Relevant bibliographies by topics / Cryogenic rocket propellants

Contents

  1. Journal articles
  2. Dissertations / Theses
  3. Books
  4. Book chapters
  5. Conference papers
  6. Reports

Academic literature on the topic 'Cryogenic rocket propellants'

Author: Grafiati
Published: 4 June 2021
Last updated: 13 February 2022

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Journal articles on the topic "Cryogenic rocket propellants"

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Elliott, J. P., J. M. McNair, M. H. Spritzer, J. A. Hurley, and S. A. Rising. "CRYOGENIC WASHOUT OF SOLID ROCKET PROPELLANTS." International Journal of Energetic Materials and Chemical Propulsion 4, no. 1-6 (1997): 221–30. http://dx.doi.org/10.1615/intjenergeticmaterialschemprop.v4.i1-6.260.

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Chehroudi, Bruce. "Recent Experimental Efforts on High-Pressure Supercritical Injection for Liquid Rockets and Their Implications." International Journal of Aerospace Engineering 2012 (2012): 1–31. http://dx.doi.org/10.1155/2012/121802.

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Pressure and temperature of the liquid rocket thrust chambers into which propellants are injected have been in an ascending trajectory to gain higher specific impulse. It is quite possible then that the thermodynamic condition into which liquid propellants are injected reaches or surpasses the critical point of one or more of the injected fluids. For example, in cryogenic hydrogen/oxygen liquid rocket engines, such as Space Shuttle Main Engine (SSME) or Vulcain (Ariane 5), the injected liquid oxygen finds itself in a supercritical condition. Very little detailed information was available on the behavior of liquid jets under such a harsh environment nearly two decades ago. The author had the opportunity to be intimately involved in the evolutionary understanding of injection processes at the Air Force Research Laboratory (AFRL), spanning sub- to supercritical conditions during this period. The information included here attempts to present a coherent summary of experimental achievements pertinent to liquid rockets, focusing only on the injection of nonreacting cryogenic liquids into a high-pressure environment surpassing the critical point of at least one of the propellants. Moreover, some implications of the results acquired under such an environment are offered in the context of the liquid rocket combustion instability problem.
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Orlin, Sergei A. "Use of cryogenic components of propellants for liquid-propellant rocket engines and in life support systems of manned space vehicles." MATEC Web of Conferences 324 (2020): 01005. http://dx.doi.org/10.1051/matecconf/202032401005.

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The cited materials show the use of oxygen, hydrogen, liquefied natural gases (methane) and fluorine as components of the fuel for liquid-propellant rocket engines (LRE). The reasons for the need to use oxygen as an oxidizing agent are indicated. The advantages and disadvantages are disclosed from the point of view of using the listed components as fuel elements for liquid-propellant rocket engines. The issues of ecology when using the considered fuels are reviewed. Shown not only the use of cryogenic components as fuel for LRE, but also in life support systems in manned spacecraft in space research.
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Guoyuan, Zhang, and Yan Xiu-Tian. "Analysis of two phase flow in liquid oxygen hybrid journal bearings for rocket engine turbopumps." Industrial Lubrication and Tribology 66, no. 1 (2014): 31–37. http://dx.doi.org/10.1108/ilt-09-2011-0072.

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Purpose – A hybrid bearing of advanced cryogenic rocket engine turbopump is designed. For cryogenic fluid propellants (such as liquid oxygen) as the lubrication of bearing, bearings operating close to liquid-vapor region (near the critical point or slightly sub-cooled) are likely to develop a two phase flow region. The paper aims to discuss these issues. Design/methodology/approach – In this paper, an all liquid, liquid-vapor mixture, and all vapor, i.e. a continuous vaporization bulk flow model of density and viscosity for mixture fluid, is presented, and the general Reynolds equation and energy equation with two phase flow as lubricants is solved. The static and dynamic performance of a 50-mm-radius hybrid bearing are obtained under 20,000 rpm speed and 10 MPa supply pressure. Findings – The results show that the variations of performance of bearing operating under cryogenic liquid oxygen are not bounded by the all liquid and all vapor cases in the liquid-vapor mixture range. There behaviours are attributed to the large change in the compressibility character of the flow. Research limitations/implications – For validating the correctness of analytical model, an experimental study on the liquid-vapor nitrogen mixture lubricated hybrid journal bearings is being carried out where low-viscosity nitrogen was selected as the lubricant for the sake of safety. Soon after, the authors will discuss the results and publish them in the new papers. Originality/value – An all liquid, liquid-vapor mixture, and all vapor, i.e. a continuous vaporization bulk flow model of density and viscosity for mixture fluid, is presented. The static and dynamic performance of hybrid bearings with two phase flow as lubricants are obtained.
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Lo, Roger E. "Modular Dissected Cryogenic Solid-Rocket Propellant Grains." Acta Astronautica 51, no. 10 (2002): 683–91. http://dx.doi.org/10.1016/s0094-5765(02)00018-8.

