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Статті в журналах з теми "Hybrid propellant rockets – Testing"
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.
Повний текст джерелаJadhav, 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.
Повний текст джерелаCasalino, Lorenzo, and Dario Pastrone. "Optimization of Hybrid Sounding Rockets for Hypersonic Testing." Journal of Propulsion and Power 28, no. 2 (March 2012): 405–11. http://dx.doi.org/10.2514/1.b34218.
Повний текст джерелаEisen, 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.
Повний текст джерелаViant, Thibaut, Pascal Forquin, Dominique Saletti, Didier Imbault, Pierre Brunet, Julien Moriceau, and Gilles Poirey. "A testing technique to investigate the tensile behavior of propellant representative material." EPJ Web of Conferences 183 (2018): 02050. http://dx.doi.org/10.1051/epjconf/201818302050.
Повний текст джерелаApel, 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.
Повний текст джерелаBarato, 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.
Повний текст джерелаPalacz, 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.
Повний текст джерелаOkninski, Adam, Pawel Surmacz, Bartosz Bartkowiak, Tobiasz Mayer, Kamil Sobczak, Michal Pakosz, Damian Kaniewski, Jan Matyszewski, Grzegorz Rarata, and Piotr Wolanski. "Development of Green Storable Hybrid Rocket Propulsion Technology Using 98% Hydrogen Peroxide as Oxidizer." Aerospace 8, no. 9 (August 24, 2021): 234. http://dx.doi.org/10.3390/aerospace8090234.
Повний текст джерелаChelaru, Teodor Viorel, Valentin Pana, and Adrian Chelaru. "Modelling and Simulation of Suborbital Launcher for Testing." Applied Mechanics and Materials 555 (June 2014): 32–39. http://dx.doi.org/10.4028/www.scientific.net/amm.555.32.
Повний текст джерелаДисертації з теми "Hybrid propellant rockets – Testing"
Fernandez, Margaret Mary. "Propellant tank pressurization modeling for a hybrid rocket /." Online version of thesis, 2009. http://hdl.handle.net/1850/10631.
Повний текст джерелаArmstrong, Isaac W. "Development and Testing of Additively Manufactured Aerospike Nozzles for Small Satellite Propulsion." DigitalCommons@USU, 2019. https://digitalcommons.usu.edu/etd/7428.
Повний текст джерелаBernard, Geneviève. "Development of a hybrid sounding rocket motor." Thesis, 2013. http://hdl.handle.net/10413/8973.
Повний текст джерелаThesis (M.Sc.Eng.)-University of KwaZulu-Natal, Durban, 2013.
Leverone, Fiona Kay. "Performance modelling and simulation of a 100km hybrid sounding rocket." Thesis, 2013. http://hdl.handle.net/10413/11422.
Повний текст джерелаM.Sc.Eng. University of KwaZulu-Natal, Durban 2013.
Chowdhury, Seffat Mohammad. "Design and performance simulation of a hybrid sounding rocket." Thesis, 2012. http://hdl.handle.net/10413/9115.
Повний текст джерелаThesis (M.Sc.)-University of KwaZulu-Natal, Durban, 2012.
Chia-ChiehMo and 莫嘉傑. "Development and Testing of Pre-decomposition H2O2 and HTPB/Paraffin Hybrid Rockets." Thesis, 2016. http://ndltd.ncl.edu.tw/handle/w6v33t.
