rocket outperforms regular rocket motors, substantially increasing its specific impulse and thrust. This investigation utilized a unique test facility capable of dynamically testing the marine ramjet. Over 20 dynamic experiments have been conducted, revealing the submerged motor characteristics at different cruise speeds, water-to-propellant mass ratios, and oxidizer-to-fuel mass ratios, thereby creating a performance map of the marine ramjet. The results were compared with static firing data and theoretical calculations, showing a good agreement with standard specific impulse improvement of about 55% compared to a regular hybrid rocket, reaching a maximum value of 380 s. The significant increase in performance demonstrates the potential of the water-breathing ramjet for propelling high-speed underwater vehicles. [ABSTRACT FROM AUTHOR]

The propulsion research lab at Utah State University has developed a minimally-intrusive optical sensing system for high-temperature/high-velocity gas-generator exhaust plumes. For this application glass fiber-optic cables, acting as radiation conduits, are inserted through the combustion chamber or nozzle wall and look directly into the flow core. The cable transmits data from the flame zone to externally-mounted spectrometers. In order to capture the full-optical spectrum, a blended dual-spectrum system was employed, with one spectrometer system tuned for best-response across the visible-light and near-infrared spectrum, and one spectrometer tuned for best-response in the near- and mid-infrared spectrum. The dual-band sensors are radiometrically-calibrated and the sensed-spectra are spliced together using an optimal Wiener filtering algorithm to perform the deconvolution. The merged spectrum is subsequently curve-fit to Planck's black-body radiation law, and flame temperature is calculated from associated curve maxima (Wien's law). The presented fiber-optic sensing systems performs a function that is analogous to Raman spectroscopy. The system non-contact, high-temperature measurement, and does not interfere with the heat transfer processes. In this report data collected from a lab-scale (200 N) hybrid rocket system are analyzed using the described method. Optically-sensed flame-temperatures are correlated to analytical predictions, and shown to generally agree within a few degrees. Additionally, local maxima in the optical spectra are shown to correspond to emission frequencies of atomic and molecular oxygen, water vapor, and molecular nitrogen; all species known to exist in the hybrid combustion plume. Presented data demonstrate that selected fiber-optics can survive temperature greater than 3000 ℃, for durations of up to 25 seconds. [ABSTRACT FROM AUTHOR]

A Hybrid Rocket Motor (HRM) is a type of chemical rocket propulsion in which the propellants are stored in different physical states. To counter the low regression rate characteristic of HRM different grain configurations with innovative port techniques has been evolved with time. It creates better mixing and combustion of fuel with oxidizer without having any special injector and energetic additives. A similar technique has been used in the present study. The fuel used in the present study is a solid grain made from polyvinyl chloride with di-butyl phthalate in the 50:50 ratios with gaseous oxygen as an oxidizer. In this paper, analyses are made to study the performance of a hybrid rocket motor by varying the axial alignment of the grain port at varying locations along the length. Due to the offset of the fuel port at varying locations, recirculation zone were created that enhanced the pressure in the combustion chamber by around 1-2 bar. The thrust generated was increased in the range of 10 to 25 N. Regression rate as well as efficiency was also observed to be increased with the use of this port alignment techniques. [ABSTRACT FROM AUTHOR]

The performance of two solid biomass wastes, date stone and jojoba solid waste, was experimentally examined for their potential application in combustion and propulsion systems. The fuels were tested in a hybrid rocket-like combustion environment, and the test result was analyzed with combustion and propulsion parameters. The performance of both fuels was comparatively evaluated and compared with a conventional hydrocarbon fuel in a hybrid rocket, with paraffin wax serving as a baseline. A compression device was introduced to compress the solid biomass wastes into a circular-shaped fuel grain compatible with a hybrid rocket combustion chamber with a hot surface ignitor. Thermogravimetric analysis (TGA) and chemical equilibrium analysis (CEA) results revealed that the performance of the biomass fuel can be comparable to conventionally used hydrocarbon paraffin-wax-based propellant within a certain range of oxidizer-to-fuel ratio, in terms of theoretical specific impulse performance. Through experimental performance tests, it was found that the compressed biomass fuel grains were successfully ignited and produced thrust. Both biomass fuels tested in a hybrid rocket combustion chamber are expected to pave the way for further developments in biomass fuels in the waste-to-energy field for their application in combustion and propulsion systems, potentially replacing fossil fuels with renewable resources. [ABSTRACT FROM AUTHOR]

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. [ABSTRACT FROM AUTHOR]

