UNIVERSITY OF CINCINNATI
Date:______August 12, 2008
I, ______,Christopher A. Harris hereby submit this work as part of the requirements for the degree of:
Master of Science in:
Aerospace Engineering
It is entitled: Acoustics and Fluid Dynamics Studies of High Speed Jet Noise Reduction Devices
This work and its defense approved by:
Chair: ______Dr. Ephraim J. Gutmark ______Dr. Paul D. Orkwis ______Dr. Steven Martens ______Dr. John P. Wojno ______
Acoustics and Fluid Dynamics Studies of High Speed Jet
Noise Reduction Devices
A thesis submitted to the Graduate School Of the University of Cincinnati
In partial fulfillment of the Requirements for the degree of
Master of Science
In the Department of Aerospace Engineering and Engineering Mechanics Of the College of Engineering
2009
By
Christopher A. Harris
B.S., Boston University, 2002
Committee Chair: Dr. Ephraim J. Gutmark Abstract
Jet noise reduction was investigated on a scale model turbofan exhaust simulator rig at a Reynold’s Number O(106) through mean and time-resolved flow and aeroacoustic measurements. Various stream-wise vorticity production devices, including conventional and modified chevron nozzles and CVG’s (Coupled Vortex
Generators), were installed to increase turbulent shear layer mixing and ultimately reduce far-field radiated noise. Simplified flow simulations using a steady RANS k- epsilon turbulence model aid to elucidate the initial vortex development for several geometries. CVG’s were installed in axisymmetric arrangements on both the core and fan streams of the exhaust simulator, and in the various boundary layers.
Measurements of the nozzle boundary layer characteristics were performed using a total pressure probe on the baseline hardware to determine appropriate mean spatial scales, and to evaluate the boundary layer momentum thickness influence on noise for a coaxial, turbulent jet. LDV of two velocity components determined the turbulence properties in the jet at various locations in the initial mixing region and past the potential core. Acoustic far-field measurements showed that high levels of peak noise reduction were possible with added high-frequency energy. One purpose was to offer an explanation of this ‘self’ noise component and mixing mechanisms, in comparison with delta tabs which also incur high-frequency noise. With properly scaled geometry design, and installation configurations, the CVG’s can achieve SPL peak noise reductions for essentially all directivity angles, with the addition of a high-frequency source that appears consistent with a self-noise induced dipole.
2
Acknowledgements
I would like to thank those who supported this work, and the great amount of work that never will be represented throughout my time at UC, and the labs at GE. Dr. Ephraim Gutmark was particularly supportive, allowing freedom to investigate these issues as I saw fit, under his guidance. Also, he provided a world- class environment with which to pursue this research effort. Dr. John Wojno at GE Aviation was my technical research supervisor and lead to whom I owe many valuable open conversations about jet noise and acoustics, and from paying close attention I received a great deal in return. I appreciate his willingness to let me take as much of a role in the technology development process as I could handle. Dr. Steve Martens, now of GE GRC, I owe sincere thanks to for building a great collaboration with UC. Sharing his particularly insightful oversight of our methods and investigations, and offering the resources of many other experts in the GE acoustics group was invaluable in producing this data, and forming analyses. Dr. Richard Cedar, besides offering his own valuable experience and insight in engine technology development, brought the driving force to make these experiments as practical as possible while pursuing a basic level of fundamental understanding. I am glad to have been exposed to his perspective. To Russ Dimicco I owe a great deal for oversight of the facility and equipment, and his tenacious distaste for substandard anything. The quality of data, and the resources to get things done was by his doing. To my colleagues Olaf Rask, Dave Munday, and Seth Harrison, I owe many sets of data, completed only with your help. Olaf provided not only the initial mentorship that helped me advance my knowledge so quickly, but was a useful backboard for discussing ideas in fluids and acoustics. I owe a good amount of my success to Olaf. But thanks also to Boo and If Then Elsa for all their data collection and processing help, both exceptional undergraduate research assistants.
i Table of Contents
List of Figures iv List of Tables x Nomenclature xi
Chapter 1 Introduction 1 1.1 Jet Noise Background 3 1.2 Motivation for Source Noise Reduction 6 1.2.1 Psychoacoustics and Aircraft Noise Effects 7 1.2.1.1 Psychoacoustics of Interest in Aircraft Noise 8 1.2.1.2 Health Effects (Non-Auditory) of Aircraft Noise 11 1.2.2 U.S. Commercial Aircraft Noise Regulations 12 1.3 Objectives 15 1.4 Thesis Outline 16
Chapter 2 Background 20 2.1 Qualitative Jet Description 21 2.1.1 Mean Subsonic Coaxial Velocity Field & Similarity 23 2.1.2 Supersonic Jet Structure 26 2.2 Aeroacoustic Theory 28 2.3 Jet Instabilities 34 2.3.1 Effects of Initial Condition on Mixing Layer and Acoustics 35 2.3.2 Influence of Jet Instability Modes on Flow and Acoustics 38 2.4 Mixing Nozzle Research 40 2.4.1 Tabbed and Chevron Mixing Nozzles 41 2.4.2 Near-Field Fluid Structures of Vortex Generators 48
Chapter 3 Experimental Facilities & Methods 52 3.1 GDPL Anechoic Aeroacoustic Test Facility 53 3.2 ATF High Bypass Model Rig & Hardware 54 3.2.1 Conventional Core and Fan Nozzles 56 3.2.2 Modified Conventional Core Nozzles 57 3.2.3 Vortex Generators on Core and Fan Nozzles 58 3.3 Acoustic and Fluid Measurement Systems 60 3.3.1 Acoustic Far-Field Microphone Array 60 3.3.2 Boundary Layer Total Pressure Pitot Measurements 62 3.3.3 Dual-Component Laser Doppler Velocimetry 63 3.3.3.1 Flow Tracker Particle Seeding 66
Chapter 4 Boundary Layer Flow and Noise Results 67
ii 4.1 Core Baseline and Chevron Nozzles Mean Initial Velocity Profiles - Clean and Tripped 68 4.2 Acoustic Far-Field Sensitivity to Boundary Layer Characteristics 74
Chapter 5 Initial Mixing Region Turbulence via LDV 80 5.1 Mean Results 81 5.2 Turbulence Statistics in Axial and Radial Velocities 86 5.3 Turbulence Velocity Spectra 92
Chapter 6 Acoustic Far-Field Noise Reduction Results 95 6.1 Baseline and Modified Chevron Root Geometry 96 6.2 Internal Chevron Core Nozzle Results 101
Chapter 7 Incompressible CFD Simulations 105 7.1 Computational Domain and Case Setup 106 7.2 Flowfield Scalar Results 109 7.3 Stream-wise Vortical Structures Development 115
Chapter 8 Conclusions 126
Bibliography 128
iii List of Figures
FIGURE 1.1 – CONTOURS OF EQUAL PERCEIVED NOISINESS AT VARIOUS NOYS LEVELS
[10] ...... 9
FIGURE 1.2 - SYSTEM NOISE COMPONENT ESTIMATES, [] ...... 13
FIGURE 1.3 - THRUST NORMALIZED AIRCRAFT NOISE LEVEL HISTORY [11] ...... 14
FIGURE 2.1 - SHADOWGRAPH FLOWFIELD OF HEATED COAXIAL MODEL JET AT CRUISE
CONDITIONS. [34] ...... 21
FIGURE 2.3 - SCHEMATIC OF SHOCK-CELL FLOW STRUCTURE DEVELOPMENT FOR
SIMPLIFIED SINGLE AXISYMMETRIC JET WITHOUT CENTERBODY PLUG [44]...... 26
FIGURE 2.4 - SUPERSONIC MIXING LAYER DETAILED SHOCK TIP SCHEMATIC [34] ...... 27
FIGURE 2.5 - SCHLIEREN NANOSPARK IMAGING OF UNDEREXPANDED CONICAL
NOZZLE CORE FLOW IN STATIC CONDITIONS AT THE UC AEROACOUSTIC TEST
FACILITY [48] ...... 27
FIGURE 2.6 - SOURCE CONVECTION EFFECT ON DIRECTIONAL RADIATION OF JET
NOISE, FOR VARIOUS JET MACH NUMBERS [52] ...... 31
FIGURE 2.7 - EARLY "TABBED" MIXER NOZZLE OF WESTLEY & LILLEY, 1952 [72]...... 41
FIGURE 3.1 - A) SCHEMATIC OF AEROACOUSTIC TEST FACILITY FAR-FIELD
MICROPHONE ARRAY, B) NOZZLE ACOUSTIC TEST RIG WITH CHEVRON NOZZLE
INSTALLED DURING LDV BACKSCATTER MEASUREMENTS ...... 53
FIGURE 3.2 - CUTAWAY VIEW & COMPONENTS OF HIGH BYPASS RATIO (8) NOZZLE
ACOUSTIC TEST RIG...... 55
FIGURE 3.3 - CORE AND FAN STREAM CONVENTIONAL NOZZLES INCLUDED IN
TESTING CONFIGURATIONS. A) BASELINE CONIC CORE NOZZLE, B) 8LP CORE
NOZZLE, C) 8LP SIN CORE NOZZLE, D) 16HP FAN NOZZLE ...... 56
FIGURE 3.4 - MODIFIED 8HP CORE CHEVRON NOZZLE WITH SINUSOIDAL ROOT A)
NOZZLE B) DIMENSIONS...... 57
iv FIGURE 3.5 - CVG GEOMETRY DEFINITIONS AND MODEL RIG INSTALLATION. A)
DELTA CVG (DVG), B) MUSHROOM CVG (MVG), C) CVG NOZZLE SECTION (NONE
INSTALLED), D) SEAT FOR CVG’S ...... 58
FIGURE 3.6 - INTERNAL CHEVRON CORE NOZZLE. A) INTERNAL SURFACE, B)
EXTERNAL SURFACE, B) COVERED ...... 59
FIGURE 3.7 - BOUNDARY LAYER TOTAL PRESSURE PROBE, PHYSICAL DIMENSIONS: A)
12" B) 11" D) 0.25" F) 0.650" M) 0.120" ...... 62
FIGURE 3.8 - SCREENSHOT OF LDV 3D TRAVERSE AND TRIGGER SYNC PROGRAM
INTERFACE ...... 65
FIGURE 3.9 – A) LASKIN SEED UNITS, B) OIL PARTICLE DIAMETER DATA, C) POLAR
LOG MIE INTENSITY PLOT ...... 66
FIGURE 4.1 – 3D SLICE VIEW OF BOUNDARY LAYER TRIP INSTALLATIONS ON BOTH
STREAMS...... 67
FIGURE 4.2 - AXIAL VELOCITY PROFILES ALONG RADIAL TRAVERSE FOR BASELINE
AND HIGH TRIP CONFIGURATIONS. SINGLE JET, COLD FLOW...... 69
FIGURE 4.3 – REPEATED PITOT MEASUREMENTS OF BASELINE NOZZLE BOUNDARY
LAYER FOR RAW DATA...... 69
FIGURE 4.4 - FAN VELOCITY PROFILE AT CORE EXIT PLANE FOR BASELINE CORE
NOZZLE WITHOUT (CLEAN) AND WITH (IMOM, IHOM) TRIPS ...... 70
FIGURE 4.5 – NORMALIZED MOMENTUM DEFICIT, CORE FLOW, BASELINE NOZZLE. ... 70
FIGURE 4.6 - RELATIVE CHANGE IN Θ FOR ALL NOZZLES AND TRIPS. A) INNER
BOUNDARY LAYER, B) OUTER ...... 73
FIGURE 4.7 – T.O.B. FAR-FIELD SPL SPECTRA, CLEAN CONFIGURATIONS. A) 70°, B) 150°
...... 74
FIGURE 4.8 – NARROWBAND FAR-FIELD SPL SPECTRA, CLEAN CONFIGURATIONS. A)
70°, B) 150°...... 74
v FIGURE 4.9 - NARROWBAND FAR-FIELD SPL SPECTRA OF CLEAN, INNER BOUNDARY
LAYER MEDIUM TRIP (IM), AND HIGH TRIP (IH) BASELINE NOZZLE. A) 70 º, B) 150º
...... 76
FIGURE 4.10 - NARROWBAND FAR-FIELD SPL SPECTRA, BASE, SINE IH, AND CHEV IH.
A) 70 º, B) 150º...... 76
FIGURE 4.11 - FAR-FIELD SPECTRA FROM 70º TO 150º FOR IMOM TRIP NOZZLES...... 77
FIGURE 4.12 - FAR-FIELD SPECTRA FROM 70º TO 150º FOR IHOH TRIP NOZZLES...... 79
FIGURE 5.1 - LDV PROBE TRAVERSE LOCATIONS...... 80
FIGURE 5.2 – CENTERLINE U/U1 VELOCITY FOR BASELINE, CHEVRON AND 12DVG
NOZZLES...... 81
FIGURE 5.3 - CENTERLINE REDUCTION IN U/U1 WITH BASELINE NOZZLE DATUM FOR
MIXER NOZZLES. C1, C1, AND V1, V2, AND V3 INDICATE REGIONS OF DISTINCT
MIXING BEHAVIOR FOR CHEVRONS AND VG’S ...... 82
FIGURE 5.4 - CENTERLINE V/U1 VELOCITY FOR ALL NOZZLES. AVERAGE VALUE OF
1.5% (CONIC NOZZLE)...... 84
FIGURE 5.5 – U/U1 VELOCITY PROFILES ALONG RADIAL TRAVERSES AT Z/D1 = 0.88 AND
1.46...... 85
FIGURE 5.6 – HISTOGRAMS OF U AT VARIOUS Z/DEQ LOCATIONS ON JET CENTERLINE
OF CONIC NOZZLE...... 86
FIGURE 5.7 – CORE NOZZLE PLUG BOUNDARY LAYER SEEDING IMAGE ...... 87
FIGURE 5.8 - HISTOGRAMS OF U AT VARIOUS Z/DEQ ON CENTERLINE FOR CHEVRON
AND 12DVG NOZZLES...... 88
FIGURE 5.9 - HISTOGRAMS OF U AT VARIOUS Y/D1 ON Z/DEQ=1.46 FOR BASELINE CONIC
AND 8HP NOZZLES...... 90
FIGURE 5.10 – AXIAL TURBULENCE INTENSITY ALONG JET CENTERLINE FOR ALL
NOZZLES...... 91
FIGURE 5.11 - TRANSVERSE TURBULENCE INTENSITY ALONG JET CENTERLINE FOR
ALL NOZZLES ...... 91
vi FIGURE 5.12 – BACKSCATTER LDV SAMPLE RATES ALONG JET CENTERLINE FOR ALL
NOZZLES...... 92
FIGURE 5.13 - SPECTRAL CENTERLINE RESULTS FROM LDV U VELOCITY COMPONENT
SFT, ALL NOZZLES...... 93
FIGURE 6.1 - 8HP CHEVRON UNMODIFIED AND SINUSOIDAL ROOT TOB SPECTRA AT 90˚
AND 150˚ ...... 96
FIGURE 6.2 – TOB SPECTRAL DELTA AT 90˚ FOR 8HP AND 8HPSIN NOZZLES, THE
ARROWS INDICATING REDUCED SOURCE TRANSITION RANGE...... 97
FIGURE 6.3 - HYBRID 8HP AND CVG CORE NOZZLE TOB SPECTRA AT 90° AND 150° ..... 99
FIGURE 6.4 - HYBRID 8HP AND CVG NOZZLE OASPL DIRECTIVITY FOR SINGLE AND
COAX-HI SETPOINTS...... 100
FIGURE 6.5 - IC CORE NOZZLE TOB SPECTRA FOR SINGLE AND COAX-HI SETPOINTS AT
70˚ AND 150˚, COMPARING NOMINALLY FABRICATED NOZZLE AND THE ADDITION
OF A SIMPLE STEEL SHEET COVERING...... 101
FIGURE 6.6 - IC CORE NOZZLE TOB SPECTRA FOR SINGLE SETPOINT COMPARING
PENETRATION EFFECT AT 90°. A) NOMINAL AND 5 PENETRATION (VALUE X 0.01
IN.), B) 5, 10, 15, 20 PENETRATION VALUES ...... 102
FIGURE 6.7 - IC CORE NOZZLE TOB SPECTRA FOR SINGLE SETPOINT COMPARING
PENETRATION EFFECT AT 150°. A) NOMINAL AND 5 PENETRATION (VALUE X 0.01
IN.), B) 5, 10, 15, 20 PENETRATION VALUES ...... 103
FIGURE 6.8 - IC CORE NOZZLE DIRECTIVITY COMPARISON WITH 8HP CHEVRON...... 104
FIGURE 7.1 – CFD GEOMETRIES. A) SOLID VG (SVG), B) CVG (NOT SHOWN), C) DELTA
TAB (MM)...... 105
FIGURE 7.2 – CFD SOLUTION DOMAIN...... 106
FIGURE 7.3 - BOUNDARY LAYER PROFILES AT SEVERAL VALUES OF Z: 0 (INLET), 10,
20, 30, AND 40MM, FOR BOTH A) SVG AND B) CVG GEOMETRIES. DASHED LINE
INDICATES THE UPPERMOST VERTEX OF THE VG’S...... 107
FIGURE 7.4 - HYBRID NEAR-WALL MESH DETAIL FOR SVG GEOMETRY...... 108
vii FIGURE 7.5 - ENTIRE DOMAIN X/H=0 VELOCITY FEATURES OF SVG GEOMETRY.
ISOSURFACE AT U0=0.20...... 109
FIGURE 7.6 - ENTIRE DOMAIN X/H=0 VELOCITY FEATURES OF CVG GEOMETRY.
ISOSURFACE AT U0=0.20...... 110
FIGURE 7.7 – A,B) STATIC CP CONTOURS FOR SVG, C) MEASURED CP FOR A JET
TRIANGLE TAB, M=0.3 [82]...... 110
FIGURE 7.8 - ENTIRE DOMAIN X/H=0 VELOCITY FEATURES OF DELTA TAB.
ISOSURFACE AT U0=0.20...... 111
FIGURE 7.9 –CP CONTOURS AT Y/H=0 (WALL). A) SVG, B) CVG (INEXACT GEOMETRY),
C) DELTA TAB ...... 112
FIGURE 7.10 - CP ISOSURFACES AT +0.25, -0.60, & STREAMLINES, FOR A) SVG, B) CVG,
C) DELTA TAB ...... 113
FIGURE 7.11 - VG SURFACE PRESSURE LEVELS. A) SVG, B) CVG (NOT SHOWN), C)
DELTA TAB...... 114
FIGURE 7.12 – PARTIAL DOMAIN GROSS VORTICITY BEHAVIOR, OVER X/H=0, Y/H=0.5
PLANES. A) SVG, B) CVG, C) DELTA TAB ...... 116
FIGURE 7.13 - CFD CROSS-PLANE Z/H LOCATIONS RELATIVE TO ENTIRE DOMAIN. ... 117
FIGURE 7.14 - NORMALIZED STREAM-WISE VORTICITY CROSS-PLANES AT VARIOUS
Z/H STATIONS, EQUALLY DIVIDED AT 25% LVG, INTO FIVE NOMINALLY LOCATED
POSITIONS. A) SVG, B) CVG, C) DELTA TAB ...... 118
FIGURE 7.15 - NORMALIZED STREAM-WISE VORTICITY CROSS-PLANES AT VARIOUS
Z/H DISPLACEMENTS RELATIVE TO TRAILING EDGE AS INDICATED. A) SVG, B)
CVG, C) DELTA TAB (SEE FIGURE 7.15 SCALES)...... 121
FIGURE 7.16 - NORMALIZED VORTICITY MAGNITUDE CROSS-PLANES AT VARIOUS Z/H
DISPLACEMENTS RELATIVE TO TRAILING EDGE AS INDICATED. A) SVG, B) CVG,
C) DELTA TAB ...... 123
viii FIGURE 7.17 - NORMALIZED AXIAL VELOCITY CROSS-PLANES AT VARIOUS Z/H
DISPLACEMENTS (NOMINALLY FOR UPSTREAM) RELATIVE TO THE TRAILING
EDGE (TE) AS INDICATED. A) SVG, B) CVG, C ) DELTA TAB ...... 124
FIGURE 7.18 - TKE CONTOURS AT 2H DOWNSTREAM OF TRAILING EDGES. A) SVG, B)
CVG, C) DELTA TAB...... 125
ix List of Tables
TABLE 3.1 - NATR EXIT PLANE CIRCULAR GEOMETRY FOR BYPASS RATIO 8 BASELINE
HARDWARE ...... 55
TABLE 3.2 - BACKSCATTER DUAL-COMPONENT LDV MEASUREMENT VOLUME
GEOMETRICAL PARAMETERS ...... 64
TABLE 4.1 – DESIGNATIONS, DIMENSIONS, AND SCALING PARAMETERS OF CORE AND
FAN TRIPS...... 68
TABLE 4.2 - INNER AND OUTER BASELINE CONIC CORE NOZZLE EXIT BOUNDARY
LAYER CHARACTERISTICS ...... 71
TABLE 4.3 - INNER SHEAR LAYER CHARACTERISTICS IN TRIPPED CONFIGURATIONS
AND CLEAN REFERENCE ...... 72
TABLE 4.4 - OUTER SHEAR LAYER CHARACTERISTICS IN TRIPPED CONFIGURATIONS
AND CLEAN REFERENCE ...... 72
TABLE 6.1 - SIMULATED CYCLE CORE AND FAN STREAM CONDITIONS AND COMBINED
ATTRIBUTES...... 95
x Nomenclature
x) streamwise coordinate r radial unit vector r radial coordinate δ f fan/ambient shear layer radial dimension δ c core/fan shear layer radial dimension δ w centerbody wake radial dimension δ boundary layer dimension δ 1 boundary layer displacement dimension Re stream Reynold’s number
Reθ momentum thickness based Reynold’s number U freestream velocity
U i velocity component, i = 1 to 3
U p velocity of primary (core) stream
U s velocity of secondary (fan) stream Vmix massflow-averaged mixed velocity ρ density
ui fluctuating component of velocity, i = 1 to 3
pij stress tensor, i = 1 to 3, j = 1 to 3 p thermodynamic pressure δ ij kronecker delta µ dynamic viscosity
eij shear stress tensor, i = 1 to 3, j = 1 to 3 a local speed of sound
a0 ambient far-field speed of sound
Tij fluctuating Reynold’s stress tensor, i = 1 to 3, j = 1 to 3 t time P acoustic power ρ 0 ambient far-field density l characteristic turbulent eddy dimension A jet exit area v e y source position vector v x observer position vector ρ' density fluctuations m eddy mach number
xi M c convection mach number M stream mach number θ momentum thickness, jet axis directivity angle α mach-wave normal effective Strouhal frequency vr mach-wave normal fluctuating velocity component RMS
Deq equivalent diameter
D1 core stream diameter AR aspect ratio δ δ 2 x , z effective 1/e optical probe width/height δ t trip diameter H shape factor, height of VG normal to surface St Strouhal number ω vorticity CD Nozzle Coefficient D jet diameter (single jet) TKE Turbulent Kinetic Energy PNL Perceived Noise Level, dB EPNL Effective Perceived Noise Level, dB CPR Core Pressure Ratio FPR Fan Pressure Ratio VG Vortex Generator CVG Coupled Vortex Generator DVG Delta Vortex Generator MVG Mushroom Vortex Generator SVG Solid Vortex Generator LDV Laser Doppler Velocimetry PIV Particle Imaging Velocimetry LES Large Eddy Simulation NATR Nozzle Acoustic Test Rig ATF Acoustic Test Facility
xii Chapter 1 Introduction
A brief account of the historical work on jet noise research is presented through
Section 1.1. Those familiar with or uninterested in these aspects should skip directly to Section 1.3 and the following section, which provide the objectives of this thesis and an outline of each chapter. This is included because much of what is currently being attempted for noise reduction has been historically attempted, in some cruder fashion possibly, and it seems to have in some respects come “full circle” with much having been attempted, much still to be understood in terms of the exact mechanisms of jet noise reduction.
