Статті в журналах: "Rocket engine nozzle"

1

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.

Повний текст джерела

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.

Стилі APA, Harvard, Vancouver, ISO та ін.

2

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.

Повний текст джерела

Анотація:

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.

Стилі APA, Harvard, Vancouver, ISO та ін.

3

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.

Повний текст джерела

Анотація:

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.

Стилі APA, Harvard, Vancouver, ISO та ін.

4

ZAGANESCU, Nicolae-Florin, Rodica ZAGANESCU, and Constantin-Marcian GHEORGHE. "Wernher Von Braun’s Pioneering Work in Modelling and Testing Liquid-Propellant Rockets." INCAS BULLETIN 14, no. 2 (June 10, 2022): 153–61. http://dx.doi.org/10.13111/2066-8201.2022.14.2.13.

Повний текст джерела

Анотація:

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).

Стилі APA, Harvard, Vancouver, ISO та ін.

5

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.

Повний текст джерела

Анотація:

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.

Стилі APA, Harvard, Vancouver, ISO та ін.

6

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.

Повний текст джерела

Анотація:

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

Стилі APA, Harvard, Vancouver, ISO та ін.

7

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.

Повний текст джерела

Анотація:

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.

Стилі APA, Harvard, Vancouver, ISO та ін.

8

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.

Повний текст джерела

Анотація:

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.

Стилі APA, Harvard, Vancouver, ISO та ін.

9

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.

Повний текст джерела

Анотація:

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.

Стилі APA, Harvard, Vancouver, ISO та ін.

10

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.

Повний текст джерела

Стилі APA, Harvard, Vancouver, ISO та ін.

11

Nae, Catalin, Irina-Carmen Andrei, Gabriela-Liliana Stroe, and Sorin Berbente. "Integration of Fuels Types and Chemical Properties with the Design of the Rocket Engine�s Bell Exhaust Nozzle and Combustion Chamber." Revista de Chimie 71, no. 1 (February 7, 2020): 436–44. http://dx.doi.org/10.37358/rc.20.1.7872.

Повний текст джерела

Анотація:

The chemical properties of the fuels are crucial for obtaining the numerical accuracy during the design and performance analysis in case of liquid fuel propelled rocket engines, as well as the trajectory optimization. In this paper, the research was primarly focused on optimizing the numerical accuracy for non-linear two-dimensional approximation the Fuel Combustion Charts; secondarily, the investigation was carried on the design of the bell-nozzle of a liquid propelled rocket engine, taking into account the variation of the coefficients which are significant for expressing the fuels chemical properties. From the Fuel Combustion Charts, the authors selected a the LOX - Kerosene combination for propelling the rocket engine, due to the most convenient matching with the technology and material specifications, safety and environmental friendly requirements; from the LOX-Kerosene Charts, the authors have originally developed a method to obtain the expression of a non-linear approximation function of two variables. The design of the bell shaped nozzle and combustion chamber for a liquid propelled rocket engine was included, in purpose to illustrate the link between the LPRE design and the fuels types and chemical properties.

Стилі APA, Harvard, Vancouver, ISO та ін.

12

O¨stlund, J., and B. Muhammad-Klingmann. "Supersonic Flow Separation with Application to Rocket Engine Nozzles." Applied Mechanics Reviews 58, no. 3 (May 1, 2005): 143–77. http://dx.doi.org/10.1115/1.1894402.

Повний текст джерела

Анотація:

The past decade has seen a qualitative advancement of our understanding of physical phenomena involved in flow separation in supersonic nozzles; in particular, the problem of side loads due to asymmetrical pressure loads, which constitutes a major restraint in the design of nozzles for satellite launchers. The development in this field is to a large extent motivated by the demand for high-performance nozzles in rocket engineering. The present paper begins with an introduction to the physical background of shock-boundary-layer interactions in basic 2D configurations, and then proceeds to internal axisymmetric nozzle flow. Special attention is given to past and recent efforts in modeling and prediction, turning physical insight into applied engineering tools. Finally, an overview is given on different technical solutions to the problem if separation and side loads, discussed in the context of rocket technology.

Стилі APA, Harvard, Vancouver, ISO та ін.

13

Abada, Omar, Abderahim Abada, and Ahmed Abdallah El-Hirtsi. "Effect of bipropellant combustion products on the rocket nozzle design." Mechanics & Industry 21, no. 5 (2020): 515. http://dx.doi.org/10.1051/meca/2020064.

Повний текст джерела

Анотація:

The focus of this research work is to investigate numerically the effect of adding the gas on the design and performance of axisymmetric MLN nozzles. A FORTRAN code was developed to design this nozzle using the characteristics method (MOC) at high temperature. The thermochemical and combustion studies of the most used liquid propellants on the satellites and launch vehicles allow to known all gases. Four engines are investigated: Ariane 5 (Vulcain 2), Ariane-5 upper stage engine (Aestus), Zenit first stage (RD-170) and Falcon 9 upper stage (Raptor). Thermodynamic analysis of parameters design MLN (such as length, Mach number, mass, thrust coefficient) was conducted. The comparison shows that the presence of 50% of H2O gas in combustion species increases the nozzle design parameters (diatomic gas including air) in the order of 25%. On the other hand, the existence of CO2 gas considerably increases approximately 35% the length and the exhaust radius. These rise depend on gases percentage in the combustion. The truncation method is applied in the MLN nozzles to optimize the thrust/weight ratio.

Стилі APA, Harvard, Vancouver, ISO та ін.

