Journal articles on the topic 'Hybrid propellant rockets – Combustion'
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Izham Izzat Ismail, Norhuda Hidayah Nordin, Muhammad Hanafi Azami, and Nur Azam Abdullah. "Metals and Alloys Additives as Enhancer for Rocket Propulsion: A Review." Journal of Advanced Research in Fluid Mechanics and Thermal Sciences 90, no. 1 (December 25, 2021): 1–9. http://dx.doi.org/10.37934/arfmts.90.1.19.
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A rocket's engine usually uses fuel and oxygen as propellants to increase the rocket's projection during launch. Nowadays, metallic ingredients are commonly used in the rocket’s operation to increase its performance. Metallic ingredients have a high energy density, flame temperature, and regression rate that are important factors in the propulsion process. There is a wide range of additives have been reported so far as catalysts for rocket propulsion. The studies show that the presence of metal additives improves the regression rate, specific impulse and combustion efficiency. Herein, the common energetic additives for rocket propulsion such as metal and light metals are reviewed. Besides the effect of these energetic particles on the regression behaviors of base (hybrid) fuel has been exclusively discussed. This paper also proposed a new alloy namely high entropy alloys (HEAs) as a new energetic additive that can potentially increase the performance of the rocket propellant system.
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Palacz, Tomasz, and Jacek Cieślik. "Experimental Study on the Mass Flow Rate of the Self-Pressurizing Propellants in the Rocket Injector." Aerospace 8, no. 11 (October 26, 2021): 317. http://dx.doi.org/10.3390/aerospace8110317.
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High vapor pressure propellants such as nitrous oxide are widely used in experimental hybrid and liquid rockets as they can be used in a self-pressurization mode, eliminating the need for external pressurization or pumps and simplifying the design of the rocket system. This approach causes the two-phase flow in the feed system and the injector orifices, which cannot be easily modeled and accounted for in the design. A dedicated test stand has been developed to better understand how the two-phase flow of the self-pressurizing propellant impacts the mass flow characteristics, enabling the simulation of the operating conditions in the rocket engine. The injectors have been studied in the range of ΔP. The flow regimes have been identified, which can be predicted by the SPI and HEM models. It has been shown that the two-phase flow quality upstream of the injector may impact the discharge coefficient in the SPI region and the accuracy of the HEM model. It has been found that the transition to the critical flow region depends on the L/D ratio of the injector orifice. A series of conclusions can be drawn from this work to design the rocket injector with a self-pressurizing propellant to better predict the mass flow rate and ensure stable combustion.
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Abdelraouf, A. M., O. K. Mahmoud, and M. A. Al-Sanabawy. "Thrust termination of solid rocket motor." Journal of Physics: Conference Series 2299, no. 1 (July 1, 2022): 012018. http://dx.doi.org/10.1088/1742-6596/2299/1/012018.
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Abstract Rocket motors are engines that create the necessary thrust for the rocket motion. There are different types of rocket motors based on the propellant state, such as solid propellant rocket motors, liquid propellant rocket motors, and hybrid propellant rocket motors. One of the biggest disadvantages of solid propellant rocket motors, in comparison to liquid and hybrid propellant rocket motors, is that they are extremely difficult to extinguish, necessitating the use of specific devices. This paper reviews various ways for thrust termination such as fluid injection, rapid increase in throat area, and sudden opening of an additional port at the forward section of the motor, which increases the depressurization rate (dp/dt) required for extinguishing. The rate of depressurization varies depending on propellant components, combustion pressure, and exhaust pressure, and may be investigated using experimental approaches. The change in the critical area for a motor can be predicted by using MATLAB code to ensure the complete extinguishing by decreasing the pressure under the deflagrationlimit with high depressurization rate.
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Bandyopadhyay, Atri, and Ankit Kumar Mishra. "Comparative Study on Hybrid Rocket Fuels for Space Launch Vehicles Moving in Higher Orbits." 4 1, no. 4 (December 2, 2022): 13–19. http://dx.doi.org/10.46632/jame/1/4/3.
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The current research is focused on understanding the propulsive parameters of hybrid rocket motors. A comparative study is prepared from various research papers. The fuels paraffin wax, hydroxyl-terminated polybutadiene (HTPB) and polymethyl methacrylate (PMMA) with added additives (Al/Mg) were combined with two oxidizers, liquid oxygen (LOX), nitrous oxide (N2O). The propulsive parameters examined were the combustion efficiency, combustion or adiabatic flame temperature, characteristics velocity and regression rate. The propellant pair paraffin-N2O provided the highest performance for all parameters studied. This study provides an advantageous propellant option for future rocket propulsion based on a comparative investigation
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Sunil, Rahul, Aditya Virkar, M. Vignesh Kumar, Iynthezhuthon Krishnamoorthy, and Vinayak Malhotra. "Combustion and propulsive characteristics of potential hybrid rocket propellant." IOP Conference Series: Materials Science and Engineering 912 (September 12, 2020): 042023. http://dx.doi.org/10.1088/1757-899x/912/4/042023.
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Whitmore, Stephen A., Cara I. Frischkorn, and Spencer J. Petersen. "In-Situ Optical Measurements of Solid and Hybrid-Propellant Combustion Plumes." Aerospace 9, no. 2 (January 23, 2022): 57. http://dx.doi.org/10.3390/aerospace9020057.
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A method for in-situ optical measurements of solid and hybrid propellant rocket plumes is developed, and results from proof of concept tests are presented. The developed method inserts fiber-optic cables acting as radiation conduits into the solid-fuel combustion port, allowing optical signals to be transmitted from the flame zone to externally-mounted spectrometers. Multiple hot-firings using a using a lab-scale gaseous-oxygen, thermo-plastic fueled hybrid rocket system were performed to validate the sensing method. Burn durations varied from 5 to 25 s, and the inserted fiber optic sensors survived for all of the hot fire tests. The obtained optical spectra were curve-fit to Planck’s black-body radiation law, and Wien’s displacement law was used to estimate the internal flame-temperature. 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 hydrogen, water vapor, and molecular nitrogen; all species known to exist in the hybrid combustion plume. Based on these preliminary test results, it is concluded that this simple in-situ measurement system operates as designed, and it shows considerable promise for future applications to a wide swath of gas-generator systems.
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Urrego, Jose Alejandro, Fabio Arturo Rojas, and Jaime Roberto Muñoz. "Variability analysis of ABS solid fuel manufactured by fused deposition modeling for hybrid rocket motors." Journal of Mechanical Engineering and Sciences 15, no. 2 (June 10, 2021): 8029–41. http://dx.doi.org/10.15282/jmes.15.2.2021.08.0633.
