Готові списки джерел за темами / Rocket engine nozzle

Добірка наукової літератури з теми "Rocket engine nozzle"

Автор: Grafiati

Опубліковано: 14 березня 2023

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Статті в журналах з теми "Rocket engine nozzle"

Strelnikov, G. A., A. D. Yhnatev, N. S. Pryadko, and S. S. Vasyliv. "Gas flow control in rocket engines." Technical mechanics 2021, no. 2 (June 29, 2021): 60–77. http://dx.doi.org/10.15407/itm2021.02.060.

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Анотація:

In the new conditions of application of launch vehicle boosters, space tugs, etc., modern rocket engines often do not satisfy the current stringent requirements. This calls for fundamental research into processes in rocket engines for improving their efficiency. In this regard, for the past 5 years, the Department of Thermogas Dynamics of Power Plants of the Institute of Technical Mechanics of the National Academy of Sciences of Ukraine and the State Space Agency of Ukraine has conducted research on gas flow control in rocket engines to improve their efficiency and functionality. Mechanisms of flow perturbation in the nozzle of a rocket engine by liquid injection and a solid obstacle were investigated. A mathematical model of supersonic flow perturbation by local liquid injection was refined, and new solutions for increasing the energy release rate of the liquid were developed. A numerical simulation of a gas flow perturbed by a solid obstacle in the nozzle of a rocket engine made it possible to verify the known (mostly experimental) results and to reveal new perturbation features. In particular, a significant increase in the efficiency of flow perturbation by an obstacle in the transonic region was shown up, and some dependences involving the distribution of the perturbed pressure on the nozzle wall, which had been considered universal, were refined. The possibility of increasing the efficiency of use of the generator gas picked downstream of the turbine of a liquid-propellant rocket engine was investigated, and the advantages of a new scheme of gas injection into the supersonic part of the nozzle, which provides both nozzle wall cooling by the generator gas and the production of lateral control forces, were substantiated. A new concept of rocket engine thrust vector control was developed: a combination of a mechanical and a gas-dynamic system. It was shown that such a thrust vector control system allows one to increase the efficiency and reliability of the space rocket stage flight control system. A new liquid-propellant rocket engine scheme was developed to control both the thrust amount and the thrust vector direction in all planes of rocket stage flight stabilization. New approaches to the process organization in auxiliary elements of rocket engines on the basis of detonation propellant combustion were developed to increase the rocket engine performance.

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Jéger, Csaba, and Árpád Veress. "Novell Application of CFD for Rocket Engine Nozzle Optimization." Periodica Polytechnica Transportation Engineering 47, no. 2 (January 10, 2018): 131–35. http://dx.doi.org/10.3311/pptr.11490.

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Numerical analyses, validation and geometric optimization of a converging-diverging nozzle flows has been established in the present work. The optimal nozzle contour for a given nozzle pressure ratio and length yields the largest obtainable thrust for the conditions and thus minimises the losses. Application of such methods reduces the entry cost to the market, promote innovation and accelerate the development processes. A parametric geometry, numerical mesh and simulation model is constructed first to solve the problem. The simulation model is then validated by using experimental and computational data. The optimizations are completed for conical and bell shaped nozzles also to find the suitable nozzle geometries for the given conditions. Results are in good agreement with existing nozzle flow fields. The optimization loop described and implemented here can be used in the all similar situations and can be the basis of an improved nozzle geometry optimization procedure by means of using a multiphysics system to generate the final model with reduced number sampling phases.

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Guram, Sejal, Vidhanshu Jadhav, Prasad Sawant, and Ankit Kumar Mishra. "Review Study on Thermal Characteristics of Bell Nozzle used in Supersonic Engine." 1 2, no. 1 (March 1, 2023): 4–14. http://dx.doi.org/10.46632/jame/2/1/2.

