Aero Nozzle Acoustic

8/9/2019 Aero Nozzle Acoustic

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CHAPTER 1

1. INTRODUCTION

Aircraft noise is one of the major contributors of noise pollution to the

environment. The growth of air transportation has been leaps and bounds and thus the

focus on aircraft noise has been heightened. Aircraft noise is mainly due to three

components:

Aerodynamic Noise

Engine and other mechanical noise

Noise from aircraft systems

There are two main sources of noise in todays commercial aircraft engines:

fan/compressor noise and jet noise. Jet noise comprises turbulent mixing noise and, in the

case of imperfectly expanded jets, shock noise.Turbulent mixing noise is very difficult to

control, and so its suppression remains a challenge. It is generally agreed that turbulent

shear flow mixing causes two types of noise: sound produced by the large-scale eddies and

sound generated by the fine scale turbulence. The former is very intense and directional

and propagates at an angle close to the jet axis. The latter is mostly uniform and affects the

lateral and upstream directions. The increase in bypass ratio over the last three decades has

resulted in a dramatic suppression in the jet noise of turbofan engines. Modern engines are

so quiet that further reduction in noise becomes extremely challenging. The success of the

high-bypass engine is offset, to some degree, by the increasing volume of aircraft

operations. This creates more environmental and political pressures for quieter aircraft.

Today the most successful technique for reducing jet noise from high-bypass engines

involves the installation of chevron mixers on the exhaust nozzles.

1.1 Noise Control (Suppression)

Jet engine noise suppression has become one of the most important fields of

research due to airport regulations and aircraft noise certification requirements. These

govern the maximum noise level aircraft are allowed to produce. Although airframe

generated noise is a factor in an aircraft's overall noise signature, the principal source of

the noise is in the engine.


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1.2 Chevrons:

Chevrons reduce the jet noise component of the engine noise. Since jet noise is

important during take-off, the benefit of chevrons is best realized during that portion of a

commercial flight.

Chevrons are zigzag or saw-tooth shapes at the end of the nacelle, with tips that are

bent very slightly into the flow. This creates vortices that form at each chevron, enhancing

the mixing rate of the adjacent flow streams. As previously mentioned, the jet noise is due

to turbulent mixing between jets and the noise generation mechanism is very complex.

When the chevrons enhance mixing by the right amount, the total jet noise reduces. If the

mixing is too much, the chevrons make the noise go up. If the mixing is too little, no noise

reduction benefits are realized.

The chevron design was successful, with a reduction of up to 4 decibels of noise at

the peak frequency and noisiest location was achieved.

How mixing is obtained?

Mixing noise is by far the most difficult to control. For greater exhaust speed, large-

scale mixing noise manifests itself primarily as Mach wave radiation, caused by the sonic

convection of turbulent eddies with respect to the ambient. In the vicinity of the potential

core, Mach waves are strong, nonlinear pressure waves that decay into weak acoustic

waves far from the jet. In high-speed hot jets, Mach wave emission is the dominant source

of sound and radiates in the aft quadrant.

The simplest way to explain Mach wave radiation is to consider the turbulent

interface between the jet and the ambient as a wavy wall propagating at a convective speed

Uc. When Uc is supersonic, Mach waves are radiated from the wall. The notion of sound

radiation from large scale flow instabilities was first confirmed in the supersonic jet

experiments of McLaughlin et al[3] and the subsequent experiments of Troutt and

McLaughlin[4]. In those experiments, the orientation, wavelength, and frequency of the

measured acoustic radiation were found to be consistent with the Mach wave concept just

outlined.

The linear stability analysis of Tam and Burton further solidified this idea by

showing that the sound emitted by a supersonic instability wave matched very well the


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aforementioned experimental data. Since then, a large volume of experimental and

theoretical works have addressed multiple aspects of this problem.

1.3 Co-flow:

The effectiveness of the co-flow in reducing Mach waves depends primarily on the

following factors:

1).The convective Mach number of the jet eddies relative to the co-flow, which

should be subsonic; the lower its value, the faster the attenuation of the pressure fluctuation

within the co-flow layer.

2).The co-flow thickness; the larger it is, the more room a disturbance has to decay

sub sonically before it is transmitted to the ambient fluid.

3).The coverage of the Mach wave emitting region of the jet by the co-flow; if the

co-flow mixes fully with the jet before the end of this region, Mach waves will still be

generated.

The underlying principle is to enhance the axial decay of the velocity in the jet

plume, hence reducing the length of the Mach wave emitting region; at the same time, all

of the other sources of noise that depend on jet velocity would also be reduced.

1.4 Decibel (Loudness) Comparison Chart

Here are some interesting numbers, collected from a variety of sources, which help

one to understand the volume levels of various sources and how they can affect our

hearing.


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Environmental Noise

Weakest sound heard 0dB

Whisper Quiet Library at 6' 30dB

Normal conversation at 3' 60-65dB

Telephone dial tone 80dB

City Traffic (inside car) 85dB

Train whistle at 500', Truck Traffic 90dB

Jackhammer at 50' 95dB

Subway train at 200' 95dB

Level at which sustained exposure may

result in hearing loss90 - 95dB

Hand Drill 98dB

Power mower at 3' 107dB

Snowmobile, Motorcycle 100dB

Power saw at 3' 110dB

Sandblasting, Loud Rock Concert 115dB

Pain begins 125dB

Pneumatic riveter at 4' 125dB

Even short term exposure can cause

permanent damage - Loudest

recommended exposure with hearing

protection

140dB

Jet engine at 100' 140dB

12 Gauge Shotgun Blast 165dB

Death of hearing tissue 180dB

Loudest sound possible 194dB

Table 1.1 various decibels (Loudness) Comparison


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2. LITERATURE REVIEW

2.1 Experimental Aero acoustics

The science of aero acoustics has seen a great deal of development over the past

fifty years through a range of analysis strategies such as experimental diagnostics,

numerical simulation and computational modeling. Numerical simulation techniques have

seen recent progress in their ability to solve the flow equations in reasonable computational

time. However, their application is limited due to the finite number of mesh points that can

be selected around the extremely thin boundary of the initial shear layers and at capturing

the small-scale structures as the jet develops. As a result, the range of Reynolds numbers