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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 (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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Domashenko, Anatoly M., and Andrey L. Dovbish. "The process of production of liquefied methane - the component of rocket propellant." MATEC Web of Conferences 324 (2020): 01004. http://dx.doi.org/10.1051/matecconf/202032401004.

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The use of new fuel components, such as LNG or liquefied methane, in rocket-space, aviation and other special-purpose engineering is promising. On the basis of these fuel components it is possible to provide a number of technical and tactical parameters of aircrafts, which are not achievable when using standard fuels. Considered were the cryogenic systems developed by PJSC "Cryogenmash" for natural gas liquefaction with liquid methane recovery by the method of low-temperature condensation, stage separation and rectification. The second method allows to reduce the content of not only low boiling but also high boiling liquids in methane liquefied.
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Dileep Kumar, N., K. Thomas Tharian, Aby Isaac, and P. V. Venkitakrishnan. "Effect of Brazing and Heat Treatment Cycle on the Mechanical Properties of Base Materials." Materials Science Forum 830-831 (September 2015): 310–13. http://dx.doi.org/10.4028/www.scientific.net/msf.830-831.310.

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Brazing is extensively used in liquid rocket engines for realizing various subsystems. In the case of cryogenic engines, brazing operation is done to realize the gas generator. Gas Generator is one of the major systems of cryogenic engine. It generates and supplies hot gases required for running turbine of main turbo pump. This uses liquid oxygen and gaseous hydrogen as propellant combination. Combustion chamber of Gas Generator is of double walled construction with the cylindrical outer shell of transition class ICSS-0716-301 austenitic-martensitic stainless steel and inner shell of ICSS-1218 321, aTi stabilized austenitic stainless steel material brazed together with Fe-Ni-Mn type braze alloy at a temperature of 1180°C. This temperature can cause the grain growth and related issues to the base material. Thus the present work focuses on the effect of the brazing/thermal cycle on mechanical properties and microstructure of the base materials in post braze condition. The results obtained on metallurgical/mechanical behavior of the material showed the different grain growth patterns in inner and outer shell materials. This helped in understanding the effect of brazing condition on the changes in mechanical properties of base materials.
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Jung, Tae-Kyu, and Soo-Yong Lee. "Development of BLDC Motor Driven Cryogenic Thrust Control Valve for Liquid Propellant Rocket Engine." Journal of the Korean Society for Aeronautical & Space Sciences 38, no. 10 (2010): 1026–30. http://dx.doi.org/10.5139/jksas.2010.38.10.1026.

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Li, Cui, Yuhan Zhuang, Yiwei Cheng, and Yanzhong Li. "Study on pressure wave propagation through cryogenic condensing two-phase flow in liquid rocket propellant feedline." Cryogenics 112 (December 2020): 103193. http://dx.doi.org/10.1016/j.cryogenics.2020.103193.

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Dissertations / Theses on the topic "Cryogenic rocket propellants"

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Ganu, Hrishikesh Vidyadhar. "A Morphological Technique For Direct Drop Size Measurement Of Cryogenic Sprays." Thesis, 2005. http://etd.iisc.ernet.in/handle/2005/1481.

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Biju, Kumar K. S. "Role Of Hydrogen Injection Temperature On The Combustion Instability Of Cryogenic Rocket Engine." Thesis, 2012. http://etd.iisc.ernet.in/handle/2005/2297.