Jet propulsion development wrought a spectrum of new environmental issues, not least of which was the high decibel sound field generated during operation.
Industry and governmental research groups in the United States, and at universities overseas such as Cranfield and Manchester, were the first to investigate rigorously the fundamental and practical aspects of aeroacoustics, as even then was aeroacoustics viewed as a stern challenge and an escalating one. In particular, spurred by the compelling notion of Britain’s assistant director of scientific research, in 1949, James Lighthill began to develop what was to become an exact theory of aerodynamically generated noise [20], based on only information inherent in the flowfield itself. Only the day following the introduction of this new challenge wrought by a classical turbulent flow, did James Lighthill sketch out the
1 main thrust of the new theory on the back of an envelope on the London train. The
“acoustic analogy” as it would be deemed in his 1952 paper, which also included a first dimensional analysis, is the seminal theoretical work in aeroacoustics, and even today still finds use with more recent modifications and extensions.
Prior to this new theoretical development, though, research had begun characterizing the noise from turbulent flows, and had achieved some preliminary steps in understanding the nature of flow generated noise. The era of the jet engine was also rapidly bearing upon increasingly large areas in terms of noise annoyance and safety, and in testing new, more powerful engines did environmental noise issues become critical. New test facilities were needed which required better acoustic attenuation and general noise control practices. Airframe attenuation practices required improvement to inhibit new levels of noise in the cabin of higher speed vehicles which was allowed by even larger jet power-plants. Eventually, the noise problem was correctly characterized as an intense research challenge that would continually develop in scope and in complexity due to the growing power of the jet engine. A paper, titled “Unsolved Military Noise Problems,” published late
1952, appositely characterizes a transition that occurred in vehicle acoustics engineering, from a focus on noise attenuation and control methods and materials, to reducing the noise emitted in the first place at the source, despite the drastically understated title.
2 1.1 Jet Noise Background
By the 1950’s the research community was beginning to grapple with jet noise, and refine the experimental methods needed to correctly characterize the noise fields and overall sound power levels produced from high thrust engines. In
1948 the P80 jet powered aircraft was the subject of one of the first directivity studies [5]. Hubbard and Lassiter in 1952 [28] produced a thorough assessment of subsonic jet noise. Their NACA study focused on the properties of a model blow- down jet flow and the nozzle geometry on the free-field radiated noise and comparison to measured turbojet engine noise. Several important aspects, clarified only later through both theoretical and experimental analysis, were nonetheless clearly identified in this work. Of note is the strong dependence of noise level on observer position, and the observance of higher low-frequency levels of noise nearer the jet axis. Also of fundamental importance was the observed dependence of SPL increase on velocity to a high power. This work was broad in the sense that also turbulence intensity, density, nozzle area, and observer distance were investigated, with an extension of the first comparison in literature of model and engine data.
Their conclusion agrees with the current belief that model scale data can well represent actual turbine engine exhaust noise. The work was also so far unique in identifying the location and character of the various source regions of the jet, with the exception of work at Cranfield by Westley and Lilley [72]. Now well known by polar correlation and other source-location techniques, the maximum noise emission region of a few diameters downstream of the nozzle exit was located approximately.
The frequency variation of the sources along the jet, pointing toward a region of
3 higher frequency noise generation near the nozzle exit, and lower frequency energy being produced many diameters downstream of the exit at greater levels. These observations correspond to extensively documented features of jets, based on the turbulence structure along the jet and instability waves that radiate noise dominated in the low frequency spectrum.
Overall, the state of experimental capabilities was improving, and honing in on the important characteristics of jet noise and other sources of noise to mostly increase the understanding of the phenomena, but also to methods to improve noise reduction. This can also be evidenced at this point in estimates of the acoustic efficiency, and even documenting the increase in efficiency with increased engine setting, however rough the pressure integration. It is interesting to note the sophistication of testing at this point. The NACA study used wooden nozzles mounted to a non-regulated air supply, which effected a velocity decrease throughout the test window. No corrections were made for atmospheric attenuation, even with nozzle diameters as small as ¾”, and ground and facility reflection was assumed negligible, which probably was a reasonable assumption for free-field measurement distances. The SPL was analyzed in six frequency bands. For the time, this setup was acceptable, and actually in light of the state of knowledge on jet noise in particular, this provided enough clarity in the measurements to provide useful, new information. Still, the experiments varied from test to test, and place to place, with highly different methods and nothing to standardize expectations, thus
Lighthill’s theory of aerodynamically generated noise became instantly extremely useful.
4 Before a summary of some of the major impetus behind jet and aircraft noise reduction research, and then developing the theoretical and experimental grounds of modern aeroacoustics, it is both insightful and interesting to discuss the approaches taken in the nascent era of jet engines toward noise as a system problem. Of course, noise is not the problem of any one group, being associated with the engine manufacturer first, who requires testing the as-built engine without worker safety concerns or community noise issues. The airframe manufacturer must provide an installed engine configuration that provides low fatigue due to noise-structure coupling, and also cater to low cabin noise and cockpit noise level demands. The airport and airlines both work with engines directly, and make on-wing adjustments, and so have similar requirements as to the engine manufacturer, but also deal with the main problem of community noise penalties during operation. The difficulty in finding the proper balance of engineering requirements, comfort and safety is highly subjective, since noise must be perceived by individuals, whom can all provide very differing responses. This was a confounding problem during the inception of the civil jet powered craft, since both the reach of the noise swaths being carved through communities was ever-expanding, and noise exposure annoyance and health effects, among other effects, was poorly understood. One certainty, that people began reacting adversely with great intensity, was well understood and communicated to governments. Therefore, it was correctly characterized early on that this was a system problem of everyone’s great interest. Accordingly, the best ways seen to tackle the system involved 1) optimizing operational aspects including flight patterns, schedules, levels, and locations, 2) planning future growth in terms
5 of aircraft anad airport requirements, respective of current technologies, economics and public polity, and 3) attempting direct scientific investigation of the noise sources, which reduces every negative aspect and may alleviate the restrictions by the other solution system components.
1.2 Motivation for Source Noise Reduction
The major factors that govern the extensive support of current research efforts in aeroacoustics are of such ubiquitous significance, that a brief outline is warranted.
As mentioned previously, early motivations for aerodynamic noise research were a necessity arising from new jet engine testing demands, which created high intensity noise environments for industrial workers and test engineers. It was soon realized that these problems were a precursor to a much larger civil aviation noise problem.
The effects of noise from aircraft soon became a strong annoyance among people subjected to the phenomenon indirectly, or at least to say without strict requirement.
During the 1968 meeting of the ICAO, a resolution was even adopted declaring aircraft noise a “great concern, and requires an urgent solution.” These concerns were based on the current scope of the aircraft noise problems, which during the
1950’s and beyond have grown into far greater impacts upon great numbers of people in terms of noise pollution annoyance, health effects, associated mental conditions, real estate devaluation, and passenger comfort during flight, not to mention the physical effects on aerostructures due to large amplitude acoustic fields.
Notable work has documented these ill effects, and provided future outlooks that consider phasing out older aircraft, increasing air traffic, population changes near
6 airports, expansion capabilities of existing airports, and an apparently increasing sensitivity to aircraft noise. In the brief analysis that follows, two distint motivations are discussed. First, the effects of jet noise on the exterior environment, that produces annoyance, health effects, and other subsequent and ancillary effects, will be outlined. Secondly, the motivation to combat the source of the cause of these environmental effects through aeroacoustics research and development will be exposed, as perhaps the most critical component of a system problem that includes many partners, and techniques of alleviating the most drastic effects.
1.2.1 Psychoacoustics and Aircraft Noise Effects
No serious debate exists as to whether jet noise is annoying. People everywhere that aircraft noise pollution is present react adversely, with few exceptions. In fact, noise is the driver behind legislation and regulation, both local and national, limiting and restriction aircraft operations, and recently by imposing direct monetary costs dependent on the noise generated by aircraft near airports. This has culminated in the national and international scope in the United Stated as FAR
(Federal Aviation Regulation) Part 36, an internationally as ICAO Annex 16 regulations. Numerous local, airport-specific regulations exist based on the local population density, land use, and governing authority. Prior to undertaking a discussion of FAR Part 36 regulations, the roots of aircraft noise as an annoyance, as an environmental health concern, and as the source of many high-cost indirect
7 effects, the nature of the noise generated by an aircraft and it’s perception by humans is briefly discussed.
1.2.1.1 Psychoacoustics of Interest in Aircraft Noise
Acoustic signals generated by aircraft travel a long path before being perceived as sound by a ground-based observer. First, broadband noise generated in the turbulent jet and by the airframe sources such as slats and slat joints, travels through the atmosphere in relative movement, to the observer. Through the atmosphere, the noise will be refracted through temperature gradients, and exchange energy with air in the form of “classical” frictional losses, as well as through the vibrational relaxation of nitrogen and oxygen. The signal then reaches the ear of the listener and is converted, by a head-related transfer function, into another signal that is then perceived. This process is open to severe subjectivity, but can be standardized and averaged for large groups of people, and converged to a set of reliable “final” data.
Besides pure annoyance, exposure to noise levels of 85-90 dBA over long periods of time lead to progressive hearing loss. It is important to distinguish, though, between the effects of jet noise, of which annoyance is direct, purely auditory effect. Other effects of jet noise include indirect effects, which may or may not be strongly dependent on actual annoyance or hearing impairment, such as medical effects, mental effects, land devaluation, land use restrictions, restricted airport operations, and large numbers of litigations and resources spent on attempting to reconcile the interests of the noise producers and those affected.
8 Psychoacoustics deals purely with a person’s perception of auditory signals, and the relationship that exists between the perceived sounds, in this case broadband noise, and the absolute metrics. Included in the set of purely psychological variables are sones or phons (loudness), mels (melody), vols (volume), and dasy (density) and noys (annoyance). This allows for using measurable acoustic data to compute perceived sound levels.
The relationship between the OASPL of a band of white noise and perception of the signal by the listener is described by a family of curves called noys. These curves reveal the average perceived loudness of a certain frequency of sound, where the actual or delivered sound pressure level is modified to maintain a constant level of perceived noisiness. The resulting curves provide a person’s perceived loudness for any given center frequency, and at many different perceived loudness levels, another key independent variable. Figure 1.1 indicates a high sensitivity of human hearing near 3150 Hz, and decreasing sensitivity above about 10,000 Hz, and below about 100 Hz. Perceived noise,
PNdB, in relation to the noys curves, is Figure 1.1 – contours of equal perceived the inverse, where the peak PNdB level noisiness at various noys levels [10]
9 corresponds to the trough in the level of noys, indicating human hearing response sensitivity rather than attenuation.
Aside from the OASPL of a signal, the duration, rise time, and tonal content play a role in determining the perceived annoyance level. Noise annoyance and the negative perception of certain types of environmental noise are strongly dependent on the characteristics of the acoustic signal. For example, habituation occurs more frequently for continuous stimuli, thus rendering jet noise more generally negatively perceived for equal exposure durations compared to relatively continuous noise sources, such as automobile traffic. This seems to work in concert with an annoyance or stimuli modification factor that is dependent on a person’s inability to control the surrounding environment, which increases negative perception of the stimulus and is not directly related to the acoustic properties of the noise [12].
In light of the median human response to noise as described, it is clear that jet noise reduction technology may be effective in several modes. One of these modes includes shifting a given band of noise at a more sensitive frequency to a much less sensitive frequency. Another is to reduce the tonal noise components associated with jet noise, which is generally only applicable to supersonic pressure ratios high enough to incur shock associated noise. Further, a combination of proven technologies, such as chevron mixer nozzles, that reduce the total acoustic power output with very small increases in some other components, may be used in conjunction with these other methods effectively to create additional reductions in the perceived noise.
10 1.2.1.2 Health Effects (Non-Auditory) of Aircraft Noise
The mechanisms of transformation of purely auditory stimuli to recognizable health effects of aircraft noise exposure are still unclear, with limitations on sample size, study lengths, and inconsistencies preventing concrete conclusions [12,14].
Some determinations have nonetheless been made. Noise annoyance has been shown to cause hypertension in adults, increased finger pulse amplitude, elevated heart rate, and sleep disturbance which leads to reduced next-day reaction time.
Speech interference is another commonly reported effect, which can lead to increased annoyance and apparently then to further aggravated indirect health effects. There also seems to be some evidence to increase self-reporting of health problems in noise exposed areas, which may be leading to an increase in noise complaints and overall assessed societal impact even though noise impact based on dBA contours surrounding airports have leveled off or decreased since the 1970’s
[15].
Some studies have linked increased suicide rates in areas surrounding airports to jet noise. Researches have determined that pre-existing mental conditions are not explicitly caused by jet noise, but are indeed exacerbated by aircraft noise pollution.
Further, it has been shown that long term memory retention and reading comprehension, along with solving difficult problems, are all impaired in children exposed to noise [11].
11 1.2.2 U.S. Commercial Aircraft Noise Regulations
The current regulations for aircraft noise certification in the U.S., the Federal
Aviation Regulations, Chapter 1, Subchapter C, Section 36, comprise Stage 4 of standards, implemented on January 1, 2006. Stage 4 applies to all new aircraft type design applications submitted afterward, but does not retroactively impose the newer, tougher noise standards on Stage 3 aircraft, nor does it provide any future date or intent to do so. The changes adopted are mirrored upon the ICAO Annex
16, Chapter 4 regulations, which essentially provides U.S. certification of all international aircraft certified to the newest Chapter 4 implementation of Volume 1,
Amendment 7, effective on March 21, 2002. Similarly, associated measurement and test analysis requirements found in part 36 are interchangeable with ICAO Annex 16 requirements. Discussing some of the most important U.S. certification regulations in Part 36 is then sufficient. One important caveat is that F.A.R. Part 36 does not regulate admissible noise levels at airports, it simply governs certification requirements based on assumed technological, economical, and other reasonable limits.
The primary certification requirements of F.A.R. Part 36 consist of three measurements of aircraft flyover noise at takeoff, sideline, and landing conditions.
Strict requirements for measurement systems, essentially compiling data on every known factor effecting noise absorption and propagation through the atmosphere, as well as flight trajectory data, is required to compute an acoustic evaluation of the flyover noise as the Effective Perceived Noise Level, measured in units of EPNdB.
To calculate the EPNL, a quantitative flyover noise metric used at each of the three
12 conditions, the 1/3 octave band spectrum is measured at ½ second increments during the flyover, and then corrected using noy curves to create Perceived Noise Level
(PNL) spectrum at 24 bands from 50 Hz to 10 kHz. Since tonal components are perceived by humans as distinctively more annoying that broadband noise, the greatest tone present in the PNL data, based on a narrowband spectral derivative threshold criterion, are then “corrected” effectively (to become PNLT) over the range of frequencies, based only on the maximum tone. The duration of the flyover is then accounted for by integration of the PNLT values, between temporal limits determined by a PNLT threshold 10 dB down from PNLTM (maximum PNLT at each time increment). The final EPNL value is then an algebraic sum of the
PNLTM values and the duration factor.
With EPNL computed at each measurement condition, it then must lie below certain limits. Since jet noise is a major contributor to the system noise of a given propulsion system and aircraft, especially on takeoff and climb, noting
Figure 1.1, noise reduction techniques that take advantage of EPNL calculation Figure 1.2 - System Noise Component may just as well reduce the overall Estimates, [] aircraft noise signature.
13 At the current level, increasing bypass ratio may lead to higher installation drag due to the required fan diameter, thus current efforts at controlling the noise
Figure 1.3 - Thrust normalized aircraft noise level history [11] emitted from the source are increasingly important [93]. This was the case with the
Airbus A380 which includes an oversized fan blade diameter compared to that for optimum SFC, due to requests by airlines for improved acoustics [12]. Another effect of the continuous de-coupling between acoustic and performance benefit is the desire to actuate noise control technologies in order to mitigate any associated acceptable level of performance degradation. It should be noted at this point, importantly, that essentially what is meant by noise reduction, is noise control. The two are used interchangeably, due to the fact that noise generating mechanisms are very robust and that a reduction in the total acoustic output of a high-reynolds number, high-speed flow is seemingly impossible without some penalty in weight or thrust loss, including Chevrons [87]. In light of the human response to noise illuminated in section 1.2.1.1 then, and physical positioning of the jet at takeoff conditions, a viable method is to alter the scales and coherence properties of the
14 turbulence such that the resulting radiated noise is lessened in the most sensitive bands, and most impactful directions.
1.3 Objectives
Several fundamental and technical goals existed for this work, in a common vein of attempting to investigate and develop practical, improved noise reduction schemes applicable to airbreathing engine exhaust systems. Modifications to chevrons, including simple geometrical variations, and both the addition of an independent testing of CVG’s (Coupled Vortex Generators) are employed to enhance jet mixing.
By employing a scale model, high bypass-ratio turbofan engine exhaust simulator rig, the value of both the fundamental and developmental work was enhanced in a practical sense. Specific goals of the research are as follows:
• Develop an understanding of the sensitivity of the various shear-layer
mean boundary layer characteristics on the far-field acoustics for both
conic and chevron nozzles, and provide scaling basis for CVG’s.
• Assess the effects of current chevron nozzles and CVG’s on the
turbulence characteristics over a large spatial extent using LDV.
• Determine the most important far-field acoustic characteristics of jet
noise for a high bypass ratio engine, with CVG’s.
• Determine the effect of some simple geometrical modifications to
chevron nozzles, and integration of CVG’s, on the far-field acoustics.
15 • Parametrically study some important geometrical variations, and
design aspects, of CVG’s and the effects on radiated noise.
• Employ incompressible CFD simulations to further explain the
possible mechanics of stream-wise vorticity formation for a CVG and
compare to other mixing devices including delta tabs.
1.4 Thesis Outline
The first three Chapters, the first being concluded here, cover the theoretical and experimental basis for the thesis. Immediately following starting with Chapter 4 are chapters providing the experimental and computational results, each pertaining to an individual effort, aimed at providing a proper piece of evidence to the goals outlined in 1.3. The execution of the experimental or numerical task is outlined in each chapter, followed by the analysis procedures and conclusions. A summary of the important conclusions from each chapter ends the thesis.
CHAPTER 1 - Introductio
A brief historical outline of the work done to reduce jet noise at the source, the turbulent jet, is presented. The motivation for jet noise reduction is considered.
Noise psychoacoustics and health effects are outlined in this vein of understanding the broad nature of the problem, and an outline of the U.S. noise metrics for jet aircraft is given. Objectives of this thesis conclude the chapter.
CHAPTER 2 – Background
16 Theory relevant to the fluid-dynamic mechanisms responsible for high Reynolds number turbulent flow noise generation are presented, and discussed in context for interpreting the acoustic and fluid dynamic results of this work. The initial development region of the jet and especially the mixing region are reviewed.
Recent research pertaining to flow dynamics, acoustic source properties, and radiated acoustics of mixing nozzle source noise reduction schemes are reviewed.