Повний текст джерела國立成功大學
航空太空工程學系
104
For the past decades, aerospace exploration has gotten lots of attentions, and the authorities of Taiwan also spent great efforts on autonomous developments of aerospace techniques. For the mission requirements in the future, a low cost, reliable and safety “green” propellant is the first priority. On the use of upper stage rockets and kick motors, liquid rockets is always a better options because of its high specific impulse, controllable thrust and multi-impulse, but the injection and system design is too complicated and expensive. Without the supports of advanced technologies, it nearly impossible to achieve the goal in the next few years. Hybrid rockets definitely a good choice to replace liquid rockets with its advantages, like controllable thrust, simpler constructions and high safety. For the last decade, the NCKU hybrid rocket team has succeed launching few N2O hybrid rockets, with thrust level from 100 kgf to 3000 kgf, and our team has also developed a composite silver catalyst used on the H2O2 monopropellant thruster, which has an excellent performance. Therefore, this thesis is going to demonstrate a pre-decomposition H2O2 hybrid rocket using our catalyst’s techniques. Key words : hybrid rocket, HTP, silver catalyst INTRODUCTION For recent years, rocket-grade hydrogen peroxide propulsion systems got a renewed interest due to its low toxicity, low cost and minor impact to the environment. For the hydrogen peroxide concentration over 92% has been called High Test Peroxide(HTP), and the decomposition temperature of 100 wt% HTP can reach 1267K. Therefore, lots of researches was developed on HTP mono-propellant, and silver, manganese based catalyst were commonly used. Hybrid rocket technology is known for more than 50 years, its separation storage of fuels and oxidizers makes it safer than other propulsion systems. Nowadays, the need for green propellants, safe storability and the use of upper stage rockets made hybrid rockets more attractive. LOX, HTP and N2O are widely used as oxidizers due to their low toxicity and low pollutant characteristics. The initiate of hybrid rocket required a heat source to gasify fuels until reaching a combustible condition, while a ignition device is needed, this makes the construction heavier and more complicated. A pre-decomposition HTP hybrid rocket which can auto-ignite the fuel is developed, and several collages and research units have already gotten tremendous results. To shorten the period of development, hybrid rockets are definitely a good replacement, our team has lots of experiences on N2O hybrid rockets and also developed a composite silver catalyst using on the HTP mono-propellant, so this thesis is going to demonstrate an auto-ignition hybrid rocket using pre-decomposition HTP. ANALYST AND DESIGN The motor design is based on the 2000N thrust requirement of upper stage rockets or kick motors, with 1/8 reduced scale, this thesis is going to verify the preliminary test of a 250 N thrust engine. The CEA (Chemical Equilibrium with Application) code provided a starting point of this investigation. Assumed the combustion chamber pressure of 380 psi and using 90 wt% H2O2 as our oxidizer and 50P (50 wt% of HTPB+50 wt% of paraffin) fuels. From this data we predicted Isp and optimum O/F ratios as showed in figure 1. The O/F ratio of 7 was selected with the best predicted Isp of 240.6s, compared to the stoichiometric O/F ratio of 7.7, this selected ratio will lead the combustion process to go through a fuel rich combustion, which can prevent the nozzle from corrosion, with the figure we can calculate the other parameters. m ̇_total=T/(Isp×g) m ̇_fuel=m ̇_total/(1+O/F) From the equations above, we can get the mass flow rate of oxidizers and fuels. Catalyst Bed And Liquid HTP Injector Design The catalyst bed design is based on the 1N mono-propellant techniques our team has developed, to decompose the H2O2 of 92.75g/s, we used the data captured in the previous experiments, and going through a scale enlargement. The composite catalyst is composed of silver flakes and γ-Al2O3 pellets due to its stronger hardness. Before entering the catalyst bed, the H2O2 was designed to go through an injector to spread liquid H2O2 uniformly into the catalyst. m ̇=C_d A√2ρ∆P The equation above was used for injector design, after assuming the pressure drop of 15 psi and the physical properties, we can estimate the total area required. Therefore, the injector was designed with 16 bores, each bore has a diameter of 1 mm. Gaseous Injector Design Injector plays an important role in hybrid rocket mixing mechanism, decomposed H2O2 will be led into combustion chamber, mixing with gaseous fuels which gasified by high temperature gaseous H2O2, the whole process is thermal-related, to reduce the energy loss is our first consideration. The mass flow rate can be expressed as m ̇=ρ_2 U_2 A Where U_2、ρ_2 can be represented as U_2=√(2γ/((γ-1))×RT_1×[1-(〖p_2/p_1 )〗^((γ-1)/γ)]) ρ_2=ρ_1 〖[p_2/p_1 ]〗^(1/γ) After decomposing, the exit gas can be regarded as a mixture of O2 and gaseous H2O, assume ideal gas, and set p2/p1 as 0.9, with previous assumption of chamber pressure 380 psi, the pressure ahead of injector can be calculated. We can also calculate the value of density using the ideal gas equation of state, with the assumption of ideal decomposition temperature to be 1067 K and γ=1.26, the ideal exit velocity of 289.106 m/s was estimated. From the calculation, the injector was designed with 8 rectangle grooves each has a length