جستجوی مقالات مرتبط با کلیدواژه « combustion chamber » در نشریات گروه « مکانیک »

In this research, the design of a can combustion chamber for a ramjet engine, the performance of this chamber and the role of flameholder have been studied. For this purpose, after talking about ramjet engine types, some general information about the combustion chamber has been explained. Then the process of chamber designing and determining its geometry has been discussed. The dimensions of the chamber were determined by using input conditions extracted from GasTurb software as well as applying a calculation code, and the geometry of the chamber have been determined by applying calculation codes. By evaluating the obtained geometry and ensuring the accuracy of the design, simulation of combustion with non-premixed liquid phase was carried out using Fluent software. While presenting the results, the effects of size, distance and number of flameholder have been investigated. It is shown that the use of flameholder in ramjet engines is essential but the use of large flameholder is not recommended.

In present research the combustion process of solid fuel ramjet(SFRJ) was modeled in numerical state. Different five geometrical cases were employed for parametric study of back step height and aft-combustor length effects on propulsion properties in this numerical model. Results showed less than 10 percentage error in compare with reference data. Extracted numerical data from this research show decreasing burning rate, increasing combustion efficiency and thrust force due to increasing the step height also decreasing combustion efficiency and thrust force as results of decreasing the aft-combustor length. The results demonstrate that variation of the aft-combustor length hasn’t sensible effect on burning rate also optimum value for specific impulse occur when dimensionless aft-combustor length is equal to 0.95.

In this study, the combustion performance of C30 microturbine combustion chamber with biogas fuel with different mass fractions of CO2 has been analyzed and investigated. By assuming the periodic geometry and the two-stage reaction of fuel and oxide, the computational cost was reduced. To simulate the flow inside the chamber, three-dimensional Navier-Stokes equations and the k-ε turbulence model have been used to model the effects of turbulence. The flow inside the combustion chamber with different fuel components was analyzed with the EDDY DISSIPATION combustion model. To validate the solution and compare the results with the fabricated sample, pure CH4 was used as fuel. In this study, it was found that in the case of a constant mass flow rate of the consumed fuel, increasing the share of CO2 in the fuel with different mass fractions, due to the decrease in the calorific value created, caused a decrease in the production temperature, and it was also found that the use of biogas as a premix caused a decrease The amount of NOx becomes desirable. According to the obtained results, it is suggested that the amount of CO2 in the fuel is at most 10%, because the temperature has decreased by only 38 Kelvin and the amount of NOx in the output is 1.4 ppm.

The purpose of this paper is to analyze the combustion characteristics in the reactive flow inside an aero gas turbine model combustion chamber under different conditions of liquid fuel spray. In order to model the two-phase flow of fuel and gas droplets, the Euler-Lagrange approach and two-way connection between the gas phase and the droplets have been used. To solve the governing equations, the finite volume structured mesh has been used. The Reynolds Navier Stokes (RANS) averaging approach and discrete ordinate model have been used to simulate flow turbulence and radiative heat transfer. The steady flamelet reaction model and the probability density function have been employed to model the combustion and interaction between turbulent and reactive flow, respectively. Also, a chemical mechanism with 26 reactions and 17 species has been used to model the combustion of Jet-A fuel. NOx modeling has been done as a post-processing by the finite rate model. The results indicate that increasing the fuel spray velocity promotes combustion in the initial region and higher flame temperature, and finally increases the amount of NO pollutant. The smaller spray cone angle causes the droplets resulting from the spray will be concentrated in a smaller area and will reduce the NO pollutant.

The field has long acknowledged the detrimental effects of entropy waves, including heightened NOx emissions, combustion chamber instability, and noise production. Entropy oscillations, when accelerated, serve as a source of indirect noise, exacerbating the combustion chamber's instability. A computational study delved into the entropy noise within a lean-premixed H2/ ethylene burner to better comprehend these phenomena. Utilizing flamelet and large Eddy Simulation, the study shed light on the behavior of entropy waves and their impact on combustion dynamics. It particularly examined how various thermal and fluid conditions affect entropy noise in both thermally convective and adiabatic combustion chambers, focusing on temperature. In-depth analysis was conducted on elements like the strength of inlet turbulence, stoichiometric ratio, and the preheating level of the unburned mixture. The findings revealed intriguing connections between these factors and the acoustic system response. Notably, an increase in equivalence ratio, and temperature led to a heightened auditory response, suggesting that such conditions foster the generation and spread of entropy noise. Conversely, the study found that high turbulence intensity at the amplifier negatively impacted the transmission of entropy oscillations, thereby reducing the auditory response. This suggests that intense turbulence can interfere with entropy wave propagation, lessening their potential to produce significant acoustic effects.