14

Wang, Yan Dong, and Hong Guang Jia. "Numerical Simulation of Laval Nozzle." Applied Mechanics and Materials 397-400 (September 2013): 266–69. http://dx.doi.org/10.4028/www.scientific.net/amm.397-400.266.

Повний текст джерела

Анотація:

Laval nozzle is the commonly used device in rocket engine and aero engine. this paper, the numerical model is derived. The convergent section subsonic flow and divergent section hypersonic flow are simulated in dimensionless method. Reverting the dimension, the result can be seen that the analytical solution, the CFX simulation solution and the numerical are in uniform.

Стилі APA, Harvard, Vancouver, ISO та ін.

15

Sellam, Mohamed, and Amer Chpoun. "Numerical Simulation of Reactive Flows in Overexpanded Supersonic Nozzle with Film Cooling." International Journal of Aerospace Engineering 2015 (2015): 1–15. http://dx.doi.org/10.1155/2015/252404.

Повний текст джерела

Анотація:

Reignition phenomena occurring in a supersonic nozzle flow may present a crucial safety issue for rocket propulsion systems. These phenomena concern mainly rocket engines which use H2gas (GH2) in the film cooling device, particularly when the nozzle operates under over expanded flow conditions at sea level or at low altitudes. Consequently, the induced wall thermal loads can lead to the nozzle geometry alteration, which in turn, leads to the appearance of strong side loads that may be detrimental to the rocket engine structural integrity. It is therefore necessary to understand both aerodynamic and chemical mechanisms that are at the origin of these processes. This paper is a numerical contribution which reports results from CFD analysis carried out for supersonic reactive flows in a planar nozzle cooled with GH2film. Like the experimental observations, CFD simulations showed their ability to highlight these phenomena for the same nozzle flow conditions. Induced thermal load are also analyzed in terms of cooling efficiency and the results already give an idea on their magnitude. It was also shown that slightly increasing the film injection pressure can avoid the reignition phenomena by moving the separation shock towards the nozzle exit section.

Стилі APA, Harvard, Vancouver, ISO та ін.

16

Zieliński, Mateusz, Piotr Koniorczyk, Janusz Zmywaczyk, and Marek Preiskorn. "Numerical simulations of temperature fields in the uncooled nozzle of a short-range anti-aircraft rocket engine with an insert in the critical section made of various materials." Bulletin of the Military University of Technology 70, no. 1 (March 31, 2021): 15–30. http://dx.doi.org/10.5604/01.3001.0015.6955.

Повний текст джерела

Анотація:

Abstract. The paper presents numerical simulations of transient heat conduction in the uncooled nozzle of a short-range anti-aircraft rocket engine. The calculations were made for the configuration of the nozzle with an insert in the critical section made of various materials. The inserts used were: POCO graphite, Al2O3 ceramics, ZrO2-3Y2O3 ceramics. For comparison, numerical simulations of the heat transfer in a nozzle made entirely of St 45 steel, the melting point of which is 1700K, were also carried out. The engine's working time was in the order of 3 s. Numerical simulations were performed using the COMSOL program. The calculation results are given in the form of temperature dependence and heat flux density as a function of time in the critical cross-section. Keywords: non-cooled nozzle, rocket engine, temperature field

Стилі APA, Harvard, Vancouver, ISO та ін.

17

Sabirzyanov, A. N., A. I. Glazunov, A. N. Kirillova, and K. S. Titov. "Simulation of a Rocket Engine Nozzle Discharge Coefficient." Russian Aeronautics 61, no. 2 (April 2018): 257–64. http://dx.doi.org/10.3103/s1068799818020150.

Повний текст джерела

Стилі APA, Harvard, Vancouver, ISO та ін.

18

Semenov, Vasiliy, Igor Ivanov, and Igor Kryukov. "Dual bell slot nozzle of a rocket engine." Perm National Research Polytechnic University Aerospace Engineering Bulletin, no. 46 (2016): 56–72. http://dx.doi.org/10.15593/2224-9982/2016.46.03.

Повний текст джерела

Стилі APA, Harvard, Vancouver, ISO та ін.

19

Yagodnikov, D. A., A. V. Voronetskii, and N. M. Pushkin. "Electrification of nozzle in a liquid rocket engine." Combustion, Explosion, and Shock Waves 31, no. 4 (1995): 450–54. http://dx.doi.org/10.1007/bf00789365.

Повний текст джерела

Стилі APA, Harvard, Vancouver, ISO та ін.

20

Sun, Dechuan, Tianyou Luo, and Qiang Feng. "New Contour Design Method for Rocket Nozzle of Large Area Ratio." International Journal of Aerospace Engineering 2019 (December 20, 2019): 1–8. http://dx.doi.org/10.1155/2019/4926413.

Повний текст джерела

Анотація:

A rocket engine for space propulsion usually has a nozzle of a large exit area ratio. The nozzle efficiency is greatly affected by the nozzle contour. This paper analysed the effect of the constant capacity ratio in Rao’s method through the design process of an apogee engine. The calculation results show that increasing the heat capacity ratio can produce an expansion contour of smaller expansion angle and exit area ratio. A simple modification of Rao’s method based on thermally perfect gas assumption was made and verified to be more effective. The expansion contour designed by this method has much thinner expansion section and higher performance. For the space engine, a new extension contour type for the end section of the nozzle is proposed. The extension curve bent outward with increasing expansion angle increases the vacuum specific impulse obviously.

Стилі APA, Harvard, Vancouver, ISO та ін.

21

Bhupendra Kumar, Mohd Shoaib, Ramanan G, and Radhakrishnan P. "Design and Computational Flow Analysis of Different Rocket Nozzle Profile." ACS Journal for Science and Engineering 2, no. 2 (September 1, 2022): 49–60. http://dx.doi.org/10.34293/acsjse.v2i2.38.

