scholarly journals Optimization of Convergent –Divergent Taper Angle with Combustion Chamber of Rocket Engine through Numerical Analysis

2020 ◽  
Vol 9 (6) ◽  
pp. 76-81
Keyword(s):  
Carbon Dioxide ◽  
Steady State ◽  
Numerical Analysis ◽  
Combustion Chamber ◽  
Main Part ◽  
Rocket Engine ◽  
Taper Angle ◽  
Optimal Values ◽  
Mach Numbers ◽  
Divergent Angle

Present days speed of vehicle is concerned in both sub -sonic vehicle( like car, bus and truck etc) super -sonic vehicles (like rockets). Convergent divergent component is the main part which decide the speed of any vehicle. A attempt is made to find the optimal convergent angle (inlet taper angle ) and divergent angle(outlet taper angle [ .) with the help of Numerical analysis through ANSYS 19.0(R3). Deciding parameter for greater thrust is angles i.e.; and [ . So, Convergent divergent nozzles with combustion chamber is designed, modelled and analysed numerical for getting optimal values of convergent and divergent angles . The various angles used are [35,15] [35,20] [35,25] [40,15] [40,20] [40,25] [45,15] [45,20] [45,25]. At these different angle the parameters such as velocity, temperature, pressure and Mach numbers are estimated with the help of steady state fluent with carbon dioxide as inlet to the combustion chamber.

Preliminary Studies on Non-Reactive Flow Vortex Cooling

2019 ◽  
Vol 12 (3) ◽  
pp. 262-271
Author(s):  
T.N. Rajesh ◽  
T.J.S. Jothi ◽  
T. Jayachandran
Keyword(s):  
Combustion Chamber ◽  
Inner Core ◽  
Rocket Engine ◽  
Injection Pressure ◽  
Chamber Pressure ◽  
Reactive Flow ◽  
Stoichiometric Ratio ◽  
Gaseous Oxygen ◽  
Mixture Ratio ◽  
Vortex Combustion

Background: The impulse for the propulsion of a rocket engine is obtained from the combustion of propellant mixture inside the combustion chamber and as the plume exhausts through a convergent- divergent nozzle. At stoichiometric ratio, the temperature inside the combustion chamber can be as high as 3500K. Thus, effective cooling of the thrust chamber becomes an essential criterion while designing a rocket engine. Objective: A new cooling method of thrust chambers was introduced by Chiaverni, which is termed as Vortex Combustion Cold-Wall Chamber (VCCW). The patent works on cyclone separators and confined vortex flow mechanism for providing high propellant mixing with improved degree of turbulence inside the combustion chamber, providing the required notion for studies on VCCW. The flow inside a VCCW has a complex structure characterised by axial pressure losses, swirl velocities, centrifugal force, flow reversal and strong turbulence. In order to study the flow phenomenon, both the experimental and numerical investigations are carried out. Methods: In this study, non-reactive flow analysis was conducted with real propellants like gaseous oxygen and hydrogen. The test was conducted to analyse the influence of mixture ratio and injection pressure of the propellants on the chamber pressure in a vortex combustion chamber. A vortex combustor was designed in which the oxidiser injected tangentially at the aft end near the nozzle spiraled up to the top plate and formed an inner core inside the chamber. The fuel was injected radially from injectors provided near the top plate and the propellants were mixed in the inner core. This resulted in enhanced mixing and increased residence time for the fuel. More information on the flow behaviour has been obtained by numerical analysis in Fluent. The test also investigated the sensitivity of the tangential injection pressure on the chamber pressure development. Results: All the test cases showed an increase in chamber pressure with the mixture ratio and injection pressure of the propellants. The maximum chamber pressure was found to be 3.8 bar at PC1 and 2.7 bar at PC2 when oxidiser to fuel ratio was 6.87. There was a reduction in chamber pressure of 1.1 bar and 0.7 bar at PC1 and PC2, respectively, in both the cases when hydrogen was injected. A small variation in the pressure of the propellant injected tangentially made a pronounced effect on the chamber pressure and hence vortex combustion chamber was found to be very sensitive to the tangential injection pressure. Conclusion: VCCW mechanism has been to be found to be very effective for keeping the chamber surface within the permissible limit and also reducing the payload of the space vehicle.

Steady-state Characterization of LLC-based Single-Stage AC/DC Converter Based on Numerical Analysis

2021 ◽  
pp. 1-1
Author(s):  
Xiang Li ◽  
Liuniu Guo ◽  
Tianchen Lang ◽  
Daorong Lu ◽  
Khalil Alluhaybi ◽  
...  
Keyword(s):  
Steady State ◽  
Numerical Analysis ◽  
Single Stage

Non-symmetric flow in Laval type nozzles

1972 ◽  
Vol 273 (1232) ◽  
pp. 185-235 ◽  
Keyword(s):  
Steady State ◽  
Combustion Chamber ◽  
Method Of Characteristics ◽  
Three Dimensions ◽  
Symmetric Flow ◽  
Gaseous Flow ◽  
The Method Of Characteristics ◽  
Lateral Forces

The equations of the steady state, compressible inviscid gaseous flow are linearized in a form suitable for application to nozzles of the Laval type. The procedure in the supersonic phase is verified by comparing solutions so obtained with those derived by the method of characteristics in two and three dimensions. Likewise, the solutions in the transonic phase are com pared with those obtained by other investigators. The linearized equation is then used to investigate the nat re of non-symmetric flow in rocket nozzles. It is found that if the flow from the combustion chamber into the nozzle is non-symmetric, the magnitude and direction of the turning couple produced by the emergent jet is dependent on the profile of the nozzle and it is possible to design profiles such that the turning couples or lateral forces are zero. The optimum nozzle so designed is independent of the pressure and also of the magnitude of the non-symmetry of the entry flow. The formulae by which they are obtained have been checked by extensive static and projection tests with simulated rocket test vehicles which are described in this paper.