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The process of fused deposition material (FDM) was used to manufacture propellant grains of Acrylonitrile Butadiene Styrene (ABS) as novel rocket fuel grain, with three types of geometry in the burning port. These solid fuel grains were used to measure the typical characteristics of combustion in rocket motors such as thrust and pressure inside the combustion chamber, seeking to obtain preliminary characteristics of operation and analyze the effect of combustion port geometry on pressure and thrust, using Multivariate Analysis of Variance (MANOVA) as statistical method. Two of the three geometries were manufactured with a helical-finocyl configuration, specially designed to be fabricated by Direct Digital Manufacturing (DDM), the other one was a straight-bore geometry also by DDM. This characterization experiment was performed on a static hybrid rocket engine, designed to inject 99.98% pure nitrous oxide into a combustion chamber with capacity to withstand 6.9 MPa of pressure, with an easy-to-exchange nozzle, avoiding erosive behavior in the throat. Statistical analyses made with the ABS fuel grains, suggest a significant effect on rocket motor pressure and thrust, due to helical geometric changes made to the combustion port of solid fuel grains made by FDM manufacture process.
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Tian, Hui, Xianzhu Jiang, Yudong Lu, Yu Liang, Hao Zhu, and Guobiao Cai. "Numerical Investigation on Hybrid Rocket Motors with Star-Segmented Rotation Grain." Aerospace 9, no. 10 (October 9, 2022): 585. http://dx.doi.org/10.3390/aerospace9100585.
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A novel fuel grain configuration comprising two star-segmented grains is proposed. The effect of the rotation, mid-chamber length, and segmented position on the fuel regression rate and the combustion efficiency in hybrid rocket motors with star-segmented grains is investigated in this paper. To this end, 90% hydrogen peroxide (H2O2) and hydroxyl-terminated polybutadiene (HTPB) are selected as the propellant combination in this research. Three-dimensional numerical simulations of the star-segmented grain configuration are conducted. A firing test of a lab-scale hybrid rocket motor was conducted to verify the accuracy of the numerical model, and the errors between simulation data and experimental results are no more than 4.5%. The case without segmented grain configuration is regarded as the base case. The simulation results demonstrate that the combustion flow field structure of the motor could be ameliorated by the segmented rotation grain configuration. Compared with the base case, the rotation of aft-section grain has little effect on the regression rate in the fore-section grain, while the average regression rate in aft-section grain increases, with a maximum increase of 25.04%. The combustion efficiency of the motor with the segmented rotation grain configuration is higher than the base case. Compared with the base case, the combustion efficiency of segmented rotation grain case with mid-chamber length 40 mm and segmented position of 1/2 is raised by 4.06%. The average fuel regression rate and the combustion efficiency of hybrid rocket motors with segmented rotation grains are higher than those in the base case during the entire period of operation, and the combustion efficiency is increased by 1.40–4.21% during the motor operation. The research findings of this paper can provide valuable guidance for the performance improvement of hybrid rocket motors with star grain.
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Lee, Dongeun, and Changjin Lee. "Fuel-rich Combustion with AP added Propellant in a Staged Hybrid Rocket Engine." Journal of the Korean Society for Aeronautical & Space Sciences 44, no. 7 (July 1, 2016): 576–84. http://dx.doi.org/10.5139/jksas.2016.44.7.576.
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D’Alessandro, Simone, Marco Pizzarelli, and Francesco Nasuti. "A Hybrid Real/Ideal Gas Mixture Computational Framework to Capture Wave Propagation in Liquid Rocket Combustion Chamber Conditions." Aerospace 8, no. 9 (September 4, 2021): 250. http://dx.doi.org/10.3390/aerospace8090250.
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The present work focuses on the development of new mathematical and numerical tools to deal with wave propagation problems in a realistic liquid rocket chamber environment. A simplified real fluid equation of state is here derived, starting from the literature. An approximate Riemann solver is then specifically derived for the selected conservation laws and primitive variables. Both the new equation of state and the new Riemann solver are embedded into an in-house one-dimensional CFD solver. The verification and validation of the new code against wave propagation problems are then performed, showing good behavior. Although such problems might be of interest for different applications, the present study is specifically oriented to the low order modeling of high-frequency combustion instability in liquid-propellant rocket engines.
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Khan, Suniya Sadullah, Ihtzaz Qamar, Muhammad Umer Sohail, Raees Fida Swati, Muhammad Azeem Ahmad, and Saad Riffat Qureshi. "Comparison of Optimization Techniques and Objective Functions Using Gas Generator and Staged Combustion LPRE Cycles." Applied Sciences 12, no. 20 (October 17, 2022): 10462. http://dx.doi.org/10.3390/app122010462.
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This paper compares various optimization techniques and objective functions to obtain optimum rocket engine performances. This research proposes a modular optimization framework that provides an optimum design for Gas Generator (GG) and Staged Combustion (SC) Liquid Propellant Rocket Engines. This process calculates the ideal rocket engine performance by applying seven different optimization techniques: Simulated Annealing (SA), Nelder Mead (NM), Cuckoo Search Algorithm (CSA), Particle Swarm Optimization (PSO), Pigeon-Inspired Optimization (PIO), Genetic Algorithm (GA) and a novel hybrid GA-PSO technique named GA-Swarm. This new technique combines the superior search capability of GA with the efficient constraint matching capability of PSO. This research also compares objective functions to determine the most suitable function for GG and SC cycle rocket engines. Three single objective functions are used to minimize the Gross Lift-Off Weight and to maximize Specific Impulse and the Thrust-to-Weight ratio. A fourth multiobjective function is used to simultaneously maximize both Specific Impulse and Thrust-to-Weight ratio. This framework is validated against a pump-fed rocket, and results are within 1% of the actual rocket engine mass. The results of this research indicate that PSO and GA-Swarm produce optimum results for all objective functions. Finally, the most suitable objective function to use while comparing these two cycles is the Gross Lift-Off Weight.
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Schulze, Moritz, and Thomas Sattelmayer. "Linear stability assessment of a cryogenic rocket engine." International Journal of Spray and Combustion Dynamics 9, no. 4 (May 4, 2017): 277–98. http://dx.doi.org/10.1177/1756827717695281.
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The linear high frequency stability of DLR’s cryogenic H2/O2 BKD test chamber is assessed using a hybrid computational fluid dynamic/computational aeroacoustic methodology, which is based on single flame simulations for the generation of an adequate mean flow and for the calibration of feedback models as well as on frequency space transformed linearized Euler equations. The application of a realistic mean flow field including combustion explains the spatial separation of transverse modes into a near face plate mode, which is found linearly unstable under certain operation conditions for the first transverse and a rear part mode. The axial mode shape length as well as eigenfrequencies is affected by propellant injection specifications and, in consequence, decisively influence pressure and transverse velocity sensitive dynamic flame response. The stability assessment procedure is finally applied to four operation conditions and the linear stability is predicted for the first transverse mode.
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Paravan, Christian, Luciano Galfetti, Riccardo Bisin, and Federico Piscaglia. "COMBUSTION PROCESSES IN HYBRID ROCKETS." International Journal of Energetic Materials and Chemical Propulsion 18, no. 3 (2019): 255–86. http://dx.doi.org/10.1615/intjenergeticmaterialschemprop.2019027834.
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Shynkarenko, Olexiy, Domenico Simone, Jungpyo Lee, and Artur E. M. Bertoldi. "Experimental and Numerical Study of the Flammability Limits in a CH4/O2 Torch Ignition System." Energies 15, no. 11 (May 24, 2022): 3857. http://dx.doi.org/10.3390/en15113857.
