microsoft word - project report
TRANSCRIPT
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CHAPTER1
INTRODUCTION
A Nozzle is a device designed to control the direction or
characteristics of a fluid flow (especially to increase velocity) as it
exits (or enters) an enclosed chamber or pipe via an orifice. A nozzle is
often a pipe or tube of varying cross sectional area and it can be used to
direct or modify the flow of a fluid (liquid or gas). Nozzles are
frequently used to control the rate of flow, speed, direction, mass,
shape, and/or the pressure of the stream that emerges from the
Frequently the goal is to increase the kinetic energy of the flowing
medium at the expense of its pressure and internal energy.
Nozzles can be described as convergent (narrowing down from a
wide diameter to a smaller diameter in the direction of the flow) or
divergent (expanding from a smaller diameter to a larger one). A de
Laval nozzle has a convergent section followed by a divergent section
and is often called a nozzle. Convergent nozzles accelerate subsonic
fluids. If the nozzle pressure ratio is high enough the flow will reach
sonic velocity at the narrowest point (i.e. the nozzle throat). In this
situation, the nozzle is said to be choked.
Increasing the nozzle pressure ratio further will not increase the
throat Mach number beyond unity. Downstream (i.e. external to the
nozzle) the flow is free to expand to supersonic velocities. Note that
the Mach 1 can be a very high speed for a hot gas; since the speed of
sound varies as the square root of absolute temperature. Thus the speed
reached at a nozzle throat can be far higher than the speed of sound at
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sea level. This fact is used extensively in rocketry where hypersonic
flows are required, and where propellant mixtures are deliberately
chosen to further increase the sonic speed.
Divergent nozzles slow fluids, if the flow is subsonic, but
accelerate sonic or supersonic fluids. Convergent can therefore
accelerate fluids that have choked in the convergent section to
supersonic speeds. This CD process is more efficient than allowing a
convergent nozzle to expand supersonically externally. The shape of
the divergent section also ensures that the direction of the escaping
gases is directly backwards, as any sideways component would not
contribute to thrust. Jet exhaust produces a net thrust from the energy
obtained from combusting fuel which is added to the inducted air. This
hot air is passed through a high speed nozzle, a propelling nozzle
which enormously increases its kinetic energy.
For a given mass flow, greater thrust is obtained with a higher
exhaust velocity, but the best energy efficiency is obtained when the
exhaust speed is well matched with the airspeed. However, no jet
aircraft can maintain velocity while exceeding its exhaust jet speed,
due to momentum considerations. Supersonic jet engines, like those
employed in fighters and SST aircraft (e.g. Concorde), need high
exhaust speeds. Therefore supersonic aircraft very typically use a CD
nozzle despite weight and cost penalties. Subsonic jet engines employ
relatively low, subsonic, exhaust velocities. They thus employ simple
convergent nozzles. In addition, bypass nozzles are employed giving
even lower speeds.
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Rocket motors use convergent-divergent nozzles with very large
area ratios so as to maximise thrust and exhaust velocity and thus
extremely high nozzle pressure ratios are employed. Mass flow is at a
premium since all the propulsive mass is carried with vehicle, and very
high exhaust speeds are desirable.
Most modern passenger and military aircraft are powered by gas
turbine engines, which are also called jet engines. There are several
different types of gas turbine engines, but all turbine engines have
some parts in common. All gas turbine engines have a nozzle to
produce thrust, to conduct the exhaust gases back to the free stream,
and to set the mass flow rate through the engine. The nozzle sits
downstream of the power turbine.
A nozzle is a relatively simple device, just a specially shaped
tube through which hot gases flow. However, the mathematics which
describes the operation of the nozzle takes some careful thought. As
shown above, nozzles come in a variety of shapes and sizes depending
on the mission of the aircraft. Simple turbojets, and turboprops, often
have a fixed geometry convergent nozzle. Turbofan engines often
employ a co-annular nozzle as shown at the top left. The core flow
exits the centre nozzle while the fan flow exits the annular nozzle.
Mixing of the two flows provides some thrust enhancement and these
nozzles also tend to be quieter than convergent nozzles. Afterburning
turbojets and turbofans require a variable geometry convergent-
divergent - CD nozzle. In this nozzle, the flow first converges down to
the minimum area or throat, and then is expanded through the
divergent section to the exit at the right. The variable geometry causes
these nozzles to be heavier than a fixed geometry nozzle, but variable
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geometry provides efficient engine operation over a wider airflow
range than a simple fixed nozzle.
Rocket engines also use nozzles to accelerate hot exhaust toproduce thrust. Rocket engines usually have a fixed geometry CD
nozzle with a much larger divergent section than is required for a gas
turbine. All of the nozzles we have discussed thus far are round tubes.
Recently, however, engineers have been experimenting with nozzles
with rectangular exits. This allows the exhaust flow to be easily
deflected, or vectored. Changing the direction of the thrust with the
nozzle makes the aircraft much more manoeuvrable.
