chekiri rafik 201406 mas thesis

Experimental Aeroacoustic Study of aLanding Gear in the Unsteady Flow

Induced by a Propeller

by

Rafik Chekiri

A thesis submitted in conformity with the requirementsfor the degree of Master of Applied ScienceGraduate Department of Aerospace Studies

University of Toronto

© April 2014 by Rafik Chekiri

Abstract

Experimental Aeroacoustic Study of a Landing Gearin the Unsteady Flow Induced by a Propeller

Rafik ChekiriMaster of Applied Science

Graduate Department of Aerospace StudiesUniversity of Toronto

An aeroacoustic study of a two-strut, two-wheel, nacelle-mounted landing gear was

conducted to investigate the effects of an upstream propeller on the radiated noise. The

development of a 1:10.8 scale model based on a Bombardier Q400 aircraft, consisting of a

propeller, motor, nacelle, and landing gear assembly is discussed. Comparisons are made

between cases with and without an actuated upstream propeller. Far-field microphone

measurements out of the airstream are presented to characterize the acoustic effects of

each model component. The main strut and wheels of the model were equipped with

surface-mounted microphones to measure unsteady pressures. It is shown that the noise

signature of the landing gear cannot be observed over the tunnel background noise in the

far-field. Unsteady surface pressures on the main strut show dominant peaks related to

vortex shedding from the drag strut for both steady and unsteady upstream conditions.

ii

Acknowledgements

This work was funded through the Green Aviation Research and Development Network(GARDN), the Natural Sciences and Engineering Research Council of Canada (NSERC),Aercoustics Engineering Ltd., and Bombardier Aerospace. Thank you to Dr. WernerRicharz of Aercoustics Engineering Ltd., as well as Stephen Collavincenzo and Dr. Ray-mond Wong of Bombardier Aerospace for their technical suggestions and guidance.

I would like to thank my good friend, Cameron Robertson, for bringing this projectto my attention and putting me in contact with my supervisor, Dr. Philippe Lavoie.Dr. Lavoie has provided the best kind of mentorship and understanding throughout theinnumerable tribulations of this work, for which I am tremendously grateful. I must alsothank Dr. Alis Ekmekci for her prompt refereeing of my thesis and useful feedback.

Joining the Flow Control and Experimental Turbulence group at the Institute forAerospace Studies has been an unforgettable experience. The attitude, knowledge, andgenerosity of the team made it a wonderful place to work and share ideas. I wouldparticularly like to thank Nicole Houser, Jason Hearst, Heather Clark, and Derrick Chowfor their insight in solving technical problems and assistance in editing this work. Thefriendships from my three years at UTIAS will always be treasured. Thank you to NicoleHouser for working late nights and early mornings with me, keeping me focussed, teachingme about life, and being an amazing partner.

Lastly, I must thank my parents for continually insisting on the importance of knowl-edge and perseverance. Your unceasing support and encouragement, even from a dis-tance, has kept me happy, productive, and sane. I love you both very much.

iii

Contents

1 Introduction 11.1 Aircraft Noise . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11.2 Objective . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21.3 Thesis Outline . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3

2 Background 42.1 Aerodynamic Sources of Sound . . . . . . . . . . . . . . . . . . . . . . . 42.2 Bluff-body Flows . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5

2.2.1 Isolated Cylinders . . . . . . . . . . . . . . . . . . . . . . . . . . . 52.2.2 Tandem Cylinders . . . . . . . . . . . . . . . . . . . . . . . . . . 8

2.3 Landing Gear Studies . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 112.3.1 Simplified Geometries . . . . . . . . . . . . . . . . . . . . . . . . 122.3.2 High Fidelity Geometries . . . . . . . . . . . . . . . . . . . . . . . 12

2.4 Propeller Flows . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 132.4.1 Flow Structures . . . . . . . . . . . . . . . . . . . . . . . . . . . . 132.4.2 Acoustics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16

3 Wind Tunnel Characterization 173.1 Facility Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 173.2 Aerodynamic Survey . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19

3.2.1 Experimental Method . . . . . . . . . . . . . . . . . . . . . . . . 193.2.2 Mean Flow Measurements and Turbulence Levels . . . . . . . . . 20

3.3 Background Acoustic Levels . . . . . . . . . . . . . . . . . . . . . . . . . 21

4 Experimental Method 254.1 Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 254.2 Model Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 254.3 Instrumentation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29

4.3.1 Far-field Microphones . . . . . . . . . . . . . . . . . . . . . . . . . 294.3.2 Surface Microphones . . . . . . . . . . . . . . . . . . . . . . . . . 294.3.3 Data Acquisition & Uncertainty . . . . . . . . . . . . . . . . . . . 31

5 Results & Analysis 375.1 Far-field Acoustics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 37

5.1.1 Static Configurations . . . . . . . . . . . . . . . . . . . . . . . . . 37

iv

5.1.2 Motor and Propeller Induced Noise . . . . . . . . . . . . . . . . . 395.2 Unsteady Surface Pressures . . . . . . . . . . . . . . . . . . . . . . . . . 39

5.2.1 Strut Surface Microphones . . . . . . . . . . . . . . . . . . . . . . 425.2.2 Wheel Surface Microphones . . . . . . . . . . . . . . . . . . . . . 44

6 Conclusions 526.1 Wind Tunnel Characterization . . . . . . . . . . . . . . . . . . . . . . . . 526.2 Model Development . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 526.3 Landing Gear Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 536.4 Recommendations for Future Work . . . . . . . . . . . . . . . . . . . . . 53

References 55

Appendices

A Wind Tunnel Layout 60

B Surface Microphone Calibration 62

C Landing Gear Model Drawings 64

v

List of Tables

2.1 Flow regimes of a disturbance-free flow around a circular cylinder. . . . . 62.2 Description of tandem cylinder spacing effects. . . . . . . . . . . . . . . . 9

4.1 Experimental test matrix for acoustic measurement tests. . . . . . . . . . 284.2 Spacing ratios at the unsteady surface pressure locations. . . . . . . . . . 30

vi

List of Figures

1.1 Bombardier Dash 8 Q400 aircraft. . . . . . . . . . . . . . . . . . . . . . . 2

2.1 Regions of disturbed flow surrounding a circular cylinder. . . . . . . . . . 52.2 Yawed cylinder schematic geometry and nomenclature. . . . . . . . . . . 72.3 Tandem cylinder configuration, definitions, and vortex shedding regimes. 82.4 Noise spectra for two 1 inch rods in a tandem configuration. . . . . . . . 92.5 Geometry for tandem cylinder configuration with yawed upstream cylinder. 112.6 Bombardier Dash-8 Q400 main landing gear assembly. . . . . . . . . . . . 122.7 Helical vortex system and slipstream tube generated by a propeller. . . . 142.8 Radial distribution of axial and tangential velocities behind a propeller. . 152.9 Far-field noise spectra of full-scale and 1/4-scale propellers. . . . . . . . . 16

3.1 Photograph of the Acoustic Wind Tunnel Facility. . . . . . . . . . . . . . 183.2 Anechoic wedge geometry. . . . . . . . . . . . . . . . . . . . . . . . . . . 193.3 Measurement grid for aerodynamic survey. . . . . . . . . . . . . . . . . . 203.4 Open jet velocity profiles of the Acoustic Wind Tunnel. . . . . . . . . . . 223.5 Power spectral density of wind tunnel jet longitudinal velocity fluctuations. 233.6 Turbulence intensity in the wind tunnel open jet vs. axial position. . . . 233.7 Turbulence intensity in the wind tunnel open jet vs. radial position. . . . 243.8 Background narrow-band SPL spectra of the acoustic wind tunnel. . . . . 24

4.1 Image of landing gear model in wind tunnel test section. . . . . . . . . . 254.2 Schematic drawing of the complete landing gear model geometry in the

wind tunnel test section. . . . . . . . . . . . . . . . . . . . . . . . . . . . 264.3 Propeller wake velocity profiles for 3-blade, 15× 7 propeller . . . . . . . 334.4 Propeller wake velocity profiles for 3-blade and 2-blade propellers. . . . . 334.5 Schematic drawing of far-field microphone test locations. . . . . . . . . . 344.6 Schematic drawing of model surface microphone locations. . . . . . . . . 344.7 Sample calibration curve for a surface microphone sensor. . . . . . . . . . 354.8 Instrumentation layout for data acquisition . . . . . . . . . . . . . . . . . 36

5.1 Far-field SPL spectra of static configurations. . . . . . . . . . . . . . . . 385.2 Far-field SPL spectra of motor configurations . . . . . . . . . . . . . . . . 405.3 Far-field SPL spectra of propeller configurations . . . . . . . . . . . . . . 415.4 Slipstream contraction for all test cases. . . . . . . . . . . . . . . . . . . 425.5 Strut unsteady surface pressure spectral levels, M∞ = 0.04 . . . . . . . . 45

vii

5.6 Strut unsteady surface pressure spectral levels, M∞ = 0.07 . . . . . . . . 465.7 Strut unsteady surface pressure spectral levels, M∞ = 0.10 . . . . . . . . 475.8 Wheel unsteady surface pressure spectral levels, M∞ = 0.04 . . . . . . . 495.9 Wheel unsteady surface pressure spectral levels, M∞ = 0.07 . . . . . . . 505.10 Wheel unsteady surface pressure spectral levels, M∞ = 0.10 . . . . . . . 51

A.1 Schematic of the Acoustic Wind Tunnel Facility. . . . . . . . . . . . . . . 61

B.1 Calibration curves for all surface microphone sensors. . . . . . . . . . . . 63

viii

1 | Introduction

1.1 Aircraft Noise

Communities and regulatory bodies have long been concerned with the noise produc-tion of the aviation industry. In the United States, the Federal Aircraft Administration(FAA) has taken the initiative to reduce aircraft noise through regulations and certifica-tions in accordance with the Federal Aircraft Regulations Part 36 [12]. The approach isthree pronged, with measures taken to control noise sources, set operational restrictions,and use effective land planning to reduce community noise exposure. Some of these ac-tions have also been adopted by the International Civil Aviation Organization (ICAO).Recent interest from the aviation industry, driven by these and other regulations fromgovernmental and international organizations, has motivated research into aircraft noise.

Since the 1970s, there has been a growing interest in aircraft noise studies. In both theUnited States and the United Kingdom, early studies were aimed at reducing the noise ofaircraft engines. However, with the advent of high bypass-ratio turbofan engines, focuswas broadened to include airframe noise. Early phases of research into understanding andpredicting airframe noise used ‘total aircraft methods.’ Aircraft parameters such as wingarea, aspect ratio, and flight velocity were used to approximate the overall noise signature(sound level and frequency spectrum) [21, 14]. This method of prediction was supersededby ‘source-component methods’ which have the advantage of dividing the analysis toindividual components that may separate the noise generation mechanisms. The airframeelements identified as sound sources include, but are not limited to, the wings, flaps, slats,stabilizers, wheel-wells, and landing gear [22]. Combining the noise signatures from eachsource, the overall airframe noise signature are determined. Although landing gear hadbeen identified as a primary contributor, early research regarding airframe noise wasdirected toward wing components (i.e. slats and flaps) primarily due to the complexgeometries of conventional gear.

Early research concerning landing gear noise wrongly identified it as an exclusively lowfrequency phenomenon [8]. This was corrected in the 1990s when wind tunnels capable of

1

Chapter 1. Introduction 2

testing full scale models became available. Dobrzynski and Buchholz [10] performed testsconcluding that early models lacked the detail required to account for higher frequencynoise apparent in actual landing gear. The tabulated noise levels peaked between 1and 20 kHz. In addition, these tests attributed low frequency noise to the large gearcomponents (wheels, struts, etc.) and high frequency noise to the smaller elements (pins,cables, nuts, etc.). Recent studies have investigated generic landing gear models such asthe four-wheel rudimentary landing gear model introduced by Spalart et al. [42, 34] andthe two-wheel LAGOON model used by Manoha et al. of Airbus [27].

Empirical approaches to the problems of predicting and modeling landing gear noiseremain dominant in industry. Past experimental studies, such as those by Guo [17] andDobrzynski et al. [9], have demonstrated that a landing gear noise signature depends onlocal flow conditions that can be influenced by the landing gear’s location on an aircraft(wing vs. fuselage) or upstream airframe components.

