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Journal articles on the topic 'Ideal rocket equation'
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1
Li, Xuge. "Comprehensive Exploration on Performance Improvement of Rocket Thruster." Highlights in Science, Engineering and Technology 88 (March 29, 2024): 847–52. http://dx.doi.org/10.54097/bnr30f66.
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Since mankind has been carrying out space research for decades, rockets are one of the most important tools used by mankind to explore space. In this article, the structure of a rocket is analyzed, and the function of each part is understood. Tsolkovsky's rocket equation is used to analyze the relationship between the initial and final mass of the rocket. Through the study of engine thrust, the calculation of engine thrust was demonstrated using the formula. Besides, the ideal specific impulse of different fuels in a vacuum is compared in a graph and some of the fuels in the chart. Ultimately, it was concluded that the performance of rockets could be improved in terms of open-edge construction, enhanced engine thrust, utilization of composite fuels and improved combustion efficiency. In the future, due to the development of rocket technology, perhaps commercial space activities will usher in a new trend.
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D’Alessandro, Simone, Marco Pizzarelli, and Francesco Nasuti. "A Hybrid Real/Ideal Gas Mixture Computational Framework to Capture Wave Propagation in Liquid Rocket Combustion Chamber Conditions." Aerospace 8, no. 9 (2021): 250. http://dx.doi.org/10.3390/aerospace8090250.
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The present work focuses on the development of new mathematical and numerical tools to deal with wave propagation problems in a realistic liquid rocket chamber environment. A simplified real fluid equation of state is here derived, starting from the literature. An approximate Riemann solver is then specifically derived for the selected conservation laws and primitive variables. Both the new equation of state and the new Riemann solver are embedded into an in-house one-dimensional CFD solver. The verification and validation of the new code against wave propagation problems are then performed, showing good behavior. Although such problems might be of interest for different applications, the present study is specifically oriented to the low order modeling of high-frequency combustion instability in liquid-propellant rocket engines.
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Yang, Yuqi, Longbin Liu, and Moyan Chen. "Study on thrust performance of small water rocket launch." Journal of Physics: Conference Series 2313, no. 1 (2022): 012021. http://dx.doi.org/10.1088/1742-6596/2313/1/012021.
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Abstract For water rocket test of low cost, quick assembly and simple, based on the single water rocket, using Bernoulli equation, the ideal gas adiabatic model and mass conservation theorem, this paper derived both the calculation method of thrust rocket the bottle under the different initial water volume ratio of water pressure, and the rule about how the water rocket thrust and pressure performance change. It is found that the pressure, the jet velocity of the water relative to the water rocket body and thrust of water rocket will decrease with the decrease of water amount under different initial charging pressure and different initial water amount. If the initial pressure is higher, the water rocket will gain more energy and launch faster; However, it doesn’t mean that the more initial water storage ratio is, the better. The maximum velocity obtained by water rocket under the same initial pressure increases first and then decreases with the initial water storage ratio and the optimal initial water storage ratio increases with the increase of initial pressure. According to data calculation, the numerical value increases with the increase of initial pressure. The optimal performance of water rocket can be obtained by taking the maximum pressure and the corresponding optimal initial injection water when the water rocket body is allowed.
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Hu, Haifeng, Xinni Gao, Yushan Gao, and Jianwen Yang. "Shock Wave and Aeroelastic Coupling in Overexpanded Nozzle." Aerospace 11, no. 10 (2024): 818. http://dx.doi.org/10.3390/aerospace11100818.
