Abstract: High-velocity oxygen fuel (HVOF) thermal spraying
uses a combustion process to heat the gas flow and coating material.
A computational fluid dynamics (CFD) model has been developed to
predict gas dynamic behavior in a HVOF thermal spray gun in which
premixed oxygen and propane are burnt in a combustion chamber
linked to a parallel-sided nozzle. The CFD analysis is applied to
investigate axisymmetric, steady-state, turbulent, compressible,
chemically reacting, subsonic and supersonic flow inside and outside
the gun. The gas velocity, temperature, pressure and Mach number
distributions are presented for various locations inside and outside
the gun. The calculated results show that the most sensitive
parameters affecting the process are fuel-to-oxygen gas ratio and
total gas flow rate. Gas dynamic behavior along the centerline of the
gun depends on both total gas flow rate and fuel-to-oxygen gas ratio.
The numerical simulations show that the axial gas velocity and Mach
number distribution depend on both flow rate and ratio; the highest
velocity is achieved at the higher flow rate and most fuel-rich ratio.
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 HVOF system design, optimization and performance
analysis.
Abstract: In this paper the supersonic ejectors are
experimentally and analytically studied. Ejector is a device that
uses the energy of a fluid to move another fluid. This device works
like a vacuum pump without usage of piston, rotor or any other
moving component. An ejector contains an active nozzle, a passive
nozzle, a mixing chamber and a diffuser. Since the fluid viscosity
is large, and the flow is turbulent and three dimensional in the
mixing chamber, the numerical methods consume long time and
high cost to analyze the flow in ejectors. Therefore this paper
presents a simple analytical method that is based on the precise
governing equations in fluid mechanics. According to achieved
analytical relations, a computer code has been prepared to analyze
the flow in different components of the ejector. An experiment has
been performed in supersonic regime 1.5
Abstract: This paper describes Nano-particle based Planar Laser
Scattering (NPLS) flow visualization of angled supersonic jets into a
supersonic cross flow based on the HYpersonic Low TEmperature
(HYLTE) nozzle which was widely used in DF chemical laser. In
order to investigate the non-reacting flowfield in the HYLTE nozzle, a
testing section with windows was designed and manufactured. The
impact of secondary fluids orifice separation on mixing was examined.
For narrow separation of orifices, the secondary fuel penetration
increased obviously compared to diluent injection, which means
smaller separation of diluent and fuel orifices would enhance the
mixing of fuel and oxidant. Secondary injections with angles of 30, 40
and 50 degrees were studied. It was found that the injectant
penetration increased as the injection angle increased, while the
interfacial surface area to entrain the freestream fluid is largest when
the injection angle is 40 degree.
Abstract: Acoustic function plays an important role in
aerodynamic mechanical engineering. It can classify the kind of
air-vehicle such as subsonic or supersonic. Acoustic velocity
relates with velocity and Mach number. Mach number relates
again acoustic stability or instability condition. Mach number
plays an important role in growth or decay in energy system.
Acoustic is a function of temperature and temperature is directly
proportional to pressure. If we control the pressure, we can control
acoustic function. To get pressure stability condition, we apply
Navier-Stokes equations.
Abstract: In this work, we try to find the best setting
of Computational Fluid Dynamic solver available for the problems in
the field of supersonic internal flows. We used the supersonic air-toair
ejector to represent the typical problem in focus. There are
multiple oblique shock waves, shear layers, boundary layers
and normal shock interacting in the supersonic ejector making this
device typical in field of supersonic inner flows. Modeling of shocks
in general is demanding on the physical model of fluid, because
ordinary conservation equation does not conform to real conditions in
the near-shock region as found in many works. From these reasons,
we decided to take special care about solver setting in this article by
means of experimental approach of color Schlieren pictures and
pneumatic measurement. Fast pressure transducers were used to
measure unsteady static pressure in regimes with normal shock in
mixing chamber. Physical behavior of ejector in several regimes is
discussed. Best choice of eddy-viscosity setting is discussed on the
theoretical base. The final verification of the k-ω SST is done on the
base of comparison between experiment and numerical results.
Abstract: Elementary particles are created in pairs of equal and opposite momentums at a reference frame at the speed of light. The speed of light reference frame is viewed as a point in space as observed by observer at rest. This point in space is the bang location of the big bang theory. The bang in the big bang theory is not more than sustained flow of pairs of positive and negative elementary particles. Electrons and negative charged elementary particles are ejected from this point in space at velocities faster than light, while protons and positively charged particles obtain velocities lower than light. Subsonic masses are found to have real and positive charge, while supersonic masses are found to be negative and imaginary indicating that the two masses are of different entities. The electron-s super-sonic speed, as viewed by rest observer was calculated and found to be less than the speed of light and is little higher than the electron speed in Bohr-s orbit. The newly formed hydrogen gas temperature was found to be in agreement with temperatures found on newly formed stars. Universe expansion was found to be in agreement. Partial mass and charge elementary particles and particles with momentum only were explained in the context of this theoretical approach.
Abstract: Numerical study of two dimensional supersonic
hydrogen-air mixing layer is performed to investigate the effect of
turbulence and chemical additive on ignition distance. Chemical
reaction is treated using detail kinetics. Advection upstream splitting
method is used to calculate the fluxes and one equation turbulence
model is chosen here to simulate the considered problem. Hydrogen
peroxide is used as an additive and the results show that inflow
turbulence and chemical additive may drastically decrease the
ignition delay in supersonic combustion.