In this research, the performance of ramjet engine combustion chamber has been evaluated by applying geometrical and physical changes such as changing the diameter of the chamber, changing the size and displacement of the secondary and dilution air holes and changing the mass flow passing through these holes. To do this, at first some typical ramjet combustion chambers were introduced and some of their features were discussed. Then using the information and guidance presented in scientific and technical references a suitable combustion chamber was designed for the defined inlet and outlet conditions. After the geometry has been determined, the simulation process has been carried out in Fluent software and while presenting the results, the performance of the chamber has been investigated. The results show that the values obtained for the geometry of the chamber and the position of the holes are reliable and changing the sizes has no global advantage and is not recommended. Furthermore, it is found that displacing air holes location could improve solution convergence.
 
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This study investigated the effect of grain fuel composition in the ramjet combustion chamber on the ignition delay time and solid fuel regression rate. In this article, a numerical investigation of ignition and stabilization of combustion of a new design of an engine equipped with solid fuel is done. The rod design consists of two solid fuels, maintaining the simple design of the classic ramjet. Numerical simulations of unsteady, turbulent, reactive and disintegration due to solid fuel by simultaneously solving combustion and turbulence models, applying numerical methods in the analysis of the governing equations that are the basis for obtaining some results including the effect of variables, functional characteristics and characteristics Combustion and finally access to the ignition delay time of fuel-rich propellant in solid fuel ramjet systems. In examining the chemical reactions of the combustion chamber, heavy polyethylene is taken into consideration and finally the ignition and combustion process continues. In line with the diagnosis for the correct prediction of the ignition delay time in the two mentioned geometries, the simulation results were in good agreement with the experimental data of the comparison and the selected model. The ignition delay time for the rod geometry was 0.28 seconds, for the classic geometry it was 0.43 seconds. And the design with the rod has increased the fuel regression rate.
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Due to the complexity of the processes governing the flow in the gas turbine, it is necessary to perform experimental tests to ensure the performance, but creating the conditions governing the components in the operational conditions is very complex. most of the combustion chamber tests are performed in atmospheric conditions and then the results are generalized to the operational conditions. This work, along with the lack of test stands with real operational conditions, has been developed based on the fact that most of the processes governing the behavior of the chamber have not undergone significant changes with the change of operating pressure, and especially if the shape of the flame does not change, it is possible to study the behavior of the chamber even at low pressure. In the present research, this issue has been numerically investigated for a real combustion chamber and it has been observed that changes are characterized by increasing pressure. Also, NOx pollutant increases at higher pressures. The overall flow structures were not affected by the pressure increase, but the pressure loss rate increased. The current results show that in the desired combustion chamber, the atmospheric test can provide acceptable results and can be generalized to high pressure conditions.

A numerical study on combustion phenomenon in a scramjet combustion chamber with fuel injection from different wedge angles (11, 12, 17 and 20 degrees) was done. In this combustion chamber, the Mach numbers of inflow air and hydrogen fuel are 2 and 1 respectively. The considered turbulent flow field was simulated by RANS equations in steady state form. In the present simulations, a realizable k   turbulence model was selected for turbulence simulation and also finite-rate/eddy-dissipation model was used for combustion simulation. A comparison was done between numerical and experimental results and accuracy of numerical method verified. These simulations were performed in Ansys Fluent commercial software. Results showed that with increase of wedge angle, the combustion efficiency rises from 63 to 67 percent. On the other hand, an increase in wedge angle results in stronger shock waves and also total pressure loss increases. Therefore, a compromise should be done by designer from both sides to get an efficient thrust.

In the present study, the effect of using a swirl nozzle and increasing the pressure of liquid fuel injection on the performance and emission of soot, nitrogen oxide and carbon monoxide combustion chamber of a propulsion system is investigated. The proposed solutions can change the intensity of cavitation and fuel spray characteristics.The effect of each of the proposed solutions with the help of a three-dimensional numerical model in EVL Fire software, which has been validated with experimental data in each section, on the performance of the diesel engine and its emissions has been investigated.The numerical results show that the creation of swirl inside the nozzle leads to a rotational flow of the fuel injector, an increase in the spray cone angle and the turbulence intensity inside the combustion chamber. Increasing the resulting spray cone angle improves the performance of the diesel engine by reducing fuel consumption and increasing power and torque, and the production of nitrogen oxide and carbon monoxide pollutants due to a better fuel-air mixture, increased turbulence intensity and average temperature inside the chamber.Combustion is reduced to an appropriate amount. Increasing the injection pressure can also improve the performance of the diesel engine.One of the negative points of increasing fuel injection pressure is the increase in nitrogen oxide production.By increasing the fuel injection pressure from 1350 bar to 2100 bar, fuel consumption decreases by 38% and its power and torque increase. Also in this case, carbon monoxide pollutants decrease by 65% and nitrogen oxide pollutants increase by 20%.