Paraffin-based fuel has a great potential for several innovative missions, including space tourism, due to its safety, low environmental impact, high performance and low cost. Despite the fact that liquefying solid fuels increases the regression rate of hybrid rocket motors, incorporating energetic materials into solid fuel can still improve the performance. The objective and scope of this study is to increase the performance characteristics of the paraffin-based fuel by using magnesium diboride (MgB2) and carbon black (CB) additives. The cylindrical-port fuel grains were manufactured with various additives percentages in mass (wt%: CB-2% and MgB2-10%) and tested using a laboratory-scale ballistic hybrid motor under gaseous oxygen. The mechanical performance results revealed that adding CB and MgB2 improved the ultimate strength and elastic modulus of paraffin-based fuels. The addition of these fillers increased the hardness of fuel by developing a strong interaction in the paraffin matrix. Thermogravimetry (TG) results showed that CB inclusion improved the thermal stability of the paraffin matrix. The average regression rates of fuels loaded with CB and MgB2 were 32% and 52% higher than those of unmodified paraffin wax, respectively. The characteristic velocity efficiency was found in the range of 68%–79% at an O/F ratio of 1.5–2.6. The MgB2 oxidation/combustion in the paraffin matrix was described by a four-step oxidation process ranging from 473 K to 1723 K. Finally, a combustion model of MgB2 in the paraffin matrix was proposed, and four-step oxidation processes were discussed in detail.

A one-dimensional model of the cooling process occurring in a hybrid-rocket nozzle working with cryogenic oxygen flowing through helical channels is addressed in this work. Although the regenerative cooling system in rocket engines involves further complexity, it is here investigated as an option for the suppression of graphite-nozzle erosion in hybrid rockets. The model is based on the solution of the conservation equations of mass, momentum, and energy of the coolant. The Modified Benedict-Webb-Rubin equation of state was used for its accuracy in both the fluid phases of oxygen. The conservation equations for the coolant flow are coupled with a one-dimensional heat transfer model for the evaluation of the nozzle wall temperature at both the cold and hot sides, which has allowed assessing the developed model prediction capabilities by means of the data collected from three engine hot firings. Thus, parametric analyses were carried out to show the effect of the number of helical channels and evaluate the performance improvement obtained in supercritical conditions. • Study of a hybrid rocket nozzle cooling system based on cryogenic oxygen flowing through helical cooling channels. • 1D modelling of the nozzle cooling with cryogenic oxygen flowing through helical channels. • Validation on firing tests by transient simulations. • Effect of the number of channels in a helical configuration. [ABSTRACT FROM AUTHOR]

High-speed submerged marine vehicles, such as torpedoes, traveling at velocities of an order of 100 m/s and above, require powerful propulsion to overcome the tremendous hydrodynamic drag. This paper aims to investigate a marine hybrid-propellant water-breathing ram rocket (marine ramjet or ducted rocket) under various underwater cruise conditions. At high underwater cruise speeds, the ram rocket outperforms regular rocket motors, substantially increasing its specific impulse and thrust. This investigation utilized a unique test facility capable of dynamically testing the marine ramjet. Over 20 dynamic experiments have been conducted, revealing the submerged motor characteristics at different cruise speeds, water-to-propellant mass ratios, and oxidizer-to-fuel mass ratios, thereby creating a performance map of the marine ramjet. The results were compared with static firing data and theoretical calculations, showing a good agreement with standard specific impulse improvement of about 55% compared to a regular hybrid rocket, reaching a maximum value of 380 s. The significant increase in performance demonstrates the potential of the water-breathing ramjet for propelling high-speed underwater vehicles.

Hybrid rockets are safe and inexpensive; however, boundary-layer combustion poses a problem in achieving a fuel regression rate equivalent to that of solid propellants. The fundamental combustion conditions, such as the fuel regression rate of LT421, a paraffin-based low-melting-point thermoplastic fuel, were investigated using a swirling-flow combustion method. Firing tests were conducted using the oxygen mass flow rate and burn time parameters. The LT fuel exhibited an ignition delay compared to polypropylene, and the pressure increased slowly relative to the thrust. The combustion pressure increased or remained constant with time, suggesting that the fuel regression rate was more dependent on the oxygen mass flow rate than the oxidizer mass flux. The shear force generated in the grain owing to the swirling flow caused fuel-grain separation when the oxygen mass flow rate exceeded 100 g/s. Fuel-grain separation was prevented by modifying the case geometry. The maximum fuel regression rate obtained in the tests was 4.88 mm/s at an oxygen mass flow rate of 190 g/s and mass flux of 72.4 kg/(m2s), which was four times higher than that of polypropylene at the same oxidizer mass flux. The fuel regression rate correlation was obtained using the oxygen mass-flow-rate-based parameter, although further modification was necessary to apply this correlation when the burning time was varied. [ABSTRACT FROM AUTHOR]

