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Design and Characterization
of the UTIAS Anechoic Wind Tunnel
by
Derrick H.F. Chow
A thesis submitted in conformity with the requirements
for the degree of Master of Applied Science
Graduate Department of Institute for Aerospace Studies
University of Toronto
© Copyright 2016 by Derrick H.F. Chow
ProQuest Number: 10043015
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Abstract
Design and Characterization
of the UTIAS Anechoic Wind Tunnel
Derrick H.F. Chow
Master of Applied Science
Graduate Department of Institute for Aerospace Studies
University of Toronto
2016
The anechoic open-jet wind tunnel facility at the University of Toronto Institute for
Aerospace Studies was updated and characterized to meet the needs of current and future
aeroacoustic experiments. The wind tunnel inlet was resurfaced and flow-conditioning
screens were redesigned to improve the freestream turbulence intensity to below 0.4% in
the test section. The circular nozzle was replaced with a square secondary contraction
that increased the maximum test section velocity to 75 m/s and improved flow uniformity
to over 99% across a usable cross-sectional area of 500 mm × 500 mm. Acoustic baffles
were installed in front of the wind tunnel inlet and foam wedges were installed in the
anechoic chamber. The overall background sound pressure levels in the chamber were
improved by 8-18 db over the range of operational freestream velocities. The anechoic
chamber cut-off frequency is 170 Hz and the reverberation time for a 60 dB sound power
decay is 0.032 s.
ii
Acknowledgements
I would like to thank Professor Philippe Lavoie for all the knowledge and wisdom he
imparted on me throughout my time in his lab. His patience and support were key to
completing the work in this thesis. I also appreciate all the professors on my research
assessment committees for lending their time and expertise to ensure my work met high
standards of quality.
I am deeply grateful to Philip McCarthy, a.k.a. “British Phil”, for the incredible dedication and effort he displayed in helping me throughout the wind tunnel retrofit process. I
am glad that he was there with me during this journey and I have learnt a lot from him.
Finally, I’d like to express my sincerest thanks to all my colleagues in the FCET and EFD
lab groups, as well as several others in the UTIAS community, who gave me the critical
support I needed to complete my work and ensure that I had a good time at UTIAS.
iii
Contents
1 Introduction
1
1.1
Motivation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
1
1.2
Airframe Noise Research . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3
1.3
Outline of Thesis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5
2 Background
2.1
7
Anechoic Wind Tunnel Design . . . . . . . . . . . . . . . . . . . . . . . . .
7
2.1.1
Aerodynamic Specifications . . . . . . . . . . . . . . . . . . . . . .
8
2.1.2
Acoustic Specifications . . . . . . . . . . . . . . . . . . . . . . . . . 10
2.2
Survey of Existing Facilities . . . . . . . . . . . . . . . . . . . . . . . . . . 13
2.3
Existing UTIAS facility . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17
3 Facility Updates
3.1
23
Objectives . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 23
3.1.1
Aerodynamic objectives . . . . . . . . . . . . . . . . . . . . . . . . 23
3.1.2
Acoustic objectives . . . . . . . . . . . . . . . . . . . . . . . . . . . 24
3.2
Flow Conditioning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26
3.3
Secondary Contraction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29
3.4
Acoustic Wedges . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32
3.5
Front-end Baffles . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 38
4 Facility Characterization
40
4.1
Flow Speed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 40
4.2
Flow Uniformity . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 41
4.3
Freestream Turbulence Intensity . . . . . . . . . . . . . . . . . . . . . . . . 45
4.4
Background Noise Levels . . . . . . . . . . . . . . . . . . . . . . . . . . . . 47
4.5
Anechoic Cut-off Frequency . . . . . . . . . . . . . . . . . . . . . . . . . . 54
4.6
Reverberation Time . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 59
iv
5 Conclusions
62
5.1 Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 62
5.2 Recommendations for Future Work . . . . . . . . . . . . . . . . . . . . . . 63
A Anechoic Foam Material Data Sheet
v
65
List of Tables
1.1
1.2
ICAO subsonic aircraft noise reduction standards. . . . . . . . . . . . . . .
ACARE and NASA aircraft noise reduction targets. . . . . . . . . . . . . .
2.1
Specifications of select aeroacoustic research facilities around the world. . . 15
3.1
Summary of aerodynamic improvement objectives for the UTIAS AWT
facility. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 23
Summary of acoustic improvement objectives for the UTIAS AWT facility. 24
Comparison of new and old wire mesh screen specifications. . . . . . . . . . 29
3.2
3.3
4.1
4.2
4.3
5.1
2
2
Summary of sound transmission distances for baffle noise attenuation test. 52
A-weighted OASPL white noise attenuation by front-end baffles and wind
tunnel components. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53
SPL error bounds for free-field assumption. . . . . . . . . . . . . . . . . . . 55
Summary of tunnel performance objectives and results from the facility
characterization. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 62
vi
List of Figures
1.1
Major sources of airframe noise. . . . . . . . . . . . . . . . . . . . . . . . .
4
2.1
Examples of anechoic wind tunnel facilities around the world. . . . . . . .
8
2.2
A-weighting SPL curve. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12
2.3
Comparison of background OASPL of select anechoic wind tunnels around
the world. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16
2.4
Views of the existing UTIAS AWT facility. . . . . . . . . . . . . . . . . . . 18
2.5
Freestream turbulence intensity plots for old AWT. . . . . . . . . . . . . . 19
2.6
Test section velocity profiles for old AWT. . . . . . . . . . . . . . . . . . . 20
2.7
Views of the fibreglass wedges in the old anechoic chamber. . . . . . . . . . 21
2.8
Background SPL spectra for old AWT. . . . . . . . . . . . . . . . . . . . . 22
2.9
Background SPL spectra for ISVR anechoic wind tunnel. . . . . . . . . . . 22
3.1
Comparison of predicted and measured SPL spectra for NACA0012 airfoil
in ISVR anechoic wind tunnel. . . . . . . . . . . . . . . . . . . . . . . . . . 25
3.2
Schematic plan view of updates to UTIAS AWT facility. . . . . . . . . . . 26
3.3
Photos of the old bell-mouth under construction. . . . . . . . . . . . . . . 27
3.4
Photo of the tunnel front-end with the resurfaced bell-mouth. . . . . . . . 27
3.5
Close-up scale comparison of the old and new honeycomb mesh screens. . . 28
3.6
Comparison of old and new AWT test section inlets with approximate
shapes of the open jets. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30
3.7
Plot of the new secondary contraction wall profile. . . . . . . . . . . . . . . 31
3.8
Photos of the secondary contraction construction process. . . . . . . . . . . 32
3.9
Views of the new polyurethane wedges in the anechoic chamber. . . . . . . 33
3.10 Acoustic absorptivity curve of fibreglass and open-cell polyurethane foam
for wedges. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 34
3.11 Acoustic absorptivity curve of eggcrate foam sheets. . . . . . . . . . . . . . 34
3.12 Chamber construction photo: Beginning to frame the bare concrete chamber. 35
3.13 Chamber construction photo: Putting up the last studs.
vii
. . . . . . . . . . 36
3.14 Chamber construction photo: Sheeting the ceiling and walls. . . . . . . .
3.15 Chamber construction photo: Installing wedges and building frame for false
door. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
3.16 Photo from inside completed anechoic chamber. . . . . . . . . . . . . . .
3.17 Schematic plan view of the front-end acoustic baffles. . . . . . . . . . . .
3.18 Photos of the new front-end acoustic baffles. . . . . . . . . . . . . . . . .
4.1
4.2
4.3
4.4
4.5
4.6
4.7
4.8
4.9
4.10
4.11
4.12
4.13
4.14
4.15
4.16
4.17
. 36
.
.
.
.
37
37
38
39
Freestream velocity versus VFD output frequency. . . . . . . . . . . . . . .
Variation in static pressure coefficient along the open-jet centreline. . . . .
Freestream velocity profiles of new AWT. . . . . . . . . . . . . . . . . . . .
Test section flow uniformity versus streamwise location. . . . . . . . . . . .
Power spectral density of streamwise velocity fluctuations. . . . . . . . . .
Comparison of centreline freestream turbulence intensity for old and new
AWT. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Plan view of free-field microphone measurement locations. . . . . . . . . .
Background SPL spectra for new anechoic chamber. . . . . . . . . . . . . .
Comparison of background OASPL of improved UTIAS AWT with select
anechoic wind tunnels around the world. . . . . . . . . . . . . . . . . . . .
A-weighted OASPL versus fan motor rotational speed. . . . . . . . . . . .
Comparison of A-weighted third-octave SPL spectra before and after baffle
attenuation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
Schematic of anechoic chamber for free-field characterization. . . . . . . . .
Chamber free-field characterization traversing 45◦ relative to the centreline.
Chamber free-field characterization traversing 90◦ to the centreline. . . . .
Chamber free-field characterization traversing 135◦ to the centreline. . . . .
Pressure response signal of impulse noise used in reverberation time test. .
Comparison of reverberation times for acoustically-wedged anechoic chamber and bare concrete chamber. . . . . . . . . . . . . . . . . . . . . . . . .
41
42
43
44
46
46
47
49
50
51
53
55
56
57
58
60
61
A.1 Material data sheet for polyurethane foam used for the anechoic wedges in
the improved UTIAS AWT. . . . . . . . . . . . . . . . . . . . . . . . . . . 66
viii
Chapter 1
Introduction
This thesis describes the improvements to the Anechoic Wind Tunnel facility (AWT)
at the University of Toronto Institute for Aerospace Studies (UTIAS) for the benefit of
future researchers interested in experimentally studying aerodynamic noise phenomena.
This first chapter opens with a brief description of the motivation behind this work, then
provides some examples of research projects that could be performed in the improved
AWT, followed by an overview of the presentation of this thesis.
1.1
Motivation
Over the past several decades, environmentally sustainable aviation has become a major topic of discussion among the public, government bodies, the aviation industry, and
academia. In particular, regulations and research have focused mainly on three environmental pollutants: carbon dioxide (CO2 ), nitrogen oxides (NOx ), and aviation noise. Of
these, aviation noise around airports was the first to be regulated by the International
Civil Aviation Organization (ICAO) in 1972 and remains a significant area of consideration by all stakeholders involved in aviation today. Table 1.1 shows a summary of ICAO’s
increasingly stringent noise standards over the last five decades. This owes in part to
the public’s direct perceptibility of and adverse reaction to aviation noise, as indicated
by the historically high proportion of community complaints that are related to airport
noise. Indeed, aircraft noise has various behavioural and physiological impacts including
sleep and work disturbance, resulting in increased stress, lower productivity, and lower
property values in communities surrounding airports [1–3].
With global commercial air travel projected to grow on average by 5% per year over the
next 20 years [5, 6], new airports will have to be built and existing ones expanded. As a
1
2
Chapter 1. Introduction
Table 1.1: Subsonic jet-powered aircraft noise reduction standards set by ICAO [4].
Year
1972
1977
2001
2014
Version
Chapter 2
Chapter 3
Chapter 4
Chapter 14
Standard
Reference level
10 dB cumulative below Ch.2
10 dB cumulative below Ch.3
7 dB cumulative below Ch.4
consequence, it is estimated that the population affected by aircraft-related noise around
airports will rise from 13.7 million in 2005 to 23.0 million people in 2035 [2,3]. To address
this growing problem, the National Aeronautics and Space Administration (NASA) and
the Advisory Council for Aerospace Research in Europe (ACARE) have individually set
targets for noise reduction for the next two to three generations of aircraft. Table 1.2
presents a summary of these targets, along with converted values for comparison to the
ICAO standards.
Table 1.2: Noise emission reduction targets set by ACARE and NASA, and their equivalent values relative to ICAO standards [7].