that can be simulated is limited, as well as the accuracy to which the turbulent motion of a

heat-conducting, viscous, compressible flow can be resolved. This, coupled with the short

run times, make it problematic to obtain full-scale converged solutions. In the light of this,

the experimental approach has the advantage of the flow equations being perfectly solved

by the physical flow itself. Given that current techniques have progressively become more

sophisticated due to the improvements in measurement technology, as well as increasingly

innovative data extraction and analysis techniques, scientists have come closer to

understanding the noise source mechanisms of turbulent jets. Experimental research in its

earliest stages focused on measuring the fluctuation in a flow field with the use of hot-

wires, hot-films and pitot- tubes. Seinerand Reethof (1974)used a hot-wire anemometer

to investigate the velocity fluctuations of a Mach 0.32 jet in an attempt to highlight the

contributions from shear noise; the interaction of the turbulent structures with the mean

flow of the jet. It was found that this noise source is the dominant mechanism, with the

transition region of the jet producing most of the noise.

However, experimental uncertainty was a serious concern when using such techniques due

to the possibility of the in-flow probes producing greater sound than the actual source that

they were designed to measure. This led to developments that aimed to minimize these

additional noise sources through the use of non-intrusive measurements. Laser Doppler

Velocimetry (LDV) was used by Schaffer (1979)to study the unbounded Mach 0.97 jet.

The observations led to the conclusion that the majority of noise measured at the

downstream angles between 20 to 30 to the jet axis is generated by the shear noise source

mechanism in the transition region, 5 to 10 diameters downstream of the nozzle exit.

Microphones have provided an effective means of exploring the dynamics of experimental


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jet aero acoustics. Arrays configured in both the axial and radially positioned azimuthal

directions provide advantages in examining the large-scale coherent structures in the jets

near and far-field. Small-scale turbulent structures however, dissipate more rapidly and can

be naturally filtered out before reaching the far-field microphones. Hence, it is complicated

to correctly interpret the measured data, as it is not evident how much information is being

filtered. The microphones pick up contributions from both the hydrodynamic and acoustic

pressure fields from the turbulent flow and if these are not distinguished, the results may

be misleading.

2.2 Jet flow Structure

Figure 2.1 illustrates the regions that constitute a model for the structure of a free

jet. The precise length of these regions is a function of the Reynolds number as well as the

nozzle exit conditions. The transition from subsonic to supersonic flow also affects the

length of these regions due to formation of complex shock cells near the nozzle exit. The

existence of a shock cell structure is a common feature of imperfectly expanded jet flows.

The region between the jets core and ambient flow experiences the steepest velocity

gradients within the jet structure and is defined as the shear layer. The shear layer is

dominated by small-scale structures, which are highly dissipative in nature and responsiblefor producing aerodynamic sound due to their high frequency, small-amplitude

oscillations. The highly unstable shear layer is also characterized by significantly higher

levels of vorticity as compared to its core. Within the mixing region lies the potential core

of the jet which is defined as a region of almost uniform mean centerline velocity from the

nozzle exit (Lilley, 1958).The gradual decay of the potential core is due to the interaction

between the jet and the surrounding stagnant or lower velocity atmosphere. A significant

development occurred in the understanding of turbulence through the studies carried out by

Crow and Champagne (1971). Large-scale turbulent structures were found to be a more

dominant factor in the overall mixing process of jets as compared to the small-scale

turbulent structures. The frequencies at which these peak noise radiating structures

propagate are known as the preferred modes of the jet .When a jet is disturbed at its

preferred mode it enhances the formation of these large-scale turbulent structures,

particularly at the end of the potential core (Panda et al., 2002).


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Witze (1974) used an extensive compilation of experimental data on supersonic

jets to derive a generalised equation for the centreline velocity decay of a jet. This can be

re-arranged to give an expression for the length of the potential core: X core=0.7/k ()0.5

where the axial length, Xcore, is non-dimensionalised by the radius of the jet and the exit

ambient density, a, normalized by the density of the jet. The proportionality constant, k, is

a function of the jet Mach number and for supersonic flows is denoted by;

k=0.063(MJ2-1)-0.15

The acoustic near field of the jet is a dominated by turbulent fluid motion which involves

the conversion of the energy from rotational hydrodynamic modes to irrational propagative

modes, a process about which little is known (Jordon, 2005). Hydrodynamic modes and

structures are characterized by non-linear and non-spherical propagation, in the near-field

region, up to 30 nozzle diameters from the jet exit (Petitjean, 2003). These noise

generating structures are believed to be non-localized, hence making it difficult to interpret

the acoustic near-field. The region beyond this 30D boundary is referred to as the acoustic

far-field, where the propagation of sound waves is spherical and the noise intensity is

proportional to the inverse square of the radial distance from the jet (Petitjean, 2003).The

initial region of a dual-stream jet consists of two shear layers: the primary shear layerbetween the primary and secondary streams and the secondary shear layer between the

secondary and ambient streams (Fig. 2.2). The primary shear layer encloses the primary

potential core. The region between the primary and secondary shear layers defines the

secondary core, which contains an initial potential region followed by a non-potential

region. This broad definition of the secondary core is essential for understanding the

acoustic benefits of certain dual-stream configurations. In many practical cases, for

instance, in coaxial jets of turbofan engines, the secondary core ends upstream of the end

of the primary core. The flow past the end of the secondary coren consists of a single shear

layer between the jet centerline and the ambient stream, thus having the characteristics of a

single-stream jet. It is useful, therefore, to divide the jet flow into two regions: the

compound region, that is, the region before the end of the secondary core, and the simple

region, which is the region past the end of the secondary core. The extent of the compound

region, relative to the length of the primary potential core, is critical for noise reduction.


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2.3 Source of jet noise

Many attempts have been made to derive a jet noise prediction theory since the

birth of aero acoustics in the 1950s. A literature survey will reveal many models and

semi-empirical theories that aim to quantify jet noise. However its understanding is

directly linked to the understanding of turbulence, which even to this day remains

incomplete.