Повний текст джерела

Анотація:

In recent days the launching value of increasing the industrial business starts with lowering prices related to the launch vehicle operation. Reducing the price begins with rising the propulsion systems, creating these vehicles a lot of economical with a restricted fuel amount. The Linear Aerospike engine has incontestable bigger thrust potency over this engines used on launch vehicles. However, unresolved problems with warming plague this standard. This study proposes an approach amendment to the rocket motor that explores a way to increase the performance of the nozzle body, and lower the chance of failure because of the acute pressure environments. The results show that however the most exhaust flow of an Aerospike is full of couture of the spike surface. Victimization in Ansys CFD software system and acting a straight forward experiment with whoosh rockets, the study explores however exhaust flow changes the performance characteristics, in the main speed close to the wall. The results with some improvement is created by CFD nozzle flow studies, to make sure the bottom pure mathematics isn't a contributory issue to the film-cooling results.

Стилі APA, Harvard, Vancouver, ISO та ін.

22

Cunningham, Carson F., Mark C. Anderson, Levi T. Moats, Kent L. Gee, Grant W. Hart, Lucas K. Hall, and Steven C. Campbell. "Acoustical measurement and analysis of an Atlas V launch without solid rocket boosters." Journal of the Acoustical Society of America 151, no. 4 (April 2022): A83. http://dx.doi.org/10.1121/10.0010730.

Повний текст джерела

Анотація:

In September 2021, an Atlas V rocket without solid rocket boosters was launched from Vandenberg Space Force Base, California, carrying the NASA/USGS Landsat 9 satellite. In this launch configuration, the plumes from the RD-180 engine’s two nozzles are unobstructed, providing the opportunity to analyze the sound generated by a liquid-fuel rocket engine with an azimuthally asymmetric nozzle geometry. Acoustical data were collected at various locations surrounding the launch pad, ranging from a few hundred meters to several kilometers. This paper discusses an overview of the measurement logistics, the types of analyses that may be performed using these data, and an initial analysis of the azimuthal variability of the overall sound pressure level and spectra.

Стилі APA, Harvard, Vancouver, ISO та ін.

23

Brykov, N. A., and K. A. Tischenko. "Computational study of gas flow characteristics when using intra-nozzle interceptors for thrust vector control." Journal of Physics: Conference Series 2388, no. 1 (December 1, 2022): 012098. http://dx.doi.org/10.1088/1742-6596/2388/1/012098.

Повний текст джерела

Анотація:

Abstract This article proposes the use of cylindrical nozzle interceptors (probes) as an alternative method of rocket engine thrust vector control, which can replace traditional thrust vector control systems. The objective of this paper was to perform a three-dimensional CFD simulation of the gas flow in a rocket engine nozzle with a rod interceptor taken out of the wall without secondary injection, analyzing the calculated flow data and estimating the lateral force. The simulations will be performed in steady-state mode. In the CFD simulations, velocity distribution contours were obtained on the symmetry plane for all positions and values of the interceptor height.

Стилі APA, Harvard, Vancouver, ISO та ін.

24

Leto, Angelo. "Investigation of a Radial Turbines Compatibility for Small Rocket Engine." E3S Web of Conferences 197 (2020): 11009. http://dx.doi.org/10.1051/e3sconf/202019711009.

Повний текст джерела

Анотація:

In the radial turbine preliminary design for an expander rocket engine, a comparison was made with axial turbine used in Pratt & Whitney RL10 engine. One of the primary requirements of a liquid propellant rocket engine is the generation of a high thrust, which depends on both the mass flow rate of the propellant and the pressure in the thrust chamber. In expander-cycle engines, which are the subject of the present study, the liquid propellant is first compressed using centrifugal turbo-pumps, then it is used to cool the combustion chamber and the nozzle and, once vaporized, it flows through the turbines used to drive the turbo-pumps. The aim was to demonstrate the greater efficiency of the radial turbine with a reduction of the pressure ratio with respect to the axial turbine.

Стилі APA, Harvard, Vancouver, ISO та ін.

25

Manski, Detlef, and Gerald Hagemann. "Influence of rocket design parameters on engine nozzle efficiencies." Journal of Propulsion and Power 12, no. 1 (January 1996): 41–47. http://dx.doi.org/10.2514/3.23988.

Повний текст джерела

Стилі APA, Harvard, Vancouver, ISO та ін.

26

Cai, Guobiao, Jie Fang, Xu Xu, and Minghao Liu. "Performance prediction and optimization for liquid rocket engine nozzle." Aerospace Science and Technology 11, no. 2-3 (March 2007): 155–62. http://dx.doi.org/10.1016/j.ast.2006.07.002.

Повний текст джерела

Стилі APA, Harvard, Vancouver, ISO та ін.

27

Bennewitz, John W., Blaine R. Bigler, Mathias C. Ross, Stephen A. Danczyk, William A. Hargus, and Richard D. Smith. "Performance of a Rotating Detonation Rocket Engine with Various Convergent Nozzles and Chamber Lengths." Energies 14, no. 8 (April 7, 2021): 2037. http://dx.doi.org/10.3390/en14082037.