Year
Target
2020
ACARE Vision 2020:
50% reduction in perceived noise
relative to year-2000 aircraft
(revised to FlightPath 2050)
2025
2035
2050
ICAO Equivalent
40 dB cumulative below Ch.4
NASA N+2:
42 dB cumulative decrease
relative to ICAO Ch.4
(equivalent)
NASA N+3:
71 dB cumulative decrease
relative to ICAO Ch.4
(equivalent)
ACARE FlightPath 2050:
65% reduction in perceived noise
relative to year-2000 aircraft
55 dB cumulative below Ch.4
Historically, the majority of aircraft noise originated from turbojet engine-related sources
such as the compressor and exhaust jet, and hence these formed the main focus of aircraft
Chapter 1. Introduction
3
noise reduction research. With the introduction of bypass turbofan engines in the 1970s
and the subsequent increases in fan bypass ratios and improvements in engine lining and
serrated nozzle technology, aircraft manufacturers have largely been able to achieve the
increasingly stringent ICAO noise regulations without sacrificing significant performance
in other areas. Today, the engine noise levels of an aircraft on approach and landing is
comparable to the noise levels generated by the turbulent flow over airframe components
[8]. However, the current engine noise reduction techniques are yielding diminishing
returns as high bypass ratio engine sizes have reached a ceiling with aerodynamic drag
and ground clearance considerations, among others. With the ambitious noise reduction
goals set by ACARE and NASA for the next three generations of aircraft, airframe noise
sources are now seen as the barrier to further significant improvements in aviation noise,
especially from aircraft flying over airport communities on approach for landing. For this
reason, an increasing number of people in industry and academia alike have turned their
attention to airframe noise reduction research. This thesis seeks to make a contribution to
those efforts by providing a high quality aeroacoustics test facility that is unique among
Canadian academic institutions.
1.2
Airframe Noise Research
The basis for airframe noise research was first formed in the theoretical aeroacoustics work
by Lighthill [9], Curle [10], and Ffowcs Williams [11]. The first efforts to quantify airframe noise came in the form of full-scale flyover measurements in the early 1970’s [12–16],
which were able to isolate the landing gear and high lift devices as the most significant
sources of airframe noise for an aircraft on approach to landing (see Fig. 1.1). In the
four decades since then, researchers have performed numerous empirical investigations of
these airframe components in aeroacoustic wind tunnels to relate the local flow physics
to the observed far-field sound characteristics. For landing gears, these research efforts
include studies on the broadband noise generated by turbulent boundary layer separation
and interaction off major gear components (e.g. main struts, wheels), and the tonal/narrowband noise arising from laminar separation and coherent vortex shedding off smaller
gear components (e.g., smaller support struts, hydraulic lines, etc. with Reynolds number below 500,000 based on diameter) [17–20]. For high-lift devices, examples include
broadband noise studies on the unsteady flow inside slat coves [21,22] and over slat tracks
and flap side-edges [23]. As well, experiments have been conducted to understand the
noise generated by the interaction between the landing gear wake and extended flaps [24].
Researchers have also identified many other sources of airframe noise, such as tonal reso-
Chapter 1. Introduction
4
Figure 1.1: Some of the major sources of airframe noise, including landing gear, slats,
flaps, and wing-tips. Image adapted from [26].
nance from pin-holes and vent cavities [25], and broadband noise from wing tip vortices
and turbulent boundary layers [8].
The majority of these studies were performed in aeroacoustic open-jet wind tunnel facilities, where detailed flow-field characterizations and far-field noise measurements could
occur in tandem. Some of the common measurement techniques include Particle Image
Velocimetry (PIV) for non-intrusive flow-field quantification, hot-wire anemometry for
local velocity spectra, microphone phased arrays to pinpoint physical noise sources, and
free-field microphones to determine model noise spectra and directivity. Using high fidelity test models, the wind tunnel far-field noise measurements could be corrected for
model characteristic dimensions and freestream Mach number to give accurate full scale
noise spectra and directivities, an ability that makes aeroacoustic facilities unique and
critical to airframe noise research. Several semi-empirical noise prediction models do exist for landing gears [27–31], but these are only useful for determining the aerodynamic
noise of major components on traditional gear architectures and depend on educated users
to apply “dressing/complexity” factors to compensate for noise from smaller details. As
well, research is well underway on computational aeroacoustic methods [32–36], but these
computationally expensive noise prediction tools are currently not self-sufficient and still
require extensive empirical validation.
The importance of aeroacoustic wind tunnels in investigating aerodynamic noise generation mechanisms and predicting far-field noise levels has been especially apparent in the
Chapter 1. Introduction
5
recent surge in efforts to develop novel airframe noise reduction methods. With the ambitious noise emission targets described previously, researchers have been experimenting
with a wide range of noise control strategies on a variety of airframe components. Passive control studies on landing gear and high-lift devices have mainly focused on using
add-on treatments. Landing gear add-on fairings have yielded up to 10 dB in noise reduction [37, 38] and splitter plates between major landing gear components have resulted
in up to 4 dB reduction [39, 40]. Flap side-edge fences and brush extensions have shown
5-10 dB improvements [41–43], while slat hook extensions and slat cove fillers, covers, and
strips have individually provided 4-5 dB reductions [44–46]. Some researchers have also
investigated active flow control strategies for noise reduction, including dielectric barrier
discharge plasma actuators and blowing/suction systems. The use of plasma actuators on
landing gear struts have suggested the possibility of up to 13 dB in noise reduction [47–49],
while blowing systems on struts have shown 5-10 dB reductions [50, 51]. Applications of
flap side-edge blowing systems have yielded 3-4 dB reductions [52], while slat cove suction/blowing systems have provided 2-4 dB reductions [53, 54]. Researchers have also
looked at cavity tone suppression using plasma actuators [55, 56].
With the varying levels of success from the airframe noise reduction research to date, it
is clear that there is still a long way to achieving the NASA and ACARE noise reduction
targets. In the future, high quality aeroacoustic wind tunnels will continue to play the
critical role of facilitating interdisciplinary research between airframe aerodynamics and
aeroacoustics, and be in demand for the validation of novel quiet landing gear architectures
and high lift device configurations.
1.3
Outline of Thesis
The UTIAS AWT facility was constructed in 1975 and underwent improvements in 1985
(replacement of honeycomb and wire mesh screens) and 2011 (repair of motor windings,
installation of a Variable Frequency Drive (VFD), and addition of a silencer aft of the
fan/motor). This facility has been used for various studies, including graduate-level research on the acoustic excitation of Tollmien-Schlichting Waves [57] and the aeroacoustics
of landing gears in the wake of propellers [58], as well as many industry-led initiatives.
The purpose of this thesis is to address some key facility deficiencies and improve the
aerodynamic and acoustic performance to keep up with the demands of high quality
aeroacoustics research. Potential future study subjects include landing gear struts and
wheels, slats, flaps, propellers, and turbojet compressor fans, among others.
Chapter 1. Introduction
6
Chapter 2 provides some background information for this thesis, beginning with a survey
of existing anechoic wind tunnels at other institutions, followed by a description of the
existing UTIAS facility. Chapter 3 discusses the specific objectives of this thesis project
and details the design rationale, construction, and installation of the various facility improvements. Chapter 4 describes the experimental procedure and presents the results of
the overall facility characterization. Finally, Chapter 5 summarizes the work of this thesis
and provides a list of recommendations for future consideration.
Chapter 2
Background
The research objectives of aeroacoustic experiments are often to characterize, understand,
and predict the aerodynamic generation of noise on a model of interest. These pose many
requirements on the capabilities of the research facility, the chief of which is the ability
to conduct high quality aerodynamic and acoustic investigations simultaneously. This
chapter discusses the rationale behind the design of anechoic wind tunnels and provides
a brief survey of some of the aeroacoustic research facilities around the world, including
the existing one at UTIAS at the time this thesis project began.
2.1
Anechoic Wind Tunnel Design
Aeroacoustic test facilities can come in many shapes and sizes, ranging from small anechoic rooms to hemi-anechoic tunnels to hangar-like facilities (see Fig. 2.1). Their general
purpose is to provide a high quality flow region to simulate operating conditions on test
models, within a test environment that is also suitable for measuring the acoustics generated by the interaction between the airflow and the models. This is usually accomplished
with a wind tunnel that has an open-jet test section enclosed within a larger anechoic
chamber. The test models are positioned inside the test section jet flow region along
with the necessary aerodynamics sensors such as pressure transducers and hot-wires. The
acoustic measurement sensors (i.e., microphones) are placed outside the test section flow
region at various specific locations in the surrounding anechoic chamber to collect data
on the far-field directivity, intensity, and frequency spectra of the aerodynamic noise generated by the test model.
Building or improving a dedicated aeroacoustics testing facility involves the consideration
of many aerodynamic- and acoustic-related parameters, most of which derive from the
7
8
Chapter 2. Background
requirements of the future experiments to be conducted in the facility. Within the test
section, aerodynamic specifications such as usable flow area, velocity, uniformity, and
turbulence intensity must be known and appropriate for the experiments at hand. In the
surrounding anechoic chamber, acoustic specifications such as the anechoic cut-off limit,
background noise levels, and useable anechoic volume are key.
(a) Florida State University facility; 1.12 m
× 0.74 m test section. Image from [59].
(b) ISVR facility; 0.45 m × 0.15 m test section. Image from [60].
(c) Daimler Chrysler AAWT hemi-anechoic
facility; 6.9 m × 4.0 m test section. Image
from [61].
(d) German-Dutch DNW-LLF facility; 8.0 m
× 6.0 m test section. Image from [62].
Figure 2.1: Examples of anechoic wind tunnel facilities around the world.
2.1.1
Aerodynamic Specifications
For airframe noise tests in a typical university laboratory setting, it is often difficult to
investigate full size components at real operating conditions due to space and wind tunnel
fan power limitations. As with aerodynamic experiments testing scaled models, aeroacoustic investigations must ensure that the flow regimes in the wind tunnel setting match
those of real operating conditions. These flow regimes can be represented by several
9
Chapter 2. Background
important similarity parameters, including the non-dimensional Reynolds, Strouhal, and
Mach numbers, Re, St, and M a, respectively.
The Reynolds number, the ratio of inertial to viscous forces in the working fluid, is given
by:
U∞ D
Re =
,
(2.1)
ν
where U∞ is the freestream velocity, D is the characteristic model dimension (e.g. airfoil
chord length or cylinder diameter), and ν is the kinematic viscosity of the working fluid.
This similarity parameter is crucial for characterizing the aerodynamic regime around the
test model because it is used to describe the boundary layer, shear layer, and wake dynamics. This information is obviously important in aeroacoustic tests as researchers aim
to understand noise generation mechanisms associated with specific aerodynamic phenomena.
The Strouhal number is a non-dimensional measure of the frequency of oscillatory flow
phenomena, such as bluff body vortex shedding or free jet shear layer instabilities. It is
given by:
fD
,
(2.2)
St =
U∞
where f is the corresponding physical frequency as measured by experimentation microphones, and D is the separating distance between opposing shear layers (i.e. the diameter
of a cylinder/bluff body or of a free jet). The Strouhal number characterizes shedding
regimes, and is used to scale the frequencies generated by shedding phenomena.
The Mach number is the ratio of inertial to elastic forces in the fluid, and is defined as:
Ma =
U∞
,
a
(2.3)
where a is the speed of sound. The importance of the Mach number lies in the dependence
of aerodynamic sound generation intensity on this similarity parameter; the intensity of
aerodynamic noise scales with Mach number to the eighth power for mechanisms related
to flow turbulence (e.g. open-jet shear layers), to the sixth power for fixed surface pressure fluctuations (e.g. periodic vortex shedding off a surface), and to the fourth power for
oscillating surface fluctuations (e.g. cavity resonance) [63].
From the above similarity parameters, D and U∞ are typically the only variables that
can be controlled by the researcher, but both of their upper ends are limited by the test
10
Chapter 2. Background
facility. Usually, new anechoic wind tunnel designers looking to accommodate a variety
of aeroacoustic tests will attempt to maximize the test section area and freestream velocity. Of course, installation and operating costs, as well as background noise levels in the
anechoic chamber, rise with test section area and velocity so consideration must be given
to all of these criteria. For modifications to existing facilities where the tunnel drive fan
and motor will not be changed, as is the case in this thesis, the maximum mass flowrate
through the test section is fixed and hence test section area and maximum velocity are
inversely proportional with one another. Therefore a balance must be struck between the
two, taking into account the types of aeroacoustics research planned for the facility.
With respect to the other major aerodynamics-related parameters, the goal is to create
laminar, uniform and steady test section flow conditions. The spatial uniformity of the
freestream velocity should be maximized across as large portion of the test section area
as possible. A statistical measure of uniformity based on the root-mean-square of local
time-averaged velocity deviations is given by:
qP
n
i (Ui
γ=
− U )2
U
,
(2.4)
where Ui is the local time-averaged velocity and U is the mean test section velocity.
The freestream turbulence intensity (TI) should be minimized across the test section to
simulate the laminar flow experienced in most real flight conditions. It is given by:
TI =
u0rms
,
U∞
(2.5)
where u0rms is the root-mean-square of time-resolved velocity fluctuations at a point in
space.