2.3.1Turbulent Mixing Noise

The three accepted sources of noise produced from supersonic jets are turbulent

mixing noise, Mach wave radiation and shock associated noise, which can either be

broadband or a discreet screech tone (Crighton, 1977). Mixing noise is a product of the

interaction between the flow structures in the shear layer region of the jet with the

surrounding ambient air. Research on jet noise at NASA has highlighted that turbulent

mixing noise of high-speed jets consists of two distinct components (Tam et al., 1996).

This claim was supported by the fact that over 1900 measured frequency spectra from

supersonic jets closely fitted the distinct spectral shape of the two principal noise sources

as shown in Figure 2.3, regardless of the jet velocity, temperature, and direction of

radiation. The first of these two distinct sources radiates principally in the downstream

direction and is consistent with Mach wave radiation, due to the large-scale turbulence

structures or instability waves of the jet flow. The second component is dominant in the

sideline and upstream directions suggesting it is noise from the fine-scale turbulence of the

jet flow. It has more recently been shown that even the noise spectra of non- axisymmetric

jets including jets from rectangular, elliptic, plug, and suppressor nozzles fits the same two

experimentally determined spectra, indicating that the noise sources of these jets are

similar to those of the circular jet (Tam 1998). Experimental work on both round and

bevelled nozzles by Viswanathan (2008) confirmed this pattern by demonstrating the

sharp contrast in the nature of these dominant sources. The acoustic spectrum characteristic

of the random incoherent sources radiating at 90, was found to be somewhat flat,

indicating low intensity noise from the rapidly dissipating fine-scale turbulence structures.

Highly coherent, large-scale turbulent structures on the other hand were shown to be

responsible for the peak noise events, which for a Mach 1.3 jet are known to radiate

towards a downstream angle of 30 to the jet axis (Hileman et al., 2002). The intensity and

directivity of the turbulent mixing noise of supersonic jets depends on the Mach number


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and the ratio of jet temperature to the ambient temperature (Tam 1995). Higher frequency

turbulent mixing noise originates from the instabilities in the shear layer close to the nozzle

exit, whereas noise related to lower frequencies originates from further downstream

(Bishop et al., 1971). The locations of large noise events have been identified using a

microphone array. It was estimated that 98% originated from between 1-12 diameters

downstream from the nozzle exit. Flow visualization techniques have been used to

illustrate that high amplitude noise events are characterized by large flow scales and that

quiet periods coincided with smaller flow scales. Lower frequency noise an event from

large-scale turbulent eddies is found to correspond to Strouhal numbers of approximately

0.15-0.25 (Tanna, 1977; Tam, 1995; Panda et al., 2002) , which equates to the Kelvin-

Helmholtz roll-up frequency, also known as the preferred mode of a jet. Kelvin-Helmholtz

instability waves, which occur between two parallel streams of varying velocity and

density such as those that exist in the shear layer of a jet, can lead to Mach wave radiation

when convected at supersonic speeds (Panda et al., 2002).The noise generated from both

large and fine-scale turbulent structures is highly directional in nature due to the effects of

convection and refraction(Ribner, 1969; Tanna, 1975). Acoustic wave refraction occurs

within the mixing region of the jet between the shear layer and the potential core of the jet.

The higher flow velocity towards the core of the jet deflects the propagating sound waves,of wavelengths equal to or smaller than the jet diameter, away from the flow direction.

This effect is more dominant in heated jets due to greater changes in temperature and

hence density. This mean shear flow refraction reduces the amount of sound radiated in the

direction of the jet flow and hence a results in relatively quieter area surrounding the jet

axis which is called the =cone of silence(Tam 1998). Emphasis on evaluating the

contribution of large-scale structures on the sound field of supersonic jets is justified as the

acoustic loads are a principal source of structural vibration (Petitjean et al., 2003)which

may contribute to structural airframe fatigue and effect the operation of guidance and

control systems and their supporting structure.

2.3.2 Screech Noise

Shock cell structures are usually formed in imperfectly expanded supersonic jets,

due to a mismatch of static pressures between the jet and surrounding atmosphere (Tam et

al., 1995). The development of oblique shocks or expansion waves in the jet plume serves

the purpose of balancing this difference of pressure for over expanded or under expanded


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flows respectively (Tam et al., 1995). Shock associated noise is considered to be a product

of the interaction between the turbulent structures induced at the nozzle and the diamond

shaped shock cell structure shown in Figure 2.1 (Henderson et al., 2008). In the presence

of the shock cells, the jet emits two additional components of noise in addition to mixing

noise, which are referred to as screech and broadband shock associated noise. Screech

noise is generally observed in acoustic spectra acquired at upstream and sideline

observation angles of cold laboratory jets (Henderson et al., 2008)and is a characteristic

found in most imperfectly expanded jets. The screech tone component is discrete in nature,

high in amplitude and is associated with various harmonics that are generated by an

acoustic feedback of the disturbances at the nozzle lip. The associated harmonics move

through the shock cell and are fed back to the nozzle lip though an external route outside

the shock cells. As a result another disturbance is created which propagates downstream,

hence giving a self-sustaining and repeating cycle (Neemah et al., 1999). The pioneering

work of Powell (1953) highlighted that there are two components of screech noise: its

fundamental and secondary harmonic tones. He derived relationships that describe the

directivity pattern of these tones by considering the shock cells as being adjacent stationary

sources of equal strength and hence determining the phase between them. He found that

the fundamental component peaked closer towards an upstream direction to the jet flow, asshown in Figure 2.3, whereas the harmonic component peaked closer to the sideline angle

of 90. Powell also presented an equation for quantifying the, frequency of the

fundamental component, fF, of screech noise, shown in equation (2.10), where UC is the

convective speed of the hydrodynamic structures, s is the shock cell spacing and MC the

convective Mach number. The accuracy of this formula depends heavily on the shock cell

spacing, s, which can be predicted numerically using the Prandtl-Pack formula shown in

equation (6) (Singh & Chatterjee, 2007).

