Relevant bibliographies by topics / Thrust chamber pressure

Contents

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

Academic literature on the topic 'Thrust chamber pressure'

Author: Grafiati
Published: 4 June 2021
Last updated: 1 February 2022

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Journal articles on the topic "Thrust chamber pressure"

1

T.N., Rajesh, T. J. Sarvoththama Jothi, and Jayachandran T. "Cold flow studies in a vortex thrust chamber." Aircraft Engineering and Aerospace Technology 91, no. 1 (2018): 69–77. http://dx.doi.org/10.1108/aeat-07-2017-0167.

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Purpose The purpose of this paper is to estimate the chamber pressure and flow behaviour in a vortex thrust chamber (VTC) during the cold flow with hydrogen and oxygen as propellants. Design/methodology/approach Experiments are carried out in a VTC with a different mixture ratio of hydrogen and oxygen. The pressures developed inside the VTC are measured. Numerical simulations are carried out to understand the flow patterns of fuel and oxidizer inside the VTC. Findings The chamber pressure is influenced by the type of injection of propellant and mixture ratio. Tangential injection of propellant
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DC, Naresh, and Rudesh M. "Design and Analysis of Combustion Chamber for HAN Based Mono Propulsion System Thruster for Spacecraft Application." International Journal of Aviation Science and Technology vm01, is02 (2020): 66–70. http://dx.doi.org/10.23890/ijast.vm01is02.0203.

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This paper presents a preliminary dimensional study of combustion chamber using Hydroxyl Ammonium Nitrate (HAN) propellants for spacecraft application. The combustion chamber consists of two parts namely thrust chamber and Convergent-Divergent (C-D) nozzle. The design for combustion chamber is very much important because the chemical energy in the propellant released within this closed volume i.e., thrust chamber and gets expanded through the C-D nozzle part. So the chamber must be designed to provide a necessary space for the propellants to react and release maximum available energy and also
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Lim, Yeerang, Jaecheong Lee, Hyochoong Bang, Hwanil Huh, and Hosung Lee. "Reduced Actuator Set for Pressure Control and Thrust Distribution for Multinozzle Propulsion Systems." International Journal of Aerospace Engineering 2017 (2017): 1–11. http://dx.doi.org/10.1155/2017/4104212.

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This study investigated an approach to reduce the number of actuators used for internal pressure control and thruster allocation in a multinozzle solid propulsion system. In the proposed design, the throat areas of four divert nozzles are controlled by only three actuators, and chamber pressure maintenance and thrust distribution are achieved by controlling the throat areas. Using the proposed actuator set, thrust allocation can be accomplished in a more efficient way than when independent actuators are employed for each nozzle.
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Asraff, A. K., S. Sheela, Krishnajith Jayamani, S. Sarath Chandran Nair, and R. Muthukumar. "Material Characterisation and Constitutive Modelling of a Copper Alloy and Stainless Steel at Cryogenic and Elevated Temperatures." Materials Science Forum 830-831 (September 2015): 242–45. http://dx.doi.org/10.4028/www.scientific.net/msf.830-831.242.

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High performance rockets are developed using cryogenic technology. High thrust cryogenic rocket engines operating at elevated temperatures and pressures are the backbone of such rockets. The thrust chamber of such engines, which produce the thrust for the propulsion of the rocket, can be considered as structural elements. Often double walled construction is employed for these chambers for better cooling and enhanced performance. The thrust chamber investigated here has its hot inner wall fabricated out of a high conductivity high ductility copper alloy and outer wall made of a ductile stainles
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Rajan, K. M., and K. Narasimhan. "Design and structural analysis of a thrust chamber for a spinning supersonic rocket – a case study." Aeronautical Journal 109, no. 1094 (2005): I—VI. http://dx.doi.org/10.1017/s0001924000000701.