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The current work is devoted to studying combustion initiation inside the methane-oxygen torch igniter for a hybrid rocket motor. The ignition system can generate a wide range of power and oxidizer-to-fuel ratios. It has a self-cooled vortex combustion chamber with one fuel jet injector and one circumferential vortex oxidizer injector. The system adjusts the mass flow rates of the propellants through the control valves and organizes cooling of the wall and flame stabilization. Experimental analysis of the ignition limits was investigated on the laboratory test bench. The propellants’ pressure and mass-flow rates, combustion temperature, ignition delay, and spark frequency were controlled during the tests. The authors executed a series of tests with different propellants’ mass flow rates. As a result, the region of stable ignition was found as well as the regions of ignition failure or unreliable ignition. A previously validated numerical model was used to analyze the flow in the reliable ignition region and the ignition failures region. Several numerical simulations of the transient three-dimensional chemically reacting flow were implemented. Consequently, the ignition delay and the thermal impact on the combustion chamber wall were determined numerically. Results of the simulations were compared with theoretical and experimental data showing good correspondence.
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Runtu, Khevinadya Ramadhani, Wahyu Sri Setiani, and Mala Utami. "Application Energetic Materials for Solid Composite Propellant to Support Defense Rocket Development." International Journal of Social Science Research and Review 6, no. 1 (January 6, 2023): 153–59. http://dx.doi.org/10.47814/ijssrr.v6i1.756.
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In its application in space technology, solid composite propellants are often used as fuel in rockets for military purposes. Increasing the energy of the propellant is carried out by observing two stages, the use of energetic materials and improvements to the process technology. The current development of propellant technology makes it possible to use new energetic materials, simple formulations, high energy, and smokeless. The purpose of this research is to find out developments related to the use of highly energetic materials as raw materials for composite propellants for defense rockets at the Rocket Technology Research Center, ORPA-BRIN. This study uses qualitative analysis methods with research designs in the literature studies and simulation results. In the context of mastering rocket propellant technology in Indonesia, the application of highly energetic materials is expected to be able to solve the problem of rocket propulsion performance. Currently, the Rocket Technology Research Center, ORPA-BRIN is developing a smokeless propellant composite with a composition based on the energetic materials AP/HTPB/Al and an oxidizing agent RDX. From the results of the combustion simulation software ProPEP and RPA, it shows that the composition of the resulting combustion gaseous (Al2O3 and HCl) shows a decrease when using RDX energetic material-based propellant. It's known that RDX can significantly reduce smoke in propellant combustion products. The application of the new highly energetic materials compound is expected to significantly solve the problem of solid rocket propulsion performance.
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ZAGANESCU, Nicolae-Florin, Rodica ZAGANESCU, and Constantin-Marcian GHEORGHE. "Wernher Von Braun’s Pioneering Work in Modelling and Testing Liquid-Propellant Rockets." INCAS BULLETIN 14, no. 2 (June 10, 2022): 153–61. http://dx.doi.org/10.13111/2066-8201.2022.14.2.13.
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This paper presents a view on how Dr. Wernher Von Braun laid the basis for realistic modelling and testing liquid-propellants rockets, by his PhD Thesis – a secret document in 1934, which remained classified until 1960. Understanding that better mathematical modelling is needed if these rockets are to become spaceflight vehicles, he clarified in his thesis essential issues like: maximum achievable rocket speed; Laval nozzle thrust gain; polytropic processes in the combustion chamber and nozzle; influence of equilibrium and dissociation reactions; original measurement systems for rockets test stand; engineering solutions adequate for series production of the combustion chamber – reactive nozzle assembly. The thesis provided a theoretical and experimental basis for a new concept of the rocket, having a lightweight structure; low tanks pressure; high-pressure pumps and injectors; low start speed; rocket stabilization by gyroscopic means or by active jet controls; longer engine burning time; higher jet speed. Numerous tests made even with a fully assembled rocket (the “Aggregate-I”), improved mathematical model accuracy (e.g., the maximum achievable altitude predicted for the “Aggregate-II” rocket was confirmed later in-flight tests).
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GARCÍA-SCHÄFER, J. E., and A. LIÑÁN. "Longitudinal acoustic instabilities in slender solid propellant rockets: linear analysis." Journal of Fluid Mechanics 437 (June 22, 2001): 229–54. http://dx.doi.org/10.1017/s0022112001004323.
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To describe the acoustic instabilities in the combustion chambers of laterally burning solid propellant rockets the interaction of the mean flow with the acoustic waves is analysed, using multiple scale techniques, for realistic cases in which the combustion chamber is slender and the nozzle area is small compared with the cross-sectional area of the chamber. Associated with the longitudinal acoustic oscillations we find vorticity and entropy waves, with a wavelength typically small compared with the radius of the chamber, penetrating deeply into the chamber. We obtain a set of differential equations to calculate the radial and axial dependence of the amplitude of these waves. The boundary conditions are provided by the acoustic admittance of the propellant surface, given by an existing analysis of the thin gas-phase reaction layer adjacent to the solid–gas interface, and of the nozzle, accounting here for the possible effect of the vorticity and entropy waves. The equations are integrated in closed form and the results provide the growth rate of the disturbances, which we use to determine the conditions for instability of the longitudinal oscillations.
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Jadhav, Shruti Dipak, Tapas Kumar Nag, Atri Bandyopadhyay, and Raghvendra Pratap Singh. "Experimental and Computational Investigation of Sounding Solid Rocket Motor." 3 1, no. 3 (December 1, 2022): 29–38. http://dx.doi.org/10.46632/jame/1/3/5.
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Experimental sounding rockets are important contributors to aerospace engineering research. However, experimental-sounding rockets are rarely used for student research projects by institutes in India. The unavailability of rocket motors, which require complex machining and explosive propellants, is a major barrier to the use of sounding rockets in student research projects. We ran into this problem while developing a sounding rocket motor for project and learning purposes. The project focuses on designing and constructing a solid rocket motor that researchers can use as the primary propulsion unit in experimental sounding rockets. Initially, basic designs were evaluated, as various concepts of observations of propellant configuration. The accessibility and ease of use of manufacturing and casting of propellants played a significant role in determining the best propellant based on these findings, the theoretical values for combustion chamber parameters were obtained. Also, materials were chosen accordingly, and a fundamental small-scale experimental design was built and extensively tested. This small-scale motor was created by combining all of the analysis and theoretical data.At experimental testing, we got to know the thrust generated is 763.47N and the motor runs for 4.1 sec, the total mass of the propellant is maxed at 1500g which gives us the max mass flow rate of 0.65Kg/sec this is the output for our solid rocket motor.
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Kurdyumov, Vadim N. "Steady Flows in the Slender, Noncircular, Combustion Chambers of Solid Propellant Rockets." AIAA Journal 44, no. 12 (December 2006): 2979–86. http://dx.doi.org/10.2514/1.21125.
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Eisen, Nachum E., and Alon Gany. "Examining Metal Additives in a Marine Hybrid-Propellant, Water-Breathing Ramjet." Journal of Marine Science and Engineering 10, no. 2 (January 20, 2022): 134. http://dx.doi.org/10.3390/jmse10020134.