1.2 Objective

The current study aims to identify the aeroacoustic signature specific to a two-strut, two-wheel landing gear when the components are in the unsteady slipstream of a propeller.The motivation of this work stems from community noise concerns related to regionalturboprop aircraft such as the Bombardier Dash-8 Q400, shown in Figure 1.1. This isthe first documented aeroacoustic investigation of the effects of an upstream propelleron the radiated noise of landing gear. Both far-field acoustic and surface pressure datawill be acquired simultaneously to identify pressure fluctuations which are manifested asfar-field noise.

Figure 1.1: Austrian Arrows Bombardier Dash 8 Q400 aircraft. Taken from [5].

Chapter 1. Introduction 3

1.3 Thesis Outline

The structure of this document is described below. Chapter 2 reviews some conceptsrelevant to the experimental design including the expected flow patterns and noise emis-sions from both landing gear components and propellers. The work done to characterizethe experimental facility is presented in Chapter 3. This was necessary prior to theaeroacoustic experiment due to many recent modifications to the wind tunnel. Bothaerodynamic and acoustic testing were performed to quantify the wind tunnel charac-teristics. Chapter 4 addresses the experimental set-up for far-field acoustic and surfacepressure measurements. The results of the landing gear study for various configura-tions and flow speeds are reported and discussed in the following chapter. Lastly, someconclusions and recommendations for future investigations are presented in Chapter 6.

2 | Background

The proposed investigation requires a variety of topics to be reviewed. This sectionbegins with a brief discussion on the theoretical sources of aerodynamic sound. Canonicalflows of single and tandem cylinder arrangements are presented as a prelude to complexlanding gear geometries. The far-field acoustics of these cylinder flows is presented priorto a discussion of aeroacoustic studies of landing gear in uniform inflow. Testing ofrudimentary and high-fidelity models are examined with key results addressed. Lastly, areview of the flow structures of propellers and their expected noise signature is presented.

2.1 Aerodynamic Sources of Sound

Lighthill’s theory of aerodynamic noise generation [26] is widely considered the startingpoint for modern aeroacoustics research. His reformulation of the Navier-Stokes equationsresulted in an inhomogeneous form of the wave equation separating acoustic propagationprocesses from aerodynamic sources. Assuming a uniform medium at rest, the equa-tions of the propagation of sound due to an externally applied fluctuating stresses usingEinstein tensor notation is

∂2ρ

∂t2− a20∇2ρ =

∂2Tij∂xi∂xj

, (2.1)

where ρ is the fluid density, a0 is the speed of sound in the fluid, and Tij is the Lighthillstress tensor:

Tij = ρuiuj + Pij − a20ρδij, (2.2)

where δij is the Kronecker delta. Thus, it can be observed that the sources of acousticradiation can be attributed to the three terms in Lighthill stress tensor, correspondingto momentum flux (ρuiuj), the stress tensor of a Stokesian fluid (Pij), and nonlinearnoise generation mechanisms (a20ρδij). The latter two of these sources are negligible forlow Mach number flows [26, 15]. It can be shown that in this case, the sound produced

4

Chapter 2. Background 5

by Tij corresponds to a quadrupole field. Lighthill’s formulation was further developedby Curle [7] to take into account the presence of solid boundaries. A notable result forlow Mach number flows is that dipole radiation dominates the sound emanating fromquadrupole sources. Ffowcs Williams and Hawkings [13] again extended this theory toinclude objects moving in flows, showing that forward motion of a source influences theradiation pattern with respect to an observer.

2.2 Bluff-body Flows

Landing gear can be decomposed into an assortment of non-streamlined (bluff) bodies.An understanding of the acoustics of bluff-body flows is beneficial in determining noisesources of complete landing gear. Flows around all bluff bodies have some similaritiesincluding large regions of flow separation and unsteadiness. The separated flow region islargely determined by the locations of boundary layer transition and separation, which areheavily influenced by Reynolds number, body shape, viscosity, and free stream turbulencelevels. Unsteadiness is manifested as fluctuating pressures on the body surface and inthe resulting wake. If pressure fluctuations are periodic, they lead to an Aeolian tone atthe vortex shedding frequency.

2.2.1 Isolated Cylinders

U∞

(i)(ii)

(iii) U > U∞

D

(iii) U > U∞

(iv) U < U∞

Figure 2.1: Regions of disturbed flow sur-rounding a circular cylinder. Adapted fromZdravkovich [48].

Many components of landing gear can becharacterized as circular cylinders of dif-ferent diameter, aspect ratio, and align-ment. Thus investigating the flow arounda circular cylinder provides a good start-ing point for landing gear studies. Flowaround circular cylinders has been exten-sively researched in the past, a compre-hensive summary of which is provided byZdravkovich [48, 49]. Surrounding a cylin-der, four areas have been identified as ar-eas where the local velocity, U , and pressure, p, vary from the freestream conditions(U∞, p∞). Figure 2.1 is a schematic representation of these regions. Region (i) is a nar-row region of retarded flow, forming unsteady flow structures in the streamwise directiondirectly ahead of the cylinder. Region (ii) marks the viscous boundary layer prior to


separation from the cylinder surface. The separation of the shear layer prescribes thedownstream boundary for region (iii). Region (iii) is an area of displaced and acceleratedflow, the extent of which is affected by the blockage of the cylinder. Lastly, region (iv) isa wide downstream region of separated flow, identified as the wake.

Perhaps the most influential parameter in defining the state of disturbance-free cylin-der flows is the Reynolds number based on the cylinder diameter, D:

ReD =ρDU∞µ

, (2.3)

where µ is the dynamic viscosity of the freestream fluid. The Reynolds number representsa ratio of inertial to viscous forces. Each cylinder flow state has a subset of flow regimesin which flow patterns are distinct. These are identified in Table 2.1. Generally, asReD increases, the point at which transition to turbulence occurs moves upstream. Atlow Reynolds numbers, viscous forces are dominant and the flow around the cylinder islaminar and attached to the cylinder surface. Inertial forces become more prevalent asReD is increased, leading to turbulence in the wake, the shear layer, and eventually theboundary layer of the cylinder.

Table 2.1: Flow regimes of a disturbance-free flow around a circular cylinder. Adaptedfrom Zdravkovich [48] and Zawodny [47].

State and Regime ReD Ranges DescriptionLaminar

No Separation 0 to 4-5 Fully-attached flowClosed Wake 4-5 to 30-48 Steady, symmetric, closed wakePeriodic Wake 30-48 to 180-200 Periodically oscillating shear layer

Transition in WakeFar-wake 180-200 to 220-250 Laminar eddy transitionNear-wake 220-250 to 350-400 Irregular eddy transition

Transition in Shear LayerLower 350-400 to 1k-2k Transition wave developmentIntermediate 1k-2k to 20k-40k Transition eddy formationUpper 20k-40k to 100k-200k Eddies burst to turbulence

Transition in Boundary LayersPrecritical 100k-200k to 300k-340k Shear layer transitionSingle Bubble 300k-340k to 380k-400k Asymmetric turbulent separationTwo-bubble 380k-400k to 500k-1M Symmetric boundary layer separationSupercritical 500k-1M to 3.5M-6M Disruption of separation bubblesPostcritical 3.5M-6M to (?) Transition advances toward stagnation line

Laminar to turbulent transition in the boundary layer, shear layer, and wake are all


sensitive to disturbances of the free-stream. These disturbances include but are not lim-ited to freestream turbulence, surface roughness, transverse and streamwise oscillations,and wall blockage. The resulting flow structures in each flow regime can be significantlymodified due to disturbances, typically initiating transition at lower Reynolds numbers.A detailed review of each flow state and regime as well as influencing parameters can befound in [48].

The Strouhal number based on cylinder diameter, StD, is a dimensionless quantityused to characterize the flow around a cylinder and is defined as

StD =fD

U∞, (2.4)

where f is the cylinder vortex shedding frequency. The shedding of vortices resultsin unsteady pressure fluctuations, which are manifested in the acoustic far-field as anAeolian tone [48]. For 300 < ReD < 2 × 105, Bearman [2] found the Strouhal numberremained approximately constant at a value of StD ≈ 0.2. For higher Reynolds numbers(3× 105 < ReD < 7× 105), StD was found to shift in behaviour from a sharp tone to abroader spectral peak, indicating the presence of turbulent flow.

L

D

Λ

U∞

Ur Un

β

Λ

Figure 2.2: Yawed cylinderschematic geometry and nomencla-ture.

A schematic depicting a yawed cylinder geome-try and definitions is shown in Figure 2.2. In thecase of a yawed circular cylinder with yaw angle Λ,the cross-section with respect to the flow becomeselliptical and freestream velocity, U∞, has compo-nents both along the span, Ur, and normal to thecylinder axis, Un. Sears [37] simplified the Navier-Stokes equations for an infinite cylinder to showthat the projected flow on the normal plane is de-scribed by the same equations as the unyawed case,and is independent of the spanwise flow. This sep-aration of variables is known as the IndependencePrinciple. It should be noted that the assumptionof a laminar boundary layer and the omission of endeffects (which are present for cylinders with finiteaspect ratio, L/D) make this an idealization. How-ever, it has been proven to be accurate for numerouscases. It follows that an equivalent Reynolds num-ber, Ren, and Strouhal number, Stn, based on the


normal component of the velocity can be defined by

Ren =ρDU∞ cos Λ

µand Stn =

fD

U∞ cos Λ. (2.5)

The validity of the Independence Principle is dependent on the flow state as definedin Table 2.1. The equivalent Reynolds and Strouhal numbers for vortex shedding wereshown to be valid in the Laminar state for Λ < 45◦ by Hanson [20]. Shirakashi et al. [38]found that in the Transition in Shear Layer state, the Independence Principle holds forΛ . 50◦. For each of these cases, the observed shedding frequency occurred at St ≈ 0.20.In his review of flow around circular cylinders, Zdravkovich reports that the IndependencePrinciple cannot correlate the onset of the Transition in Boundary Layer state [49].

2.2.2 Tandem Cylinders

Parallel Tandem Cylinders

S

(a)

(b)

(c)

U∞

D

Figure 2.3: Tandem cylinder configuration,definitions, and vortex shedding regimes.

Analyzing the flow around landing gear ge-ometries requires some understanding offlow interactions between bluff bodies. Acanonical case for this flow interaction isthe flow around tandem cylinders, withone cylinder behind the other, as definedin Figure 2.3. The separation betweenthe cylinders is typically defined using thespacing ratio, S/D, equal to the ratioof distance between cylinders to cylinderdiameter. In the case of non-identicalcylinders, the spacing ratio is typicallydefined using the upstream cylinder di-ameter. A review of the interactions oftandem cylinders based on the separationis presented in Table 2.2, as describedby Zdravkovich [49]. The spacing ratioregimes are affected by the flow Reynoldsnumber.

The effect of a tandem arrangement on Strouhal number has been examined in theTransition in Shear Layer state and results in three distinct ranges as indicated in Ta-


Table 2.2: Description of tandem cylinder spacing effects.

Spacing Ratio, S/D Description St

1 < S/D < 1.1-1.3 · Forward cylinder shear layers 0.24-0.28envelope downstream cylinder

1.1-1.3 < S/D < 3.5-3.8 · Reattachment of forward shear

0.12-0.15

layers on downstream cylinder

3.0 < S/D < 4.0 · Intermittent vortex sheddingbehind forward cylinder

3.8 < S/D < 5-6 · Synchronized vortex sheddingof both cylinders

S/D > 5-6 · Independent (uncoupled) vortex 0.17-0.19shedding

ble 2.2. The first flow regime (a) involves two closely spaced cylinders producing anarrow wake, resulting in a high St (0.24-0.28). For larger cylinder spacings (b), theupstream cylinder shear layer reattaches to the downstream cylinder, yielding a low St

behind the downstream cylinder (0.12-0.15). Finally for large enough spacing ratios (c),both cylinders behave in vortex street regime exhibiting a slowly rising Strouhal number(0.17-0.19).

S

Figure 2.4: Noise spectra for two 1 inch rodsin a tandem configuration (M∞ = 0.13).Dashed line represents measurements takenwith turbulence generated from an upstreamgrid. Figure adapted from Hutcheson andBrooks [23].

Essentially, the differences in theseshedding patterns can be attributed tothe difference in the local flow state ofeach cylinder. The upstream cylinder hasdisturbance-free oncoming flow and thusbehaves as an isolated cylinder. The down-stream cylinder contends with a velocitydeficit due to the wake and the turbulentshear layer of the upstream cylinder. Asthe spacing between the cylinders is in-creased, the effect of the tandem arrange-ment decreases.