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The nozzle is an important component of the rocket motor system, and a rocket’s overall performance is highly reliant on its aerodynamic design. The nozzle contour can be meticulously shaped to improve performance significantly. The design and shape of rocket nozzles have evolved over the last several decades as a result of extensive research. The nozzle design is composed of two components, an integrated throat, an entry and an exit cone, and a thermal protection system. The Bell Nozzle is designed to provide clearance space for placing the ITE and exit cone, with a cone inflection angle of 16 and a thermal protection system. This paper intends to review and summarize all such developments. Small-scale engine testing allows for the analysis of rocket nozzle materials, but the history of nozzle surface temperature and thermal stress may be adversely affected by side effects. The review focuses primarily on the nozzle shape which has the largest radiative flux past the neck, but the nozzle shape has the highest heat flux in the throat due to the mass-flow rate per unit area. The distribution of nozzle wall pressure is strongly influenced by the Mach number of the injected secondary flow, leading to undesirable side loads. Finally, future development possibilities are suggested.

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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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Bogoi, Alina, Radu D. Rugescu, Valentin Ionut Misirliu, Florin Radu Bacaran, and Mihai Predoiu. "Inviscid Nozzle for Aerospike Rocket Engine Application." Applied Mechanics and Materials 811 (November 2015): 152–56. http://dx.doi.org/10.4028/www.scientific.net/amm.811.152.

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A computational method for the steady 2-D flow in axially symmetrical rocket nozzles with a given profile is developed, in order to determine the Maximum thrust contour of rocket engine nozzles with large expansion ratio. The optimized nozzles proved a more than 10% increase in the integral specific impulse recorded during the variable altitude atmospheric flight of rocket vehicles. The method is well suited for application in the design of the optimum contour for axially-symmetric nozzles for atmospheric rocket ascent, specifically for aerospike type nozzles, as for other similar industrial applications in gas and steam turbine technology.

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Sultanov, T. S., та G. A. Glebov. "Numerical Computation of Specific Impulse and Internal Flow Parameters in Solid Fuel Rocket Motors with Two-Phase Сombustion Products". Herald of the Bauman Moscow State Technical University. Series Mechanical Engineering, № 3 (138) (вересень 2021): 98–107. http://dx.doi.org/10.18698/0236-3941-2021-3-98-107.

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Eulerian --- Lagrangian method was used in the Fluent computational fluid dynamics system to calculate motion of the two-phase combustion products in the solid fuel rocket motor combustion chamber and nozzle. Condensed phase is assumed to consist of spherical particles with the same diameter, which dimensions are not changing along the motion trajectory. Flows with particle diameters of 3, 5, 7, 9, and 11 μm were investigated. Four versions of the engine combustion chamber configuration were examined: with slotted and smooth cylindrical charge channels, each with external and submerged nozzles. Gas flow and particle trajectories were calculated starting from the solid fuel surface and to the nozzle exit. Volumetric fields of particle concentrations, condensed phase velocities and temperatures, as well as turbulence degree in the solid propellant rocket engine flow duct were obtained. Values of particles velocity and temperature lag from the gas phase along the nozzle length were received. Influence of the charge channel shape, degree of the nozzle submersion and of the condensate particles size on the solid propellant rocket engine specific impulse were determined, and losses were estimated in comparison with the case of ideal flow

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Bruce Ralphin Rose, J., and J. Veni Grace. "Performance analysis of lobed nozzle ejectors for high altitude simulation of rocket engines." International Journal of Modeling, Simulation, and Scientific Computing 05, no. 04 (September 29, 2014): 1450019. http://dx.doi.org/10.1142/s1793962314500196.

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Ejectors are used in high altitude testing of rocket engines to create vacuum for simulating the engine test in vacuum conditions. The performance of an ejector plays a vital role in creating vacuum at the exit of the engine nozzle and the nozzle design exit pressure at the time of ignition. Consequently, the performance of ejectors has to be improved to reduce the consumption of active fluid. In this investigation, the performance of an ejector has been improved by changing the exit shear plane of the nozzle. Conventionally, conical nozzles are used for creating the required momentum. Lobes of 4 no's, 6 no's and 8 numbers for an equivalent area ratio = 5.88 are used to increase the shear area. The influence of shear plane variation in the suction pressure is studied by a detailed CFD analysis.

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Shustov, S. A., I. E. Ivanov, and I. A. Kryukov. "Numerical study of the separation of a turbulent boundary in rocket engine nozzles with an optimized supersonic part." Journal of Physics: Conference Series 2308, no. 1 (July 1, 2022): 012015. http://dx.doi.org/10.1088/1742-6596/2308/1/012015.