Hybrid rocket engine (HRE) performance is dependent on fuel/oxidizer selection and fuel grain geometry. A literature review was performed to identify key trends and findings related to the application of the additive manufacturing (AM) of fuel systems for HREs. The effects of complex combustion port geometries, embedded structures, and end-burning systems, along with the use of metallic additives, turbulators, diaphragms, gel-like fuels, powdered fuels, liquid fuels, and liquifying fuels and their impact on regression rates, combustion efficiencies, and/or mechanical strength are thoroughly documented here. In general, the application of AM to HRE fuels can be implemented to increase regression rates and combustion efficiency, and tailor HRE designs. Chemical equilibrium analysis computations were completed to characterize the theoretical performance of HTPB and common AM fuels (ABS, PLA, PC, PMMA, Nylon 6, and a UV-based fuel) with common oxidizers (LOX and N2O). AM fuels exhibit a similar theoretical performance as the commonly used HTPB fuel, and proper selection of the fuel can yield improved performance and design metrics. Development of AM approaches for HRE fuel design have significantly expanded their design trade space and should enable the competitive application of HREs for future propulsion missions. [ABSTRACT FROM AUTHOR]

A three-dimensional-printed solid fuel grain with a star-fractal swirl port was developed to improve the fuel regression rate of hybrid rockets. The printing material used for the grains was an acrylonitrile-butadiene-styrene resin. Nitrous oxide was employed as the oxidizer. Combustion experiments on the hybrid rocket engine with the developed grain showed that the star-fractal swirl port appeared to improve the thrust and specific impulse compared to the classical circular and starfractal ports when tested using the same oxidizer mass flow rate that is a controllable parameter of hybrid rocket engines. After the experiments, the burned grains were cut and the grain web thicknesses were measured to evaluate the local regression rate. It was found that the fuel surface regressed more rapidly near the concave parts of the star fractal rather than the convex parts. Next, the measured local regression rates were axially averaged. The axially-averaged local regression rates of the star-fractal swirl port were much higher than those of the star-fractal and circular ports. Finally, a numerical approach using the level set method for evaluating time- and space-averaged regression rate was developed. This approach was validated by comparing the space-averaged regression rates with the axially-averaged local rates. [ABSTRACT FROM AUTHOR]

Hybrid rockets have distinct advantages over their pure solid or liquid propellant counterparts, and their performance can be improved by inclusion of metal additives. Several metallic additives (micro-Al, micro-Ti, micro-Mg, micro-Zr, nano-Al, nano-B, and Mg-coated nano-B) were selected as potential candidates for hybrid rocket applications and characterized by applicable microscopy techniques. The regression rates and combustion efficiencies of plain HTPB and HTPB loaded with each additive at various concentrations (10%, 20%, and 30% by mass) burning in GOX were evaluated at moderate oxidizer mass fluxes (10–150 kg/m2-s) and pressures ( ≤ 0.86 MPa, 125 psia). In general, the inclusion of any of the metallic additives led to a reduction in the regression rate and did not significantly change the combustion efficiency. The only exceptions were fuel formulations containing micro-Zr, which yielded a moderate (10–20%) increase in the regression rate at a concentration of 10%. The observed trends were more prevalent at higher oxidizer mass fluxes and higher additive loadings. The reductions in regression rate were attributed to heat transfer blocking effects derived from accumulation of additive particles on the fuel surface layer. These phenomena were especially prevalent in highly loaded fuel formulations containing the nano-additives that exhibited unstable combustion and periodic surface-layer shedding. Zirconium appears to be the best metallic additive available since it can yield the highest theoretical density-specific impulse under the lowest O/F operation ratio without resulting in decrements to overall performance. Notably, combustion efficiency data for all fuel formulations were well correlated to the combustion residence time, and high combustion efficiencies (>95%) were achievable when a satisfactory residence time (~75 ms) was realized. [ABSTRACT FROM AUTHOR]

