Search results for: Supersonic combustor.

Authors: R. Kanmaniraja, R. Freshipali, J. Abdullah, K. Niranjan, K. Balasubramani, V. R. Sanal Kumar

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The present environmental issues have made aircraft jet noise reduction a crucial problem in aero-acoustics research. Acoustic studies reveal that addition of chevrons to the nozzle reduces the sound pressure level reasonably with acceptable reduction in performance. In this paper comprehensive numerical studies on acoustic characteristics of different types of chevron nozzles have been carried out with non-reacting flows for the shape optimization of chevrons in supersonic nozzles for aerospace applications. The numerical studies have been carried out using a validated steady 3D density based, k-ε turbulence model. In this paper chevron with sharp edge, flat edge, round edge and U-type edge are selected for the jet acoustic characterization of supersonic nozzles. We observed that compared to the base model a case with round-shaped chevron nozzle could reduce 4.13% acoustic level with 0.6% thrust loss. We concluded that the prudent selection of the chevron shape will enable an appreciable reduction of the aircraft jet noise without compromising its overall performance. It is evident from the present numerical simulations that k-ε model can predict reasonably well the acoustic level of chevron supersonic nozzles for its shape optimization.

Keywords: Supersonic nozzle, Chevron, Acoustic level, Shape Optimization of Chevron Nozzles, Jet noise suppression.

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Authors: Y. B. Galerkin, E. Y. Popova, K. V. Soldatova

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There is an evident trend to elevate pressure ratio of a single stage of a turbo compressors - axial compressors in particular. Whilst there was an opinion recently that a pressure ratio 1,9 was a reasonable limit, later appeared information on successful modeling tested of stages with pressure ratio up to 2,8. The authors recon that lack of information on high pressure stages makes actual a study of rational choice of design parameters before high supersonic flow problems solving. The computer program of an engineering type was developed. Below is presented a sample of its application to study possible parameters of the impeller of the stage with pressure ratio 3,0. Influence of two main design parameters on expected efficiency, periphery blade speed and flow structure is demonstrated. The results had lead to choose a variant for further analysis and improvement by CFD methods.

Keywords: Supersonic stage, impeller, efficiency, flow rate coefficient, work coefficient, loss coefficient, oblique shock, direct shock.

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Authors: Sang-Lae Lee, Ji-Hwan Kim

Abstract:

In this study, it is investigated the stability boundary of Functionally Graded (FG) panel under the heats and supersonic airflows. Material properties are assumed to be temperature dependent, and a simple power law distribution is taken. First-order shear deformation theory (FSDT) of plate is applied to model the panel, and the von-Karman strain- displacement relations are adopted to consider the geometric nonlinearity due to large deformation. Further, the first-order piston theory is used to model the supersonic aerodynamic load acting on a panel and Rayleigh damping coefficient is used to present the structural damping. In order to find a critical value of the speed, linear flutter analysis of FG panels is performed. Numerical results are compared with the previous works, and present results for the temperature dependent material are discussed in detail for stability boundary of the panel with various volume fractions, and aerodynamic pressures.

Keywords: Functionally graded panels, Linear flutter analysis, Supersonic airflows, Temperature dependent material property.

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Authors: R. Haoui

Abstract:

The aim of this work is to determine the supersonic nozzle profiles used in propulsion, for the launchers or embarked with the satellites. This design has as a role firstly, to give a important propulsion, i.e. with uniform and parallel flow at exit, secondly to find a short length profiles without modification of the flow in the nozzle. The first elaborate program is used to determine the profile of divergent by using the characteristics method for an axisymmetric flow. The second program is conceived by using the finite volume method to determine and test the profile found connected to a convergent.

Keywords: Characteristic method, nozzles, supersonic flow, propellers.