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Abstract The design of a thrust chamber for a rocket propulsion system is a challenging task. The thrust chamber has to be designed for minimum structural weight with an adequate factor of safety. This calls for a thorough knowledge of various structural loads, both internal and external, and the behaviour of the structure in flight. This paper presents the design and structural analysis of a pressure vessel used as thrust chamber for a rocket propulsion unit. The effects of kinetic heating, thermal stress, spinning and various aerodynamic loads and their mutual interactions are accounted for
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Sofyan, Sofyan, and Vicky Wuwung. "RX-320 Rocket Static Pressure Combustion Chamber Prediction and Validation by Using Invers Method." Jurnal Teknologi Dirgantara 16, no. 1 (2018): 45. http://dx.doi.org/10.30536/j.jtd.2018.v16.a2866.

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The static pressure data of the combustion chamber which can generally be obtained by performing direct measurements when static test is performed on the rocket is an important parameter in predicting the thrust and design of the combustion chamber of the rocket. However, there is a model rocket for flight test that is used in static test. Thus, there is no mounting for static pressure sensors (for measurement) are made. To solve the problem, then the inverse method is used as an iterative solution for the basic equations of the rocket thrust force in the nozzle by guessing the value of the st
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Rajesh, T. N., T. J. S. Jothi, and T. Jayachandran. "Preliminary Studies on Non-Reactive Flow Vortex Cooling." Recent Patents on Mechanical Engineering 12, no. 3 (2019): 262–71. http://dx.doi.org/10.2174/2212797612666190510115403.

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Background: The impulse for the propulsion of a rocket engine is obtained from the combustion of propellant mixture inside the combustion chamber and as the plume exhausts through a convergent- divergent nozzle. At stoichiometric ratio, the temperature inside the combustion chamber can be as high as 3500K. Thus, effective cooling of the thrust chamber becomes an essential criterion while designing a rocket engine. Objective: A new cooling method of thrust chambers was introduced by Chiaverni, which is termed as Vortex Combustion Cold-Wall Chamber (VCCW). The patent works on cyclone separators
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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 (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 geomet
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Kim, Hanseup, Khalil Najafi, and Luis P. Bernal. "Helmholtz Resonance Based Micro Electrostatic Actuators for Compressible Gas Control: A Microjet Generator and a Gas Micro Pump." Journal of Microelectronics and Electronic Packaging 7, no. 1 (2010): 1–9. http://dx.doi.org/10.4071/1551-4897-7.1.1.

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This paper reports Helmholtz-resonance based micro electrostatic actuators to control compressible gaseous fluids in the micro scale. Particularly, it discusses design, fabrication, and testing results of two electrostatic actuators: a micro jet generator and an integrated peristaltic multistage micro pump. These electrostatic actuators vibrate a micro membrane in a micro chamber at a high frequency (>10 kHz), and easily induce the resonant behavior of compressible gases in the chamber. Such resonant behavior, often called the Helmholtz resonance, can repeatedly create instantaneous pre
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Niino, M., A. Kumakawa, R. Watanabe, and Y. Doi. "Evaluation of cold isostatic pressing of high-pressure thrust chamber closeout." Journal of Propulsion and Power 2, no. 1 (1986): 25–30. http://dx.doi.org/10.2514/3.22841.

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Dissertations / Theses on the topic "Thrust chamber pressure"

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Weber, Fabian. "Optical Analysis of the Hydrogen Cooling Film in High Pressure Combustion Chambers." Thesis, Luleå tekniska universitet, Rymdteknik, 2019. http://urn.kb.se/resolve?urn=urn:nbn:se:ltu:diva-76872.

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For performance optimisation of modern liquid cryogenic bipropellant rocket combustion chambers, one component which plays an important role in reducing the wall side heat flux, is the behaviour of the cooling film. At the Institute of Space Propulsion of the German Aerospace Center (DLR) in Lampoldshausen, hot test runs have been performed using the experimental combustion chamber BKM, to investigate the wall side heat flux which is -- among other factors -- dependent on cooling film properties. To gain more insight into the film behaviour under real rocket-like conditions, optical diagnostic
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Books on the topic "Thrust chamber pressure"

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Kumakawa, Akinaga. Characteristics of heat transfer to nickel plated chamber walls of high pressure rocket combustors. National Aerospace Laboratory, 1991.