Gel fuel droplets exhibit disruptive burning due to the rupture of their gellant shell, which causes the release of unreacted fuel vapors from the droplet interior to the flame in the form of jets. In addition to pure vaporization, this jetting allows convective transport for fuel vapors, which accelerates gas-phase mixing and is known to improve droplet burn rates. Using high-magnification and high-speed imaging, this study found that the viscoelastic gellant shell at the droplet surface evolves during the droplet's lifetime, which causes the droplet to burst at different frequencies, thereby triggering a time-varying oscillatory jetting. In particular, the continuous wavelet spectra of the droplet diameter fluctuations show that the droplet bursting exhibits a nonmonotonic (hump-shaped) trend, where the bursting frequency first increases and then decreases to a point where the droplet stops oscillating. The changes in the shell structure are captured by tracking the temporal variation of the area of rupture sites, spatial movement of their centroid, and the degree of overlap between the rupture areas of successive cycles. During the initial period (immediately following its formation) when the shell is newly formed, it is weak and flexible, which causes it to burst at increasingly high frequencies. This is because the area at and around the rupture site becomes progressively weaker with each rupture in an already weak shell. This is shown by a high degree of overlap between the areas of successive ruptures. On the other hand, the shell flexibility during the initial period is demonstrated by a reversal in the motion of rupture site centroids. However, at later stages when the droplet has undergone multiple ruptures, the depletion of the fuel vapor causes accumulation of gellant on the shell, thus causing the shell to become strong and rigid. This thick, strong, and rigid shell suppresses droplet oscillations. Overall, this study provides a mechanistic understanding of how the gellant shell evolves during the combustion of a gel fuel droplet and causes the droplet to burst at different frequencies. This understanding can be used to devise gel fuel compositions that result in gellant shells with tailored properties, and therefore, control the jetting frequencies to tune droplet burn rates. [ABSTRACT FROM AUTHOR]

The performance of two solid biomass wastes, date stone and jojoba solid waste, was experimentally examined for their potential application in combustion and propulsion systems. The fuels were tested in a hybrid rocket-like combustion environment, and the test result was analyzed with combustion and propulsion parameters. The performance of both fuels was comparatively evaluated and compared with a conventional hydrocarbon fuel in a hybrid rocket, with paraffin wax serving as a baseline. A compression device was introduced to compress the solid biomass wastes into a circular-shaped fuel grain compatible with a hybrid rocket combustion chamber with a hot surface ignitor. Thermogravimetric analysis (TGA) and chemical equilibrium analysis (CEA) results revealed that the performance of the biomass fuel can be comparable to conventionally used hydrocarbon paraffin-wax-based propellant within a certain range of oxidizer-to-fuel ratio, in terms of theoretical specific impulse performance. Through experimental performance tests, it was found that the compressed biomass fuel grains were successfully ignited and produced thrust. Both biomass fuels tested in a hybrid rocket combustion chamber are expected to pave the way for further developments in biomass fuels in the waste-to-energy field for their application in combustion and propulsion systems, potentially replacing fossil fuels with renewable resources.

A hydrogen-oxygen torch igniter was developed and utilized to study its effects on a solid acrylonitrile butadiene styrene thermoplastic (ABS)-gaseous oxygen (GOx) hybrid rocket motor. The effects thereof are not presently found in literature. An increase of the hydrogen mass flow rate, at constant GOx mass flow rate, led to an increase in the average linear regression rate and a decrease in characteristic ignition time for fuel-lean and stoichiometric hydrogen-oxygen flames. Additionally, extending the duration of hydrogen flow from 10% to 50% of the total burn time led to an increase in the regression rate. Results suggest that the hydrogen-oxygen torch igniter can be used as a reliable starter with reignition capabilities and also for regression rate and ignition delay enhancement that targets the rate-limiting step, pyrolysis, for hybrid rocket combustion at stoichiometric and fuel-lean operating conditions of the torch igniter. • An increase in hydrogen mass flow rate can increase average linear regression rate. • An increase in hydrogen mass flow rate can decrease characteristic ignition time. • At fuel-rich torch operation, there is a competition between hydrogen and the ABS. • Extending hydrogen duration led to an increase in average linear regression rate. [ABSTRACT FROM AUTHOR]

The study intends to investigate the physical, chemical and thermal characteristics of paraffin blended fuels to determine their suitability as a solid fuel in a hybrid rocket. Wax fuels are a viable and efficient alternative to conventional rocket fuels, having excellent structural strength and thermal and mechanical properties. By utilizing both axial and swirl injection technique, the combustion performance of paraffin -- beeswax blended fuels have been tested with a fabricated cylindrical grain in a laboratory-scale rocket setting along with oxygen. The test outcomes revealed solid fuel compositions of more beeswax content in paraffin wax on an oxygenated gaseous environment with a swirl-flow injection method has the highest average regression rate of 1.649 mm/sec at 181 kg/m²s mass flux. Axially injected oxygen with pure paraffin wax has the lowest value of 0.85 mm/sec at 96 kg/m²s. The regression rate comparisons revealed that oxygen injection by a swirl injector increased the regression rates by 40% for mass fluxes greater than 80 kg/m²s. Compared to other studies, the combustion efficiencies have been obtained in this study are good. Blended fuels can manage and increase combustion efficiencies for axial and swirl flow conditions. Swirl injectors outperform axial injectors for oxygen injection and allow for a higher proportion of Beeswax combined with paraffin. This study exclusively designed and manufactured an axial injector and swirl injector, according to the required dimensions of a lab-scale hybrid rocket's combustion chamber, injector, and exhaust nozzle, and their performances have been evaluated. [ABSTRACT FROM AUTHOR]