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Authors: M. Mojtahedpoor, M. M. Doustdar

Abstract:

A numerical study on the effect of side-dump angle on fuel droplets sizing and effective mass fraction have been investigated in present paper. The mass of fuel vapor inside the flammability limit is named as the effective mass fraction. In the first step we have considered a side-dump combustor with dump angle of 0o (acrossthe cylinder) and by increasing the entrance airflow velocity from 20 to 30, 40 and 50 (m/s) respectively, the mean diameter of fuel droplets sizing and effective mass fraction have been studied. After this step, we have changed the dump angle from 0o to 30o,45o and finally 60o in direction of cylinderand also we have increased the entrance airflow velocity from 20 up to 50 (m/s) with the amount of growth of 10(m/s) in each step, to examine its effects on fuel droplets sizing as well as effective mass fraction. With rise of entrance airflow velocity, these calculations are repeated in each step too. The results show, with growth of dump-angle the effective mass fraction has been decreased and the mean diameter of droplets sizing has been increased. To fulfill the calculations a modified version of KIVA-3V code which is a transient, three-dimensional, multiphase, multicomponent code for the analysis of chemically reacting flows with sprays, is used.

Keywords: Side-Dump combustor, Droplets sizing, Side-Dump angle, KIVA-3V

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Authors: R. Kamali, A. R. Binesh, S. Hossainpour

Abstract:

To understand working features of a micro combustor, a computer code has been developed to study combustion of hydrogen–air mixture in a series of chambers with same shape aspect ratio but various dimensions from millimeter to micrometer level. The prepared algorithm and the computer code are capable of modeling mixture effects in different fluid flows including chemical reactions, viscous and mass diffusion effects. The effect of various heat transfer conditions at chamber wall, e.g. adiabatic wall, with heat loss and heat conduction within the wall, on the combustion is analyzed. These thermal conditions have strong effects on the combustion especially when the chamber dimension goes smaller and the ratio of surface area to volume becomes larger. Both factors, such as larger heat loss through the chamber wall and smaller chamber dimension size, may lead to the thermal quenching of micro-scale combustion. Through such systematic numerical analysis, a proper operation space for the micro-combustor is suggested, which may be used as the guideline for microcombustor design. In addition, the results reported in this paper illustrate that the numerical simulation can be one of the most powerful and beneficial tools for the micro-combustor design, optimization and performance analysis.

Keywords: Numerical simulation, Micro-combustion, MEMS, CFD, Chemical reaction.

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Authors: Vladislav A. Nazukin, Valery G. Avgustinovich, Vakhtang V. Tsatiashvili

Abstract:

The most important part of modern lean low NOx combustors is a premixer where swirlers are often used for intensification of mixing processes and further formation of required flow pattern in combustor liner. Swirling flow leads to formation of complex eddy structures causing flow perturbations. It is able to cause combustion instability. Therefore, at design phase, it is necessary to pay great attention to aerodynamics of premixers. Analysis based on unsteady CFD modeling of swirling flow in production combustor swirler showed presence of large number of different eddy structures that can be conditionally divided into three types relative to its location of origin and a propagation path. Further, features of each eddy type were subsequently defined. Comparison of calculated and experimental pressure fluctuations spectrums verified correctness of computations.

Keywords: DES simulation, swirler, vortical structures.

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Authors: Lei Zhang, Wen-jun Ruan, Hao Wang, Peng-xin Wang

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The source of the jet noise is generated by rocket exhaust plume during rocket engine testing. A domain decomposition approach is applied to the jet noise prediction in this paper. The aerodynamic noise coupling is based on the splitting into acoustic sources generation and sound propagation in separate physical domains. Large Eddy Simulation (LES) is used to simulate the supersonic jet flow. Based on the simulation results of the flow-fields, the jet noise distribution of the sound pressure level is obtained by applying the Ffowcs Williams-Hawkings (FW-H) acoustics equation and Fourier transform. The calculation results show that the complex structures of expansion waves, compression waves and the turbulent boundary layer could occur due to the strong interaction between the gas jet and the ambient air. In addition, the jet core region, the shock cell and the sound pressure level of the gas jet increase with the nozzle size increasing. Importantly, the numerical simulation results of the far-field sound are in good agreement with the experimental measurements in directivity.

Keywords: Supersonic gas jet, Large Eddy Simulation(LES), acoustic noise, Ffowcs Williams-Hawkings (FW-H) equations, nozzle size.