The frequency for the first harmonic component of screech noise, fH, can be estimated as

being twice that of the fundamental (Raman, 1999): Earlier research in relation to shock

associated noise was concentrated on screech tone abatement, due to its potential to cause

structural damage if amplified to above a certain level. Several methods have been

investigated for the purpose of screech tone abatement which include: porous plugs, co-

annular jets, tabs and nozzle asymmetry. Even today, 60 years from the discovery of

screech noise, it plays a vital role in the design of aircraft propulsion systems due to the

ability of its harmonic component, which can reach as high as 170dB, to cause sonic


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fatigue failure as occurred on the British Aircraft Corporation VC-10 and on the F-15 and

B1-B of the United States Air Force (Raman, 1999). The jet Mach number, temperature,

nozzle lip thickness, and the existence of noise reflecting surfaces around the immediate

vicinity of the jet have been all shown to have an effect on screech noise intensity. An

increase in the nozzle lip thickness can increase screech intensity by up to 10dB (Norum,

1983), simply by acting as a sound-reflecting surface and hence amplifying the effects of

the acoustic feedback loop from the shock cells. Acoustic measurements by Kaji et al.,

(1996) have provided evidence to suggest that shock noise is most intense somewhere

between the third and fourth shock cell structures from the nozzle exit. However due to the

non-linearity of the feedback loop, an empirical equation to describe the screech intensity

is not yet available (Kandula, 2008).

2.3.3 Broadband Shock Noise

The characteristics of broadband shock associated noise have been explored for

over 30 years since the pioneering work of Harper-Bourne and Fisher (1973) who

developed the first empirical model for such noise. Converging-diverging nozzles,

operating at design Mach number with fully and perfectly expanded supersonic jets,

exhibit noise related to turbulent mixing only. For imperfectly expanded regimes however,

there exist shock cell structures in the flow for the purpose of correcting the flow pressure

as described earlier. The interaction of these shock cells with other structures in the mixing

region of the jet, such as instability waves (Mach wave radiation) and small-scale turbulent

eddies, leads to the amplification and generation of intense acoustic waves of which a

component is broadband shock noise (Kandula, 2008). The resultant acoustic spectrum,

like the one shown in Figure 2.4, highlights this interaction noise as sharp peaks at the

sideline and upstream angles of a jets acoustic field.

Broadband shock noise is evidently the dominant characteristic of the upstream

spectra of a jets acoustic field. Tam (1995) highlighted other distinct features of

broadband shock noise based on experimental evidence from a converging-diverging Mach

1.67 jet by Norum and Seiner (1982). Firstly, the frequency that corresponds to the

broadband shock peak, increases with the observer angle from the jet axis and is highest at

the upstream angles. Secondly, the broadband peak consists of several sub-scaled mini


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peaks, which are most distinct at the sideline angle of 90. Thirdly, the broadband shock

noise spectral bandwidth increases with observer angle and is once again highest at the

upstream angles beyond the sideline position. Finally, the sound pressure level of the

broadband shock peak is independent of the jet temperature; however the corresponding

frequency increases with an increase in jet temperature. Recent experimental and

theoretical work has shown that another type of broadband shock-noise, which radiates

primarily in the downstream direction, is also present in dual stream (co-annular)

supersonic jets. This form of the broadband shock cell noise has low intensity at the

sideline (90) observer angle and is believed to be generated by the interaction of large

turbulence structures and the shock cells in the inner shear layer of the jet (Tam et al.,

2008).

2.4 Passive Methods for Jet Noise Reduction

2.4.1 Swirling Jets

Before the 1980s the primary application of swirling jets was in combustion chambers for

the purpose of enhancing mixing efficiency (Lilley, 1977). However, experimental studies

on full-scale turbojet engines by Schwartz (1975) proposed the use of swirl as a potential

application for jet noise reduction in an era coinciding with the Concorde noise reduction

programme. Swirl was generated through stationary vanes circumferentially placed around

an annular cross-section just upstream of the nozzle exit. Changes in overall sound

pressure levels (OASPL) of -4dB and +1dB were recorded at angles of 30 and 90 to the

downstream jet axis respectively, with 40% of the overall mass flow ratio of the jet passing

through the swirl generating vanes. However, the mean centreline velocity of the

supersonic jet (~650m/s) was reduced by the introduction of a tangential velocity

component, equating to a loss of thrust by almost 2%. The interaction of centrifugal forces

with free turbulent shear flows was found to increase proportionally with the strength of

the swirl. The strength of swirl was controlled by changing blade angle of the swirl

generating vanes from 40 to 60 to the flow axis, but no quantitative measure of its

strength was presented. Carpenter & Johannesen (1975) used the classical one-

dimensional theory for choked, convergentdivergent nozzles and validated previous

experiments in relation to swirl having a negative impact on engine performance. They

also pointed out that the key to analysing swirling propulsive jets is the choice of which

parameters to keep constant between swirling and non-swirling cases (Carpenter &


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Johannesen, 1975). It was convincingly argued that the nozzle pressure ratio be maintained

such that the thrust parameter be fairly observed when the jet is subjected to swirl. This

will therefore be taken as the justification for maintaining constant plenum pressure in this

study. The Swirl number, S, is the parameter used to indicate the contribution of the

tangential swirling component in an axial flow and can be defined as the ratio of the

rotational to axial momentum as shown in equation(Beer & Chigier, 1972). Relatively low

levels of swirl, corresponding to S


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elimination of screech noise. DNS studies have shown that the dynamics of large-scale

structures are strongly affected by the degree of swirl (Kollmann et al., 2001). This

supports the concept of introducing swirl, as the resultant changes to shear layer dynamics

should affect noise production. RANS and LES simulations have shown swirl to affect

noise levels by increasing turbulent kinetic energy within the flow profile, as well as

reducing the potential core length of the jet (Tucker et al., 2003). Gilchrist & Naughton

(2005) investigated the effect of swirl on the near-field flow of an incompressible swirling

jet of Reynolds numbers ~1.010. For low levels of swirl, characterised by Swirl numbers


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lobe and have been found to reduce low frequency noise at the expense of increased high

frequency noise in the near-field (Callender et al., 2007). Mechanical chevrons were

initially developed by Boeing, General Electric and NASA specifically for reducing jet

noise from the aft angles towards the rear of the engine as shown in Figure 2.5. These

devices suppress the overall noise by promoting rapid mixing of the propulsive jet with the

surrounding flow (Knowles et al., 2005). However, alterations to the nozzle geometry such

as these may incur a weight penalty, as well as increased drag and turbulence during cruise

due to the induced mixing by the serrated shape of the nozzle. Hence, the design of

compact, lightweight and actively controllable noise reduction methods is the focus of

considerable recent research.