Повний текст джерела

Анотація:

A rotating detonation rocket engine (RDRE) with various convergent nozzles and chamber lengths is investigated. Three hundred hot-fire tests are performed using methane and oxygen ranging from equivalence ratio equaling 0.5–2.5 and total propellant flow up to 0.680 kg/s. For the full-length (76.2 mm) chamber study, three nozzles at contraction ratios ϵc = 1.23, 1.62 and 2.40 are tested. Detonation is exhibited for each geometry at equivalent conditions, with only fuel-rich operability slightly increased for the ϵc = 1.62 and 2.40 nozzles. Despite this, counter-propagation, i.e., opposing wave sets, becomes prevalent with increasing constriction. This is accompanied by higher number of waves, lower wave speed Uwv and higher unsteadiness. Therefore, the most constricted nozzle always has the lowest Uwv. In contrast, engine performance increases with constriction, where thrust and specific impulse linearly increase with ϵc for equivalent conditions, with a 27% maximum increase. Additionally, two half-length (38.1 mm) chambers are studied including a straight chamber and ϵc = 2.40 nozzle; these shortened geometries show equal performance to their longer equivalent. Furthermore, the existence of counter-propagation is minimized. Accompanying high-fidelity simulations and injection recovery analyses describe underlying injection physics driving chamber wave dynamics, suggesting the physical throat/injector interaction influences counter-propagation.

Стилі APA, Harvard, Vancouver, ISO та ін.

28

Dîrloman, F.-M., L.-C. Matache, T. Rotariu, T.-V. Țigănescu, D. Zvîncu, M.-I. Ungureanu, and O. Iorga. "Computational fluid dynamics simulations for composite rocket propellant optimization." IOP Conference Series: Materials Science and Engineering 1182, no. 1 (October 1, 2021): 012017. http://dx.doi.org/10.1088/1757-899x/1182/1/012017.

Повний текст джерела

Анотація:

Abstract When designing a rocket engine configuration, both in terms of propellant grain, combustion chamber and nozzle geometry, one of the most convenient approach is using Computational Fluid Dynamics (CFD) Simulation. Numerical simulation is an alternative method of scientific investigation, which substitutes large number of experiments that often imply high financial burden and are also dangerous for the personnel involved. The numerical approach is often more useful than consecrated experimental method because it provides complete data that cannot be directly observed or measured, or it is difficult to highlight by other means. In this study we focused on applying CFD simulation to composite rocket propellants in a rocket engine with convergent-divergent nozzle configuration using Ansys Fluent Software. An ammonium nitrate (AN) based composite rocket propellant having four components was analyzed: oxidizer, metallic fuel, binder and catalyst agent. Explo5®, a thermochemical software, was also used to calculate the equilibrium compositions of the combustion products in the combustion chamber. It turned out that the results obtained on the basis of the simulation are consistent with those of the experimental testing. The data collected so far will be used to optimize the grain configuration of the composite rocket propellant.

Стилі APA, Harvard, Vancouver, ISO та ін.

29

Hasegawa, Keiichi, Akinaga Kumakawa, Kazuo Kusaka, Masahiro Sato, Makoto Tadano, Akira Konno, Hiroshi Aoki, Eijiro Namura, and Masahiro Atsumi. "Fundamental Study of Extendible Nozzle and Dual-Bell Nozzle for Reusable Rocket Engine." SPACE TECHNOLOGY JAPAN, THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 2 (2003): 25–34. http://dx.doi.org/10.2322/stj.2.25.

Повний текст джерела

Стилі APA, Harvard, Vancouver, ISO та ін.

30

Zhao, Na, Yong Gang Yu, and Yu Qiang Wang. "Numerical Simulation of the Spray Characteristics in Small Scale Liquid Rocket Engine Combustion Chamber." Advanced Materials Research 383-390 (November 2011): 7729–33. http://dx.doi.org/10.4028/www.scientific.net/amr.383-390.7729.

Повний текст джерела

Анотація:

The mathematical and physical model of the liquid propellant spray in straight nozzle was proposed for studying the performance characteristics of the small-scale liquid rocket engine. With the Fluent software, the numerical simulation was carried out. Sauter mean diameter (SMD) of the HAN-based liquid propellant (LP1846) in the engine combustor changing with spray pressure, nozzle diameter and the liquid surface tension were analyzed. The results indicate that: in the spray pressure region of 1.8MPa~3.0MPa, at a fixed spray pressure, the smaller is the nozzle diameter, the smaller is the droplets’ SMD and the relationship between the SMD and the nozzle diameter is approximately linearity; for the same nozzle diameter and spray pressure, the larger is the surface tension, the larger is the liquid droplets’ SMD.

Стилі APA, Harvard, Vancouver, ISO та ін.

31

LEE, YOUNG-SHIN, JAE-HOON KIM, HYUN-SOO KIM, DUCK-HOI KIM, SEONG-HOI KU, and SOON-IL MOON. "A STUDY ON THE THERMAL VIBRATION ANALYSIS OF THE GRAPHITE DISK UNDER THERMAL SHOCK." International Journal of Modern Physics B 20, no. 25n27 (October 30, 2006): 4105–10. http://dx.doi.org/10.1142/s0217979206040933.

Повний текст джерела

Анотація:

Graphite is applied to structural material of the high temperature reactor and nozzle of high energy rocket engine. The excessive vibration and stress field can be occurred for this material due to the severe thermal condition. In this study, the thermal stress and vibration characteristics of ATJ graphite under high temperature condition are investigated by finite element analysis (FEA). The specimen is designed as a disk shape in order to simulate the rocket nozzle combustion condition. The experiment of thermal heat is also conducted using by CO 2 laser.

Стилі APA, Harvard, Vancouver, ISO та ін.

32

Arzhannikov, Andrey, and Alexey Beklemishev. "An Electro-Jet Rocket Engine With Big Thrust At Helical Corrugated Magnetic Field." Siberian Journal of Physics 11, no. 1 (March 1, 2016): 107–18. http://dx.doi.org/10.54362/1818-7919-2016-11-1-107-118.