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Physical mechanism for high frequency instability in cryogenic engines at low hydrogen injection temperature has been a subject of debate for long time. Experimental and early developmental studies revealed no instabilities and it was only much later when liquid hydrogen at lower initial temperature (~50 to 100 K) was injected into the combustion chamber that instabilities were detected. From the compilations of the experimental data related to the instability of cryogenic engines by Hulka and Hutt, it was found that the instability was strongly connected to the temperature of hydrogen. Experiments conducted with hydrogen temperature ramping from a higher value to lower values indicated that the temperatures in excess of 90 K favor stability under most practical operating conditions. Even though this has been known for over forty years, there has been no clear and simple explanation for this. Many physical mechanisms have been hypothesized to explain how temperature ramping causes instability, but all appear to have limited range of applicability. Current understanding of cryogenic engine combustion instability has been achieved through a combination of experimental investigation and approximate analytical models as well as CFD tools. Various researchers have tried to link the low hydrogen injection temperature combustion instability phenomena with various potential mechanisms for combustion instability. They involve coupling of combustion acoustics with atomization, vaporization, mixing, chemical kinetics or any combination of these processes. Various studies related to the effect of recess, injector hydrodynamics, acoustic damping of gas liquid scheme injectors and effect of drop size distribution on the stability characteristics of cryogenic engines were compiled in the thesis. Several researchers examined fuel droplet vaporization as the rate controlling mechanism. Recently a new method for the evaluation of stability characteristics of the engine using model chamber were proposed by Russians and this is based on mixing as the rate controlling mechanism. Pros and cons of this method were discussed. Some people examined the combustion instability of rocket engines based on chemistry dynamics. A considerable amount of analytical and numerical studies were carried out by various researchers for finding out the cause of combustion instability. Because of the limitations of their analysis, they could not successfully explain the cause of combustion instability at low hydrogen injection temperature. A compilation of previous numerical studies were carried out. A number of researchers have applied CFD in the study of combustion instabilities in liquid propellant rocket engines. In the present thesis, a theoretical model has been developed based on the vaporization of droplets to predict the stability characteristics of the engine. The proposed concept focuses on three dimensional simulation of combustion instability for giving some meaningful explanations for the experimental work presented in the literature. In the present study the pressure wave corresponding to the transverse modes were superimposed on a three dimensional steady state operating conditions. Steady state parameters were obtained from the three dimensional combustion modeling. The conservation equations for mass, momentum and energy are non dimensionalized for facilitating the order of magnitude analysis. In order to do the stability analysis, variables are represented as the sum of their steady values and deviation from the steady state. A harmonic time dependence is assumed for the perturbations. For the transverse mode of oscillations independent variables of the zeroth order equations are r and θ only and the dependant variables are not functions of the axial distance. The axial dependence comes only through the first order equations. In this analysis, the wave motion in the combustion chamber is assumed to be linear, confining the nonlinearity to the vaporization process only. The reason behind making this assumption is that the vaporization process is the major mechanism driving the instability. Vaporization histories of liquid oxygen drops in a combustor with superimposed transverse oscillations were computed and stability characteristics of the engine were estimated. The stability characteristics of the engine are accessed from the solutions of first order equations. Effects of various parameters like droplet diameter, hydrogen injection temperature and hydrogen injection area on the stability characteristics of cryogenic engines are studied. A comparison of predicted and published experimental results was made which showed general agreement between experiment and computation. The present study and experimental results show clearly that hydrogen injection velocity is the critical parameter for instability rather than hydrogen injection temperature. What has happened in actual experiments when hydrogen injection temperature is varied is an effective alteration of the injection velocity that leads to the situation of instability. For higher relative velocity between hydrogen and liquid oxygen, the response of the vaporization rate in the presence of pressure wave is minimum compared to lower relative velocity. Due to this cryogenic engines will go to unstable mode at lower relative velocity.
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Rakesh, P. "A Study of the Characteristics of Gas-On-Liquid Impinging Injectors." Thesis, 2014. http://hdl.handle.net/2005/3126.