An interesting three point criterion was suggested by Alkislar & Krothapalli (2007) to

maximize the effectiveness of the stream-wise vortices on the aero acoustics of the jet in

terms of their location, strength and persistence.

The strength of the vortices should be enough to disrupt only the large-scale

coherent structures.

Any additional turbulent kinetic energy produced by the vortices should be aimed

mostly at the low-speed, outer edge of the shear layer.

There are optimum values for the initial strength and separation of the stream wise

vortices.

2.5 Effect of chevron count and penetration on the acoustic characteristics of chevron

nozzles

Experimental investigations 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. 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 0 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 reduction at low nozzle pressure ratios(P.S.Tide, K.Srinivasan, 2009).


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2.6 Turbulence and heat excited noise sources in single and coaxial jets

The generation of noise in subsonic high Reynolds number single and coaxial

turbulent jets is analyzed by a hybrid method. The computational approach is based on

large-eddy simulations (LES) and solutions of the acoustic perturbation equations (APE).

The method is used to investigate the acoustic fields of one isothermal single stream jet at

a Mach number 0.9 and a Reynolds number 400, 000 based on the nozzle diameter and two

coaxial jets whose Mach number and Reynolds number based on the secondary jet match

the values of the single jet. One coaxial jet configuration possesses a cold primary flow,

whereas the other configuration has a hot primary jet(Matthias Meinke, 2009).

2.7 Navier-Stokes analysis methods for turbulent jet flows with application to aircraft

exhaust nozzle.

This article presents the current status of computational fluid dynamics (CFD)

methods as applied to the simulation of turbulent jet flow fields issuing from aircraft

engine nozzles. For many years, Reynolds-averaged Navier-strokes (RANS) methods have

been used routinely to calculate such flows, including very complex nozzle configurations.

RANS method replaces all turbulent fluid dynamic effects with a turbulence model. Such

turbulent models have limitations jets with significance three- dimensionality,

compressibility and high temperature streams. In contrast to the RANS approach, direct

numerical simulation (DNS) methods calculate the entire turbulent energy spectrum by

resolving all turbulent motion down to the Kolmogorov scale. Although this avoids the

limitations associated with turbulence modeling. DNS methods will remain

computationally impractical in the foreseeable future for all but the simplest

configurations. Large-Eddy simulation (LES) methods, which directly calculate the large-

scale turbulent structures and reserve modeling only for the smallest scales, have been

pursued in recent years and may offer the best prospects for improving the fidelity of

turbulent jet flow simulations. A related approach is the group of hybrid RANS/LES

methods, where RANS is used to model the small-scale turbulence in wall boundary layers

and LES is utilized in regions dominated by the large-scale jet mixing. The advantages,

limitations, and applicability of each approach are discussed and recommendations for

further research are presented (Nicholas J.Georgiadis, James R. DeBonis 2007).


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2.8 Numerical predictions of noise in nozzle with and without chevrons

Numerical simulation of round, compressible, turbulent jets using the shear stress

transport (SS k-x) 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 case is 0.75. Overall sound pressure levels

(SPL) at far-field observer locations have been calculated using Ffowcs Williams-

Hawkings 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 centerline 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 clod and hot

jets.

2.9 Summary

A literature review of past and present understanding of jet noise and the

techniques used to reduce it, reveals the lack of experimental data in relation to tangential

air injection for the purpose of active jet noise control, and hence makes an ideal starting

point for the present study. Jet noise is characterized by two principal noise sources,

turbulent mixing noise shock associated noise. It is accepted that turbulent mixing noise

consists of two distinct noise components. The first is a product of rapidly dissipating fine-

scale turbulent structures in the immediate shear layer of the jet that are highly radioactive

in the sideline and upstream directions. The second is a result of the growth of fine-scale

disturbances into large-scale turbulent eddies that exist towards the end of the potential

core and propagate towards the downstream aft angles, dominating the mixing process and

far- field acoustic spectra. Mach wave radiation occurs in the mixing region of a jet as a

result of the supersonic convection of large-scale eddies, which for cold laboratory jets

travel at approximately 70% of the jet exit velocity. Screech noise is a high amplitude

discrete tone commonly found in imperfectly expanded jets and is a product of the

feedback loop induced by the turbulence-shockwave interaction around the third to fourth

shock cell in the potential core. It has fundamental and harmonic components both of


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which radiate in opposing directions and have been known to cause sonic fatigue failure.

Broadband shock noise, thought to be sourced from the same turbulence-shockwave

mechanism as screech noise, has a larger bandwidth and occurs at higher frequencies in

both near and far-field upstream acoustic spectra. Promoting enhanced levels of mixing

within the shear layer disrupts the foundations of all noise source mechanisms and is

shown to reduce jet noise. In supersonic jets the application of swirling component of

velocity has been proven to successfully reduce screech noise, whereas for subsonic flow

regimes, low levels of swirl have been found to be more efficient in reducing mixing noise

and turbulence levels. However, most studies have focused on passive methods with fixed

geometries such as guiding vanes and vortex chambers for producing swirl.


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3. COMPUTATIONAL WORKS

3.1 Modeling

Two-dimensional and three-dimensional chevron nozzles are designed by using

ANSYS-CFX software which is flexible one and user friendly. Co-flow analysis of

chevron nozzle is carried out with the help of FLUENT and CFX Post software. From the

figure 3.1 consist of baseline geometry with zero penetration and chevron count.

Fig 3.1 Isometric view of co-flow nozzle Geometry

Fig 3.2 Outer nozzle with base plate


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Fig 3.3 Base plate Fig 3.4 Inner nozzle

Fig 3.5 Inner nozzle with chevron

Fig 3.6 Position of tab in the nozzle


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3.1.1 Nozzle Specifications

Specifications Dimensions in mm

Inlet diameter of inner core 30.4

Outer diameter of inner core 11.6

Inlet diameter of outer cone 54.5

Outlet diameter of outer cone 17.85

Nozzle Length 35.5

Chevron length for eight count chevron 4.5

Chevron Penetration 0

Tab size Width - 1.5mm,

height - 1mm

Table3.1 Nozzle Specifications

3.2 Preprocessing

ICEM CFD is the meshing tool used to mesh the model in order to get the accurate result.