Повний текст джерела

Анотація:

A fundamentally new electro-jet rocket engine having a big thrust with a high specific impulse is described in this paper. The acceleration mechanism of magnetized plasma along the axis of a cylindrical chamber with a helical corrugated magnetic field is put in the basis of such engine. The plasma acceleration is achieved during its drift motion by applying a radial electric field. The analytical description of the plasma motion process gives a visual representation of how the diamagnetic forces provide the process of the continuous acceleration of plasma ions along the axis of the helical corrugated magnetic field. As the result of this process, the accelerated plasma stream flows through the expanding cross section of a magnetic nozzle and the thrust of the rocket engine is created. Estimated calculations showed the ability of the new electro-jet rocket engine to achieve the big trust (in the range 102 –104 Newton) with the high specific impulse (from the level 3·104 to 103 seconds, respectively) at a reasonable efficiency. This set of parameters is fundamentally unattainable for another jet engines operating on the basis of other physical mechanisms.

Стилі APA, Harvard, Vancouver, ISO та ін.

33

Zhao, Wei Guo, Ji Hong Dong, Wei Li, Hai Ping Wang, and Quan Feng Guo. "Research on Defect Detection Technology of C/C Composite." Advanced Materials Research 295-297 (July 2011): 264–69. http://dx.doi.org/10.4028/www.scientific.net/amr.295-297.264.

Повний текст джерела

Анотація:

Carbon matrix/Carbon fibre reinforced composite(C/C composite) has many special characteristics such as high strength, high elastic modulus, high temperature resistance and ablation resistance. So it is used to manufacture rocket engine nozzle in aerospace filed. But its security will be seriously effected during rocket engine working due to the manufacturing defects. Based on analyzing defect types and structure characteristic of C/C composite, the ultrasonic and infrared testing methods on C/C composite are further studied in this thesis. And effective defect detection has been achieved.

Стилі APA, Harvard, Vancouver, ISO та ін.

34

Ryzhkov, V. V., and I. I. Morozov. "Technology of computational analysis of the working process parameters of low-thrust rocket engines running on gaseous oxygen-hydrogen fuel with the use of ANSYS CFD." VESTNIK of Samara University. Aerospace and Mechanical Engineering 18, no. 2 (July 2, 2019): 62–74. http://dx.doi.org/10.18287/2541-7533-2019-18-2-62-74.

Повний текст джерела

Анотація:

The paper presents the description of a mathematical model of the working process of a low-thrust rocket engine operating on gaseous oxygen-hydrogen fuel and some fragments of the technology of computational analysis of distribution of gas-dynamic parameters in the engine duct. We present the results of calculating the stream line distribution, the distribution of total temperature profile along the flow path of the engine chamber and at its characteristic cross sections, the axial component of (total) speed of combustion products in the Laval nozzle output section. The results of calculating the temperature in the area of the rocket engine’s inner wall are presented. It is shown that the distribution of the combustion products’ stagnation temperature has a significant impact on the efficiency of fuel conversion in the engine chamber, its thermal state and makes it possible to identify the ways of improving the workflow of the low-thrust rocket engine.

Стилі APA, Harvard, Vancouver, ISO та ін.

35

Verma, S. B., and Oskar Haidn. "Unsteady Side-Load Evolution in a Liquid Rocket Engine Nozzle." Journal of Spacecraft and Rockets 57, no. 2 (March 2020): 391–97. http://dx.doi.org/10.2514/1.a34556.

Повний текст джерела

Стилі APA, Harvard, Vancouver, ISO та ін.

36

Forde, Scott, Mel Bulman, and Todd Neill. "Thrust augmentation nozzle (TAN) concept for rocket engine booster applications." Acta Astronautica 59, no. 1-5 (July 2006): 271–77. http://dx.doi.org/10.1016/j.actaastro.2006.02.052.

Повний текст джерела

Стилі APA, Harvard, Vancouver, ISO та ін.

37

Vinod, G., S. Renjith, and V. Thaddeus Basker. "Thermo Structural Analysis of Carbon-Carbon Nozzle Exit Cone for Rocket Cryo Engines." Applied Mechanics and Materials 877 (February 2018): 320–26. http://dx.doi.org/10.4028/www.scientific.net/amm.877.320.

Повний текст джерела

Анотація:

Launch and space vehicle structures are required to be extremely weight efficient. The need to achieve the performance required for the engine in the upper stage of a launch vehicle, increase the payload capacity drives rocket engine manufacturers to seek higher thrust level, specific impulse and thrust to weight ratio. The use of high temperature C-C composite materials is an efficient way to reach these objectives by allowing use of high expansion ratio. Nozzle extensions benefiting of the outstanding thermal, mechanical and fatigue resistance of these materials to decrease mass and featuring high temperature margins. A three-directionally reinforced (3D) carbon-carbon (c-c) material nozzle exit cone is selected for the current study. C-C composite exit nozzle must possess excellent stability and strength under extreme conditions for a specified amount of time. Carbon-carbon composites are appropriate materials for applications that require high specific strength at elevated temperatures. The paper describes the thermo structural analysis of a typical c/c nozzle exit cone.

Стилі APA, Harvard, Vancouver, ISO та ін.

38

KOSTYUSHIN, Kirill V. "NUMERICAL INVESTIGATION OF UNSTEADY GASDYNAMIC PROCESSES AT THE LAUNCH OF SOLID-PROPELLANT ROCKETS." Vestnik Tomskogo gosudarstvennogo universiteta. Matematika i mekhanika, no. 67 (2020): 127–43. http://dx.doi.org/10.17223/19988621/67/12.