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Authors: Mohammad Reza Soltani, Mohammad Farahani, Javad Sepahi Younsi

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Heat source addition to the axisymmetric supersonic inlet may improve the performance parameters, which will increase the inlet efficiency. In this investigation the heat has been added to the flow field at some distance ahead of an axisymmetric inlet by adding an imaginary thermal source upstream of cowl lip. The effect of heat addition on the drag coefficient, mass flow rate and the overall efficiency of the inlet have been investigated. The results show that heat addition causes flow separation, hence to prevent this phenomena, roughness has been added on the spike surface. However, heat addition reduces the drag coefficient and the inlet mass flow rate considerably. Furthermore, the effects of position, size, and shape on the inlet performance were studied. It is found that the thermal source deflects the flow streamlines. By improper location of the thermal source, the optimum condition has been obtained. For the optimum condition, the drag coefficient is considerably reduced and the inlet mass flow rate and its efficiency have been increased slightly. The optimum shape of the heat source is obtained too.

Keywords: Drag coefficient, heat source, performanceparameters, supersonic inlet.

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Authors: Guohua Tu, Zhi Fu, Zhiwei Hu, Neil D Sandham, Jianqiang Chen

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Tripping of boundary-layers from laminar to turbulent flow, which may be necessary in specific practical applications, requires high amplitude disturbances to be introduced into the boundary layers without large drag penalties. As a possible improvement on fixed trip devices, a technique based on vortex shedding for enhancing supersonic flow transition is demonstrated in the present paper for a Mach 1.5 boundary layer. The compressible Navier-Stokes equations are solved directly using a high-order (fifth-order in space and third-order in time) finite difference method for small-scale cylinders suspended transversely near the wall. For cylinders with proper diameter and mount location, asymmetry vortices shed within the boundary layer are capable of tripping laminar-turbulent transition. Full three-dimensional simulations showed that transition was enhanced. A parametric study of the size and mounting location of the cylinder is carried out to identify the most effective setup. It is also found that the vortex shedding can be suppressed by some factors such as wall effect.

Keywords: Boundary layer instability, boundary layer transition, vortex shedding, supersonic flows, flow control.

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Authors: Siddharth Ahuja, T. M. Muruganandam

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An analytical investigation is performed to study the unsteady response of a one-dimensional, non-linear dynamic shock system to external downstream pressure perturbations in a supersonic flow in a varying area duct. For a given pressure ratio across a wind tunnel, the normal shock's location can be computed as per one-dimensional steady gas dynamics. Similarly, for some other pressure ratio, the location of the normal shock will change accordingly, again computed using one-dimensional gas dynamics. This investigation focuses on the small-time interval between the first steady shock location and the new steady shock location (corresponding to different pressure ratios). In essence, this study aims to shed light on the motion of the shock from one steady location to another steady location. Further, this study aims to create the foundation of the Unsteady Gas Dynamics field enabling further insight in future research work. According to the new pressure ratio, a pressure pulse, generated at the exit of the tunnel which travels and perturbs the shock from its original position, setting it into motion. During such activity, other numerous physical phenomena also happen at the same time. However, three broad phenomena have been focused on, in this study - Traversal of a Wave, Fluid Element Interactions and Wave Interactions. The above mentioned three phenomena create, alter and kill numerous waves for different conditions. The waves which are created by the above-mentioned phenomena eventually interact with the shock and set it into motion. Numerous such interactions with the shock will slowly make it settle into its final position owing to the new pressure ratio across the duct, as estimated by one-dimensional gas dynamics. This analysis will be extremely helpful in the prediction of inlet 'unstart' of the flow in a supersonic engine intake and its prominence with the incoming flow Mach number, incoming flow pressure and the external perturbation pressure is also studied to help design more efficient supersonic intakes for engines like ramjets and scramjets.

Keywords: Analytical investigation, compression and expansion waves, fluid element interactions, shock trajectory, supersonic flow, unsteady gas dynamics, varying area duct, wave interactions.