Because the nozzle conducts the hot exhaust back to the free
stream, there can be serious interactions between the engine exhaust
flow and the airflow around the aircraft. On fighter aircraft, in
particular, large drag penalties can occur near the nozzle exits. As with
the inlet design, the external nozzle configuration is often designed by
the airframes and subjected to wind tunnel testing to determine the
performance effects on the airframe. The internal nozzle is usually the
responsibility of the engine manufacturer.
1.1 INTRODUCTION TO JET
Jet is a free shear flow driven by momentum introduced at the
nozzle exit of, usually, a nozzle or am orifice which exhibits a
characteristic that, the ratio of width to axial distance is a constant.
The jet may also define as a continuous fluid flow issuing from an
orifice into a medium of lower speed fluid. As the jet fluid travels
further away from its origin, it slows down due to mixing with slower
speed ambient fluid. This is due to boundary layer at the nozzle exit
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which develops roll up structure, or ring vortices, which grow in size
when they move downstream, due to the entrainment of ambient fluid
into jet stream. Thus, mass flow at any cross-section of the jet
progressively increases along the downstream direction. Hence, to
converse momentum the centreline velocity decreases with
downstream distance. The resulting centreline velocity decay as
proportional to gradient across the shear layer and is a strong function
of distance downstream of the exit. The vast quanta of knowledge
presently available and continuous research currently being carried out
stand testimony to the importance associated with the jet flows.
1.1.1 CLASSIFICATION OF JETS
Basically jets can be classified into two categories namely;
incompressible and compressible jets Fig.1.1. The jets with Mach
number less than 0.3, up to which the compressibility effects are
negligible are called incompressible jets. Compressible jets can be
again subdivided into subsonic, sonic and supersonic jets. Jets with
Mach number 1.0 are called sonic jets, which can be correctly
expanded or under expanded. Supersonic jets are the jets with Mach
number more than one. These can be further classified into over
expanded, correctly expanded and under expanded jets.
1.1.2REGIMES OF JETS
A Schematic diagram of a typical subsonic jet and the different
flow zones are shown in Fig. 1.2 the flow regimes in the subsonic jets
are classified as follows:
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Fig 1.1 Classification of jets.
a. Potential core region: This the region consists of a coreconstant axial velocity close to the jet exit velocity surrounded
by a rapidly growing and strongly sheared annulus of mixing
layer or shear layer with intense turbulence. Potential core region
extends about 5 times the nozzle exit diameter (D) downstream
from the nozzle exit. This is because, the mixing initiated at the
jet boundaries has not yet permeated into the entire flow field,
thus leaving a region that is characterized by a constant axial
velocity.
b. Transition region: This is the region where the centerlinevelocity begins to decay. This characteristic decay zone extends
from about 5D to 10D downstream over which the turbulence
JETS
INCOMPRESSIBLE
0
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changes from its annular to a somewhat pseudo-cylindrical
distribution. As a result, the velocity difference between the
ambient fluid and the high speed core of the jet decrease and
attenuates the shear that supports the vortical rings in the jet and
thus the velocity profiles become smoother with jet propagation.
c. Fully developed region: Beyond the transition region the jetbecomes similar in appearance to a flow of fluid from a source of
infinitely small thickness In reality the jet velocity becomes
insignificant after about 30D
Fig.1.2 Schematic of different zones in a subsonic jet.
1.2 NEED OF JET CONTROL
There are numerous system, especially in the aerospace, where
the ability to enhance the mixing characteristics of a jet will greatly
improve their performance. For example, by increasing the rate of
mixing between air and fuel, the efficiency of a combustion cycle can
be improved. Other examples of technological application requiring
control of mixing in compressible flows include thrust augmenting
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ejectors, thrust vector control, metal deposition, and gas dynamic
lasers.
1.3 TYPES OF CONTROLS
Control may be defined as the ability to modify the flow
characteristics of jets. Jet controls can be broadly classified into active
and passive controls. Both active and passive controls mainly aim at
modifying the flow and noise characteristics.
1.3.1 ACTIVE CONTROL
In active control, an auxiliary power source (like micro jets) is
used to control the jet characteristics. Many active jet control methods
use energized actuators to dynamically manipulate flow phenomena.
Pulsed jets, piezoelectric actuators, micro jets and oscillating jets are
among the most effective controls for active mixing enhancement.
1.3.2 PASSIVE CONTROL
In passive control the controlling energy is drawn directly from
the flow to be controlled. Passive controls are mostly desired because
no external power source is required. Passive control methods use
geometrical modifications which alter the flow structure. Some of the
commonly used passive control methods are shown below.
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Fig.1.3 Schematic of different type of passive controls.