Повний текст джерела

Анотація:

The paper presents the results of the methodology developed for calculating unsteady gasdynamic processes occurring at the launch of missiles, in the gas-dynamic paths of rocket engines, and in the external regions. The method accounts for the variation in the geometry of the solidpropellant charge in the course of solid-propellant rocket engine operation and in the geometry of the computational domain at the rocket launch. The analysis of the unsteady force impact of the supersonic jet on the launch surface is carried out. It is shown that the maximum force action is located in the vicinity of the Mach disks of the unperturbed jet. Numerical studies of gasdynamic processes at the launch of a model solid-propellant booster rocket are implemented including the case when the nozzle plug opening is taken into account. The contribution of the thrust force components at the stage of bootstrap operation is assessed. The presence of the plug at the initial stage of the engine start leads to an abrupt change in the thrust and minor fluctuations, which are damped as the pressure in the combustion chamber rises.

Стилі APA, Harvard, Vancouver, ISO та ін.

39

Mironov, Daniil, and Aleksey Salnikov. "DYNAMIC BEHAVIOR OF SOLID FUEL ROCKET ENGINE DURING OPERATION (REVIEW)." Perm National Research Polytechnic University Aerospace Engineering Bulletin, no. 70 (2022): 7–17. http://dx.doi.org/10.15593/2224-9982/2022.70.01.

Повний текст джерела

Анотація:

One of the problems that has been solved for more than 80 years is the dynamic behavior both during the development of a rocket engine and during operation, these are oscillations that occur in the combustion chamber. Gas vibrations and mechanical vibrations of the rocket engine elements can cause vibration loads, which, under certain conditions, leads to resonance phenome-na. This can cause engine failures. An analysis of the behavior of a rocket engine as a dynamic system with an assessment of frequency interactions over the entire time of its operation has not been completely resolved today. Various options for studying the dynamic behavior of a rocket engine, algorithms for determining natural oscillation frequencies are considered. The analysis of existing approaches for solving the problem of determining the dynamic behavior of the rocket engine was carried out. In various works, the mechanical vibrations of the engine housing or its elements, such as a sliding nozzle, are calculated using various methods. In a number of works, the structure is considered as a model of discrete masses, where the elements are connected through the stiffness and viscosity coefficients. In other cases, fluctuations of the gas flow during combustion in the combustion chamber are considered, methods of numerical simulation of the process are developed that take into account the features of vor-tex formation and instability of the gas flow, as well as dependence on the shape of the charge. However, the joint problem has not been solved; in the presented works, the mutual influence of the vibrations of the rocket engine case with the fuel and gas flow during operation is not considered. To get a complete picture of the dynamic loads experienced in a solid propellant rocket engine, this interaction must be taken into account.

Стилі APA, Harvard, Vancouver, ISO та ін.

40

Somov, V. V. "DETERMINATION OF THE TYPE OF A SINGLE-USE GRENADE LAUNCHER BASED ON ITS COMPOSITE PARTS AND FRAGMENTS OF REACTIVE GRENADE FOUND AT THE PLACE OF ACCIDENT." Theory and Practice of Forensic Science and Criminalistics 17 (November 29, 2017): 245–52. http://dx.doi.org/10.32353/khrife.2017.31.

Повний текст джерела

Анотація:

In carrying out an investigation into the explosion, among others, the investigative version of the use of a single-use reactive grenade launcher is being considered. The most common for criminal explosions are applied grenade launchers RPG-18, RPG-22, RPG-26. Their use is due to a number of such properties as small size and weight, which makes it possible to transfer them covertly, the range of the shot significantly exceeding the range of the hand grenade throw, the high detonating effect of the rocket grenade explosion. The single-use rocket launchers are generally of the same design. Their differences are in the features of the components construction and dimensional characteristics, which are given in the article. On the basis of expert practice, details ofgrenade launchers that remain at the site of the explosion and have the least damage are determined. These details are the objects of investigation of the explosion technical expertise. These objects include launchers of grenade launchers and rocket parts ofjet grenades. The design features of the launchers, their dimensional characteristics and marking symbols make it possible to determine their belonging to a specific type of jet grenade launchers. Missile parts of jet grenades differ in the form of the combustion chamber of the jet engine, nozzle, in the size ofthe outlet section of the nozzle, in the form and size of the stabilizerfeathers. To determine the belonging of the rocket part of the grenade to a specific type ofjet grenade launcher, it’s necessary to establish a set of structural features and dimensional characteristics. At considerable damage of the combustion chamber of the jet engine, as a rule, the nozzle block remains intact that allows to define diameter of critical section of a nozzle, and on it to establish type of the used single-use grenade launcher.

Стилі APA, Harvard, Vancouver, ISO та ін.

41

Wolański, P. "RDE research and development in Poland." Shock Waves 31, no. 7 (October 2021): 623–36. http://dx.doi.org/10.1007/s00193-021-01038-2.

Повний текст джерела

Анотація:

AbstractA very short survey of research conducted in Poland on the development of the rotating detonation engine (RDE) is presented. Initial studies conducted in cooperation with Japanese partners lead to development of a joint patent on RDE. Then, an intensive basic and applied research was started at the Institute of Heat Engineering of the Warsaw University of Technology. One of the first achievements was the demonstration of performance of the rocket engine with an aerospike nozzle utilizing continuously rotating detonation (CRD), and research was directed into development of a small turbofan engine utilizing such a combustion regime. These activities promoted international cooperation and stimulated RDE development not only in Poland but also in other countries. A research directed to measure and calculate flow parameters as well as to analyze the use of liquid fuels was conducted. In the Institute of Aviation in Warsaw, research on the application of the CRD to turbine engines as well as rocket, ramjet, and combined cycle engines was carried out. In the paper, a special emphasis is given to international cooperation in this area with partners from many countries engaged in the development of the pressure gain combustion to propulsion systems.