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Authors: Mushfiqul Alam, Kashyapa Narenathreyas

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The demand for efficient transonic transport has been growing every day and may turn out to be the most pressed innovation in coming years. Oblique wing configuration was proposed as an alternative to conventional wing configuration for supersonic and transonic passenger aircraft due to its aerodynamic advantages. This paper re-demonstrates the aerodynamic advantages of oblique wing configuration using open source CFD code. The aerodynamic data were generated using Panel Method. Results show that Oblique Wing concept with elliptical wing planform offers a significant reduction in drag at transonic and supersonic speeds and approximately twice the lift distribution compared to conventional operating aircrafts. The paper also presents a preliminary conceptual aircraft sizing which can be used for further experimental analysis.

Keywords: Aerodynamics, asymmetric sweep, oblique wing, swing wing.

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Authors: Selvakuma Kumaresh, Man Young Kim

Abstract:

Combustion phenomenon will be accomplished effectively by the development of low emission combustor. One of the significant factors influencing the entire Combustion process is the mixing between a swirling angular jet (Primary Air) and the non-swirling inner jet (fuel). To study this fundamental flow, the chamber had to be designed in such a manner that the combustion process to sustain itself in a continuous manner and the temperature of the products is sufficiently below the maximum working temperature in the turbine. This study is used to develop the effective combustion with low unburned combustion products by adopting the concept of high swirl flow and motility of holes in the secondary chamber. The proper selection of a swirler is needed to reduce emission which can be concluded from the emission of Nox and CO2. The capture of CO2 is necessary to mitigate CO2 emissions from natural gas. Thus the suppression of unburned gases is a meaningful objective for the development of high performance combustor without affecting turbine blade temperature.

Keywords: Combustion, Emission, Can-type Combustion Chamber, CFD, Motility of Holes, Swirl Flow.

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Authors: Ibrahim Mnafeg, Azgal Abichou, Lotfi Beji

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Traditional mechanical control systems in thrust vectoring are efficient in rocket thrust guidance but their costs and their weights are excessive. The fluidic injection in the nozzle divergent constitutes an alternative procedure to achieve the goal. In this paper, we present a 3D analytical model for fluidic injection in a supersonic nozzle integrating an orifice. The fluidic vectoring uses a sonic secondary injection in the divergent. As a result, the flow and interaction between the main and secondary jet has built in order to express the pressure fields from which the forces and thrust vectoring are deduced. Under various separation criteria, the present analytical model results are compared with the existing numerical and experimental data from the literature.

Keywords: Flow separation, Fluidic thrust vectoring, Nozzle, Secondary jet, Shock wave.

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Authors: Man Young Kim

Abstract:

The objective of this work is to investigate the turbulent reacting flow in a three dimensional combustor with emphasis on the effect of inlet swirl flow through a numerical simulation. Flow field is analyzed using the SIMPLE method which is known as stable as well as accurate in the combustion modeling, and the finite volume method is adopted in solving the radiative transfer equation. In this work, the thermal and flow characteristics in a three dimensional combustor by changing parameters such as equivalence ratio and inlet swirl angle have investigated. As the equivalence ratio increases, which means that more fuel is supplied due to a larger inlet fuel velocity, the flame temperature increases and the location of maximum temperature has moved towards downstream. In the mean while, the existence of inlet swirl velocity makes the fuel and combustion air more completely mixed and burnt in short distance. Therefore, the locations of the maximum reaction rate and temperature were shifted to forward direction compared with the case of no swirl.

Keywords: Gaseous Fuel, Inlet Swirl, Thermal Radiation, Turbulent Combustion

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Authors: Rabah Haoui

Abstract:

The aim of this work is to analyze a viscous flow around the axisymmetric blunt body taken into account the mesh size both in the free stream and into the boundary layer. The resolution of the Navier-Stokes equations is realized by using the finite volume method to determine the flow parameters and detached shock position. The numerical technique uses the Flux Vector Splitting method of Van Leer. Here, adequate time stepping parameter, CFL coefficient and mesh size level are selected to ensure numerical convergence. The effect of the mesh size is significant on the shear stress and velocity profile. The best solution is obtained with using a very fine grid. This study enabled us to confirm that the determination of boundary layer thickness can be obtained only if the size of the mesh is lower than a certain value limits given by our calculations.