1.4. CHEVRONS
Chevrons are saw tooth-like patterns at the trailing edge of jet engine
nozzles that help reduce noise from the ensuing jet. It has been known
from past experimental studies with laboratory-scale jets that small
protrusions at the nozzle lip, called tabs, would suppress screech
tones. In the 1980s and 1990s the tabs were explored extensively for
mixing enhancement in jets. These studies advanced the understanding
of the flow mechanisms and suggested that the technique might have a
potential for reduction of turbulent mixing noise that is the
dominant component of jet noise for most aircraft. Driven by stringent
noise regulations, such a potential first received serious attention
on an application level in the mid 1990s. Engine companies expressed
interest and some proposed their own concepts for tests. In 1996-97,
concepts from General Electric Aircraft Engines (GEAE), Pratt &
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Whitney (P&W) and others were combined into a test program under
NASAs Advanced Subsonic Technology (AST) Program. Various
tab/chevron configurations were evaluated for noise reduction with
models of separate flow nozzles in free-jet tests and encouraging results
were obtained. However, scepticism lingered and there was reluctance to
embrace the technology primarily out of concerns about thrust penalty.
In 1998 the impact on thrust was evaluated and found to be less than
0.25%. This was the turning point in the development of the technology
when industry started to invest heavily with product development
programs. The effort under AST culminated in flight tests in 2001 on
NASAs Lear jet 25 and Honeywells Falcon 20 test aircraft proving the
noise reduction.
Today, chevrons are implemented on various engines,
However, as stated, the evolution of the technology can be traced back
to decades of fundamental studies with tabs and similar devices at
universities, NASA as well as in industry. The concerted NASA /
industry studies in the 1990s eventually led to designs that produced
significant noise reduction while keeping the thrust loss within acceptable
limits. The objective of this paper is to provide an account of this
evolution, starting with a summary of the earlier fundamental studies.
1.4.1. EARLIER STUDIES ON THE EFFECT OF TABS
It has been known for a long time that tabs, small protrusions placed near
the nozzle exit, suppress screech noise. Screech is a phenomenon typical of
small, clean, laboratory jets that, under imperfectly expanded supersonic
condition, involve a feedback loop to produce a sharp tone. In laboratory
experiments the curious suppression effect is readily demonstrated by
inserting a small obstacle, such as the tip of a pencil, near the nozzle exit.
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One of the earliest studies of noise suppression by such devices is that of
Westley & Lilley . A picture of the teeth patterns used in their
experiment, in the then newly established program of jet noise research at
Cranfield, UK.
The authors observed large reduction of supersonic jet noise by these
devices apparently in part due to suppression of screech. Later experiments
usually deployed a single tab or two tabs that were sufficient to suppress
screech. Suppression of screech was desired in order to allow a clearer
study of other components of jet noise.
With regards to the effect of tabs on the jet flow field, The authors of this
work noted that the insertion of small rectangular tabs into the jet flow on
the nozzle perimeter had a profound effect; the apparent potential core
length was reduced to about two diameters followed by a rapid decay of
the centerline mean velocity.
With previous studies when compared with the rectangular protrusion at
the nozzle exit, it was soon recognized that a triangular tab with same
base width worked just as well. Moreover, when the apex of the triangular
tab was tilted downstream it appeared to work even better.
Thats why here we have chosen the triangular protrusion in addition to
that we have added 8 tabs for better reduction of noise.
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CHAPTER 2
LITERATURE SURVEY
2.1EFFECT OF CHEVRON COUNT AND PENETRATION
Experimental investigations (1) have been carried out on chevron
nozzles to assess the importance of chevron parameters such as the
number of chevrons (chevron count) and chevron penetration. Acoustic
measurements such as overall sound pressure level, spectra, directivity,
acoustic power, and broadband shock .Noise has been made over a
range of nozzle pressure ratio from sub-critical to under expansion
levels. Shadowgraph images of the shock-cell structure of jets from
various chevron nozzles have also been captured for different nozzle
pressure ratios. The results indicate that a higher chevron count with a
lower Level of penetration yields the maximum noise suppression for
low and medium nozzle pressure ratios. Of all the geometries studied,
chevron nozzle with eight lobes and 0degree penetration angle gives
the maximum noise reduction. Chevron nozzles are found to be free
from screech unlike regular nozzles. Acoustic Power index has been
calculated to quantitatively evaluate the performance of the various
chevron nozzles. Chevron count is the pertinent parameter for noise
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reduction at low nozzle pressure ratios, whereas at high nozzle pressure
ratios, chevron penetration is crucial. The results illustrate that by
careful selection of chevron parameters substantial noise reduction can
be achieved.
2.2 NUMERICAL PREDICTIONS OF NOISE IN
NOZZLESWITH AND WITHOUT CHEVRONS
Numerical simulations (2) of round, compressible, turbulent jets
using the Shear Stress Transport (SST kx)model have been carried
out. The three-dimensional calculations have been done on a
tetrahedral mesh with 0.9 million cells. Two jets, one cold and hot,
have been simulated. The Mach number for both the Cases is 0.75.