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2

Jankowsky, Robert S. Experimental performance of a high-area-ratio rocket nozzle at high combustion chamber pressure. National Aeronautics and Space Administration, Office of Management, Scientific and Technical Information Program, 1996.

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Smith, Tamara A. Comparison of theoretical and experimental thrust performance of a 1030:1 area ratio rocket nozzle at a chamber pressure of 2413 kN/m(2) (350 psia). Lewis Research Center, 1987.

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4

George C. Marshall Space Flight Center., ed. Pressure fed thrust chamber technology: Test plan. Aerojet Propulsion Division, 1990.

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Pressure fed thrust chamber technology program: Contract number NAS8-37365 final report. Aerojet Propulsion Division, 1992.

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United States. National Aeronautics and Space Administration., ed. Pressure fed thrust chamber technology program: Contract NAS 8-37365, final report. National Aeronautics and Space Administration, 1992.

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7

M, Kazaroff John, Pavli Albert J, and United States. National Aeronautics and Space Administration. Scientific and Technical Information Program, eds. Experimental performance of a high-area-ratio rocket nozzle at high combustion chamber pressure. National Aeronautics and Space Administration, Office of Management, Scientific and Technical Information Program, 1996.

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M, Kazaroff John, Pavli Albert J, and United States. National Aeronautics and Space Administration. Scientific and Technical Information Program., eds. Experimental performance of a high-area-ratio rocket nozzle at high combustion chamber pressure. National Aeronautics and Space Administration, Office of Management, Scientific and Technical Information Program, 1996.

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M, Kazaroff John, Pavli Albert J, and United States. National Aeronautics and Space Administration. Scientific and Technical Information Program., eds. Experimental performance of a high-area-ratio rocket nozzle at high combustion chamber pressure. National Aeronautics and Space Administration, Office of Management, Scientific and Technical Information Program, 1996.

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10

A performance comparison of two small rocket nozzles. National Aeronautics and Space Administration, 1996.

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Book chapters on the topic "Thrust chamber pressure"

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Chemnitz, Alexander, and Thomas Sattelmayer. "Calculation of the Thermoacoustic Stability of a Main Stage Thrust Chamber Demonstrator." In Notes on Numerical Fluid Mechanics and Multidisciplinary Design. Springer International Publishing, 2020. http://dx.doi.org/10.1007/978-3-030-53847-7_15.

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Abstract The stability behavior of a virtual thrust chamber demonstrator with low injection pressure loss is studied numerically. The approach relies on an eigenvalue analysis of the Linearized Euler Equations. An updated form of the stability prediction procedure is outlined, addressing mean flow and flame response calculations. The acoustics of the isolated oxidizer dome are discussed as well as the complete system incorporating dome and combustion chamber. The coupling between both components is realized via a scattering matrix representing the injectors. A flame transfer function is applied to determine the damping rates. Thereby it is found that the procedure for the extraction of the flame transfer function from the CFD solution has a significant impact on the stability predictions.
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Arun Kumar, P., C. Rajeev Senan, B. Ajith, and Aishwarya Shankhdhar. "Thrust Prediction Model for Varying Chamber Pressure for a Hypergolic Bipropellant Liquid Rocket Engine." In Lecture Notes in Mechanical Engineering. Springer Singapore, 2018. http://dx.doi.org/10.1007/978-981-13-2697-4_52.

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Mejia, Guilherme Lourenço. "Solid Rocket Motor Internal Ballistics Simulation Considering Complex 3D Propellant Grain Geometries." In Energetic Materials Research, Applications, and New Technologies. IGI Global, 2018. http://dx.doi.org/10.4018/978-1-5225-2903-3.ch007.