Keywords: Supersonic flow, viscous flow, finite volume, blunt body.

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Authors: Rabah Haoui

Abstract:

Hypersonic flows around spatial vehicles during their reentry phase in planetary atmospheres are characterized by intense aerothermal phenomena. The aim of this work is to analyze high temperature flows around an axisymmetric blunt body taking into account chemical and vibrational non-equilibrium for air mixture species. For this purpose, a finite volume methodology is employed to determine the supersonic flow parameters around the axisymmetric blunt body, especially at the stagnation point and along the wall of spacecraft for several altitudes. This allows the capture shock wave before a blunt body placed in supersonic free stream. The numerical technique uses the Flux Vector Splitting method of Van Leer. Here, adequate time stepping parameter, along with CFL coefficient and mesh size level are selected to ensure numerical convergence, sought with an order of 10-8

Keywords: Chemical kinetic, dissociation, finite volumes, frozen, hypersonic flow, non-equilibrium, Reactive flow, supersonicflow , vibration.

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Authors: Rabah Haoui

Abstract:

The aim of this work is to analyze a viscous flow in the axisymmetric nozzle taken into account the mesh size both in the free stream and into the boundary layer. The resolution of the Navier- Stokes equations is realized by using the finite volume method to determine the supersonic flow parameters at the exit of convergingdiverging nozzle. The numerical technique uses the Flux Vector Splitting method of Van Leer. Here, adequate time stepping parameter, along with CFL coefficient and mesh size level is selected to ensure numerical convergence. The effect of the boundary layer thickness is significant at the exit of the nozzle. The best solution is obtained with using a very fine grid, especially near the wall, where we have a strong variation of velocity, temperature and shear stress. This study enabled us to confirm that the determination of boundary layer thickness can be obtained only if the size of the mesh is lower than a certain value limits given by our calculations.

Keywords: Supersonic flow, viscous flow, finite volume, nozzle

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Authors: Muhammad Yamin Younis, Muhammad Amjad Sohail, Tawfiqur Rahman, Zaka Muhammad, Saifur Rahman Bakaul

Abstract:

The performance of Advection Upstream Splitting Method AUSM schemes are evaluated against experimental flow fields at different Mach numbers and results are compared with experimental data of subsonic, supersonic and hypersonic flow fields. The turbulent model used here is SST model by Menter. The numerical predictions include lift coefficient, drag coefficient and pitching moment coefficient at different mach numbers and angle of attacks. This work describes a computational study undertaken to compute the Aerodynamic characteristics of different air vehicles configurations using a structured Navier-Stokes computational technique. The CFD code bases on the idea of upwind scheme for the convective (convective-moving) fluxes. CFD results for GLC305 airfoil and cone cylinder tail fined missile calculated on above mentioned turbulence model are compared with the available data. Wide ranges of Mach number from subsonic to hypersonic speeds are simulated and results are compared. When the computation is done by using viscous turbulence model the above mentioned coefficients have a very good agreement with the experimental values. AUSM scheme is very efficient in the regions of very high pressure gradients like shock waves and discontinuities. The AUSM versions simulate the all types of flows from lower subsonic to hypersonic flow without oscillations.

Keywords: Subsonic, supersonic, Hypersonic, AUSM+, Drag Coefficient, lift Coefficient, Pitching moment coefficient, pressure Coefficient, turbulent flow.

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Authors: A. A. Amer, H. M. Gad, I. A. Ibrahim, S. I. Abdel-Mageed, T. M. Farag

Abstract:

This paper represents an experimental study of LPG diffusion flame at elevated preheated air temperatures. The flame is stabilized in a vertical water-cooled combustor by using air swirler. An experimental test rig was designed to investigate the different operating conditions. The burner head is designed so that the LPG fuel issued centrally and surrounded by the swirling air issues from an air swirler. There are three air swirlers having the same dimensions but having different blade angles to give different swirl numbers of 0.5, 0.87 and 1.5. The combustion air was heated electrically before entering the combustor up to a temperature about 500 K. Five air to fuel mass ratios of 15, 20, 30, 40 and 50 were also studied. The effect of preheated air temperature, swirl number and air to fuel mass ratios on the temperature maps, visible flame length, high temperature region (size) and exhaust species concentrations are studied. Some results show that as the preheated air temperature increases, the volume of high temperature region also increased but the flame length decreased. Increasing the preheated air temperature, EINOx, EICO2 and EIO2 increased, while EICO decreased. Increasing the preheated air temperature from 300 to 500 K, for all air swirl numbers used, the highest increase in EINOx, EICO2 and EIO2 are 141, 4 and 65%, respectively.