Jetting in burning gel fuel droplets is an important process which, in addition to pure vaporization, enables the convective transport of unreacted fuel vapors from the droplet interior to the flame envelope. This aids in accelerating the fuel efflux and enhancing the mixing of the gas phase, which improves the droplet burn rates. In this study, Schlieren imaging was used to characterize different jetting dynamics that govern the combustion behavior of organic-gellant-laden ethanol gel fuel droplets. To initiate jetting, the gellant shell of the burning gel fuel droplet was subjected to either oscillatory bursting or isolated bursting, or both. However, irrespective of the jetting mode, the jets interacted with the flame envelope in one of three possible ways. Based on the velocity and the degree to which a jet disrupts the flame envelope, it is classified as either a flame distortion, a fire ball outside the flame or a pin hole jet (localized flame extinction), where the pin hole jets have the highest velocity (1000–1550 mm/s), while the flame distortion events have the lowest velocity (500–870 mm/s). Subsequently, the relative number of the three types of jetting events during the droplet lifetime was analyzed as a function of the type of organic gellant. It was demonstrated that the combustion behavior of gel fuels (hydroxypropyl methylcellulose: HPMC at 3 wt.%) that tend to form thin-weak-flexible shells is dominated by low-velocity flame distortion events, while the gel fuels (methylcellulose: MC at 9 wt.%) that facilitate the formation of thick-strong-rigid shells are governed by high-velocity fire ball and pin hole jets. Overall, this study provides critical insights into the jetting behavior and its characterization, which can help us to tune the droplet gasification and the gas phase mixing to achieve an effective combustion control strategy for gel fuels. [ABSTRACT FROM AUTHOR]

Human spaceflight, space tourism, and the launch of microsatellites are all expected to grow in the future. In fact, with the recent increase of the satellite market, almost 1000 smallsats per year are foreseen to be launched over the next decade. Hybrid rocket propulsion has received significant attention for military and commercial applications due to its potential safety, throttleable, and restart ability when compared with solid rockets, economicity, simplicity, and compactness features when compared to liquid rockets. However, in order to make this new technology the future of the next generation of rockets, some drawbacks of the heterogeneous combustion in hybrid rockets such as low fuel regression rate and varying oxidizer-to-fuel ratio during the combustion process must be well addressed. The diffusion-limited combustion in hybrid rocket motor is responsible for the low regression and poor combustion efficiency of fuels such as Hydroxyl-terminated polybutadiene (HTPB), Poly methylmethacrylate (PMMA), and other polymeric binder-fuels. The paraffin-based solid fuel represents a potential solution to the slow regression rate of current solid polymeric fuels. However, paraffin-based fuels suffer from poor mechanical properties and rapid volatilization, preventing their full development and applications for a space mission. In this work, a review of various techniques to improve hybrid rocket fuel's ballistic and mechanical performance is presented.

Hybrid rockets are safe and inexpensive; however, boundary-layer combustion poses a problem in achieving a fuel regression rate equivalent to that of solid propellants. The fundamental combustion conditions, such as the fuel regression rate of LT421, a paraffin-based low-melting-point thermoplastic fuel, were investigated using a swirling-flow combustion method. Firing tests were conducted using the oxygen mass flow rate and burn time parameters. The LT fuel exhibited an ignition delay compared to polypropylene, and the pressure increased slowly relative to the thrust. The combustion pressure increased or remained constant with time, suggesting that the fuel regression rate was more dependent on the oxygen mass flow rate than the oxidizer mass flux. The shear force generated in the grain owing to the swirling flow caused fuel-grain separation when the oxygen mass flow rate exceeded 100 g/s. Fuel-grain separation was prevented by modifying the case geometry. The maximum fuel regression rate obtained in the tests was 4.88 mm/s at an oxygen mass flow rate of 190 g/s and mass flux of 72.4 kg/(m2s), which was four times higher than that of polypropylene at the same oxidizer mass flux. The fuel regression rate correlation was obtained using the oxygen mass-flow-rate-based parameter, although further modification was necessary to apply this correlation when the burning time was varied.