In the Transition in Boundary Layerstate, any changes are only relevant tothe upstream cylinder, as the downstreamcylinder is fully in its turbulent wake. Oka-jima [30] examined two tandem cylinder spacing ratios for ReD < 4.5× 105. His resultsshow approximately constant values of Strouhal number of StD = 0.14 for S/D = 3 andStD = 0.19 for S/D = 5.


The acoustic emissions of tandem parallel cylinder correspond to the vortex shed-ding frequencies and thus the corresponding Strouhal numbers. Some studies have beenconducted in order to quantify the far-field aeroacoustic signature of tandem cylinderconfigurations. Hutcheson and Brooks [23] examined single and multiple rod configura-tions for 3.8× 105 < ReD < 105. Considering the tandem configuration, spacing ratiosof 1,2,3, and 4.5 were investigated for D = 1 inch. The acoustic spectra for each ofthese spacing ratios for smooth cylinders at a Mach number, M∞ = U∞/a0, of 0.13 ispresented in Figure 2.4. In this figure, acoustic emissions are shown in terms of soundpressure levels, SPL, defined as

SPL = 10 log10

(p2rms

p2ref

), (2.6)

where prms is the root mean square sound pressure being measured and pref is a refer-ence pressure (typically 20 µPa). The shedding frequency for a smooth 1 inch diametercylinder at this Mach number was determined in these experiments to be approximately300 Hz (StD = 0.19).

For the tandem cylinder case with the minimum spacing, a single tone was observedat a frequency nearly double the expected shedding frequency of an isolated rod. Forspacing ratios of 2 and 3, the observed primary tone is lower than the expected singlecylinder frequency by approximately 20%. Each of these cases emit additional tones atthe second and third harmonics of the dominant tone frequency. For the largest spacing(S/D = 4.5), a single tone approaching the single cylinder frequency is observed. Thecorresponding Strouhal numbers for the observed tones in Figure 2.4, correspond well tothe regimes defined by Zdravkovich [49].

Yawed Tandem Cylinders

Wilkins and Hall [46] have investigated tandem yawed cylinders in the configurationshown in Figure 2.5. They observed that this arrangement of cylinders resulted in abroadened peak at the shedding frequency for S > 4.5D. It is suggested that the tur-bulence produced by the upstream cylinder disrupts organized vortex shedding leadingto highly three-dimensional flow. This three-dimensionality also resulted in decreasingthe regularity of vortex shedding from the rear cylinder. The study examined yaw an-gles Λ = 10◦, 20◦, and 30◦, finding that the yaw angle was of greater influence to vortexshedding than the cylinder spacing.


D

U∞

UrUn

Λ

Λ

Sn

S

Λ

Figure 2.5: Geometry for tandem cylinder configuration with yawed upstream cylinder.

2.3 Landing Gear Studies

The use of cylinder studies in engineering applications of landing gear noise has itslimitations due to the inherently three-dimensional nature of a landing gear flow field.Wind tunnel tests on landing gear geometries of various scales and fidelities have beenconducted to examine the aerodynamic and aeroacoustic behaviour of representativemodels. A depiction of a typical landing gear and its components are shown in Figure 2.6.Wind tunnel models may include any combination of struts, torque links, bogies, wheels,and finer details such as hydraulic hoses and brakes. Flow measurement techniques suchas hot-wire anemometry and Particle Image Velocimetry (PIV) are used to find regions offluctuating velocity, indicative of noise sources. Unsteady pressure transducers locatedon model surfaces have also aided in the detection of sound production regions. Theradiated noise is measured using microphones positioned around the model. Microphoneslocated at different angles allow for acoustic directivity patterns to be determined. Phasedmicrophone arrays are commonly used to localize noise sources through a processes knownas beamforming. Beamforming results in a series of spatial maps depicting the soundlevels for given frequency ranges. The data from these studies has been used to validatecomputational aeroacoustic codes as well as provide references for noise prediction toolsthat can be used in aircraft design. These studies are briefly summarized below.


Main (Shock)

strut

Drag strut

Stabilizer

brace

Main gear pin

Auxiliary

actuator

Main actuator

Fairings

Proximity

sensors

U∞

Figure 2.6: Bombardier Dash 8 Q400 mainlanding gear assembly. Figure adapted fromwww.smartcockpit.com [6].

Early studies investigating airframenoise performed in the 1970’s indicatedthat landing gear are important contribu-tors to overall noise levels. The first studyto isolate landing gear for noise measure-ments was conducted by Heller and Do-brzynski [22], using a rudimentary two-wheel landing gear consisting of only astrut, axle, and wheels. Landing gear noisewas identified as a predominantly low fre-quency phenomenon. This was later cor-rected after Dobrzynski and Buchholz’sfull-scale testing of Airbus A320 and A340landing gear in 1997 [10], which showed abroadband noise signature spanning from200 Hz - 10 kHz.

Despite the fact that it is necessary to include the fine geometric details of a landinggear to acquire a thorough description of its resulting sound field, many conclusions canbe drawn by investigating simplified geometries.

2.3.1 Simplified Geometries

Testing of simplified landing gear geometries is commonplace due to the difficulties as-sociated with replicating fine details down to model scale sizes. Some important resultshave been reported from testing of these geometries. In the tests of Dobrzynski andBuchholz [10] as well as those of Guo et al. [18], it was shown that the addition of detailsonto a rudimentary model manifests itself as a broadband increase in sound levels withthe greatest gains in the high frequency range. Flow structures developed due to wakebody interactions between closely spaced components have been identified as importantairframe noise contributors. These regions include the gaps between wheels [25, 32] andthe junctions of the strut and axle [32].

2.3.2 High Fidelity Geometries

The full-scale landing gear tests of Dobrzynski and Buchholz were the first to demonstratethat landing gear noise is fundamentally broadband in nature [10]. These noise levelsgenerally had a sixth power scaling relative to flow velocity. In the middle frequencies


of the two wheel gear emissions, the behaviour differed which was attributed to velocity-independent noise generation between the wheels. In 2006, Guo et al. [18] tested ahigh-fidelity model-scale Boeing 737 main landing gear, identifying dominant sourcesdownstream of the main strut and front wheels. Small components were removed for sometests showing broadband decrease in acoustic power (most noticeably at high frequencies).

Turbulence generated from the landing gear assembly was proposed as a primarysource of high frequency noise as seventh-power velocity scaling was the best fit in thatregion of the acoustic spectrum. Other high-fidelity tests on Boeing landing gear byRavetta et al. [33] showed that the removal of the drag brace produced a reduction inoverall sound levels. Ringshia et al. [35] revisited this work and showed high turbulence inthe wake region of the drag brace and that the vortex formed in the strut wake impingeson the front face of the rear drag brace which may be a significant noise source.

Experimental investigations of models of various fidelities by Guo et al. [19] showedthat the noise spectrum of landing gear can be decomposed into three frequency ranges.Each range represents contributions from the wheels (low frequencies), struts (mid fre-quencies), and small details such as cut-outs and hoses (high frequencies). In a predictionmodel based on the Strouhal numbers of major components, Guo et al. were able to in-troduce a ‘complexity factor’ to account for the effects of small details, with satisfactoryresults.

2.4 Propeller Flows

The current work aims to investigate the effects of an upstream propeller on the flowover a landing gear. Since Guo et al. [17] showed that the local flow conditions affect theradiated noise, a review of the flow structures and acoustics related to propeller flows arediscussed below.

2.4.1 Flow Structures

A propeller consists of equally spaced blades rotated by the torque of an engine. Thedirect result of the rotating blades is a pressure increase behind the propeller and adecrease of pressure ahead of it. This pressure difference is a measure of propeller thrust.In the wake region aft of an actuating propeller, vortices are shed downstream in almosthelical paths and represent the propeller slipstream. Vortices are shed both by the trailingedge of each blade as well as the blade tips.

In his work regarding the propeller-wing aerodynamic interference, Veldhuis provides


an excellent summary of the characteristics of propeller slipstream flow [44]. Resultsconcerning the axial and swirling velocity profiles as well as slipstream contraction arepresented below. Figure 2.7 is a schematic of the major flow structures associated withthe propeller slipstream.

U∞ tip vortex

blade vortex sheet

slipstream tube

root vortex

rolled up

vortex system

Figure 2.7: Helical vortex system and slipstream tube generated by a propeller. Figureadapted from Veldhuis [44].

A propeller is functionally used to provide thrust in most applications and thus onewould expect the axial velocity behind a propeller to be greater than the freestreamvalue. Just behind the propeller, the axial flow velocity is equal to U∞ + Ua, the sum offreestream velocity and the propeller-induced velocity. Far downstream of the propeller,the axial flow velocity in the propeller slipstream tube reaches U∞ + 2Ua.

A typical velocity profile aft of a propeller is shown in Figure 2.8 for both the axialand tangential velocities, Ua and Ut, respectively. The induced axial velocity is shownto reach a maximum near 3

4R, where R is the propeller radius, typically at a location

of maximum blade loading. The tangential (or swirl) velocity distribution is dependenton the loading conditions of the blades and is susceptible to change for different advanceratios [44]. The swirl angle of a blade vortex sheet, θsw, can be computed using

θsw = tan−1(

Ut

Ua + U∞

). (2.7)

Since the axial velocity is a function of the streamwise position and the tangential velocityis not, the swirl angle changes along the length of the slipstream. Thus, the incident flow


angle on any downstream components should be calculated based on their location behinda propeller.

U

a

t ∞

U

UU / , U / Ua ∞

t

0.05

Figure 2.8: Typical radial distribution of axial velocity, Ua, and tangential velocity, Ut,directly behind a propeller. Figure adapted from Veldhuis [44].

The slipstream of a propeller contracts to maintain a constant mass flow as velocityis increased through the propeller. The diameter of the slipstream tube downstream ofthe propeller should be computed to identify which regions are affected by the propeller.Veldhuis [44] presents an approximation of the contraction ratio,

Rs

R=

√√√√ 1 + a

1 + a(

1 + x√R2+x2

) , (2.8)

where Rs is the radius of the slipstream at location x, and a is the axial inflow factorrepresenting the velocity at the propeller plane.

A dimensionless quantity frequently used to compare propeller operational settingsand performance is the advance ratio, J , defined as

J =U∞nDp

, (2.9)

where n is the rotational rate of the propeller and Dp is the propeller diameter. Thisquantity indicates the ratio between the distance the propeller moves forward per rev-olution and Dp. The advance ratio is a measure of effective pitch and may be used todetermine the incoming angle of the fluid relative to a propeller blade.


2.4.2 Acoustics

Increased interest in the aviation industry with respect to turboprop aircraft and otheropen rotor technologies, has resulted in research concerning propeller noise reduction.Trebble et al. [43] have performed some aeroacoustic studies of both model- and full-scale propellers. The authors report both near- and far-field acoustic measurements forvarious operating conditions, modifying rotational rates, blade-angle settings, and windtunnel flow speeds. In order to compare the full-scale and model-scale datasets, thepropeller blade tip Mach numbers, Mtip, were matched for the two cases.

Mtip =vtipa

=

√U2tip + U2

∞

a(2.10)

Presented in Figure 2.9 is a noise spectra comparison between full-scale and 1/4-scale models normalized to 1600 RPM = 26.7 Hz. This corresponds to a blade passingfrequency of 106.7 Hz (four-bladed propeller), which can be observed as the location ofthe first, and largest tonal peak. Good correlation exists between the observed value ofsound pressure levels of the tonal harmonics in the spectra for both propellers. Thougha discrepancy exists in the broadband levels, this is attributed to contamination fromthe wind tunnel background noise. The number of measurable harmonic tones was foundto increase with the propeller speed. The sound pressure level of the fundamental tonealso increased with propeller rotational speed and propeller blade-setting angle (increasedthrust and torque). When the propellers were lightly loaded, they exhibited a rise in thehigh frequency content of the broadband noise levels.

120

100

80

60

0 0.2 0.4 0.6 0.8 1.0 1.2 1.4

dB

SPL

1.6Equivalent full-scale frequency (kHz)

Full-scale

¼-scale

m=1

m=2

m=3

m=4

115.0

106.1

100.9

95.5

116.8

106.4

98.9

91.5

118.9

103.9

100.3

96.9

SPL

Full-scale¼

scale MEASURED CORRECTED

Figure 2.9: Far-field noise spectra comparison between full-scale and 1/4-scale propellermodels normalized to 1600 RPM. Figure adapted from Trebble et al. [43].