2.1.2
Acoustic Specifications
While it is theoretically possible to conduct aeroacoustic tests on a model inside typical
hard-walled wind tunnels, there are several key factors that make high quality acoustic
measurements difficult to obtain within these facilities [64]. Specifically, the closed test
section creates a reverberant environment in which individual microphones have difficulty
distinguishing the test model incident noise from reflections off the walls. Placing the
microphones within the closed test section also exposes them to airflow and causes high
levels of microphone self-noise from turbulent flow over the microphone grid and casing,
which in cases of high flow velocities can decrease the signal-to-noise ratio to unusable
11
Chapter 2. Background
levels. As well, there is usually limited space within closed test sections and microphones
will be within the acoustic near-field of the test model, which is unacceptable in the case
that measurements are needed for the study and prediction of far-field noise levels. Nevertheless, some limited forms of aeroacoustics research can be performed inside closed test
section wind tunnels using microphone phased arrays and coherence/correlation analysis
techniques to determine noise source locations and frequency spectra, even in the high
background noise environments [65]. However, noise intensity and directivity, which are
used to predict full-scale far-field noise emission levels, are best evaluated in anechoic
open-jet wind tunnels.
Acoustic researchers often express noise intensity in terms of Sound Pressure Level (SPL)
in the logarithmic unit decibels (dB) via the equation:
SPL = 20 × log10
p0
pref
,
(2.6)
where p0 is the measured fluctuating pressure signal, and pref is a reference pressure level,
which is usually chosen to be 2×10−5 Pa (approximately the lower limit of sensitivity
of the human ear). In spectral analysis, these SPLs correspond to sound intensities over
particular narrow bands of frequency, and may be integrated over a large range of frequencies to yield an overall measure of sound intensity, called Overall Sound Pressure Level
(OASPL). The frequency limits of integration depend on the particular study at hand; a
commonly used range is 20 Hz to 20 kHz, which forms the approximate limits of human
hearing. Prior to integration, researchers may also apply a correction to the SPL spectra
known as A-weighting, which adjusts SPL values for the frequency-dependent loudness
perception levels of the human ear (the ear is most sensitive to noise in the 1 kHz to
6 kHz frequency range). The A-weighting adjustment levels for frequencies from 20 Hz
to 20 kHz are shown in Fig. 2.2.
The purpose of having an anechoic chamber surrounding the open-jet test section is to
provide a non-reverberant, quiet, still-air environment for noise measurement. The anechoic property is crucial for simulating an acoustic free-field environment, and is typically
provided by wedge-shaped sound absorbent material covering all chamber interior surfaces (minus the floor for hemi-anechoic chambers). Materials for wedge construction
include polyurethane foam, fibreglass, rock-fibre, and proprietary acoustically-absorbent
materials, among others. The shape of these wedges are designed such that most incident acoustic waves do not reflect back directly into the test section measurement region;
any sound waves with one-quarter wavelength smaller than the depth of the wedges are
12
Chapter 2. Background
10
SPL attenuation [dB]
0
−10
−20
−30
−40
−50
10
2
10
3
10
4
Frequency [Hz]
Figure 2.2: A plot of the A-weighting SPL across the frequency range 20 Hz to 20 kHz.
Data from [66].
absorbed. That is, the theoretical frequency cutoff limit above which the chamber is
considered ‘anechoic’, is determined from the wedge depth, Lw , via:
fc =
a
.
4 × Lw
(2.7)
Aeroacoustic researchers are often interested in surveying noise frequency spectra for
interesting sound intensity peaks or other phenomena, in search for connections to aerodynamic phenomena. There may also be motivation to determine how the noise under
study may affect humans if it is within their sensible frequency range. The anechoic cutoff
frequency is important in determining the bandwidth of frequencies permitted into this
spectral analysis, and should be made as low as possible. However, increasing wedge depth
increases installation cost and, in the case of retrofitting an existing facility, decreases the
usable anechoic volume as described later in this section.
Even with acoustic reverberations eliminated, high background noise levels in the anechoic
chamber will significantly affect the ability of microphones to discern noise generated by
the test model from that of other sources. The signal-to-noise ratio of measurement microphone systems should be at least 10 dB above the background noise to ensure high
quality sound intensity measurements [67]. Some typical background noise sources include
the wind tunnel drive fan and other components, model and apparatus support stands
in the flow, the open-jet shear layer, vibrating surfaces, and general environmental noise.
Chapter 2. Background
13
It is crucial in the design and construction of anechoic wind tunnels to minimize these
extraneous noise sources so as to maximize the signal-to-noise ratio of the microphone
measurements. Some strategies employed by past researchers to minimize background
noise include baffles and silencers to stop the transmission of fan noise into the test section and heavily insulated jet collectors to decrease the noise generated from turbulent jet
shear layer impingement [59,68,69]. Unwanted vibrations must be eliminated by securing
or stiffening parts, and environmental noise should be sealed out using thick concrete wall
construction for the chamber. As well, model and apparatus support struts that extend
into the test section flow should have cross-sections in the form of the M-H airfoil, which
was specially designed to exhibit minimal shedding (thus low self-noise) while providing
stiffness through a 30% thickness-to-chord ratio [64].
In order to accommodate far-field noise measurements in the anechoic chamber, the physical dimensions of the chamber must be large enough to accommodate an arc of far-field
microphones positioned around the test model. These microphones must be in the acoustic
far-field of the model, which is defined as the region that is at least one acoustic wavelength
and two noise source dimensions away from the test model [64]. These microphones must
be within the usable anechoic volume of the facility, away from the acoustically absorbent
wedges on the walls to avoid interference with the noise measurements. The angular position of the measurement arc would depend on the test at hand, but typical arcs may
extend from 20◦ to 135◦ relative to the jet centreline downstream of the test model.
One more acoustic specification that some, but not all, researchers quote is RT60 , defined
as the time it takes for the SPL of an impulse noise and subsequent reverberations to
decrease by 60 dB from the peak of the impulse. This is a measure used mostly in regular
rooms and concert halls, and is related to the size and shape of the room and the quality
of the anechoic treatment on surfaces. Lowering reverberation time means improving the
anechoic property of the test facility.
2.2
Survey of Existing Facilities
Anechoic wind tunnel facilities are used in various settings around the world, including
at aircraft and auto manufacturer test facilities, national research centres, and universities. These facilities can be categorized into open- and closed-circuit types. Open-circuit
tunnels draw in and expel out air to the surrounding environment, while closed-circuit
tunnels continuously recirculate air with limited exchange with the surrounding envi-
Chapter 2. Background
14
ronment. Open-circuit tunnels require less space and incur substantially lower costs to
construct. Closed-circuit tunnels have more parts in the form of turning vanes and air
return sections, but are able to offer higher flow quality and achieve test section velocities
equivalent to open-circuit tunnels using less power and generating less background noise.
Tests involving combustion or smoke visualization should be performed in open-circuit
tunnels.
The type of tunnel driver also has a profound impact on tunnel operation, and existing
facilities can be categorized into blowdown, blower, or suction driver types. Blowdown
tunnels supply air from a pressurized reservoir, while the blowing and suction type tunnels drive air continuously using a fan upstream and downstream of the test section,
respectively. The blowdown tunnels are limited in test runtime, but usually offer superior
background noise levels to the continuously driven tunnels because there is no fan and
motor noise. Blower type open-jet test sections have inferior flow quality in the test section and may exhibit low-frequency flow instabilities from collector spillage. Suction type
open-jet test sections operate below atmospheric pressure, but exhibit less jet spreading
and offer higher flow quality than the blower type tunnels do [70]. For the tunnels with
variable pitch fans and variable rotational speed motors, electrical power consumption
and background noise may be improved via an adjustment of these settings for the lowest
motor RPM [64].
In the past two decades, a handful of new facilities have been built and existing ones
improved [67,68]. The auto industry has taken the lead in constructing larger and quieter
facilities, due in large part to their ability to pay for the enormous costs of installation
and operation. For example, the Daimler Chrysler Aeroacoustic Wind Tunnel (AAWT)
in Detroit, U.S.A., which has one of the largest test section areas among anechoic wind
tunnels at 28 m2 and is capable of velocities up to 71 m/s, cost US$37.5 million to construct [71]. Universities have also pushed the envelope with new aeroacoustic test facilities
that have impressive specifications, albeit on a smaller scale. An example is the tunnel
at the Institute for Sound and Vibration Research in Southampton, U.K., which offers
a maximum test section velocity of 100 m/s within an anechoic chamber that has one
of the lowest background noise levels in the world [60]. Table 2.1 compares key facility
dimensions and specifications for the aerodynamic and acoustic capabilities of leading
facilities around the world.
Tunnel type
Open-jet,
closed-circuit
Open-jet,
closed-circuit
Open-jet ,
closed-circuit
Open-jet,
closed-circuit
Hessert Center,
Open-jet,
Notre Dame
open-circuit, suction
University of
Open-jet,
Florida
open-circuit, suction
ISVR
Open-jet,
Southampton
blowdown
UTIAS
Open-jet,
(old facility)
open-circuit, suction
Florida State
Open-jet,
University
open-circuit, suction
Virginia Tech
Kevlar wall,
closed-circuit
Brandenburg Univ. Open-jet
of Tech. at Cottbus open-circuit, blower
Universitat
Open-jet,
Siegen
open-circuit, blower
NASA Langley
Quiet Flow Facility
Ford DTF
WT8
Daimler Chrysler
AAWT
Audi
Facility name
Maximum
velocity
[m/s]
58
54
71
83
28
76
100
60
75
57
72
24
Test section
dimensions
[m×m]
0.91 × 0.61
5.56 × 3.34
6.9 × 4.0
3.94 × 2.8
0.61 × 0.61
1.12 × 0.74
0.45 × 0.15
0.7 diameter
1.22 × 0.91
1.83 × 1.83
0.15 × 0.12
0.13 × 0.13
0.40
0.20
0.03
0.12
0.5
0.1
0.07
0.04
0.30
0.16
0.34
4.5 × 3.2 × 2.9
5.0 × 4.5 × 3.9
7.0 × 4.2 × 2.4
4.8 × 4.5 × 2.7
7.33 × 7.33
× 5.50
4.9 × 4.9 × 2.4
5.5 × 5.0 × 2.3
9.1 × 7.3 × 3.7
-
21 × 15 × 10
-
Turbulence
Chamber
intensity
dimensions
[%]
[m×m×m]
9.1 × 6.1 × 7.6
125
500
190
250
140
80
100
125
-
-
-
Anecoic freq.
cut-off
[Hz]
-
Chapter 2. Background
15
Table 2.1: Survey of select aeroacoustic research facilities around the world. Data from
[59, 60, 73–76].
16
Chapter 2. Background
In order for background noise level data from all the different facilities to be comparable,
they must be converted to a common basis with respect to microphone measurement
distance from the test section centreline as well as the open-jet cross-sectional area. The
conversion is given by [60]:
OASPLconv = OASPLfac + 20 × log10
rfac
rref
− 10 × log10
Afac
Aref
,
(2.8)
where OASPLfac is the reported background sound pressure level, rfac is the original microphone measurement distance from the open-jet centreline, rref is the reference microphone
distance from the open-jet centreline (set at 2 m in this case), Afac is the open-jet crosssectional area of the facility, and Aref is the reference cross-sectional area (set at the new
test section inlet area of 0.36 m2 , for ease of comparison with results later). Figure 2.3
summarizes the background noise levels over a range of test section velocities for the
aeroacoustic facilities of interest.
Background A−weighted OASPL [dB]
80
70
60
Audi
DaimlerChrysler
Ford DTF
Florida State
ISVR Southampton
Univ. Cottbus
Univ. Florida
Univ. Siegen
Univ. Notre Dame
Old UTIAS
50
40
30
20
0
10
20
30
40
50
60
Test section velocity [m/s]
70
80
90
100
Figure 2.3: Comparison of overall background sound pressure levels of select anechoic
wind tunnels around the world, over their range of capable test section velocities. Data
adapted from [59, 60, 72–74].
Chapter 2. Background
2.3
17
Existing UTIAS facility
The existing aeroacoustic research facility at UTIAS consists of an open-circuit, suctiontype wind tunnel with an open-jet test section enclosed inside an anechoic chamber. The
tunnel is driven by a nine-blade axial fan powered by a 150 h.p. A/C induction motor
that is capable of rotational speeds up to 1800 RPM. The test section inlet diameter is
0.70 m, with a maximum velocity of 60 m/s and freestream turbulence intensity of less
than 0.5%. The anechoic chamber is lined with fibreglass wedges with tip-to-base depth
of 0.61 m, yielding a theoretical anechoic cut-off frequency of 140 Hz. The chamber inner
dimensions as measured from wedge-tip to wedge-tip are 4.9 m × 4.9 m × 2.4 m (L × W
× H).