Стилі APA, Harvard, Vancouver, ISO та ін.

42

Beyer, Steffen, Stephan Schmidt, Franz Maidl, Rolf Meistring, Marc Bouchez, and Patrick Peres. "Advanced Composite Materials for Current and Future Propulsion and Industrial Applications." Advances in Science and Technology 50 (October 2006): 174–81. http://dx.doi.org/10.4028/www.scientific.net/ast.50.174.

Повний текст джерела

Анотація:

Various technology programmes in Europe are concerned with preparing for future propulsion technologies to reduce the costs and increase the life time of components for liquid rocket engine components. One of the key roles to fulfil the future requirements and for realizing reusable and robust engine components is the use of modern and innovative materials. One of the key technologies which concern various engine manufacturers worldwide is the development of fibrereinforced ceramics – CMC's (Ceramic Matrix Composites). The advantages for the developers are obvious – the low specific weight, the high specific strength over a large temperature range, and their good damage tolerance compared to monolithic ceramics make this material class extremely interesting as a construction material. Different kind of composite materials are available and produced by EADS ST, the standard material SICARBON® (C/SiC made by Liquid Polymer Infiltration) and the new developed and qualified composite materials SICTEX® (C/SiC made by Liquid Silicon Infiltration) and CARBOTEX® (C/C made by Rapid Chemical Vapour Infiltration). The composites are based on textile techniques like weaving, braiding, stiching and sewing to produce multiaxial preforms, the SICTEX® material is densificated by the cost effective Liquid Silicon Infiltration (LSI). Over the past years, EADS Space Transportation (formerly DASA) has, together with various partners, worked intensively on developing components for airbreathing and liquid rocket engines. Since this, various prototype developments and hot firing-tests with nozzle extensions for upper and core stage engines and combustion chambers of satellite engines were conducted. MBDA France and EADS-ST have been working on the development of fuel-cooled composite structures like combustion chambers and nozzle extensions for future propulsion applications.

Стилі APA, Harvard, Vancouver, ISO та ін.

43

Gao, Yuhang, and Jian Zheng. "Noise simulation of wake field of solid rocket motor." Journal of Physics: Conference Series 2364, no. 1 (November 1, 2022): 012004. http://dx.doi.org/10.1088/1742-6596/2364/1/012004.

Повний текст джерела

Анотація:

Abstract When the rocket engine works, it will produce high-temperature and high-speed gas, which will be rapidly mixed with the surrounding medium, and will produce huge jet noise. This paper studies jet flow field and sound field of rocket engine using Fluent. The broadband noise simulation shows that the noise distribution is related to the nozzle jet. The sound pressure data of transient FW-H observation points show that the noise frequency distribution is wide. At the same time, the size of the noise will increase with the increase of the inlet pressure, and the influence range will become larger with the increase of area expansion ratio.

Стилі APA, Harvard, Vancouver, ISO та ін.

44

Khan, Sohaib, Muhammad Umer Sohail, Ihtzaz Qamar, Muzna Tariq, and Raees Fida Swati. "Effect of Secondary Combustion on Thrust Regulation of Gas Generator Cycle Rocket Engine." Applied Sciences 12, no. 20 (October 19, 2022): 10563. http://dx.doi.org/10.3390/app122010563.

Повний текст джерела

Анотація:

Thrust regulation is applied to maintain the performance of the liquid propellant rocket engine. The thrust level of a rocket engine can be readily controlled by adjusting the number of propellants introduced into the combustion chamber. In this study, a gas generator design is proposed in which thrust regulation is maintained by performing secondary combustion in the divergent section of the nozzle of a gas generator. Tangential and normal injection techniques have also been studied for better combustion analyses. A normal injection technique is used for the experiment and CFD results are validated with the experimental data. Chemical equilibrium analyses are also performed by minimizing Gibbs free energy with the steepest descent method augmented by the Nelder–Mead algorithm. These equilibrium calculations give the combustion species as obtained through the CFD results. Performance evaluation of the rocket engine, with and without secondary combustion in the gas generator, led to an increase of 42% thrust and 46.15% of specific impulse with secondary combustion in the gas generator.

Стилі APA, Harvard, Vancouver, ISO та ін.

45

Mosolov, S. V., I. G. Lozino-Lozinskaya, D. M. Pozvonkov, and D. F. Slesarev. "Test Results of a Model Additively Manufactured Oxygen-Methane Combustion Chamber of a Liquid Rocket Engine." Herald of the Bauman Moscow State Technical University. Series Mechanical Engineering, no. 3 (138) (September 2021): 60–79. http://dx.doi.org/10.18698/0236-3941-2021-3-60-79.

Повний текст джерела

Анотація:

The paper focuses on an experimental unit developed for modeling combustion characteristics in a model oxygen-methane combustion chamber of a liquid rocket engine. The key components of the unit, i.e., the mixing head of the combustion chamber and the regeneratively cooled nozzle, were manufactured using advanced methods of additive manufacturing. The paper emphasizes the specific character of the combustion chamber components made with the use of additive technology and introduces hot-fire test results of the model combustion chamber as part of the experimental unit. The study shows the durability of the mixing head and combustion chamber nozzle under hot-fire test conditions, as well as the reliable operation of the experimental unit as a whole, which confirms the selected design and technological solutions. Within the study, we analyzed the cooling system of the experimental unit for the test conditions, estimated the thermal state of the nozzle, with account for the features of the additively manufactured cooling path. To increase the cooling system’s reliability and expand the combustion chamber pressure application, it is recommended to apply a heat-shielding coating on the firewall of the nozzle. Using new experimental data, we analyzed the parameters of improving the efficiency of the model combustion chamber with the additively manufactured components and corresponding in scale and consumption characteristics to the combustion chamber of the liquid rocket engine

Стилі APA, Harvard, Vancouver, ISO та ін.