Overall sound pressure levels (SPL) at far-field observer locations have
been calculated using fowcs WilliamsHawkins equation. The
numerical predictions have been compared with experimental results
available in the literature. Axial and radial variation of the mean axial
velocity and overall SPL levels are compared. The potential core
length is predicted well, but the predicted centreline velocity decay is
faster than the measured value. The URANS calculations are not able
to predict the absolute values for the overall SPL, but predict the trends
reasonably well. The calculations predict the trends and absolute values
of the variations of the spectral amplitude well for the aft receivers, but
not for the forward receivers. Effect of chevrons on the noise from the
jet is also investigated for cold and hot jets. In each case, two chevron
taper angles, namely, 0degree and 5degree are considered. The latter
nozzle produces the most significant modification to the baseline
spectra and is less effective at high frequencies in abating the noise.
The present calculations predict a reduction in the overall SPL for the
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chevron nozzle with 0degree taper angle and a slight increase for
chevron nozzle with 5degree taper angle, for both cold and hot jets.
2.3 LARGE-EDDY SIMULATION OF CHEVRON JET FLOWWITH NOISE PREDICTIONS
Hybrid large-eddy(3) type simulations for chevron nozzle jet
flows are performed at Mach 0.9 and Re = 1.03*10^6. Without using
any sub grid scale model (SGS), the numerical approach applied in the
Present study is essentially implicit large-eddy simulation (ILES).
However, a Reynolds-averaged NavierStokes (RANS) solution is
patched into the near wall region. This makes the overall solution
strategy Hybrid RANSILES. The disparate turbulence length scales,
implied by these different modelling Approaches are matched using a
HamiltonJacobi equation. The complex geometry features of the
chevron Nozzles are fully meshed. With numerical fidelity in mind,
high quality, hexahedral multi-block Meshes of 12.5*10^6 cells are
used. Despite the modest meshes, the novel RANSILES approach
shows Encouraging performance. Computed mean and second-order
fluctuating quantities of the turbulent near field compare favourably
with measurements. The radiated far-field sound is predicted using the
Ffowcs Williams and Hawkins (FWH) surface integral method.
Encouraging agreement of the predicted far field sound directivity and
spectra with measurements is obtained.
Key purpose of this study is to highlight the penetration effects
of chevrons. Hence, the most severe 18.2bend SMC006 nozzle is
chosen as our primary focus. For comparative purposes the 5SMC001
is also fully studied. To explore the altered chevron shear layer mixing
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by the chevrons, the near nozzle instantaneous flow field is examined.
Only the results for the much more strongly penetrating SMC006 are
presented here gives computational snapshots of density Gradient
magnitude or Numerical Schlieren on both cut planes In spite of
some density variations within it, the potential core is clearly visible,
its length being slightly shorter than 5D. Strong density Variations are
located in the shear layer next to the potential Core. The outer edge of
the shear layer can also be easily identified.
The numerical simulation confirms the experiment with respect
to the influence of chevron penetration. With a bend angle of 18.2,the
potential core is shorter, downstream Reynolds stresses lower, but the
sound source distribution wider in the side line direction. The
characteristics of far-field sound for the 18.2 bend are therefore to the
bend angle of 5. Namely, the more server bend has louder Side line but
quieter downstream noise. It also increases the high-frequency sound
intensity but decreases the low-frequency. This consistency of the
current numerical study with measurements encouraging and
suggesting a reliable numerical frame work has been developed.
2.4 AN EXPERIMENTAL INVESTIGATION OF FLOW WITH
A SINGLE DELTA TAB
A single inverted(4) delta tab attached to the trailing edge of a
splitter plate in a two-stream mixing layer has been examined
experimentally using a three-component laser-Doppler anemometer.
Detailed mean flow and turbulence measurements were obtained at a
velocity ratio of 2:1 between the two co-flowing streams. The results
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showed that, when the tab was tilted to the high-speed side, stream
wise vortices generated and the subsequent mixing were stronger and
more intense than when tilting it to the low-speed side. The strength of
stream wise vorticity appeared to have a direct correlation with the
level of turbulence generated in the cross-stream directions. Attempts
were also made to quantify the effect of each (streamwise vorticity)
production term in the streamwise vorticity transport equation.
The effects of a single inverted delta tab in a two-stream mixing
flow situation have been investigated using a three component laser-
Doppler anemometer. The distortion in the streamwise mean velocity
flow fields and the generation of the streamwise vorticity that resulted
from the introduction of tab.
The tab has been established via measurements of three mean velocity
components and six Reynolds stress components in the present
investigation. The results have clearly shown that when the tab is tilted
to the high-speed side, it enhances mixing between the two streams
better than when tilted on the side of the low-speed flow. When the tab
is placed on the high-speed side, both sources for stream wise vorticity
generation are operative. When the tab is placed on the low-speed side,
only source 2 is operative. This observation suggests that the
production of turbulence in the lateral directions is strongly associated
with stream wise vorticity and that the imbalance of production and
dissipation leads to significant diffusion in the stations further
downstream.