3 | Wind Tunnel Characterization

The following sections detail the work done to characterize the flow and acoustic prop-erties of the acoustic wind tunnel at the University of Toronto Institute for AerospaceStudies (UTIAS). These tests were completed in order to quantify the performance of thewind tunnel, as multiple modifications have taken place since its previous experimentaluse. The wind tunnel was constructed in 1975 and has since been used for various stud-ies ranging from investigations of Tollmien-Schlichting instabilities to marine propellerperformance. Since 2011, changes include:

• repair of the motor windings,

• adjustments of the fan blades,

• installation of a variable frequency drive (VFD) to adjust the motor rotational rate,

• installation of anechoic wedges in the test chamber,

• application of acoustic treatments to the tunnel contraction and nozzle, and

• construction of a downstream silencer.

3.1 Facility Overview

The experiments presented in the current work were conducted at UTIAS in the AcousticWind Tunnel Facility, a photograph of which is shown in Figure 3.1. This facility featuresan open-circuit, open jet, suction-type wind tunnel. The nozzle diameter, DN , of theopen jet is 0.70 m, and the nozzle-to-collector distance is 2.78 m. The wind tunnel canbe operated at a maximum flow speed of U∞ ≈ 60 m/s.

The wind tunnel is powered by a 150 h.p. motor driving a nine-bladed axial fanwith a maximum RPM of 1800. Air is drawn from a geodesic dome 55 m in diameterto mitigate atmospheric effects on the inlet conditions. The flow is then conditionedthrough an aluminum hexagonal honeycomb of 1.59 mm cell width by 50 mm deep, a

17

Chapter 3. Wind Tunnel Characterization 18

Figure 3.1: Photograph of the Acoustic Wind Tunnel Facility (right) and the adjacentgeodesic dome (left).

120 mesh (120 strands per inch) screen and four 54 mesh screens for turbulence reductionas described by Ball [1]. The conditioning devices lead into a 10:1 contraction which isimmediately followed by a square to circular transition and the nozzle to the open jettest section. Flow continues from the jet collector into the diffuser toward the motorand fan. On either side of the motor and fan, silencers attenuate noise from propagatinginto the tunnel test section where measurements are taken or outside near the facility.A plenum with catcher screens protects the fan blades from debris flowing through thetunnel. Figure A.1 in Appendix A shows a schematic of the entire Acoustic Wind TunnelFacility.

The issues associated with wall reflections in closed-section wind tunnels for aeroa-coustic measurements can be resolved by placing the test section in an anechoic environ-ment. In this case, open jet wind tunnels are preferable to closed section facilities since anacoustically treated room can be easily built around an open test section. The open jetof the Acoustic Wind Tunnel Facility is located within a chamber treated with fibreglasswedges. A schematic of the stepped wedge geometry used in the anechoic chamber isshown in Figure 3.2. Each wedge block is rotated 90◦ to its adjacent wedges. The wedgescan effectively absorb wavelengths greater than four times the height of the wedge. As-suming the speed of sound in air, a0, is constant (343 m/s at standard temperature andpressure), the frequency, f , of a sinusoid and its wavelength, λ, are related by

a0 = fλ, (3.1)

and given a quarter-wavelength equal to the height of the wedge, 0.610 m (24 inches),the resulting theoretical cutoff frequency is equal to approximately 140 Hz. Thus it isadvised that experiments conducted in the UTIAS acoustic wind tunnel be designed toexamine frequencies greater than 140 Hz since investigations of lower frequencies willrequire reflected sound waves to be taken into account.


61122

23

6146

30

5.1

Figure 3.2: Schematic of the anechoic wedge geometry. Dimensions are in centimeters.

3.2 Aerodynamic Survey

3.2.1 Experimental Method

A survey was conducted in the open jet of the wind tunnel test section to examinethe spread of the jet shear layer, the shape of the jet potential core (region with nearlyuniform flow velocity), and the turbulence levels within the jet. Two linear traverses wereplaced under the test section allowing for movement in the x and r directions as definedin Figure 3.3. The traverses were operated using MATLAB through a Motion Group 4-channel motion controller via a serial cable. A pitot-static tube and single hot-wire probewere near the jet centerline and were traversed to the grid locations shown in Figure 3.3.Pressure measurements were acquired using an MKS Instruments Type 120AD Baratron100 Torr differential pressure transducer, which has an accuracy of 0.12% of the reading.Pressure measurements from a pitot-static tube correspond to the dynamic pressure, q∞and can be related to the velocity, U∞, through the incompressible form of Bernoulli’sequation:

p0 − p∞ =1

2ρU2∞ = q∞ (3.2)

The hot-wire used was a 5 µm copper-plated tungsten wire with a sensing lengthof 1 mm. Data were taken through a DISA (Dantec) 55M10 Constant TemperatureAnemometer (CTA) operating at an overheat ratio of 0.6. Prior to each test run, themeasurement hot-wire was calibrated in-situ against a pitot-static tube. This was done


placing the pitot-static tube and hot-wire probes about the tunnel centreline, 1 cm apartand equating measurements from the CTA with the measured dynamic pressures fora range of 26 flow speeds between 5 m/s and 55 m/s. A J-type thermocouple wasmounted in the tunnel collector to provide a temperature correction to hot-wire data.Temperature fluctuations for the duration of each test were typically ±2◦C. Hot-wiredata and pressure transducer data were acquired at a rate of 4 kHz for 30 seconds usinga National Instruments PCI-6361 A/D data acquisition system at each measurementlocation. The overall uncertainty in the mean velocity profiles was estimated to be lessthan 1%.

x

r

278.1

120.0

10.0 typ.

48.080.0

4.0 typ.

1.0 typ.

Figure 3.3: Measurement grid for aerodynamic survey. Dimensions are in centimeters.

3.2.2 Mean Flow Measurements and Turbulence Levels

The mean flow velocity profiles at the measured downstream stations for Mach numbersof 0.10 (U∞ = 33 m/s) and 0.15 (U∞ = 51 m/s) are presented in Figures 3.4a and 3.4b,respectively. For M∞ = 0.10, the accelerated flow near the lip of the nozzle presentsan uneven velocity profile until approximately x/DN = 1. However, the flow remainsuniform over the majority of the jet width beginning at x/DN ≈ 0.3. A similar trendis observed for the survey conducted at M∞ = 0.15. The jet shear layer appears togrow identically for both test cases, reaching the extent of the measurement grid atx/DN = 0.71.

The power spectral density of turbulent fluctuations along the jet axis for M∞ = 0.10

and M∞ = 0.15 have been estimated from hot-wire data and are presented in Figure 3.5.Both Figures 3.5a and 3.5b show a peak in the spectra for x/DN > 0.99 occuring ata Strouhal number based on DN between 0.7 and 0.8. This is identical to the peak inthe turbulent power spectral density shown by Ball for this facility in 1983. Michel andFroebel [29] identify a broad hump at St = 1 as a perturbation inherent to free jets,


where turbulent eddies in the shear layer create near-field pressure gradients which causefluctuations in the flow. This spectral peak is shown to increase with distance from thejet nozzle.

Turbulence intensity, defined as

Tu = u′/U∞ (3.3)

where u′ is the root mean square of the velocity fluctuations, was determined for M∞ =

0.10 and 0.15. The power spectral density of the turbulent fluctuations were integratedabove 10 Hz to compute u′. Turbulence intensity along the jet axis is presented in Figure3.6. In the potential core, the turbulence intensity remains under 0.5% for x/DN lessthan 1.4. The radial variations in turbulence intensity for M∞ = 0.10 are presented inFigure 3.7. Though irregular variations are evident near the nozzle, a consistent profileis achieved by x/DN = 0.57. By x/DN = 0.99 the region of Tu ≤ 0.1% is restricted to0.6DN .

3.3 Background Acoustic Levels

A Bruel & Kjaer 4192 free field condenser microphone was placed 0.7 m downstream ofthe nozzle and 2.1 m away from the jet centreline to measure background noise levelsat different jet exit velocities. The background sound pressure level (SPL) narrowbandspectra for a range of jet exit velocities between M∞ = 0.04 and 0.16 are presented inFigure 3.8. The measurements were taken at a sampling rate of 32 kHz for 180 secondsusing a National Instrument PCIe-6259 A/D data acquisition card. Flow velocities weredetermined using a pitot-static tube located in the jet collector. The broad peaks occuringat the low frequencies of the spectra can be attributed to jet shear layer impinging onthe wind tunnel collector and are similarly located to the peaks found in the turbulencespectra. Strouhal scaling of the sound pressure level data results in a collapse of thehigh frequency data and a shift of the low frequency peak to lower Strouhal numbers asvelocity is increased. Excessive background noise due to recirculating flows induced bythe jet may be due to the omission of wind screens on far-field microphones [41]. Between100 Hz and 1 kHz the broadband background noise of this acoustic wind tunnel drops ata rate of 20 dB per decade. This result is comparable to those presented by Mathew atthe University of Florida Aeroacoustic Flow Facility for speeds between 18 and 46 m/s[28] .


020

40

-0.5

-0.250

0.25

0.5

0

.14

020

40

020

40

0.2

8

020

40

0.4

3

020

40

0.5

7

020

40

0.7

1

ve

loc

ity

, m

/s

020

40

0.8

5

020

40

0.9

9

020

40

1.1

4

020

40

1.2

8

020

40

1.4

2

020

40

1.5

6

020

40

1.7

00

.00

x/D

N =

r/DN

(a)M∞

=0.10

0.1

4

020

40

60

020

40

60

0.2

8

020

40

60

0.4

3

020

40

60

0.5

7

020

40

60

0.7

1

020

40

60

0.9

9

020

40

60

1.1

4

020

40

60

1.2

8

020

40

60

1.4

2

020

40

60

1.5

6

ve

loc

ity

, m

/s

020

40

60

0.8

5

020

40

60

1.7

0

020

40

60

-0.5

-0.250

0.25

0.5

x/D

N =

0.0

0

r/DN

(b)M∞

=0.15

Figure3.4:

Ope

njetvelocity

profi

lesof

theAcousticW

indTu

nnel

forMachnu

mbe

rsof

(a)0.10

and(b)0.15

.


10-1

100

101

101

102

103

10-7

10-6

10-5

10-4

10-3

10-2

x/DN

= 0.00

x/DN

= 0.28

x/DN

= 0.57

x/DN

= 0.85

x/DN

= 1.13

x/DN

= 1.42

x/DN

= 1.70

Po

we

r S

pe

ctra

l De

nsi

ty

St = fDN/U

∞

Frequency, Hz

(a) M∞ = 0.10

101

102

103

10-7

10-6

10-5

10-4

10-3

10-2

x/DN

= 0.00

x/DN

= 0.28

x/DN

= 0.57

x/DN

= 0.85

x/DN

= 1.13

x/DN

= 1.42

x/DN

= 1.70

Po

we

r S

pe

ctra

l De

nsi

ty

Frequency, Hz

10-1

100

101

St = fDN/U

∞

(b) M∞ = 0.15

Figure 3.5: Power spectral density of wind tunnel jet longitudinal velocity fluctuations,measured on-axis, for (a) M∞ = 0.10 and (b) M∞ = 0.15.

0 0.2 0.4 0.6 0.8 1 1.2 1.4

0.005

0.01

0.015

0.02

M∞

= 0.10

M∞

= 0.15

Tu

= u

’/U

∞

x/DN

Figure 3.6: Turbulence intensity in the wind tunnel open jet vs. axial position, measuredon-axis, for M∞ = 0.10 and M∞ = 0.15.


-0.6 -0.5 -0.4 -0.3 -0.2 -0.1 0 0.1 0.2 0.3 0.4 0.5 0.6

10-3

10-2

10-1

x/DN

= 0.00

x/DN

= 0.28

x/DN

= 0.57

x/DN

= 0.99

x/DN

= 1.42

Tu

= u

’/U

∞

r/DN

Figure 3.7: Turbulence intensity in the wind tunnel open jet vs. radial position forM∞ = 0.10.

100

101

102

103

104

30

40

50

60

70

80

90

100

110

120

M∞

= 0.04

M∞

= 0.07

M∞

= 0.10

M∞

= 0.13

M∞

= 0.16

SP

L, d

B r

e 2

0 µ

Pa

Frequency, Hz

(a)

10-2

10-1

100

101

30

40

50

60

70

80

90

100

110

120

M∞ = 0.04

M∞ = 0.07

M∞ = 0.10

M∞ = 0.13

M∞ = 0.16

SP

L, d

B r

e 2

0 µ

Pa

St = fDN/U

∞

(b)

Figure 3.8: Background narrow-band sound pressure level spectra of the acoustic windtunnel plotted against (a) frequency in hertz and (b) Strouhal number based on nozzlediameter.