The near-field flow characteristics of vortex generators are also reviewed.
CHAPTER 3 – Experimental Facilities & Methods
Details of the anechoic chamber where the experimental rig measurements were performed, and the base hardware configurations and permutations examined, are provided. The physical arrangement and capabilities of each of the measurement systems employed, and the associated computer DAQ and pre-processing, up to the recording of the raw data file will be outlined in this section. General experimental setup requirements and details will be addressed, while specific experiment configuration information, such as probe traversing locations, spatial and temporal resolutions, etc., for a given set of measurements will be provided in the corresponding chapter. The measurement systems, for both the acoustic pressure measurements, and the fluid velocity measurements, include:
• Acoustic far-field microphone array
• Total pressure boundary-layer type pitot probe
• Dual-Component (u,v) 400mW CW laser doppler velocimeter
17
CHAPTER 4 – Boundary Layer Flow and Noise Results
Prior to collection of far-field acoustic data, a study of the mean properties of the boundary layer at the exit plane of conic and chevron nozzles was conducted. This prepared for evaluation and correlation of the importance of each boundary layer in effecting the far-field acoustics, and designing the scaling parameters for the CVG’s geometry. Far-field acoustic measurements accompany each configuration.
CHAPTER 5 – Initial Mixing Region Turbulence via LDV
The turbulence properties for both standard conic and chevron nozzles in coaxial flow are presented. This included high shear-layer spatial resolution radial traverses of a dual-component LDV probe, and centerline velocity measurements. For the centerline measurements, results for a first-generation CVG’s nozzle are included.
CHAPTER 6 – Acoustic Far-Field Noise Reduction Results
Far-field investigation of the noise properties of various core chevrons and core and fan CVG’s was conducted. Initially, simple geometric changes to existing noise chevrons was performed to estimate any sources of high-frequency noise, and further improve noise reduction performance. Internal and external shear layers of the core nozzle were included. A second geometry, also amenable to SMA actuation, involving chevrons internal to the nozzle, was performed. The main result of this investigation was to expose the sensitivity of the most important
18 parameters for affecting noise reduction by using vorticity producing mixing devices, as well as providing insight into future improvements.
CHAPTER 7 – Incompressible CFD Simulations
Identification of some fundamental physical properties of the mean flowfield generated by the CVG’s was carried out through a steady, 3D hybrid-mesh simulation of three simplified nozzle components. The domain, with about 9x105 nodes, using a steady RANS solver and κ − ε turbulence closure was modeled in
Fluent. Primarly, the sought-after insights into the flowfield were the importance of an additional design parameter on the developing flow region immediately downstream of the CVG’s, and to assess the first order character of the highly vortical flowfield immediately downstream of the devices.
CHAPTER 8 – Conclusions
A summary of the most important finding of this work, aimed at explaining the noise reduction potential for these devices in practical application, and recommendations for future research are reviewed.
19 Chapter 2 Background
The statistical fluid-dynamic properties of single stream and coaxial jets, and the dynamics of fluid structures important in the axial development of the jet, especially stream-wise vortices generated by mixing tabs, and the axisymmetric and azimuthally varying mixing-layer, are described in some detail. These highly dynamic aspects of the jet development, in terms of mean flow, turbulence properties, and quasi-periodic decomposition, will shed light on the re-organization of the turbulence and mean-flow structure by use of mixing enhancements, in departure from the baseline conic coaxial and single-flow jet hardware studied in comparison. This will provide the basis for constructing the appropriate form of the theory to aid in describing some of the aeroacoustic features observed, and the effects by CVG’s. A review of existing literature focused on the relevant developments related to subsonic and supersonic jet mixing noise through use of stream-wise vorticity enhancement devices completes the section.
20 2.1 Qualitative Jet Description
Flow structure of a coaxial jet, typical of that found during cruise, is very complex.
The core stream is issuing from the nozzle with a very high temperature and velocity, but subsonic based on stream Mach number, and the bypass stream with a much lower temperature yet low supersonic Mach number. A shadowgraph of this condition for a coaxial jet in static test is shown in Figure 2.1. In the heated core stream, it is clear that the spatial scales of the density fluctuations are fairly small, and of greater magnitude than those found in the secondary stream. Small, weak
Figure 2.1 - Shadowgraph flowfield of heated coaxial model jet at cruise conditions. [34] standing expansion shocks are perceived at the exit plane of the core stream, and two additional standing shocks are present near the latter part of the core nozzle plug. Inspection of the shear layers reveals sharp gradient very near the core and fan nozzle exit planes, which rapidly diminish in the downstream direction. This shear interface also reveals the growth of the shear layer. In the fan stream, larger standing shocks are evident, which diminish in relation to the underexpanded pressure ratio and also the external flow Mach number [48]. Mach wave radiation
21 appears emanating from the strongest shock in the train, and is directed upstream at approximately a 135° angle to jet axis. The high shear velocity of the fan stream by the quiescent surroundings has a strong effect on the development of the outer fan shear layer, which shows rapid downstream growth. The major features that affect jet noise can be detailed to some extent separately, as to a first order the effects are largely independent, and have been treated so by both experimentalists and theoreticians since Lighthill’s analogy was first developed to handle the complex fluctuating stress tensor. These fluid dynamic and spatio-temporal observations can be summarized as,
• Jet velocity field (mean and turbulence) and similarity regions
• Entropy gradients and fluctuations
• Supersonic flow quasi-stationary flow structures and density variations
• Installation and flight effects
Mainly, in terms of fluid mechanics, this work investigates the mean flow qualities in the initial development region of the jet and the turbulence characteristics and their axial evolution. Flight effects are explored in simulated capacity. The noise reduction potential for supersonic jets is also investigated, for a single jet, as due to the large mass flow rates of the bypass stream (up to ~12lb/s), the fan velocity was limited at sonic.
22 2.1.1 Mean Subsonic Coaxial Velocity Field & Similarity
Considering either a single stream or coaxial jet, three regions, the first and last exhibiting scaling similarity, are distinct in the axial development of the flow in a high Re , O(106), turbulent jet. As the flow exits the exhaust nozzle exit plane, a typical profile of which is shown in Figure 2.2, the initial conditions are characterized by top-hat velocity profiles with thin, turbulent boundary layers, with
Pr imary Potential Fan Potential U (r, x ) U (r, x ) Core 1 s1 Core 1 s2
δ f (x)
δ c (x) r δ (x) x w
Transitional Development Region Region 2nd Similarity Region
Small Scale Large Coherent Mixing Noise Structure Radiation
Figure 2.2 - Subsonic coaxial, high Reynolds number jet mean flow structure details momentum thickness on the order of 1% of the equivalent nozzle diameter. Viscous forces, within only the first several diameters downstream of the exit planes, convert almost ¼ of the thrust power of the stream into mainly turbulent motions [22].
In static ambient laboratory conditions, the coaxial jet is composed of five axisymmetric boundary layers, two for each stream. These can be separated into an
23 ) “inner” and “outer” boundary layer dictated by r . The outer boundary layers are further generally much thinner than the inner in a typical modern converging, conical nozzle due to highly converging internal contour and favorable pressure gradients. Also, in a separate flow fan and core exhaust nozzle exit plane configuration, as is typical of high bypass-ratio engines such as the GE90 and
GENX, the separation distance causes increased inner boundary layer growth of the fan stream by the time it reaches the core stream exit plane. The shear layer initial instability at such a high Reynolds number flow is very high in frequency, being inversely dependent on the boundary layer thickness. Across the core/fan shear layer, there exists also a thermal mixing layer due to the heated core stream.
Depending on the simulated engine cycle conditions, meaning the exit temperature and nozzle pressure ratio of both the fan and core streams, for a given simulated engine setting, the absolute stream velocities may exhibit either a normal or inverted velocity profile. The normal velocity profile is constituted by a core stream with a higher velocity relative to the fan stream, and vice versa for the inverted profile. The acoustic benefits of each scenario, as investigated both experimentally [36] and by modeling of the instability wave characteristics and varying the stream velocity profile and density conditions [37], show that the major differences in radiated noise are experienced when the phase velocity of the instabilities is supersonic. Further, variations in the ratio of the fan stream velocity
to core stream velocity, U p /U s , determine the noise amplitude significance due to the inner and outer shear layer instabilities, with the trade occurring at a value of about 0.5. Experiments determined that slightly supercritical core pressure ratio
24 could lead to the largest decrease in downstream directed noise. Although this velocity criterion may be difficult to coordinate with real engine operating conditions, this illustrates the crucial dependence of instability wave growth on the spatial distribution of the momentum flux for equal thrust levels.
At about 2-4 diameters downstream of the exit plane a transitional region of no simple similarity occurs, due to highly non-linear motions of vortex merging as the potential flow regions decay completely [49]. Axial mode switching from the preferred mode of the jet column instability, to the dominance of the azimuthal or helical modes occurs in this region [38]. Thus, important flow features that lead to increasing jet mixing of the potential flow with the surrounding flow and reorganize the structures that radiate acoustic density fluctuations are prescribed highly by the flow development in this region and upstream.
After the transition region, where any separate streams have mixed well enough so that the average velocity profile is once again axisymmetric, the jet reaches a state of similarity [49]. Past this point, the entire flow is subsonic, whether is began subsonic or not, and scaling laws reduce the mean flow and turbulence statistics in a simple way. This occurs about 8-10 diameters downstream of the nozzle exit, where noise generation is a relatively insignificant byproduct of the turbulent flow compared to upstream regions.
25 2.1.2 Supersonic Jet Structure
The central regions of a supersonic jet contains a repeating train of oblique shocks, in high pressure ratios extending past the potential core, due to under-expanded flow issuing from a nozzle, such as at cruise conditions, where the turbofan bypass flow may be slightly supersonic. In this case, the flow consists broadly of a superposition of nearly frozen small-scale turbulence, larger mixing-layer structures and quasi-stationary discontinuities in pressure and temperature. Figure 2.3 shows a simplified schematic of the gross flowfield development for a single jet. Near the boundary of the mixing layer at the sonic line, the shock cells interact with the vortical structures, which radiate sound. A more detailed diagram of this mixing
Figure 2.3 - Schematic of shock-cell flow structure development for simplified single
axisymmetric jet without centerbody plug [44] layer is shown in Figure 2.4. By modification of the mixing layer, which highly interacts with the noise-producing shock-cell tip regions of the shock-cell structure, it will be shown that significant reductions in the shock-associated broadband and tonal noise components can be realized. Schlieren imaging similar to the
26
Figure 2.4 - Supersonic mixing layer detailed shock tip schematic [34] experimental conditions used in this work reveals alternating compression and expansion wave structure in a typical underexpanded jet at low supersonic Mach number. The external flow plays a crucial role in the development of the shock cell
Figure 2.5 - Schlieren nanospark imaging of underexpanded conical nozzle core flow in static
conditions at the UC Aeroacoustic Test Facility [48] structure decay rate, and thus the radiated noise, as shown by many researchers employing both theory and experiment [49,50,48].
27 2.2 Aeroacoustic Theory
The foundation of aeroacoustics lies principally in the relation, proposed by James
Lighthill in 1952 [35], that relates in exact form from first principles, the influence of turbulent flowfield fluctuations on the radiated acoustic field. In making the connection between the actual fluctuating stress field and forcing of the acoustic medium, the analogy of an imaginary, external forcing superimposed on a globally stationary acoustic medium was made. To construct this analogy, the fundamental equations of motion for a fluid, the conservation of mass, and conservation of momentum, in differential form, neglecting sources and body forces, are written as
∂ ∂ ρu + (ρu u + p ) = 0 , ∂ i ∂ i j ij (2.5) t x j
∂ρ ∂ + ρu = 0 ∂ ∂ i (2.6) t xi
where pij is comprised of normal and viscous shear stresses, as
= δ − µ − 1 δ pij p ij 2 (eij 3 ekk ij ) . (2.7)
By acknowledgement of what is meant by the analogy, the “stresses” inherent
2 ρδ in the quiescent acoustic field, a0 ij , are simply subtracted from the stress field imparted by the actual flowfield, and then by rearranging (2.5) and (2.7) into
∂ ∂ρ ∂T ρu + a 2 = − ij ∂ i 0 ∂ ∂ (2.8) t xi x j where Lighthill’s stress tensor,
28 = ρ + − 2 ρδ Tij uiu j pij a0 ij (2.9) is the external forcing field, imposed by the real fluctuating fluid stresses. In this fashion, then, substituting (2.6) into (2.8) produces an inhomogeneous wave equation displaying a quadrupole source distribution as the forcing function, as,
∂ 2 ρ ∂T − a 2∇2 ρ = − ij ∂ 2 0 ∂ ∂ (2.10) t xi x j which, when through application of a general solution to (2.10) in simplified form, ignoring negligible viscous terms by consideration of the high Reynolds number prevalent in jets of interest, and avoiding surface presence terms, dimensional analysis may be performed. Taking the Strouhal number of O(1) based on jet diameter and mean velocity U, leads to a dependency of acoustic power output,
P ∝ ρ U 8a 5l 2 0 0 (2.11) which displays the thoroughly studied dependence on the mean flow velocity of about the 8th power. This result aided in prompting further increases in commercial turbofan bypass ratio, beside the most prominent reason of increased efficiency, which is approaching 10 currently. An obvious feature of the high dependence of radiated acoustic energy on jet velocity is that it must be eventually constrained by
ρ 3 the total kinetic energy flux of the flow, AeU , as mach number increases.
By referring back to the wave equation (2.10), a tractable solution requires simplifications that, although based on this exact form of the wave equation, exert important effects that are important for modifying the source distribution through the introduction of stream-wise vorticity at the nozzle lip. Before simplifying the exact integral form due to jet flowfield assumptions, first the condition of the far-
29 field, that the source region is compact compared to the observer location as v v y << x , is simply applied and a solution to the unmodified inhomogeneous wave
equation may clearly be re-expressed in terms of Tij still evaluated at the retarded time as v v 2 ⎛ − ⎞ v 1 xi x j ∂ v x y ρ'(x,t) = T ⎜ y,t − ⎟ . (2.12) π 4 v 3 ∫ ∂ 2 ij ⎜ ⎟ 4 a0 x J t ⎝ a0 ⎠
A first simplification of (2.7) is to discard viscous terms since the far-field density fluctuations, when considered independently by this source term, depend at most on the eddy mach number as m5 , and on Reynolds number to the inverse power [51]. Clearly, for the high speed jets used in this study the effect is negligible. Secondly, the form given first by Lighthill, equation (2.12), is that of the stationary source case neglecting what has been termed the acoustic-flow interaction. This includes the propagation of sources within the jet through the mean flow and turbulent eddies that refract and scatter sound. It has been shown that isotropic turbulence by itself generates acoustic radiation [53], which reveals the importance of treating the effects of mean flow gradients on the inherent “self” noise of the flow, and thus regarding the complex mean structure as a separate aeroacoustic phenomenon. Decomposition of the Lighthill stress tensor to singularize the complex effects can offer low order estimation on the importance of each.
Before doing so, it is worthwhile to investigate incorporating the effects of convecting sources, by directly incorporating the Doppler modification term corrected from Lighthill’s initial low-mach number expression as proposed in the
30 following form [52], considering the explicit case of supersonic eddy convection velocity in the direction of Mach wave radiation, by decomposition (that turbulent eddies possess unsteadiness relative to the average convection frame),
1/ 2 αv 2 []− θ 2 + r (2.13) 1 M c cos( ) 2 . a0
The effect of convection on the directional character of the radiated noise is shown explicitly in Figure 2.6, for Mach numbers from 0.2 to 1.0. This source modification is directly implemented into (2.12) for inclusion into the far-field radiated density fluctuation, creating a large amplification of the downstream directed noise, and attenuating less so the upstream radiated noise. This, as should be suspected, is frequency dependent, as well as depending on the spatial distribution of the noise generating sources. The directionality shown in Figure 2.6 is typical of the nearly frozen fine scale, random turbulence, and as Figure 2.6 - Source convection effect shown agrees well at lower mach numbers [54]. on directional radiation of jet noise, One immediately obvious deviation from this for various jet Mach numbers [52] convection picture occurs at close angles to the jet axis at about less than 20°, where a “cone of silence” has been observed showing a relative region of 20dB quiet compared to more upstream levels. This refraction effect is a function of the mean flow profile and the density ratio of the jet with the surrounding fluid. Attempts have been made to modify Lighthill’s original 31 formulation to account for mean flow refraction effects, namely that by Lilley, where modification of the wave operator resulted in a third order equation, which
Lilley himself, as well as others, have had difficulty in the justification thereof [23,
30]. Regardless, employing the third order axisymmetric Lilley’s equation in terms of its linearized-variable sources has been used with some success to illuminate the contributions of refraction and density gradients on the individual right hand side terms of
1 d ∂p dU ∂ 2 p D3 p − D∇ 2 p − (log a 2 )D + 2 = 2 ∂ ∂ ∂ a dr r dr x1 xr ∂ 2 (u u ) dU ∂ 2 (u u ) (2.14) ρ i j + 2ρ 2 j + H.O.T. ∂x ∂x dr ∂x ∂x 142i 43j 1424 4314j self shear which correspond to turbulence-turbulence interaction “self” noise and the interaction of the turbulence components with the radial velocity gradient, termed the “shear” noise. Linearization creates higher order source terms that are unimportant to the total noise production [55], which are indicated on the right hand side of (2.14). Although this method has helped highlight the differences in the shear and self noise terms in flows of practical interest, and indicate the sensitivity of radiated noise on the turbulence properties to the degree of anisotropy, the utility in an overall sense of spectral accuracy, even in simpler flows, is limited.
Decomposition of the Lighthill tensor, however, coupled with DNS, has shed some light recently on the statistical identity for the fourth-order space time covariances that are the crucial sources of the far-field sound production in the acoustic analogy. By recognizing the acoustic-flow interaction directly in
Lighthill’s source term as a decomposition of turbulent-turbulent interaction, and
32 turbulent-mean flow interaction, using a typical Reynolds decomposition on the velocity component vector, a desirable separation of these effects is made, in terms of the nature of the source term. Estimating Lighthill’s fluctuating stress tensor as
T '= ρ(u u′ + u′u ) + ρu′u′ + ( p′ − a 2 ρ′)δ . ij 142i4j 43i4j 12i3j 1424 0 434ij (2.15) shear self entropy
This form neglects viscosity, as previously mentioned would be discarded.
The first term represents the interaction of the various first order turbulence quantities (as long as the region considered exhibits similarity so that the mean flow can be considered a parameter) with the mean flow. Turbulence-turbulence interaction is represented by the self-noise second order term. Entropy interactions are expressly retained since they have been shown to be correlated to a significant degree with other source terms, and thus may be important to noise production even in isothermal jets [57]. This type of source evaluation has shown that the directivity of the self noise is much more omnidirectional than the shear noise term, which falls off dramatically at emission angles perpendicular to the jet. Introduction of turbulence by modifications to the nozzle exit will undoubtedly create important self noise terms, and due to the stream-wise vortex structure, may introduce additional dipole sources from vortex unsteadiness. A coaxial flow around the structure will then likely create the additional shear noise term which will affect downstream angles.
33 2.3 Jet Instabilities
Crow and Champagne instigated research into the development of organized turbulence inherent in high Reynolds number flows in the form of large scale structures, or instability waves, that have been shown to directly radiate noise at supersonic phase velocities [58, 50, 52]. Since their experimental verification of organized turbulence in jet flows, and investigation into the response of these structures and the impact on jet mean flow and turbulence development, great amounts of research have been conducted to identify concrete effects of jet instability on flow and noise, confirming the main effects on even very high
Reynolds number flows. The impact of these structures on supersonic noise radiation was observed by Phillips in departure from Lighthill’s all-inclusive solution, deeming the “eddy Mach wave” a by-product of using Lilly’s convective wave equation, which revealed the shear noise terms uniquely (although neglected important self noise terms), although it was actually a predictable consequence of
Lighthill’s theory, buried in the complicated stress tensor [51]. Since then, the conditions which affect the development of large scale structures, mixing layer structures development has been extensively researched. Important aspects related to this work are the effects of near-field flow conditions, by description of the momentum thickness existing at the nozzle exit, and turbulence levels in the mixing layer and jet centerline, on the development of large-scale structures and ultimately on far-field noise. The modification of instability modes by influences on the mixing layer typical of CVG’s should be considered, although no direct
34 measurements have been attempted in this work due to recent experiments indicating weak effects of chevron nozzles on the energetic modes.
2.3.1 Effects of Initial Condition on Mixing Layer and Acoustics
Transition from laminar to turbulent boundary layers at the nozzle exit plane has been shown to have a significant effect on the evolution of the shear layer & large- scale coherent structures that comprise much of the turbulent energy [39, 40, 41].
The impact on radiated noise has been less well studied. Simply modifying the initial momentum thickness of a laminar shear layer changes the spatial onset of large growth rates in shear-layer instabilities [42, 43], so that vortical roll-up occurs farther downstream as initial boundary layer thickness increases. The distribution of energy from larger scale convecting toroidal vortices is dependent on the location of roll-up, which in turn governs the onset of saturation in the growth rate that signals spatially the energy transfer from larger scale structures to fine-scale turbulent motion. Altering the initial momentum thickness of a laminar shear layer therefore changes the spatial distribution of turbulence scales and spectrally the energy distribution, and naturally the noise radiated from the sources convected therein. When laminar versus turbulent conditions are examined, different results are found. Most prominently, sensitivity to Reynolds number based on momentum thickness is insignificant in either state alone, but the transition to turbulent state increases the spreading rate and delays the onset of self-preservation [41, 43, 45].