The wind tunnel is situated outdoors, with its front-end attached to the side of a geodesic
dome that is 55 m in diameter. Fig. 2.4 provides views of the existing facility from outside
and inside the geodesic dome, as well as a schematic plan view. The geodesic dome is used
by the UTIAS robotics research groups as a Martian surface simulator, and also doubles
as a sort of supply plenum for the wind tunnel. Air from inside the dome is guided into
the tunnel by a square bell-mouth section measuring 2.74 m by 2.74 m and then sent
through a series of flow conditioning screens. Unfortunately, activity in the dome creates
large amounts of airborne dust, which over time clogs up the flow conditioning screens
and renders them ineffective. The fine dust particles that are not caught in the screens enter the test section, where they may accumulate on and damage sensitive instrumentation.
The first of the flow conditioning screens is an aluminum honeycomb mesh with hexagonal
cells of 1.59 mm width and 50 mm depth used to straighten the flow. This honeycomb
screen is bound by two 120-mesh (120 openings per inch length) stainless steel wire screens
to keep it in place. The air then passes through a 200 mm settling passage before entering a series of 4 stainless steel, 54-mesh, 0.10 mm-diameter wire screens, with 50 mm
streamwise separation between each of them. The purpose of these wire screens is to reduce freestream turbulence intensity and promote spatial flow uniformity. The issue with
these existing honeycomb and wire mesh screens is that their physical specifications are
too small by the standards of generally accepted wind tunnel design rules. As a result,
they are less effective at reducing freestream turbulence intensity and incur excess static
pressure drops, which negatively impact the maximum achievable test section velocity.
After the conditioning screens, the flow accelerates through a square contraction section
with a contraction ratio of 9.3, followed by a square-to-circle conversion section. After
18
Chapter 2. Background
(a) Outdoors view of the faciity.
(b) Inside the geodesic dome looking at the
existing tunnel bellmouth inlet.
(c) Schematic plan view.
Figure 2.4: Views of the existing UTIAS AWT facility.
19
Chapter 2. Background
a circular tube section, the flow accelerates through a cone-shaped nozzle that sharply
converges the flow from 0.9 m to 0.7 m in diameter just before entering the test section
inside the anechoic chamber. The open-jet section from nozzle exit to collector entrance
is 2.78 m long. After entering the 0.9 m diameter collector, the flow passes through diffuser and silencer sections, followed by the driving fan section and finally out into the
atmosphere. An aerodynamic characterization of the existing test section open-jet was
performed by Chekiri [58]. Figure 2.6 shows time-averaged freestream velocity surveys
across the diameter of the open-jet, for test section streamwise locations from 0 mm to
1190 mm (i.e., from 0 to 1.70 nozzle diameters). The wake surveys of the first several
streamwise locations clearly show accelerated regions near the shear layers at the edges
of the jet, and this is due to the contracting nozzle just before the entrance of the test
section. As a result, the spatially uniform flow velocity area is reduced, and the usable
test section diameter is restricted to 0.5 m. Figure 2.5a shows the freestream turbulence
intensity along the jet centreline for freestream Mach numbers 0.10 and 0.15. Figure 2.5b
shows the variation of freestream turbulence intensity with radial position in the openjet, for various streamwise positions in the test section at freestream Mach of 0.10. It is
clear that as the flow moves downstream from the nozzle exit, the converging open-jet
shear layer causes a significant increase in freestream turbulence intensity in the test section. Improving the test section area, flow velocity, and flow quality are crucial to future
experimental research in the facility and thus is one of the main areas of focus in this thesis.
(a) Centreline TI versus streamwise location at
freestream Mach 0.15 and 0.10.
(b) TI versus radial location at various streamwise locations at freestream Mach 0.10.
Figure 2.5: Freestream turbulence intensity plots for old AWT. Figures from [58].
Chapter 2. Background
20
Figure 2.6: Test section time-averaged velocity surveys over a range of streamwise locations, at freestream velocities of (a) Mach 0.10 and (b) Mach 0.15 for old AWT. Figure
from [58].
21
Chapter 2. Background
The anechoic chamber housing the test section was constructed from 300 mm thick concrete walls, and all hard interior surfaces were covered in wedge-like fibreglass blocks
(see Fig. 2.7). These wedge blocks were constructed from individual 51 mm thick yellow fibreglass batts friction-fit into “step-function” wedge-like shapes, instead of the more
commonly installed triangular prism wedge shape. The fibreglass wedges were not secured
by any form of metal mesh or cloth material and over time many of them lost their form
or completely dislodged from the ceiling and walls due to gravity and perturbations from
recirculating chamber flow during tunnel operation. As well, the bare fibreglass wedges
released many loose fibres into the air, which posed a serious health risk to researchers
who worked inside the chamber.
(a) Inside the anechoic chamber looking at
the test section.
(b) Drawing of two pairs of fibreglass wedges.
Dimensions in centimetres. Figure from [58].
Figure 2.7: Views of the fibreglass wedges in the old anechoic chamber.
Some background SPL characterization was performed on the old anechoic chamber and is
summarized in the SPL spectra in Fig. 2.8. For comparison, the background SPL spectra
of the ISVR anechoic wind tunnel is given in Fig. 2.9. At the maximum freestream velocity
shown for the existing UTIAS facility, 61.2 m/s, the background noise levels in the UTIAS
facility are between 15-25 dB higher than those in the ISVR facility across the range of
frequencies from 100 Hz to 10,000 Hz. This relatively poor acoustic performance in the
existing tunnel limits the signal-to-noise ratio of microphone measurements and hampers
the ability of researchers to collect high quality acoustic data. For this reason, chamber
acoustics is another main area of focus in the current facility improvement efforts.
22
Chapter 2. Background
80
19.0 m/s
29.6 m/s
37.7 m/s
44.9 m/s
48.3 m/s
54.4 m/s
61.2 m/s
60
SPL [dB ref 2x10
−5
Pa]
70
50
40
30
20
10 2
10
10
3
10
4
Frequency [Hz]
Figure 2.8: Background SPL narrowband frequency spectra of old AWT, for select
freestream velocities from 19.0 m/s to 61.2 m/s. Data adapted from [58].
Figure 2.9: Background SPL narrowband frequency spectra of ISVR anechoic wind tunnel,
for select freestream velocities from 33.1 m/s to 99.6 m/s. Figure from [60].
Chapter 3
Facility Updates
3.1
Objectives
The specific thesis objectives and associated performance metrics are discussed below.
One of the overarching objectives was to keep the total cost below CA$40,000.
3.1.1
Aerodynamic objectives
Table 3.1 lists the aerodynamics-related targets of the facility update efforts. The test
section freestream velocity and dimension targets were driven by minimum Reynolds number requirements and maximum model flow blockage limits. The freestream turbulence
intensity and spatial uniformity targets were motivated by the desire to provide as high
quality flow in the test section as possible.
Table 3.1: Summary of aerodynamic improvement objectives for the UTIAS AWT facility.
Objective
Target
Test section maximum
freestream velocity
Increase to at least 70 m/s from current 60 m/s
Test section
dimensions
The usable cross-sectional area should be at least 0.5 m
in diameter and the streamwise length be at least 1.2 m
Test section freestream Below current value of 0.5% within first 1.5 test
turbulence intensity
section widths downstream of inlet
Test section spatial
flow uniformity
At least 99%
23
24
Chapter 3. Facility Updates
3.1.2
Acoustic objectives
Table 3.2 lists the acoustics-related targets of the facility update efforts.
Table 3.2: Summary of acoustic improvement objectives for the UTIAS AWT facility.
Objective
Target
Background noise level
15 dB SPL reduction from current levels
across frequencies 200 Hz - 20 kHz
Anechoic limit
cut-off frequency
Below 200 Hz
Usable anechoic area
around test section
At least enough space to place an arc of
microphones 2 m from the jet centreline
Reverberation time
Below 0.2 s to 60 dB SPL decay
The expected experimental microphone measurements across the frequency bandwidths
of interest must achieve a signal-to-noise ratio of at least 10 dB above the tunnel background noise level. The NACA0012 airfoil self-noise is commonly used as a benchmark
to approximate experimental noise levels, and a semi-empirical noise prediction method
for this airfoil was developed by Brooks et al. [77]. Figure 3.1 shows both predicted and
experimental (from ISVR) noise spectra for a NACA0012 airfoil with chord length 0.15 m
at velocities 33.3 m/s and 80 m/s, values which approximate the freestream velocity lower
and upper bounds for future UTIAS AWT experiments. Comparisons between these airfoil noise spectra and the UTIAS AWT background SPL spectra in Fig. 2.8 reveals that
the UTIAS AWT requires a large background noise level improvement of around 15 dB
across all frequencies in order to achieve the 10 dB target for signal-to-noise ratio.
The target for the anechoic cut-off frequency is controlled by the depth of the acoustic
wedges, which was determined with primary consideration given to the space available
inside the existing anechoic chamber and the cost of acoustic wedges (they were by far
the most expensive item on the budget). The upper limit of 200 Hz is within the range
of other facilities around the world, and provides adequate bandwidth for future scaledmodel aeroacoustic tests in the facility. The anechoic space requirement was set to ensure that an arc of free-field microphones can be situated in the acoustic far-field of any
aerodynamic noise with frequencies above 200 Hz generated from sources within the test
section. Finally, the reverberation time should be as low as possible, with 0.2 s considered
Chapter 3. Facility Updates
25
Figure 3.1: Comparison of predicted and measured narrowband SPL spectra for
NACA0012 airfoil with chord length 0.15 m at freestream velocities of 33.3 m/s and
80 m/s in the ISVR anechoic wind tunnel. Also shown is the ISVR background narrowband SPL at 33.3 m/s and 80 m/s. Figure from [60].
a reasonable upper limit for anechoic chambers of this size. In comparison, the bare concrete room housing the anechoic chamber has a reverberation time of approximately 1.5 s.
In order to achieve these thesis objectives, several key areas in the UTIAS AWT facility
were targeted for improvement (see Fig. 3.2). First, the bell-mouth intake surfaces were
refinished and the flow conditioning screens were replaced. Second, the square-to-circle
flow section and nozzle were replaced with a square secondary contraction. Third, the
fibreglass wedges in the test chamber were replaced with polyurethane foam wedges, and a
false door was installed inside the existing chamber entrance. Finally, baffles were added
to the front-end of the facility. The following sections describe in detail the rationale,
design, and implementation of these facility updates.
Chapter 3. Facility Updates
26
Figure 3.2: Schematic plan view of updates to UTIAS AWT facility. Front-end baffles
were added, flow conditioning screens were replaced, existing nozzle was replaced with a
secondary contraction, and new anechoic wedges and false door were installed.
3.2
Flow Conditioning
The wind tunnel front-end, which consists of a bell-mouth, a honeycomb screen, and three
wire mesh screens, performs important flow intake and conditioning functions on the air
heading into the test section. The bell-mouth can be separated into four corner parts and
is made of painted steel sheets that are about 2 mm thick (see Fig. 3.3). The surfaces of
the old bell-mouth pieces were very rough as a result of paint stripping and severe rusting. As a first step towards improving the flow conditioning, these surfaces were ground
down to the bare steel and then repainted with three coats of Tremclad exterior rust-proof
paint. Figure. 3.4 shows an image of the front-end with the resurfaced bell-mouth.
The honeycomb screen primarily serves to straighten the flow going into the test section.
This is achieved as long as there are at least 150 cell openings across the width of the
flow section [78]. As well, the honeycomb screen can help reduce freestream turbulence
intensity if the cell depths are approximately 6-8 cell diameters. In order to minimize the
pressure drop, the cross-sectional shape for each cell should be hexagonal [79]. Following
these guidelines, the new honeycomb design has aluminum hexagonal cells with 9.53 mm
cell width and 76.2 mm cell depth. The honeycomb screen was installed with wire mesh
screens (discussed later) attached on both sides to keep it in place. Figure 3.5 provides a
scale comparison of the old honeycomb cells with the new ones.
Chapter 3. Facility Updates
Figure 3.3: Photos of the old bell-mouth under construction.
Figure 3.4: Photo of the tunnel front-end with the resurfaced bell-mouth.
27
28
Chapter 3. Facility Updates
(a) Old honeycomb screen with 1.59 mm cell
width and 50.8 mm cell depth.