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The growing demand for increasing the engine power of a liquid rocket is driving the development of high-power De-Laval nozzles, which is primarily achieved by increasing the expansion ratio. A high-expansion-ratio for De-Laval nozzles can cause flow separation, resulting in unsteady, asymmetric forces that can limit nozzle life. To enhance nozzle performance, various separation control methods have been proposed, but no methods have been fully implemented thus far due to the uncertainties associated with simulating flow phenomena. A numerical study of a high-area-ratio rocket engine is performed to analyze the aeroelastic performance of its structure under flow separation conditions. Based on numerical methodology, the flow inside a rocket nozzle (the VOLVO S1) is analyzed, and different separation patterns are comprehensively discussed, including both free shock separation (FSS) and restricted shock separation (RSS). Since the location of the flow separation point strongly depends on the turbulence model, both the single transport equation and two-transport-equation turbulence models are simulated, and the findings are compared with the experimental results. Therefore, the Spalart–Allmaras (SA) turbulence model is the ideal choice for this rocket nozzle geometry. A wavelet is used to analyze the amplitude frequencies from 0 to 100 Hz under various pressure fluctuation conditions. Based on a clear understanding of the flow field, an aeroelastic coupling method is carried out with loosely coupled computational fluid dynamics (CFD)/computational structural dynamics (CSD). Some insights into the aeroelasticity of the nozzle under separated flow conditions are obtained. The simulation results show the significant impact of the structural response on the inherent pressure pulsation characteristics resulting from flow separation.
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Akiki, M., and J. Majdalani. "Compressible integral representation of rotational and axisymmetric rocket flow." Journal of Fluid Mechanics 809 (November 9, 2016): 213–39. http://dx.doi.org/10.1017/jfm.2016.654.
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This work focuses on the development of a semi-analytical model that is appropriate for the rotational, steady, inviscid, and compressible motion of an ideal gas, which is accelerated uniformly along the length of a right-cylindrical rocket chamber. By overcoming some of the difficulties encountered in previous work on the subject, the present analysis leads to an improved mathematical formulation, which enables us to retrieve an exact solution for the pressure field. Considering a slender porous chamber of circular cross-section, the method that we follow reduces the problem’s mass, momentum, energy, ideal gas, and isentropic relations to a single integral equation that is amenable to a direct numerical evaluation. Then, using an Abel transformation, exact closed-form representations of the pressure distribution are obtained for particular values of the specific heat ratio. Throughout this effort, Saint-Robert’s power law is used to link the pressure to the mass injection rate at the wall. This allows us to compare the results associated with the axisymmetric chamber configuration to two closed-form analytical solutions developed under either one- or two-dimensional, isentropic flow conditions. The comparison is carried out assuming, first, a uniformly distributed mass flux and, second, a constant radial injection speed along the simulated propellant grain. Our amended formulation is consequently shown to agree with a one-dimensional solution obtained for the case of uniform wall mass flux, as well as numerical simulations and asymptotic approximations for a constant wall injection speed. The numerical simulations include three particular models: a strictly inviscid solver, which closely agrees with the present formulation, and both $k$–$\unicode[STIX]{x1D714}$ and Spalart–Allmaras computations.
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Knoetze, J. H. "Die berekening van die stukrag van ’n vuurpylmotor." Suid-Afrikaanse Tydskrif vir Natuurwetenskap en Tegnologie 12, no. 3 (1993): 67–71. http://dx.doi.org/10.4102/satnt.v12i3.565.
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Traditionally the thrust of a rocket motor is calculated by first calculating the thrust coefficient and then multiplying it by the product of the throat area and pressure. The thrust coefficient is calculated using a standard gas dynamics equation. This equation assumes that the combustion products are a single component, non-reacting ideal gas and that the flow through the nozzle is isentropic. The thrust coefficient is a function of the ratio of specific heats, y, the area ratio of the nozzle and the motor and ambient pressures. Standard methods exist for calculating the tosses due to deviations from the assumed flow. The combustion products of modern composite propellants contain a significant portion of condensed species (primarily A1₂O₃), while the composition of the combustion products changes continuously as the products move throught the nozzle. Some uncertainty therefore exists with regard to which value of y to use and how to handle the condensed species. The assumption o f an ideat, non-reacting gas can be el iminated hy as.mming the process to he isentropic and to calculate the thrust hy using the thermodynamic state and composition of the combustion products in the motor and nozzle exit. This can be achieved by using any of the standard thermochemistry programs available in the rocket industry. It is thus possible to use the results of a standard thermochemistry program directly in an alternative method for calculating thrust. Using this method only the mass flow rate (which is a function of pressure, throat area and effective caracteristic velocity) and the results from the thermochemistry program are needed to calculate the thrust. The advantages of the alternative method are illustrated by comparing the results of the two methods with a measured thrust curve.