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Solid rocket motors (SRM) are extensively employed in satellite launchers, missiles and gas generators. Design considers propulsive parameters with dimensional, manufacture, thermal and structural constraints. Solid propellant geometry and computation of its burning rate are essential for the calculation of pressure and thrust vs time curves. The propellant grain geometry changes during SRM burning are also important for structural integrity and analysis. A computational tool for tracking the propagation of tridimensional interfaces and shapes is then necessary. In this sense, the objective of this work is to present the developed computational tool (named RSIM) to simulate the burning surface regression during the combustion process of a solid propellant. The SRM internal ballistics simulation is based on 3D propagation, using the level set method approach. Geometrical and thermodynamic data are used as input for the computation, while simulation results of geometry and chamber pressure versus time are presented in test cases.
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"CARS Measurements at High Pressure in Cryogenic LOX/GH2 Jet Flames." In Liquid Rocket Thrust Chambers. American Institute of Aeronautics and Astronautics, 2004. http://dx.doi.org/10.2514/5.9781600866760.0369.0404.

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Kumakawa, Akinaga, Nobuyuki Yatsuyanagi, and Hiroshi Sakamoto. "ZrO2/Ni composite plating for high pressure thrust chambers." In Advanced Materials '93. Elsevier, 1994. http://dx.doi.org/10.1016/b978-0-444-81991-8.50083-2.

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6

"coating layer itself, an d at the interface between the coating and the substrate, causes instant fracturing and separation of coating material from the surface. In general, if a coating or contaminant is CHEMICALLY bonded to a surface, dry ice particle blasting will NOT effectively remove the coating. If the bond is PHYSICAL o r MECHANICAL in nature, such as a coating of rubber residue which is "anchored" into the porous surface of an aluminum casting, then there is a good chance that dr y ice blasting will work. Contaminants which are etched, or stained into the surfaces of metals, ceramics, plastics, or other materials typically cannot be removed with dry ice blasting. If the surface of the substrate is extremely porous or rough, providing strong mechanical "anchoring" for the contaminant or coating, dr y ice blasting may not be able to remove all of the coating, or the rate of removal may be too slow to allow dry ice blasting to be cost effective. The classic example of a contaminant that does NOT respond to dry ice blast-ing is RUST. Rust is both chemically and strongly mechanically bonded to steel substrate. Advanced stages of rust must be "chiseled" away with abrasive sand blasting. Only the thin film of powderized "flash" rust on a fresh steel surface can be effectively removed with dry ice blasting. 4.2.1.1. Inductio n (venturi) and direct acceleration blast systems - the effect of the typ e of system on available kinetic energy In a two-hose induction (venturi) carbon dioxide blastin g system, the medium particles are moved from the hopper to the "gun" chamber by suction, where they drop to a very low velocity before being induced into the outflow of the nozzle by a large flow volume of compressed air. Some more advanced two-hose systems employ a small positive pressure to the pellet delivery hose. In any type of two-hose system, since the blast medium particles have only a short distance in which to gain momentum and accelerate to the nozzle exit (usually only 200 to 300 mm), the final particle average velocity is limited to between 60 and 120 meters per second. So, in general, two-hose systems, although not so costly, are limited in their ability to deliver contaminant removal kinetic energy to the surface to be cleaned. When more blasting energy is required, these systems must be "boosted" a t the expense of much more air volume required, and higher blast pressure is re-quired as well, with much more nozzle back thrust, and very much more blast noise generated at the nozzle exit plane. The other type of solid carbon dioxide medium blasting system is like the "pressurized pot" abrasive blasting system common in the sand blasting and Plas-ti c Media Blasting industries. These systems use a single delivery hose from the hopper to the "nozzle" applicator in which both the medium particles and the compressed air travel. These systems are more complex and a little more costly than the inductive two-hose systems, but the advantages gained greatly outweigh the extra initial expense. In a single-hose solid carbon dioxide particle blasting system, sometimes referred to as a "direct acceleration " system, the medium is introduced from the hopper into a single, pre-pressurized blast hose through a sealed airlock feeder. The particles begin their acceleration and velocity increase." In Surface Contamination and Cleaning. CRC Press, 2003. http://dx.doi.org/10.1201/9789047403289-25.