Heat transfer plays a major role in the design and performance of missiles and vehicles that have rocket driving. The main importance of this discussion is related to the safety range missile materials, especially in areas such as the combustion chamber and nozzle, which are under critical heat. Only one failure is enough to interfere with the operation of missiles. It may also be due to excessive heat in the nozzle throat, a portion of the sheet metal locally damaged. In which case the exhaust gas flow becomes asymmetrically, which led to the creation of inappropriate force vector direction. This leads to an error in the missile guidance to the target and a lot of fuel consumed on the spacecraft to correct its path by secondary driving system. In this project the amount of heat transferred to the chamber for different sections of engine during the simulation by software is calculated, also the coolant conditions and temperatures have been calculated in accordance with existing heat and cooling operation performed in each section. For this purpose, the engine divided into four parts, the combustion chamber, the convergent nozzle, throat nozzle and the divergent nozzle. The rate of heat transferred from the gas chamber (according to the conditions of gas in different sections) as well as the chamber temperature in the vicinity of the coolant and coolant temperature calculated for different sections of each section. Notably, in this research, ANSYS software package is used to create geometry, mesh and simulations.

In the design of the combustion chamber, various parameters should be considered. These parameters include uniform temperature distribution at the outlet of the chamber, more flame stability, lower pollution, higher combustion efficiency, lower wall temperature, and lower pressure drop in the chamber. Regarding to the complex condition of the flow in the combustion chamber due to the various effects of turbulence and mixing of flows as well as the behavior of turbulent flames, predicting the performance of flow in the combustion chambers is very complicated. In this paper, it is tried to study and optimize the combustion chamber of Amirkabir University of Technology in terms of swirler. It is done by using the numerical method and finally the selected swirler in the numerical method is tested in the experimental setup to investigate optimization method .According to the studies, swirler with an angle of 60 degrees, 12 blades, and a thickness of 0.75 mm is selected as the final case. In the experimental results, the amount of CO pollution has significantly reduced. The output temperature, the pattern factor and unburned hydrocarbon have reduced in the final case. However, the temperature uniformity inside the chamber has increased.

V94.2 gas turbine is one of the most widely used gas turbines for power generation in Iran. In this study, a brief literature review of damage analysis and life estimation of V94.2 combustor is presented accompanied with aerothermal and thermomechanical analyses in the real operation condition. Since creep failure is one of the limiting factors in the life time of combustion chambers, creep life prediction is insvestigated in this paper. Creep time rupture prediction is one of the most important steps of the creep life calculation. Therefore, first, aerothermal and structural analyses are performed considering the effect of the creep failure on the combustion chamber. Next, creep time rupture is calculated using two well-known methods Larson-Miller and Sherby-Dorn parameters. These parameters for IN617 are extracted as a function of stress based on creep experimental data. The accumulated creep damages of 40’000 hours of full-load operation in the combustion chamber are reported for both above methods. Critical nodes limiting the component life are diagnosed from the detected critical creep damage region. According to the results, the creep rupture time prediction technique plays a vital role in the specification of critical creep damage region, temporal damage variation and consequently, creep life estimation. Therefore, the calculated damage from one technique can be three times greater than the results from the other technique.

Extensive research has been done to improve the quality and reduce the propulsion weight of aircraft, which has led to the invention of new technologies and the construction of new engines. In all these studies, the study of temperature distribution and heat transfer in the body of the combustion engine in order to evaluate the characteristics and performance of the engine is considered important and necessary. One of the most innovative designs is the vane rotary motor design, which is considered to be a revolution in the motor industry. In this article, initial design of the seals of this engine is discussed based on thermal analysis. The method of increasing the air temperature in the combustion chamber has been investigated and the temperature distribution has been calculated by two methods of engineering estimation and numerical analysis. In engineering estimation, the basic relations of heat transfer, radiative, conduction, and energy balance equation are used, and the energy survival equations are satisfied, and the combustion process is modeled as a diesel cycle. Combustion is considered as a heat source and heat transfer coefficients are calculated in the form of computer code. The results of numerical analysis of this issue are compared and validated with the results of the estimation program and the appropriate cover and material for the components of the combustion chamber as well as the appropriate material and location of the seals are selected according to the temperature distribution