(b) New honeycomb screen with 9.53 mm cell
width and 76.2 mm cell depth. Photo taken
with 18-mesh, 0.229 mm diameter wire screen
in front of honeycomb.
Figure 3.5: Close-up scale comparison of the old and new honeycomb mesh screens.
The purpose of the wire mesh screens is to cause a static pressure drop that makes the
flow velocity profile more spatially uniform and reduces boundary layer growth [78]. The
freestream turbulence intensity can also be reduced, given a minimum settling spacing of
500 wire diameters downstream of each screen to allow the screen-generated turbulence
to dissipate [80]. Mesh screens are characterized by their open-area ratio,
β = (1 − M d)2 ,
(3.1)
where M is the number of mesh openings per unit length and d is the wire diameter, and
by their pressure drop coefficient as formulated by Wieghardt’s empirical formula [81],
− 1
∆P
1 − β Ud 3
K=
= 6.5
,
q
β
βν
(3.2)
where ∆P is the static pressure drop, q is dynamic pressure, U is freestream velocity entering the screen, and ν is the kinematic viscosity. Wind tunnel design guidelines [78, 82]
suggest β > 57% to avoid instabilities associated with jets forming from the individual
screen openings. As well, a series of 3-5 screens with K = 1.6 − 2.0 for each is suggested
to achieve flow uniformity while maintaining an acceptable overall pressure drop. The
maximum streamwise velocity of air entering the bell-mouth inlet was measured to be
approximately 3.7 m/s.
29
Chapter 3. Facility Updates
The old set of screens consisted of M = 120-mesh preliminary screens on both sides of
the honeycomb grid, followed 200 mm downstream by four 54-mesh main screens spaced
50 mm apart from each other. This old set-up needed to be replaced because the low
open-area ratio of the 120-mesh screens compounded with the heavy accumulation of dirt
on all of the screens caused severe flow blockage. Following the design guidelines, a new
set of five screens were made. Two 18-mesh, 0.229 mm diameter wire preliminary screens
were installed on both sides of the honeycomb, followed 200 mm downstream by three
34-mesh, 0.165 mm diameter wire main screens spaced 76 mm apart from each other.
Table 3.3 lists the specifications of the old and new screens.
Table 3.3: Comparison of new and old wire mesh screen specifications.
Screen type
3.3
M (per inch)
d [mm]
β [%]
K
New preliminary
18
0.229
70.2
0.92
Old preliminary
120
0.067
47.0
4.83
New main
34
0.165
60.7
1.73
Old main
54
0.086
66.8
1.55
Secondary Contraction
The facility has a square primary contraction situated in the concrete wall separating the
anechoic chamber from the interior of the geodesic dome. In the old set-up, the primary
contraction connected to a square-to-circle conversion section, followed by a straight circular section, and capped off by a conic nozzle at the test section inlet. This nozzle produced
an undesirable region of accelerated flow near the shear layers, with a contracted uniform
velocity test section area about 500 mm in diameter. The new design maintains a square
cross-section throughout; the primary contraction is followed by a straight square section,
which is in turn followed by a secondary contraction that ends in a tangential exit to get
the desired “top-hat” velocity profile in the test section. Figure 3.6 shows images of the
old nozzle and new secondary contraction sections.
There are several important factors to consider in the design of a wind tunnel contraction,
including flow separation, crossflow, and flow uniformity [75, 82, 83]. Flow separation at
the entrance of the contraction can be caused by an abrupt increase in the wall pressure
Chapter 3. Facility Updates
30
(a) Old upstream flow section with square-to-circle conversion, then circular straight
section, capped off with conic nozzle at test section inlet.
(b) New upstream flow section with straight square section followed by secondary
contraction with tangential exit lip at test section inlet.
Figure 3.6: On the left, comparison of new and old flow sections leading to the test section.
On the right, renders of the shear layer and test section velocity profiles expected from
the two different test section inlets.
coefficient in the convex (fore) section of the contraction. This can be avoided by using
a longer convex section and a lower order polynomial to shape the walls in that section.
Crossflow occurs due to having a non-axisymmetric cross-section, but this can be reduced
via a longer contraction length and having a square cross-section. Since the contraction
is the last tunnel component that the flow passes through before entering the test section,
it is critical that flow uniformity at the contraction exit be maximized. This can be done
using a higher order polynomial to shape the walls in the concave (aft) section of the
contraction, and ensuring the exit lip is tangential to the desired direction of the test
section freestream flow.
31
Chapter 3. Facility Updates
The wall profile chosen for the secondary contraction consists of two matched-polynomial
curves, with a 3rd order polynomial for the fore section and a 8th order polynomial for
the aft section (see Fig. 3.7). The convex- and concave-section match point, where the
first- and second-derivatives of the polynomial curves were set equal to each other, was
set at 70% of the contraction length. The contraction ratio for this new section alone is
1.91, while the overall ratio for the primary and secondary contractions combined is 20.9.
A side benefit of this relatively high contraction ratio is a reduction in the longitudinal
freestream turbulence intensity going into the test section.
500
Wind tunnel height/width [mm]
400
300
200
100
0
-100
-200
-300
-400
-500
0
200
400
600
800
1000
Streamwise location [mm]
Figure 3.7: Plot of the secondary contraction wall profile. Each curve is constructed from
3rd and 7th order polynomials with the matching point at the 70% streamwise location.
The secondary contraction section was manufactured in-house via a fibreglass hand-layup
method on a positive mould. This mould consisted of a wooden frame supporting Owens
Corning extruded polystyrene rigid foam insulation sheets, which were formed into the
contraction wall profile using a computer-guided hot-wire foam cutter. The formed foam
sheets were sealed with three coats of a thick epoxy mixture, consisting of West Systems 105 Resin, 205 Fast Hardener, and 406 Colloidal Silica, to yield a hard and smooth
moulding surface. To help with mould release, two layers of TR104 carnauba blend wax
were applied to the mould surface using a soft cloth, followed by a light spray of Polytek
Pol-Ease 2300 multipurpose silicone release agent. The actual contraction section walls
were made from a sandwich composite of 5.8 oz. woven fibreglass ‘E’ cloths and 1.59 mm
32
Chapter 3. Facility Updates
thick balsa sheets. These were formed on the mould by first curing two layers of thick
epoxy mixture, followed by a fibreglass cloth and epoxy, then a balsa sheet and epoxy, and
then another fibreglass cloth and epoxy, and then sealed with two layers of epoxy. Two
100 mm wide, 12.7 mm thick plywood ribs in the shape of the contraction profile were
epoxied onto the outer surface of each wall for structural stiffening. As well, a 50 mm
wide, 12.7 mm thick plywood flange was epoxied onto the downstream opening of the
contraction to reinforce the square shape and provide a smooth tangential exit into the
test section. Finally, a 100 mm wide, 3.18 mm thick steel flange was epoxied onto the
upstream opening of the contraction so that it could be bolted onto the square flow section
extending from the primary contraction. Figure 3.8 shows some images of the secondary
contraction construction process.
(a) Surface polishing for half the mould.
(b) Full fibreglass and balsa contraction on
mould with metal flange on back. Reinforcing
joints with fibreglass strips.
Figure 3.8: Photos of the secondary contraction construction process.
3.4
Acoustic Wedges
Covering the walls with anechoic wedges serves to minimize acoustic reverberation in the
chamber. This is achieved via two mechanisms: one is the pseudo-random wall geometry
that minimizes the reflection of sound waves back into the chamber, and the other is the
large surface area of anechoic material that absorbs the sound waves. Fig. 3.9a shows the
new anechoic wedges fully installed in the chamber.
33
Chapter 3. Facility Updates
(a) Photo of inside the anechoic chamber looking at the test section inlet.
(b) Drawing of two pairs of polyurethane
wedges. Dimensions in milimetres.
Figure 3.9: Views of the new polyurethane wedges in the anechoic chamber.
The shape and dimension of the wedges depend on several design factors. Ideally, the
shape would be conical so that there are no parallel surfaces to produce acoustic reflections.
However, due to wedge stiffness, cost, manufacturing, and installation considerations, a
triangular prism shape was chosen (see Fig. 3.9b). The wedge dimensions were chosen to
minimize the anechoic cut-off frequency, while satisfying the anechoic volume and material cost constraints. With the chosen wedge dimensions shown in Fig. 3.9b, the anechoic
chamber dimensions ended up being 4.88 m × 4.88 m × 1.83 m as measured from wedgetip to wedge-tip, and the theoretical anechoic cut-off frequency is approximately 150 Hz.
The acoustic absorption properties of the wedge material can be described by its sound
absorption coefficient, which is defined as the ratio of absorbed sound energy to incident
sound energy on the material surface [84]. Anechoic wedges are commonly made of porous
materials due to their ability to absorb incident sound waves by incurring viscous and
momentum losses on the oscillating air molecules within the numerous irregular pores in
the material [85]. A polyurethane open-cell foam was sourced from Foamite, a local foam
mattress manufacturer, for the new wedges (see Appendix A for a material data sheet
for this foam). This foam material poses a much lower health risk because it does not
shed minuscule fibres into the air. An acoustic absorptivity test conducted at Bombardier
Inc. by Raymond Wong using the Impedence Tube Method (ASTM C384) found that the
foam material has a superior acoustic absorption coefficient compared to the old fibreglass
material (see Fig. 3.10). Polyurethane open-cell foam sheets with a thickness of 38 mm
were also purchased to line the chamber wall edges, wind tunnel flow sections, and other
Chapter 3. Facility Updates
34
hard surfaces in the chamber that could not be covered with wedges. These foam sheets
had ‘egg-crate’-shaped surfaces to help reduce acoustic reverberation (see Fig. 3.11 for
the absorption coefficient spectrum).
Figure 3.10: Acoustic absorptivity spectra for fibreglass and open-cell polyurethane foam.
Data acquired by Bombardier using Impedance Tube Method (ASTM C384).
Figure 3.11: Acoustic absorptivity spectrum for open-cell polyurethane foam eggcrate
sheet, 38 mm thickness. Data from [86] using Reverberation Room Method (ASTM
C423).
Chapter 3. Facility Updates
35
Before installing the anechoic wedges, the bare anechoic chamber was first framed with
beams across the ceiling and studs along the walls (see Figs. 3.12-3.13). All contact
points between wood and concrete were protected by plastic vapour barrier sheets. The
wall studs were secured at their top- and base-plates, with an approximately 15 mm air
gap between the studs and concrete walls to prevent vibration noise. The ceiling was
then sheeted with 6.35 mm thick oriented strand board and the walls with 6.35 mm thick
plywood strips (see Fig. 3.14). Finally, the anechoic wedges were attached to the sheeted
ceiling and walls, and the eggcrate foam sheets were attached to the remaining exposed
surfaces, both using 3M Super-77 spray adhesive (see Fig. 3.15).
A false door covered with wedges was also installed just inside of the existing metal door
entranceway. In comparison to the old design that only had one layer of fibreglass covering
the metal door, the false door provides additional acoustic attenuation of outside noise
sources as well as decreased acoustic reverberation from inside noise sources. This false
door swings inward and has wedges positioned on it such that they interlock with wedges
on the wall when the door is fully opened (see Fig. 3.16).
Figure 3.12: Chamber construction photo: Beginning to frame the bare concrete chamber.
Chapter 3. Facility Updates
Figure 3.13: Putting up the last studs.
Figure 3.14: Sheeting the ceiling and walls.
36
Chapter 3. Facility Updates
Figure 3.15: Installing wedges and building frame for false door.
Figure 3.16: View of the completed chamber and false door.
37
Chapter 3. Facility Updates
3.5
38
Front-end Baffles
The final major addition to the aeroacoustic facility was the set of acoustic baffles installed
upstream of the wind tunnel bellmouth inside the geodesic dome. These baffles serve as
silencers by blocking direct line of sight from the interior of the geodesic dome into the
anechoic chamber. Seven baffles were installed in a staggered manner (see Fig. 3.17) to
maximize the total airflow area into the tunnel bellmouth. Each baffle is 100 mm thick
and extends from the gravel floor to the ceiling of the geodesic dome. They consist of
wooden frames filled with 100 mm thick fibreglass batts that were recycled from the old
anechoic chamber, and covered in black broadcloth to keep the fibreglass in place. Figure 3.18 shows the completed set of baffles.
Figure 3.17: Schematic plan view of the front-end acoustic baffles. Dimensions are in
metres.
39
Chapter 3. Facility Updates
(a) Left-hand side view.
(b) Head-on view.
(c) Right-hand side view.
Figure 3.18: Photos of the new front-end acoustic baffles as viewed from inside the geodesic
dome looking towards the wind tunnel inlet.