4 | Experimental Method

4.1 Overview

The purpose of the current work is to determine the baseline aeroacoustic behaviour fora configuration involving a model landing gear in the unsteady wake of a propeller. Thelanding gear geometry considered includes an inclined drag strut ahead of the main strutwith wheels. Of primary interest to this study are the measurements of the far-fieldacoustics and the unsteady surface pressures. When acquired simultaneously, this datacan aid in deducing fluctuations which result in far-field sound emissions from the model.Coherence and cross-correlations between the far-field and surface microphones can pro-vide insight into noise generation as shown by Siddon [39]. Experiments attempting toidentify the contributions of various geometries and configurations to the surface andfar-field (acoustic) pressures are described below.

4.2 Model Description

Figure 4.1: Image of landing gear model inwind tunnel test section.

An image of the tested model is shown inFigure 4.1. The model geometry is basedon the main landing gear of the Bom-bardier Dash-8 Q400, a regional turbopropaircraft, at a scale of 1:10.8. This scalingensures that an upstream model propellerand the landing gear can be located en-tirely within the jet potential core of thewind tunnel. The model consists of anelectric motor driving a 15 inches (38.1 cm)diameter propeller, a tubular nacelle, amain strut, a drag strut, an axle, and two

25

Chapter 4. Experimental Method 26

28°

Ø 70.0

Ø 8.0

1.8

7.8

Ø 38.1

27.0

6.0

Ø 9.0

1.0 28.7

42.0

1.4

Figure 4.2: Schematic drawing of the complete landing gear model geometry with anupstream propeller in the wind tunnel test section. The propeller is rotated 45◦ betweenimages for clarity. A two-bladed propeller was used for tests at M∞ = 0.10, whereas athree-bladed propeller was used for M∞ = 0.04 and 0.07. Dimensions are in centimeters

wheels. The landing gear components are primarily constructed of aluminum. The treadof each wheel and a removable strip on the main strut were made of ABS plastic to allowfor the complicated cut-outs that accommodate the surface sensors. Although small-scalefeatures such as hoses, cables, and cut-outs are not replicated, it is possible to predictthe additional noise of such small features by introducing a ‘complexity factor’ in futureprediction models [16]. A schematic drawing of the landing gear geometry with respectto the open jet nozzle is shown in Figure 4.2.

A Power 46 Brushless Outrunner Motor was used to turn the propeller along using a60-Amp Pro Switch-Mode BEC Brushless Electronic Speed Controller, both from E-flite.Power for the motor was provided by a TRC Electronics SP-750-12 (12V) single outputpower supply. Throttle was controlled through USB communication with an Arduinomicrocontroller. An optical tachometer was constructed using an phototransistor switch(OPB606A) to count propeller revolutions and was placed within the nacelle, adjacent tothe motor. The back of the propeller hub was covered completely with black tape save forone reflective strip, such that the phototransistor switch would provide a signal once perrevolution. An Eagle Tree Systems eLogger V4 was also used to monitor power consump-tion of the system, throttle levels, temperature, and motor RPM (through a brushlessmotor sensor). The constructed tachometer and the brushless motor RPM sensor werefound to be within ±50RPM for all tests. This corresponds to a maximum relative errorof ± 2% for the lowest RPM case and an absolute uncertainty of ± 0.83Hz. Since the av-eraging method of the eLogger was unknown, the RPM of the optical tachometer is usedas the reference value and verified against the provided eLogger data recorder softwareto ensure a consistent reading.

Table 4.1 summarizes the model geometries, configuration names, and flow speeds for


all test cases. Configuration elements were incrementally added to the model assemblyto give additional information regarding the contributions of each isolated component tothe total sound emissions.

The motor speed in each case was adjusted to achieve maximum thrust while main-taining motor current at safe levels. Two propellers of diameter Dp = 38.1 cm, wereused for the tests described in this section. At M∞ = 0.04 and 0.07, a Master Airscrew3-blade 15 × 7 propeller was used in order to achieve a blade passing frequency (BPF)of the propeller exceeding the cutoff frequency of the anechoic chamber (140Hz). AtM∞ = 0.10, the 3-blade propeller was unable to produce a net positive thrust, so it wasreplaced with a Master Airscrew 2-blade 15 × 8 propeller. At this wind tunnel speed,the 2-blade propeller was able to operate at a suitable RPM such that the BPF exceedthe chamber frequency cut-off limit. A radial survey of the slipstream for each propelleroperating condition was conducted in the wind tunnel using a pitot-static tube locatedat a position 0.25Dp downstream of the actuation plane as was done by Beveridge [4].An MKS Type 120AD Baratron differential pressure transducer was used to determinethe dynamic pressure at each survey location and thus the velocity. Figure 4.3 shows thetime-averaged velocity profile behind the three-blade propeller at the maximum operat-ing condition for all test speeds. The values of the advance ratio, J , for the propelleroperating at each freestream velocity are marked in the legend of Figure 4.3. Thrustvalues, T , were calculated using

T = 2πρ

∫ ∞r=r0

(UaU∞ − U2

a

)dr, (4.1)

which is a result from the momentum equation. Integration was done using the trape-zoidal method and each thrust coefficient, CT , was computed using

CT =T

ρn2D4p

. (4.2)

The computed thrust coefficients for the three-blade propeller at M∞ = 0.04, 0.07, and0.10 are 0.049, 0.013, and -0.007, respectively. The negative value is an indication thatthe generated thrust could not overcome the drag of the propeller blades in this case.The two-blade was also tested at M∞ = 0.10 and was shown to have a thrust coefficientof 0.013 operating at maximum throttle. Figure 4.4 compares the velocity profiles of thetwo-blade and three-blade propellers for M∞ = 0.10.


Table 4.1: Experimental test matrix for acoustic measurement tests. *All tests conductedat M∞ = 0.10 used a 2-blade propeller.

Config. Speed Main Strut Drag Rotating Propeller SchematicName (M∞) & Wheels Strut Motor

S-1 0.04, 0.07, 0.10*

S-2 0.04, 0.07, 0.10* X

S-3 0.04, 0.07, 0.10* X X

M-1 0.04, 0.07, 0.10* X

M-2 0.04, 0.07, 0.10* X X

M-3 0.04, 0.07, 0.10* X X X

P-1 0.04, 0.07, 0.10* X X

P-2 0.04, 0.07, 0.10* X X X

P-3 0.04, 0.07, 0.10* X X X X


4.3 Instrumentation

4.3.1 Far-field Microphones

Far-field acoustic spectra were acquired using two Bruel & Kjaer 1/2 ” 4192 free-fieldcondenser microphones, mounted on a linear traverse in the anechoic chamber. Far-field measurements were taken along the wind tunnel centerline at locations F1 to F5,and also 15◦ above the model nacelle at locations G1 to G5, as indicated in Figure 4.5.The farthest upstream measurement locations possible due to space restrictions werepositions F1 and G1, which are 12 cm behind the propeller plane. Position F3 is collinearwith the model main strut when it is installed. To be considered in the acoustic far-field, measurements should be taken at least one wavelength, λ, from the source. Forthese experiments, the far-field acoustic measurement plane is situated 2.10m from thetunnel centreline. Assuming the speed of sound, a = 343m/s, this results in a minimumfrequency of 163Hz. Any frequencies below this value cannot be considered to be in theacoustic far-field.

Pistonphone Calibration

A Bruel & Kjaer Type 4428 pistonphone was used to calibrate the far-field microphones.The pistonphone produces a 250Hz tone at a known amplitude that is corrected tothe local atmospheric pressure conditions. Each far-field microphone was separatelycalibrated with the pistonphone and their sensitivities recorded. The sensitivities forthe microphones measuring at locations F1 - F5 and G1 - G5 were determined to be10.63mV/Pa and 12.02mV/Pa, respectively.

4.3.2 Surface Microphones

Unsteady Surface Pressure Sensor Locations

Measuring the unsteady surface pressures that manifest themselves in the acoustic far-field as sound is paramount in this investigation. The small size of the landing gearmodel allowed for only a small number of surface locations to be investigated. Studieson tandem cylinders have demonstrated that downstream components experience muchhigher pressure fluctuations than those experienced by upstream ones [24]. The primarycause of this was determined to be the impingement of the shear layer of the upstreamcomponent onto the downstream one. Thus, the effects of the drag strut on the oncomingpropeller flow were of primary interest in this study. Sensors with a high dynamic range


were selected to enable the measurement of the large pressure fluctuations anticipatedwith this configuration.

Unsteady surface pressure measurements were acquired on the model at the locationsindicated in Figure 4.6. The spacing ratios, S/D, at each microphone location basedon the upstream drag strut diameter and normal distance from the yawed cylinder arelisted in Table 4.2. Knowles Electronics EK-26899-P03 microphones were affixed withepoxy into cut-outs on the wheel tread and the main strut to provide this pressurefluctuation data. There were a total of 8 sensor locations: 4 along the span of the mainstrut facing the oncoming flow, and 2 on each wheel, one directly facing oncoming flowand another 45◦ below it on the wheel tread surface. These locations were strategicallychosen to investigate the presence of flow structures associated with the propeller andtheir interactions with the drag strut. Note that the positions of the wheel microphonesare adjustable by rotation of the wheel. However, only the stated locations were usedin this study. Microphone wires have been routed through the model and out of theairstream, so as not to influence the flow-field and acoustics.

Table 4.2: Spacing ratios, Sn/D, at the unsteady surface pressure sensor locations basedon drag strut diameter and normal distance between cylinders.

Microphone Spacing Ratio,Location Sn/D

A 7.92B 6.89C 5.72D 4.54

Calibration of Surface Sensors

Characterization of the surface (measurement) microphones was conducted in-situ withthe use of a function generator, speaker, and calibrated free-field (reference) microphone.The reference microphone was placed adjacent the surface sensor with their diaphragmsparallel to one another and 0.2 cm apart. The centre of the speaker cone was aligned withthe centre of the two microphones, approximately 5 cm away. The function generator wasused to perform a logarithmic sweep through a frequency range from 100Hz to 8 kHz at2Vpp. Data were then acquired from both sensors in order to calibrate the measurementmicrophone relative to the reference. This method is similar to the linear system approachpresented by Snarski [40]. Details of the calibration procedure used in the flow controllaboratory can be found in the internal report by Zhao [50]. The sensitivity of the surfacesensor, Sm, was computed using


Sm =SrefAref|Gxy|AmGxx

(4.3)

where Sref is the sensitivity of the reference microphone, Aref is the gain of the referencemicrophone, Gxx is the estimated power spectrum of the reference microphone, Am is thegain of the measurement microphone, and Gxy is the estimated cross spectrum of the twosignals.

A representative calibration curve from the surface sensor calibration procedure isshown in Figure 4.7. A nearly flat region in the sensitivity was observable for all mea-surement microphones in the range of 200Hz to 1 kHz. The average value of the sensitivityover this frequency interval was selected as the nominal sensitivity for all acoustic tests.

4.3.3 Data Acquisition & Uncertainty

In addition to microphone measurements, flow temperature and speed were collectedduring each test run to determine the flow Mach number and fluid properties, such asviscosity. Atmospheric pressure was obtained prior to each test from the website of En-vironment Canada [11]. Flow speeds were measured via a pitot-static tube in the windtunnel collector connected to an MKS Type 120AD Baratron differential pressure trans-ducer operating in 10 Torr mode, since the flow speeds of these experiments were less thanthat of the aerodynamic survey described in Section 3.2.1. Temperature measurementswere made using a J-type thermocouple positioned in the wind tunnel collector.

Surface microphone signals were passed through a circuit designed to increase thedynamic range of sensors. The unsteady pressure signals were then passed throughamplifiers with nominal gains of 20 dB and built-in bandpass filters of 20Hz - 10 kHz. Allunsteady signals were passed through 10 kHz antialiasing filters and were AC-coupled tothe data acquisition unit to remove any DC bias. A schematic of the instrumentationlayout is provided in Figure 4.8.

All data were sampled for 180 seconds at 32,878Hz using a 16-bit National Instru-ments PCIe-6259 A/D data acquisition card, yielding a total number of samples per testrun of N = 5.982× 106. In order to resolve the tonal peaks associated with the propellernoise emissions, a frequency resolution of ∆f = 2Hz was achieved by dividing the datainto Nfft = 16, 384 samples per block. The number of records averaged was nrec=361.