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Abstract At present, propulsion systems are being developed for flights of aerospace vehicles both in the Earth’s atmosphere and beyond. At the stage of the propulsion system design, it is necessary to be able to reliably determine the energy and thrust characteristics of rocket engines in regimes with flow separation in the nozzle for a wide range of nozzle pressure ratio (NPR) (1 ≤ m ≤ 30, m = ph /pa , p—pressure, indices h and a refer, respectively, to the parameters of the external environment and at the nozzle exit) characterizing the flow in nozzles and jets. In this case, the nozzle, as a rule, has a thrust optimized contour (TOC) supersonic part. In this regard, we present some results of a numerical study of flows with separation of the turbulent boundary layer in the TOC nozzles of main rocket engines in the above range of the NPR based on the Reynolds Averaged Navier–Stokes (RANS) system of equations.

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Vasyliv, S. S., and H. O. Strelnykov. "Rocket engine thrust vector control by detonation product injection into the supersonic portion of the nozzle." Technical mechanics 2020, no. 4 (December 10, 2020): 29–34. http://dx.doi.org/10.15407/itm2020.04.029.

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For solving non-traditional problems of rocket flight control, in particular, for the conditions of impact of a nuclear explosion, non-traditional approaches to the organization of the thrust vector control of a rocket engine are required. Various schemes of gas-dynamic thrust vector control systems that counteract impact actions on the rocket were studied. It was found that the dynamic characteristics of traditional gas-dynamic thrust vector control systems do not allow one to solve the problem of counteracting impact actions on the rocket. Appropriate dynamic characteristics can provide a perturbation of the supersonic flow by injecting into the nozzle the detonation products with the main shock wave propagating in the supersonic flow. This way to perturb the supersonic flow in a rocket engine nozzle is investigated in this paper. In order to identify the principles of producing control forces and provide a perturbation of the supersonic flow by injecting into the nozzle the detonation products with the main shock wave propagating in the supersonic flow, a computer simulation of the nozzle flow was performed. The nozzle of the 11D25 engine developed by Yuzhnoye State Design Office and used in the third stage of the Cyclone-3 launch vehicle was taken as a basis. The thrust vector control scheme relies on the use of the main fuel component detonation. The evolution of the detonation wave in the supersonic flow of the combustion chamber nozzle was simulated numerically. According to the nature of the perturbation propagation in the nozzle, the lateral force from the perturbation has an alternating character with the perturbation stabilization in sign and magnitude when approaching the critical nozzle section. The value of the relative lateral force is sufficient for counteracting large disturbing moments of short duration. Thus, the force factors that can be used to control the rocket engine thrust vector are identified. Further research should focus on finding the optimal location of the detonation product injection in order to prevent mutual compensation of force factors.

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Kumar, S. Senthil, and M. Arularasu. "Advanced Computational Flow Analysis - Rocket Engine Nozzle." Asian Journal of Research in Social Sciences and Humanities 6, no. 11 (2016): 1219. http://dx.doi.org/10.5958/2249-7315.2016.01265.x.

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Дисертації з теми "Rocket engine nozzle"

Östlund, Jan. "Supersonic flow separation with application to rocket engine nozzles." Doctoral thesis, KTH, Mechanics, 2004. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-3793.

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The increasing demand for higher performance in rocketlaunchers promotes the development of nozzles with higherperformance, which basically is achieved by increasing theexpansion ratio. However, this may lead to flow separation andensuing instationary, asymmetric forces, so-called side-loads,which may present life-limiting constraints on both the nozzleitself and other engine components. Substantial gains can bemade in the engine performance if this problem can be overcome,and hence different methods of separation control have beensuggested. However, none has so far been implemented in fullscale, due to the uncertainties involved in modeling andpredicting the flow phenomena involved.

In the present work the causes of unsteady and unsymmetricalflow separation and resulting side-loads in rocket enginenozzles are investigated. This involves the use of acombination of analytical, numerical and experimental methods,which all are presented in the thesis. A main part of the workis based on sub-scale testing of model nozzles operated withair. Hence, aspects on how to design sub-scale models that areable to capture the relevant physics of full-scale rocketengine nozzles are highlighted. Scaling laws like thosepresented in here are indispensable for extracting side-loadcorrelations from sub-scale tests and applying them tofull-scale nozzles.