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This research focuses on theoretically and experimentally evaluating the performance of a metallized hybrid-propellant, water-breathing ramjet. The aluminum and/or magnesium particles added to a polymeric (polyester) fuel grain are hydro-reactive, using the surrounding water as an oxidizer, in addition to a source of gas. Theoretically, the metal additives significantly increase the specific impulse of the motor, and as the percentage of the hydro-reactive ingredient increases, the theoretical performance increases as well. Additionally, aluminum is more energetic than magnesium. However, it was experimentally discovered that the addition of aluminum beyond 20% resulted in a slag formation and did not increase the specific impulse. Adding 30% of magnesium was relatively favorable to aluminum due to its better reactivity, enabling the achievement of an actual specific impulse of up to 485 s at standard conditions, approximately double the performance of common solid rockets.
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KOSTYUSHIN, Kirill V. "NUMERICAL INVESTIGATION OF UNSTEADY GASDYNAMIC PROCESSES AT THE LAUNCH OF SOLID-PROPELLANT ROCKETS." Vestnik Tomskogo gosudarstvennogo universiteta. Matematika i mekhanika, no. 67 (2020): 127–43. http://dx.doi.org/10.17223/19988621/67/12.
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The paper presents the results of the methodology developed for calculating unsteady gasdynamic processes occurring at the launch of missiles, in the gas-dynamic paths of rocket engines, and in the external regions. The method accounts for the variation in the geometry of the solidpropellant charge in the course of solid-propellant rocket engine operation and in the geometry of the computational domain at the rocket launch. The analysis of the unsteady force impact of the supersonic jet on the launch surface is carried out. It is shown that the maximum force action is located in the vicinity of the Mach disks of the unperturbed jet. Numerical studies of gasdynamic processes at the launch of a model solid-propellant booster rocket are implemented including the case when the nozzle plug opening is taken into account. The contribution of the thrust force components at the stage of bootstrap operation is assessed. The presence of the plug at the initial stage of the engine start leads to an abrupt change in the thrust and minor fluctuations, which are damped as the pressure in the combustion chamber rises.
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Santos, L. M. C., L. A. R. Almeida, A. M. Fraga, and C. A. G. Veras. "EXPERIMENTAL INVESTIGATION OF A PARAFFIN BASED HYBRID ROCKET." Revista de Engenharia Térmica 5, no. 1 (July 31, 2006): 08. http://dx.doi.org/10.5380/reterm.v5i1.61658.
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Hybrid rockets are known to be simpler, safer, environmentally friend, and, more importantly, cheaper than most of the technologies for propulsion devices used today. Hybrid rockets can be applied as the propulsion system in satellites launch vehicles, micro-satellites and tactical missiles. This paper deals with combustion of ultra-high molecular weight polyethylene (UHMWPE) and paraffin as the solid fuels burning with gaseous oxygen (GOX) as well as N O as the oxidizer in lab scale hybrid rocket motors. A test 2 stand was built to carry out the experiments. The main objectives were to investigate the ignition of the solid fuels, burning performance and regression rates for different operating conditions. With paraffin-based fuel the hybrid motor had the regression rate enhanced two to three folds compared to the UHMWPE, as reported in the literature. The overall performance of the motor, with paraffin as the fuel, is comparable to other technologies. Paraffin-based hybrid rockets can, then, be a safer and cheaper alternative to satellite launch vehicles for the Brazilian space program.
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Matsuoka, Tsuneyoshi, and Harunori Nagata. "Combustion characteristics of the end burning hybrid rockets in laminar flow." Acta Astronautica 68, no. 1-2 (January 2011): 197–203. http://dx.doi.org/10.1016/j.actaastro.2010.07.009.
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Carmicino, Carmine. "Special Issue “Advances in Hybrid Rocket Technology and Related Analysis Methodologies”." Aerospace 6, no. 12 (November 26, 2019): 128. http://dx.doi.org/10.3390/aerospace6120128.
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Hybrid rockets are chemical propulsion systems that, in the most common configuration, employ a liquid oxidizer (or gaseous in much rarer cases) and a solid fuel; the oxidizer, stored in tanks, is properly injected in the combustion chamber where the solid fuel grain is bonded [...]
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Apel, Uwe, Alexander Baumann, Christian Dierken, and Thilo Kunath. "AQUASONIC – A Sounding Rocket Based on Hybrid Propulsion." Applied Mechanics and Materials 831 (April 2016): 3–13. http://dx.doi.org/10.4028/www.scientific.net/amm.831.3.
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The AQUASONIC project is aimed to develop a sounding rocket including a hybrid propulsion system based on the propellant combination nitrous oxide and polyethylene. It takes place in the frame of the STERN (Student Experimental Rockets) programme founded by the German Space Agency (DLR) in order to promote students in the area of launch vehicles. Main element of the project is the AQUASONIC rocket, which shall reach a flight altitude of 5-6 km and a velocity of MACH 1. All major activities like design, manufacturing, verification and, finally, the launch campaign will be performed by students. The rocket shall be launched at Esrange Space Centre (Sweden) in 2016. Thus, students are able to apply their skills and knowledge to a real project like it is conducted by the space industry or research organisations.
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Barato, Francesco, Elena Toson, and Daniele Pavarin. "Variations and Control of Thrust and Mixture Ratio in Hybrid Rocket Motors." Advances in Astronautics Science and Technology 4, no. 1 (April 18, 2021): 55–76. http://dx.doi.org/10.1007/s42423-021-00076-3.
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AbstractHybrid rocket motors have several attracting characteristics such as simplicity, low cost, safety, reliability, environmental friendliness. In particular, hybrid rockets can provide complex and flexible thrust profiles not possible with solid rockets in a simpler way than liquid rockets, controlling only a single fluid. Unfortunately, the drawback of this feature is that the mixture ratio cannot be directly controlled but depends on the specific regression rate law. Therefore, in the general case the mixture ratio changes with time and with throttling. Thrust could also change with time for a fixed oxidizer flow. Moreover, propellant residuals are generated by the mixture ratio shift if the throttling profile is not known in advance. The penalties incurred could be more or less significant depending on the mission profile and requirements. In this paper, some proposed ways to mitigate or eliminate these issues are recalled, quantitatively analysed and compared with the standard case. In particular, the addition of energetic additives to influence the regression rate law, the injection of oxidizer in the post-chamber and the altering-intensity swirling-oxidizer-flow injection are discussed. The first option exploits the pressure dependency of the fuel regression to mitigate the shift during throttling. The other two techniques can control both the mixture ratio and thrust, at least in a certain range, at the expense of an increase of the architecture complexity. Moreover, some other options like pulse width modulation or multi-chamber configuration are also presented. Finally, a review of the techniques to achieve high throttling ratios keeping motor stability and efficiency is also discussed.
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Murachman, Bardi, Sajono Sajono, Fauzan Afandi, and Johan Khaeri. "Optimization Study of the Solid Propellant (Rocket Fuel) Based on Extracted Bitumen of Indonesian Natural Buton Asphalt." ASEAN Journal of Chemical Engineering 13, no. 2 (September 17, 2014): 57. http://dx.doi.org/10.22146/ajche.49732.