The unsteady pressure measurements at the model surface and in the far-field aresubject to both bias and random error. The root mean square error of spectral data isexpressed by


ε =√ε2r + ε2b . (4.4)

For each sensor, the normalized random error in the autospectrum, Gxx, of the measureddata can be calculated using the following equation from Bendat and Piersol [3]:

εr =1√nrec

. (4.5)

The normalized bias error is estimated using the equation below, also from [3]:

εb =b[Gxx]

Gxx

≈ ∆f 2

24

G′′xxGxx

, (4.6)

where b[Gxx] is the bias in the autospectral estimate. Since this equation is difficult tocompute for a random system and the bias error is expected to be small relative to therandom component, the bias error is neglected. This yields a total error, ε, equal to5.26% for the spectral estimates of the measured data.


0.4 0.6 0.8 1 1.2 1.4 1.6 1.8

r p /

Dp

Ua / U

∞

0

0.1

0.2

0.3

0.4

0.5

0.6M∞ = 0.10, J = 0.52

M∞ = 0.07, J = 0.54

M∞ = 0.04, J = 0.34

Figure 4.3: Propeller wake time-averaged velocity profiles for the 3-blade, MasterAirscrew 15× 7 propeller at M∞ = 0.04, 0.07 and 0.10, at maximum throttle.

0.4

r p /

Dp

Ua / U

∞

0

0.1

0.2

0.3

0.4

0.5

0.6

1.20.6 0.8 1

3-blade, J = 0.52

2-blade, J = 0.54

Figure 4.4: Propeller wake time-averaged velocity profiles for the 3-blade, MasterAirscrew 15×7 propeller and the 2-blade, Master Airscrew 15×8 propeller atM∞ = 0.10and maximum throttle.


216FLOW

15 °

F1

F2

F3

F4

F5

G1

G2

G3

G4

G5

1020

20

20

58

Figure 4.5: Schematic drawing of far-field microphone test locations. Dimensions are incentimeters.

11.113.7

16.218.7

28.731.9

L1R1

L2R2

A

D

B

C

Figure 4.6: Schematic drawing of model surface microphone locations. Drag strut omittedfor clarity. Dimensions are in centimeters.


F requenc y, kHz

Se

ns

itiv

ity

, m

V/P

a

1 2 3 4 57.5

7.6

7.7

7.8

7.9

8

8.1

8.2

8.3

8.4

8.5Mic rophone A

Figure 4.7: Sample calibration curve for a surface microphone sensor (Microphone A).Solid line represents the mean and dashed line represents one standard deviation of themean between 200Hz and 1 kHz, respectively.


OPB606A Re�ective

Object SensorPhototransistor

Switch

EK-26899-P03

Surface Microphone

Sensors

E-�ite Power 46

Brushless Outrunner

MotorE-�ite 60 Amp

Electronic Speed

Controller

ESC Temperature

Sensor

Brushless Motor

RPM SensorEagle Tree Systems

eLogger v4

Arduino Duemilanova

Microcontroller

High SPL

Circuit

20 dB Signal

Ampli!er

Antialiasing

Filter

National Instruments

PCIe-6259 A/D

Data Acquisition CardAeroacoustics

Laboratory

Computer

The Motion Group

MMC-43-3C

Motor Controller

Streamwise Traverse

Bruel & Kjaer

Type 4192, 1/2”

Condenser Microphones

Larson-Davis

2200C Microphone

Power Supply

Figure 4.8: Instrumentation layout for data acquisition .

5 | Results & Analysis

5.1 Far-field Acoustics

As indicated in Table 4.1, far-field acoustics were measured for nine different configura-tions tested at three wind tunnel Mach numbers. In the following sections, spectra frommeasurements taken at location F3, as shown in Figure 4.5, are presented.

5.1.1 Static Configurations

Figure 5.1 shows the far-field SPL spectra for all the configurations at each wind tunnelflow speed. Background far-field measurements of the empty test section are also plottedfor comparison. Without the propeller or motor actuating, the levels for each configura-tion are nearly indistinguishable from one another as shown in Figure 5.1a. Figure 5.1bfocuses on the range of 100Hz to 2000Hz, where sharp tonal peaks are observed near1 kHz. It is suspected that these tones result from whistling across small gaps around thetest chamber door. Despite numerous efforts to completely seal the door, narrow peaksin the range of 800Hz to 1.3 kHz occasionally are apparent in the far-field spectra. In thecase ofM∞ = 0.10, two small broad peaks rising 4 dB over background levels are observedcentred at 0.8 kHz and 1.1 kHz respectively for S-1 and S-3. Further investigations areneeded to identify the cause of these peaks as they do not directly correspond to vortexshedding from any model component at St = 0.2 or any of the peaks observed in theunsteady surface pressure spectra.

For M∞ = 0.04 to 0.10, the expected equivalent Reynolds number range for the dragstrut and main strut is 5k < Ren < 30k. This flow state is characterized by transitionin the shear layer [48]. Thus, the expected equivalent Strouhal number range of thetandem cylinder system at microphone locations A, B, and C is 0.17 < Stn < 0.19. Formicrophone location D, the spacing ratio of Sn/D = 4.54 yields an expected value of0.12 < Stn < 0.15. Vortex shedding behind an isolated main strut is expected to occurat StD = 0.2. No peaks corresponding to the expected vortex shedding are observed inthe far-field acoustic spectra.

37

Chapter 5. Results & Analysis 38

F requency , Hz

SP

L,

dB

re

20

µP

a

100

101

102

103

104

30

40

50

60

70

80

90

100

110S -1- M

∞ = 0.04

S -2- M∞ = 0.04

S -3- M∞ = 0.04

B kgd - M∞ = 0.04

S -1- M∞ = 0.07

S -2- M∞ = 0.07

S -3- M∞ = 0.07

B kgd- M∞ = 0.07

S -1- M∞ = 0.10

S -2- M∞ = 0.10

S -3- M∞ = 0.10

B kgd- M∞ = 0.10

(a)

F requency , Hz

SP

L,

dB

re

20

µP

a

500 1000 200030

40

50

60

70S -1- M

∞ = 0.04

S -2- M∞ = 0.04

S -3- M∞ = 0.04

B kgd - M∞ = 0.04

S -1- M∞ = 0.07

S -2- M∞ = 0.07

S -3- M∞ = 0.07

B kgd- M∞ = 0.07

S -1- M∞ = 0.10

S -2- M∞ = 0.10

S -3- M∞ = 0.10

B kgd- M∞ = 0.10

(b)

Figure 5.1: Far-field SPL spectra of configurations S-1, S-2, and S-3. for frequency rangesof (a) 1Hz-20 kHz and (b) 100Hz-2 kHz


5.1.2 Motor and Propeller Induced Noise

The far-field acoustic SPL spectra for the M and P configurations are shown in Figures5.2 and 5.3. The spectra typically show behaviour identical to the background noiseof the tunnel, with tones of the motor fundamental frequency, fm, and propeller bladepassing frequency, BPF. Considering when the motor is spinning and the propeller isabsent (M configurations), the dominant peak of all far-field spectra occurs at the secondharmonic of the motor fundamental frequency, fm. The peak expected at fm only exceedsthe background levels for M∞ = 0.10 and can be seen as a small peak at 122Hz inFigure 5.2b. For all values of M∞, the strength of the harmonics of fm are greatestin configuration M-2, when the main strut is cantilevered from the nacelle, without thesupport of the drag strut. This may be indicative of structural vibrations of the modelsupport at the motor fundamental frequency. Upon inclusion of the drag strut (M-3),the main strut is more rigidly supported, and modes 3-7 of fm are attenuated as shownin Figure 5.2b. Reductions of 2-3 dB SPL are evident from the fourth harmonic onwarduntil the broadband background noise overcomes the peaks in the spectrum.

When the propeller is added (configurations P-1, P-2, and P-3), the dominant peakis the BPF for all test cases. For M∞ = 0.04 and M∞ = 0.07, the propeller is 3-bladedand thus the BPF = 3fm, whereas for M∞ = 0.10, the propeller has only two bladesso the BPF = 2fm. Contrary to the spectra of the M configurations (Figure 5.2b), thestrengths of the upper fm harmonics are not reduced when the drag strut supports themain strut. Natural harmonics of the BPF are stronger than their neighbouring BPFsubharmonics. For M∞ = 0.10, the two strongest tones of the far-field acoustic spectraoccur at BPF and 2BPF, which is in contrast to the lower speed cases where the strongesttone is found at the BPF but the next largest peak is a subharmonic (2/3BPF = 2fm).For the complete landing gear configuration, the isolated motor noise at the BPF is atleast 15 dB less than when the propeller is present. The inclusion of a rotating propellerincreases the levels of the expected tonal peak at fm, which was not prominent in the Mconfigurations. A broadband increase in levels of approximately 3 dB for all frequenciesgreater than the BPF is also observable.

5.2 Unsteady Surface Pressures

For each case outlined in Table 4.1 that included the main strut and wheels (S-2, S-3, M-2, M-3, P-2, and P-3), unsteady surface pressure measurements were acquiredsimultaneously with the far-field acoustic data as described in Section 4.3.2.


F requency , Hz

SP

L,

dB

re

20

µP

a

100

101

102

103

104

30

40

50

60

70

80

90

100

110M-1- M

∞ = 0.04

M-2- M∞

= 0.04

M-3- M∞

= 0.04

B kgd- M∞

= 0.04

M-1- M∞

= 0.07

M-2- M∞

= 0.07

M-3- M∞

= 0.07

B kgd- M∞

= 0.07

M-1- M∞

= 0.10

M-2- M∞

= 0.10

M-3- M∞

= 0.10

B kgd- M∞

= 0.10

(a)

F requency , Hz

SP

L,

dB

re

20

µP

a

500 1000 200030

40

50

60

70M-1- M

∞ = 0.04

M-2- M∞

= 0.04

M-3- M∞

= 0.04

B kgd- M∞

= 0.04

M-1- M∞

= 0.07

M-2- M∞

= 0.07

M-3- M∞

= 0.07

B kgd- M∞

= 0.07

M-1- M∞

= 0.10

M-2- M∞

= 0.10

M-3- M∞

= 0.10

B kgd- M∞

= 0.10

(b)

Figure 5.2: Far-field SPL spectra of configurations M-1, M-2, and M-3 for frequencyranges of (a) 1Hz-20 kHz and (b) 100Hz-2 kHz.


F requency , Hz

SP

L,

dB

re

20

µP

a

100

101

102

103

104

30

40

50

60

70

80

90

100

110P -1 - M

∞ = 0.04

P -2 - M∞ = 0.04

P -3 - M∞ = 0.04

B kgd - M∞ = 0.04

P -2 - M∞ = 0.07

P -2 - M∞ = 0.07

P -3 - M∞ = 0.07

B kgd - M∞ = 0.07

P -1 - M∞ = 0.10

P -2 - M∞ = 0.10

P -3 - M∞ = 0.10

B kgd - M∞ = 0.10

(a)

F requency , Hz

SP

L,

dB

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20

µP

a

500 1000 200030

40

50

60

70

80

90

P -1 - M∞ = 0.04

P -2 - M∞ = 0.04

P -3 - M∞ = 0.04

B kgd - M∞ = 0.04

P -2 - M∞ = 0.07

P -2 - M∞ = 0.07

P -3 - M∞ = 0.07

B kgd - M∞ = 0.07

P -1 - M∞ = 0.10

P -2 - M∞ = 0.10

P -3 - M∞ = 0.10

B kgd - M∞ = 0.10

(b)

Figure 5.3: Far-field SPL spectra of configurations P-1, P-2, and P-3 for frequency rangesof (a) 1Hz-20 kHz and (b) 100Hz-2 kHz.


a = 0.5 (M∞

= 0.04)

a = 0.2 (M∞

= 0.07)

a = 0.1 (M∞

= 0.10)

0.2

0.4

0.6

0.8

1.0

1.2

x/Rp

Rs/R

p

0 0.5 1.0 1.5 2.0 2.5

A

B

D

C

Figure 5.4: Slipstream contraction for all test cases with strut microphone locationsindicated. Computed using (2.8) for inflow ratios based on Figures 4.3 and 4.4.

5.2.1 Strut Surface Microphones

Figures 5.5, 5.6, and 5.7 show the surface SPL spectra for each strut microphone at eachwind tunnel flow speed. Each subfigure investigates the case with and without the dragstrut for a given propeller mode.