46

Simmons, J., and Richard Branam. "Parametric Study of Dual-Expander Aerospike Nozzle Upper-Stage Rocket Engine." Journal of Spacecraft and Rockets 48, no. 2 (March 2011): 355–67. http://dx.doi.org/10.2514/1.51534.

Повний текст джерела

Стилі APA, Harvard, Vancouver, ISO та ін.

47

Otsubo, Naoyuki, Kunio Hirata, and Hirotaka Otsu. "659 Enhancement of Rocket Engine Thrust using Variable Area Ratio Nozzle." Proceedings of Conference of Tokai Branch 2008.57 (2008): 439–40. http://dx.doi.org/10.1299/jsmetokai.2008.57.439.

Повний текст джерела

Стилі APA, Harvard, Vancouver, ISO та ін.

48

Trzun, Zvonko, Milan Vrdoljak, and Hrvoje Cajner. "The Effect of Manufacturing Quality on Rocket Precision." Aerospace 8, no. 6 (June 4, 2021): 160. http://dx.doi.org/10.3390/aerospace8060160.

Повний текст джерела

Анотація:

The effect of manufacturing quality on rocket impact point dispersion is analyzed. The approach presented here applies to any type of rocket. Here, manufacturing quality is demonstrated for the unguided rocket, and by simulating four typical manufacturing errors: erroneously manufactured warhead, misalignment between the warhead and engine chamber, asymmetrically installed propellant, and error in nozzle manufacturing. A new methodology is proposed, which combines a 3D CAD model of the asymmetrical projectile (due to manufacturing errors) and the improved Six-degrees-of-freedom (6DOF) model of its flight into a comprehensive Monte-Carlo simulation. In that way, the rocket trajectory dispersion is correlated directly to the imperfection of the manufacturing process. Three quality levels are simulated (low, standard, and high quality), and each of the analyzed manufacturing errors depends on the chosen quality. The results show how important it is to impose the highest quality on nozzle manufacturing, and if this condition is not met, reveal if strict tolerances applied to other steps of the manufacturing process can compensate for the consequential drop of precision.

Стилі APA, Harvard, Vancouver, ISO та ін.

49

Ma, Yarui, Jiwen Cui, Hui Wang, and Jiubin Tan. "Impacts of Micro-Deviations of Aperture on the Characteristics of Collision Atomization Field." Applied Sciences 12, no. 9 (May 6, 2022): 4685. http://dx.doi.org/10.3390/app12094685.

Повний текст джерела

Анотація:

As the final flow channel of the liquid rocket engine, the manufacturing of impinging atomization nozzles has become a critical link in the manufacture of impinging atomization components. At present, the high-precision machining of millimeter nozzles in large-scale production is quite difficult, which inevitably leads to the diversity of internal flow field and atomization field parameters. In this investigation, the influence of the diameter deviation of impinging nozzles on the atomization field is analyzed by experiment. The transparent nozzle is installed in a typical colliding atomizer. The ultra-precision measurement is carried out by the optical fiber measurement system and the flow field in the nozzle is visualized. The information of atomized droplets in the atomization field is assembled by the laser interferometric particle imaging technology (IPI). The experimental results indicate that the micro-deviations of the collision aperture have a profound influence on the cavitation state of the flow field in the nozzles and the atomization characteristics (droplet diameter and atomization cone angle) of the atomization field.

Стилі APA, Harvard, Vancouver, ISO та ін.

50

Lin, Binbin, Hongliang Pan, Lei Shi, and Jinying Ye. "Effect of Primary Rocket Jet on Thermodynamic Cycle of RBCC in Ejector Mode." International Journal of Turbo & Jet-Engines 37, no. 1 (March 26, 2020): 61–70. http://dx.doi.org/10.1515/tjj-2017-0013.

Повний текст джерела

Анотація:

AbstractThrust augmentation in ejector mode is important to improve engine performance so as to enable more propulsion applications of RBCC. Usually, the internal flow-path is configured mainly for combustion organization in ramjet and scramjet modes, thus the RBCC performance in ejector mode will mainly depend on the primary rocket operating parameters and jet expansion state. Considering such restrictions and requirements on the primary rocket, this paper studies the effects of the primary rocket jet on the thermodynamic cycle performance of ejector mode, in which incoming air is dominated by ejector-suction by the primary rocket jet at low flight Mach numbers. It is found that the engine performance in ejector mode will be improved by increasing the primary rocket chamber pressure and nozzle expansion ratio. Furthermore, for better design of primary rocket, primary rocket chamber pressure should be maximized on the basis of ensuring a complete expansion of primary rocket to the ambient pressure at the design point of flight conditions after 3-D CFD simulations for a full flow-path of RBCC including fore-airframe and aft-airframe of a flight vehicle. The results of 3-D numerical simulations show that the bypass ratio and specific impulse increase by 35.5 % and 12.5 % respectively at sea-level static condition.

Стилі APA, Harvard, Vancouver, ISO та ін.

Статті в журналах з теми "Rocket engine nozzle"

Оформте джерело за APA, MLA, Chicago, Harvard та іншими стилями