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Costa, Fernando S., and Gustavo A. A. Fischer. "Propulsion and Thermodynamic Parameters of van der Waals Gases in Rocket Nozzles." International Journal of Aerospace Engineering 2019 (August 14, 2019): 1–11. http://dx.doi.org/10.1155/2019/3139204.
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Propellants or combustion products can reach high pressures and temperatures in advanced or conventional propulsion systems. Variations in flow properties and the effects of real gases along a nozzle can become significant and influence the calculation of propulsion and thermodynamic parameters used in performance analysis and design of rockets. This work derives new analytical solutions for propulsion parameters, considering gases obeying the van der Waals equation of state with specific heats varying with pressure and temperature. Steady isentropic one-dimensional flows through a nozzle are assumed for the determination of specific impulse, characteristic velocity, thrust coefficient, critical flow constant, and exit and throat flow properties of He, H2, N2, H2O, and CO2 gases. Errors of ideal gas solutions for calorically perfect and thermally perfect gases are determined with respect to van der Waals gases, for chamber temperatures varying from 1000 to 4000 K and chamber pressures from 5 to 35 MPa. The effects of covolumes and intermolecular attraction forces on flow and propulsion parameters are analyzed.
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TAKAHASHI, Ryuji, Nobuyuki TSUBOI, Takashi TOKUMASU, and Shin-ichi TSUDA. "Validation of Soave–Redlich–Kwong equation of state coupled with a classical mixing rule for sound speed of non-ideal gas mixture of oxygen-hydrogen as liquid rocket propellants." Journal of Thermal Science and Technology 18, no. 1 (2023): 22–00365. http://dx.doi.org/10.1299/jtst.22-00365.
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Дубровський, І. Д., and В. Л. Бучарський. "THE APPLICATION OF THE EXTENDED CELLS METHOD TO SIMULATE THE FLOW OF COMBUSTION GASES IN THE LPRE CHAMBER." Journal of Rocket-Space Technology 31, no. 4 (2023): 32–39. http://dx.doi.org/10.15421/452305.
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 Abstract. The numerical modeling of the process of two-dimensional axisymmetric flow of combustion gases in the chamber of a liquid rocket engine is considered in this study. In general, when solving such problems, meshes are used which lines coincide with the boundaries of the computational domain. However, an alternative solution is proposed here, which is to apply the extended cells method. It allows using rectangular Cartesian grids, which lines do not coincide with the boundaries of the computational domain, without reducing the stability of the numerical solution due to the fractional finite volumes. This also simplifies the setting of boundary conditions in such volumes. The advantage of the proposed approach over the generally accepted one is the absence of the global geometric transformations during the entire modelling process, which leads to a reduction in its duration. To perform the numerical modelling, an inviscid ideal compressible gas of constant chemical composition was chosen as a basic model of a continuum. It is described by a system of the unsteady Euler equations in integral form, which was closed by the Mendeleev-Clapeyron equation of state. For the numerical solution of this system, the finite volume method was used with the reconstruction of the flow parameters by the WENO algorithm of the third order of accuracy. The solution of the Riemann problem was carried out using the Lax-Friedrichs relations. Time integration of the system of equations was performed using the explicit Runge-Kutta method of the third order of accuracy. All calculations were carried out on a uniform rectangular Cartesian mesh, which lines did not coincide with the boundaries of the computational domain. The results were compared with the solution of the same problem using the ANSYS Fluent on an unstructured mesh coinciding with the boundaries of the computational domain. The value of the relative error obtained as a result of comparing both solutions did not exceed 0.05.