The expected radius of the slipstream relative to the main strut microphones basedon (2.8) are shown in Figure 5.4. For all test conditions, microphones A, B, and C areexpected to be within the propeller slipstream. However, microphone D approaches thelimit of the slipstream for M∞ = 0.04. Thus, the strut is expected to experience boththe influence of propeller tip vortex when and strong axial shear at that location whenthe flow is unimpeded.

Isolated Main Strut

For the isolated main strut, unsteady surface pressure spectra at microphone locations A- D are presented as the dashed lines on the left side of in Figures 5.5, 5.6, and 5.7. TheStrouhal number presented in these cases is based on the nominal freestream velocity,U∞, and the main strut diameter. For all tests omitting the propeller (i.e. S-2 and M-2),


a broad hump spans the lower frequencies and is typically capped by the vortex sheddingfrequency of the main strut (St ≈ 0.2). For M∞ = 0.04 and 0.07, at frequencies greaterthan 1 kHz, the pressure signal quality deteriorates as electrical noise dominates. Thesteep roll-off above 10 kHz is due to the antialiasing filter employed.

The spinning motor (Configuration M-2) increases the tonal content of the spectrawith a strong peak at the motor fundamental frequency, fm, reaching SPL values be-tween 95-105 dBSPL. Harmonics of fm are also evident with the tonal spectral levelsmonotonically decreasing. Upon the introduction of the motor, the peak associated withthe vortex shedding frequency increases by 10-15 dBSPL for M∞ = 0.04 and 0.07. AtM∞ = 0.10, peak levels at St = 0.2 remain unchanged. This may be due to broadbandnoise produced by the motor or possibly vibrations emanating throughout the model andsupport structure. Further investigation is necessary to identify the cause of this levelincrease. Microphone A is most heavily influenced by the motor, exhibiting a broadeningof the peak at the cylinder vortex shedding frequency, which is indicative of increasedturbulence levels and less coherent vortex shedding.

Adding the propeller to the isolated main strut (configuration P-2) causes the BPFharmonics to rise above the neighbouring tones (harmonics of fm). The BPF tones aremost prominent on microphones C and D, which are closest to the plane of the propellertip. More than 20 harmonics of fm are observed on microphones C and D for M∞ = 0.07

in this case.

Tandem Configuration

The unsteady surface pressure spectra at microphone locations A - D upon inclusionof the drag strut (configurations S-3, M-3 and P-3) are shown on the right hand side ofFigures 5.5, 5.6, and 5.7. In these cases, the indicated Strouhal number is calculated basedon the normal component of the velocity, Un = U∞ cos Λ, and the drag strut diameter.The turbulent wake of the drag strut is manifested as a broadband SPL increase over theentire measured bandwidth of each of the surface unsteady pressure sensors. An increaseby 20-30 dB relative to the case of the isolated main strut is observed and is of a generallyconsistent shape between configurations for a given flow speed.

Peaks observed at the vortex shedding frequency of the drag strut (St ≈ 0.2) arebroader, and for configurations S-3 and M-3 are only observed at microphone locations Cand D. This is consistent with the work of Wilkins [45], which shows a weakening of thespectral peaks associated with vortex shedding of a tandem yawed cylinder arrangementat high spacing ratios (S/D > 4.5) for Re = 56k. He suggests that as the spacing ratio isincreased, the shear layer of the upstream cylinder no longer impinges on the downstream


cylinder at which point coherent vortex shedding ceases. The spacing ratio where thisbehaviour occurs in the current study is 5.72 < Sn/D < 6.89 for 6.4k< Ren <19.6k.

In addition to the primary vortex shedding peak, the inclusion of the drag strutyields a secondary peak at twice the shedding frequency, St ≈ 0.4, in configurationsS-3 and M-3. This peak has been widely observed at subcritical Reynolds numbers formeasurements not far off the wake centreline [36] and corresponds to the impingement ofboth shear layers of the upstream cylinder on the main strut. These secondary peaks areobserved on all surface pressure spectra which exhibit the peak at the shedding frequencyfor configurations S-3 and M-3. This behaviour is primarily observed on microphones Cand D.

Typically, the unsteady surface pressure spectra of microphones A - D do not exhibitdramatic changes between cases with and without the propeller (i.e. M-3 and P-3). ForM∞ = 0.04 and 0.07, the influence of the propeller is observed on the surface unsteadypressure spectra of all microphones as shown on the right side of Figures 5.5c and 5.6c.In these cases, the peak associated with vortex shedding appears to shift along the spanof the main strut. At microphone locations A and B, a peak is observed near Stn = 0.25.This peak reduces in strength and shifts toward Stn = 0.2 at microphone locations A andB. This may be an indication of categorically different vortex shedding regimes along thedrag strut. Further investigation should be undertaken to examine the flow structuresassociated with the spectral peaks and their spanwise distributions.

The BPF tone and its first two harmonics peak over the background levels in con-figuration P-3 and are easily distinguishable from the peaks related to vortex shedding.The peaks at the BPF are comparable in magnitude to the vortex shedding peak of thetandem configuration. However, in comparison to the BPF tone in the isolated strut case(P-2), the magnitude is decreased by approximately 5 dB. This may be due to a shieldingeffect of the drag strut, breaking the structures associated with the blade vortex sheetprior to them impinging on the main strut.

5.2.2 Wheel Surface Microphones

Figures 5.8, 5.9, and 5.10 show the unsteady wheel surface pressures at locations L1, L2,R1, and R2 for all test configurations. Since the wheels are not in the slipstream of thepropeller, or the wake of the drag strut, the changes in unsteady pressures from caseto case are minimal. Each wheel surface microphone has levels that are nearly identicalto those of the sensor located symmetrically across the x − z plane (e.g. L1 and R1).Compared to the levels of unsteadiness experienced by the main strut, the wheels do not


F requenc y, Hz

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B

C

D

S -2

F requenc y, Hz

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10-1

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B

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S -3

(a) Static

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B

C

D

M-2 fm

F requenc y, Hz

S tn

SP

L,

dB

re

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a

100

101

102

103

104

10-2

10-1

100

101

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A

B

C

D

M-3 fm

(b) Motor

F requenc y, Hz

S t

SP

L,

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a

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10-3

10-2

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C

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P -2 B P F

F requenc y, Hz

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10-2

10-1

100

101

40

50

60

70

80

90

100

110

120

130

140

A

B

C

D

P -3B P F

(c) Propeller

Figure 5.5: Strut unsteady surface pressure spectral levels, M∞ = 0.04


F requenc y, Hz

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S -2

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A

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S -3

(a) Static

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µP

a

100

101

102

103

104

10-3

10-2

10-1

100

101

40

50

60

70

80

90

100

110

120

130

140

A

B

C

D

M-2f

m

F requenc y, Hz

S tn

SP

L,

dB

re

20

µP

a

100

101

102

103

104

10-3

10-2

10-1

100

101

40

50

60

70

80

90

100

110

120

130

140

A

B

C

D

M-3f

m

(b) Motor

F requenc y, Hz

S t

SP

L,

dB

re

20

µP

a

100

101

102

103

104

10-3

10-2

10-1

100

101

40

50

60

70

80

90

100

110

120

130

140

A

B

C

D

P -2B P F

F requenc y, Hz

S tn

SP

L,

dB

re

20

µP

a

100

101

102

103

104

10-3

10-2

10-1

100

40

50

60

70

80

90

100

110

120

130

140

A

B

C

D

P -3

B P F

(c) Propeller



F requenc y, Hz

S t

SP

L,

dB

re

20

µP

a

100

101

102

103

104

10-3

10-2

10-1

100

40

50

60

70

80

90

100

110

120

130

140

A

B

C

D

S -2

F requenc y, Hz

S tn

SP

L,

dB

re

20

µP

a

100

101

102

103

104

10-3

10-2

10-1

100

40

50

60

70

80

90

100

110

120

130

140

A

B

C

DS -3

(a) Static

F requenc y, Hz

S t

SP

L,

dB

re

20

µP

a

100

101

102

103

104

10-3

10-2

10-1

100

40

50

60

70

80

90

100

110

120

130

140

A

B

C

D

M-2f

m

F requenc y, Hz

S tn

SP

L,

dB

re

20

µP

a

100

101

102

103

104

10-3

10-2

10-1

100

40

50

60

70

80

90

100

110

120

130

140

A

B

C

DM-3f

m

(b) Motor

F requenc y, Hz

S t

SP

L,

dB

re

20

µP

a

100

101

102

103

104

10-3

10-2

10-1

100

40

50

60

70

80

90

100

110

120

130

140

A

B

C

D

P -2B P F

F requenc y, Hz

S tn

SP

L,

dB

re

20

µP

a

100

101

102

103

104

10-3

10-2

10-1

100

40

50

60

70

80

90

100

110

120

130

140

A

B

C

DP -3B P F

(c) Propeller



appear to play an important role in the flow interactions with the propeller or drag strut.Reorienting the wheels such that the surface microphones are in a separated region ofthe flow may give indications of wake interactions. Future studies involving landing gearin the unsteady wake of a propeller should concentrate efforts on the interaction of theslipstream and strut(s).


F requenc y, Hz

SP

L,

dB

re

20

µP

a

100

101

102

103

104

40

60

80

100

120

140

L 1

L 2

R 1

R 2

(a) Static

F requenc y, Hz

B P F

SP

L,

dB

re

20

µP

a

100

101

102

103

104

10-2

10-1

100

101

40

60

80

100

120

140

L 1

L 2

R 1

R 2

fm

(b) Motor

F requenc y, Hz

B P F

SP

L,

dB

re

20

µP

a

100

101

102

103

104

10-2

10-1

100

101

40

60

80

100

120

140

L 1

L 2

R 1

R 2

B P F

(c) Propeller

Figure 5.8: Wheel unsteady surface pressure spectral levels, M∞ = 0.04


F requenc y, Hz

SP

L,

dB

re

20

µP

a

100

101

102

103

104

40

60

80

100

120

140

L 1

L 2

R 1

R 2

(a) Static

F requenc y, Hz

B P F

SP

L,

dB

re

20

µP

a

100

101

102

103

104

10-2

10-1

100

101

40

60

80

100

120

140

L 1

L 2

R 1

R 2

fm

(b) Motor

F requenc y, Hz

B P F

SP

L,

dB

re

20

µP

a

100

101

102

103

104

10-2

10-1

100

101

40

60

80

100

120

140

L 1

L 2

R 1

R 2

B P F

(c) Propeller



F requenc y, Hz

SP

L,

dB

re

20

µP

a

100

101

102

103

10440

60

80

100

120

140

L 1

L 2

R 1

R 2

(a) Static

F requenc y, Hz

B P F

SP

L,

dB

re

20

µP

a

100

101

102

103

104

10-2

10-1

100

101

40

60

80

100

120

140

L 1

L 2

R 1

R 2

fm

(b) Motor

F requenc y, Hz

B P F

SP

L,

dB

re

20

µP

a

100

101

102

103

104

10-2

10-1

100

101

40

60

80

100

120

140

L 1

L 2

R 1

R 2

B P F

(c) Propeller


6 | Conclusions

6.1 Wind Tunnel Characterization

Prior to conducting the aeroacoustic testing described in the current work, it was nec-essary to characterize the aerodynamic and acoustic properties of the Acoustic WindTunnel Facility at UTIAS due to many recent modifications. Irregularities in the meanflow profile were observed to be sufficiently reduced by x/DN = 0.3 for test velocitiesof 33 and 55 m/s. The maximum attainable flow velocity in the current configurationwas found to be approximately 60 m/s. Turbulence intensity in the potential core wasmeasured at less than 1% for x/DN ≤ 1.4. As pertaining to the tunnel acoustics, far-fieldnoise spectra were presented showing levels comparable to similarly sized wind tunnels.This study provides a basis for future wind tunnel testing in the facility and an indicationof the aerodynamic and acoustic limits that are imposed for such testing.

6.2 Model Development

The development of the model constituted a significant portion of this project. The use ofmodel-scale parts required thrust test data based on slipstream measurements in order todetermine axial inflow factors. The model set-up allowed for tests including combinationsof the nacelle, main strut, wheels, drag strut, and propeller. Configurations including themain strut and wheels were equipped with unsteady surface pressure sensors to measurethe interactions with the drag strut and propeller. Descriptions of the sensor calibrationprocedure and the data acquisition instrumentation are given followed by a discussion ofthe uncertainties in the presented spectral estimates.

52

Chapter 6. Conclusions 53

6.3 Landing Gear Testing

The first ever far-field noise and surface pressure measurements taken for a two-strut,two-wheeled landing gear with an actuated upstream propeller have been compared anddiscussed for various configurations at Mach numbers of 0.04, 0.07, and 0.10. The pro-peller and model support were shown to dominate the far-field acoustic signature formost test cases.