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The work presented here pertains to investigations on gas-on-liquid type of impinging injectors with a generic approach with prospective applications in several areas, and at places with particular emphasis on cryogenic or semi-cryogenic liquid propellant rockets. In such rockets, one of the components arrives at the injector in a gaseous phase after passing through the regenerative coolant passages or a preceding combustion stage. Most often, the injectors in such systems are of shear coaxial type. The shear coaxial injectors suffer from several disadvantages like complexity in design, manufacture and quality control. Adoption of impinging jet configuration can alleviate these problems in addition to providing further benefits in terms of cost, robustness in high temperature environment and manifolding. However, there is very little literature on gas-on-liquid injectors either in this context or in any other Even for the simplest form of impinging injectors such as like-on-like doublets, literature provides no conclusive direction at describing a spray from the theoretical models of physical mechanisms. Empirical approach is still the prime mode of obtaining a proper understanding of the phenomena. Steady state spray characterization includes mainly of describing the spatial distribution of liquid mass and drop size distribution as a function of geometric and injection parameters. The parameters that are likely to have an impact on spray characteristics are orifice diameter, ratio of orifice length to diameter, pre-impingement length of individual jets, inter orifice distance, impingement angle, jet velocity and condition of the jet just before impingement. The gas-on- liquid configuration is likely to experience some qualitative changes because of the expansion of the gas jet. The degree to which each one of the above variables influences the drop size and mass distribution having implication to combustion performance forms the core theme of the thesis. A dedicated experimental facility has been built, calibrated and deployed exhaustively. While spray drop size measurement is done largely by a laser diffraction instrument, some of the cases warranted an image processing technique. Two different image processing algorithms are developed in-house for this purpose. The granulometric image processing method developed earlier in the group for cryogenic sprays is modified and its applicability to gas-on-liquid impinging sprays are verified. Another technique based on the Hough transform which is feature extraction technique for extracting quantitative information has also been developed and used for gas-on-liquid impinging injectors. A comparative study of conventional liquid-on-liquid doublet with gas-on-liquid impinging injectors are first made to establish the importance of studying gas-on-liquid impinging injectors. The study identifies the similarities and differences between the two types and highlights the features that make such injectors attractive as replacements to coaxial configuration. Spray structure, drop-size mass distributions are quantified for the purpose of comparison. This is followed by a parametric study of the gas-on-liquid impinging injectors carried out using identified control variables. Though momentum ratio appeared to be a suitable parameter to describe the spray at any given impingement angle, the variations due to impingement angle had to be factored in. It was found that normal gas momentum to liquid mass is an apt parameter to generalize the spray characteristics. It was also found that using identical nozzles for desired mass ratio could lead to rather large deflection of the spray which may not be acceptable in combustion chamber design. One way of overcoming this is to work with unequal orifice sizes for gas and liquid. It was found that using smaller gas orifice for a given liquid orifice resulted in lower SMD (Sauter Mean Diameter of the spray) for constant gas and liquid mass flow rates. This is attributable to the high dynamic pressure of gas in the case of smaller gas orifices for the same mass flow rate. The impinging liquid jets with unequal momentum in the doublet configuration would result in non-uniform mass and mixture ratio distribution within the combustion chamber which may have to operate under varying conditions of mass flow rates and/or mixture ratio. The symmetrical arrangement of triplet configuration can eliminate this problem at the same time generating finely atomized spray and a homogeneous mixture ratio. In view of the scanty literature available in this field, the atomization characteristics of the spray generated by liquid centered triplet jets are examined in detail. It was found that as in the case of gas-on-liquid impinging doublets, normal gas momentum to liquid mass is an ideal parameter in describing the spray. Variants of this configuration are studied recently for many other applications too. As done in the case of doublets, efforts have also been made to compare gas centered triplet to liquid-liquid triplet. It was found that the trend of SMD of gas centered triplet is different from that of liquid-liquid triplets, thus pointing to a different mechanism in play. The SMD in the case of liquid-liquid triplets decreases monotonically with increasing specific normal momentum. It is to be noted that specific normal momentum is an ideal parameter for describing the spray characteristics of liquid-liquid triplets and doublets. In the case of gas centered triplet the SMD first increases and then decreases with specific normal momentum, the inversion point depends on the gas mass flow rate for a constant specific normal momentum. The thesis concludes with a summary of the major observations of spray structures for all the above injector configurations and quantifies the parametric dependencies that would be of use to engineering design
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Books on the topic "Cryogenic rocket propellants"

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Ko, William L. Thermocryogenic buckling and stress analyses of a partially filled cryogenic tank subjected to cylindrical strip heating. National Aeronautics and Space Administration, Office of Management, Scientific and Technical Information Program, 1994.

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Ko, William L. Thermocryogenic buckling and stress analyses of a partially filled cryogenic tank subjected to cylindrical strip heating. National Aeronautics and Space Administration, Office of Management, Scientific and Technical Information Program, 1994.

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Kearns, Kimberly A. The revamping of an ignition test facility. National Aeronautics and Space Administration, Glenn Research Center, 2002.