Three main types of side-load mechanisms have been observedin the test campaigns, due to: (i) intermittent and randompressure fluctuations, (ii) transition in separation patternand (iii) aeroelastic coupling. All these three types aredescribed and exemplified by test results together withanalysis. A comprehensive, up-to-date review of supersonic flowseparation and side-loads in internal nozzle flows is givenwith an in-depth discussion of different approaches forpredicting the phenomena. This includes methods for predictingshock-induced separation, models for predicting side-loadlevels and aeroelastic coupling effects. Examples are presentedto illustrate the status of various methods, and theiradvantages and shortcomings are discussed.

A major part of the thesis focus on the fundamentalshock-wave turbulent boundary layer interaction (SWTBLI) and aphysical description of the phenomenon is given. Thisdescription is based on theoretical concepts, computationalresults and experimental observation, where, however, emphasisis placed on the rocket-engineering perspective. This workconnects the industrial development of rocket engine nozzles tothe fundamental research of the SWTBLI phenomenon and shows howthese research results can be utilized in real applications.The thesis is concluded with remarks on active and passive flowcontrol in rocket nozzles and directions of futureresearch.

The present work was performed at VAC's Space PropulsionDivision within the framework of European spacecooperation.

Keywords:turbulent, boundary layer, shock wave,interaction, overexpanded,rocket nozzle, flow separation,control, side-load, experiments, models, review.

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Östlund, Jan. "Flow Processes in Rocket Engine Nozzles with Focus on Flow Separation and Side-Loads." Licentiate thesis, KTH, Mechanics, 2002. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-1452.

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Wahlström, Dennis. "Probabilistic Multidisciplinary Design Optimization on a high-pressure sandwich wall in a rocket engine application." Thesis, Umeå universitet, Institutionen för fysik, 2017. http://urn.kb.se/resolve?urn=urn:nbn:se:umu:diva-138480.

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A need to find better achievement has always been required in the space industrythrough time. Advanced technologies are provided to accomplish goals for humanityfor space explorer and space missions, to apprehend answers and widen knowledges. These are the goals of improvement, and in this thesis, is to strive and demandto understand and improve the mass of a space nozzle, utilized in an upperstage of space mission, with an expander cycle engine. The study is carried out by creating design of experiment using Latin HypercubeSampling (LHS) with a consideration to number of design and simulation expense.A surrogate model based optimization with Multidisciplinary Design Optimization(MDO) method for two different approaches, Analytical Target Cascading (ATC) and Multidisciplinary Feasible (MDF) are used for comparison and emend the conclusion. In the optimization, three different limitations are being investigated, designspace limit, industrial limit and industrial limit with tolerance. Optimized results have shown an incompatibility between two optimization approaches, ATC and MDF which are expected to be similar, but for the two limitations, design space limit and industrial limit appear to be less agreeable. The ATC formalist in this case dictates by the main objective, where the children/subproblems only focus to find a solution that satisfies the main objective and its constraint. For the MDF, the main objective function is described as a single function and solved subject to all the constraints. Furthermore, the problem is not divided into subproblems as in the ATC. Surrogate model based optimization, its solution influences by the accuracy ofthe model, and this is being investigated with another DoE. A DoE of the full factorial analysis is created and selected to study in a region near the optimal solution.In such region, the result has evidently shown to be quite accurate for almost allthe surrogate models, except for max temperature, damage and strain at the hottestregion, with the largest common impact on inner wall thickness of the space nozzle. Results of the new structure of the space nozzle have shown an improvement of mass by ≈ 50%, ≈ 15% and ≈ -4%, for the three different limitations, design spacelimit, industrial limit and industrial limit with tolerance, relative to a reference value,and ≈ 10%, ≈ 35% and ≈ 25% cheaper to manufacture accordingly to the defined producibility model.

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Bulut, Jane. "Design and CFD analysis of the demonstrator aerospike engine for a small satellite launcher application." Master's thesis, Alma Mater Studiorum - Università di Bologna, 2020.