 
 
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Fu, Jia, and Chaoqi Xia. "Microstructure Evolution and Mechanical Properties of X6CrNiMoVNb11-2 Stainless Steel after Heat Treatment." Materials 14, no. 18 (2021): 5243. http://dx.doi.org/10.3390/ma14185243.
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X6CrNiMoVNb11-2 supermartensitic stainless steel, a special type of stainless steel, is commonly used in the production of gas turbine discs in liquid rocket engines and compressor disks in aero engines. By optimizing the parameters of the heat-treatment process, its mechanical properties are specially adjusted to meet the performance requirement in that particular practical application during the advanced composite casting-rolling forming process. The relationship between the microstructure and mechanical properties after quenching from 1040 °C and tempering at 300–670 °C was studied, where the yield strength, tensile strength, elongation and impact toughness under different cooling conditions are obtained by means of mechanical property tests. A certain amount of high-density nanophase precipitation is found in the martensite phase transformation through the heat treatment involved in the quenching and tempering processes, where M23C6 carbides are dispersed in lamellar martensite, with the close-packed Ni3Mo and Ni3Nb phases of high-density co-lattice nanocrystalline precipitation created during the tempering process. The ideal process parameters are to quench at 1040 °C in an oil-cooling medium and to temper at 650 °C by air-cooling; final hardness is averaged about 313 HV, with an elongation of 17.9%, the cross-area reduction ratio is 52%, and the impact toughness is about 65 J, respectively. Moreover, the tempered hardness equation, considering various tempering temperatures, is precisely fitted. This investigation helps us to better understand the strengthening mechanism and performance controlling scheme of martensite stainless steel during the cast-rolling forming process in future applications.
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Vasiliev, Igor, Boris Kiforenko, and Yaroslav Tkachenko. "COMPARATIVE ANALYSIS OF THE EFFICIENCY OF CONSTANT POWER THROTTLED ROCKET ENGINES FOR INTERORBITAL FLIGHTS TO GEOSTATIONAR." Journal of Automation and Information sciences 6 (November 1, 2021): 66–77. http://dx.doi.org/10.34229/1028-0979-2021-6-7.
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Carrying out low-thrust transfers of spacecrafts in the near-earth space from intermediate elliptic to the geostationary orbit using electric rocket engines seems to be one of the most important tasks of modern cosmonautics. Electric rocket engines, whose specific impulse of the reactive jet is an order of magnitude more than in chemical RD, are preferable for interorbit flights with a maximum payload in the case when a significant increase in the duration of the maneuver is permissible. Ability to throttling the rocket engine thrust is traditionally considered as one of the ways to reduce both the engine mass and the required fuel assumptions for performing the specified maneuver. Using the concept of an ideal-rocket engine provides the upper estimates of the payload mass of interborbital flights for the given power level. Accounting for the properties of real engines leads to the need of considering the mathematical models with more strict limits on control functions. A study of the efficiency of three modes of thrust control of an electric propulsion rocket engine was carried out when performing practically interesting spacecraft flights from highly elliptical intermediate near-earth orbits to geostationary orbits. A mathematical model of constant power relay rocket engine has been built. The formulation of the variational problem of the Maer type is given about the execution of a given dynamic maneuver for the throttled and unregulated electric rocket engines of constant power. Using the Pontryagin maximum principle, an analysis of the optimal control functions was carried out, for which the final relations were written out, which allowed to write down the system of differential equations of the optimal movement of the spacecraft, equipped with relay electric rocket engine. The obtained numerical and quality results of the study of the effectiveness of various modes of thrust control of an electric propulsion engine to increase the payload of a given orbital maneuver confirmed the correctness of mathematical models of throttled and relay engines and, in general, the efficiency of using solutions of the averaged equations of optimal motion of a spacecraft for numerical solution of the corresponding boundary value problems in an exact formulation.