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Chato, David J. Ground testing on the nonvented fill method of orbital propellant transfer: Results of initial test series. National Aeronautics and Space Administration, 1991.

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United States. National Aeronautics and Space Administration., ed. Advanced small rocket chambers option II: Fundamental processes and material evaluation : final report. National Aeronautics and Space Administration, 1993.

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Scott, Adams, Palaszewski Bryan, and United States. National Aeronautics and Space Administration., eds. Nanoparticulate gellants for metallized gelled liquid hydrogen wth aluminum. National Aeronautics and Space Administration, 1996.

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Preliminary assessment of using gelled and hybrid propellant propulsion for VTOL/SSTO launch systems. National Aeronautics and Space Administration, Lewis Research Center, 1998.

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Bryan, Palaszewski, O'Leary Robert, and Lewis Research Center, eds. Preliminary assessment of using gelled and hybrid propellant propulsion for VTOL/SSTO launch systems. National Aeronautics and Space Administration, Lewis Research Center, 1998.

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United States. National Aeronautics and Space Administration., ed. An overview of the British Aerospace "Hotol" transatmospheric vehicle. National Aeronautics and Space Administration, 1986.

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United States. National Aeronautics and Space Administration., ed. Performance tests of a liquid hydrogen propellant densification ground system for the X33/RLV. National Aeronautics and Space Administration, 1997.

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Book chapters on the topic "Cryogenic rocket propellants"

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Vishnu, S. B., and Biju T. Kuzhiveli. "Effect of Roughness Elements on the Evolution of Thermal Stratification in a Cryogenic Propellant Tank." In Low-Temperature Technologies [Working Title]. IntechOpen, 2021. http://dx.doi.org/10.5772/intechopen.98404.

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The cryogenic propulsion era started with the use of liquid rockets. These rocket engines use propellants in liquid form with reasonably high density, allowing reduced tank size with a high mass ratio. Cryogenic engines are designed for liquid fuels that have to be held in liquid form at cryogenic temperature and gas at normal temperatures. Since propellants are stored at their boiling temperature or subcooled condition, minimal heat infiltration itself causes thermal stratification and self-pressurization. Due to stratification, the state of propellant inside the tank varies, and it is essential to keep the propellant properties in a predefined state for restarting the cryogenic engine after the coast phase. The propellant’s condition at the inlet of the propellant feed system or turbo pump must fall within a narrow range. If the inlet temperature is above the cavitation value, cavitation will likely to happen to result in the probable destruction of the flight vehicle. The present work aims to find an effective method to reduce the stratification phenomenon in a cryogenic storage tank. From previous studies, it is observed that the shape of the inner wall surface of the storage tank plays an essential role in the development of the stratified layer. A CFD model is established to predict the rate of self-pressurization in a liquid hydrogen container. The Volume of Fluid (VOF) method is used to predict the liquid–vapor interface movement, and the Lee phase change model is adopted for evaporation and condensation calculations. A detailed study has been conducted on a cylindrical storage tank with an iso grid and rib structure. The development of the stratified layer in the presence of iso grid and ribs are entirely different. The buoyancy-driven free convection flow over iso grid structure result in velocity and temperature profile that differs significantly from a smooth wall case. The thermal boundary layer was always more significant for iso grid type obstruction, and these obstructions induces streamline deflection and recirculation zones, which enhances heat transfer to bulk liquid. A larger self-pressurization rate is observed for tanks with an iso grid structure. The presence of ribs results in the reduction of upward buoyancy flow near the tank surface, whereas streamline deflection and recirculation zones were also perceptible. As the number of ribs increases, it nullifies the effect of the formation of recirculation zones. Finally, a maximum reduction of 32.89% in the self-pressurization rate is achieved with the incorporation of the rib structure in the tank wall.
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"Fundamentals of Supercritical Mixing and Combustion of Cryogenic Propellants." In Liquid Rocket Thrust Chambers. American Institute of Aeronautics and Astronautics, 2004. http://dx.doi.org/10.2514/5.9781600866760.0339.0367.

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"Simulation and Analysis of Thrust Chamber Flowfields: Cryogenic Propellant Rockets." In Liquid Rocket Thrust Chambers. American Institute of Aeronautics and Astronautics, 2004. http://dx.doi.org/10.2514/5.9781600866760.0527.0551.