The low toxicity hypergolic solid fuels (HSFs) and hydrogen peroxide (H 2 O 2) oxidizer have been demonstrated as environment-friendly hybrid propellants. In the present study, a new HSF containing ammonia borane (AB), Pd–C, and paraffin wax, was fabricated in the form of the hypergolic ignitor for the hybrid rocket. Furthermore, the atmospheric combustion test (ACT) and hot-fire test (HFT) with HSF ignitor, PMMA or PE fuel, and 95% H 2 O 2 oxidizer were performed using fuel-exposed-hybrid-combustor; studied the ignition reliability and combustion characteristics of HSF. The hypergolic ignition of HSF initiated through Pd–C followed by AB combustion with H 2 O 2 was observed by using the high-speed camera images of ACT. A HFT with a ground thrust of 120 N exhibited the rising time and pressure instability of 137 ms and ±4.00%, respectively. Moreover, smooth combustion transition from HSF to PE was achieved with stable chamber pressure and no hard-start. The results suggest that the HSF ignitor containing AB may be the suitable preference for the hypergolic hybrid rocket. • Demonstration of a new ammonia borane-based hypergolic solid fuel ignitor for the hybrid rocket. • Steady-state combustion of hybrid fuel with 95 wt% H 2 O 2 at rocket conditions. • Hypergolic hybrid fuel ignitor: environmentally-friendly, low-cost, and alternative for toxic rocket propellants. [ABSTRACT FROM AUTHOR]

Hybrid rocket engine (HRE) performance is dependent on fuel/oxidizer selection and fuel grain geometry. A literature review was performed to identify key trends and findings related to the application of the additive manufacturing (AM) of fuel systems for HREs. The effects of complex combustion port geometries, embedded structures, and end-burning systems, along with the use of metallic additives, turbulators, diaphragms, gel-like fuels, powdered fuels, liquid fuels, and liquifying fuels and their impact on regression rates, combustion efficiencies, and/or mechanical strength are thoroughly documented here. In general, the application of AM to HRE fuels can be implemented to increase regression rates and combustion efficiency, and tailor HRE designs. Chemical equilibrium analysis computations were completed to characterize the theoretical performance of HTPB and common AM fuels (ABS, PLA, PC, PMMA, Nylon 6, and a UV-based fuel) with common oxidizers (LOX and N2O). AM fuels exhibit a similar theoretical performance as the commonly used HTPB fuel, and proper selection of the fuel can yield improved performance and design metrics. Development of AM approaches for HRE fuel design have significantly expanded their design trade space and should enable the competitive application of HREs for future propulsion missions.

Gel fuel droplets exhibit disruptive burning due to the rupture of their gellant shell, which causes the release of unreacted fuel vapors from the droplet interior to the flame in the form of jets. In addition to pure vaporization, this jetting allows convective transport for fuel vapors, which accelerates gas-phase mixing and is known to improve droplet burn rates. Using high-magnification and high-speed imaging, this study found that the viscoelastic gellant shell at the droplet surface evolves during the droplet’s lifetime, which causes the droplet to burst at different frequencies, thereby triggering a time-varying oscillatory jetting. In particular, the continuous wavelet spectra of the droplet diameter fluctuations show that the droplet bursting exhibits a nonmonotonic (hump-shaped) trend, where the bursting frequency first increases and then decreases to a point where the droplet stops oscillating. The changes in the shell structure are captured by tracking the temporal variation of the area of rupture sites, spatial movement of their centroid, and the degree of overlap between the rupture areas of successive cycles. During the initial period (immediately following its formation) when the shell is newly formed, it is weak and flexible, which causes it to burst at increasingly high frequencies. This is because the area at and around the rupture site becomes progressively weaker with each rupture in an already weak shell. This is shown by a high degree of overlap between the areas of successive ruptures. On the other hand, the shell flexibility during the initial period is demonstrated by a reversal in the motion of rupture site centroids. However, at later stages when the droplet has undergone multiple ruptures, the depletion of the fuel vapor causes accumulation of gellant on the shell, thus causing the shell to become strong and rigid. This thick, strong, and rigid shell suppresses droplet oscillations. Overall, this study provides a mechanistic understanding of how the gellant shell evolves during the combustion of a gel fuel droplet and causes the droplet to burst at different frequencies. This understanding can be used to devise gel fuel compositions that result in gellant shells with tailored properties, and therefore, control the jetting frequencies to tune droplet burn rates.