3.2.1 Three dimensional meshing

Initially open the ICEM CFD software. Then import the model in ICEM CFD using

import options. The import file should be in the form of iges file and step file. After

importing the model is clean up with the help of cleanup option to avoid the unwanted

edges. According to our requirements we can decompose the domain and mesh, for the

three-dimensional case we can go for tetra mixed yields good result. Then give the

boundary conditions, save and export the model.


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Fig 3.7 Mesh image.

3.3 Solver

ANSYS CFX is analyzing software used to simulate the model for this application.

3.3.1 Steps to solve

Initially have to open the CFX-Pre software. Then have to import our meshed

model into the CFX-Pre by read mesh file. After reading the case file have to check thegrid and zones. After checking grid and zone we have change the units as mm or cm or m

as what we want. Then select define, models, solver keep the defaults as it is. Then select

turbulence model as SST.

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 CFX-POST for post processing works.

3.4. Formula used to find the total pressure

P0/P = {1+ ((-1)/2) M2}/-1

P = 101325 Pa

For Mach no 0.6,

P0/P = {1+ ((1.4-1)/2) 0.62}1.4/ (1.4)-1

P0 = 129240.420 pa


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4. RESULTS AND DISCUSION

In this chapter we discuss about the final results obtained from the CFX PRE solver

and CFX post which are compared with the existing journals to make sure that we are onthe right path. The results are taken for co-flow jets with nozzle, chevron nozzle and

chevron nozzle with tab in two different locations of 31mm and 32mm.The base line

geometry for the nozzle are taken from the reference [1] Pinnam Lovaraju, E.

Rathakrishnan paper and has been tested under various Mach number.

Co-flow analysis of chevron nozzle with eight counts is done with the help of CFX

PRE software and CFX post, and the acoustic characteristics of co - flow nozzle and co-

flow nozzle with chevron and tab are measured.

Both results in comparison shows instead of co-flow nozzle, co-flow nozzle with

chevron and tabs yields minimum noise level, and it reduces 10db of noise level in average

at various Mach number 0.6, 0.8 and 1.0. Co-flow analysis of chevron nozzle with tab

yields the good results too. Using of Shear Stress Transport (SST) model in CFX we can

obtain the turbulence result which is compare with the existing results.

4.1 Centreline Velocity Decay:

The centerline velocity decay is an authentic measure of jet propagation, that is,

faster the decay, the faster is the jet mixing with the entrained fluid mass and so on. The

centerline velocity decay shows the extent of jet core clearly. In other words, it can be

stated that the core of a jet, either subsonic or supersonic, is the distance from nozzle exit

at which the characteristic decay begins.

4.2 Effect of tab on jet decay:

Two pairs of stream wise vortices were generated when tabs were introduced at the exit of

the nozzle. These vortices are small and have a long life span. Therefore they act as

effective mixing promoters and enhance mixing. This faster mixing result in rapid jet

decay of tabbed jets compared to a jet from a plain nozzle. According to past research

smaller the size of vortex, the better is the mixing because small size vortices travel longer

distance and are stable. The present investigation centered on tabs had explored the merits

of small vortex and effectiveness of the tab on mixing promotion. The perforations in the

tab seem to induce vortices of smaller size that travel right through the jet centerline. The


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vortices from the tab combined with the vortices from the circular perforations had

severely affected the potential core of the jet thereby increasing the entrainment of the

ambient into the jet core. The low pressure region behind the tab has facilitated greater

mixing of the jet with the ambient.

4.3 Mach number profiles along jet direction

Center line Mach number decay was measured for corresponding Mach numbers

0.6, 0.8 and 1.0. Faster the decay of Mach number, faster the jet mixing with the

surrounding fluid flowLovaraju.P and Rathakrishnan.E. [7]. The local Mach number (M)

is normalized with the jet Mach number (Me) and axial distance (X) is non-dimensional

with the inner nozzle exit diameter (D). Below figure gives the potential core for Mach

number 0.6, 0.8 and 1.0.

When there is no chevron or tabbed chevron potential core is extended up to

X/D=5.22 for Mach number 0.6. In the presence chevron and tabbed chevron X/D reduced

to 3.66 and 3.13. This results gives reduction in potential core implies faster the jet mixing

with nearby fluid. For 0.8 and 1.0 Mach number only baseline nozzle gives the potential

core of X/D 5.75 and 7.31 respectively. But in chevron and tabbed chevron for 0.8 Mach

potential core reduced to X/D= 4.18 and 3.9. For Mach number 1.0, baseline, chevron and

tabbed chevron nozzle potential core length X/D are 7.31, 5.22, and 2.94. These stream

wise vortices produced by the Chevron and Tabbed chevron causing reduction in jet

potential core and enhance mixing.


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FIG.4.1 BASELINE(M=0.6) FIG.4.2 BASELINE(M=0.8)

FIG.4.3 BASELINE(M=1.0) FIG.4.4 CHEVRON(M=0.6)

FIG.4.5 CHEVRON(M=0.8) FIG.4.6 CHEVRON(M=1.0)

FIG.4.7 TABBED CHEVRON(M=0.6) FIG.4.8 TABBED CHEVRON(M=0.8)

FIG.4.9 TABBED CHEVRON(M=1.0)


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FIG.4.10 BASELINE(NPR=3) FIG.4.11 BASELINE(NPR=5)

FIG.4.12 BASELINE(NPR=7) FIG.4.13 CHEVRON(NPR=3)

FIG.4.14 CHEVRON(NPR=5) FIG.4.15 CHEVRON(NPR=7)

FIG.4.16 TABBED CHEVRON(NPR=3) FIG.4.17 TABBED CHEVRON(NPR=5)

FIG.4.18 TABBED CHEVRON(NPR=7)


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GRAPH-4.1 SUPERIMPOSED GRAPH (M=0.6)



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GRAPH-4.4 SUPERIMPOSED GRAPH (NPR=3)


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GRAPH-4.7 BASELINE RADIAL DIRECTION (M=0.6)