2.5 CHARACTERISTICS OF SONIC JETS WITH TABS
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The result of an experimental investigation(5) on the effect of
vortex generation in the form of a mechanical tab placed at the nozzle
exit on the evolution of jet and its decay are reported in this paper. Jets
from a sonic nozzle with and without tabs operated at nozzle pressure
ratio from 2 to 7 were studied. Tabs with two combinations of length -
to-width ratio were investigated by keeping the blockage area constant.
The tabs offered a blockage of 10.18% of the nozzle exit area. The
centreline pitot pressure decay shows that for the tabbed jet a
maximum core reduction of about 75% can be achieved at a nozzle
pressure ratio (NPR) 7 compared to an uncontrolled jet. So that the
tabs drastically weaken the shock structure in the jet core and disperse
the supersonic zone of the flow making them occupying a greater zone
of the flow field compare to the plane nozzle. This causes the waves to
become weaker and the jet to spread faster. The tabs are found to shed
counter-rotating vortices all along the edges, resulting in enhanced
mixing. Isobaric contours reveal that the streamwise vortices cause an
inward indentation of the entrained mass into the jet core and an
outward ejection of core flow. To understand the distortion introduced
by tabs on the jet cross section and its growth leading to bifurcation of
the jet, a surface coating visualization method was developed and
employed.
Vortex generators in the form of tabs have been found to be quite
effective in influencing jet evolution and mixing. The tabs are found to
be effective in bringing down the centreline pitot pressure oscillations
for all nozzle pressure ratios. Shadow graph pictures reveal that the
tabs diffuse the shocks, resulting in weaker shocks in the core region.
Further, it is found that the tabs disperse the supersonic zone of the
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flow field compared to the plane nozzle. The tab is found to generate a
pair of counter-rotating vortices, which result in enhanced mixing.
Isobaric contours reveal that generated stream wise vortices cause an
inward indentation of entrained mass into jet core and an outward
ejection of core flow. Surface visualization explains the distortion
introduced by the tabs on the jet cross section. The distortion produced
by the two tabs grows which downstream distance and result
essentially in bifurcation of the jet. The tab length is found to be more
effective in mixing enhancement then the width for the same blockage
area and the limit for tab length are the nozzle radius and not the
boundary layer thickness.
2.6 EFFECT OF TAB ON MIXING CHARACTERISTICS OF
SUBSONIC AND SONIC JETS
The effect of tabs placed (6) at the exit of a circular nozzle of
10mm diameter on the near flow field of the jet was investigated
experimentally for subsonic mach numbers. The tab used was a hollow
semi circular tube of diameter 1.5mm and length 2mm. The near jet
flow field was studied for three configurations of the tab, namely, the
concave surface facing the flow exiting nozzle (arc tab facing -in) and
convex surface facing the flow (arc tab facing -out) for the blockage
ratio of 7.64 %.the center line mach number decay shows that for the
jet with arc tab facing in, a maximum reduction in core length of
about 80% was achieved at all subsonic and sonic correctly expanded
conditions of jet. Arc tab facing-out and rectangular tab configuration
reduces the core length to about 50%. The decay of arc tab controlled
jet was compared with that obtained for a plain rectangular tab of same
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blockage and a plain circular nozzle. The jet was found to decay at a
faster rate in the case of arc- tab facing in configuration as compared to
the facing out and rectangular tab configurations. Mach number
profiles show that the arc tab facing in distorts the jet efficiently by
spreading the jet wider in the plane normal to the tab and the effect of
spread is more pronounced in the jet with arc tab facing in as compared
to the arc tab facing out. The effect of tab orientation and shape seem
to have a profound influence on the development of the jet in the near
field.
Among the three configurations arc tab facing in was found to be
more efficient in reducing the potential core length and distorting three
jet structure compared to the other two configurations in all subsonic
and sonic correctly expanded mach numbers .arc tab facing in is
showing consistently better performance for all the mach numbers
compared to other two configurations. It may be due to the fact that the
arc tab facing in creates stronger pressure gradient shedding stronger
stream wise vortices causing reduction in potential core and enhanced
mixing.
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CHAPTER 3
COMPUTATIONAL WORK
3.1 MODELLING
Designing three dimensional nozzle with different configurations
at exit we used CATIA modelling software which is flexible one anduser friendly.
3.1.1 THREE DIMENSIONALMODELLING
For this project, the model of nozzle has been created by using
modelling software CATIA. Initially the sketch has been created
with the given dimension, and at the exit the domain for the
visualization of the potential core region has been designed with the
respective dimensions.
NOZZLE SPECIFICATIONS:
Specifications Dimensions
Length 30mm
Inlet diameter 30mm
Exit diameter 16mm
Chevron length 5.44mm
Wedge 3.08mm
Table 3.1 Nozzle specifications
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Fig 3.1 Two dimensional free jet nozzle
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Fig 3.2Three dimensional free jet nozzle
Fig 3.3 Drafted views of free jet nozzle
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Fig 3.4 Chevron nozzle with 8 count
Fig 3.5 Drafted views of chevron nozzle
Fig 3.3 is the three dimensional diagram of nozzle with chevron. Todraw the nozzle with chevron follows the same as nozzle. After
developing the free jet nozzle then draw chevron which seems like
saw tooth. To draw chevron first choose the appropriate plane and
go to sketcher. Then draw the chevron with correct dimensions for our
specifications. After finishing sketch exit the sketcher. Then select the
chevron and choose the pocket option to pocket the chevron. After
pocketing the chevron choose the circular pattern in the transformation
features. Thus the chevrons are created in the exit of nozzle with our
dimensions and the chevron count is 8.