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Starting with a brief overview of thrust generation for launchers, this study focuses on the design process of the demonstrator aerospike engine, DEMOP-1, of the Pangea Aerospace's commercial grade engine and its flow field analysis. The primary goal of the study is to obtain the plug nozzle design delivers 30 kN thrust using cryogenic liquid oxygen (LOX) as the oxidizer and cryogenic liquid methane (LCH4) as the fuel, with the mixture ratio of 3.4. Design parameters considered as 30 bar of combustion chamber pressure (Po) and expansion ratio as 15 for an optimum expanded nozzle. On the basis of decided design characteristics, Angelino's method is used to design the nozzle contour through MATLAB. The flow field over the aerospike analyzed using commercial CFD program FLUENT for sea level, optimum expansion and vacuum conditions. Flow simulations are carried out for air (specific heat ratio, gamma= 1.4), and afterwards based on the obtained thrust values at each altitude for air, expected thrust values for the real propellant, LOX/LCH4 (specific heat ratio, gamma = 1.1664), are calculated. Finally, the study is concluded with the comparison of trend in thrust and specific impulse for conventional bell nozzle and aerospike. For the conventional bell engine the values obtained in commercial computational simulation of chemical rocket propulsion and combustion software RPA for bell nozzle with same characteristics with aerospike, Po = 30 bar and expansion ratio = 15, are taken as reference for sea level, optimum expansion level and vacuum condition performance. Due to its ability to adopt the altitude, aerospike delivers higher performance at the low altitudes with respect to the conventional bell nozzle which has the same expansion ratio and combustion chamber pressure. Last in order but not in importance, after obtaining the flow field on plug of the aerospike, the shock wave impingement on the nozzle surface at sea level has been investigated.

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Garby, Romain. "Simulations of flame stabilization and stability in high-pressure propulsion systems." Phd thesis, Toulouse, INPT, 2013. http://oatao.univ-toulouse.fr/9706/1/garby.pdf.

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Denton, Brandon Lee. "Design and analysis of rocket nozzle contours for launching pico-satellites /." Online version of thesis, 2008. http://hdl.handle.net/1850/6003.

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(6927776), Alexis Joy Harroun. "Investigation of Nozzle Performance for Rotating Detonation Rocket Engines." Thesis, 2019.

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Progress in conventional rocket engine technologies, based on constant pressure combustion, has plateaued in the past few decades. Rotating detonation engines (RDEs) are of particular interest to the rocket propulsion community as pressure gain combustion may provide improvements to specific impulse relevant to booster applications. Despite recent significant investment in RDE technologies, little research has been conducted to date into the effect of nozzle design on rocket application RDEs. Proper nozzle design is critical to capturing the thrust potential of the transient pressure ratios produced by the thrust chamber. A computational fluid dynamics study was conducted based on hotfire conditions tested in the Purdue V1.3 RDE campaign. Three geometries were investigated: nozzleless/blunt body, internal-external expansion (IE-) aerospike, and flared aerospike. The computational study found the RDE's dynamic exhaust plume enhances the ejection physics beyond that of a typical high pressure device. For the nozzleless geometry, the base pressure was drawn down below constant pressure estimates, increasing the base drag on the engine. For the aerospike geometries, the occurrance of flow separation on the plug was delayed, which has ramifications on nozzle design for operation at a range of pressure altitudes. The flared aerospike design, which has the ability to achieve much higher area ratios, was shown to have potential performance benefits over the limited IE-aerospike geometry. A new test campaign with the Purdue RDE V1.4 was designed with instrumentation to capture static pressures on the nozzleless and aerospike surfaces. These results were used to validate the results from the computational study. The computational and experimental studies were used to identify new flow physics associated with a rocket RDE important to future nozzle design work. Future computational work is necessary to explore the effect of different parameters on the nozzle performance. More testing, including with an altitude simulation chamber, would help quantify the possible benefit of new aerospike nozzle designs, including the flared aerospike geometry.

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Книги з теми "Rocket engine nozzle"

United States. National Aeronautics and Space Administration., ed. Comparison of two procedures for predicting rocket engine nozzle performance. [Washington, DC]: National Aeronautics and Space Administration, 1987.

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Marable, R. W. Design, fabrication, and test of the RL10 derivative II chamber/primary nozzle. [West Palm Beach, Fla: Pratt and Whitney Aircraft, 1989.