0.00

0.20

0.40

0.60

0.80

1.00

1.20

0.00 0.50 1.00 1.50 2.00 2.50 3.00

M/Me

Y/D

0.6 MACH BASELINE RADIAL DIRECTION

X/D=0 X/D=2 X/D=5.5 X/D=8 X/D=12 X/D=15 X/D=18

0

0.2

0.4

0.6

0.8

1

1.2

0 0.5 1 1.5 2 2.5 3

M/Me

Y/D


X/D=0 X/D=2 X/D=5.5 X/D=8 X/D=12 X/D=15 X/D=18


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GRAPH-4.10 CHEVRON RADIAL DIRECTION (M=0.6)

0.00

0.20

0.40

0.60

0.80

1.00

1.20

0 0.5 1 1.5 2 2.5 3

M/Me

Y/D


X/D=0 X/D=2 X/D=5.5 X/D=8 X/D=12 X/D=15 X/D=18

0

0.2

0.4

0.6

0.8

1

1.2

0 0.5 1 1.5 2 2.5 3

M/Me

Y/D

0.6 MACH CHEVRON RADIAL DIRECTION

X/D=0 X/D=2 X/D=5.5 X/D=8 X/D=12 X/D=15 X/D=18


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0.00

0.20

0.40

0.60

0.80

1.00

1.20

0 0.5 1 1.5 2 2.5 3

M/Me

Y/D


X/D=0 X/D=2 X/D=5.5 X/D=8 X/D=12 X/D=15 X/D=18

0.00

0.20

0.40

0.60

0.80

1.00

1.20

0 0.5 1 1.5 2 2.5 3

M/Me

Y/D


X/D=0 X/D=2 X/D=5.5 X/D=8 X/D=12 X/D=15 X/D=18


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GRAPH-4.12 TABBED CHEVRON RADIAL DIRECTION (M=0.6)


0.00

0.20

0.40

0.60

0.80

1.00

1.20

0 0.5 1 1.5 2 2.5 3

M/Me

Y/D

0.6 MACH TABBED CHEVRON RADIAL IN Y/D

X/D=0 X/D=2 X/D=5.5 X/D=8 X/D=12 X/D=15 X/D=18

0.00

0.20

0.40

0.60

0.80

1.00

1.20

0 0.5 1 1.5 2 2.5 3

M/Me

Z/D

0.6 MACH TABBED CHEVRON RADIAL IN Z/D

X/D=0 X/D=2 X/D=5.5 X/D=8 X/D=12 X/D=15 X/D=18


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0.00

0.20

0.40

0.60

0.80

1.00

1.20

0 0.5 1 1.5 2 2.5 3

M/Me

Y/D


X/D=0 X/D=2 X/D=5.5 X/D=8 X/D=12 X/D=15 X/D=18

0.00

0.20

0.40

0.60

0.80

1.00

1.20

0 0.5 1 1.5 2 2.5 3

M/Me

Z/D


X/D=0 X/D=2 X/D=5.5 X/D=8 X/D=12 X/D=15 X/D=18


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0.00

0.20

0.40

0.60

0.80

1.00

1.20

0 0.5 1 1.5 2 2.5 3

M/Me

Y/D


X/D=0 X/D=2 X/D=5.5 X/D=8 X/D=12 X/D=15 X/D=18

0.00

0.20

0.40

0.60

0.80

1.00

1.20

0 0.5 1 1.5 2 2.5 3

M/Me

Z/D


X/D=0 X/D=2 X/D=5.5 X/D=8 X/D=12 X/D=15 X/D=18


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4.4 Radial Mach number profiles

Below figurepresents the radial Mach number variation of Mach 0.6, 0.8 and 1.0

jet at axial distances of X/D =0, 2.0, 5.5, 8.0, 12.0, 15.0 and 18.0, with baseline, chevron

and tabbed chevron. These plots clearly show the effect of chevron and tabbed chevron of

co-flow on central jet in the radial direction.. According to past research smaller the size

of vortex, the better is the mixing because small size vortices travel longer distance and are

stable. The present investigation centered on tabs had explored the merits of small vortex

and effectiveness of the tab on mixing promotion.Graphs show the radial distribution of

Mach number for baseline, chevron and tabbed chevron. The effect of co-flow is clearly

shown in below graph for 0.6, 0.8 and 1.0 Mach number. In this project tabs placed

perpendicular to the earths surface, for this reason only flow spreader perpendicular to

earths surface. The stream wise vortices introduced by chevron and tabbed chevron at the

primary core of the nozzle exit may be due the presence of pressure gradient upstream and

downstream of the chevron and tabbed chevron. Stream wise vortices strength is the strong

function of pressure gradient. While introduction of tabs off centered peak will attain at the

potential core. Here, off centered peak is achieved at Z/D= 0.7. This shows peak get shifted

away from the centerline. And this once again shows that the vortices move away from the

nozzle exit as they travel downstream.4.5 Nozzle pressure ratio

To understand the concept of under expanded cases, nozzle pressure ratio

calculations were done and super imposed graphs were plotted. These profiles clearly show

the effect of co-flow on the central jet characteristics at all axial locations. These effects of

co-flow shock interaction were studied in [2] Experimental Studies on Co-flowing

Subsonic and Sonic Pinnam Lovaraju, E. Rathakrishnan. While introduction of chevron

and tabbed chevron shock cell strength reduces when compared to baseline nozzle.

Following graphs shows comparison of nozzle pressure ratio at NPR=3, 5 and 7. Tab

weakens the shock cell structure drastically in the jet core. For NPR=3 is observed, weak

shock waves are clearly seen in baseline nozzle. While raising the nozzle pressure ratio to

5 and 7 in baseline nozzle strong shocks are captured. It is easy to see barrel shocks are

formed at NPR= 5and 7 in baseline nozzle. While introduction of chevron, shock cell gets

weak and it distracts one shock cell structure. In this paper new methodology tabbed

chevron reduces further one more shock cell. These will clearly shows introduction of

chevron and tabbed chevron promotes mixing.


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4.6 Sound pressure level

In order to find sound pressure level (dB), sound pressure level formulae is used for

baseline, chevron and tabbed chevron below table indicates the value of sound pressure

level at the nozzle exit.