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Fig 3.6 Chevron nozzle with wedge thickness 1mm
Fig 3.7 Drafted view
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Fig 3.8 Chevron nozzle with wedge thickness 2mm
Fig 3.9 Drafted view
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3.2PREPROCESSING
ICEM CFX is the pre-processor tool used to mesh the model in
order to get the accurate result.
3.2.2THREE DIMENSIONAL MESHING
Import the Three dimensional iges file or step file into the ICEM
CFX. Then do any clean up operations if necessary. Then create the
domain according to requirement which may be rectangular brick or
cylinder for our convenient. The zone height is to be 5D to 10D and
length to be 30D to 40D range or to be greater than that too. After
creating the zone we place the domain at the exit of the nozzle by using
copy option. We should unite the two volumes as a single volume
.Then the domain can be decomposed according to the requirement.
Then the mesh has been made with the size of 1 which can be reduced
according to our requirement for better results.
MESH TYPE
Combination of Quadra and Hexa mesh for nozzle and for
boundary domain tetra mesh type.
3.2.3 CHEVRON NOZZLE WITH AND WITHOUT WEDGES
In this follow the same as in above instead of uniting the
volumes here subtract the volumes by using the subtract option i.e.,
subtracting the nozzle from domain. After subtracting there will be one
volume only. Before subtracting place the nozzle inside the domain for
the convenience. Then mesh the faces using the face mesh operation by
tria elements under pave scheme. We can mesh the edges of the
volume and then do face and volume mesh if needed. In the edge mesh
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select each edge of the volume and to give number of nodes in each
edge based on the need. After finishing edge mesh then select the face
mesh. In the face mesh we have to select each face and mesh. In the
face mesh we may select the elements as QUAD or TRIA.After
finishing the face mesh then go for the zones operation. In the Zones
we have to select the INLET, OUTLET,WALL, AXIS, PRESSURE
INLET, PRESSURE OUTLET, etc,. After applying the zones we may
check the quality of the mesh by using quality operation. After
finishing all above steps have to save the mesh and to export the mesh.
3.2.4BOUNDARY CONDITIONS
The below diagram illustrates the boundary conditions of 3dimensional
models. In ICEM CFX select the pre-processor modules and feed the
input for the respective Mach number.
Fig 3.10boundary conditions
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Fig 3.11 Three Dimensional free jet Mesh
Fig 3.11 is the meshing model of base line nozzle. The meshing
process for this model follows the above procedure with little variation.
In which nozzle volume and zone volume are united to a single
volume. Then this volume is meshed by using volume mesh operation
using tetra and hex elements scheme and it contains 835710 elements.
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Fig 3.12Chevron nozzle mesh
Fig 3.12 is the meshing model of chevron nozzle. The meshing
process for this model follows the above procedure with little variation.
In which nozzle volume and zone volume are subtracted to a single
volume. Then this volume is meshed by using volume mesh operation
using tetrahedral elements under and it contains 363094 elements.
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Fig 3.13 Chevron nozzle with wedge thickness1mm mesh
Fig 3.13 is the meshing model of base line nozzle. The meshing
process for this model follows the above procedure with little variation.
In which nozzle volume and zone volume are subtracted to a single
volume. Then this volume is meshed by using volume mesh operation
using tetrahedral (boundary) & hex(nozzle) elements under scheme and
it contains 453731 elements.
Fig 3.14 Chevron nozzle with wedge thickness 2mm mesh
The above meshing model contains 456208 tetrahedral elements
and meshing under scheme. The procedure is same as for the previous
one.
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3.3SOLVER
ICEM CFX has a post processor or analyzing software used to
simulate the model for their application.
3.3.1STEPS TO SOLVE
After finishing the pre-processor and then have to import our
meshed model into the software by read the case file such as mesh file.
After reading the case file have to check the grid and zones. Then go to
define select the model and select the solver. Then click the viscous
model and select the appropriate viscous model for your problem. Here
we selected k-omega (used for turbulent analysis).
Then click materials in define option and change the material to
ideal gas and click ok. Then click operating condition and give the
operating condition value in Pascal or any other unit. Then click
Boundary conditions and give the pressure values for boundary
conditions such as pressure inlet, pressure outlet, pressure far field, etc,
Then go to solve select initialize and select initialize in which select the
inlet condition and click apply. Then go to monitors in solve and select
residual and click ok for print and plot.