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Conference papers on the topic "Thrust chamber pressure"

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Manikanda Kumaran, RajaGopal, Thirumalachari Sundararajan, and D. Raja Manohar. "Pressure Variation in Thrust Chamber During High Altitude Simulation." In 47th AIAA Aerospace Sciences Meeting including The New Horizons Forum and Aerospace Exposition. American Institute of Aeronautics and Astronautics, 2009. http://dx.doi.org/10.2514/6.2009-798.

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Ji, Jialong, and Bing Sun. "Research on Structural Optimization for Regenerative-Cooling Thrust Chamber." In ASME/JSME/KSME 2015 Joint Fluids Engineering Conference. American Society of Mechanical Engineers, 2015. http://dx.doi.org/10.1115/ajkfluids2015-09332.

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A regenerative-cooling liquid rocket engine chamber is applied as the study object, and the heat transfer models are established to get the distributions of chamber parameters. A thermal-fatigue simulation is carried out go get the strain of the throat section with the thermal field and pressure field applied as the boundary conditions of a two-dimensional finite-element model. A usage factor is used for the life estimation of chamber, and the structural parameters of cooling channels are optimized aiming for a smallest usage factor which means a longest life. The influence of parameter change
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Armbruster, Wolfgang, Justin Hardi, Dimitry Suslov, and Michael Oschwald. "High-Speed Flame Radiation Imaging of Thermoacoustic Coupling in a High Pressure Research Thrust Chamber." In 2018 Joint Propulsion Conference. American Institute of Aeronautics and Astronautics, 2018. http://dx.doi.org/10.2514/6.2018-4951.

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Lefor, Dominik, Jan Kowalski, Boris Kutschelis, Thomas Herbers, and Ronald Mailach. "Optimization of Axial Thrust Balancing Swirl Breakers in a Centrifugal Pump Using Stochastic Methods." In ASME 2014 4th Joint US-European Fluids Engineering Division Summer Meeting collocated with the ASME 2014 12th International Conference on Nanochannels, Microchannels, and Minichannels. American Society of Mechanical Engineers, 2014. http://dx.doi.org/10.1115/fedsm2014-21262.

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Hydraulic axial thrust balancing in single-stage radial pumps is a frequently applied procedure to reduce the bearing load to a reasonable level. It leads to considerable efficiency loss, which gives reason to investigate the optimization potential of the common balancing methods. This paper’s focus is on so-called casing ribs, which are used in the default design of an examined industrial pump. Radial vanes on the casing wall of the front impeller side chamber work as swirl breakers by decreasing rotating flow whereby the static pressure at the shroud increases and counteracts the resulting a
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Choudhuri, Ahsan R., Benjamin Baird, and S. R. Gollahalli. "A Numerical and Experimental Study of a Microthruster Performance." In ASME 2001 Engineering Technology Conference on Energy. American Society of Mechanical Engineers, 2001. http://dx.doi.org/10.1115/etce2001-17015.

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Abstract Performance and flow characteristics of a conical microthruster operated in cold thruster mode are presented. Nitrogen, Argon and Carbon dioxide were used as the medium. The performance parameters (specific impulse, impulse efficiency, thrust coefficient and characteristic velocity) are measured. The chamber pressure was varied from 2.8 bar to 12 bar. It is found that the high viscous loss severely degrades the performance of the thruster at a pressure below 5 bar. However, the mass flow remains unaffected even at a low pressure. The computed Mach number shows that at low pressures fl
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Shimura, Takashi, Satoshi Kawasaki, Masaharu Uchiumi, Toshiya Kimura, Mitsuaki Hayashi, and Jun Matsui. "Stability of an Axial-Thrust Self-Balancing System." In ASME 2012 Fluids Engineering Division Summer Meeting collocated with the ASME 2012 Heat Transfer Summer Conference and the ASME 2012 10th International Conference on Nanochannels, Microchannels, and Minichannels. American Society of Mechanical Engineers, 2012. http://dx.doi.org/10.1115/fedsm2012-72196.