Chapter 4
Facility Characterization
4.1
Flow Speed
The wind tunnel test section freestream velocity was calibrated against the frequency
output of the tunnel motor VFD. The VFD was controlled via an analog voltage output
from a National Instruments (NI) BNC-2120 terminal block connected to a National Instruments PCIe-6361 data acquisition card (DAQ). This voltage signal was in the range of
0-10 V, corresponding linearly to the limits of the VFD frequency output range of 0-60 Hz.
The frequency output is also controllable via a keypad terminal on the VFD unit. The
freestream flow velocity was measured using a 3 mm diameter pitot-static tube situated
at the centre of the test section inlet plane. This pitot tube was connected via Tygon
tubing to a MKS Baratron 120AD differential pressure transducer, which has a maximum
measurement range of 100 torr with an accuracy of 0.25% of reading. Freestream air
temperature was acquired using a T-type thermocouple. All signals were acquired by the
NI PCIe-6361 DAQ via a BNC-2120 terminal block.
The wind tunnel axial fan has a hub diameter of 1.38 m and a tip diameter of 2.10 m.
There are a total of 9 blades with hub-to-tip twist angles of approximately 20◦ . Each blade
can be individually set to equally-spaced pitch angle settings labelled number “-3” to “6”,
corresponding roughly to pitch angles of -12◦ to 42◦ at three-quarters blade radius relative
to the fan plane. The freestream velocities were measured at the range of blade angle
settings between -3 and 5. Setting number 6 could not be physically obtained because
the trailing edges of the blades come into contact with the motor housing behind the fan
hub. The differences in freestream velocity between the different blade angle settings were
negligible, but the motor rotational rate did decrease by 2% at the highest blade angle
compared to the lowest angle. This would result in slightly lower power consumption when
40
41
Chapter 4. Facility Characterization
Mean freestream velocity [m/s]
operating the wind tunnel, as well as slightly lower fan and motor noise levels transmitted
into the anechoic chamber (discussed later). Thus, the fan blades were fixed at setting
number 5 for the remainder of the tests in this thesis. Fig. 4.1 shows the calibration points
and linearly fitted curve for the test section freestream velocity.
80
60
40
20
0
0
10
20
30
40
50
60
VFD setting [Hz]
Figure 4.1: Freestream velocity versus VFD output frequency for fan blades at angle setting number ‘5’. Equation for line of best fit is: U∞ = (1.27 m s−1 Hz−1 )VFD−0.14 m s−1 ,
with R2 of the linear fit at 0.9998.
4.2
Flow Uniformity
The spatial uniformity of the test section freestream flow was characterized using the
flow velocity measurement system described above. The pitot-static tube in this case was
mounted on top of a two-axis linear traverse system capable of placing the probe tip anywhere within an area from 0 mm to 2200 mm downstream of the test section inlet and up
to 400 mm cross-stream distance from the test section centreline. The two traverses were
driven by stepper motors powered by a Motion Group MMC 4S-SC Motor Controller.
The spatial resolution of the streamwise traverse was approximately 0.032 mm/full motor
step, while that of the cross-stream traverse was approximately 0.0064 mm/full motor
step.
As a first step, the static pressure along the test section centreline was measured to
check if there were any large pressure gradients in the test section (e.g., caused by the
jet collector). The pitot-static tube was placed on a linear streamwise traverse with the
42
Chapter 4. Facility Characterization
probe tip at the jet centreline pointing upstream. One end of the differential pressure
transducer was connected to the static-pressure port of the traversing probe, while the
other end was connected to a static pressure port at the test section inlet. The probe
was traversed the 2200 mm length of the open jet from inlet to collector, in 100 mm
increments. Pressure data were acquired at a freestream velocity of 55 m/s for 60 s at a
sampling rate of 4000 Hz at each measurement location, and are presented in Fig. 4.2 in
terms of coefficient of pressure, given by:
Cp =
Ps − P∞
,
q∞
(4.1)
where Ps is the static pressure reading at the traversing probe, P∞ is the freestream static
pressure at the inlet, and q∞ is the freestream dynamic pressure at the inlet. As expected,
there is a slight adverse pressure gradient moving downstream from the inlet due to the
streamlines expanding towards the collector. The pressure coefficient within 2 test section
widths downstream of the inlet remains below an acceptable value of 1%.
0.05
0.04
0.03
Cp
0.02
0.01
0
-0.01
-0.02
0
0.5
1
1.5
2
2.5
3
3.5
4
Streamwise location x/D
Figure 4.2: Variation in static pressure coefficient along the open-jet centreline from inlet
to collector.
The freestream velocity surveys were taken in the mid-height plane of the test section in
the form of 6 cross-stream profiles, every 300 mm from 0 mm to 1500 mm downstream of
the inlet. Each of the 6 profiles consisted of measurement points along a 800 mm line centred on the test section centreline, with 10 mm cross-stream spacing between each point
from 0 mm to 200 mm, 40 mm spacing from 200 mm to 600 mm, and then 10 mm spacing
from 600 mm to 800 mm. These profiles for freestream velocities of 75 m/s, 55 m/s, and
35 m/s are presented in Fig. 4.3. The profiles at each streamwise location are generally
consistent across all three freestream velocities, with the exception of the profiles at the
43
Chapter 4. Facility Characterization
Cross−stream location
normalized by test section width
0.8
0.6
0.4
0.2
0
−0.2
x = 0D
0.5D
1.0D
1.5D
2.0D
2.5D
−0.4
−0.6
−0.8
−0.2
0
0.2
0.4
0.6
0.8
1
1.2
0.8
1
1.2
0.8
1
1.2
Freestream velocity normalized by mean inlet velocity
(a) U∞ = 75 m/s.
Cross−stream location
normalized by test section width
0.8
0.6
0.4
0.2
0
−0.2
x = 0D
0.5D
1.0D
1.5D
2.0D
2.5D
−0.4
−0.6
−0.8
−0.2
0
0.2
0.4
0.6
Freestream velocity normalized by mean inlet velocity
(b) U∞ = 55 m/s.
Cross−stream location
normalized by test section width
0.8
0.6
0.4
0.2
0
−0.2
x = 0D
0.5D
1.0D
1.5D
2.0D
2.5D
−0.4
−0.6
−0.8
−0.2
0
0.2
0.4
0.6
Freestream velocity normalized by mean inlet velocity
(c) U∞ = 35 m/s.
Figure 4.3: Freestream velocity profiles at 75, 55, and 35 m/s measured at a range of
streamwise locations between x = 0 - 2.5 test section widths downstream of the inlet.
44
Chapter 4. Facility Characterization
inlet plane. At U∞ = 35 m/s, there is a 3% undershoot in freestream velocity near the
centre of the jet, within 0.3 test section widths of the centreline. At U∞ = 75 m/s, there
is a 5% overshoot near the edges of the jet, around 0.4-0.5 test section widths from the
centreline. These deviations from the mean freestream velocity are not present in the
profiles at the other streamwise locations and hence should not pose a problem for future
experiments.
A test section flow uniformity of at least 99% (i.e., non-uniformity = 1−uniformity ≤ 1%)
was chosen as an acceptable threshold. For each of the three measured test section
velocities, the useable test section area was estimated by finding the maximum test section
width that satisfied this uniformity threshold at each of the six streamwise measurement
locations. The variation of this usable test section width with streamwise location is
presented in Fig. 4.4; there is no substantial difference for the three measured velocities.
Future experiments will mainly be conducted in between 0.5 - 2.0 test section widths
downstream of the inlet and will have a minimum usable cross-sectional area of 500 mm
× 500 mm.
580
Usable test section width [mm]
570
560
550
540
530
520
510
500
490
480
0
0.5
1
1.5
2
2.5
x/D, Streamwise location normalized by test section inlet diameter
Figure 4.4: Minimum test section width that satisfies the 99% uniformity threshold, as a
function of streamwise location, for test section velocities of 75, 55, and 35 m/s.
Chapter 4. Facility Characterization
4.3
45
Freestream Turbulence Intensity
The freestream turbulence intensity along the centreline of the test section was measured
via hot-wire anemometry. A 5 µm diameter tungsten wire was soldered onto a 2 mm wide
single-prong hot-wire probe and coated with copper to yield a 0.5 mm sensing length on
the wire. The probe was connected to a Dantec 56C17 Constant Temperature Anemometer and calibrated against a pitot-static tube in the test section from 0 m/s to 75 m/s. For
the turbulence intensity measurements, the hot-wire probe was positioned in the centre of
the test section via an aluminum sting that was fastened to a linear streamwise traverse.
The hot-wire voltage signal was passed through a 10.5 kHz low-pass filter and acquired
by a NI PCIe-6361 DAQ for 20 s at a sampling rate of 21 kHz.
The power spectra of the streamwise velocity fluctuations at various streamwise locations
for a freestream velocity of 55 m/s are given in Fig. 4.5. The spectral frequency is
expressed in terms of Strouhal number based on the test section width, Dts :
St =
f Dts
.
U∞
(4.2)
The velocity fluctuations observed within one test section width downstream of the inlet
are mainly due to vorticity passed into the test section from the upstream tunnel components (i.e., dome environment, screens and contraction). Moving farther downstream
from the test section inlet, the velocity fluctuations increase due to shear layer-induced
pressure fields on the free-jet, as indicated by the broad hump in the spectra around St =
0.5 for x/D = 1.0 to 2.5 [87]. The narrow peaks on the humps are generated by resonant
pressure fluctuations between the test section inlet and the jet collector.
The centreline freestream turbulence intensity value at each of the streamwise measurement locations was calculated via integration of the velocity fluctuation power spectral
densities over a frequency range from a certain lower cut-off value up to where the noise
floor begins. The lower cut-off frequency, fc,low , was found based on the wavelength, λ, of
the largest longitudinal turbulent flow structures that could exist in the test section:
U∞
,
λ
(4.3)
λ = 2 × Lts ,
(4.4)
fc,low =
where
and Lts is the length of the test section. At U∞ = 55 m/s and Lts = 2.2 m, this translated
46
Chapter 4. Facility Characterization
Power Spectral Density of
Velocity Fluctuations [m 2 s −2Hz −1]
10
10
10
10
10
10
−2
x/D = 0.5
x/D = 0.75
x/D = 1.0
x/D = 1.5
x/D = 2.0
x/D = 2.5
−3
−4
−5
−6
−7
10
−1
10
0
10
1
St
Figure 4.5: Power spectral density of streamwise velocity fluctuations along test section
centreline at freestream velocity of 55 m/s, for streamwise locations between 0.5 and 2.5
test section widths downstream of inlet. Frequency is given in terms of Strouhal number
based on the test section width. All spectra have reached their noise floors by St = 10.
to a lower cut-off frequency of 12.5 Hz. All the velocity fluctuation spectra had reached
their noise floors by 1 kHz, so this value was chosen as the upper cut-off frequency.
Figure 4.6 shows the centreline freestream turbulence intensity values for both the old
and new AWT test sections. The comparison shows that the new tunnel flow conditioning
screens, as well as the increased contraction ratio, helped improve the turbulence intensity
values in the test section.
−2
TI
10
Old AWT
Improved AWT
10
−3
0
0.5
1
1.5
2
2.5
x/D, streamwise location normalized by test section inlet width
Figure 4.6: Comparison of centreline freestream turbulence intensity values for old and
new AWT test sections at freestream velocity of 55 m/s over a range of streamwise locations from 0.5 to 2.5 test section widths downstream of the inlet.
Chapter 4. Facility Characterization
4.4
47
Background Noise Levels
The background noise levels in the anechoic chamber were measured using a 0.5” diameter
G.R.A.S. 40AC 200 V externally polarized free-field microphone connected to a G.R.A.S.
26AS preamplifier via a 0.5 in -to- 0.25 in adapter. The microphone and preamplifier were
powered by a G.R.A.S. 12AA power module that also provided 20 Hz high-pass filtering
and optional 20 dB and 40 dB amplification to the microphone signals. The voltage signal
from the 12AA module was acquired by the NI PCIe-6361 DAQ, at a sampling rate of
42 kHz for 60 s for each data point. This measurement system had a SPL dynamic range
of -5 dB to 134 dB when using the 40 dB amplification setting. The pressure response
of the microphone and preamplifier is designed to be flat over the frequency range from
3.15 Hz to 40 kHz, with an SPL uncertainty of ±2 dB. Using a G.R.A.S. 42AB sound
calibrator that emits a 1 kHz tone at 114 dB, the sensitivity of the microphone and preamplifier pairing was determined to be 13.55 mV/Pa.