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Longmire, Nelson P., and Daniel T. Banuti. "Limits of Fluid Modeling for High Pressure Flow Simulations." Aerospace 9, no. 11 (2022): 643. http://dx.doi.org/10.3390/aerospace9110643.
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Flows in liquid propellant rocket engines (LRE) are characterized by high pressures and extreme temperature ranges, resulting in complex fluid behavior that requires elaborate thermo-physical models. In particular, cubic equations of state and dedicated models for transport properties are firmly established for LRE simulations as a way to account for the non-idealities of the high-pressure fluids. In this paper, we review some shortcomings of the current modeling paradigm. We build on the common study of property errors, as a direct measure of the density or heat capacity accuracy, to evaluate the quality of cubic equations of state with respect to pseudo boiling of rocket-relevant fluids. More importantly, we introduce the sampling error as a new category, measuring how likely a numerical scheme is to capture real fluid properties during a simulation, and show how even reference quality property models may lead to errors in simulations because of the failure of our numerical schemes to capture them. Ultimately, a further evolution of our non-ideal fluid models is needed, based on the gained insight over the last two decades.
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Jung, Taekyu, and Sejin Kwon. "Design and performance evaluation of a bellows-type mixture ratio stabilizer for a liquid bipropellant rocket engine." Proceedings of the Institution of Mechanical Engineers, Part C: Journal of Mechanical Engineering Science 223, no. 3 (2008): 723–31. http://dx.doi.org/10.1243/09544062jmes1216.
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A stabilizer that maintains a steady propellant mixture ratio in a liquid bipropellant rocket engine was introduced. First, a design criterion for the ideal performance of a general stabilizer was derived. A new stabilizer with bellows (bellows-type stabilizer) was proposed in the present study and relevant design parameters were identified by a mathematical model as well as a theoretical analysis. Governing equations were established to predict the static behaviour of the stabilizer. A bellows-type stabilizer was fabricated and its performance was measured. The performance predicted by the mathematical model showed satisfactory agreement with measurements and this validates the adequacy of the mathematical model proposed in the present study.
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Li, Guoyu, Yongxing Yang, Zijie Mao, Mengxue Chen, and Xueyu Lu. "Research on kinematic characteristics of spatial crank rocker mechanism based on MATLAB." Journal of Physics: Conference Series 2562, no. 1 (2023): 012081. http://dx.doi.org/10.1088/1742-6596/2562/1/012081.
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Abstract Many students trained in carbon-free trolley engineering to design smoothly but found that it was difficult to achieve the ideal results with a lot of time and effort in trajectory debugging. To deeply explore the characteristics of space crank rocker, through repeated experiments and theoretical analysis of carbon-free trolleys, and using MATLAB, SolidWorks, SPSS, and other software analysis, combined with projection geometric analysis to derive the equations of motion and derive the displacement, velocity, and acceleration characteristic curves. The conditions for the existence of the space crank rocker mechanism with an interlaced angle of 90° are derived, and the relationship between corner amplitude and main structural parameters is obtained. The size of the influence factor of each structural parameter was compared. In the commissioning, according to the characteristics of these parameters, the commissioning cycle of the carbon-free trolley can be greatly shortened.
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Askerov, Adil A., Anastasiya V. Chervakova, and Kirill V. Kostyushin. "Investigation of acoustic characteristics of a single supersonic jet flowing into a flooded space." Vestnik Tomskogo gosudarstvennogo universiteta. Matematika i mekhanika, no. 78 (2022): 49–59. http://dx.doi.org/10.17223/19988621/78/4.