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Leonard, Schoenman, and United States. National Aeronautics and Space Administration., eds. Advanced small rocket chambers option 3: 110 1bf Ir-Re rocket. [Washington, DC]: National Aeronautics and Space Administration, 1995.

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J, Sovie Amy, Haag Thomas W, and United States. National Aeronautics and Space Administration., eds. Arcjet nozzle design impacts. [Washington, DC]: National Aeronautics and Space Administration, 1989.

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Milton, Lamb, and United States. National Aeronautics and Space Administration. Scientific and Technical Information Division., eds. Aeropropulsive characteristics of isolated combined turbojet/ramjet nozzles at Mach numbers from 0 to 1.20. [Washington, D.C.]: National Aeronautics and Space Administration, Scientific and Technical Information Division, 1988.

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M, Kim Y., Shang H. M, and United States. National Aeronautics and Space Administration., eds. Turbulence modelling of flow fields in thrust chambers: Final technical report for the period June 10, 1991 through September 13, 1992. [Huntsville, Ala.]: Research Institute, the University of Alabama in Huntsville, 1993.

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J, Pavli Albert, Kacynski Kenneth J, and United States. National Aeronautics and Space Administration. Scientific and Technical Information Office., eds. Comparison of theoretical and experimental thrust performance of a 1030:1 area ratio rocket nozzle at a chamber pressure of 2413 kN/m℗ø(350 psia). [Washington, D.C.]: National Aeronautics and Space Administration, Scientific and Technical Information Office, 1987.

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George C. Marshall Space Flight Center., ed. Flight motor set 360H005 (STS-28R). Brigham City, UT: Thiokol Corp., Space Operations, 1990.

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United States. National Aeronautics and Space Administration., ed. Calculation of propulsive nozzle flowfields in multidiffusing chemically recating environments. [Washington, DC]: National Aeronautics and Space Administration, 1994.

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Частини книг з теми "Rocket engine nozzle"

Ludescher, Sandra, and Herbert Olivier. "Film Cooling in Rocket Nozzles." In Notes on Numerical Fluid Mechanics and Multidisciplinary Design, 65–78. Cham: Springer International Publishing, 2020. http://dx.doi.org/10.1007/978-3-030-53847-7_4.

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Abstract In this project supersonic, tangential film cooling in the expansion part of a nozzle with rocket-engine like hot gas conditions was investigated. Therefore, a parametric study in a conical nozzle was conducted revealing the most important influencing parameter on film cooling for the presented setup. Additionally, a new axisymmetric film cooling model and a method for calculating the cooling efficiency from experimental data was developed. These models lead to a satisfying correlation of the data. Furthermore, film cooling in a dual-bell nozzle performing in altitude mode was investigated. The aim of these experiments was to show the influence of different contour inflection geometries on the film cooling efficiency in the bell extension.

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Alamu, Samuel O., Marc J. Louise Caballes, Yulai Yang, Orlyse Mballa, and Guangming Chen. "3D Design and Manufacturing Analysis of Liquid Propellant Rocket Engine (LPRE) Nozzle." In Advances in Intelligent Systems and Computing, 968–80. Cham: Springer International Publishing, 2019. http://dx.doi.org/10.1007/978-3-030-32523-7_73.

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Barfusz, Oliver, Felix Hötte, Stefanie Reese, and Matthias Haupt. "Pseudo-transient 3D Conjugate Heat Transfer Simulation and Lifetime Prediction of a Rocket Combustion Chamber." In Notes on Numerical Fluid Mechanics and Multidisciplinary Design, 265–78. Cham: Springer International Publishing, 2020. http://dx.doi.org/10.1007/978-3-030-53847-7_17.