Background noise measurements were taken over the range of tunnel freestream velocities
from 0 to 75 m/s. The microphone was placed at two different locations in the acoustic
far-field region of the test section jet (see Fig. 4.7). Both positions were at the height of
the test section centreline, with the microphone diaphragm pointed towards an imaginary
point on the jet centreline 0.6 m downstream of the test section inlet. The diaphragm
was oriented 90◦ to the centreline in the first position, and 60◦ to the centreline facing
upstream in the second position.
Figure 4.7: Plan view of free-field microphone locations for tunnel background noise
measurements. Figure not to scale.
Chapter 4. Facility Characterization
48
The narrowband SPL spectra measured at these two microphone locations for a range
of test section velocities from 15 - 75 m/s are shown in Fig. 4.8. The noise spectra for
velocities below 15 m/s were omitted because the microphone measurement system had
reached its noise floor. This is not a significant issue as the large majority of future
aeroacoustics investigations in this tunnel will be conducted at velocities above 15 m/s.
As expected from an open-jet tunnel, the noise spectra are representative of broadband
noise. Through careful sealing of the anechoic chamber and wind tunnel components, all
noticeable stray whistling noises and cavity resonance tones that would have caused peaks
in the spectra have been removed.
Compared to the background SPL spectra of the old UTIAS tunnel, the new spectra
present reductions in background noise levels in the test chamber across all frequencies of
interest. For example, at a freestream velocity of 60 m/s, the reduction is approximately
10 dB at the lowest frequencies between 100-400 Hz, and as much as 15 dB above 500 Hz.
This substantial reduction means that microphones in the improved chamber can achieve
at least 10 dB signal-to-noise ratio for many airframe noise studies, including landing
gear and airfoil trailing edge noise. If an array of microphones and beamforming signal
processing algorithms are used, the signal-to-noise ratio may be further increased by at
least 5-6 dB, maybe up to 30 dB depending on the algorithm [88–90].
The background noise in the anechoic chamber mainly comes from the fan and motor,
open-jet shear layer impingement on the collector, freestream turbulence, and noise transmitted into the open-circuit tunnel from the surrounding up- and downstream environments. The motor contributes high frequency tonal electrical noise (some noticeable peaks
in the spectra at freestream velocities of 15-30 m/s) and low frequency mechanical vibrations, the latter of which could be reduced through motor and fan blade balancing, as
well as bearing greasing. Shear layer impingement on the collector would produce low frequency broadband noise, and previous researchers (e.g. [59,75,91]) have had some success
reducing this noise by expanding the collector or padding it with acoustically absorbent
material like foam or fibreglass. Turbulence in the shear layer emits broadband noise.
Environmental sources are unpredictable and could contribute both broadband and tonal
noise to the spectra, depending on the source. The main method of attenuating this type
of noise is through silencer/baffle noise blockage.
49
Chapter 4. Facility Characterization
80
15 m/s
20 m/s
30 m/s
40 m/s
50 m/s
60 m/s
70 m/s
75 m/s
Background SPL [dB ref 2x10 Pa]
70
−5
60
50
40
30
20
10
0
−10 2
10
3
10
Frequency [Hz]
10
4
(a) Microphone diaphragm oriented 90◦ to tunnel centreline.
80
15 m/s
20 m/s
30 m/s
40 m/s
50 m/s
60 m/s
70 m/s
75 m/s
Background SPL [dB ref 2x10 −5 Pa]
70
60
50
40
30
20
10
0
−10 2
10
3
10
Frequency [Hz]
10
4
(b) Microphone diaphragm oriented 60◦ to tunnel centreline, facing upstream.
Figure 4.8: Narrowband background SPL spectra in anechoic chamber, measured with
free-field microphone positioned at height of test section centreline, with diaphragm oriented at two different angles towards a point on the centreline that is 0.6 m downstream
of inlet.
50
Chapter 4. Facility Characterization
As mentioned in Section 4.1, the angle of attack of the tunnel fan blades only had a slight
effect on the motor rotational rate and hence the motor noise. The chosen blade angle
setting number ‘5’ was the quietest, with the other blade angle settings about 1-2 dB
louder (within measurement error of microphone). Figure 4.9 shows the overall sound
pressure level in the anechoic chamber measured at the two microphone locations, over a
range of test section velocities from 15-75 m/s with the fan blade angle at setting number
‘5’. The curves for the old UTIAS tunnel and several of the quietest tunnels around the
world are also provided for comparison. There is a clear improvement in the background
noise levels in the UTIAS tunnel, with the microphone measurements at 90◦ to the centreline achieving OASPL reductions between 8-18 dB over the whole range of operating
freestream velocities. As well, the noise levels do not change substantially between the
two microphone measurement angles, which is important for future far-field directivity
measurements.
Background A−weighted OASPL [dB]
80
70
Audi (52 deg)
DaimlerChrysler (32 deg)
Ford DTF (49 deg)
Florida State (90 deg)
ISVR Southampton (90 deg)
Univ. Cottbus (45 deg)
Univ. Florida (65 deg)
Univ. Siegen (45 deg)
Univ. Notre Dame (60 deg)
Old UTIAS (90 deg)
Improved UTIAS (90deg)
Improved UTIAS (60deg)
60
50
40
30
20
0
10
20
30
40
50
60
Test section velocity [m/s]
70
80
90
100
Figure 4.9: Comparison of overall background sound pressure levels between new and
old UTIAS AWT and select anechoic wind tunnels around the world, over their range of
capable test section velocities. Microphone orientation angle relative to tunnel centreline
is shown beside the facility name in the legend. Comparison data adapted from [59, 60,
72–74].
51
Chapter 4. Facility Characterization
The OASPL versus freestream velocity data points can be fitted with a logarithmic curve
to give some indication of the types of contributing noise sources in the tunnel. Past
empirical studies have shown that fan noise OASPL in dB versus fan rotational speed can
be described by the following relation [92, 93]:
∆ OASP L = 10 × log10
R2
R1
n
,
(4.5)
where R2 and R1 are motor rotational speeds, and n is some empirically-determined number. For dipole-type noise sources related to pressure fluctuations on axial fan blades, n is
typically between 5 and 6. For quadrupole noise sources related to freestream turbulence,
n is approximately 7-8. The motor rotational rate of the UTIAS tunnel is approximately
directly proportional to the freestream velocity. The A-weighted OASPL measured by the
microphone at the 90◦ orientation is plotted versus motor rotations per minute (RPM) in
Fig. 4.10 along with a logarithmic curve fitted through the data points. The curve of best
fit has an exponent n = 6.17, indicating that most of the background noise comes from
dipole-type sources such as the fan and shear layer interaction with the jet collector, as
well as some noise from quadrupole sources such as freestream turbulence.
Figure 4.10: A-weighted OASPL measured by far-field microphone in the 90◦ orientation
versus fan motor rotational speed. Logarithmic curve of best fit is ∆OASP L = 10 ×
log10 (R2 /R1 )6.17 .
Prior to the addition of front-end baffles, noise disturbances from inside the reverberant
geodesic dome entered the test section through the tunnel bellmouth inlet. This noise
was from activity in the dome and thermal creaking of the dome walls, and it could not
be filtered out of acoustic measurements due to its randomness. On the back-end of the
tunnel, there are already baffles built into the diffuser section that help reduce the level
52
Chapter 4. Facility Characterization
of fan and motor noise and outdoor disturbances propagating back upstream into the
anechoic chamber. After the addition of the front-end baffles, an acoustic characterization was performed to estimate their effectiveness in attenuating sound waves travelling
towards the tunnel inlet. This characterization also determined the sound attenuation by
the wind tunnel front-end components (i.e., honeycomb, mesh screens, and contractions).
The process involved generating white noise using an Altec Lansing ACS22 computer
speaker inside the geodesic dome and measuring the sound transmission loss through the
left, centre, and right side of the baffles. For each of the three test directions, microphones measured the sound pressure level in front of the baffles, behind the baffles at the
centre of the bellmouth inlet, and inside the anechoic chamber at the 90◦ far-field noise
measurement location described earlier. The distances between each of these microphone
locations and the speaker are summarized in Table 4.1.
Table 4.1: Summary of sound transmission distances between the speaker and each of the
three microphones at the baffle, the bellmouth inlet, and in the anechoic chamber. Transmission directions are defined facing the baffles when viewing from inside the geodesic
dome (same directions as in Fig. 3.18).
Transmission
Direction
Baffle mic
[m]
Bellmouth mic
[m]
Chamber mic
[m]
Leftside of baffles
2.9
7.6
13.0
Centre / head-on
2.9
5.2
10.6
Rightside of baffles
2.9
7.6
13.0
The noise measurements were corrected for spherical transmission spreading loss via the
equation:
SPLloss = 20 × log10
r2
r1
,
(4.6)
where r2 and r1 are the distances between the speaker and the two microphone locations
under comparison. The atmospheric attenuation of sound at 20 ◦ C and 60% relative
humidity across the frequency range of 100 Hz to 10 kHz is between 0.0002 dB/m and
0.1 dB/m, which, given the short transmission distances at hand, was considered negligible. The A-weighted OASPL attenuation by the baffles and wind tunnel front-end
components are summarized in Table 4.2.
53
Chapter 4. Facility Characterization
Table 4.2: Summary of white noise attenuation by the front-end baffles and tunnel frontend components, in terms of A-weighted OASPL. Transmission directions are defined
facing the baffles when viewing from inside the geodesic dome (same directions as in
Fig. 3.18).
Transmission Direction
Baffles [dB]
Tunnel Front-end [dB] Total [dB]
Leftside of baffles
-11
-14
-25
Centre / head-on
-16
-14
-30
Rightside of baffles
-8
-13
-21
There are considerable gains in front-end noise level attenuation by the addition of the
baffles in all three directions. The centre of the baffles offer the strongest attenuation
due to the wood sheeting on the front of the baffles and minimal airflow openings. The
righthand side of the baffles have the weakest attenuation because of the relatively large
opening for air to flow into the bellmouth inlet. While there is no direct line-of-sight
into the tunnel inlet from any direction outside of the baffles, it is conjectured that lower
frequency sound waves are able to bend around the baffles. Figure 4.11 compares the
third-octave A-weighted OASPL spectra measured in front and behind the righthand side
of the baffles. It is clear that at frequencies below around 100 Hz, the baffle attenuation
is weaker than at frequencies above 100 Hz.
80
A−weighted OASPL [dB]
75
70
65
60
55
50
45
40
Behind baffles at tunnel inlet
In front of baffles pointed at speakers
35
30 1
10
2
10
3
10
4
10
5
10
Third−octave frequency bands [Hz]
Figure 4.11: Third-octave A-weighted SPL of white noise generated by the speaker in
front of the baffles, as measured by microphones in front of and behind the baffles on the
right hand side of the wind tunnel inlet.
54
Chapter 4. Facility Characterization
4.5
Anechoic Cut-off Frequency
As discussed earlier, the theoretical cut-off frequency based on the depth of the wedges
is approximately 150 Hz, but construction details and wind tunnel components protruding into the chamber can increase this value. In order to determine the actual cut-off
frequency, a series of microphone traverses were performed using the tonal noise discrete
sampling method found in ISO-3745:2012 for ten tonal frequencies from 1 kHz down to
125 Hz.
A speaker was set up on the tunnel centreline, 0.6 m downstream of the test section inlet,
and was driven by an Agilent 33210A function generator. Three free-field microphone
traverses were performed on paths extending from the speaker to the chamber walls at
45◦ , 90◦ , and 135◦ relative to the downstream tunnel centreline (see schematic of set-up
in Fig. 4.12), with the microphone diaphragm facing the speaker at all times. The microphone hung under a traverse line made of 0.3 mm diameter fishing line and was pulled
by a take-up line attached to a spool on a rotary stage driven by a stepper motor. The
whole microphone traversing system had a linear resolution of 0.5 mm per full step of the
motor. In order to resolve the spatial characteristics of the tonal noise, the separation
distance between measurement locations along a traverse must not exceed 15% of the
acoustic wavelength of interest [94]. This translated to 0.05 m for the shortest wavelength
that was tested (i.e., at frequency of 1 kHz). Traverses for all the tonal frequencies were
performed from 0.3 m to 2.2 m away from the speaker diaphragm, at 0.05 m increments.
A stationary reference microphone was placed on the tunnel centreline 0.3 m from the
speaker to monitor the stability of the sound source during the traverses.