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In this paper, the acoustic characteristics of a single supersonic jet flowing from a nozzle of a rocket engine into a flooded space at different pressure ratios are studied. A system of the Favre-averaged Navier-Stokes equations is used to describe the unsteady flow of a viscous compressible heat-conducting gas in a supersonic Laval nozzle and an outflowing jet. The system is enclosed by the ideal gas law. The implementation of the physical and mathematical model and the numerical studies are carried out using the OpenFOAM Extended open platform based on the modified dbnsTurbFoam solver. A conical nozzle with an opening angle of 45° at the Mach number of 3 at the nozzle exit is considered in this study. Air is used as the working gas. Amplitude-frequency spectra of acoustic radiation at the point located at a distance from the nozzle outlet are obtained at different pressure ratios of the outflowing supersonic jet. Analysis of the amplitude-frequency characteristics of the jet under study shows that the maxima occur mainly at low frequencies. The maximum oscillation amplitude for the considered jet configurations is revealed at a pressure ratio of 1 on a frequency of 787 Hz. The maximum sound pressure level is 149 dB.
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Hedemann, Samuel R. "Correction to the Traditional Ideal Rocket Thrust Equation." May 19, 2017. https://doi.org/10.5281/zenodo.3596173.
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The accepted ideal rocket equation for thrust, widely used in rocketry literature, is proved to be incorrect. The correct replacement is then presented and shown to obey the full form of Newton's second law \(F={\dot p}\) , so the sum of external forces on the rocket is the time derivative of its momentum. The accepted velocity equation is correct, which may explain why this error has been overlooked.
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"Theoretical studies of parabolic and hyperbolic equations of motion to refine the design parameters of processes in a liquid rocket engine." Vestnik Mashinostroeniya, March 2023, 221–24. http://dx.doi.org/10.36652/0042-4633-2023-102-3-221-224.
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The problems of operation of a liquid-propellant rocket engine (LRE) on real gases with viscosity and compressibility are considered. The equations for laminar flow, which take into account the viscoelastic term characterizing the real flow, are analyzed and compared with the Euler equation for an ideal gas. The possibility of profiling supersonic nozzles of liquidpropellant rocket engines and other rocket engines is shown, taking into account the viscoelastic properties of real gas. Keywords:second order equations, laminar flow, liquid rocket engine, parameters, numerical methods, supersonic flow. mger_97@mail.ru
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Kovalev, K. E., K. V. Fedotova, and O. A. Vorozheeva. "Computational study of the system efficiency of supplying components in the model low-thrust rocket engine on oxygen-methane." Engineering Journal: Science and Innovation, no. 10 (130) (October 2022). http://dx.doi.org/10.18698/2308-6033-2022-10-2217.
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The paper considers a model low-thrust rocket engine using the environmentally friendly gaseous oxygen–methane components as a scientific and technical lead obtained by the authors in the course of preliminary experimental studies. The engine design makes it possible to investigate the influence of the mixing unit configuration, namely position of the supply holes and presence or absence of the components swirling on the mixing process efficiency. Numerical simulation was carried out in a three-dimensional stationary formulation of “cold” mixing of gaseous oxygen and methane and was based on the Favre-averaged Navier — Stokes equations solution closed by the k–ω-SST turbulence model and the ideal gas state equation. Calculation results are provided for various configurations of the mixing unit. It is shown that the most efficient method for the considered model low-thrust rocket engine is the method of supplying gaseous components with swirling in a single direction.
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El-Gamal, M., E. Gutheil, and J. Warnatz. "The Structure of Laminar Premixed H2-Air Flames at Elevated Pressures." Zeitschrift für Physikalische Chemie 214, no. 4 (2000). http://dx.doi.org/10.1524/zpch.2000.214.4.419.
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In high-pressure flames that occur in many practical combustion devices such as industrial furnaces, rocket propulsion and internal engine combustion, the assumption of an ideal gas is not appropriate. The present paper presents a model that includes modifications of the equation of state, transport and thermodynamic properties. The model is implemented into a Fortran program that was developed to simulate numerically one-dimensional planar premixed flames. The influence of the modifications for the real gas behavior on the laminar flame speed and on flame structure is illustrated for stoichiometric H
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Priamadi, Eko, Arif Nur Hakim, and Romie O. Bura. "DESAIN NOSEL ROKET CAIR RCX250 MENGGUNAKAN METODE PARABOLIK DENGAN MODIFIKASI SUDUT EKSPANSI." Jurnal Teknologi Dirgantara 9, no. 1 (2012). http://dx.doi.org/10.30536/j.jtd.2011.v9.a1622.