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The asphalt propellant for rockets has been investigated since 1960. This material has been developed with the variation of fuels, oxidizer, binders, metal elements and additives. As solid propellant, it has some advantages and disadvantages during the implementation. At present, Extracted Buton asphalt has been studied as an alternative propellant fuels. It is a natural asphalt, extracted from Buton island asphalt rock. When the extract of buton asphalt is mixed with oxidizer, binder, and metal powder, it can be functioned as propellant which is able to release high intensity of energy, have strong thrust and power to fly the rocket. This optimization study of solid propellant was conducted by mixing the Buton asphalt as fuel, oxidizer, metal element and other additives to form a solid propellant. The oxidizer consisted of potassium nitrate (KNO3) and potassium perchlorate (KClO4). The variations of KClO4/KNO3, propellant density and the ratio of the nozzle diameter were also conducted in order to find the best propellant composition and the optimum operating conditions to produce enough power while maintain the integrity of the rocket. The main parameters such as the propellant’s thrust (F) and the specific impulse (Isp) were examined. The results showed that higher composition of KClO4/KNO3 gave the higher value of the thrust and the specific impulse. KClO4/KNO3 levels above the 1:1 ratio produced an explosive properties at the time of ignition. The tendency of propellant to explode during ignition process was also observed. The optimum condition was obtained at the KClO4/KNO3 ratio of 1:1 , the propellant density was 1.900 gram/cm3 and Ae/A* was 3.33. These conditions generated impulse value that last for 277.07 seconds, average thrust of 14.082 N, and average rate of combustion of 0,24 cm/second. Therefore, it can be concluded that propellant with fuel from extracted of Buton asphalt can be used as an alternative propellant for rocket.
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Mahottamananda, Sri Nithya, Yash Pal, Mengu Dinesh, and Antonella Ingenito. "Beeswax–EVA/Activated-Charcoal-Based Fuels for Hybrid Rockets: Thermal and Ballistic Evaluation." Energies 15, no. 20 (October 14, 2022): 7578. http://dx.doi.org/10.3390/en15207578.
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Beeswax (C46H92O) is a naturally derived substance that has the potential to be used as a solid fuel for hybrid rocket applications and as a substitute for paraffin wax fuel in hybrid rockets. BW burns more efficiently than paraffin wax because of the oxygen molecule it contains. The low thermal stability and poor mechanical properties of BW limit its practical use for upper-stage propulsion applications, and these issues are rarely addressed in the literature on hybrid rockets. This study investigates the thermal stability and ballistic properties of BW using ethylene-vinyl acetate (EVA) and activated charcoal (AC) as an additive. The thermal stability of BW–EVA/AC fuel compositions was analyzed using a thermogravimetric analyzer (TGA). The thermal stability of the blended BW compositions improved significantly. A laboratory-scale hybrid rocket motor was used to evaluate such aspects of ballistic performance as regression rate, characteristic velocity, and combustion efficiency. The results revealed that the pure BW exhibited a higher regression rate of 26.5% at an oxidizer mass flux of 96.4 kg/m2-s compared to BW–EVA/AC blends. The addition of EVA and AC to BW was found to increase the experimental characteristic velocity and combustion efficiency. The combustion efficiency of BW-based fuel was improved from 62% to 94% when 20 wt.% EVA and 2 wt.% AC were added into the fuel matrix.
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Freesmeier, Jon J., and P. Barry Butler. "Analysis of a Hybrid Dual-Combustion-Chamber Solid-Propellant Gas Generator." Journal of Propulsion and Power 15, no. 4 (July 1999): 552–61. http://dx.doi.org/10.2514/2.5478.
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Marquardt, Timothy, and Joseph Majdalani. "Review of Classical Diffusion-Limited Regression Rate Models in Hybrid Rockets." Aerospace 6, no. 6 (June 20, 2019): 75. http://dx.doi.org/10.3390/aerospace6060075.
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In this article, we undertake a concise review of several milestone developments in classical regression rate models of hybrid rocket motors. After a brief description of the physical processes entailed in hybrid rocket combustion, Marxman’s diffusion-limited theory is re-constructed and discussed. Considerations beyond the scope of basic convection-driven models, which address disparate forms of the blowing correction, variable fluid properties, and pressure and radiation effects, are also given. Finally, a selection of kinetically-limited models is presented, with the aim of comparing the characteristics of several competing theories that become applicable under particular circumstances.
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Kim, Youngin, and Jeongho Cho. "Surface Erosion Analysis for Thermal Insulation Materials of Graphite and Carbon–Carbon Composite." Applied Sciences 9, no. 16 (August 13, 2019): 3323. http://dx.doi.org/10.3390/app9163323.
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A rocket uses fuel and oxidizers to generate propulsion by combustion and ejection, and is used for space exploration aircrafts, weapons, and satellite launches. In particular, the nozzle generating thrust of solid-propellant rockets is exposed to a high-temperature and high-pressure environment with erosion occurring from the combustion gas. When erosion occurs on the nozzle throat of such a rocket, it has a great impact on the flight performance such as reaching distance and flight speed. Many studies have been conducted to characterize erosion based on the thermochemical erosion model, since it has become important to choose nozzle materials suitable for such environments having robustness against combustion gasses of high temperature and high pressure. However, there is a limit to fully analyze the erosion characteristics only by the thermochemical erosion model. In this paper, we thus consider the mechanical erosion model with the thermochemical model for better understanding of erosion characteristics and investigate the thermochemical and mechanical erosion characteristics of nozzle throat heat-resistant materials made of graphite and carbon–carbon composites; the main factors affecting erosion are discussed by comparing the results of the experimental and theoretical models.
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Gańczyk-Specjalska, Katarzyna, Paulina Paziewska, Rafał Bogusz, Rafał Lewczuk, Katarzyna Cieślak, and Michał Uszyński. "Performance and Sensitivity Properties of Solid Heterogeneous Rocket Propellant Based on a Binary System of Oxidizers (PSAN and AP)." Processes 9, no. 12 (December 7, 2021): 2201. http://dx.doi.org/10.3390/pr9122201.
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Solid heterogeneous rocket propellants (SHRP) containing ammonium perchlorate (AP) emit a lot of hydrogen chloride (HCl) during combustion, which poses various environmental issues and makes the detection of the rockets easier. Part of the AP can be replaced by ammonium nitrate (V) (AN), which does not lead to the production of HCl. AN is a commonly used environmentally friendly oxidizer, but it is not usually applied in SHRP due to its disadvantages. One of these disadvantages is a phase transition near room temperature, which causes the density change of AN. Three types of phase stabilized ammonium nitrate (V) (PSAN) with inorganic potassium salts were obtained in order to shift this transition into higher temperatures (above the temperature range of the storage and the usage of SHRP). The SHRP with the PSAN were obtained, and the measurements of the heat of combustion, density, hardness, the sensitivity to mechanical stimuli and the thermomechanical properties were performed. The obtained propellants were characterized by similar operational parameters or were slightly lower than those without the PSAN. This means that AP can be partially replaced without significantly compromising the handling, safety or functionality of the propellants, while increasing the environmental performance of the solution.