Further, compared to laminar jets, the turbulence distribution in a broad gamut of wavenumbers hampers the formation of large-scale structures. Into the fully
35 turbulent state, the axisymmetric free shear flow then exhibits various linear regions of growth corresponding to the nature of the coherent structures before and after the transition region.
While evoking fundamental understanding of various components of vortical development and shear layer instabilities, generally in lower Reynolds number flows where issues such as vortex pairing may be a significant contribution to the radiated noise, real exhaust systems issue turbulent flow at high Reynolds numbers.
Experiments have attempted to characterize the influence of variations in mean and turbulence properties for the evolution of initially top-hat profile flows with thin, turbulent boundary layers, more recently with direct evaluation of radiated acoustic fields as a primary result [47].
For an fully turbulent axisymmetric shear layer, boundary layer momentum thickness is a function of the nozzle exit Reynolds number, and the upstream surface contour, assuming a subsonic converging nozzle typical of a modern gas turbine engine. A plane mixing layer can be fully characterized by the Reynolds number based on the initial momentum thickness, the average velocity across the initial mixing layer, and the turbulent state (intensity, spectrum) of the flow.
Axisymmetric mixing layers introduce azimuthal curvature as another parameter.
High speed subsonic and supersonic boundary layers form very thin sheets of vorticity essentially independent of the nozzle diameter (except in the sense of curvature), that are subject to the Kelvin-Helmholtz instability and exhibit rapid, but also initially linear, downstream growth [49, 61]. At its beginning just beyond the nozzle exit plane, the shear layer exhibits a bandwidth of naturally unstable
36 frequencies from Strouhal numbers of about 0.011 to 0.017 [63], that are excited by the turbulence present in, and convecting with the boundary layer. This high- frequency instability wave then governs the formation of large-scale structures that grow and merge, governing entrainment and the production of turbulent kinetic energy at smaller scales across the shear layer similar to jets of lower Reynolds numbers. Introducing stream-wise vortical structures, then, will drastically alter this process. Immediately beyond the nozzle exit, the transfer of axial momentum into rotational energy in the vortices will lead to in increase in the mixing layer growth rate through increased entrainment. This may be analogous to altering the location of virtual origin in initially laminar jet shear layers. Injection of larger turbulent energy into the mixing layer will affect the coalescence of structures farther downstream in the transitional region. Although it appears that indeed jets with and without forcing in the mixing layer will achieve similar properties far downstream at about 8-10 D. Near-field acoustic correlation measurements have shown that a majority of the radiated noise is produced within several diameters downstream of the end of the potential core, which terminates at about 4-5 D [34].
Thus, the initial properties of the vortical structures, their interaction with the mixing layer and its growth rate, and the growth rate of the stream-wise vortical structures themselves, should they have a long enough coherence time, will influence the radiated acoustics. This will manifest in changes in the directivity pattern for the peak noise, in spectral content of reorganized turbulence structure, and potentially in new sources introduced by the stream-wise structures, or through impact of these with nozzle surfaces.
37 2.3.2 Influence of Jet Instability Modes on Flow and Acoustics
Several properties of the jet instability modes and mode numbers, and their impact on noise generation, including flows emanating from chevron nozzles have been reported [66]. The preferred mode in high Reynolds number jets is associated with the diameter of the nozzle exit, distinguishable from the higher frequency K-H shear layer instability, or otherwise the annulus height for concentric axisymmetric flows
[64]. For a circular jet, the preferred mode has a Strouhal number of about 0.3, while non-dimensional vortex-passage frequency locking for planar jets asymptotes to a maximum of about 0.44, dependent on the ratio of jet height to initial momentum thickness. For a circular jet this corresponds to the most lightly damped axisymmetric mode, associated with the jet column for the m=0 eigenvalue. Linear instability analysis based on a weakly non-parallel (slowly diverging) jet mean velocity profile has shown that initially the jet supports an effectively infinite number of unstable modes, which is altered by the development of the mixing layer surrounding the jet so is a function of Mach number [68]. For a high-subsonic jet, the axisymmetric mode exhibits greatest growth rates until the end of the potential core region, at a Strouhal number roughly the same as the first helical mode m = ±1
[67]. Near the end of the potential core, the helical mode becomes the most unstable [64], and further only modes up to m = ±2 contain a very substantial portion of the energy spectrum. Initially, the role of these instability waves and the convecting structures, in view of axial evolution, is to entrain fluid into the vortices in between the chaotic pairing and tearing events that occur at high Reynolds numbers [49]. These instability wave packets, in the form of large-scale coherent
38 (relatively) structures, quickly growing, pairing, and tearing from nozzle exit to the vicinity of the potential core termination, may exist for long periods of time, and influence largely the spread rate of the jet or equivalently the entrainment. As the flow becomes supersonic, the helical mode appears earlier in the jet, and dominates the energy spectrum of the azimuthal modes [49, 61].
At very high Reynolds numbers studied in this work, the coherence of these structures is subjected to substantial jitter compared to the lower Mach numbers typically found in the literature, thus have lower coherence peaks, but can be well- correlated spatially for up to several jet diameters. These observations, and prior flow visualizations of the instability waves have directly lead to empirical observation of acoustic similarity profiles for both subsonic and supersonic single jets, composed of the higher frequency, less directional, fine-scale turbulence self noise, and the lower frequency, highly directional, instability-wave generated noise
[68]. Recent analytical models of instability-wave generated noise have confirmed these general directional and scaling trends [70]. However, extension of similarity to experimental coaxial jet data has shown that the similarity profile is inconsistent for aft-directed low-frequency radiated noise, assumed to be produced by instability wave mechanisms [71]. The complex structure of a coaxial jet, which has an inner shear layer (see Figure 1.1), is not simply amenable to linear instability wave similarity assumptions. Still, of particular interest is the role instability waves have on the radiated acoustics, and how current technology, including Chevron nozzles, modifies these waves dynamically and acoustically. Recently, experiments were conducted to answer the question of the impact of proven noise-reduction methods
39 using Chevron core mixer nozzles on these noise-producing structures in isolation, focusing on the lower energy-containing modes [66]. Based on the modes investigated based on the fluctuating axial velocity component, no large differences in energy redistributions between modes were obtained, and further that the turbulent fluctuation modes behaved similarly for a conic baseline nozzle and a
Chevron nozzle. Acoustic modes revealed similarly that the distribution of modal energy was not large, and that two different Chevron nozzles produced similar results. Consequently, although the mean velocity statistics and TKE from Chevron nozzles has been well documented [109] in connection with aft-directed noise reduction, the interaction with instability wave noise generating mechanisms is weak or unclear at best, and further experimentation is needed to clarify effects on the most energetic modes.
2.4 Mixing Nozzle Research
Since the “teeth” nozzles of Westley and Lilley [72], and the more industrial corrugated mixer studies of Greatrex [73], the variations of nozzle geometry tested for the purpose of turbulent mixing noise reduction has become vast. Revelations on the instabilities of the jet mixing layers, the relevant supported jet modes and their impact on noise production, and the clarification of coherent structures in mixing and noise radiation for both subsonic and supersonic phase velocities, have urged more complex (non-static) methods for control of the jet turbulence.
Generally, mixing schemes are classified into static, pseudo-active, and active control. Chevrons, lobed mixer nozzles and mixing tabs have been intended for
40 static use historically. Pseudo-active technologies may be classified by either slow actuation that relies on open-loop “control”, including SMA materials, or respond fast but dependent upon flow velocities in the jet, such as flexible filaments [74].
Fully active methods actuate to a prescribed response on demand, and provide means of optimizing noise reduction. Fluidic injection, water/polymer injection, synthetic jets, “morphing” chevrons, or acoustic forcing exemplify active technologies of interest in jet noise reduction. Primarily, trends in the design, scaling, fluid and acoustic mechanisms for static turbulent jet mixing nozzles found in the literature have provided the basis for investigating CVG’s. This survey will recover the salient research for vorticity producing mixer nozzles, such as chevrons and mixing tabs.
2.4.1 Tabbed and Chevron Mixing Nozzles
Nozzles with the addition of triangular or rectangular surfaces impinging into the jet flow at or near the exit plane, either normally or at some angle to the stream vector, are generally referred to as tabbed nozzles.
The size of the tab is usually small compared to the diameter of the nozzle, when considering single jets, Figure 2.7 - Early "Tabbed" or the annulus height for a coaxial or plug flow jet. Mixer Nozzle of Westley & Some of the first studied on a model scale jet exhaust Lilley, 1952 [72] rig are shown in Figure 2.7. Due to the highly spatially coherent vorticity field, tabs have been also termed vortex generators,
41 similar to those used extensively in passive flow control over wings and bodies.
Chevrons nozzles are similar except that the base triangular section is well aligned with the flow, the impingement into the nominal streamlines on the order of 1% the nozzle diameter.
The study of the tabbed devices, some shown in Figure 2.7, is one of the earliest for the sole purpose of noise reduction. In those static, blow-down type rig tests, both subsonic and choked operating conditions were used to assess the effect of tabs on the noise field. A maximum spectral 12dB reduction in shock noise, and a 5dB reduction in subsonic noise was measured, with the configuration noted in
Figure 2.7.c.2, with three tabs directed normally into the flow, and three at a 30° angle to the nominal streamlines. These tests supposed the effect of “corrugation” of the mixing layer to be the main effect, based on the relatively large obstruction of the flow (about 0.25D2) which would be unacceptable today. This and other research led to the corrugated and lobed mixer nozzles of the sixties and seventies, which were similarly studied extensively on low bypass ratio engines. As jet engines have asymptotically moved to higher bypass ratios, the weight, complexity, cost, and performance losses of these nozzles was not warranted, and especially desirable were modifications (“hush-kits”) that could attach to existing engines with a very low performance penalty. Chevron nozzles were essentially the first commercial solution in this respect, and proved to reduce the total PNL signature of exhaust systems with marginal increases in any noise component [77], which was similarly desirable for new engine integration.
42 Before commercial research and development of Chevron nozzles, it was shown, motivated by the work at Cranfield, that triangular “delta” tabs projecting into a jet flow at the nozzle exit plane dramatically altered the turbulence properties.
Measured contours of lateral velocities near the tabs confirmed existence of a pair of spatially coherent stream-wise vortices [78]. Increased maximum axial turbulence intensities and shear stresses in the first few diameters of the round nozzle were traded for lower downstream levels, and correlation length scales were reduced near the jet centerline. Little effect was seen on jet centerline velocity decay.
Extension of these low Mach number and Reynolds number tests was made to flows of a significantly more practical regime for jet exhausts, showing that substantial far-field acoustic reductions in both broadband shock-associated noise and screech tones was possible [79]. These tests comprised a large range of supersonic pressure ratios from sonic to about Mach 2, and static temperature ratio up to about 2.5.
Only one relevant jet angle to screech noise, at 90º, was reported. Subsonic jet mixing enhancement and turbulence development was studied using pairs of 1/16D square tabs, as well as faired wedge tabs with equivalent end shape, installed at the nozzle exit plane [76]. Notable is that a large distortion of the jet by two tabs even in subsonic high Reynolds number flow of about 6x105. Any number of tabs accelerates centerline velocity decay, the fewest having the greatest effect, and although turbulence intensity is higher along the jet centerline, the energy spectra show a redistribution of large peaks near the end of the potential core to both lower and higher frequencies, suggesting severe modification (suppression) of organized structures.
43 Tabbed nozzles of varying configurations (number of tabs) were studied in greater detail with respect to the global flow structure, in terms of mixing layer development, and mean flow centerline velocity measurements [80] of low supersonic jets. With an impingement of 0.20D into the Mach 1.12 jet, these relatively large devices were certainly effective, in either one, two, three or four tabs azimuthally equally spaced around the nozzle perimeter, at mixing the jet drastically. Cross-stream total pressure surveys far downstream at 9D indicate almost bifurcated jet distortion in some configurations, and showed enhanced entrainment. Considering the both the change in the maximum rate of velocity decay was largely increased, and the virtual origin of the jet shifted several diameters upstream, it should be unsurprising that screech tones were completely eliminated in near-field acoustic measurements. Although a dual configuration of smaller projection tabs was tested, with impingement of about 3/16” or 0.10D, diametrically opposed on the nozzle exit, the centerline velocity showed little decay in comparison with the unforced jet. No acoustic measurements were reported for this case.
Further details on the flow structure effects on jets were warranted, and
“vortex generators” (VG’s), in the form of triangular tabs were measured in regard to the influence on entrainment of Mach 1.63 supersonic jets (in a preliminary study to [82] by Zaman et al). Of two orientations of the triangular tabs evaluated, one with apex directed upstream at 135º to the flow, and another at 45º, a counter- rotating pair of vortices was documented from the leeward edge, aiding in an increase of 50% greater mass flux in the far downstream region of the jet. Similar
44 results of highly increased entrainment were found for a subsonic rectangular free jet [83]. Various observations on the detailed flow structure clarified the mechanisms of the increased entrainment and screech noise reduction [82] in the subsequent related study. In both, somewhat more realistically sized tabs, about
1.5% of the jet area in blockage, showed that large distortions of the jet could be produced, and further measured predictably large thrust losses penalties associated with flow blockage and CD. Effects of initial conditions showed that the flow structure appeared to be unaffected by reasonable changes in momentum thickness and initial turbulence levels. Additional quantitative hot-wire data in a low Mach
Number (0.3) exit condition supported the original findings of a tradeoff in turbulence and shear levels in [78], and showed the expected result, also indicated in
[78], that stream-wise vortices are formed, counter-rotating pairs by each tab, embedded within the azimuthal vorticity sheet.
Parametric study was initiated to clarify mixing and acoustic sensitivities in a subsonic square jet [83] anechoic facility. The aspect ratio, scale, angle of attack, and spatial configurations of half delta-wing vortex generators were assessed. At short downstream distances on the order of 4D, the mixing and entrainment effectiveness of a simple pair of half VG’s was very effective. Low frequency noise in the downstream direction was found to be well attenuated by most configurations, but pairs with an axial offset, assumed to interfere in creating a common vortex, produced larger high-frequency noise. All configurations were set back from the trailing edge of the rectangular nozzle, and very likely produced dipole surface sources. A single pair of two VG’s, separated span-wise by several VG lengths
45 produced good downstream reductions in noise, and only incurred slight increases at side angles. Noise reduction increased with Mach number in this case.
Axial positioning of a triangular tab was varied in an under-expanded round
Mach 1.63 jet [84], showing large distortions in flow-field for an upstream tab displacement of 1.5 times the tab length, including inverting the vortex pair rotational sense, but decreased effectiveness in altering screech suppression.
Further detailed study of a triangular tab, at a 45º to the flow, although in very low
Reynolds number mixing flow O(1000), that span-wise tab spacing for an array of tabs, and also the pitch angle of the tab, greatly affect the maximum stream-wise vorticity and small-scale eddy density. Other experimenters found that mixing was more dependent on tab area and number than on shape and pitch [92] for more realistic Reynolds numbers. They also noted apparent independence on Mach number for flow bifurcation, indicating that the stream-wise vorticity mechanism was relevant through a large flow regime. Further, for a subsonic jet of Mach 0.92, with a circular nozzle and per-tab blockage of about 1.8%, recent tests [89] have shown similar decreases in peak noise as the number of tabs increases.
Tab mixing noise acoustics, which has been almost exclusively investigated for reduction of screech noise, on basic rigs representing idealized flows, was recently studied on a medium bypass ratio mixed exhaust model [86]. The exit plane of the mixed exhaust was located approximately 4.2 equivalent core diameters
(plug nozzle configuration) of the core nozzle exit plane. Combinations of triangular, 45° tabs, in number and blockage ratio were examined. In the subsonic range tested, core-only tabs produced a 2-4dB decrease in peak noise, remarkably
46 without significant high-frequency tab-induced noise. Applied to the mixed exhaust flow, minimal peak noise reduction was attained, and tab self-noise produced large high-frequency broadband increases in noise. Tabs were also applied to the mixed exhaust with a 12-lobe core stream mixer nozzle, in two different clockings, one with the tabs azimuthally centered on the “hot spots”, or along the bisector of the core lobe, and the other on the bisector of the fan chute. Slightly greater reduction in low frequency noise was achieved, while increases in higher frequency noise led to EPNL levels approximately equivalent with the 12-lobe nozzle alone for a fairly large range of thrust levels tested. Turbulence data showed that high turbulence levels at the mixed exhaust plane may have been responsible for generating loading self noise for those tab configurations.
These results are particular to the mixed exhaust configuration, then, and the specific tab design, but these results and those prior have shown that in certain configurations, stationary stream-wise vortical structures can reduce the noise. The mechanisms proposed have included an increase in the “interfacial” area of the mixing layer, and separately the enhanced entrainment of fluid surrounding the primary flow as a direct result of energizing the stream-wise structures through geometrical changes. Another recent analysis [90], employing energy balance turbulence decomposition involving slow and more quickly varying (smaller scale) fluctuating components, has speculated that “deturbulization” of the flow can occur by these vortices, by inhibiting the transfer of energy from the main flow to large scale eddies. Relating each of these conclusions, critically, is the formation of the stream-wise structures. Integration of these flow structures into a particular exhaust
47 configuration, and thus the potential acoustics benefit, will then depend on details of the formation and the mixing layer into which they evolve.
Chevrons are in large result a descendent of the tabbed and lobed mixer nozzle designs [93]. Now in successful commercial service, and being implemented on bypass streams which, at high bypass ratios, are now large generators of noise, current research and development is being focused on application of active control and optimized shaping [88]. The knowledge accumulated by detailed investigation of the turbulent flow-field, and the acoustic fields, has provided the groundwork for further development of chevrons investigated in this work [87, 109].
2.4.2 Near-Field Fluid Structures of Vortex Generators
The axial location of vortex generation takes place at the wake of a splitter plate (nozzle lip) which then quickly becomes a developing mixing layer, posing many complex interactions between the azimuthally oriented vorticity sheet and the stream-wise structures embedded in the developing shear. However, in light of the general details of vortex dynamics embedded in turbulent flow, and sufficient details about the initial mixing region, some qualified observations on the experimental data can be made. Although a clear modification to the jet is evoked through study of global development of the jet itself, by centerline velocity measurement and statistics, for instance, the development of the bound vortical structures themselves are crucial to this end. Some clarification of the flow-field development by these structures in elementary flows will provide insight, as in
48 Chapter 7, investigating the averaged flowfield unique to the CVG geometry embedded in a plane developing boundary layer.
Both conventional and newer generations of vortex generators, diversely applied for local addition of highly spatially coherent regions of circulation (e.g., improved mixing for combustion control [94]), or freestream momentum extraction and spatial transfer (e.g., flow separation control [95, 96]), have been characterized by the functional parameters of interest in promoting long-lived stream-wise vortices [96]. Of most interest are the effects of VG self-geometry parameters, such as overall physical scaling, planform shape, aspect ratio, and inclination angle.
Arrangement in patterns, such as pairing, stacking in the flow direction and other configurations has illustrated combined effects on vorticity generation. The effects of flow parameters, mostly Reynolds number, Mach number, and boundary layer conditions have further clarified the optimal use of VG’s, and provided valuable insight to the dynamic vortical qualities for various flow applications.
Distillation of the prominent results from a selection of the large array of VG geometries of stream penetration O(δ ) explored have been summarized recently
[95]. Most of these devices were applied to boundary layer control for airfoil stall alleviation and lift enhancement, and inlet separation and distortion control. VG height, given as H /δ , is essentially the critical parameter describing the function nature, that is, whether the VG is more of a producer of vorticity than an reorganizer of existing oblique turbulent structures within a boundary layer [96, 100]. At a height of slightly less than the boundary layer thickness, peak vorticity initially remains constant and then decreases approximately like 1/ x 2 in a subsonic turbulent
49 boundary layer [101, 102]. Similar results are found for the subsonic plane mixing layer [103]. The effects of VG geometry on the initial vorticity and circulation were studied analytically in a quasi two-dimensional fashion compared to experiment for boundary layer flow [101]. Airfoil aspect ratio, for values under about 2, greatly influences the maximum vorticity, with a local maximum at AR ≈ 1. Circulation, however, decreases as 1 AR at high, but probably non-separated angle of attack.
The effect of VG height, as it approached the boundary layer thickness, evoked a linear increase in circulation, as would be expected, but the peak vorticity reached a maximum near H = δ . Both angle of attack and stream velocity showed a linear increase in both circulation and peak vorticity, also as would be expected. These results are in line with results from vane-type VG’s, which in recent computational studies and experiments showed similar evolutions of the circulation and vorticity
[104, 105].
Some flow development details of triangular delta tabs have been investigated for their mixing properties and vortical development in jets and plane mixing layers [85, 82]. These insightful but limited studies either concentrated on gross flow-field details, or used tabs that were so large that jet bifurcation took place. Some limited data on maximum stream-wise vorticity decay was provided, for a single geometry. More basic results were provided on similar tab geometry, focusing on vortex development and spatial decay, of multiple tab configurations at lower Reynolds numbers to allow hot-wire measurements [97, 98]. The most important results were on the additive nature of the vorticity sources based on the tab inclination angle to the flow, and the significant dependence of a properly
50 spaced array of tabs on mutual stabilization such that the downstream peak vorticity and circulation decay rate is attenuated. It was also shown that small-scale structures, generated in the unstable core region of larger-scale vortices can be enhanced by addition of stream-wise vortices [97], which although, for delta tabs, exhibit a spatially consistent peak stream-wise to azimuthal vorticity ratio of about
20% (relatively weak) [85], by inducing premature instability in the larger scale structures enhancing gross mixing in addition to larger scale mass transport [99].