Keywords: Preheated air temperature, air swirler, flame length, emission index.

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Authors: A. Haddad, H. Kbab

Abstract:

A supersonic expansion cannot be achieved within a convergent-divergent nozzle if the flow velocity does not reach that of the sound at the throat. The computation of the flow field characteristics at the throat is thus essential to the nozzle developed thrust value and therefore to the aircraft or rocket it propels. Several approaches were developed in order to describe the transonic expansion, which takes place through the throat of a De-Laval convergent-divergent nozzle. They all allow reaching good results but showing a major shortcoming represented by their inability to describe the transonic flow field for nozzles having a small throat radius. The approach initially developed by Kliegel & Levine uses the velocity series development in terms of the normalized throat radius added to unity instead of solely the normalized throat radius or the traditional small disturbances theory approach. The present investigation carries out the application of these three approaches for different throat radiuses of curvature. The method using the normalized throat radius added to unity shows better results when applied to geometries integrating small throat radiuses.

Keywords: De-Laval nozzles, transonic calculations, transonic flow, supersonic nozzle.

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Authors: Hosein Faramarzpour

Abstract:

This work presents the optimum operational conditions for a hydrogen-based micro-scale power source, using a verified mathematical model including fluid dynamics and reaction kinetics. Thereafter, the stable operational flame regime is pursued as a key factor in optimizing the design of micro-combustors. The results show that with increasing velocities, four H2 flame regimes develop in the micro-combustor, namely: 1) periodic ignition-extinction regime, 2) steady symmetric regime, 3) pulsating asymmetric regime, and 4) steady asymmetric regime. The first regime that appears in 0.8 m/s inlet velocity is a periodic ignition-extinction regime which is characterized by counter flows and tulip-shape flames. For flow velocity above 0.2 m/s, the flame shifts downstream, and the combustion regime switches to a steady symmetric flame where temperature increases considerably due to the increased rate of incoming energy. Further elevation in flow velocity up to 1 m/s leads to the pulsating asymmetric flame formation, which is associated with pulses in various flame properties such as temperature and species concentration. Further elevation in flow velocity up to 1 m/s leads to the pulsating asymmetric flame formation, which is associated with pulses in various flame properties such as temperature and species concentration. Ultimately, when the inlet velocity reached 1.2 m/s, the last regime was observed, and a steady asymmetric regime appeared.

Keywords: Thermophotovoltaic generator, micro combustor, micro power generator, combustion regimes, flame dynamic.

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Authors: Sh.Khalilarya, H.Oryani, S.Jafarmadar, H.Khatamnezhad, A.Nemati

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Understanding of how and where NOx formation occurs in industrial burner is very important for efficient and clean operation of utility burners. Also the importance of this problem is mainly due to its relation to the pollutants produced by more burners used widely of gas turbine in thermal power plants and glass and steel industry. In this article, a numerical model of an industrial burner operating in MILD combustion is validated with experimental data.. Then influence of air flow rate and air temperature on combustor temperature profiles and NOX product are investigated. In order to modification this study reports on the effects of fuel and air dilution (with inert gases H2O, CO2, N2), and also influence of lean-premixed of fuel, on the temperature profiles and NOX emission. Conservation equations of mass, momentum and energy, and transport equations of species concentrations, turbulence, combustion and radiation modeling in addition to NO modeling equations were solved together to present temperature and NO distribution inside the burner. The results shows that dilution, cause to a reduction in value of temperature and NOX emission, and suppresses any flame propagation inside the furnace and made the flame inside the furnace invisible. Dilution with H2O rather than N2 and CO2 decreases further the value of the NOX. Also with raise of lean-premix level, local temperature of burner and the value of NOX product are decreases because of premixing prevents local “hot spots" within the combustor volume that can lead to significant NOx formation. Also leanpremixing of fuel with air cause to amount of air in reaction zone is reach more than amount that supplied as is actually needed to burn the fuel and this act lead to limiting NOx formation

Keywords: Mild combustion, Flameless, Numerical simulation, Burner, CFD.