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Abstract Rocket engine nozzle structures typically fail after a few engine cycles due to the extreme thermomechanical loading near the nozzle throat. In order to obtain an accurate lifetime prediction and to increase the lifetime, a detailed understanding of the thermomechanical behavior and the acting loads is indispensable. The first part is devoted to a thermally coupled simulation (conjugate heat transfer) of a fatigue experiment. The simulation contains a thermal FEM model of the fatigue specimen structure, RANS simulations of nine cooling channel flows and a Flamelet-based RANS simulation of the hot gas flow. A pseudo-transient, implicit Dirichlet–Neumann scheme is utilized for the partitioned coupling. A comparison with the experiment shows a good agreement between the nodal temperatures and their corresponding thermocouple measurements. The second part consists of the lifetime prediction of the fatigue experiment utilizing a sequentially coupled thermomechanical analysis scheme. First, a transient thermal analysis is carried out to obtain the temperature field within the fatigue specimen. Afterwards, the computed temperature serves as input for a series of quasi-static mechanical analyses, in which a viscoplastic damage model is utilized. The evolution and progression of the damage variable within the regions of interest are thoroughly discussed. A comparison between simulation and experiment shows that the results are in good agreement. The crucial failure mode (doghouse effect) is captured very well.

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Sansica, A., J. Ch Robinet, Eric Goncalves, and J. Herpe. "Three-Dimensional Instability of Shock-Wave/Boundary-Layer Interaction for Rocket Engine Nozzle Applications." In 31st International Symposium on Shock Waves 2, 523–30. Cham: Springer International Publishing, 2019. http://dx.doi.org/10.1007/978-3-319-91017-8_67.

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Sansica, A., J. Ch Robinet, Eric Goncalves, and J. Herpe. "Correction to: Three-Dimensional Instability of Shock-Wave/Boundary-Layer Interaction for Rocket Engine Nozzle Applications." In 31st International Symposium on Shock Waves 2, C1. Cham: Springer International Publishing, 2019. http://dx.doi.org/10.1007/978-3-319-91017-8_148.

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Decher, Reiner. "More Components: Inlets, Mixers, and Nozzles." In The Vortex and The Jet, 137–54. Singapore: Springer Singapore, 2022. http://dx.doi.org/10.1007/978-981-16-8028-1_13.

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AbstractTheintegrationof a gas turbine engine into a functioning jet propulsion engine for an airplane requires more components: inlets and nozzles. For the inlet, the special care exercised to avoid ingestion of boundary layers air is described. The design features of nozzles are described and extended to include discussion of more extreme configurations such as those found on rocket engines.

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Morgenweck, Daniel, Jutta Pieringer, and Thomas Sattelmayer. "Numerical Determination of Nozzle Admittances in Rocket Engines." In Notes on Numerical Fluid Mechanics and Multidisciplinary Design, 579–86. Berlin, Heidelberg: Springer Berlin Heidelberg, 2010. http://dx.doi.org/10.1007/978-3-642-14243-7_71.

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"Rocket Engine Nozzle Concepts." In Liquid Rocket Thrust Chambers, 437–67. Reston ,VA: American Institute of Aeronautics and Astronautics, 2004. http://dx.doi.org/10.2514/5.9781600866760.0437.0467.

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Тези доповідей конференцій з теми "Rocket engine nozzle"

MCAMIS, R., D. LANKFORD, and W. PHARES. "Theoretical liquid rocket engine nozzle flow fields." In 28th Joint Propulsion Conference and Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1992. http://dx.doi.org/10.2514/6.1992-3730.

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Robles, Luis R., Johnny Ho, Bao Nguyen, Geoffrey Wagner, Jeremy Surmi, Khulood Faruqui, Ashley Carter, et al. "Conceptual Regenerative Nozzle Cooling Design for a Hydroxyl-Terminated Polybutadiene and Oxygen Hybrid Rocket Engine." In ASME 2017 International Design Engineering Technical Conferences and Computers and Information in Engineering Conference. American Society of Mechanical Engineers, 2017. http://dx.doi.org/10.1115/detc2017-68396.