Whether this mechanism holds for high Reynolds number jets is unknown.
51 Chapter 3 Experimental Facilities & Methods
The main facility used in this research was the anechoic Aeroacoustic Test Facility
(ATF), located in the Gas Dynamics and Propulsion Laboratory (GDPL) at the
University of Cincinnati. The unique air storage capabilities (22,000lb @ 1800psig,
1µ filtered) and the recent installation of a 360 SCFM compressor providing recharging times under 8 hours, allows for high mass flow rates of up to 20 lb/s, for long run times, to the Nozzle Acoustic Test Rig (NATR) housed in the ATF.
Additionally, a low-pressure air delivery system (4500lb @ 170psig) was employed to allow for very stable flow conditions (within 0.1% RMS).
52 3.1 GDPL Anechoic Aeroacoustic Test Facility
A schematic of the plan configuration of the anechoic ATF is shown in Figure 3.1.
The chamber occupies 24’x25’x11’ of volume, and the walls are treated with fiberglass fill to provide a lower cutoff frequency of ~500 Hz. Due to the sizing of the model rig, this frequency is well below the characteristic peak frequency at all
Exhaust Wall Control Room
FF Mic Array 70º
150º
180º
24’
Figure 3.1 - a) Schematic of Aeroacoustic Test Facility far-field microphone array, b) Nozzle
Acoustic Test Rig with chevron nozzle installed during LDV backscatter measurements directivity angles, and provides adequate headroom for determining low frequency noise reduction characteristics. The far-field (FF) microphone array is also
indicated, at a distance of about 35 Deq , sufficient for far-field measurements for this rig and similar to others [107]. Acoustic tiles similar in construction to the walls are used to cover the floor during FF measurements, and removed to facilitate other
53 instrumentation. The exhaust wall is also treated with batting to prevent noise. The ambient conditions in the room are monitored on a low frequency data system to provide time histories and average conditions relevant for long and short measurements. Relative humidity, pressure and temperature are measured so that absorption characteristics can be corrected for via Shields and Bass method [108].
The air supply system and controls were designed to provide heated air, up to
350°F, by a local heat exchanger on the ceiling of the room to reduce heat loss. The core stream plenum is safely operated up to 26psig, for a core Mach number of 1.3.
The bypass stream was designed for a Mach number up to sonic, delivered at
“ambient” temperature, usually about 75ºF. Further details of the general facility capabilities and the acoustic characteristics of anechoic chamber can be found in
109 and 110.
3.2 ATF High Bypass Model Rig & Hardware
Several versions of model hardware were examined on the model turbofan NATR, simulating a modern separate flow exhaust system of bypass ratio 8. A cross- section of the assembled rig is shown in Figure 3.2. Major features shown are the interchangeable conical core and fan nozzles, large contraction sections curved with fifth order polynomials, and the centerbody-supported plug in the core nozzle.
Nozzle curves are also fitted with fifth order polynomials.
54 Fan Supply Ports (4) Nozzle Detail Flow Conditioners NACA 0012 Supports Probe Access Ports Core Nozzle
Fan Cowl
Core Cowl
Core Muffler Plug Core Contraction Fan Contraction Fan Nozzle
Figure 3.2 - Cutaway view & components of high bypass ratio (8) Nozzle Acoustic Test Rig
Upstream of the core stream pipe spool connecting the supply air to the core plenum, a pipe noise muffler is installed and was found to greatly improve the pipe noise emanating from the core nozzle. Downstream of the spool, the flow enters a conditioning segment composed of a honeycomb and then a wire distribution cone and fine mesh screen. The fan supply air enters the plenum surrounding the core hardware at the four quadrants to achieve high mass flows. These are similarly
Table 3.1 - NATR exit plane circular geometry for bypass ratio 8 baseline hardware
Hardware Element Diameter Circular Area Equivalent Area Annulus Height Deq Coaxial Deq Core 3.03 7.19 3.51 0.43 2.11 4.67 Fan 6.04 28.65 13.63 0.83 4.17 Plug 2.16 3.67 - - - - Core Cowl 4.37 15.02 - - - - located upstream of conditioning elements, before entering the settling chamber.
Contraction section geometry is also included in Table 3.1.
55 The fan and core stream plenums are instrumented with mechanical transducers to measure, at redundant points, the static pressure and temperature for setpoint calculation at a sampling rate of 4Hz. Measured values of stream conditions, specifically average temperature and velocity across the potential flow region at the nozzle exit plane, were previously measured and found to agree within a few percent of isentropic calculations. Similar results were calculated in data acquired in this study; see for example, Sections 4.1 and 5.1, for exit plane radial traverse pitot and LDV results respectively.
3.2.1 Conventional Core and Fan Nozzles
The nozzles used in this test have been successfully implemented on the NATR for both subsonic and supersonic noise reduction on the bypass ratio 8 hardware used in this investigation [48, 111, 112]. Four core stream conventional chevron nozzles were tested, though not for all studies involved in this thesis: 1) a baseline, conic nozzle (conic in the last significant portion of the interior contour), 2) an 8-lobe, low-penetration chevron nozzle with right angled tips, 3) an 8-lobe low-penetration sinusoidal nozzle, and 4) a high-penetration chevron nozzle with right angled tips a) b) c) d)
Figure 3.3 - Core and fan stream conventional nozzles included in testing configurations. a) baseline conic core nozzle, b) 8LP core nozzle, c) 8LP sin core nozzle, d) 16HP fan nozzle
56 (not shown, similar to Figure 3.3b). Penetration is defined as the distance perpendicular to a linear extension of the nozzle contour at the chevron tip. Two fan stream nozzles were tested, including 1) a baseline, conic nozzle, and 2) a 16- lobe high penetration chevron nozzle with right angled tips. See Figure 3.3.
3.2.2 Modified Conventional Core Nozzles
Conventional chevrons were modified in several variations, in light of general acoustic scaling conventions and TKE mappings from previous works, showing high turbulence levels very near the nozzle exit in the laterally-directed flow (a.k.a. jetlet) centered azimuthally at the chevron root. For the 8-lobe chevron, several changes to the root geometry were made. These are shown in Figure 3.4, showing a sinusoidal root of 0.50” radius, and the same chevron nozzle with a 4mm CVG
a) b) R 1”
0.25”
Figure 3.4 - Modified 8HP core chevron nozzle with sinusoidal root a) nozzle b) dimensions positioned at each root. See Section 3.2.3 for a description of the CVG geometry and scaling. The high levels of turbulence thought to be generated by the small geometrical length scales of the root geometry, forming the lateral outflow characteristics, was reorganized into a potentially superior mixing tradeoff by these variations. The effect on acoustics is presented in Chapter 6.
57
3.2.3 Vortex Generators on Core and Fan Nozzles
The most extensive far acoustic field investigation considered a type of VG used in limited experimental configurations previously published [94, 99], and used in previous jet noise experiments for initial characterization of the devices in limited extent for single flow conditions. The simple CVG geometry is shown in Figure a) DVG b) MVG c) d)
0.01” shelf
Figure 3.5 - CVG geometry definitions and model rig installation. a) delta CVG (DVG), b)
mushroom CVG (MVG), c) CVG nozzle section (none installed), d) seat for CVG’s
3.5, in two different variations, depending on direction of the induced lateral velocity in the central region of the counter-rotating vortex pair. In the case of
Figure 3.5a, the coupled delta vortex generators are positioned such that the resultant induced velocity has a downward orientation. The counter-rotating structures act to produce a high outward radial velocity component in the region between, so that fluid from the high-speed flow is distributed into surrounding fluid.
Conversely, mushroom VG’s, so named for the opposite vorticity orientation producing a cross-flow velocity field similar to that shape, acts to entrain surrounding fluid into the primary stream. The difference between the two, for lower Reynolds number flows, is previously explained [99]. Variations of these
58 basic geometries, including stream penetration, number of CVG’s per nozzle, sweep angle, positioning with respect to the nozzle exit, are examined directly on the far field acoustics.
The nomenclature system employed is as follows:
{number of CVG’s}{location}{design}VG{secondary design}-{fan nozzle…}
An example, for 12 DVG’s evenly distributed around the nozzle azimuth installed on the core nozzle, with the baseline conic fan nozzle: 12DVG-Base. As another example, 10 half delta VG’s installed on the exterior of the core nozzle operated in single flow conditions: 10eHVG. Since not all hardware was tested in every study, the complete listing of the CVG geometries (and others) used in each will be made per chapter.
Another nozzle concept examined involved a core nozzle geometry similar to chevrons but located well upstream of the nozzle trailing edge. This would allow for greater flexibility in implementation of SMA actuation in the volume occupied by the nozzle body. The test nozzle was fabricated by creating 8 equally spaced cutouts in the shape of the chevrons, and then a milled seat for a ring of chevrons was used with small welds for attachment. Figure 3.6 shows the internal and external surfaces near the nozzle lip, with the adjustable chevron tabs, positioned with the downstream chevron vertex about 1/8” upstream of the nozzle exit plane. a ) b) c)
Figure 3.6 - Internal chevron core nozzle. a) internal surface, b) external surface, b) covered 59 The penetration of the chevrons was manually bent until the desired level was attained, while the nozzle was heated and installed on the rig. The penetration values were measured with a caliper before and after a change in penetration level.
Due to previous measurements with this nozzle indicating high levels of high- frequency noise generation, the outer surface of the nozzle was covered to compensate for the manufacturing method, which left a gap surrounding each chevron that was found to be the main source. The covering, composed of 0.005” steel shim stock, and surrounded with aluminum tape, produced slight protuberances in the outer nozzle surface.
3.3 Acoustic and Fluid Measurement Systems
The physical arrangement and capabilities of each of the measurement systems employed, and the associated computer DAQ and pre-processing, up to the recording of the raw data file are outlined in this section.
3.3.1 Acoustic Far-Field Microphone Array
In the anechoic chamber, 8 B&K #4939, ¼” free-field pre-polarised condenser microphones are located along a circular arc, approximately 35 equivalent diameters from the vertical centerline of the bypass nozzle exit plane. The microphone has a very good pressure-field response up to 80kHz (acceptably up to 100kHz), and for operational safety, are routinely used with the gridcap installed, for which a frequency correction is applied in post-processing in addition to the free-field
60 correction and microphone response correction. The diaphragms are matched with a
B&K pre-amp #2670. A single point frequency calibration at 1000Hz at 94dB was performed and recorded before each set of measurements, and used in shifting the entire frequency calibration curve up or down. The history of the single frequency calibration is tracked over the life of the microphone to ensure data quality.
Further, a factory re-calibration of the microphones was performed during the data acquisition, which established a high degree of confidence in the measurement system. During data acquisition, the microphone is pre-positioned pointing at the tip of the plug, where most high-frequency noise originates. This is due to the polar response of the microphone which is much more sensitive at high frequencies. The microphones are mounted on thin, adjustable, but permanently affixed holders that are positioned to avoid reflection and scattering, which were inflicted upon the earliest data collected at the ATF.
Two B&K Nexus 2692 conditioning amplifiers, with low-pass filters at
100kHz, pass the eight channels of analog data to an NI data acquisition system, coupled with Labview, that samples the channels at 200kHz. Ten seconds of high bandwidth acoustic and corresponding low data-rate are recorded and stored locally.
After each measurement set, several key variables are considered for determination of the goodness of data. A typical set of acoustic data used in this investigation has a high flow accuracy in pressure ratio of less than 0.15% deviation, and RMS error less than 0.4% for the core stream, and about double that for the fan stream, which is relatively much more unsteady. Temperature is extremely steady, and typically
61 the temperature was within 2° of setpoint. Thus, no flow or thrust corrections were applied to acoustic data, but were re-evaluated to ensure consistency.
All data contained in this thesis are presented in relative scales, showing the levels, generally in dB in reference to 20µPa, with no absolute datum. The data for each plot is offset at a unique datum. This should be considered when comparing upstream and downstream acoustic data, for example, that may show very different features on the same scale, yet may appear similar when plotted independently.
3.3.2 Boundary Layer Total Pressure Pitot Measurements
A small head section total pressure probe was employed to measure the issuing velocity profiles at the jet core stream exit plane. Figure 3.7 describes the geometry of the probe. The ratio of the smaller dimension of the probe orifice to the core stream annulus height is ~17. Due to the slender, fragile nature of the probe, it was first tested for suitability to the high-subsonic hot stream conditions to determine adequacy in strength for resisting flow deflection, and also dynamic response to forcing by disturbances in the flow. The probe was found to be suitable at the exit plane of the core nozzle, and across the shear layer into the fan stream, but fluttered in downstream shear layers. A
Validyne Model DP15-46 differential pressure transducer was calibrated up to sonic conditions Figure 3.7 - Boundary Layer Total Pressure using a Druck Model DPI 510 Probe, Physical Dimensions: A) 12" B) 11" D)
0.25" F) 0.650" M) 0.120"
62 pressure calibrator. A 2-dimensional traverse was used to collect data at 500Hz in a sweep mode, continuously moving the probe at a very slow rate to ensure probe stability and provide enough data to provide good spatial averaging, about 1/8” per second. Electrical continuity using a small LED light helped dock the probe to the surface of the test rig, and the extent of the measurement range was through the inner fan shear layer. The fan exit plane was characterized separately since it is a separate flow nozzle design.
3.3.3 Dual-Component Laser Doppler Velocimetry
A multiline Ar+ CW laser, with a maximum power output of 400mW was used to provide the signal for dual-component backscatter LDV measurements. The maximum safe level considering temperature limits was used, about 380mW. The power output ratio of the green (514.5nm) to blue (488nm) line at the probe output was about 1.5, and determined the green interference fringes for measuring the axial velocity, while blue was used for radial velocity. The Dantec Fiberflow 300mm focal length backscatter probe was aligned carefully using a microscopic lens, backlit with the violet laser line through the multiple mode receiving optical fiber, so that both planar interference fringes would be formed, and axial/radial signal coherence would be maximized. The probe had an expanded beam diameter of about 1.35mm, which combined with the beam separation, wavelengths and probe focal length, determine the optical probe geometry listed in Table 3.2, calculated by standard geometrical methods [113]. As the probe length to height ratio is about 16, the probe was traversed vertically, in the plane normal to the probe axis to minimize
63 non-linear errors in flow gradients. A 40MHz frequency shift of one of each the green and blue beams was accomplished with a Bragg cell to eliminate velocity ambiguity possible in the radial direction. Beam separation was set at a maximum distance of 38mm to reduce the probe volume and increase intensity of the fringes
Table 3.2 - Backscatter dual-component LDV measurement volume geometrical parameters
Beam Probe *Probe Fringe Laser Half-Angle Wavelength Probe Width, Length Volume Fringe Separation
Component (deg) (nm) δx or δz (mm) (mm) (mm3) Count (µm) Axial (green) 514.5 0.150 2.45 0.23 4.2 3.51 36 Radial (blue) 488 0.143 2.33 0.20 4.0 *prolate spheroid approximation and the fringe spacing for high-speed measurements.
Processing of the variable input current PMT transducer voltage signal was accomplished with a Dantec BSA P80 real-time burst spectrum analyzer operated in coherence windowing mode between the two data channels. Measurements in the potential core agreed well with previous measurements, and PIV, within about 3% deficit of the isentropic expected velocity. The system power combined with the seeding system provided a data rate, variable throughout the flow, of up to almost
8kHz. In shear layers near the nozzle exit, where the seeding was less well mixed, the data rate was consequently lower. Traverses of the probe were made throughout the flowfield along radial vectors, and along the jet centerline. At the centerline, even step increments of ¼” were chosen to provide high precision, and along the radii, uneven step increments were used to increase data density in the shear layer.
The estimated uncertainty in the velocity measurement was about 1 m/s for the axial velocity, and 0.4 m/s for the radial velocity.
64 A 3D traverse code was developed in LabView to traverse the probe over arbitrary grids with increased resolution in the shear layers. Triggering the LDV acquisition via TTL line output, and synchronizing flow control parameter data collection allowed mean scaling corrections. A screenshot of the user interface in is shown in Figure 3.8. The grid could be set using even steps, or formatted text files could be chosen to define the grid spacing. Measurement status was monitored in real-time since sample times of 2 seconds at each point required long run times.
Figure 3.8 - Screenshot of LDV 3D traverse and trigger sync program interface
65 3.3.3.1 Flow Tracker Particle Seeding
Oil droplet seeding of the flow was accomplished via Laskin nozzle olive oil seeders and swirl chambers. The mean particle size output from the device was about 1µm, about twice the wavelength of the green laser, and thus exhibited Mie scattering character. Figure1a shows the oil seeder array, a histogram of previously
PDPA measured particle sizing [109] at ~5psig differential pressure across the nozzle, and a Mie scattering intensity plot for reference. Due to operation in
90 120 60
150 30
180 0
210 330
240 300 270
Figure 3.9 – a) Laskin seed units, b) oil particle diameter data, c) polar log Mie intensity plot backscatter mode for LDV, a large tradeoff of intensity of light is made for simplicity of traversing the probe and alignment accuracy. In the ATF, the oil was injected into up to three ports in the fan plenum and one in the core stream plenum well upstream of the nozzle (and upstream of the flow conditioning elements) to encourage mixing. The injectors in the fan stream were simply injection orifaces along the plenum sidewall, whereas in the core stream a seeding bar was used with
1/8” holes along the length of the bar to disperse the seed more evenly since only one port was available on the core.
66 Chapter 4 Boundary Layer Flow and Noise Results
To characterize the boundary layer mean characteristics, pitot measurements using a boundary layer probe were carried out at the exit plane of a baseline conic nozzle, and two chevron nozzles. One of the chevrons was the 8LP, and the other was the
8LPsin nozzle. To assess the effects of varying the initial momentum thickness on acoustics, both the core outer and fan inner boundary layers surrounding the core nozzle surface were tripped independently and also simultaneously to increase the momentum thickness at the exit plane. Besides extending the knowledge of the effect of initial conditions to a coaxial jet, if the far acoustic field is insensitive to changes in θ for the baseline nozzle, then any active noise control methods that impose small disturbances to the flow when turned off may probably be negligible.
Chevrons and other mixing nozzles would desirably be insensitive to increased θ to simplify application and design.
The wire trips and orifice ring Round increased the momentum thickness of the Wire Trip core boundary layer up to 58%, and up to
88% in the fan stream. Due to the Alternating converging exit contour in the core nozzle, Flip-Tab Ring and the diverging wall on the outer face of (plug shown removed for clarity) the core nozzle, the trips were not equally Figure 4.1 – 3D slice view of boundary
layer trip installations on both streams.
67 Table 4.1 – Designations, dimensions, and scaling parameters of core and fan trips.
Nominal Size
Trip Designation Application δt (in) δt/D (%) IM Inside Medium Core 0.063 1.8 IH Inside High Core 0.125 3.6 OM Outer Medium Fan 0.063 1.1 OH Outer High Fan 0.100 1.7 effective, despite the large physical size of the flip-tab ring used in the core stream.
Thus, the trips used for the core stream of annulus height only 54% of the fan annulus were actually about 67% and 100% larger than those in the fan stream based on the flow scaling. Figure 4.1 shows the positions where the trips were installed at the core nozzle interface. Table 4.1 details the sizes of trips used, and the relevant scaling parameters. The core stream of the coaxial flow was heated to 220ºF, with the 1.423 CPR creating ideal nozzle velocity of 931 ft/s by isentropic relations. The
FPR was at 1.258 for a fan velocity of 663 ft/s and a shear velocity across the streams of 268 ft/s, and ratio of Up1/Us1 of 1.40.
4.1 Core Baseline and Chevron Nozzles Mean Initial Velocity Profiles - Clean and Tripped
A sample repeatability test of the unscaled raw pressure profile in Figure 4.3 indicates repeatability within the shear layer, which was very good in this instance shown. The x-axis represents the number of samples collected as the computer- controlled traverse scanned the probe from the centerbody through the fan stream.
Due to this method though, of continually moving the probe, variations in plenum pressure caused the small variations across a single sweep. Temperature changed
68 negligably during the measurements. 1.2
Baseline 1 1.0 Multiple runs were used to smooth out Baseline 2 0.8 CORE the inconsistencies and provide a very 0.6
0.4 dense amount of data points across the FAN 0.2
0.0 Normalized Transducer Output Transducer Normalized shear layer, which contains 6000 10000 14000 18000 22000 26000 30000 34000 -0.2 approximately 230 points from the 97% No. Samples velocity point to the lowest point, for Figure 4.3 – Repeated pitot measurements of baseline nozzle boundary layer for raw data each individual sweep. Data were normalized by max()U , and adjusted radially using the boundary layer data along the plug to account for probe positioning errors, and were then averaged. This showed the general repeatability as well as the
U (ft/s) similarity of the profiles for small deviations in 900 800 set point (<1% U). 700 600 500 The trips applied to the core stream had 400 300 the effect of increasing the momentum thickness 200
100 r-rp (in.) by acting over a large portion of the jet annulus. 0 0.00 0.10 0.20 0.30 0.40 0.50 Defect in velocity extending well into the core Figure 4.2 - Axial velocity profiles stream is clearly visible in Figure 4.2, showing a along radial traverse for baseline comparison of the untripped velocity profile of and high trip configurations. core nozzle exit plane. The core traverse is from Single jet, cold flow. a radial position at a datum of zero at the plug surface, outward in radius until the edge of the nozzle lip at about 0.43”. A profile of the fan velocity, with an OM
(Outer Medium) trip applied, and in the clean configuration is shown in Figure 4.4.