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Wada, Yutaka, Yoshio Seike, Makihito Nishioka, Nobuyuki Tsuboi, Toru Shimada, Katsuya Hasegawa, Kiyokazu Kobayashi, and Keiichi Hori. "COMBUSTION MECHANISM OF TETRA-OL GLYCIDYL AZIDE POLYMER AND ITS APPLICATION TO HYBRID ROCKETS." International Journal of Energetic Materials and Chemical Propulsion 8, no. 6 (2009): 555–70. http://dx.doi.org/10.1615/intjenergeticmaterialschemprop.v8.i6.70.
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Barato, Francesco. "Challenges of Ablatively Cooled Hybrid Rockets for Satellites or Upper Stages." Aerospace 8, no. 7 (July 14, 2021): 190. http://dx.doi.org/10.3390/aerospace8070190.
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Ablative-cooled hybrid rockets could potentially combine a similar versatility of a liquid propulsion system with a much simplified architecture. These characteristics make this kind of propulsion attractive, among others, for applications such as satellites and upper stages. In this paper, the use of hybrid rockets for those situations is reviewed. It is shown that, for a competitive implementation, several challenges need to be addressed, which are not the general ones often discussed in the hybrid literature. In particular, the optimal thrust to burning time ratio, which is often relatively low in liquid engines, has a deep impact on the grain geometry, that, in turn, must comply some constrains. The regression rate sometime needs to be tailored in order to avoid unreasonable grain shapes, with the consequence that the dimensional trends start to follow some sort of counter-intuitive behavior. The length to diameter ratio of the hybrid combustion chamber imposes some packaging issues in order to compact the whole propulsion system. Finally, the heat soak-back during long off phases between multiple burns could compromise the integrity of the case and of the solid fuel. Therefore, if the advantages of hybrid propulsion are to be exploited, the aspects mentioned in this paper shall be carefully considered and properly faced.
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BANNO, Ayana, Yo KAWABATA, Yutaka WADA, Nobuji KATO, and Keiichi HORI. "Temperature Field Measurement of Low Melting Point Fuel in Boundary Layer Combustion Type Hybrid Rockets." Proceedings of Mechanical Engineering Congress, Japan 2017 (2017): S1920107. http://dx.doi.org/10.1299/jsmemecj.2017.s1920107.
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Rofiq, Muhammad, Lalu Saefullah, and Mohammad Ali. "The DESIGN OF A TRANSMISSION SYSTEM ON A BALL MILL MACHINE TO PROCESS HOMOGENIZATION OF BLACK POWDER USING AN AUTOMATIC SYSTEM." Jurnal Otoranpur 2, Oktober (October 27, 2021): 63–70. http://dx.doi.org/10.54317/oto.v2ioktober.195.
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Abstract: In the military world, battle technology has experienced very rapid development, both close-range combat and long-range combat. In long-distance combat the technology that is often used is rocket technology where this rocket technology can paralyze the enemy in large numbers. Black powder is a mixture of potassium nitrate (KNO3), carbon powder (C) and sulfur (S). (Evie Lestariana, LAPAN). The black powder itself functions as an igniter filling material for the propellant combustion process. This research is to design a tool to support the manufacture of black powder in the igniter filling to help the combustion of rockets, especially those owned by the Indonesian Army. This study uses an experimental method using empirical calculations to get a tool with the desired specifications. Before carrying out the calculations, the author collects research data to carry out calculations, the data collected are as follows: 1. Roller shaft rotation: 246 rpm, 2. Roller shaft torque: 4.0697 Nm, 3. Roller shaft diameter: 25 mm, 4. Pulley shaft diameter top: 100mm. The results obtained after carrying out the calculations are the diameter of the driving pulley 98 mm, the belt contact angle against the pulley 174.6º, and the length of the V-belt 453.14 mm, moment of inertia on the gearbox output shaft 1884,8 mm4.
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Kearney, Sean P., and Daniel R. Guildenbecher. "Temperature measurements in metalized propellant combustion using hybrid fs/ps coherent anti-Stokes Raman scattering." Applied Optics 55, no. 18 (June 20, 2016): 4958. http://dx.doi.org/10.1364/ao.55.004958.
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JANISZEWSKI, Jacek, and Ryszard WOŹNIAK. "HYPERSONIC PROPULSION SYSTEMS – A REVIEW OF DESIGN SOLUTIONS." PROBLEMY TECHNIKI UZBROJENIA 161, no. 3 (November 29, 2022): 7–35. http://dx.doi.org/10.5604/01.3001.0016.1105.
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Abstract: The development of space rocket technology in the mid-twentieth century intensified research in the field of engineering materials used both for rockets and protection of space vehicles against damage, e.g., by meteorites. This development, in turn, forced the development of laboratory propulsion systems with hypersonic velocities, enabling the study and modelling of phenomena occurring at high impact velocities. Currently, to accelerate projectiles to hypersonic velocities, launching systems are applied that use the energy of the explosion of explosives or the combustion of liquid fuels, plasma, the energy of the electromagnetic field or light gases expanding. The work presents a review of the design solutions of various accelerators that enable the projectile to reach muzzle velocities above 3000 m/s. Particular attention has been paid to two-stage gas system, which uses light gases such as hydrogen and helium. The paper also presents the history and design of the first, and so far the only, Polish two-stage light gas gun developed at the Military University of Technology, with the help of which in 1973 a projectile weighing 0.5 g was fired at a speed of 4500 m/s. The paper ends with a description of the two-stage propellant system currently under construction at the Institute of Armament Technology, Faculty of Mechatronics, Armaments and Aviation, the Military University of Technology - with the considerable help of the HSW S.A. company from Stalowa Wola.
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Timoshenko, V. I., V. P. Halynskyi, and Yu V. Knyshenko. "Theoretical studies on rocket/space hardware aerogas dynamics." Technical mechanics 2021, no. 2 (June 29, 2021): 46–59. http://dx.doi.org/10.15407/itm2021.02.046.
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This paper presents the results of theoretical studies on rocket/space hardware aerogas dynamics obtained from 2016 to 2020 at the Department of Aerogas Dynamics and Technical Systems Dynamics of the Institute of Technical Mechanics of the National Academy of Sciences of Ukraine and the State Space Agency of Ukraine along the following lines: rocket aerodynamics, mathematical simulation of the aerogas thermodynamics of a supersonic ramjet vehicle, jet flows, and the hydraulic gas dynamics of low-thrust control jet engines. As to rocket aerodynamics, computational methods and programs (CMPs) were developed to calculate supersonic flow past finned rockets. The chief advantage of the CMPs developed is computational promptness and ease of adding wings and control and stabilization elements to rocket configurations. A mathematical simulation of the aerogas thermodynamics of a supersonic ramjet vehicle yielded new results, which made it possible to develop a prompt technique for a comprehensive calculation of ramjet duct flows and generalize it to 3D flow past a ramjet vehicle. Based on marching methods, CMPs were developed to simulate ramjet duct flows with account for flow past the airframe upstream of the air inlet, the effect of the combustion product jet on the airframe tail part, and its interaction with a disturbed incident flow. The CMPs developed were recommended for use at the preliminary stage of ramjet component shape selection. For jet flows, CMPs were developed for the marching calculation of turbulent jets of rocket engine combustion products with water injection into the jet body. This made it possible to elucidate the basic mechanisms of the effect of water injection, jet–air mixing, and high-temperature rocket engine jet afterburning in atmospheric oxygen on the flow pattern and the thermogas dynamic and thermalphysic jet parameters. CMPs were developed to simulate the operation of liquid-propellant low-thrust engine systems. They were used in supporting the development and ground firing tryout of Yuzhnoye State Design Office’s radically new system of control jet engines fed from the sustainer engine pipelines of the Cyclone-4M launch vehicle upper stage. The computed results made it possible to increase the informativity of firing test data in flight simulation. The CMPs developed were transferred to Yuzhnoye State Design Office for use in design calculations.