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.
Abstract: A parametric study of a mixed-compression
supersonic inlet is performed and reported. The effects of inlet Mach
Numbers, varying from 4 to 10, and angle of attack, varying from 0
to 10, are reported for a constant inlet dynamic pressure. The paper
looked at the variations of mass flow rates through the inlet, gain in
entropy through the inlet, and the angles of the external oblique
shocks. The mass flow rates were found to decrease monotonically
with Mach numbers and increase with angle of attacks. On the other
hand the entropy gain through the inlet increased with increasing
Mach number and angle of attack. The variation in static pressure
was found to be identical from the inlet throat to the exit for Mach
number values higher than 6.
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.
Abstract: The flow field within the combustor of scramjet
engine is very complex and poses a considerable challenge in the
design and development of a supersonic combustor with an optimized
geometry. In this paper comprehensive numerical studies on flow
field characteristics of different cavity based scramjet combustors
with transverse injection of hydrogen have been carried out for both
non-reacting and reacting flows. The numerical studies have been
carried out using a validated 2D unsteady, density based 1st-order
implicit k-omega turbulence model with multi-component finite rate
reacting species. The results show a wide variety of flow features
resulting from the interactions between the injector flows, shock
waves, boundary layers, and cavity flows. We conjectured that an
optimized cavity is a good choice to stabilize the flame in the
hypersonic flow, and it generates a recirculation zone in the scramjet
combustor. We comprehended that the cavity based scramjet
combustors having a bearing on the source of disturbance for the
transverse jet oscillation, fuel/air mixing enhancement, and flameholding
improvement. We concluded that cavity shape with
backward facing step and 45o forward ramp is a good choice to get
higher temperatures at the exit compared to other four models of
scramjet combustors considered in this study.
Abstract: Along with increasing development of generation of supersonic planes especially fighters and request for increasing the performance and maneuverability scientists and engineers suggested the delta and double delta wing design. One of the areas which was necessary to be researched, was the Aerodynamic review of this type of wings in high angles of attack at low speeds that was very important in landing and takeoff the planes and maneuvers. Leading Edges of the wings,cause the separation flow from wing surface and then formation of powerful vortex with high rotational speed which studing the mechanism and location of formation and also the position of the vortex breakdown in high angles of attack is very important. In this research, a double delta wing with 76o/45o sweep angles at high angle of attack in steady state and incompressible flow were numerically analyzed with Fluent software. With analaysis of the numerical results, we arrived the most important characteristic of the double delta wings which is keeping of lift at high angles of attacks.
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.
Abstract: Supersonic open and closed cavity flows are investigated experimentally and computationally. Free stream Mach number of two is set. Schlieren imaging is used to visualise the flow behaviour showing stark differences between open and closed. Computational Fluid Dynamics (CFD) is used to simulate open cavity of flow with aspect ratio of 4. A rear wall treatment is implemented in order to pursue a simple passive control approach. Good qualitative agreement is achieved between the experimental flow visualisation and the CFD in terms of the expansion-shock waves system. The cavity oscillations are shown to be dominated by the first and third Rossister modes combining to high fluctuations of non-linear nature above the cavity rear edge. A simple rear wall treatment in terms of a hole shows mixed effect on the flow oscillations, RMS contours, and time history density fluctuations are given and analysed.
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
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.
Abstract: In this study, an optimization of supersonic air-to-air ejector is carried out by a recently developed single-objective genetic algorithm based on adaption of sequence of individuals. Adaptation of sequence is based on Shape-based distance of individuals and embedded micro-genetic algorithm. The optimal sequence found defines the succession of CFD-aimed objective calculation within each generation of regular micro-genetic algorithm. A spring-based deformation mutates the computational grid starting the initial individualvia adapted population in the optimized sequence. Selection of a generation initial individual is knowledge-based. A direct comparison of the newly defined and standard micro-genetic algorithm is carried out for supersonic air-to-air ejector. The only objective is to minimize the loose of total stagnation pressure in the ejector. The result is that sequence-adopted micro-genetic algorithm can provide comparative results to standard algorithm but in significantly lower number of overall CFD iteration steps.
Abstract: This article presents the boundary conditions for the problem of turbulent supersonic gas flow in a plane channel with a perpendicular injection jets. The non-reflection boundary conditions for direct modeling of compressible viscous gases are studied. A formulation using the NSCBC (Navier- Stocks characteristic boundary conditions) through boundaries is derived for the subsonic inflow and subsonic non-reflection outflow situations. Verification of the constructed algorithm of boundary conditions is carried out by solving a test problem of perpendicular sound of jets injection into a supersonic gas flow in a plane channel.
Abstract: Plasma Wind Tunnels (PWT) are extensively used for screening and qualification of re-entry Thermel Protection System (TPS) materials. Proper design of a supersonic diffuser for plasma wind tunnel is of importance for achieving good pressurerecovery (thereby reducing vacuum pumping requirement & run time costs) and isolating downstream stream fluctuations from propagating costs) and isolating downstream stream fluctuationnts the details of a rapid design methodology successfully employed for designing supersonic diffuser for high power (several megawatts)plasma wind tunnels and numerical performance analysis of a diffuser configuration designed for one megawatt power rated plasma wind tunnel(enthalpy ~ 30 MJ/kg) using FLUENT 6.3® solver for different diffuser operating sub-atmospheric back-pressures.
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.