69 Three configurations are actually shown, U (ft/s) 800 two being with the same outer trip 700
(OM), but with differing inner trips (IM, 600 500 IH), yet show very good agreement, 400 IHOM 300 IMOM which should be expected. 200 CLEAN
Probe effects (the finite head 100 r-rp (in.) 0 dimension) limit the useful range of 0.40 0.50 0.60 0.70 data, especially against the nozzle plug Figure 4.4 - Fan velocity profile at core exit surface, but in the outer core shear layer plane for baseline core nozzle without (clean) and with (IMOM, IHOM) trips of interest, the velocity is unaffected until very close to the nozzle lip, about 50% of the probe width, since the probe was positioned downstream of the nozzle lip by about 5 mils. Since the boundary layer quantities of displacement thickness and momentum thickness are given by,
r=δ ⎛ ⎞ r=δ ⎛ ⎞ δ = − u θ = u − u 1 ∫ ⎜1 ⎟dr , ∫ ⎜1 ⎟dr (4.1), (4.2) 0 ⎝ U ⎠ 0 U ⎝ U ⎠ then the differences in neglecting the outer portions of the boundary layer, where u/U is small, will have small effects on
u/U(1-u/U) momentum thickness. To improve the 1.0
δ 0.8 calculation of 1 , the boundary layer was 0.6 measured as shown in single flow conditions 0.4 when the probe is most steady in the shear layer 0.2
0.0 to verify the linearity of the shear layer very 0.00 0.10 0.20 0.30 0.40 0.50 Radial Distance, in. near the wall. Thus, deviations in linearity close Figure 4.5 – Normalized momentum
deficit, core flow, baseline nozzle.
70 to the wall due to probe interference within about ~50% of the face height of the probe to the wall, are discarded and replaced with a linear extrapolation. To confirm this procedure, and especially the effect on the momentum thickness, the non-dimensional momentum defect across the shear layer was plotted versus the radial distance in Figure 4.5. As the outer core nozzle wall is approached, it is easily seen that sensitivity to displacement and momentum thickness based on velocity errors near the wall will be small, at most du/U according to (4.1).
The properties of the baseline conic nozzle calculated as described are stated in Table 4.2. The inner boundary layer is very thin, about 1.3% of the core stream
Table 4.2 - Inner and outer baseline conic core nozzle exit boundary layer characteristics
Shear Layer δ (0.97Ue) δ (%Hann) δ* (in.) δ* (%Hann) θ (in.) θ (%Hann) H Reθ Inner (Core) 0.038 8.8% 0.011 2.4% 0.0059 1.4% 1.78 2740 Outer (Fan) 0.180 22.8% 0.033 4.2% 0.0232 2.9% 1.44 8324
annulus height by momentum thickness. Comparing the shape factor to a value of
2.7 for a Blasius layer, the profiles have a greater amount of axial momentum near the wall, as should be expected. It is also seen that the probe width is about 30% of the boundary layer thickness, but in contrast to the situation where the probe approaches the face of a surface, sweeping the probe past the nozzle lip reduced the interference effects near the wall considerably.
The reduced boundary layer data for the untripped and tripped baseline conic,
8LP “Chevron” and sinusoidal 8LP “Sine” core nozzles are tabulated in Table 4.3 for the inner shear layer and Table 3.1 for the outer shear layer. Data does not exist for the for the chevrons in the clean untripped configuration, due to a data loss issue
71 imposed by a faulty data cable, but the characteristics should be similar to the sinusoidal chevron, and the existing data confirm this.
Table 4.3 - Inner shear layer characteristics in tripped configurations and clean reference
INNER SHEAR LAYER
δ1 θ
Nozzle Trip δ1 (in.) (%Hann) θ (in.) θ/θ0 (%Hann) H Baseline Clean 0.0099 2.3% 0.0059 1.00* 1.4% 1.68 Medium 0.0114 2.7% 0.0074 1.25 1.7% 1.55 High 0.0143 3.3% 0.0096 1.63 2.2% 1.49 Sine Root Clean 0.0122 2.8% 0.0064 1.00 1.5% 1.90 Medium 0.0168 3.9% 0.0098 1.53 2.3% 1.72 High 0.0174 4.1% 0.0117 1.84 2.7% 1.48 Chevron Root Medium 0.0138 3.2% 0.0092 1.44 2.1% 1.51 High 0.0154 3.6% 0.0109 1.71 2.5% 1.40
* Reference value for baseline only. Sine nozzle initial momentum thickness used otherwise.
Table 4.4 - Outer shear layer characteristics in tripped configurations and clean reference
OUTER SHEAR LAYER
δ1
Nozzle Trip δ1 (in.) (%Hann) θ (in.) θ/θ0 θ (%Hann) H Baseline Clean 0.0310 3.9% 0.0224 1.00* 5.2% 1.39 Medium 0.0382 4.8% 0.0382 1.71 8.9% 1.00 High 0.0545 6.9% 0.0420 1.88 9.8% 1.30 8LP Root Clean 0.0356 4.5% 0.0228 1.00 5.3% 1.56 Medium 0.0621 7.9% 0.0441 1.93 10.3% 1.41 High 0.0682 8.6% 0.0487 2.13 11.3% 1.40 8LPsin Root Medium 0.0575 7.3% 0.0425 1.86 9.9% 1.35 High 0.0590 7.5% 0.0445 1.95 10.3% 1.33
Because of the shape of the Sine and Chevron nozzles, the data was measured located axially at the root of each chevron, which imposed a slight axial offset between the absolute locations versus the baseline geometry, in the upstream
δ θ direction. This probably explains the slightly higher values of 1 and calculated for the Chevron and Sine nozzles compared to the baseline.
72 A plot of the relative variations in momentum thickness obtained is shown in
Figure 4.6. The inner shear layer exhibits good modulation with the trips, accruing nearly a linear increase in momentum thickness proceeding from the medium to the high trips. The fan layer was modulated nearly to the same value for either trip, besides a small increase. This is because, referring to Figure 4.4, the structure of the boundary layer was composed of two linear ranges, one in the viscous sublayer, one associated with the divergence of the wall. The second linear range was affected greatly by tripping since the outer nozzle surface is highly divergent, the half cone angle of the outer nozzle surface being about 25 degrees. This created a large initial change in momentum thickness, even with moderate variation in shape factor, and subsequent smaller increase in momentum thickness associated with virtually no change in the shape factor.
a )b)
2.0 2.5
2.0 1.5
1.5 1.0 1.0
0.5 0.5 Chev Sine 0.0 0.0 Base Clean Clean Medium Medium High High
Figure 4.6 - Relative change in θ for all nozzles and trips. a) inner boundary layer, b) outer
73 4.2 Acoustic Far-Field Sensitivity to Boundary Layer Characteristics
The clean untripped acoustic spectrum in third octave band are shown in Figure 4.7 at 70° and 150° degrees to the jet axis. Corresponding narrowband data are shown
Base Clean Sine Clean Chev Clean 85 95
90 rk rk a 80 a M M r 85r e e B P B P d 75 80d B)5 - B) - 5 d 75d 70 PL ( PL ( S 70S Frequency (Hz) Frequency (Hz) 65 65 100 1000 10000 100000 100 1000 10000 100000
Figure 4.7 – T.O.B. far-field SPL spectra, clean configurations. a) 70°, b) 150° in Figure 4.8. Levels are relative and on each scale is independent, typical for all acoustics plots. A high degree of similarity is found between the spectrum of the baseline, Sine and Chevron nozzles, which show only a slight decrease at low
80 95 75 90
rk 85 rk a 70 a
M 80 M r r e 65 75e B P 70B P d 60 d 65 B)5 - 55 60B) - 5 d d 50 55 PL ( PL PL ( PL S 50S 45 Frequency (Hz) 45 Frequency (Hz) 40 40 100 1000 10000 100000 100 1000 10000 100000
Figure 4.8 – Narrowband far-field SPL spectra, clean configurations. a) 70°, b) 150°
74 frequencies, almost not measurable, and a very slight increase in noise at high frequencies above the baseline nozzle.
The cycle point chosen corresponds to a fairly low shear velocity (in terms of chevron nozzle effectiveness), and it is seen that only small differences in the TOB and narrowband spectrum can be found between configurations so that the baseline configuration may be used for all comparisons with trips. The cycle point was selected based on similarity to an existing engine-scale test condition. Should the flow conditions change, however, it is almost certain the non-linear initial shear layer development in the highly three-dimensional flow region near the nozzle exit plane should exhibit a different sensitivity to the boundary layer. However, previous work has shown (and generally focused upon) that the greatest acoustic effect in modification of the initial condition occurs from the laminar to turbulent transition for single jets. Then these results first serve to extend the acoustic effect for changes in momentum thickness to that of a more realistic coaxial jet, with initially turbulent boundary layer. The second purpose is to obtain some knowledge of the degree of sensitivity. And another is to understand how small, (though these trips are relatively large) distributed perturbations to the boundary layers, on both sides of a nozzle, might influence the far acoustic field.
The sensitivity of the acoustics to the inner shear layer trips, IM and IH, is seen to be almost nil in Figure 4.9 for the baseline conic nozzle. That there are not only no tones present in the spectra, but also no change in the broadband character of the spectra, implies that even the large flip-tab trip does not create any self-noise, nor does it alter the momentum distribution enough to affect the jet in any
75 Base Clean Base IM Base IH 80 95 90 75 85 rk rk a 70a 80 M M r r e 65e 75 B P B P 70 d 60d 65 B)5 - 55 - 5 B) 60 d d 50 55 PL ( PL PL ( PL S S 50 45 Frequency (Hz) 45 Frequency (Hz) 40 40 100 1000 10000 100000 100 1000 10000 100000
Figure 4.9 - Narrowband far-field SPL spectra of clean, inner boundary layer medium trip
(IM), and high trip (IH) baseline nozzle. a) 70 º, b) 150º significant way. The initial transition region, which controls the growth of the azimuthal structures, must not be altered enough to affect the axial cascade of energy to smaller scale turbulence. This is distinct from lower Reynolds numbers
(104) where significant effects on the vortical development and axial saturation rate and acoustic source distribution, have been shown by LES simulations [42].
Similar character was seen for the Sine and Chevron nozzles, where almost identical spectrum were found regardless of the inner shear layer momentum
80 95 90 See 4.7 for 75 85 rk
rk symbols a 70a 80 M M r r e 65e 75 B P B P 70 d 60d 65 B) - 5 55B) - 5 60 d d 55 50 PL ( PL ( S S 50 45 Frequency (Hz) 45 Frequency (Hz) 40 40 100 1000 10000 100000 100 1000 10000 100000
Figure 4.10 - Narrowband far-field SPL spectra, base, Sine IH, and Chev IH. a) 70 º, b) 150º
76 thickness, throughout all directivity angles. Figure 4.10 shows the noise spectra of the high trips on these two nozzles compared to the baseline without trips.
Once trips were installed on the fan stream, however, discernable differences in the aeroacoustic properties of the jet were found. Figure 4.11 shows the approximate 9.5dB maximum increase in high-frequency content in a broadband 70° manner. The noise increases in the upstream direction, which is consistent 90° with both a refracted quadrupole source embedded within the inner layer of the coaxial stream, or possibly a self noise 110° component of the trip and the resulting acoustic scattering at the trailing edge of 130° the sharp nozzle lip. The fact that the source decays fairly constantly across the directivity range indicates this may be the 150° case [115, 116]. Based on the trip diameter, maximum fan velocity of about
630 ft/s and St = 0.13 (assuming small compressibility effects), approximately the 2D shedding frequency found for Figure 4.11 - Far-field spectra from 70º to recent trailing edge acoustic simulations 150º for IMOM trip nozzles.
77 of flow past a trailing edge with a blunt body upstream [116], the dimensional frequency corresponds to about 10.4kHz. This is certainly near the onset of the additional source, but noting that the general directivity pattern of a scattered trailing edge dipole varies as sin 2 θ 2 , a plot of the TOB directivity at 25kHz 76
(approx. peak difference) strongly shows 74 otherwise. Considering that the high trips 72 B) d (
L placed on the inside boundary layer yield P
S 70 no noticeable additional source 68 components, tonal or broadband, indicates Directivity Angle (Degrees) 66 60 80 100 120 140 160 that surface or trailing edge dipoles are 4.3 - Directivity of 25kHz TOB unlikely the cause. Unless the geometry of component of IMOM trip nozzle the nozzle coupled with mean flow refraction effects could modify the dipole directivity that strongly, this phenomenon seems entirely related to the jet aeroacoustics, in which case additional measurements of the fluctuating velocities at the nozzle lip would be necessary to verify the nature of the source.
The large trips seem to, verify that the source is indeed a broadband quadrupole one, imposed by the stream turbulence downstream of the mixing nozzle. Figure 4.12 shows the spectra for the IHOH configurations at a range of directivity angles. Analysis of the OH configuration, which had no trips installed on the core stream, appeared similar to the IHOH baseline or chevron nozzles, except
78 for a small reduction in the noise at very high frequencies about 60kHz across all nozzles.
The arrows in the figure indicate two small near-tonal components, centered at frequencies of 16.4kHz and 23.4kHz. These correspond to St = 0.21 and
St = 0.30 respectively. These values are somewhat higher than a value of 0.13, but for a higher Reynolds number flow, a more appropriate velocity would be 70° the convection velocity of the trailing 70° structures, not the maximum stream 90° velocity [116]. In this case the Strouhal 90° number would be increased from 0.13 to
0.20, based on a convection velocity of 110°
0..65U This is well in line with the 110° prominent peak in the data, and seems to 130° indicate that only the small portion of 130° the energy in that tonal “bump” is 150° directly related to the trip self-noise.
Further, that the peak directivity is 150° slightly aft of 90°, also is fairly well in line with the normal of (the majority of) the nozzle surface, which points along Figure 4.12 - Far-field spectra from 70º to the 115° degree directivity vector. 150º for IHOH trip nozzles.
79 Chapter 5 Initial Mixing Region Turbulence via LDV
Investigation of some of the turbulence statistics in the jet under the influence of various mixing nozzles was accomplished using the dual component backscatter
LDV setup. A non-intrusive measurement was desired that provided high accuracy, with reasonable temporal resolution to provide characteristics of the turbulence and for LES validation purposes as well. An extensive amount of acoustic data is available at the simulated engine cycle point used in this test, which drove the decision to focus on a coaxial flow for which the chevrons were effective in reducing noise. The heated core flow of pressure ratio 1.80, and fan flow pressure ratio of 1.26, corresponded to Mach numbers of 0.96 and 0.58 respectively.
Velocities were 1224 ft/s and 663 ft/s, with a shear of 561 ft/s.
Core nozzles used in this study were the baseline conic nozzle, the conventional 8HP chevron, and some initial data was collected for the 12DVG nozzle on the jet centerline. Two radial Y/D1 traverses were made in the development region for both the conic and 8HP nozzle, to 5.80 starting at the nozzle exit plane and at two Z/Deq other axial locations with the last at the
tip of the core nozzle plug, as shown in 0.88 1.46
Figure 5.1. The locations corresponded to Figure 5.1 - LDV probe traverse locations
80 existing single flow PIV data for the three configurations, and extended from the core plug surface and through the fan stream to the outer shear layer. Note that the axial coordinate is normalized by Deq, while the transverse (radial) coordinate is normalized by D1. Azimuthally, the traverse was aligned with the root vertex of the chevron in one case, and with the tip of the chevron in a second case.
5.1 Mean Results
The centerline velocities for the three nozzles, normalized by the maximum velocity in the potential core U1 are shown in Figure 5.2. Each point represents the
U/U1 0.65
0.55
0.45 Base Chev 12DVG
X/Deq 0.35 1.5 2.0 2.5 3.0 3.5 4.0 4.5 5.0 5.5 6.0
Figure 5.2 – Centerline U/U1 velocity for baseline, chevron and 12DVG nozzles mean value of approximately 20,000 samples at an average rate of nearly 6500 Hz, and a high point density was intended to produce smooth spectral turbulence data.
The core stream mixer nozzles increase largely the lateral flux of axial momentum very early into the downstream development, resulting in a far-reaching reduction in centerline velocity, which has been well documented. The gross effect is similar for both the chevron and the 12DVG nozzles upstream of approximately Z/Deq=4.5.
81 The relative reduction in normalized centerline velocity in departure from the baseline conic nozzle is also shown annotated in Figure 5.3. The peak reduction is about 9.5% for each mixer, and occurs at roughly the same axial position, although the high-penetration chevron shows slightly increased effects upstream. The difference curves share a common trend until the chevron nozzle begins to asymptote slowly toward the baseline conic datum. At this point, the centerline velocity quickly reverts for the 12DVG nozzle, approaching or slightly increasing over the conic nozzle in only about 0.5Deq.
∆U/U 0.111 C1 0.08 C2
0.05 V2 0.02 V1 X/Deq V3 -0.01 1.5 2.5 3.5 4.5 5.5
Figure 5.3 - Centerline reduction in U/U1 with baseline nozzle datum for mixer nozzles. C1,
C1, and V1, V2, and V3 indicate regions of distinct mixing behavior for Chevrons and VG’s
In approximating the overall structure of the curves for the chevrons and the
12DVG nozzle, distinct zones are indicated in Figure 5.3 by the curve segments. C1 marks the initial nonlinear development of the flows between the chevrons.
Ejecting large amounts of axial momentum away from the centerline, the centerline velocity increases nearly monotonically until an abrupt change near Z/Deq=3.2. Past this point, the chevron centerline velocity assumes nearly the same form as the baseline nozzle. The appearance of the long approximately linear velocity
82 difference indicated by C2 in Figure 5.3 suggests that the virtual origin of the remaining portion of the jet has simply been shifted by the chevron flows. This bolsters the idea that most of the reduction in noise by chevrons through overall velocity reduction is accomplished through bulk momentum transfer, and not the vorticity mechanism. This can be explained simply by noting that the chevron centerline velocity exhibits large changes that were vorticity the primary mechanism, probably would not induce such effects. Further, the similarity in the downstream of the chevron and baseline velocities renders the notion that all of the laterally ejected momentum merges back together and then develops into a free flow that appears as though it had originated much farther upstream. The three regions shown V1, V2 and V3, in the case of the CVG nozzle however, indicates a much different process, dominated by much slower changes in axial velocity. V1 may correspond to the growth period of the vortices, V2 to the decay of the transverse gradient or merging of the individual structures, and V3 to the fully developed flow, downstream of any rapid decay in mixing due to vortex decay. It is clear that the processes which decrease the centerline velocity appear much different for a chevron and the CVG devices, however, more data for strongly vortical mixing devices should be compared to support these findings. Although great care was taken to repeat the baseline cases and increase the temporal frequency resolution of the measurements, the nature of the jet mixing was not investigated at other
Reynolds numbers or shear velocities between the two streams.
Explanation for the relative importance of the two sources of centerline velocity decay may be found in the fundamental differences in the mixing
83 mechanisms between the two nozzles. For chevrons, the large lateral flux of axial momentum seen in varying degrees of chevron penetration appears to have sheets of
(largely axially-oriented) vorticity shedding from the sharp edges that encompass the lateral flow [117]. Thus, the chevron cutout might be regarded as both trading mean axial momentum directly for lateral momentum through flow “escaping” through the root area, and also in the form of axial vorticity, as a superimposed secondary flow. The efficiency of this trade is fairly good since vectoring the axial momentum as cosine leads to an addition in lateral momentum varying as sine, while the vorticity generation depends on the local pressure gradients.
In contrast, the CVG’s use the “pressure-hill” upstream of the devices to add larger-scale vorticity, or deflection, to the flow, which is then accelerated through the central split of the device and creates high levels of smaller-scale vorticity [85].
The crucial difference seems to be that the main source of mixing in the chevrons is
V/U1
0.10
0.05
0.00 1.5 2.5 3.5X/Deq 4.5 5.5
Figure 5.4 - Centerline V/U1 velocity for all nozzles. Average value of 1.5% (conic nozzle). the lateral transport of axial momentum by the jetlets, whereas the CVG’s probably use very energetic vorticity rather than bulk transport as the dominant mechanism of mixing. This will be illuminated further in Chapter 7, but could lead to the
84 departure in the mean centerline velocity decay between the mixer nozzles, since the two flow structures would have very different stability characteristics. A plot of the normalized radial velocity along the traverse confirms the true jet centerline was adequately located, since the mean velocity is nearly constant and small, especially in the downstream regions (where the centerline U decay rates differ), where transverse gradients diminish.
The radial mean axial velocity for the two axial locations Z/Deq of 0.88 and
1.46, corresponding to about 50% and 100% of the axial distance between the core
U/U1 U/U1 Base Chev Root Chev Tip 1.0 1.0
Z/Deq=0.88 Z/Deq=1.46
0.8 0.8
0.6 0.6
0.4 0.4
Y/D1 Y/D1 0.2 0.2 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0 0.00.10.20.30.40.50.60.70.8
Figure 5.5 – U/U1 velocity profiles along radial traverses at Z/D1 = 0.88 and 1.46. nozzle exit plane and the plug tip, are shown in Figure 5.5. The upstream location shows a lobe of high axial velocity projecting outward radially into the fan stream, with a very small associated decrease in core stream velocity. The ratio of the peak lobe velocity to the baseline velocity at the same radial location is about 4/3, a large increase in axial momentum in such a short downstream distance. Chevron tip data shows an essentially unmodified velocity in the potential core of the core stream,
85 relative to the baseline nozzle core flow, and decreased velocities across the shear layer. The downstream data is similar, with the extensive initial wake from the plug body extending nearly 0.10Deq radially. In the core stream, the peaks velocities are still comparable, with even the chevron root region showing little decrease in velocity.