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Authors: Payam Haghparast, Mikhail V. Sorin, Hakim Nesreddine

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The complex oblique shock phenomenon can be simply assumed as a normal shock at the constant area section to simulate a sharp pressure increase and velocity decrease in 1-D thermodynamic models. The assumed normal shock location is one of the greatest sources of error in ejector thermodynamic models. Most researchers consider an arbitrary location without justifying it. Our study compares the effect of normal shock place on ejector dimensions in 1-D models. To this aim, two different ejector experimental test benches, a constant area-mixing ejector (CAM) and a constant pressure-mixing (CPM) are considered, with different known geometries, operating conditions and working fluids (R245fa, R141b). In the first step, in order to evaluate the real value of the efficiencies in the different ejector parts and critical back pressure, a CFD model was built and validated by experimental data for two types of ejectors. These reference data are then used as input to the 1D model to calculate the lengths and the diameters of the ejectors. Afterwards, the design output geometry calculated by the 1D model is compared directly with the corresponding experimental geometry. It was found that there is a good agreement between the ejector dimensions obtained by the 1D model, for both CAM and CPM, with experimental ejector data. Furthermore, it is shown that normal shock place affects only the constant area length as it is proven that the inlet normal shock assumption results in more accurate length. Taking into account previous 1D models, the results suggest the use of the assumed normal shock location at the inlet of the constant area duct to design the supersonic ejectors.

Keywords: 1D model, constant area-mixing, constant pressure-mixing, normal shock location, ejector dimensions.

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Authors: Esam I. Jassim, Mohamed M. Awad

Abstract:

Natural gas flow contains undesirable solid particles, liquid condensation, and/or oil droplets and requires reliable removing equipment to perform filtration. Recent natural gas processing applications are demanded compactness and reliability of process equipment. Since conventional means are sophisticated in design, poor in efficiency, and continue lacking robust, a supersonic nozzle has been introduced as an alternative means to meet such demands. A 3-D Convergent-Divergent Nozzle is simulated using commercial Code for pressure ratio (NPR) varies from 1.2 to 2. Six different shapes of nozzle are numerically examined to illustrate the position of shock-wave as such spot could be considered as a benchmark of particle separation. Rectangle, triangle, circular, elliptical, pentagon, and hexagon nozzles are simulated using Fluent Code with all have same cross-sectional area. The simple one-dimensional inviscid theory does not describe the actual features of fluid flow precisely as it ignores the impact of nozzle configuration on the flow properties. CFD Simulation results, however, show that nozzle geometry influences the flow structures including location of shock wave. The CFD analysis predicts shock appearance when p01/pa>1.2 for almost all geometry and locates at the lower area ratio (Ae/At). Simulation results showed that shock wave in Elliptical nozzle has the farthest distance from the throat among the others at relatively small NPR. As NPR increases, hexagon would be the farthest. The numerical result is compared with available experimental data and has shown good agreement in terms of shock location and flow structure.

Keywords: CFD, Particle Separation, Shock wave, Supersonic Nozzle.

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Authors: Yi Liu

Abstract:

In this paper, computational fluid dynamics (CFD) is utilized to characterize a prototype biolistic delivery system, the biomedical device based on the contoured-shock-tube design (CST), with the aim at investigating shocks induced flow instabilities within the contoured shock tube. The shock/interface interactions, the growth of perturbation at an interface between two fluids of different density are interrogated. The key features of the gas dynamics and gas-particle interaction are discussed

Keywords: Simulation, Shock wave, Particle, Interface, Supersonic, Richtmyer-Meshkov Instability

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Authors: A.Hasanpour, M. Farhadi, H.R. Ashorynejad

Abstract:

In this paper a numerical technique is used to predict the metal temperature of a gas turbine vane. The Rising combustor exit temperatures in gas turbine engines necessitate active cooling for the downstream turbine section to avoid thermal failure. This study is performed the solution of external flow, internal convection, and conduction within the metal vane. Also the trade-off between the cooling performances in four different hole shapes and configurations is performed. At first one of the commonly used cooling hole geometry is investigated; cylindrical holes and then two other configurations are simulated. The average temperature magnitude in mid-plan section of each configuration is obtained and finally the lower temperature value is selected such as best arrangement.