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Conference papers on the topic "Cryogenic rocket propellants"

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CLAPP, MITCHELL, and MAXWELL HUNTER. "A single stage to orbit rocket with non-cryogenic propellants." In 29th Joint Propulsion Conference and Exhibit. American Institute of Aeronautics and Astronautics, 1993. http://dx.doi.org/10.2514/6.1993-2285.

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Carrick, Patrick. "Theoretical performance of high energy density cryogenic solid rocket propellants." In 31st Joint Propulsion Conference and Exhibit. American Institute of Aeronautics and Astronautics, 1995. http://dx.doi.org/10.2514/6.1995-2892.

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Tao, Yujing, Wu Jiping, and Wang Zhenguo. "Design and Application of Throttling Venturi for Cryogenic Propellants in Tripropellant Rocket Engine." In 42nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit. American Institute of Aeronautics and Astronautics, 2006. http://dx.doi.org/10.2514/6.2006-4883.

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Lee, Sun-Kyung, Sang Yeop Han, Dong-Soon Shin, and R. Ray Taghavi. "Thermal-Fluidic Numerical Analysis for the Development of Heat Exchanger in the Propellant Tank Pressurization System of KSLV-II Upper Stage." In ASME/JSME/KSME 2015 Joint Fluids Engineering Conference. American Society of Mechanical Engineers, 2015. http://dx.doi.org/10.1115/ajkfluids2015-03742.

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KSLV-II (Korea Space Launch Vehicle - II) launch vehicle is a three staged satellite launch vehicle using a liquid propellant propulsion system in all three stages. It will deliver 1,500 kg satellite to Sun Synchronous Orbit (SSO, 700 km, 98.2°) or 2,600 kg satellite to Low Earth Orbit (LEO, 300 km, 80.3°). Propellants for KSLV-II are kerosene as a fuel and liquid oxygen as an oxidizer for propelling. Those fuel and oxidizer are stored in on-board tanks separately. To run a liquid propellant rocket engine on ground or in flight, those propellants should be supplied to LRE’s using so-called Propellant Pressurizing Sub-system, which makes propellants be pressurized in tanks using pressurant. A pressurant for PPSS of KSLV-II is helium, which is stored in tanks located in an oxidizer tank. The stored He is under cryogenic condition (50 K) as gaseous state. Such He is heated and expanded through heat exchanger, which is using a combustion gas coming out from gas generator for turbo-pump as an energy source, to be used as pressurant. This paper contains the results of performance analysis and thermal-fluidic numerical analysis to develop the above-mentioned heat exchanger for KSLV-II upper stage (the 2nd stage). The technical requirements for such heat exchanger are as follows: pressurant mass flow rate for oxidizer tank - 0.127 kg/sec; and for fuel tank - 0.043 kg/sec. The outlet temperature of He from heat exchanger is 550±10 K.
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Vermes, A. G., and C. Lettieri. "Source Term Based Modeling of Rotating Cavitation in Turbopumps." In ASME Turbo Expo 2018: Turbomachinery Technical Conference and Exposition. American Society of Mechanical Engineers, 2018. http://dx.doi.org/10.1115/gt2018-75194.

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The recent growth of private options in launch vehicles has substantially raised price competition in the space launch market. This has increased the need to deliver reliable launch vehicles at reduced engine development cost, and has led to increased industrial interest in reduced order models. Large-scale liquid rocket engines require high-speed turbopumps to inject cryogenic propellants into the combustion chamber. These pumps can experience cavitation instabilities even when operating near design conditions. Of particular concern is rotating cavitation, which is characterized by an asymmetric cavity rotating at the pump inlet, which can cause severe vibration, breaking of the pump and loss of the mission. Despite much work in the field, there are limited guidelines to avoid rotating cavitation during design and its occurrence is often assessed through costly experimental testing. This paper presents a source term based model for stability assessment of rocket engine turbopumps. The approach utilizes mass and momentum source terms to model cavities and hydrodynamic blockage in inviscid, single-phase numerical calculations, reducing the computational cost of the calculations by an order of magnitude compared to traditional numerical methods. Comparison of the results from the model with experiments and high-fidelity calculations indicates agreement of the head coefficient and cavity blockage within 0.26% and 5% respectively. The computations capture rotating cavitation in a 2D inducer at the expected flow coefficient and cavitation number. The mechanism of formation and propagation of the instability is correctly reproduced.
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Ferraiuolo, M., A. Martucci, F. Battista, and D. Ricci. "Thermostructural Analyses Supporting the Design of the HYPROB Heat Sink Subscale Breadboard." In ASME 2014 International Mechanical Engineering Congress and Exposition. American Society of Mechanical Engineers, 2014. http://dx.doi.org/10.1115/imece2014-36882.