5.2 Turbulence Statistics in Axial and Radial Velocities
Histograms of u, shown in Figure 5.6, indicate the unsteadiness in the wake region of the centerline measurements, by the broad width, and a highly non-Gaussian distribution which represents the large scale structures that amplify toward the end
potential core after Z/Deq = 3.11. The Z/Deq = 1.46 2500 u Z/Deq = 1.46 2000 location is directly after the plug, about 1/8” 1500
1000 downstream, and shows relatively low velocities, and 500 0 a broad but smooth distribution because of the large 2500 u Z/Deq = 2.01 2000 wake of smaller, less energetic turbulence scales, and 1500 1000 relatively low large-scale instability levels. The 500
0 slight misalignment of the probe volume with respect 2500 u Z/Deq = 3.11 2000 1500 to the centerline causes the asymmetry in the velocity 1000 distribution. At a very short distance downstream, 500 0 the histograms show large numbers of very high 2500 u Z/Deq = 5.75 2000 1500 velocity records that are much larger than the mean. 1000 500 After the end of the potential core the cluster of 0 0.0 0.1 0.3 0.4 0.5 0.6 0.7 0.9 1.0 1.1 Figure 5.6 – Histograms of u
at various Z/Deq locations on 86 jet centerline of conic nozzle. similar high-velocity measurements spreads out over a larger velocity range, signifying the transition into small-scale, broadband turbulent motion.
The histograms also serve as a check on the measurements, in terms of frequency filtering so that unwanted noise is not captured and included in the measurement. Bandpass signal filtering was performed prior to the measurements, and not as a CORE EXIT post-processing as some researchers have done [118]. Figure 5.7 – Core nozzle plug boundary layer seeding image Although the potential for velocity bias also exists, which the histograms cannot capture, due to unequal sampling of non- homogeneously seed particles in wake flow, that effect has been shown to be very small even in a high shear jet flow with good seeding. A seeding image from previous PIV measurements, Figure 5.7, confirms that seeding is adequate within the boundary layer (and throughout the entire core flow) along the plug, just downstream of the core nozzle, and so should remain well-dispersed along the jet centerline.
87 The histograms of u for the chevron and 12DVG nozzle are shown in Figure
5.8. It is apparent that very close to the nozzle plug, there is some centerline misalignment, the 12DVG the best of the three configurations. However, after a fairly short downstream distance, the widening of the jet alleviates the issue.
Despite careful alignment of the probe volume using a microscopic lens to identify the longitudinal bounding extents, this inconsistency must be taken into account when comparing the farthest upstream data turbulence levels and spectra. The physics of the large-scale structures and the cascade into more random turbulence, nevertheless, seem to be borne out by the velocity distributions.
The radial traverses are similarly useful to provide quantitative feel for the
Chevron 2500 u Z/Deq = 1.46 2500 u Z/Deq = 1.46 CVG 2000 2000 8HP 1500 1500 12DVG 1000 1000
500 500
0 0 2500 u Z/Deq = 2.01 2500 u Z/Deq = 2.01 2000 2000 1500 1500 1000 1000 500 500 0 0 2500 u Z/Deq = 3.11 2500 u Z/Deq = 3.11 2000 2000 1500 1500 1000 1000 500 500 0 0 2500 u Z/Deq = 5.75 2500 u Z/Deq = 5.75 2000 2000 1500 1500 1000 1000 500 500 0 0 0.0 0.1 0.3 0.4 0.5 0.6 0.7 0.9 1.0 1.1 0.0 0.1 0.3 0.4 0.5 0.6 0.7 0.9 1.0 1.1
Figure 5.8 - Histograms of u at various Z/Deq on centerline for chevron and 12DVG nozzles
88 makeup of the turbulence statistics, but are less facilitating in diagnosing spatial differences, since any misalignment of the probe in either the tangential direction
(longitudinal axis of the probe) or the axial direction are displaced across smaller gradients. Small changes should not help identify misalignments, but are also inherently less problematic. The u velocity histograms of the core flow potential core, the inner shear layer, the fan potential core, and the outer shear layer are presented in Figure 5.9 at Z/Deq = 1.46 (plug tip) for the baseline and chevron (at the root). The core stream potential core at Y/D1 = 0.17 shows very small velocity bandwidth in the time history, with an apparent symmetry, as expected. Along the next radial location in the shear layer, the histogram of the axial velocity appears as a broader distribution as expected, but a large asymmetry biased toward lower velocities is found in the chevron shear layer at the root. The velocities at the edge of the potential core of the fan stream just beyond the edge of the inner shear layer
(as would be apparent from the mean velocity) see unsteadiness from the chevron lobes. The outer shear layer shows essentially identical distributions for the two nozzles.
89 The axial RMS velocity fluctuations for the range of the centerline traverse are shown in Figure 5.10. All three nozzles show high turbulence levels, due to
12000 u Y/D1 = 0.17 12000 u Y/D1 = 0.17 10000 10000 8000 8000 6000 6000 4000 4000 Baseline 2000 2000 Chevron 0 0 Conic 5000 u Y/D1 = 0.33 5000 u Y/D1 = 0.33 8HP 4000 4000
3000 3000
2000 2000
1000 1000
0 0 12000 u Y/D1 = 0.53 12000 u Y/D1 = 0.53 10000 10000 8000 8000 6000 6000 4000 4000 2000 2000 0 0 5000 u Y/D1 = 0.73 5000 u Y/D1 = 0.73 4000 4000 3000 3000 2000 2000 1000 1000 0 0 0.2 0.3 0.4 0.5 0.6 0.7 0.8 1.0 1.1 0.1 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.1
Figure 5.9 - Histograms of u at various Y/D1 on Z/Deq=1.46 for baseline conic and 8HP nozzles. being in the plug wake. The baseline conic nozzle reaches a maximum turbulence intensity near Z/Deq=3.1 downstream of the fan nozzle at a level of 19.1%, while both mixer nozzles substantially increase the axial turbulence for the entire measurement range. The form of the two mixer nozzles is similar, but the 12DVG nozzle begins a slow decrease near the maximum location for the baseline nozzle, then increases again at a saddle point at 1 Deq downstream. Transverse turbulence levels are a fraction of even the minimum axial turbulence of the baseline nozzle,
90 u'RMS/U1 0.30
0.25
0.20
0.15
0.10 1.5 2.5 3.5X/Deq 4.5 5.5
Figure 5.10 – Axial turbulence intensity along jet centerline for all nozzles shown in Figure 5.11. Here, the chevron nozzle shows a small peak level near the nozzle exit, perhaps associated with the transverse small-scale unsteadiness of the jetlets, or the superimposed instabilities they support. It is interesting from a mixing perspective to note that all figures Figure 5.4, Figure 5.10, and Figure 5.11 indicate the small axial gradients for the vortex generators, with the centerline radial turbulence intensity varying only slowly and smoothly over the measured range compared to either of the other two nozzles.
v'RMS/U1 0.08
0.06
0.04
0.02 1.5 2.5 3.5X/Deq 4.5 5.5
Figure 5.11 - Transverse turbulence intensity along jet centerline for all nozzles
91 5.3 Turbulence Velocity Spectra
A prime motivation in using LDV is to resolve the time history to some significant extent, besides the inherent accuracy, and small optical probe dimensions and other important features particularly desirable in turbulent jets. Although backscatter
LDV reduces the amount of photons scattered to the receiving optics by orders of magnitude, Figure 5.12 shows that reasonable frequencies can be obtained. These
Hz
8000
6000
4000
2000 Base Chev 12DVG
X/Deq 0 1.5 2.0 2.5 3.0 3.5 4.0 4.5 5.0 5.5 6.0
Figure 5.12 – Backscatter LDV sample rates along jet centerline for all nozzles data were all measured using the same velocity filter width of about twice the width of the value of the mean value of the mean axial velocity along the centerline, centered at the mean velocity, for each component. This was indicated in the previous histograms. This ensured capturing all the relevant velocities, while neglecting the most officious outliers generated by signal noise. Between the incident laser intensity, the backscatter setup, and the maximum beam spacing of the probe, the highest SNR for the given equipment resulted in the data rates shown.
For a more physical estimate, the restrictions of processing non-uniformly sampled data require that a higher sample rate (times a modified Nyquist factor of πN) is
92 required before the spectra are attenuated unnaturally by the FFT, corresponding to resolved frequencies in the flow of about 640-1100 Hz, based on the lower and higher sample rate ranges of about 4000 and 7000 in Figure 5.12. The relevant
Strouhal numbers, based on Vmix and Deq, are 0.31 and 0.52 for the resolved frequency range, within a likely reasonable range for the most energetic features.
Figure Figure 5.13 shows the results of processing the limited number of
Figure 5.13 - Spectral centerline results from LDV u velocity component SFT, all nozzles. averages available for 4096 samples per block with a frequency resolution of 8 Hz.
The upstream portions of the spectra must be interpreted carefully, for the reasons discussed regarding misalignment issues. This is probably cause of the fairly sharp change in the chevron turbulence spectra at about Z/Deq=2.4. Recall from Figure 5.8
93 that the histograms showed fairly good consistency after about this point. This should allow for consideration of the downstream spectra. At about the location where the chevron turbulence intensity reaches its peak value (Figure 5.10), the spectra indicate dissipation across the lower frequencies, and some gain in the higher frequencies, which may be difficult to judge due to the high frequency attenuation. Nevertheless, the lower frequencies are losing energy, whereas in the case of the 12DVG nozzle, the peak levels in the lower frequencies seem to be translated downstream, corresponding approximately to the location (Z/Deq=2.8) where the second linear velocity reduction region begins (V2). Delaying the growth of the large structure could be attributed to this observation, which would lead to low-frequency noise reductions. The baseline turbulence spectra, although much less energetic, concentrate the most energy in the low frequencies centered in a region at about Z/Deq = 2.5. A region of intense lower-frequency turbulence appears also for the chevron in this location. The CVG nozzle shows a very different character, with a void at this location. Whether the correlation with the centerline velocity decay is causal is uncertain, but it seems to support the idea of greatly different mixing mechanisms, or relative importance of the axial efflux and vorticity sources, for chevrons and CVG’s.
94 Chapter 6 Acoustic Far-Field Noise Reduction Results
Measurement of the far acoustic field for permutations of several base configurations of high bypass ratio core nozzles was performed to assess the overall noise reduction potential for each, and determine the sensitivity of influential parameters. For the conventional 8-lobe chevron core nozzle, sinusoidal root geometry reduced the high-frequency noise generated by the small flow scales existing at the chevron root. CVG’s were also installed at the root of the device in order to further increase noise reduction and expose potential differences in mixing mechanisms between the CVG’s and the chevrons. The internal chevron nozzle, tested in both an improved configuration compared to the baseline fabricated design, and at various penetration levels, showed that similar trends to conventional chevrons exist, with the potential to be actuated more easily, and possibly further improve noise reduction Stream conditions included heated core flow at the cycle point designations found in Table 6.1.
Table 6.1 - Simulated cycle core and fan stream conditions and combined attributes
Core Stream Fan Stream Combined
Designation CPR T1 (˚F) M1 U1 (ft/s) FPR T2 (˚F) M2 U2 (ft/s)Ushear Umix Single 1.50 220 0.78 1003 - - - - 1003 1003 Coax-LO 1.80 220 0.96 1224 1.26 75 0.58 663 561 778 Coax-HI 1.80 220 0.96 1224 1.50 75 0.78 890 334 958
95
6.1 Baseline and Modified Chevron Root Geometry
The SPL spectra for the baseline conic core nozzle, the unmodified 8HP chevron, and the 8HPsin sinusoidal root chevron, Figure 6.1, reveal the main effect of the root geometry on the high-frequency noise produced by chevrons. Data for the
Base-Base 8HPsin-Base 8HP-Base 90 105 90˚ 150˚ 100 85 Coax-HI 95 Coax-HI B) B) d d ( (
80 90 PL PL S S 85 75 Single Single 80 Frequency (Hz) Frequency (Hz) 70 75 100 1000 10000 100000 100 1000 10000 100000
Figure 6.1 - 8HP chevron unmodified and sinusoidal root TOB spectra at 90˚ and 150˚
Coax-HI and Single cycles are shown, the effects being much less dramatic for the coaxial flow case. The sinusoidal root produces, at the upstream angle of 90˚, a nearly identical low-frequency character. A broadband decrease in the higher- frequency noise, just beyond the peak in the forward-directed spectra at a frequency of about 10 kHz, indicates a reduced strength in the quadrupole turbulence source associated with chevrons. A spectral delta plot, using the baseline conic nozzle spectra as the datum, at the 90˚ directivity angle of Figure 6.2, shows that the reduction of the source has a similar nature for both stream conditions. For each flow, the 8HPsin nozzle curve bifurcates with the 8HP nozzle spectrum and, as the
96 frequency increases, monotonically 8HP-Base 8HPsin-Base 2 increases both the absolute noise 1 B) reduction, and the noise reduction d 0 PL ( S
-1
relative to the 8HP nozzle. It is Delta -2 believed that the physical mechanisms Frequency (Hz) -3 100 1000 10000 100000 for the reduction are the same for each Figure 6.2 – TOB spectral delta at 90˚ for flow and nozzle. These include a 8HP and 8HPsin nozzles, the arrows modification of the turbulence intensity, indicating reduced source transition range. spectrum, and correlation lengths of the chevron jetlets, as well as a change (decrease) in the lateral transport of axial momentum in the 8HPsin case due to the radius covering part of the chevron serration. The change to a sinusoidal root also affects the low-frequency peak noise levels similarly at aft angles, supporting the notion that the mechanisms are the same.
Thus, the frequencies of interest are similar for each flow, being only those associated with the start of the deviation of the 8HPsin spectra from the 8HP spectra. The single flow case reveals this deviation more clearly, clearly showing a noise source transition over a frequency range bounded by (indicated with arrows) about 10 and 30 kHz, with a peak (minimum noise reduction) of about 20 kHz. This range marks the reduction of acoustic sources by eliminating small scales, and should correspond loosely to the largest length scale in the change from sharp chevron root to a sinusoidal root. This should be true for both coaxial and single jets, with the spectral delta curves for the 8HP and 8HPsin nozzles appearing
97 “scaled” at high frequencies, likely due to the higher levels of noise in the coaxial case. This makes the reduced high frequency noise “transition” of the 8HPsin essentially occur at a discrete frequency (~20kHz) in the coaxial case, but corresponding to the same center frequency of the transition region for the single jet. This frequency coincidence seems to further validate the belief that the physics are similar for the flows, but simply more apparent in the single flow case, and also implies the flow scaling frequency of interest is that associated with the flow velocity through the chevron root, independent of the fan stream velocity. Scaling on the core velocity, covered length of the chevron root of 0.25”, and dimensional frequency of 10 kHz (20 kHz) from Figure 6.2, corresponds to a Strouhal number of
0.21 (0.42). This might be taken as the peak noise frequency of the jetlet portion eliminated by the smooth chevron root, which roughly correlates in the range of
0.21 to 0.42 with the peak noise of a small isothermal jet [69]. In this way, eliminating a small portion of the lateral jetlet reduces largely the high-frequency noise, and even shows that chevrons can be made to produce noise reduction beyond these frequencies. Some loss in the typical aft low-frequency noise reduction is sacrificed, though.
By altering the root geometry in a completely new way, by placing CVG vortex generators just upstream of the nozzle centered on the chevron root vertex, a very different use of the kinetic energy available in the lateral jetlets was made. As will be described in greater detail in Chapter 7, the two phenomena associated with the CVG’s are the fluid deflection up into the high-speed stream, and the relatively strong stream-wise vorticity generation. By installing 8 vortex generators at the
98 roots of the chevrons of the 8HP nozzle (1 per each), the intent was to forestall the lateral turning of the fluid to a point well past the chevron root to attempt a reduction in high-frequency noise relative to the chevron, and also to produce a more vigorous mixing structure in the form of enhanced stream-wise vortices.
The upstream and aft spectra for the 8HPVG nozzle, which incorporates the
90 Base 8HPVG 8HP 90 100 90˚ 150˚ 85 95 Coax-HI Coax-HI B)
80 B) 90 d d PL ( PL 75 PL (
S 85 S Single Single 70 80 Frequency (Hz) Frequency (Hz) 65 75 100 1000 10000 100000 100 1000 10000 100000
Figure 6.3 - Hybrid 8HP and CVG core nozzle TOB spectra at 90° and 150°
CVG’s at the chevron roots, show that significant reduction relative to 8HP noise levels across nearly the entire spectrum was attained. Figure 6.3 shows the spectral properties of the hybrid chevron and CVG nozzle at directivity angles of 90° and
150°. Some interesting features of the CVG’s are apparent from the upstream directed far-field spectra, related to the CVG self noise. For the Single flow conditions, the high-frequency noise created by the VG’s is less than that of the chevrons, true for all directivities actually. More extensive flowfield data will reveal the mechanisms more clearly, but in light of the results in Chapter 7, the likely cause of the additional acoustic benefit is exactly that stated previously.
Since the CVG’s create a strong stream-wise vorticity field, it appears that this
99 encourages greater mixing rates. In the coaxial case, some additional noise is created by the CVG’s, whereas the 8HP practically replicates the baseline conic nozzle spectra at 90° since the shear velocity is relatively low, a key parameter for chevron effectiveness in coaxial flow. By introducing the secondary stream, which interacts with the jetlets, the greater the transverse velocities, such as inherent in concentrated counter-rotated vortex pairs, the shearing and fluctuating Reynolds stresses should become large relative to the case where the jetlet energy is mainly composed of axial momentum flux. In order that this effect should be consistent with a static fluid surrounding the single jet, which according to Figure 6.3, produces no more noise at high frequencies than the 8HP, either a source of noise is replaced by the CVG’s, or the vortical structures from the CVG’s behave dynamically dissimilar when the jetlets are stretched axially by the bypass flow.
This is still somewhat unclear, the proportional significance of dynamic fluid mechanisms that are important noise generators in the near-field of the jet when coaxial flow is present. Further study is warranted, though, since despite the small
1
0 Coax-HI B) d -1 PL ( AS O -2 Single Directivity Angle (degrees) -3 60 80 100 120 140 160
Figure 6.4 - Hybrid 8HP and CVG nozzle OASPL directivity for Single and Coax-HI setpoints.
100 increase at high frequencies, the directional characteristics of the jet with the hybrid chevron and CVG nozzle show promising noise reduction benefits (Figure 6.4).
6.2 Internal Chevron Core Nozzle Results
A preliminary acoustic investigation of the IC nozzle, described in Section 3.2.2, revealed a large source of high-frequency broadband noise caused by manufacturing method. A plot of the covered nozzle, which reduced greatly the noise caused by the air jets through the fabrication slots for the chevrons, is shown in Figure 6.5.
105 Baseline Conic IC Nominal IC Covered 100 90 105 70˚ 150˚ 100 85 Coax-HI 95 Coax-HI
B) 80 B)
d 90 d PL ( 75 PL ( 85 S Single S Single 80 70 75 Frequency (Hz) Frequency (Hz) 65 70 100 1000 10000 100000 100 1000 10000 100000
Figure 6.5 - IC core nozzle TOB spectra for Single and Coax-HI setpoints at 70˚ and 150˚,
comparing nominally fabricated nozzle and the addition of a simple steel sheet covering.
Both in the single jet case and the coaxial flow, the slots used for installation of the chevrons within the internal surface of the nozzle generate large levels of noise beyond the baseline conic reference. Covering the slots on the outer surface of the nozzle, and imposing a penetration level of 0.00”, flush with the interior contour, the baseline nozzle acoustic character is closely approached. Some significant differences in the spectra still exist, namely the high-frequency noise generation at 101 upstream angles. Figure 6.5 implies that two sources exist, composed of a core stream source associated with the cavity through which the chevron projects, and a source present in coaxial flow associated with perturbation of the outer core boundary layer. However, based on the extensive acoustic reversion with a fairly simple improvement, and considering the relative insensitivity of the inner core nozzle boundary layer as shown in Chapter 4, which must be greatly disturbed by the large pockets to generate additional noise, a more sophisticated fabrication method would probably eliminate most if not all of the unintended high-frequency component.
The penetration levels of the IC nozzle was varied using the Single cycle setpoint, Figure 6.6 representing a large range of variation at 90º directivity, to determine the extent of any noise reduction and the nature of the additional high-
85 Base 0 (Cover) 5 10 15 20 85 85 a) 90˚ b) 90˚
80 80 B) B) d d PL ( PL ( S S 75 75
Frequency (Hz) Frequency (Hz) 70 70 100 1000 10000 100000 100 1000 10000 100000
Figure 6.6 - IC core nozzle TOB spectra for Single setpoint comparing penetration effect at
90°. a) Nominal and 5 penetration (value x 0.01 in.), b) 5, 10, 15, 20 penetration values frequency noise. Comparison of the lowest penetration level of 0.05 in. shows a negligible increase in high-frequency noise with a moderate broadband decrease
102 across the peak noise frequency at the jet side angle. By doubling the penetration to a value of 0.10 in., the high frequency noise generation is no longer small compared to the nominal covered IC nozzle, and further increases in penetration accompany increasingly large source generation that even supersedes the original baseline peak noise level.