The paper presents a useful engineering model for the design of preliminary hybrid rocket engines. This model involves the experimentally obtained equation regarding the speed of the burning interface between solid fuel and gaseous phase, called the regression rate. This regression rate characterizes the mass rate of the burning fuel and additionally the mass rate of the oxidizer through the imposed ratio of these mass rates during combustion. The preliminary design is applied to combustion using pure HTPB and gaseous oxygen and was developed for three cases, constant regression rate, constant oxygen mass flow rate and constant O/F (ratio of mass rates of oxygen and of fuel). The design evaluates the initial fuel port geometry, the initial mass of fuel and oxygen, the combustion time, the thrust at sea level, and the time-dependent functions of regression rate, of fuel and oxygen mass rates, and of thrust. It was assumed that the regression rate formula applies at any time of combustion. These evaluations can also be applied for other fuel/oxygen couples by knowing the correct formula for the regression rate. [ABSTRACT FROM AUTHOR]

In hybrid rockets, low regression rate and low combustion efficiency are major drawbacks restricting its utilization in practical applications. Inserting a protrusion in a hybrid rocket combustion chamber is one of the methods to enhance the regression rate and combustion efficiency. In the present study, while utilizing the protrusion method the combustion stability of the hybrid rocket was investigated. The protrusions used for this study were of two different types; one was made of graphite and another one was nylon. The nylon protrusion can regress during the combustion along with the fuel whereas, graphite cannot. Before using the protrusion method, the combustion stability of the hybrid rocket motor (base case) was characterized. During the combustion of the base case motor, intense pressure oscillations were observed for a definite period after the beginning of the combustion. These oscillations were found to be the result of the interaction between the vortex shedding and fundamental longitudinal acoustic modes of the chamber at the fuel grain exit. Upon inserting the protrusion near the fuel grain exit, significant suppression in the pressure oscillations was observed. A 96 % reduction in amplitude of acoustic modes relative to the base case was observed. There was no substantial difference in combustion stability of the hybrid rocket when using graphite or nylon protrusions. Further, it was realized that the combustion stability of the rocket motor remains unchanged even after varying the thickness of the protrusion at 0.8X/L location. [ABSTRACT FROM AUTHOR]

In this paper, the experimental and numerical investigations of both swirling hybrid rocket motors without and with the design of regenerative cooling (RC hereafter), using 90%-wt H 2 O 2 and polypropylene (PP) as the liquid oxidizer and the solid fuel, respectively, were performed and compared. We implemented a new design of regenerative cooling using the oxidizer to transfer the heat generated from combustion. A series of ground horizontal hot-fire tests, which include the catalytic-bed only and the motor without/with RC were performed to measure the propulsion performance related parameters. The resulting C* efficiency of the RC design with catalytic-bed only remained in comparison with the one without RC. With a similar initial combustion chamber geometry, the RC implementation resulted in a better engine Isp efficiency from 96.3 to 98.6 with a lower O/F ratio from 6.49 to 5.94. In addition, the topologies of the burned fuel grain surface were optically measured by a 3D laser scanner. The resulting variation of regression rates between the upstream and downstream fuel for the chamber with the RC was smaller than case without the RC. Moreover, a simplified semi-empirical computational fluid dynamics (CFD) model based on the measured regression rates was constructed for investigating the physical phenomenon inside the hybrid rocket combustion chamber. The simulations show that a higher swirling number in the later portion of the port was found for the RC case. By concerning about the different O/F ratios and port size of the burned fuel, a series of simplified numerical CFD investigation with a fixed O/F and a fixed port size varying the regression rates from upstream to downstream fuel systematically with the same average region rate was performed to complement the semi-empirical CFD simulations. The results showed that there were two distinct regions including an upstream non-classical region dominated by the gas impingement of gas and a downstream classical region caused by the swirling effect of gases. In the non-classical region, the calculated distributions of the swirling numbers of the former are essentially the same for all the cases no matter what the magnitude of the regression rate is, while in the classical region higher swirling numbers are found for the case of more uniform distribution of regression rates. Heat generated near the upstream of the fuel grain was transferred more efficiently through the design of RC, which makes the resulting fuel regression more uniform than that without RC. The resulting swirling efficiency in the RC case was higher than that without RC. The HRM with RC hence had a better engine Isp efficiency with a lower O/F ratio. • Hybrid rocket motors with the design of regenerative cooling improves the propulsion efficiency. • Optical 3D scanner precisely measures the solid fuel topology. • The results of the proposed numerical model for hybrid rocket combustion are in good agreement with experiments. [ABSTRACT FROM AUTHOR]