Computational study of unsteady gas flow in the combustion chamber of a ramjet engine with heat transfer

2021 ◽  
pp. 50-61
Author(s):  
Надежда Петровна Скибина
Keyword(s):  
Heat Transfer ◽  
Numerical Simulation ◽  
Supersonic Flow ◽  
Combustion Chamber ◽  
Gas Flow ◽  
Stokes Equations ◽  
Heat Conducting ◽  
Measuring Device ◽  
Ramjet Engine ◽  
Mach Numbers

Проведено численное исследование нестационарного турбулентного сверхзвукового течения в камере сгорания прямоточного воздушно-реактивного двигателя. Описана методика экспериментального измерения температуры на стенке осесимметричного канала в камере сгорания двигателя. Математическое моделирование обтекания исследуемой модели двигателя проводилось для скоростей набегающего потока M = 5 ... 7. Начальные и граничные условия задачи соответствовали реальному аэродинамическому эксперименту. Проанализированы результаты численного расчета. Рассмотрено изменение распределения температуры вдоль стенки канала с течением времени. Проведена оценка согласованности полученных экспериментальных данных с результатами математического моделирования. Purpose. The aim of this study is a numerical simulation of unsteady supersonic gas flow in a working path of ramjet engine under conditions identical to aerodynamic tests. Free stream velocity corresponding to Mach numbers M=5 ... 7 are considered. Methodology. Presented study addresses the methods of physical and numerical simulation. The probing device for thermometric that allows to recording the temperature values along the wall of internal duct was proposed. To describe the motion of a viscous heat-conducting gas the unsteady Reynolds averaged Navier - Stokes equations are considered. The flow turbulence is accounted by the modified SST model. The problem was solved in ANSYS Fluent using finite-volume method. The initial and boundary conditions for unsteady calculation are set according to conditions of real aerodynamic tests. The coupled heat transfer for supersonic flow and elements of ramjet engine model are realized by setting of thermophysical properties of materials. The reliability testing of numerical simulation has been made to compare the results of calculations and the data of thermometric experimental tests. Findings. Numerical simulation of aerodynamic tests for ramjet engine was carried out. The agreement between the results of numerical calculations and experimental measurements for the velocity in the channel under consideration was obtained; the error was shown to be 2%. The temperature values were obtained in the area of contact of the supersonic flow with the surface of the measuring device for the external incident flow velocities for Mach numbers M = 5 ... 7. The process of heating the material in the channel that simulated the section of the engine combustion chamber was analyzed. The temperature distribution was studied depending on the position of the material layer under consideration relative to the contact zone with the flow. Value. In the course of the work, the fields of flow around the model of a ramjet engine were obtained, including the region of supersonic flow in the inner part of axisymmetric channel. The analysis of the temperature fields showed that to improve the quality of the results, it is necessary to take into account the depth of the calorimetric sensor. The obtained results will be used to estimate the time of interaction of the supersonic flow with the fuel surface required to reach the combustion temperature.

Effect of Divergence Angle on Flow Through Inlet Section of Diffuser of a Gas Turbine Combustion Chamber: The CFD Approach

2006 ◽  
Author(s):  
Digvijay B. Kulshreshtha ◽  
S. A. Channiwala ◽  
Jitendra Chaudhary ◽  
Zoeb Lakdawala ◽  
Hitesh Solanki ◽  
...  
Keyword(s):  
Combustion Chamber ◽  
Gas Turbine ◽  
Total Pressure ◽  
Divergence Angle ◽  
Velocity Head ◽  
Turbine Engine ◽  
Inlet Velocity ◽  
Pressure Losses ◽  
Flow Through ◽  
Mach Numbers

In the combustor inlet diffuser section of gas turbine engine, high-velocity air from compressor flows into the diffuser, where a considerable portion of the inlet velocity head PT3 − PS3 is converted to static pressure (PS) before the airflow enters the combustor. Modern high through-flow turbine engine compressors are highly loaded and usually have high inlet Mach numbers. With high compressor exit Mach numbers, the velocity head at the compressor exit station may be as high as 10% of the total pressure. The function of the diffuser is to recover a large proportion of this energy. Otherwise, the resulting higher total pressure loss would result in a significantly higher level of engine specific fuel consumption. The diffuser performance must also be sensitive to inlet velocity profiles and geometrical variations of the combustor relative to the location of the pre-diffuser exit flow path. Low diffuser pressure losses with high Mach numbers are more rapidly achieved with increasing length. However, diffuser length must be short to minimize engine length and weight. A good diffuser design should have a well considered balance between the confliction requirements for low pressure losses and short engine lengths. The present paper describes the effect of divergence angle on diffuser performance for gas turbine combustion chamber using Computational Fluid Dynamic Approach. The flow through the diffuser is numerically solved for divergence angles ranging from 5 to 25°. The flow separation and formation of wake regions are studied.

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