Based on the assumption that the anechoic chamber approximates an acoustic free-field,
the SPL readings at distances r1 and r2 from the sound source should adhere to the
following relation:
SPL(r1 ) = SPL(r2 ) − 20 × log10
r1
r2
.
(4.7)
This relation is valid for measurement locations that are in the acoustic far-field region
of the sound source, which is given by [94]:
95 m/Hz
,
r ≥ MAX 0.5 m,
f
(4.8)
where f is the noise frequency of interest in Hz. The acceptable bounds for any deviation
of the traverse SPL readings from the free-field assumption are defined in ISO-3745:2012
55
Chapter 4. Facility Characterization
Figure 4.12: Schematic of speaker and reference microphone locations and three microphone traverse lines for the free-field characterization test. All traverse lines are in the
mid-plane of the anechoic chamber coinciding with the tunnel centreline.
and summarized in Table 4.3 for the frequencies of interest. The SPL measurements for
the three traverse directions, expressed as deviation from the free-field assumption, are
given in Figs. 4.13-4.15 for the measurement locations that are in the far-field of the
speaker.
Table 4.3: Acceptable SPL error bounds for free-field characterization of anechoic chambers. Values defined in ISO-3745:2012.
Frequency range
SPL error bound
≤ 630 Hz
± 1.5 dB
800 Hz - 5 kHz
± 1.0 dB
56
Chapter 4. Facility Characterization
SPL deviation from acoustic freefield assumption [dB]
2
0
-2
2
0
-2
5
0
-5
2
0
-2
2
0
-2
2
0
-2
2
0
-2
2
0
-2
2
0
-2
1
0
-1
f = 125
0
0.5
1
f = 150 1.5
2
2.5
0
0.5
1
f = 160 1.5
2
2.5
0
0.5
1
f = 170 1.5
2
2.5
0
0.5
1
f = 180 1.5
2
2.5
0
0.5
1
f = 190 1.5
2
2.5
0
0.5
1
f = 200 1.5
2
2.5
0
0.5
1
f = 250 1.5
2
2.5
0
0.5
1
f = 500 1.5
2
2.5
0
0.5
1 f = 1000 1.5
2
2.5
0
0.5
1
2
2.5
1.5
Distance from speaker [m]
Figure 4.13: Deviation of SPL measurements from free-field assumption along the traverse
path at 45◦ relative to the downstream centreline (towards the door), for various speaker
frequencies from 125 Hz to 1 kHz. Acceptable SPL error bounds from ISO-3745:2012.
57
Chapter 4. Facility Characterization
SPL deviation from acoustic freefield assumption [dB]
2
0
-2
2
0
-2
2
0
-2
2
0
-2
2
0
-2
2
0
-2
2
0
-2
2
0
-2
2
0
-2
1
0
-1
f = 125
0
0.5
1
f = 150 1.5
2
2.5
0
0.5
1
f = 160 1.5
2
2.5
0
0.5
1
f = 170 1.5
2
2.5
0
0.5
1
f = 180 1.5
2
2.5
0
0.5
1
f = 190 1.5
2
2.5
0
0.5
1
f = 200 1.5
2
2.5
0
0.5
1
f = 250 1.5
2
2.5
0
0.5
1
f = 500 1.5
2
2.5
0
0.5
1 f = 1000 1.5
2
2.5
0
0.5
1
2
2.5
1.5
Distance from speaker [m]
Figure 4.14: Deviation of SPL measurements from free-field assumption along the traverse
path at 90◦ relative to the downstream centreline, for various speaker frequencies from
125 Hz to 1 kHz. Acceptable SPL error bounds from ISO-3745:2012.
58
Chapter 4. Facility Characterization
SPL deviation from acoustic freefield assumption [dB]
2
0
-2
2
0
-2
2
0
-2
2
0
-2
2
0
-2
2
0
-2
2
0
-2
2
0
-2
5
0
-5
2
0
-2
f = 125
0
0.5
1
f = 150 1.5
2
2.5
0
0.5
1
f = 160 1.5
2
2.5
0
0.5
1
f = 170 1.5
2
2.5
0
0.5
1
f = 180 1.5
2
2.5
0
0.5
1
f = 190 1.5
2
2.5
0
0.5
1
f = 200 1.5
2
2.5
0
0.5
1
f = 250 1.5
2
2.5
0
0.5
1
f = 500 1.5
2
2.5
0
0.5
1 f = 1000 1.5
2
2.5
0
0.5
1
2
2.5
1.5
Distance from speaker [m]
Figure 4.15: Deviation of SPL measurements from free-field assumption along the traverse
path at 135◦ relative to the downstream centreline, for various speaker frequencies from
125 Hz to 1 kHz. Acceptable SPL error bounds from ISO-3745:2012.
Chapter 4. Facility Characterization
59
For the 45◦ traverse, the anechoic chamber and false door provide a good acoustic freefield environment at frequencies down to 170 Hz and within 2.2 m of the speaker. At
160 Hz and 150 Hz, and measurement locations greater than 1.8 m from the speaker,
the SPL values exceed the acceptable error bounds and therefore the chamber cannot
be considered anechoic in those conditions. For the 90◦ traverse, the SPL values for all
measured frequencies and at measurement points up to 2 m from the speaker are within
the allowable error bounds. Beyond 2 m from the speaker, the wedges on the closest wall
begin to have an adverse effect on the free-field quality, pushing the SPL deviations at
frequencies of 190 Hz and 200 Hz slightly above acceptable limits. Finally, for the 135◦
traverse, the chamber largely adheres to the free-field assumption at frequencies down to
170 Hz and within 2.2 m of the speaker, except for small (≤0.5 dB) deviations beyond
allowable limits at frequencies of 500 Hz and 1 kHz. This is most likely due to acoustic reflections from the protruding tunnel contraction and flow sections, which were only
covered in one layer of egg-crate foam sheets. A potential improvement to the acoustic
treatment on these tunnel sections could be miniature wedges with a depth of 0.15-0.2 m,
but this is non-critical and the additional wedges would be expensive.
In summary, the anechoic cut-off frequency of the chamber is around 170 Hz, and the
free-field assumption is valid for a minimum distance of 2 m from a point on the tunnel
centreline that is 0.6 m downstream of the test section inlet, thus meeting the relevant
thesis objectives.
4.6
Reverberation Time
The integrated impulse response method found in ISO-3382-1:2009 was used to determine
the reverberation time of the anechoic chamber. The standard normally requires an impulse noise source that has a peak SPL of at least 75 dB above the noise floor in order
to directly measure RT60 . However, this was not possible with the readily available lab
equipment, so instead an approximation method within the standard was used. RT20 , the
time for an impulse response to decrease by 20 dB, was calculated from an impulse noise
with a peak SPL at least 35 dB above the noise floor. RT20 was then linearly extrapolated
to get RT60 .
A microphone was set-up at the 90◦ far-field measurement location as previously described, and a loud clap in the middle of the anechoic chamber served as the noise source.
The pressure response signal from this impulse noise is shown in Fig. 4.16.
60
Chapter 4. Facility Characterization
8
6
Pressure signal [Pa]
4
2
0
−2
−4
−6
−8
0
200
400
600
800
1000
1200
1400
1600
1800
2000
Time [s]
Figure 4.16: Pressure response signal of a single clap impulse noise in the centre of the
anechoic chamber, as measured by a microphone at the 90◦ far-field measurement location.
Using cumulative trapezoidal integration on the impulse response pressure signal, noise
decay curves were estimated for octave bands with centre frequencies between 250 Hz
and 8 kHz . The reverberation time for each octave band was then approximated using a
least-squares fit on the corresponding decay curve between the points -5 dB and -25 dB
below the peak of the impulse noise to get RT20 , and then linearly extrapolated to get
RT60 . Figure 4.17 compares the reverberation times for the improved anechoic chamber
and the bare concrete chamber. The presence of acoustic wedges greatly reduces the reverberation time from an average of 1.5 s for the bare concrete chamber to 0.032 s for the
acoustically-treated chamber. In general, the reverberation time of the anechoic chamber
increases as frequency decreases towards the anechoic cut-off frequency. Interestingly,
there is also a slight increase in the RT60 of the treated anechoic chamber around 4 kHz,
which corresponds to the dip in the acoustic absorption coefficient of the foam used for
the wedges (see Fig. 3.10).
As mentioned in Section 2, the literature does not usually discuss reverberation time for
acoustic wind tunnels so comparisons cannot be made. However, the values here all satisfy the thesis objective of RT60 ≤ 0.2 s and meet the expectations of future aeroacoustic
experiments in this facility.
61
Chapter 4. Facility Characterization
10
New anechoic chamber
Bare concrete chamber
0
RT
60
[s]
101
10-1
10-2 2
10
103
104
Frequency [Hz]
Figure 4.17: Comparison between the RT60 of the improved anechoic chamber and of the
bare concrete chamber, for octaves with centre frequencies between 250 Hz and 8 kHz.
Chapter 5
Conclusions
5.1
Summary
This thesis discussed the motivation, design, construction, and characterization of the
improvements to the UTIAS Anechoic Wind Tunnel facility. Reduction of aircraft noise
forms a key part of the sustainable aviation agenda currently being pushed by international governing bodies. Facilities like the UTIAS AWT are rare and play a critical role
in the aeroacoustic research efforts led by both industry and academia. The work performed in this thesis helped to update the UTIAS AWT for current and foreseeable future
research needs.
At the beginning of this thesis, a set of aerodynamic and acoustic performance objectives
were laid out. A full characterization of the facility has shown that these objectives were
either met or exceeded, with the exception of the background noise levels at frequencies
below 500 Hz. Table 5.1 summarizes the results.
Table 5.1: Summary of tunnel performance objectives and results from the facility characterization.
Objective
Target
Result
Test section maximum
freestream velocity
≥70 m/s
75 m/s
Test section
dimensions
Cross-sectional diameter ≥0.5 m,
streamwise length ≥1.2 m
0.5 m × 0.5 m,
1.2 m
62
63
Chapter 5. Conclusions
Test section freestream
turbulence intensity
≤ 0.5% for x/D ≤ 1.5
≤ 0.4%
Test section spatial
flow uniformity
≥ 99% across test section
≥ 99%
Background noise level
-15 dB SPL, 200 Hz - 20 kHz
-10 dB (200-500 Hz),
-15 dB (500-20,000 Hz)
Anechoic limit
cut-off frequency
≤ 200 Hz
170 Hz
Usable anechoic area
around test section
2 m arc about a point on centreline & 0.6 m downstream of inlet
Satisfied, 45◦ -135◦
relative to centreline
Reverberation time
Below 0.2 s
0.032 s
5.2
Recommendations for Future Work
There are several immediate recommendations for future aeroacoustics tests in the UTIAS
AWT.
First, if the interest is in the low frequency noise range (200-500 Hz) and the current background levels are too high, then a few things could be investigated. The signal-to-noise
ratio of far-field microphone measurements at these frequencies may be improved by using windscreens to reduce microphone self-noise generated by the out-of-jet recirculating
flows. The jet-collector may require improvement to reduce the noise generated by shear
layer impingement, possibly by adding more/better padding material or by increasing the
flow capture area of the collector. If the latter is chosen, care must be taken not to increase
the area too much and cause the static pressure in the test section to exceed tolerances.
Another source of low frequency noise that may require attention is the vibrational noise
from the motor and fan. If balancing and greasing does not sufficiently reduce vibrations,
a thin circumferential section may need to be cut out of the diffuser and replaced with a
rubber membrane to isolate the vibrations.
Chapter 5. Conclusions
64
Second, if the test model is long (i.e., extends past 2 test section widths downstream of
the inlet), and far-field noise measurements are required at azimuthal angles less than 60◦
to the centreline pointing upstream, there may not be enough length in the chamber to
set up these microphones (the false door and/or collector wall would get in the way). In
this case, the wind tunnel straight flow section that currently sits between the primary
and secondary contractions may be omitted to shift the test section (and hence model
and far-field microphone arc) upstream by 0.6 m. A characterization would have to be
performed to determine how this extra 0.6 m length of open-jet affects the freestream flow
profiles.
Finally, for all future aeroacoustics experiments in the AWT, there should be a way of
measuring the local atmospheric pressure and humidity inside the chamber to give an
accurate calculation of certain aerodynamic and acoustic parameters (e.g., air density,
sound absorptivity in air, etc.).
Appendix A
Anechoic Foam Material Data Sheet
65
Appendix A. Anechoic Foam Material Data Sheet
66
Figure A.1: Material data sheet for polyurethane foam used for the anechoic wedges in
the improved UTIAS AWT.
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