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Today’s rocket engines regeneratively cooled using high energy cryogenic propellants (e.g. LOX and LH2, LOX and LCH4) play a major role due to the high combustion enthalpy (10–13.4 kJ/kg) and the high specific impulse of these propellants. In the frame of the HYPROB/Bread project, whose main goal is to design build and test a 30 kN regeneratively cooled thrust chamber, a breadboard has been conceived in order to: • investigate the behavior of the injector that will be employed in the full scale final demonstrator, • to obtain a first estimate of the heat flux on the combustion chamber for models validation, • to implement a “battleship” chamber for a first verification of the stability of the combustion The breadboard is called HS (Heat Sink) and it is made of CuCrZr (Copper Chromium Zirconium alloy), Inconel 718 and TZM (Titanium Zirconium Molybdenum alloy). The aim of the present paper is to illustrate the thermostructural design conducted on the breadboard by means of a Finite Element Method code taking into account the viscoplastic behavior of the adopted materials. An optimization process has been carried out in order to keep the structural integrity of the breadboard maximizing the life cycles of the component. Heat fluxes generated by combustion gases have been evaluated by means of CFD quick analyses, while convection and radiation with the external environment have not been considered in order to be as conservative as possible from a thermostructural point of view. Transient thermal analyses and static structural analyses have been performed by means of ANSYS code adopting an axisymmetric model of the chamber. These analyses have demonstrated that the Breadboard can withstand the design goal of 3 thermo-mechanical cycles with a safety factor equal to 4 considering a firing time equal to 3 seconds.
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Behruzi, Philipp, Joerg Klatte, Nicolas Fries, Andreas Schuette, Burkhard Schmitz, and Horst Koehler. "Cryogenic Propellant Management Sounding Rocket Experiments on TEXUS 48." In 49th AIAA/ASME/SAE/ASEE Joint Propulsion Conference. American Institute of Aeronautics and Astronautics, 2013. http://dx.doi.org/10.2514/6.2013-3904.

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Soller, Sebastian, Nico Rackemann, and Gerhard Kroupa. "Laser Ignition Application to Cryogenic Propellant Rocket Thrust Chambers." In Laser Ignition Conference. OSA, 2017. http://dx.doi.org/10.1364/lic.2017.lfa4.3.

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Kelly, Sean, and Corin Segal. "Experiments in Thermosensitive Cavitation of a Cryogenic Rocket Propellant Surrogate." In 50th AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition. American Institute of Aeronautics and Astronautics, 2012. http://dx.doi.org/10.2514/6.2012-1283.

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Naraghi, Mohammad H. N., and Edmundo M. Nunes. "Effects of Gas Radiation on the Thermal Characteristics of Regeneratively Cooled Rocket Engines." In ASME 2002 International Mechanical Engineering Congress and Exposition. ASMEDC, 2002. http://dx.doi.org/10.1115/imece2002-33920.

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Abstract:
This paper studies the effects of radiative heat transfer on the thermal characteristics of regeneratively cooled rocket engines. A conjugated radiative, conductive and convective model is used to analyze the effects of radiative heat transfer in two regeneratively cooled rocket engines. One engine has liquid hydrogen and liquid oxygen as the propellant and liquid hydrogen as the coolant. The other engine has RP1 (a hydrocarbon fuel) and liquid oxygen as the propellant and liquid oxygen as the coolant. It is shown that gas radiation has some effect on the wall temperature of the LH2-LO2 engine and a small effect on its coolant flow characteristics. For the RP1-LO2 engine, however, gas radiation significantly increases the coolant pressure drop, temperature and Mach number. It is also shown that radiation effects must be addressed in cooling channel design, so that wall temperatures and cryogenic coolant flow temperature/pressure are at suitable levels.
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Reports on the topic "Cryogenic rocket propellants"

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Rice, Eric E., Edward Bangsund, Martin J. Chiaverini, and Daniel J. Gramer. ORBITEC Advanced Cryogenic Solid Hybrid Rocket Engine and Propellant Developments: A 1998 Status Report. Defense Technical Information Center, 1998. http://dx.doi.org/10.21236/ada409805.

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