Keywords: Forced Convection, Gas Turbine Blade, Hole Configuration

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Authors: Chien-Song Chyang, Yu-Chi Wang

Abstract:

For sludge disposal, incineration is considered to be better than direct burial because of regulations and space limitations in Taiwan. Additionally, burial after incineration can effectively prolong the lifespan of a landfill. Therefore, it is the most satisfactory method for treating sludge at present. Of the various incineration technologies, the fluidized bed incinerator is a suitable choice due to its fuel flexibility. In this work, sludge generated from industrial plants was treated in a pilot-scale vortexing fluidized bed. The moisture content of the sludge was 48.53%, and its LHV was 454.6 kcal/kg. Primary gas and secondary gas were fixed at 3 Nm³/min and 1 Nm³/min, respectively. Diesel burners with on-off controllers were used to control the temperature; the bed temperature was set to 750±20 °C, and the freeboard temperature was 850±20 °C. The experimental data show that the NO emission increased with bed temperature. The maximum NO emission is 139 ppm, which is in agreement with the regulation. The CO emission is low than 100 ppm through the operation period. The mean particle size of fly ash collected from baghouse decreased with operating time. The ration of bottom ash to fly ash is about 3. Compared with bottom ash, the potassium in the fly ash is much higher. It implied that the potassium content is not the key factor for aggregation of bottom ash.

Keywords: Sludge incineration, fluidized bed combustion, fly ash, bottom ash.

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Authors: J. Dumrongsak, A. M. Savill

Abstract:

Non-premixed turbulent combustion Computational Fluid Dynamics (CFD) has been carried out in a simplified methanefuelled coaxial jet combustor employing Large Eddy Simulation (LES). The objective of this study is to evaluate the performance of LES in modelling non-premixed combustion using a commercial software, FLUENT, and investigate the effects of the grid density and chemistry models employed on the accuracy of the simulation results. A comparison has also been made between LES and Reynolds Averaged Navier-Stokes (RANS) predictions. For LES grid sensitivity test, 2.3 and 6.2 million cell grids are employed with the equilibrium model. The chemistry model sensitivity analysis is achieved by comparing the simulation results from the equilibrium chemistry and steady flamelet models. The predictions of the mixture fraction, axial velocity, species mass fraction and temperature by LES are in good agreement with the experimental data. The LES results are similar for the two chemistry models but influenced considerably by the grid resolution in the inner flame and near-wall regions.

Keywords: Coaxial jet, reacting LES, non-premixed combustion, turbulent flow.

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Authors: Caihong Su

Abstract:

Transition prediction of boundary layers has always been an important problem in fluid mechanics both theoretically and practically, yet notwithstanding the great effort made by many investigators, there is no satisfactory answer to this problem. The most popular method available is so-called e-N method which is heavily dependent on experiments and experience. The author has proposed improvements to the e-N method, so to reduce its dependence on experiments and experience to a certain extent. One of the key assumptions is that transition would occur whenever the velocity amplitude of disturbance reaches 1-2% of the free stream velocity. However, the reliability of this assumption needs to be verified. In this paper, transition prediction on a flat plate is investigated by using both the improved e-N method and the parabolized stability equations (PSE) methods. The results show that the transition locations predicted by both methods agree reasonably well with each other, under the above assumption. For the supersonic case, the critical velocity amplitude in the improved e-N method should be taken as 0.013, whereas in the subsonic case, it should be 0.018, both are within the range 1-2%.

Keywords: Boundary layer, e-N method, PSE, Transition

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