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Kuo, Kenneth K., Grant A. Risha, Brian J. Evans, and Eric Boyer. "Potential Usage of Energetic Nano-sized Powders for Combustion and Rocket Propulsion." MRS Proceedings 800 (2003). http://dx.doi.org/10.1557/proc-800-aa1.1.
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ABSTRACTNano-sized energetic metals and boron particles (with dimensions less than 100 nanometers) possess desirable combustion characteristics such as high heats of combustion and fast energy release rates. Because of their capability to enhance performance, various metals have been introduced in solid propellant formulations, gel propellants, and solid fuels. There are many advantages of incorporating nano-sized materials into fuels and propellants, such as: 1) shortened ignition delay; 2) shortened burn times, resulting in more complete combustion in volume-limited propulsion systems; 3) enhanced heat-transfer rates from higher specific surface area; 4) greater flexibility in designing new energetic fuel/propellants with desirable physical properties; 5) nano-particles can act as a gelling agent to replace inert or low-energy gellants; 6) nano-sized particles can also be dispersed into high-temperature zone for direct oxidation reaction and rapid energy release, and 7) enhanced propulsive performance with increased density impulse. In view of these advantages, numerous techniques have been developed for synthesizing nano-particles of different sizes and shapes. To reduce any possible hazards associated with the handling of nano-sized particles as well as unwanted particle oxidation, various passivation procedures have been developed. Some of these coating materials could enhance the ignition and combustion behavior, others could increase the compatibility of the particles with the surrounding material. Many researchers have been actively engaged in the characterization of the ignition and combustion behavior of nano-sized particles as well as the assessment of performance enhancement of propellants and fuels containing energetic nano-particles. For example, solid fuels could contain a significant percentage of nano-sized particles to increase the mass-burning rate in hybrid rocket motors, the regression rate of solid propellants can be increased by several times when nano-sized particles are incorporated into the formulation. Specifically, hybrid motor data showed that the addition of 13% energetic aluminum powders can increase the linear regression rate of solid HTPB-based fuel by 123% in comparison to the non-aluminized HTPB fuel at a moderate gaseous oxidizer mass flow rate. Strand burner studies of two identical solid propellant formulations (one with 18% regular aluminum powder and the other with 9% aluminum replaced by Alex® powder) showed that nano-sized particles can increase the linear burning rate of solid propellants by 100%. In addition to solid fuels and propellants, spray combustion of bipropellants has been conducted using gel propellants impregnated with nano-sized boron particles as the fuel in a rocket engine. High combustion efficiencies were obtained from burning nano-sized boron particles contained in a non-toxic liquid-fuel spray. Materials characterization such as chemical analyses to determine the active aluminum content, density measurements, and imaging using an electron microscope have been performed on both neat nano-sized particles and mixtures containing the energetic materials. In general, using energetic nano-sized particles as a new design parameter, propulsion performance of future propellants and fuels can be greatly enhanced.
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Pal, Yash, Anthony Raja, and Kavitha Gopalakrishnan. "Theoretical and Experimental Heat of Combustion Analysis of Paraffin-Based Fuels as Preburn Characterization for Hybrid Rocket." Journal of Aerospace Technology and Management, no. 12 (October 9, 2020). http://dx.doi.org/10.5028/jatm.v12.1180.
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The energy characteristics and theoretical performance of the hybrid rocket fuels are discussed in this paper. Aluminum (Al) and boron (B) metal additives were used to increase the energy density of the paraffin-based solid fuels. To predict the energy characteristics, the heat of combustion was evaluated by adiabatic bomb calorimetry. Theoretical performance parameters such as specific impulse (Isp), flame temperature, and characteristic velocity were obtained with NASA Chemical Equilibrium with Applications (CEA) code. Calorimetric test results revealed that paraffin/polyethylene/boron (P/PE/B)-based fuel formulations exhibited the highest heat of combustion among all the tested fuels. The heat of combustion value of the P/PE/B sample at 25 wt% B loading was found to be 9612 ±16 cal/g and 9293±17 cal/g for the P/PE/Al fuel formulation. The CEA results showed that the addition of Al to paraffin is noneffective in improving specific impulse performance. When B loading increased from 5 to 25 wt% in the P/PE/B, the Isp increased by 47 s compared to pure paraffin. A specific impulse increase implies the possible propellant mass saving. The reduction of the oxidizer and fuel masses may yield increased payload performance for given boundary conditions. The P/PE/B25 formulation has reported the highest value of characteristics velocity (C*) compared to other paraffin-based formulation.
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Yasuda, Kazuki, Daisuke Nakata, Masaharu Uchiumi, Kugo Okada, and Ryoji Imai. "Fundamental Study on Injector Flow Characteristics of Self-Pressurizing Fluid for Small Rocket Engines." Journal of Fluids Engineering 143, no. 2 (November 4, 2020). http://dx.doi.org/10.1115/1.4048688.
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Abstract Nitrous oxide is a suitable propellant for rocket engines and has been widely used in various countries, given its high saturated vapor pressure, which enables self-pressurization. Because nitrous oxide is in a state of vapor–liquid equilibrium in tanks, it is easy to form a gas–liquid two-phase flow by cavitation in feed line. Since accurately estimating the performance of rocket engines requires evaluating the characteristics of propellant flows, tests reported in this paper were conducted using hybrid rocket engines under three conditions: cold flow test, hot firing test at low back pressure, and hot firing test at high back pressure. With consideration to the subcooling degrees, nitrous oxide may be in an unsteady superheated state in the upstream flow of the injector. In a comparison of the pressure ratios between the injector in each test condition, it is observed that a critical two-phase flow was formed in the injector in the cold flow test and in the low backpressure firing test. In the high backpressure hot firing test, the injector flow may be choked, but the large oscillations were observed in chamber pressure and thrust. According to the FFT analysis results, these oscillations were caused by chugging and acoustic oscillation. In light of these experimental results, it is suggested that when the chamber pressure fluctuates due to combustion instability such as chugging and acoustic oscillation, it may affect the injector flow characteristics and the critical two-phase flow.
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Urrego, J. Alejandro P., Fabio A. M. Rojas, and Jaime R. L. Muñoz. "Combustion Performance Comparison of Propellant Grain for Hybrid Rocket Motors Manufactured by Casting and Fused Deposition Modeling." International Journal of Mechanical Engineering and Robotics Research, 2019, 960–65. http://dx.doi.org/10.18178/ijmerr.8.6.960-965.