An attendant decrease in the low frequency noise, however, is significant at the aft angles, to a maximum of about 3.4dB SPL. The 150º spectra in Figure 6.7
95 95 a) 90˚b) 90˚ 90 90 B) 85 B) 85 d d PL ( 80 ( PL
S 80 S
75 75 Frequency (Hz) Frequency (Hz) 70 70 100 1000 10000 100000 100 1000 10000 100000
Figure 6.7 - IC core nozzle TOB spectra for Single setpoint comparing penetration effect at
150°. a) Nominal and 5 penetration (value x 0.01 in.), b) 5, 10, 15, 20 penetration values shows a small broadband decrease for the lowest penetration value, and increasingly large aft noise reduction as the penetration increases, similar to a conventional chevron. The high-frequency noise though, is particularly high-level, compared to a chevron, and contributes to an overall lower noise reduction in the current configuration compared to the 8HP nozzle. In considering the OASPL, Figure 6.8 shows that very low penetration values can approach the benefit of the chevron.
Potentially, the IC configuration could be far easier to implement SMA materials or other actuators within as opposed to conventional chevron nozzles. To mitigate or
103 2
1 B) d PL ( 0 AS O
-1 8HP Delta
Directivity Angle (degrees) -2 60 80 100 120 140 160
Figure 6.8 - IC core nozzle directivity comparison with 8HP chevron. eliminate entirely the high-frequency noise associated with these nozzles that deteriorates the overall noise reduction, possible further study should vary the placement of the chevrons axially within the nozzle, possible even achieving a hybrid configuration between the IC nozzle and conventional chevron nozzle. It was shown that simple root geometry modifications drastically reduced the high- frequency noise, and considering the CFD data evaluated in Chapter 8, the vortical structures that evolve past the trailing edge of a delta tab inclined to a developing wall flow may be a large cause of the noise. Repositioning the chevrons downstream then could eliminate or reduce much of the noise generated by these structures impacting the nozzle wall or causing scattering as the turbulence convects past the nozzle trailing edge.
104 Chapter 7 Incompressible CFD Simulations
Three test cases were computed using a steady RANS flow solver and two-equation k − ε turbulence model with Fluent. Two of the test geometries are shown in Figure
a) c) a)
c)
Figure 7.1 – CFD geometries. a) Solid VG (SVG), b) CVG (not shown), c) delta tab (mm)
7.1. The basis for Figure 7.1a is a typical CVG geometry, not shown, with modifications from actual hardware to allay computational grid and flow stability concerns. Deviations include small variations in angles and retaining a thickness to the geometry that is greater than that used to fabricate the actual hardware, which was composed of 0.01 in. steel sheet stock. These issues are believed to have a small order effect on the flow development. The SVG geometry was computed to quantify the important difference in vorticity generation mechanisms for a simpler geometry similar to a triangular tab, and with as much similarity to the CVG as possible. Computation of a typical delta tab geometry in Figure 7.1c was performed to provide a reference case for which some experimental data exists, both in literature and for the internal chevron nozzle acoustics in this work, and to further compare the resulting vorticity contributions and development. The delta tab height
105 was reduced to 5mm due to the large exposed sharp edge length, attempting to approximate a better comparison to Figure 7.1a. The details of the computational setup were specified to create a solution that would accurately describe this wall- bounded flow with regions of moderate local turbulence, shear and vorticity. A degree of accuracy allowing the development and characteristics of the flowfield downstream of the vortex generators in a boundary layer flow with the height H on the order of δ was the desired result, presented in the following sections.
7.1 Computational Domain and Case Setup
The geometry of the computational domain was a function of the penetration of the
CVG into the stream, or the height of the CVG above the lower bounding plane, H.
This determined the Reynolds number based on CVG height of 6mm, to be 10,000, reasonable considering the vortical structure has been shown to be independent of Mach number, despite the experimental Reynolds number on the order of 105. The physical arrangement y and dimensions of the numerical domain x z are illustrated in Figure 7.2. As a first study, the difficulties in modeling a shear Figure 7.2 – CFD solution domain. layer with strong stream-wise vortices
106 was avoided, instead continuing the lower bounding surface through the extent of the domain. The k − ε model was used to obtain reasonable vorticity fields.
A uniform inflow of 100 m/s was prescribed at the entrance plane, with the boundary layer developed using a semi-empirical enhanced wall treatment. This method applies the k − ε model to the out region of the boundary layer, but increasing resolution using an unstructured mesh nearest the wall. A one-equation turbulence model, by Wolfstein, is applied nearest to the wall to model the buffer layer and the viscous sublayer. For the inflow velocity, and the higher resolution grid nearest the wall, the y+ was usually over 70 nearest the wall, which should allow more accurate modeling of the buffer and viscous sublayer regions below y+~60. The development of the boundary layer for the SVG and CVG are shown for several locations on the centerline of the domain in Figure 7.3, indicating the height of the VG geometries for reference.
a) b)
Figure 7.3 - Boundary layer profiles at several values of z: 0 (inlet), 10, 20, 30, and 40mm, for both a) SVG and b) CVG geometries. Dashed line indicates the uppermost vertex of the VG’s.
107 A flux-conserving BC was applied to the exit plane, and no-slip wall condition on the lower plane. Slip conditions were applied otherwise. To improve the resolution at the boundary layer a hybrid mesh was composed of an irregular mesh at the lower bounding wall, while a hex mesh was used throughout the remaining domain. A detailed view of the closed VG geometry is shown in
Figure 7.4. The total domain consisted of about
905k computational cells. A grid resolution study was not performed, but similar computations Figure 7.4 - Hybrid near-wall studies using similar techniques have found mesh detail for SVG geometry. adequate results with grids of this size104.
108 7.2 Flowfield Scalar Results
An overview of the velocity midplane flowfield features is shown in Figure 7.5for the SVG case. The region of circulation behind the SVG is fairly strong, extending
Figure 7.5 - Entire domain X/H=0 velocity features of SVG geometry. Isosurface at U0=0.20. about 2H downstream, and produces an extended wake, with the rotational flow directly behind the SVG directed downward into the boundary layer indicated by stream ribbons, which curl up several times initially behind the SVG. Axial momentum is abruptly transferred mainly in the lateral directions, and at no point is there a noticeable influx of this momentum back along the centerline of the wake as might occur from stable streamwise vortices of moderate circulation strength.
109 The surface pressure contours are compared to measured values in Figure 7.7, from a triangular tab installed with the base at x / w = 0, and at 45° angle to the flow.
Similar rapid falloff of upstream pressure influence is shown, and obvious a) b) c)
Figure 7.7 – a,b) Static Cp contours for SVG, c) measured Cp for a jet triangle tab, M=0.3 [82]. differences are likely from the shaping of the tab and the transverse wall curvature of the jet. This creates the transverse pressure gradient, that when combined with the transverse pressure gradient behind the blockage, will determine the strength of
Figure 7.6 - Entire domain X/H=0 velocity features of CVG geometry. Isosurface at U0=0.20.
110 the “pressure hill” source of vorticity. The alignment of the relatively strong, compact vorticity shedding from the sharp trailing edge will combine in a positive or destructive way with the transverse pressure gradient source. The overview of
CVG solution gross velocity magnitude field at X/H=0 is shown in Figure 7.6, with some noticeably large downward streamline deflections, a bifurcated separated region, and increased twisting as the stream ribbon follows downstream. The delta tab case is shown in Figure 7.8, with an adjustment in the X/H direction to compensate for a slightly different computational domain size, aligning the devices at the most downstream point on each. The delta tab vorticity vector orientation clearly induces upwash downstream of the body. A large region of velocity defect persists for a relatively much greater region downstream compared to the other two geometries. This might indicate additional flow noise produced by this type of geometry due to unsteadiness inherent in the velocity gradients.
Figure 7.8 - Entire domain X/H=0 velocity features of delta tab. Isosurface at U0=0.20.
111 A comparison of the pressure contours for all cases, Figure 7.9, shows that the influence extends several H upstream in each. The major difference is the lateral extent of the high pressure region at the base of the geometry, and the markedly different downstream pressure gradients. The gradients are especially high near the leading edge of the CVG, and are also fairly high at the lateral vertices
a) b) c)
Figure 7.9 –Cp contours at Y/H=0 (wall). a) SVG, b) CVG (inexact geometry), c) delta tab of the delta tab. The pressure contours near the trailing edge of the CVG are approximately the same as for the SVG on the pressure side, but the downstream low pressure contours are much less steep, owing to the yaw angle of the front face.
Reduction of the low-pressure lobe along the CVG centerline region is seen due to the split. This increased pressure lobe on the centerline, and slightly upstream of the trailing edge of the CVG is an indication of the rapid stream-wise addition of downwash. As installed on a trailing edge plate, this could be a source of loading noise or trailing edge noise dipoles making up part of the increased high-frequency far-field noise spectrum. The delta tab has a much different pressure profile along the front edge, showing a high, even pressure across the entire face, and a large low- pressure region that extends significantly farther downstream, creating much higher
112 levels of pressure drag in comparison, even at a low angle of attack of 30º to the surface.
To better estimate the performance of the devices in the mechanisms of generating and combining vorticity sources in a qualitative way, 3D pressure isosurfaces are shown in Figure 7.10 in conjunction with streamlines arrayed upstream to pass along the trailing edges. The effects of the strongest pressure gradients, shown as two discrete isosurfaces, is to deflect the flow around the object in such a way that there is a cascade of vorticity scales depending on the magnitude of the gradient, which obviously varies continuously throughout the flowfield. The streamlines curve slowly in the regions of smaller gradients, but “injected” sufficiently upstream, will generally not enter the momentum sparse circulation region. Distortion of the low-pressure region by the CVG’s compared to the other two geometries, however, by injecting momentum through the center of the split, allows for a larger portion of the flow energy to interact with the low pressure region. At the downstream leaning trailing edge, the flow is already reorienting braids into stream-wise filaments, and a larger structure evolves.
a) b) c)
Figure 7.10 - Cp isosurfaces at +0.25, -0.60, & streamlines, for a) SVG, b) CVG, c) delta tab
113 The surface pressures on the tabs revealed some additional insight into the stream-wise momentum generation process and device drag. Figure 7.11 compares the pressure coefficient for all geometries. The trailing edge of all geometries is similar in transition of pressure values to the backward face. The leading edge of the CVG, however, displays a remarkable difference. The yaw angle of the VG induces a high pressure gradient across the edge. The inner faces of the CVG even show the lowest pressure of all the geometries. This creates a large gradient along this edge that seems to be an important source of converting the axial momentum to lateral swirling motion in a positive coordination with the less tightly vortical flow away from the edges into the freestream, which is not present in the other cases. In the delta tab case, the low pressure region maintains distinction from the “pressure hill” generated vorticity source, which shows no noticeable vortical structure, probably due to the orientation of the vortices. In the case of a delta tab on a splitter plate trailing edge or nozzle, the flow would have a crucial freedom to
a) c)
Figure 7.11 - VG surface pressure levels. a) SVG, b) CVG (not shown), c) delta tab deflect outwards, into the shearing stream, which is captured in this simulation.
Although the resolution of this simulation was not very high, it does indicate that
114 drag coefficient would likely be similar for these devices, and noting that the CVG is yawed to the flow angle decreases the drag effects of the large low pressures.
7.3 Stream-wise Vortical Structures Development
Details of the vorticity crossplanes were examined to explain the potential mechanics of the near-field stream-wise vortex formation, and what the important features of the initial development might imply for shear-mixing flows. Although these computations were performed on a relatively coarse grid, various computations using both SA (1-eqn. Spalart-Allmaras) and SST (2-eqn. Shear Stress
Transport) turbulence models [104], and also using two degrees of grid resolution, provided similar initial flow regions. The major differences were generally in the peak vorticity evolution and associated vortical flow quantities, which should be expected for regions of high shear with the SA model. Turbulence models also caused the largest discrepancy in the initial flowfield. Thus, the following data has been analyzed with regard to offering explanations of near-field shear layer mixing mechanics, attendant mixing effectiveness, and potential noise sources introduced by generation of turbulence at small scales.
115 First, an overview of the vorticity magnitude features in each case introduces r ω H /U the gross variations in the VG wakes. Isocontours of 0 at various values, corresponding to the approximately equivalent volumes along the trailing edge, show significant differences in magnitude. The delta tab is at a relatively small obstruction angle, which explains the reduced magnitude in that case. On comparing the vorticity just downstream in each case, the two lobes of the CVG seem to exist at greater a) strength downstream.
To evaluate the stream-wise vorticity development, cross-planes are plotted in plan view b) over a fairly close range of
Z/H values in Figure 7.14.
The planes begin at the leading edge of the device, the first three representing c) equal Z/H displacements,
where the third is aligned at the device vertex of each. The next two are similarly spaced to reach the trailing edge of the Figure 7.12 – Partial domain gross vorticity behavior, over X/H=0, Y/H=0.5 planes. a) SVG, b) CVG, c) delta tab 116 device, thus the first five planes represent the flow through the model geometry.
The downstream planes are then spaced past the device trailing edge by Z/H of 0.5,
1, 2, 3, 4, and 6. Figure 7.13 represents the cross-stream locations in the full computational domain. The delta tab, which has a vertex located at the trailing edge, was divided in an equivalent manner preserving the number of cross-planes.
Thus a total of five are located within the Z/H limits of the VG, and six are located downstream at equivalent displacements.
In the way of stream-wise vorticity production, the SVG has very small levels, even by the trailing edge point. This is due
to the geometrical alignment of the trailing Figure 7.13 - CFD cross-plane Z/H edge, mostly in the Y-Z plane, as is shown locations relative to entire domain. by the initially strong bound vorticity in
Figure 7.12. In the case of the CVG, it is apparent that some interaction is occurring between the lateral momentum and the device, which would be altered for an essentially two-dimensional geometry, but by the 4th station, has detached from the wall of the device. The delta tab shows little device self-interaction.
117 The most interesting feature of the initial region is the development of the
a) SVG b) CVG c) delta tab
1
2
3
4
5
Figure 7.14 - Normalized stream-wise vorticity cross-planes at various Z/H stations, equally divided at 25% LVG, into five nominally located positions. a) SVG, b) CVG, c) delta tab pair of counter-rotating stream-wise vortices in the central region of the CVG and delta tab geometries, and their respective orientations. The initial rotation of the
118 CVG is such that it creates a vortex-pair within the high-speed stream away from the wall and shearing forces. The delta tab geometry, on the other hand, introduces an opposite orientation, so that the initial rotation acts to shear the transverse flow along the wall before being reinforced by the oppositely formed vortex. This is important, since the vortices will interact with high levels of shear immediately after formation which may destabilize and strain excessively leading to breakdown.
The stream-wise vorticity development downstream of the trailing edges of the devices is presented in Figure 7.15. Clearly, the stream-wise vorticity is very weak in the SVG case, although a pair of counter-rotating vortices can be distinguished, the overall weakness does not seem to improve the downstream preservation or vorticity through stabilization. The vorticity vector is mainly oriented in the cross-plane, and appears to remain mostly in this plane throughout the vortex lifetime. Although initial strength of the delta tab stream-wise vortices is somewhat higher, due to the bound vorticity shedding from the sharp trailing edge being mostly aligned in this case with the downstream direction, shearing tends to disintegrate the stream-wise rotation over a fairly short distance. Still, some evidence of stream-wise vorticity is left at the farthest downstream station of 6 Z/H past the trailing edge. The CVG vorticity in the X-Y plane remains very strong over this initial development range, diffusing at a small rate through a range of about 4-6
H.
A view of the vorticity magnitude in some corresponding planar regions shows that although the overall vorticity may be quite strong in the case of the SVG
119 and delta tab, the transverse swirling energy is the most dominant component in the case of the CVG. Figure 7.16 shows contours in planes downstream of the trailing
120 a) SVG b) CVG c) delta tab
0.5
1
2
3
4
6
Figure 7.15 - Normalized stream-wise vorticity cross-planes at various Z/H displacements relative to trailing edge as indicated. a) SVG, b) CVG, c) delta tab (see Figure 7.15 scales)
121 edges. For the CVG, generally the areas of maximum stream-wise vorticity correspond strongly with the regions of maximum vorticity magnitude. This is not true in the case of the delta tab and SVG, where the vorticity does not reorient itself due to the larger scale upstream pressure hill vorticity source. The vorticity shed near the peak of the delta tab further does not seem to contribute to the stream-wise vorticity as is convects downstream. The likely source of the re-orientation of the vorticity is the mean flow vector, which for the CVG is very different than both the
SVG and CVG. Secondarily, the vorticity that is being shed from the dowstream trailing edge, due to the fact that it is tapered (not a typical “half delta wing” triangle with a vertical trailing edge), also reorients the vorticity to the same downstream sense.
Contours of axial velocity in Figure 7.17 at selected cross-planes downstream of the VG show the nature of momentum distribution in relation to the vorticity of
Figure 7.15 and Figure 7.16. The separated regions of the SVG and delta tab are somewhat similar in form, although the delta tab flow separation is much larger and extends farther downstream. The CVG however, shows significant velocities in the center region of the flow-field, with some acceleration past the tips in the second cross-plane. This central flow bifurcates the wake into two distinct regions of velocity deficit. Comparing the vorticity magnitude in the planes downstream in the developing wake, the stream-wise vorticity for both the SVG and delta tab are subjected to the axial shearing. Thus, energy which seems to be injected constructively with respect to the developing vortex field of the CVG, seems to weaken the stream-wise vorticity of the other two.
122 a) SVG b) CVG c) delta tab
0
0.5
2
3
Figure 7.16 - Normalized vorticity magnitude cross-planes at various Z/H displacements relative to trailing edge as indicated. a) SVG, b) CVG, c) delta tab
Plots of normalized TKE shown in Figure 7.18, shown at a downstream location 2H past the trailing edge seem to confirm this. The interaction zone where the separated region meets the strong spanwise vorticity shedding from the tip region of the delta tab shows large turbulence generation. The SVG has the lowest turbulence levels, and appears to be associated with the region where the mean flow is reconnecting past the separation zone. The low levels of TKE near the tip of the
123
a) SVG b) CVG c) delta tab
-1
-.5
0 (TE)
.5
1
3
Figure 7.17 - Normalized axial velocity cross-planes at various Z/H displacements (nominally for upstream) relative to the trailing edge (TE) as indicated. a) SVG, b) CVG, c ) delta tab
124 SVG, as opposed to the concentrated region for the delta tab, may arise from the relative lack of vorticity generated at the tip of the SVG. The CVG has a concentration of vorticity near the core of the stream-wise vorticity, with some additional turbulent energy apparently related to the vorticity generated at the outer spanwise corner of the VG where a high spanwise velocity gradient also persists downstream. This implies that the mean flow developed by the VG outside of the highly rotational region works to reduce the shearing of the vortices, and possibily stabilizing the stream-wise cores in addition as they are carried downstream.
a) b) c)
Figure 7.18 - TKE contours at 2H downstream of trailing edges. a) SVG, b) CVG, c) delta tab
125
Chapter 8 Conclusions
Broadband jet noise reduction was investigated with respect to the influence of the various shear layers. VG’s, modified chevrons, internal chevrons, and hybrid nozzles were shown to provide good peak noise reduction of jet noise up to practically all directivity angles. The root modifications of the chevron nozzles revealed a large acoustic tradeoff at both peak noise levels and for excess noise sources. The VG configurations offered the best noise reduction potential in terms of OASPL at all most directivity angles, and increased the effective noise benefit of chevron nozzles by a significant amount. High frequency broadband noise, a common by-product of mixing devices of all types, was also observed in these tests for some configurations, generally those with the greatest peak noise reductions, as a quadrupole source.
These mixing devices were investigated in their impact on the jet development with time-resolved LDV measurements for a large portion of the main turbulent energy along the centerline of the jet. Large generation of turbulence across a sizeable range of the frequency spectrum, and in great spatial extent well past the core and fan potential cores at ~6Deq was observed. The mean results indicated some alternate form of mixing mechanics may be the primary mode of axial momentum flux in the radial direction for the VG’s compared to chevrons.
126 CFD simulations of a simplified boundary layer VG’s and delta tab provided details on the near-field development of the fluid structures responsible for mixing and entrainment of secondary flow into the primary stream across the shear layer. Pairs of streamwise counter-rotating vortices were predominant in the flowfield of the
CVG, but were essentially a background feature compared to the large separation region of the SVG which promoted an extensive wake. The delta tab vorticity shed from the trailing edge of the tab appeared to develop much differently also from that of the CVG, in terms of translating lower toward the bounding plane, and also exhibiting a relatively large axial rate of vorticity magnitude decrease for the streamwise component. Turbulent energy in the immediate wake of the delta tab was also up to 40% greater than that of the other geometries, throughout a larger volume.
127 Bibliography
1 2 3 4 56 7 8 9 10 11 12 13 14 15 16 17 18 19 20
21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 37 38 39
40 41 42 43 44 45 46 47 48 49 50 51 52 53 54 55 56 57 58 59
60 61 62 63 64 65 66 67 68 69 70 71 72 73 74 75
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96 97 98 99 100 101 102 103 104 105 106 107 108 109
110 111 112 113 114 115 116 117 118
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