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Rocket pump is characterized by high speed and high delivery pressure. Therefore, balancing of axial thrust acting on the rotor assembly is one of the most important factors. To realize complete axial thrust balancing, a balance piston-type axial-thrust self-balancing system is often used in rocket pumps. Such a system is comprised of an inlet orifice (#1) located at the outlet part of the impeller, outlet orifice (#2) located at the small-radius position of the back shroud and a chamber between these two orifices. Those orifices made by edges of the casing and the impeller shroud look like ri
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Gori, Fabio, Riccardo Pecorari, and Marco Mastrapasqua. "Numerical Simulation of the Coupling Between Vortex Shedding and Acoustic Field in a Solid Propellant Engine." In ASME 2003 International Mechanical Engineering Congress and Exposition. ASMEDC, 2003. http://dx.doi.org/10.1115/imece2003-42680.

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The paper investigates the numerical simulation of vortex shedding in the flow field of solid-propellant rocket motors. This phenomenon, resulting from the strong coupling between shear-layer instability and acoustic waves in the chamber, produces thrust and pressure oscillations. Numerical simulations are performed on the combustion chamber of the Ariane 5 MPS P230 (Solid Rocket Motor) with the commercial code Fluent CFD for conditions corresponding to 89 s of combustion time. The objective of the study is to reproduce the pressure oscillations frequencies and magnitudes, to compare the avail
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Nageswara Reddy, Pereddy. "Theoretical Assessment of Turbocharged Pulse Detonation Engine Performance at Various Flight Conditions." In ASME Turbo Expo 2020: Turbomachinery Technical Conference and Exposition. American Society of Mechanical Engineers, 2020. http://dx.doi.org/10.1115/gt2020-14451.

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Abstract A typical Pulse Detonation Engine (PDE) cycle of operation includes three basic processes: initiation and propagation of detonation wave in the Detonation Chamber (DC); a quasi-steady exhaust of detonation products from the DC at varying pressure through the supersonic nozzle; and a steady exhaust of remained detonation products at constant pressure through the nozzle while filling the DC with fresh air. In the present work, a novel method of Turbo-charging is proposed to increase the inlet pressure/density of fresh air fed into the DC in each cycle so as to increase the thrust develo
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Suslov, D. I., J. S. Hardi, B. Knapp, and M. Oschwald. "Hot-fire testing of liquid oxygen/hydrogen single coaxial injector at high-pressure conditions with optical diagnostics." In Progress in Propulsion Physics – Volume 11. EDP Sciences, 2019. http://dx.doi.org/10.1051/eucass/201911391.

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Injector behavior is of utmost importance for the performance and stability of liquid rocket engines (LREs). A major problem is getting a highly efficient homogeneous mixture and effective chemical reaction of fuels at minimum chamber length. Despite substantial progress in numerical simulations, a need for experimental data at representative conditions for development and validation of numerical design tools still exists. Therefore, in the framework of the DLR-project “ProTau,” the authors have performed tests to create an extended data base for numerical tool validation for high-pressure liq
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Pace, Raymond M., and Jason Ritzel. "Elimination of BWR Mark I Program Primary Containment Drywell-to-Wetwell Differential Pressure." In ASME 2020 Pressure Vessels & Piping Conference. American Society of Mechanical Engineers, 2020. http://dx.doi.org/10.1115/pvp2020-21001.

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Abstract The Boiling Water Reactor (BWR) Mark I (Mk I) Containment Program was implemented by the Owner’s Group circa 1970 to mitigate the effects of small, intermediate and design bases, Loss of Coolant Accident (LOCA), pipe break accidents on the torus, supports, internal structures and attached piping [1 & 2]. One of the significant mitigation attributes implemented during the Mk I Program is Normal Operation (NO) with a drywell-to-wetwell Differential Pressure (dP) maintained by the plant Nitrogen Make-up System used to inert the Primary Containment (PC) [3]. The differential pressure
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