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Regenerative rocket nozzle cooling technology is well developed for liquid fueled rocket engines, but the technology has yet to be widely applied to hybrid rockets. Liquid engines use fuel as coolant, and while the oxidizers typically used in hybrids are not as efficient at conducting heat, the increased renewability of a rocket using regenerative cycle should still make the technology attractive. Due to the high temperatures that permeate throughout a rocket nozzle, most nozzles are predisposed to ablation, supporting the need to implement a nozzle cooling system. This paper presents a proof-of-concept regenerative cooling system for a hybrid engine which uses hydroxyl-terminated polybutadiene (HTPB) as its solid fuel and gaseous oxygen (O2) as its oxidizer, whereby a portion of gaseous oxygen is injected directly into the combustion chamber and another portion is routed up through grooves on the exterior of a copper-chromium nozzle and, afterwards, injected into the combustion chamber. Using O2 as a coolant will significantly lower the temperature of the nozzle which will prevent ablation due to the high temperatures produced by the exhaust. Additional advantages are an increase in combustion efficiency due to the heated O2 being used for combustion and an increased overall efficiency from the regenerative cycle. A computational model is presented, and several experiments are performed using computational fluid dynamics (CFD).

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Amano, Ryoichi S., and Yi-Hsin Yen. "Design of Solid Rocket Engine." In ASME 2015 International Design Engineering Technical Conferences and Computers and Information in Engineering Conference. American Society of Mechanical Engineers, 2015. http://dx.doi.org/10.1115/detc2015-48092.

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This paper presents both experiment and simulation of alumina molten flow in a solid rocket motor (SRM), when the propellant combusts, the aluminum is oxidized into alumina (Al2O3) which, under the right flow conditions, tends to agglomerate into molten droplets, impinge on the chamber walls, and then flow along the nozzle wall. Such agglomerates can cause erosive damage. The goal of the present study is to characterize the agglomerate flow within the nozzle section by studying the breakup process of a liquid film that flows along the wall of a straight channel while a high-speed gas moves over it. We have used an unsteady-flow Reynolds-Averaged Navier-Stokes code (URANS) to investigate the interaction of the liquid film flow with the gas flow, and analyzed the breakup process for different flow conditions. The rate of the wave breakup was characterized by introducing a breakup-length-scale for various flow conditions based on the Volume Fraction (VF) of the liquid, which is an indicator of a two-phase flow liquid breakup level. A smaller breakup-length-scale means that smaller drops have been created during the breakup process. The study covers the breakup and fluid behaviors based on different gas-liquid momentum flux ratios, different surface tension and viscosity settings, different Ohnesorge numbers (Oh), and different Weber numbers. Both water and molten aluminum flows were considered in the simulation studies. The analysis demonstrates an effective method of correlating the liquid breakup with the main flow conditions in the nozzle channel path.

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Hall, Joshua, Carl Hartsfield, Joseph Simmons, and Richard Branam. "Optimized Dual-Expander Aerospike Nozzle Upper Stage Rocket Engine." In 49th AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2011. http://dx.doi.org/10.2514/6.2011-419.

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Manski, Detlef, and Gerald Hagemann. "Influence of rocket design parameters on engine nozzle efficiencies." In 30th Joint Propulsion Conference and Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1994. http://dx.doi.org/10.2514/6.1994-2756.

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Stewart, Kyle J., Periklis Papadopoulos, and Jordan Pollard. "Nuclear Thermal Rocket Engine with a Toroidal Aerospike Nozzle." In AIAA Propulsion and Energy 2020 Forum. Reston, Virginia: American Institute of Aeronautics and Astronautics, 2020. http://dx.doi.org/10.2514/6.2020-3841.

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Matveev, Valeriy, Vasilii Zubanov, Leonid Shabliy, and Anastasia Korneeva. "Optimization of Nozzle Shape of Hydrogen-Oxygen Rocket Engine." In 8th International Conference on Simulation and Modeling Methodologies, Technologies and Applications. SCITEPRESS - Science and Technology Publications, 2018. http://dx.doi.org/10.5220/0006890003650370.

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Kozlov, Alexander, Jose Hinckel, and Adalberto Comiran. "Investigation of a nozzle tap-off liquid rocket engine scheme." In 32nd Joint Propulsion Conference and Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1996. http://dx.doi.org/10.2514/6.1996-3118.

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Besnard, Eric, Hsun Hu Chen, Tom Mueller, and John Garvey. "Design, Manufacturing and Test of a Plug Nozzle Rocket Engine." In 38th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2002. http://dx.doi.org/10.2514/6.2002-4038.

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DAVIDIAN, KENNETH. "Comparison of two procedures for predicting rocket engine nozzle performance." In 23rd Joint Propulsion Conference. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1987. http://dx.doi.org/10.2514/6.1987-2071.

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