After finishing all above steps click iterate in solve and give
value for number of iterations and start the iteration. After the solution
is converged and iteration will stop and we see the results. Go to the
display menu select the contours and select which one to display such
as velocity, pressure, etc, and click display and the contour will display
in different colours for variations of values. We can also see the XY
plot for velocity, pressure, etc, by selecting plot menu and select XY
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plane and select pressure ,velocity ,etc,. Thus the meshed model is
analyzed using CFX following the above procedure.
3.3.2FORMULA TO FIND THE TOTAL PRESSURE
PO/P M2)/-1
P= 101325 Pascal
PO/P 0.42
)1.4/1.4-1
PO= 113134.6279 Pascal
3.3.3 BOUNDARY CONDTION PRESSURE VALUES
Mach number Total pressure
(pascal)
Gauge pressure
(pascal)
Operating
pressure
(pascal)
0.4 113134.6279 101325 0
0.6 129240.4201 101325 0
0.8 154453.7514 101325 0
Table 3.2 Boundary conditions pressure values
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CHAPTER 4
RESULT AND DISCUSSION
Potential core region in the subsonic flow is the region where thecenter line velocity is same as the velocity at the exit of the nozzle. The
Mach number along the center line was calculated from the total
pressure and static pressure. The static pressure across the jet is
assumed to be the same as the surrounding ambient pressure. The
assumption is perfectly valid for subsonic jets.
In this chapter we are discussed about the centreline mach
number decay for chevron, chevron with 1mm wedge and chevron with
2mm wedge. Here we plot dimensional and non-dimensional graphs
for all subsonic mach number for all configuration. The plots are
obtained from the post processor ICEM CFX. We discussed and
compared the results of center line Mach number decay for all subsonic
mach number to each configurations.
Streamwise vortices are generated when wedges and chevron are
introduced at the exit of the nozzle. They act as effective mixing
promoters and enhance the jet mixing. These faster mixing results in
rapid jet decay of chevron with wedge compared to a jet from a plane
nozzle. The chevron increase vortices in the exit nozzle along with
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wedges also induce vortices and increase the jet mixing and help to
reduce the potential core.
The contour plot obtained from the CFD is dimensional and
convert it in to non dimensional by excel. The graph is plotted between
X/Dj and M/Mj. The length of the potential core is divided by the exit
diameter and Mach number is divided by the jet Mach number. Both in
dimensional and non dimensional plot we can view the Mach number
decay clearly.
4.1 EFFECT OF CENTER LINE MACH NUMBER FOR FREE
JET
Fig 4.1Mach number plot for free jet M= 0.4
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Fig 4.2Centerline Mach number decay of free jet
The above graph shows the centerline Mach number decay for
the Mach number 0.4 and is plotted between X/Dj and M/Mj. Here the
potential core decays at 5.5D.
Fig 4.3Mach number plot for free jet M= 0.6
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Fig4.4Centerline Mach number decay of free jet
The above graph shows the centerline Mach number decay for
the Mach number 0.6 and is plotted between X/Dj and M/Mj. Here the
potential core decays at 5.81D. When comparing with 0.4 Mach
number of free jet the potential core length is high.
Fig 4.5Mach number plot for free jet M= 0.8
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Fig 4.6Centerline Mach number decay of free jet
The above graph shows the center line Mach number decay for
the Mach number 0.8 and is plotted between X/Dj and M/Mj. Here the
potential core decays at 6.12D. When comparing with 0.4 and 0.6
Mach number of free jet the potential core length is high.
4.2 CENTER LINE MACH NUMBER DECAY OF CHEVRON
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Fig 4.7Mach number plot for chevron M= 0.4
Fig 4.8Centerline Mach number decay of chevron
The above centerline Mach number decay graph is plotted
between M/Mj and X/Dj for the chevron nozzle for the Mach number
0.4. The potential core decay starts from 3.88D. When comparing with
0.4 Mach number of free jet the potential core length is 30 % reduced.
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Fig 4.9Mach number plot for chevron M= 0.6
Fig 4.10Centerline Mach number decay of chevron
The above centerline Mach number decay graph is plotted
between M/Mj and X/Dj for the chevron nozzle for the Mach number
0.6. The potential core decay starts from 3.93D. When comparing with
0.6 Mach number of free jet the potential core length is 33 % reduced.
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Fig 4.11Mach number plot for chevron M= 0.8
Fig 4.12Centerline Mach number decay of chevron
The above centerline Mach number decay graph is plotted
between M/Mj and X/Dj for the chevron nozzle for the Mach number
0.8. The potential core decay starts from 3.93D. When comparing with
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0.8 Mach number of free jet the potential core length is 36 % reduced.
4.3 CENTER LINE MACH NUMBER DECAY OF CHEVRON
WITH WEDGE
Fig 4.13Mach number plot for chevron with wedge thickness 1mm
M= 0.4
Fig 4.14Centerline Mach number decay of chevron with
wedge1mm
The above centerline Mach number decay graph is plotted
between M/Mj and X/Dj for the chevron nozzle with 1mm wedge
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thickness for the Mach number 0.4. The potential core decay starts
from 3.17D. When comparing with 0.4 Mach number of free jet and
chevron the potential core length is reduced to 42 % and 18%
respectively.