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Eisen, Nachum E., and Alon Gany. "Novel Testing of a Water-Breathing Naval Ramjet at Underwater Cruise Conditions." Journal of Propulsion and Power, December 23, 2022, 1–7. http://dx.doi.org/10.2514/1.b38920.
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This paper presents, for the first time, firing tests and results measured during a hybrid-propellant marine water-breathing ramjet operation at underwater cruise conditions. The firing test demonstrated ignition under water, stable combustion, and a significant increase in chamber pressure due to the introduction of water by the ram effect. The exclusive test facility used to accommodate this unique experiment is described along with a detailed explanation of the motor design. In addition, a theoretical model to predict the performance of the motor is shown along with the numeric process used to solve the model. Measurements taken during the dynamic firing test reveal good compatibility with the theoretical model. The theoretical model was also used to evaluate values that could not be measured during the experiment. The results found during this work fit previous results attained by static firing tests, demonstrating significant improvement in specific impulse due to addition of water. An improvement of 50% in specific impulse was demonstrated when comparing the results of an underwater dynamic firing test of a marine ramjet to those of a parallel rocket test conducted under identical conditions in the water basin.
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Lee, Jungpyo, Heejang Moon, and Jinkon Kim. "Thermal–Combustion Coupled Instability in Hybrid Rockets with Fuel Blowing." Journal of Propulsion and Power, November 30, 2022, 1–10. http://dx.doi.org/10.2514/1.b38888.
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Thermal–combustion coupled instability creates one of the natural frequencies in hybrid rocket systems and is commonly observed as the dominant hybrid oscillation frequency. This type of instability is caused by the coupling between the thermal lag in the solid fuel and the combustion transients in the boundary layer, and it is unique to hybrid rocket systems. This study investigates the nature of thermal–combustion coupled instability, considering various test variables with a laboratory-scale hybrid rocket motor. The main parameters affecting the thermal–combustion coupled instability frequency are the fuel length and the freestream velocity in the fuel port, whereas the chamber pressure and the oxidizer mass flow rate have no significant direct effect. A new thermal–combustion coupled model is proposed that can account for changes in flow variables caused by the blowing by considering the blowing mass from the solid fuel surface. The new model’s thermal–combustion coupled instability frequencies are in good agreement with the experimental data over a wide range of test conditions.
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Risha, Grant A., Eric Boyer, Brian Evans, Kenneth K. Kuo, and Rafaat Malek. "Characterization of Nano-Sized Particles for Propulsion Applications." MRS Proceedings 800 (2003). http://dx.doi.org/10.1557/proc-800-aa6.6.
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ABSTRACTEnergetic nano-sized particles have been shown to have a great potential for use in the aerospace propulsion applications. Some of the unique combustion properties of nano-particles such as very rapid ignition and short combustion times make them particularly valuable for propulsion systems; they can be included in solid fuels, solid propellants, or even as energetic gellant in liquid systems. However, due to the novelty of the application and rapid development of production techniques, there is no comprehensive understanding of what characteristics of a nano-sized particle are important in contributing to desirable performance and ease of processing into a final usable form. Previous studies have shown that HTPB-based solid fuels containing various types of nano-sized particles showed differing performance results when tested in the same hybrid rocket motor under identical conditions. Many of these particles have data available only on the basic composition (aluminum, boron, boron carbide, etc.), average diameter, and/or BET surface area. In order to better understand and correlate observed combustion behavior with intrinsic material properties, the particles of interest need to be better characterized. A variety of standard particle characterization techniques were applied to the fifteen types of particles examined in this study and the results tabulated. Some of the parameters measured were average particle diameter, specific surface area, amount of active content, and oxide layer thickness. Trends in propulsion performance measured using a parameter of great interest to the hybrid rocket community (fuel mass burning rate) in general matched trends in particle characteristics (i.e. active content, surface area), but there were some noticeable exceptions. This study indicates that there is still much more to learn about the correlation between physical and chemical properties and measured combustion performance. Other parameters that should be examined in the future include particle size distribution, degree of agglomeration, reactivity and thermal effects (oxidation rate, onset temperature for oxidation exotherm, heat release associated with any excess stored energy), etc.
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Borkowska, Małgorzata, Jan Mężyński, Maciej Moskalewicz, Tomasz Rasztabiga, and Marek Tulik. "Thermal insulating cover for the metal body of a rocket motor." Materiały Wysokoenergetyczne / High Energy Materials, December 6, 2019, 31–38. http://dx.doi.org/10.22211/matwys/0091e.
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Many years of experience in the design and manufacturing technology of ammunition and rockets by BUMAR AMUNICJA S.A. enabled undertaking, together with the Institute IMPiB Department Elastomers and Rubber Technology in Piastów, the task of developing an upgraded insulation cover for the metal body of the rocket motor of the GROM-M system. The insulating layer of the metal body of the motor is characterised by the fact that local thermal exposure occurring during rocket propellant combustion, often causes the motor body to burn through locally. Applying an insulating composition matched to the temperature of combustion of the rocket motor, brings about a carbonisation process applicable to this temperature. Thus, it provides a sufficient layer of carbon-carbon phase composite at the site to enhance the strength of the motor and protect the above-mentioned metal body from exposure to heat during combustion of the rocket propellant. The thermal insulating layer is a double layered coating consisting of a layer of rubber compound and a layer of impregnated carbon cloth. To obtain uniformity of the insulating layer and hence its heat resistance, similar components are used in the composition of the rubber compound for impregnating a composite based on carbon cloth: the same type of rubber, the same phenol-formaldehyde resin and the same vulcanising agents.
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Rüttgers, A., and A. Petrarolo. "Local anomaly detection in hybrid rocket combustion tests." Experiments in Fluids 62, no. 7 (June 12, 2021). http://dx.doi.org/10.1007/s00348-021-03236-1.
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AbstractLocal anomaly detection was applied to image data of hybrid rocket combustion tests for a better understanding of the complex flow phenomena. Novel techniques such as hybrid rockets that allow for cost reductions of space transport vehicles are of high importance in space flight. However, the combustion process in hybrid rocket engines is still a matter of ongoing research and not fully understood yet. Since 2013, combustion tests with different paraffin-based fuels have been performed at the German Aerospace Center (DLR) and the whole process has been recorded with a high-speed video camera. This has led to a huge amount of images for each test that needs to be automatically analyzed. In order to catch specific flow phenomena appearing during the combustion, potential anomalies have been detected by local outlier factor (LOF), an algorithm for local outlier detection. The choice of this particular algorithm is justified by a comparison with other established anomaly detection algorithms. Furthermore, a detailed investigation of different distance measures and an investigation of the hyperparameter choice in the LOF algorithm have been performed. As a result, valuable insights into the main phenomena appearing during the combustion of liquefying hybrid rocket fuels are obtained. In particular, fuel droplets entrained into the oxidizer flow and burning over the flame are clearly identified as outliers with respect to the main combustion process. Graphic abstract