SL

NUMBER

AT EXIT 0.6

MACH

(dB)

0.8

MACH

(dB)

1.0

MACH

(dB)

1 NOZZLE 169.84 175.92 181.86

2 CHEVRON 164.71 169.12 173.40

3 TABBED

CHEVRON

163.84 167.17 166.00

TABLE-4.1 SOUND PRESSURE LEVEL AT NOZZLE EXIT


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CONCLUSIONS

Vortex generators in the form of chevron and tabbed chevron have been found to

be quite effective in co-flow jet mixing. Tabbed chevron found effective in all the nozzle

pressure ratios (NPR).Chevron and tabbed chevron diffuse shocks. . At high NPR chevron

and tabbed chevron reduces shock strength and shock cell formation. Comparing all the

result in NPR Tabbed chevron is effective to diffuse shocks, resulting in weaker shock in

core region. Chevron and tabbed chevron is found to modify the characteristics of co-flow

central jet development Jet mixing is augmented by the chevron and tabbed chevron by the

formation of vortices. The length of the potential core of the subsonic and correctly

expanded sonic jets decreases in the presence of chevron and tabbed chevron results in

good mixing with the atmospheric flow in radial direction. Finally for 0.6 Mach number

introduction of chevron and tabbed chevron reduces 29.90% and 40.03% of potential core.

For 0.8 Mach number introduction of chevron and tabbed chevron reduces 27.30% and

32.17% of potential core. For 1.0 Mach number introduction of chevron and tabbed

chevron reduces 28.50% and 59.78% of potential core. It shows faster the centerline decay

faster the mixing with surrounding flow in the co-flowing jets. This chevron and tabbed

chevron is effective to modify mixing characteristics in co-flowing jets.


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6. REFERENCES

1. Pinnam Lovaraju E. Rathakrishnan Experimental Studies on Co-flowing

Subsonic and Sonic Jets. Flow, Turbulence and Combustion. An International

Journal published in association with ERCOFTAC.

2. E. Rathakrishnan Corrugated Tabs for Supersonic Jet Control(Keynote Paper)

3. McLaughlin, D. K., Morrison, G. D., and Troutt, T. R., Experiments on the

Instability Waves in a Supersonic Jet and Their Acoustic Radiation, Journal of

Fluid Mechanics, Vol. 69, Pt. 1, 1975, pp. 7395.

4. Troutt, T. R., and McLaughlin, D. K., Experiments on the Flow and Acoustic

Properties of a Moderate Reynolds Number Supersonic Jet, Journal of Fluid

Mechanics, Vol. 116, March 1982, pp. 123156.

5. K.B.M.Q. Zaman, J.E. Bridges and D.L. Huff Evolution from 'Tabs' to 'Chevron

TechnologyNASA Glenn Research Center - Cleveland, OH, USA.

6.

Malcolm Gibson The Chevron Nozzle: A Novel Approach to Reducing Jet Noise.

NASA Headquarters, Washington, DC

7. Effect of tabs on mixing characteristics of subsonic and supersonic jets. Author:

S.Thanigaiarasu, S.Elangovan, and E.Rathakrishnan.

8.

Lovaraju.P, Agarwal.V. and Rathakrishnan.E (2005),Mixing Enhancements of

Subsonic and Transonic Jets with Cross wire, 8th ASV paper 18

9.

Chiranjeevi Phanindra.Band Rathakrishnan.E (2010), Corrugated tabs for

supersonic jet control, AIAA Journal, Vol.48, No.2

10.

Clement.S and Rathakrishnan.E (2006), Characteristics of sonic jets with tabs,

Springer Link-Shock waves.


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11.Alkislar, M.B., Krothapalli, A., and Lourenco. L. M., Structure of a screeching

rectangular jet: a stereoscopic particle image velocimetry study,Journal of Fluid

Mechanics, Vol. 489, 2003

12.

Avital, E. J., Alonso, M., and Supontisky, V., Computational aeroacoustics: The

low speed jet,The Aeronautical Journal, Vol. 112, No.1133, July 2008

13.

Bishop, K. A., Fowcs-Williams, J. E. and Smith, W., On the noise sources of the

unsuppressed high-speed jet, Journal of Fluid Mechanics, Vol. 50, No. 1, 1971

14.

Callender, B., Gutmark, E., and Martens, S., A comprehensive study of fluidic

injection technology for jet noise reduction, AIAA 2007

15.Carpenter, P. W. and Johannesen, N. H., An extension of one -dimensional theory

to inviscid swirling flow through choked nozzles, Aeronautics Quarterly, Vol. 26,

1975

16.Tide, P.S., Srinivasan., Effect of chevron count and penetration on the acoustic

characteristics of chevron nozzles, Journal of applied acoustic 2009

17.

Nicholas, J., Georgiadis., James R.DeBonis., Navier-stokes analysis method forturbulent jet flows with application to aircraft exhaust nozzles, Journal of

ScienceDirect 2006

18.Lardeau, S., Collin, E., Analysis of jet-mixing layer interaction, Journal of

ScienceDirect 2003

19.Seong Ryong Kho, Wolfgang Schroder., Turbulence and heat excited noise

source in single and coaxial jets, Journal of Sound and Vibration 2008

20.Groschel, E., Schroder, W., Renze, P., Noise prediction for a turbulent jet using

different hybrid methods, Journal of Science Direct 2007

21.T.Ph. Bui, W. Schroder, M. Meinke., Numerical analysis of the acoustic field of

reacting flows via acoustic perturbation equations, Science Direct 2007.


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TEXT BOOK REFERENCES

1.

Grinstein F.F,Glauser.M&George.W.K (1995), Vorticity in jets, Chapter 3 of

FLUID VORTICES, Ed.S.Green, pp.65-94, Kluwer Academic Publishing.

2.

Rathakrishnan.E (2010), Applied Gas Dynamics, Wiley Publications ISBN 978-

0-470-82576.

WEBSITE REFERENCES

http://www.fluent.com

http://courses.cit.cornell.edu/fluent/

http://combust.hit.edu.cn:8080/fluent/Gambit13_help/gambit.http
http://www.fluent.com/http://courses.cit.cornell.edu/fluent/http://courses.cit.cornell.edu/fluent/http://www.fluent.com/

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