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The present research is conducted to design the optimum nozzles for RCX250 engine, that is designed to produce maximum thrust of 250 kgf with combination of LOX and Kerosene as its propellant. The new nozzles were determined to be parabolic nozzle, with conical nozzle as its comparison. The parabolic nozzle was designed using Thrust Optimized Parabolic (TOP) method invented by G.V.R.Rao. TOP nozzle design method is performed by approximating a Thrust Optimized Contoured (TOC) Nozzle using parabolic equation. The method would result more efficient nozzle than conical or ideal bell nozzle. Further, the parabolic nozzle were modified in its initial and exit angle to create uniform velocities distribution at nozzle exit. A Computational Fluid Dynamics Method (CFD) is used to simulate the nozzle designs. The simulation was carried out in axis-symmetric condition using commercial CFD software. The simulation results show that MOD 1 nozzle, with initial angle (θN) 26 deg and exit angle (θe) 12 deg, gives maximum thrust, which is 4.67 % higher than reference conical nozzle. Key words:Liquid rocket, Parabolic nozzle, Thrust, CFD
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Wang, Chen, Li-bo Ding, and Tan Lu. "Variable-structure trajectory correction control of small-caliber ammunition based on cascade fuzzy active disturbance rejection." Proceedings of the Institution of Mechanical Engineers, Part C: Journal of Mechanical Engineering Science, November 25, 2024. http://dx.doi.org/10.1177/09544062241296918.
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In this study, a low-cost two-dimensional (2D) trajectory correction mechanism is proposed to address the large trajectory deviation of small-caliber extended-range ammunitions. By controlling the projectile’s roll, a single-channel 2D trajectory correction function can be realized. Moreover, based on the ideal trajectory, a five-degree-of-freedom approximate rigid-body trajectory equation is proposed. A high-order system is transformed into two low-order cascading subsystems using fuzzy control and the self-disturbance rejection control theory, which reduces the parameter tuning complexity. The entire control process is divided into three stages which realizes the second-order variable structure control of non-affine nonlinear subsystems. A low-altitude atmospheric disturbance model is established, and as an example, numerical calculations are performed on the control process and projectile impact point dispersion of a 35 mm small-caliber rocket. The results verify the performance of the proposed trajectory correction under atmospheric disturbances. This study provides a control approach for coupled underactuated nonlinear systems and provides guidance for the design of a low-cost 2D trajectory correction controller for small-caliber ammunitions.
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"Optimal Interception Technique Variation Calculus Task of Multidimensional Visual Ballistic Missiles." International Journal of Innovative Technology and Exploring Engineering 8, no. 12 (2019): 2273–79. http://dx.doi.org/10.35940/ijitee.l3376.1081219.
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A strategy of manoeuvrating using inverted flight is presented to determine the optimal path for the dive stage of a hypersonic missile. The fighting incident in this document is taken into account when the hypersonic missile strikes the goal on the floor. In particular, the hypersonic rockets are first implemented by a maneuvering type called the reversed plane. The ideal route is later conceived by minimizing the attack time with the current terminal route angle, taking into consideration the limitations of the attack angle, fluid pressure, thermal transfer rate and ordinary overload. A better pseudo-spectral hp-adaptive technique with a mesh size decrease is introduced to resolve the constructed trajectory optimization issue. The paper shows the application of variation calculations in the process of anti-ballistic missile interception, to achieve the optimal concept of the open-switch control law. It provides analytical outcomes in the shape of suitable Euler-Lagrange Equations, relying on the preservation of the wind concept, for three distinct indices, using a simple racket and missile model. It also offers the computer program to simulate the intercom system quickly, with chosen parameters, parametrical assessment and simple alteration possibilities by other scientists, as published as m-Function of MATLAB in accessible code.