This numerical research conducted using CONVERGE Computational Fluid Dynamic (CFD) code and devoted to assessing the simultaneous and separate impacts of Diesel Direct Injection Timing (DDIT) (16 to 6 Crank Angle (CA) Before Top Dead Center (BTDC) with 2 CA steps), combustion chamber geometry (re-entrant (baseline), cylindrical, and wide-shallow chamber), and applying syngas (20 and 40% of total energy per cycle) in a heavy-duty off-road RCCI engine. In the case of combustion simulation, the SAGE combustion model was used coupled with a detailed chemical kinetic mechanism consist of 72 species and 360 reactions. Results showed that under baseline operating conditions (DDIT of 10 CA BTDC and using re-entrant piston bowl) increasing the syngas to diesel ratio up to 40% caused a 3.4% rise in heat transfer loss and simultaneous reduction in Nitrogen Oxides (NOx) about 12%, Particulate Matter (PM) up to 88%, and Hydro-Carbons (HCs) nearly 82% compared to Pure Diesel Combustion (PDC) conditions. Besides, utilizing the wide-shallow combustion chamber along with diesel injection at 16 CA BTDC at diesel- 40% syngas combustion operating conditions led to the increment of heat transfer loss (7%), combustion loss (2.5%), and also, simultaneous reduction of NOx (3%), PM (37%), HC (62%), and gross indicated efficiency (4.7%) compared to baseline PDC case.

In this study, the evaporation of liquid oxidizer inside a combustion chamber during engine start-up in a low-pressure environment is numerically investigated. In thruster starting stage, first, a portion of oxidizer which has been evaporated inside the injector capillaries fills the void inside the combustion chamber and causes a pressure rise (up to 0.2 bars). This makes the injection of oxidizer as fluid possible and now the evaporation rate can be investigated. In the investigated thruster the mass flow of injected liquid is 3 grams per second and the type of injector is pressure-swirl. Numerical simulation in this study is based on an Eulerian- Lagrangian method known as Discrete Phase Method (DPM), which investigates the interaction of two phases using Navier-Stokes Equations. A verification is done to support the results of the method used in order to obtain quantitative variables essential to this study. The tendency of fluids to reach a stable state after an abrupt process of flashing is visible in the results of this study. After a small period of time around 20 milliseconds a stable temperature around 264 kelvins is reached which causes a stable pressure of 7 kPa in the combustion chamber.

Combustion chambers are one of the main parts of the power generation systems such as gas turbines and internal combustion engines and affect their efficiency and environmental pollutions. To reduce the high amount of the pollutions e.g. NOx, lean premixed combustion was introduced to be used instead of traditional non-premixed flames, however, this method has more tendency to become unstable. In fact, the thermal and acoustics interactions in the combustion chambers can amplify the acoustic waves and thereby produce noise and increase the vibration level of the liner. The continuation of large amplitude vibrations can lead to failure due to fatigue. Therefore, the vibration modeling of the liner is very important. In the present research, the vibration of a liner in a laboratory combustion chamber is investigated. First step to model the vibration of the liner is to extract its modal parameters in the cold and hot states. Finite element (FE) model updating is utilized to modify the initial FE model of the liner based on the experimental modal parameters. Then, using the computational fluid dynamics (CFD), the flow analysis is performed to obtain the pressure and velocity fluctuations during the analysis time. These data are then used to model the flame as a monopole acoustic source. To find the structural response of the liner due to this acoustic source, the transient analysis is performed. The results show the effectiveness of the updated FE model to predict the modal parameters and the vibration amplitude of the liner with acceptable accuracy.

In this research, the combustion chamber of the liquid propellant engine has been simulated numerically by the computational fluid dynamics (CFD) method. After the simulation, several methods have been proposed to improve the combustion. These include using baffle inside the combustion chamber, increasing the equilibrium ratio and the nano catalyst. In each of these methods, the combustion temperature, mass fraction of fuel and oxidizer, mass fraction of combustion products and mass fraction of pollutants were calculated and compared with a simple chamber. The results showed that using these methods increases the combustion heat by 28.36%, reduces fuel mass fraction by 27.91%, increases mass fraction of complete combustion products including water and nitrogen, and reduces NOx pollutant mass fraction by 26.85 as it is known that NOx pollutant is a product of incomplete combustion and generally reducing it results in better combustion.

Combustion of a mixture of inhomogeneity reactants is a type of combustion in which the mixing of fuel and air is not complete. This study investigated the numerical effect of the pilot location on the turbulent jet flame lift off height with different levels of inhomogeneity in the combustion chamber. The effect of different equevalence ratios has also been investigated. In this numerical study, the configuration modeling of Navier Stokes equations using Reynolds averaging method, standard k-ε turbulence and reactive flow modeling of Eddy Dissipation Concept (EDC) method have been used. Observations show that the flame lift off height decreases with increasing inhomogeneity length and the flame attach during a certain inhomogeneity length, which depends on the equevalence ratio and pilot location. In all of the cases studied, a range of inhomogeneity lengths was found during which the flame was attached. The results show that the flame lift off height in nonpremixed mode is higher than in the case where the fuel-air mixture ignites almost premixed in the chamber. The higher the equevalence ratio, the lower the flame lift off height. Based on the three locations studied in the side wall for the pilot location and the results obtained, the pilot in the middle area of the side wall had a lower flame lift off height and a higher attached flame range. Also, the flame area inside the chamber changes as the inhomogeneity changes.