Fig 4.15Mach number plot for wedge thickness1mm M= 0.6
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Fig 4.16Centerline Mach number decay of chevron with wedge
1mm
The above graph is plotted between M/Mj and X/Dj for the
chevron nozzle with 1mm Wedge thickness for the Mach number 0.6.
The potential core decay starts from 3.05D. When comparing with 0.6
Mach number of free jet and chevron the potential core length is
reduced to 47 % and 22% respectively.
Fig 4.17Mach number plot for wedge thickness1mm M= 0.8
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Fig 4.18Centerline Mach number decay of chevron with wedge
1mm
The above centerline Mach number decay graph is plotted
between M/Mj and X/Dj for the chevron nozzle with 1mm Wedge
thickness for the Mach number 0.8. The potential core decay starts
from 3.05D. When comparing with 0.8 Mach number of free jet and
chevron the potential core length is reduced to 50% and 22%
respectively.
Fig 4.19Mach number plot for wedge thickness 2mm M= 0.4
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Fig 4.20Centerline Mach number decay of chevron with wedge
2mm
The above centerline Mach number decay graph is plotted
between M/Mj and X/Dj for the chevron nozzle with 2mm Wedge
thicknessfor the Mach number 0.4. The potential core decay starts from
2.0D. When comparing with 0.4 Mach number of free jet and chevron
and Wedgewith 1mmthickness the potential core length is reduced to
63%, 48% and 40% respectively.
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Fig 4.21Mach number plot for wedge thickness 2mm M= 0.6
Fig 4.22Centerline Mach number decay of chevron with wedge
2mm
The above centerline Mach number decay graph is plotted
between M/Mj and X/Dj for the chevron nozzle with 2mm Wedgethickness for the Mach number 0.6. The potential core decay starts
from 2.0D. When comparing with 0.6 Mach number of free jet and
chevron and Wedge with 1mmthickness the potential core length is
reduced to 65%, 49% and 34% respectively.
Fig 4.23Mach number plot for wedge thickness 2mm M= 0.8
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Fig 4.24Centerline Mach number decay of chevron with wedge
2mm
The above centerline Mach number decay graph is plotted
between M/Mj and X/Dj for the chevron nozzle with 2mm Wedge
thickness for the Mach number 0.8. The potential core decay starts
from 2.0D. When comparing with 0.8 Mach number of free jet and
chevron and Wedge with 1mmthickness the potential core length is
reduced to 67%, 49% and 34% respectively.
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CHAPTER 5
CONSTRAINTS AND RECTIFICATION
To evaluate the Reduction in noise we need experimental setupbut we have preceded in such a way that if there is a reduction in
pressure, noise will reduce.
CONCLUSION
The Wedges and chevrons were used in the presents study,
namely chevron, chevron with 1mm wedge thickness, chevron with
2mm Wedge thickness were found to be effective in distorting the jet
structure. From the figure 5.1, fig 5.2, fig 5.3 among the three
configurations, chevron with 2mm Wedge was found to be more
efficient in reducing the potential core length and distorting the jet
structure compared to the other two configurations in all subsonic
mach numbers. From the above figures, Nozzle with the chevron with
2mm Wedge , the maximum core length reduction achieved was 67 %
of uncontrolled jet , followed by rapid decay of jet center line mach
number, where the chevron with 1 mm Wedge reduce the core length
to 50 % and chevron reduce the core length to 40 % at all subsonic
mach numbers. Chevron with 2 mm Wedge is showing consistently
better performance for all mach numbers compare to other two
configurations. It may due to that fact the chevron with 2mm Wedge
creates stronger pressure gradient shedding stronger streamwise
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vortices causing reduction in potential core and enhanced mixing.
Mach number profiles at different X/D locations shows the chevron
with 2mm Wedge is more effective spreading of the jet normal to the
wedge compare to other two configurations.
Chevron technology has provided a modest relief. Unfortunately,
a complete understanding of jet noise mechanisms is still not in our
grasp. The insight of fundamental experiments coupled with
application of CFD allowed the development of the subject
technology with tools slightly better than cut-and-try. Hope forfurther control and reduction of jet noise hinges on advancement of
our understanding of the relevant mechanisms.
Fig 5.1 Comparison of Centerline Mach number decay
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Fig 5.2 Comparison of Centerline Mach number decay
Fig 5.3 Comparison of Centerline Mach number decay
COMPARISON OF POTENTIAL CORE LENGTH
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MACH
NO
FREE JET CHEVRON WEDGE 1mm WEDGE 2mm
0.4 5.50D 3.88D 3.17D 2.00D
0.6 5.81D 3.93D 3.05D 2.00D
0.8 6.12D 3.93D 3.05D 2.00D
Table 5.1Comparison of potential core length
CHAPTER6REFERENCES
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