of 5 mm and width of 2 mm. Fuel Grain NCKU rocket team has made a great effort on N2O hybrid rocket, and developed the 50P fuel, which composed of 50% paraffin and 50% HTPB, though 50P fuels was chosen as our solid fuels, with the data acquired from previous experiments, a regression rate of 1.5 mm/s was assumed. m_fuel=ρ×A×L With the specific weight of 0.9, we can estimate the dimension of the fuel, a hollow cylinder fuel grain was made, with the inner diameter of 20 mm, outside diameter of 53 mm and the length of 180 mm. Nozzle From CEA calculations, the total mass flow rate of 106 g/s and the reaction temperature 2700 K during combustion process was known, with γ=1.14 m ̇=p_c A_t γ√(〖[2/(γ+1)]〗^((γ+1)/(γ-1))/√(γRT_c )) With the equation above, a nozzle throat area of 64.615 mm2 graphite IG-11 was machined as our nozzle. RESULTS AND DISCUSSION Characteristics of Composite Silver Catalyst To verified the feasibility of auto-ignition, the attempt of using pre-decomposition to ignite the fuel was tried. The catalyst was designed and went through a scale enlargement from the 1N monopropellant we developed. Silver flakes and zeolite supports was used, but from the preliminary tests, zeolite was replaced byγ-Al2O3 due to its lack of hardness, and a terrific outcome was obtained. During decomposition process, the temperature over 800K was reached, and the catalyst chamber pressure was maintained steady through whole process, and the structure of the catalyst bed still remain stable after tests. Combustion Characteristics of 50P Fuels The motor was successfully auto-ignited in the tests, and from one of the tests listed below, the average thrust of 205 N was measured, the H2O2 flow rate of 92.05g and average chamber pressure of 352.18 psi was also measured in that test, from the figure below, a stable curvature was obtained. And from the definition of specific impulse, a value of 192.12s was reached in case 4. Isp=F ̅/(m ̇_f+m ̇_o ) Swirl Effects in Combustion Chamber From the earlier studies, applied swirl injection to the hybrid rocket would increase the performance of combustion efficiency, and the intensity of swirling was defined as the swirl number S=2/3((1-〖(R_h/R)〗^3)/(1-〖(R_h/R)〗^2 ))tanα Thus a swirl injector was applied, with different swirl number, the table listed below shows the data measured in the tests, with stronger swirling injection, the combustion efficiency was contrarily decreased, these results differed from the studies, and from the flame observation, the causes was concluded. Due to insufficient gasified fuels, the oxidizer/fuel mixing was poor, the un-gasified fuels was carried to the downstream of combustion chamber. CONCLUSION This thesis was tend to demonstrate the auto-ignition of hybrid rocket using decomposition H2O2, a thrust level of 250 N was set, and the swirling injection was applied as well, few results we got was listed below: Development of high capacity catalyst bed: γ-Al2O3 was used as catalyst support, and zeolite was replaced due to its weaker hardness. During the tests, the catalyst was able to sustain the high H2O2 flow rate and the high temperature. Establishment of firing system and procedure: This thesis was dedicated in establishing a procedure and method on auto-firing a H2O2 hybrid rocket, and successfully conform the design of H2O2 catalyst bed and hybrid rocket motor. Succeed auto-firing a hybrid rocket motor: During tests, the time delay of auto-firing are less than 0.08 second, and succeed producing thrust. The key on promoting the mixing of oxidizer and fuels: In the tests, all combustion conditions were in fuel rich, but from the observation of internal ballistics, a low combustion efficiency condition was observed. With the use of swirling injection, from the previous studies, it should improve the combustion efficiency. In our tests, the regression rate did improved, but the internal ballistics and the flame still shows a low combustion efficiency, this may due to low viscosity of paraffin based fuels, and cause an entrainment effect.
Книги з теми "Hybrid propellant rockets – Testing"
Dranovsky, Mark L. Combustion instabilities in liquid rocket engines: Testing and development practices in Russia. Reston, Va: American Institute of Aeronautics and Astronautics, 2007.
Знайти повний текст джерелаRocker, M. Modeling on nonacoustic combustion instability in simulations of hybrid motor tests. Marshall Space Flight Center, Ala: National Aeronautics and Space Administration, Marshall Space Flight Center, 2000.
Знайти повний текст джерелаSati︠u︡kov, V. A. Tekhnologicheskai︠a︡ mekhanika toplivnykh magistraleĭ zhidkostnykh raketnykh dvigateleĭ. Moskva: Fizmatlit, 2009.
Знайти повний текст джерелаUnited States. Congress. House. Committee on Science, Space, and Technology. Subcommittee on Space Science and Applications. Results of the development motor 8 test firing: Hearing before the Subcommittee on Space Science and Applications of the Committee on Science, Space, and Technology, House of Representatives, One Hundredth Congress, first session, September 16, 1987. Washington: U.S. G.P.O., 1987.
Знайти повний текст джерелаUnited States. Congress. House. Committee on Science, Space, and Technology. Subcommittee on Space Science and Applications. Tests of the redesigned solid rocket motor program: Hearing before the Subcommittee on Space Science and Applications of the Committee on Science, Space, and Technology, U.S. House of Representatives, One Hundredth Congress, second session, January 27, 1988. Washington: U.S. G.P.O., 1988.