As a green oxidizer, hydrogen peroxide is an earth‐storable, non‐toxic liquid oxidizer. This study shows a possible combination of wax‐based aluminized fuel with hydrogen peroxide as an oxidizer, which can be used to improve the specific impulse and density specific impulse of the hybrid rockets. In this study, various methods have been tried to ignite the hybrid rocket and measure the regression rates and ballistic performance. This research presents sustainable combustion of hydrogen peroxide with aluminized wax‐based fuel without the need for any catalyst bed. This combination of hybrid propellant demonstrates superiority over other fuels with hydrogen peroxide and shows promises in terms of an option as the future green propellant. The theoretical specific impulse of about 315 s at 100 bar with a nozzle exit area ratio of 50 is shown by this fuel and oxidizer combination. Another benefit of this combination of hybrid propellant is that the peak specific impulse is obtained at an O/F ratio of almost unity which makes the system more compact. [ABSTRACT FROM AUTHOR]

Jetting in burning gel fuel droplets is an important process which, in addition to pure vaporization, enables the convective transport of unreacted fuel vapors from the droplet interior to the flame envelope. This aids in accelerating the fuel efflux and enhancing the mixing of the gas phase, which improves the droplet burn rates. In this study, Schlieren imaging was used to characterize different jetting dynamics that govern the combustion behavior of organic-gellant-laden ethanol gel fuel droplets. To initiate jetting, the gellant shell of the burning gel fuel droplet was subjected to either oscillatory bursting or isolated bursting, or both. However, irrespective of the jetting mode, the jets interacted with the flame envelope in one of three possible ways. Based on the velocity and the degree to which a jet disrupts the flame envelope, it is classified as either a flame distortion, a fire ball outside the flame or a pin hole jet (localized flame extinction), where the pin hole jets have the highest velocity (1000–1550 mm/s), while the flame distortion events have the lowest velocity (500–870 mm/s). Subsequently, the relative number of the three types of jetting events during the droplet lifetime was analyzed as a function of the type of organic gellant. It was demonstrated that the combustion behavior of gel fuels (hydroxypropyl methylcellulose: HPMC at 3 wt.%) that tend to form thin-weak-flexible shells is dominated by low-velocity flame distortion events, while the gel fuels (methylcellulose: MC at 9 wt.%) that facilitate the formation of thick-strong-rigid shells are governed by high-velocity fire ball and pin hole jets. Overall, this study provides critical insights into the jetting behavior and its characterization, which can help us to tune the droplet gasification and the gas phase mixing to achieve an effective combustion control strategy for gel fuels.

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.

High thrust propulsion systems that can be used in small spacecraft are urgently needed to expand the use of small spacecraft. This study proposes a hybrid thruster for small spacecraft using wire-shaped magnesium as a fuel and water as an oxidizer. "Hybrid thruster" means that the in-space propulsion system generates thrust using a chemical reaction between a solid fuel and a vapor oxidizer, such as those utilized for hybrid rockets. Both magnesium and water are very safe, highly available, and very storable. To assess the feasibility of a hybrid thruster, we carried out experiments utilizing magnesium-wire combustion in water vapor at a pressure lower than the atmosphere, and estimated the input power for ignition, ignitability, and the average magnesium mass consumption rate. Then, propulsion performances were calculated using the experimental results, which show there is an input power of 2-3W, with higher ignitability and a lower average mass consumption rate for thinner magnesium wires. The calculated specific impulse achieved its maximum at 323 s, with a mass mixture ratio of 1.25. At that mass mixture ratio, the calculated thrust using magnesium wires having diameters of 0.5mm and 0.8mm was 49.8mN and 68.5 mN, respectively. Additionally, the applicability of the hybrid thruster was shown through a case study considering an existing deep-space misson. [ABSTRACT FROM AUTHOR]

Most hybrid rocket propellants comprise solid fuel and liquid oxidizer. The attractive characteristics of hybrid rockets are their high safety and controllable thrust. However, despite such benefits, hybrid rockets are not yet put into use because of their low regression rate and combustion efficiency. Therefore, many studies specifically examine improvement of the regression rate and combustion efficiency. Early efforts started in the 1960s; then both ammonium perchlorate and ammonium nitrate were employed as fuel additives. This study is conducted to evaluate ignition characteristics, theoretical calculation, and pyrolysis of a microcrystalline-wax based solid fuel adding ammonium nitrate as a solid oxidizer. We report whether the improvement can expect by the addition ammonium nitrate in this paper. [ABSTRACT FROM AUTHOR]