Знайти повний текст джерелаUnited States. Congress. House. Committee on Science, Space, and Technology. Subcommittee on Space Science and Applications. Tests of the redesigned solid rocket motor program: Hearing before the Subcommittee on Space Science and Applications of the Committee on Science, Space, and Technology, U.S. House of Representatives, One Hundredth Congress, second session, January 27, 1988. Washington: U.S. G.P.O., 1988.
Знайти повний текст джерелаUnited States. Congress. House. Committee on Science, Space, and Technology. Subcommittee on Space Science and Applications. Tests of the redesigned solid rocket motor program: Hearing before the Subcommittee on Space Science and Applications of the Committee on Science, Space, and Technology, U.S. House of Representatives, One Hundredth Congress, second session, January 27, 1988. Washington: U.S. G.P.O., 1988.
Знайти повний текст джерелаD, Cruit W., Smith A. W, George C. Marshall Space Flight Center., and AIAA/ASME/SAE/ASEE Joint Propulsion Conference (32nd : 1996 : Lake Buena Vista, Fla.), eds. Cold-flow study of hybrid rocket motor flow dynamics. [Huntsville, AL]: NASA Marshall Space Flight Center, 1996.
Знайти повний текст джерелаD, Cruit W., Smith A. W, and George C. Marshall Space Flight Center., eds. Cold-flow study of hybrid rocket motor flow dynamics: 32nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference, July 1-3, 1996, Lake Buena Vista, FL. [Huntsville, Ala: NASA Marshall Space Flight Center, 1996.
Знайти повний текст джерелаD, Cruit W., Smith A. W, and George C. Marshall Space Flight Center., eds. Cold-flow study of hybrid rocket motor flow dynamics: 32nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference, July 1-3, 1996, Lake Buena Vista, FL. [Huntsville, Ala: NASA Marshall Space Flight Center, 1996.
Знайти повний текст джерелаЧастини книг з теми "Hybrid propellant rockets – Testing"
Kara, Ozan, and Arif Karabeyoglu. "Hybrid Propulsion System: Novel Propellant Design for Mars Ascent Vehicles." In Propulsion - New Perspectives and Applications [Working Title]. IntechOpen, 2021. http://dx.doi.org/10.5772/intechopen.96686.
Повний текст джерелаТези доповідей конференцій з теми "Hybrid propellant rockets – Testing"
Werthman, W., and Christine Schroeder. "A preliminary design code for hybrid propellant rockets." In 32nd Aerospace Sciences Meeting and Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1994. http://dx.doi.org/10.2514/6.1994-6.
Повний текст джерелаHOLLMAN, S., and R. FREDERICK, JR. "Labscale testing techniques for hybrid rockets." In 29th Joint Propulsion Conference and Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1993. http://dx.doi.org/10.2514/6.1993-2409.
Повний текст джерелаCasalino, Lorenzo, and Dario Pastrone. "Optimization of Hybrid Sounding Rockets for Hypersonic Testing." In 45th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2009. http://dx.doi.org/10.2514/6.2009-4841.
Повний текст джерелаWhitmore, Stephen A., Zachary S. Spurrier, Jerome K. Fuller, and John D. Desain. "A Survey of Additively Manufactured Propellant Materials for Arc-Ignition of Hybrid Rockets." In 51st AIAA/SAE/ASEE Joint Propulsion Conference. Reston, Virginia: American Institute of Aeronautics and Astronautics, 2015. http://dx.doi.org/10.2514/6.2015-4034.
Повний текст джерелаNguyen, Nam, Victor Ong, Alan Villanueva, Dehwei Hsu, Nathan Nguyen, Navdeep Dhillon, and Praveen Shankar. "Design and testing of solid propellant rockets towards NASA Student Launch and Intercollegiate Rocket Engineering Competitions." In 2018 Joint Propulsion Conference. Reston, Virginia: American Institute of Aeronautics and Astronautics, 2018. http://dx.doi.org/10.2514/6.2018-4864.
Повний текст джерелаRoux, Vincent, and Shawn Duan. "Characterizing Potential Damage to Landers and Their Payloads Caused by Regolith Ejecta During Operations on or Near the Surface of the Moon, Mars, and Other Worlds." In ASME 2021 International Mechanical Engineering Congress and Exposition. American Society of Mechanical Engineers, 2021. http://dx.doi.org/10.1115/imece2021-70923.
Повний текст джерелаKobald, M., C. Schmierer, U. Fischer, K. Tomilin, A. Petrarolo, and M. Rehberger. "The HyEnD stern hybrid sounding rocket project." In Progress in Propulsion Physics – Volume 11. Les Ulis, France: EDP Sciences, 2019. http://dx.doi.org/10.1051/eucass/201911025.
Повний текст джерела