Unsteady surface pressure measurements on the landing gear main strut and wheelswere also recorded. For an isolated main strut, the pressure fluctuations on the main strutwere observed to exhibit the strongest fluctuations at the blade passing frequency andits harmonics. When the drag strut was added, these pressure fluctuations were maskedby a broadband increase in levels due to the wake of the upstream yawed cylinder. Twopeaks corresponding to vortex shedding from the drag strut were observed above thebroadband levels for the main strut surface microphones located nearest to the edge ofthe propeller slipstream.

The recorded sound pressure level spectra can be used as an indication of what toexpect for full-scale testing thereby enabling the development of more accurate noiseprediction tools. The results of this study also provide a new benchmark case for com-putational aeroacoustic codes that are used in design and analysis of aircraft landinggear.

6.4 Recommendations for Future Work

The current work represents the first attempt at investigating noise from landing gearin a propeller slipstream. In order to continue the analysis of these configurations andimprove the quality of the data, some suggestions for future work are relayed below.

• The structure of the motor support and the landing gear model should be decou-pled for further testing. Efforts were made to reduce to the vibrations propagatingthrough the structure for the current work. However, oscillations at motor fre-quency, fm, were evident in all surface pressure measurements. Though the modeldid not appear to move by visual inspection, vibrations could be felt by placing ahand on the wheel while the motor was spinning.

• In order to more adequately represent the full-scale configuration, acquisition ofan accurately scaled down propeller should be considered. An investigation usinghigher fidelity components as such would give valuable insight to the drawbacks

Chapter 6. Conclusions 54

of using remote control model parts. Operating at lower Reynolds numbers thanactual flight conditions may invoke difficulties with respect to matching both aero-dynamic and acoustic parameters for propellers and may constitute an additionalexperimental program.

• Future studies should concentrate efforts on strut interactions as the influence of thedrag strut appears to dominate the surface pressure spectra across the span of themain strut. A study examining the observed spanwise variation in vortex sheddingis of particular interest; It is advised that hot-wire or PIV investigations of thetandem yawed cylinder configuration be performed to identify the flow structuressurrounding the cylinders and quantify the local flow conditions. These localizedvariations of the flow field may have a significant effect on prediction models asshown by Guo [17].

• Techniques such as those of Siddon [39] and Pan [31] could be used to characterizethe acoustics of the fluid-surface interactions. These cross-spectral analysis methodsof the far-field and surface pressure signals may identify the regions of greatestinfluence to the far-field acoustics, allowing for more focussed efforts to reducenoise emissions.

References

[1] N. G. Ball. Acoustic Excitation of Tollmien-Schlichting Waves in a Laminar Bound-ary Layer. Ph.D. dissertation, University of Toronto, 1987.

[2] P. W. Bearman. On vortex shedding from a circular cylinder in the critical Reynoldsnumber régime. Journal of Fluid Mechanics, 37(3):577–585, 1969.

[3] J. S. Bendat and A. G. Piersol. Engineering Applications of Correlation and SpectralAnalysis. John Wiley and Sons, New York, 2nd edition, 1993.

[4] J. L. Beveridge. The Radial Distribution of Propeller Thrust from Model Wake Mea-surements. Technical report, Department of the Navy, Hydromechanics Laboratory,1958.

[5] Bombardier Aerospace. Q400 Customer Images. http://media.bombardiercms.

com/q400/medias/gallery/q400customerImages/high/06.jpg.

[6] Bombardier Aerospace. Smart Cockpit: Bombardier Dash-8-400 Landing Gear.http://www.smartcockpit.com/docs/Q400-Landing\_Gear\_1.pdf.

[7] N. Curle. The influence of solid boundaries upon aerodynamic sound. Proceed-ings of the Royal Society A: Mathematical, Physical and Engineering Sciences,231(1187):505–514, Sept. 1955.

[8] W. M. Dobrzynski. Almost 40 years of airframe noise research: What did we achieve?Journal of Aircraft, 47(2):353–367, Mar. 2010.

[9] W. M. Dobrzynski, M. Pott-Pollenske, D. Foot, and M. Goodwin. Landing gearsaerodynamic interaction noise. In European Congress on Computational Methods inApplied Sciences and Engineering 2004, pages 115–135, Jyväskylä, Finland, 2004.Multi-Science.

55

References 56

[10] W. M. D. Dobrzynski and H. Buchholz. Full-scale noise testing on Airbus land-ing gears in the German Dutch Wind Tunnel. In 3rd AIAA/CEAS AeroacousticsConference, Atlanta, Georgia, 1997. AIAA Paper 1997–1597, American Institute ofAeronautics and Astronautics.

[11] Environment Canada. Toronto, ON - 7 Day Forecast - Environment Canada. http://www.smartcockpit.com/docs/Q400-Landing\_Gear\_1.pdf.

[12] Federal Aviation Administration. Part 36: Noise Standards: Aircraft Type andAirworthiness Certification. In Code of Federal Regulations, Title 14: Aeronauticsand Space. U.S. Government Printing Office, Washington, D.C., 2011.

[13] J. E. Ffowcs Williams and D. L. Hawkings. Sound generation by turbulence andsurfaces in arbitrary motion. Philosophical Transactions of the Royal Society ofLondon. Series A, Mathematical and Physical Sciences, 264(1151):321–342, May1969.

[14] J. S. Gibson. Non-engine Aerodynamic Noise Investigation of a Large Aircraft.Technical Report October, NASA Langley, Hampton, Virginia, 1974.

[15] M. E. Goldstein. Aeroacoustics. McGraw-Hill International Book Company, NewYork, 1976.

[16] Y. Guo. A statistical model for landing gear noise prediction. Journal of Sound andVibration, 282(1-2):61–87, Apr. 2005.

[17] Y. Guo. Effects of local flow variations on landing gear noise prediction and analysis.Journal of Aircraft, 47(2):383–391, Mar. 2010.

[18] Y. Guo, R. W. Stoker, and K. J. Yamamoto. Experimental study on aircraft landinggear noise. Journal of Aircraft, 43(2):306–317, Mar. 2006.

[19] Y. P. Guo, K. J. Yamamoto, and R. W. Stoker. An empirical model for landing gearnoise prediction. In 10th AIAA/CEAS Aeroacoustics Conference, Manchester, U.K.,2004. AIAA Paper 2004–2888, American Institute of Aeronautics and Astronautics.

[20] A. Hanson. Vortex shedding from yawed cylinders. AIAA Journal, 4(4):738–740,1966.

[21] G. J. Healy. Measurement and Analysis of Aircraft Far-Field Aerodynamic Noise.Technical Report December, NASA Langley, Hampton, Virginia, 1974.

References 57

[22] H. H. D. Heller and W. M. D. Dobrzynski. Sound radiation from aircraft wheel-well/landing-gear configurations. Journal of Aircraft, 14(8):768–774, 1977.

[23] F. V. Hutcheson and T. F. Brooks. Noise radiation from single and multiple rodconfigurations. In 12th AIAA/CEAS Aeroacoustics Conference, number May, Cam-bridge, Massachusetts, 2006. AIAA Paper 2006–2629, American Institute of Aero-nautics and Astronautics.

[24] L. N. Jenkins, D. H. Neuhart, C. B. McGinley, M. M. Choudhari, and M. R. Khor-rami. Measurements of unsteady wake interference between tandem cylinders. In36th AIAA Fluid Dynamics Conference, Monterey, California, 2006. AIAA Paper2006–3202, American Institute of Aeronautics and Astronautics.

[25] B. S. Lazos. Mean flow features around the inline wheels of four-wheel landing gear.AIAA Journal, 40(2):193–198, Feb. 2002.

[26] M. J. Lighthill. On sound generated aerodynamically. I. General theory. Pro-ceedings of the Royal Society A: Mathematical, Physical and Engineering Sciences,211(1107):564–587, Mar. 1952.

[27] E. Manoha, J. Bulté, and B. Caruelle. LAGOON: an experimental database forthe validation of CFD/CAA methods for landing gear noise prediction. In 14thAIAA/CEAS Aeroacoustics Conference, number May, Vancouver, British Columbia,2008. AIAA Paper 2008–2816, American Institute of Aeronautics and Astronautics.

[28] J. Mathew. Design, Fabrication, and Characterization of an Anechoic Wind TunnelFacility. Ph.D. dissertation, University of Florida, 2006.

[29] U. Michel and E. Froebel. Lower limit for the velocity fluctuation level in windtunnels. Experiments in Fluids, 6(1):49–54, 1988.

[30] A. Okajima. Flows around two tandem circular cylinders at very high Reynoldsnumbers. Bulletin of the Japan Society of Mechanical Engineers, 22(166):504–511,1979.

[31] Y. S. Pan. Cross-correlation methods for studying near- and farfield noise character-istics of flow-surface interactions. The Journal of the Acoustical Society of America,58:586–594, 1975.

[32] A. R. Quayle, A. P. Dowling, H. Babinsky, W. R. Graham, and Y. Liu. Phasedarray measurements from landing gear models. In 13th AIAA/CEAS Aeroacoustics

References 58

Conference, Rome, Italy, May 2007. AIAA Paper 2007–3463, American Institute ofAeronautics and Astronautics.

[33] P. A. Ravetta, R. A. Burdisso, and W. F. Ng. Wind tunnel aeroacoustic measure-ments of a 26%- scale 777 main landing gear model. In 10th AIAA/CEAS Aeroa-coustics Conference, Manchester, U.K., 2004. AIAA Paper 2004–2885, AmericanInstitute of Aeronautics and Astronautics.

[34] R. Reger, F. Liu, and L. Cattafesta. An experimental study of the rudimentary land-ing gear. In 51st AIAA Aerospace Sciences Meeting, number January, Grapevine,Texas, 2013. AIAA Paper 2013–0464, American Institute of Aeronautics and Astro-nautics.

[35] A. K. Ringshia, P. A. Ravetta, W. F. Ng, and R. A. Burdisso. Aerodynamic mea-surements of the 777 main landing gear model. In 12th AIAA/CEAS AeroacousticsConference, number May, Cambridge, Massachusetts, 2006. AIAA Paper 2006–2625,American Institute of Aeronautics and Astronautics.

[36] A. Roshko. Experiments on the flow past a circular cylinder at very high Reynoldsnumber. J. Fluid Mech, pages 345–356, 1961.

[37] W. R. Sears. The boundary layer of yawed cylinders. Journal of the AeronauticalSciences, 15(1):49–52, 1948.

[38] M. Shirakashi, S. Ueno, Y. Ishida, and S. Wakiya. Vortex excited oscilation of acircular cylinder in a uniform ulow. Bulletin of the Japan Society of MechanicalEngineers, 27(228):1120–1126, 1984.

[39] T. E. Siddon. Surface dipole strength by cross-correlation method. The Journal ofthe Acoustical Society of America, 53:619–633, 1973.

[40] S. R. Snarski. Relation Between the Fluctuating Wall Pressure and the TurbulentStructure Of a Boundary Layer on a Cylinder In Axial Flow. Technical ReportAugust, Naval Undesea Warfare Center, New London, Connecticut, 1993.

[41] P. T. Soderman and C. S. Allen. Microphone Measurements In and Out of Airstream.In T. J. Mueller, editor, Aeroacoustic Measurements, pages 1–61. Springer-Verlag,Berlin, Germany, 2002.

[42] P. R. Spalart, M. L. Shur, M. K. Strelets, and A. K. Travin. Towards noise predic-tion for rudimentary landing gear. In Procedia IUTAM, volume 1, pages 283–292,Southampton, U. K., 2010.

References 59

[43] W. J. G. Trebble, J. Williams, and R. P. Donnelly. Propeller noise at model- andfull-scale. Journal of Aircraft, 20(1):34–41, 1983.

[44] L. L. M. Veldhuis. Propeller Wing Aerodynamic Interference. Ph.D. dissertation,Technische Universiteit Delft, 2005.

[45] S. J. Wilkins. Vortex-Shedding in Yawed-Tandem Circular Cylinder Arrangements.Ph.D. dissertatioin, University of New Brunswick, 2011.

[46] S. J. Wilkins and J. W. Hall. Experimental investigation of a tandem cylindersystem with a yawed upstream cylinder. Journal of Pressure Vessel Technology,136:011302–1–8, Feb. 2014.

[47] N. S. Zawodny. Aeroacoustic Characterization of Scaled Canonical Nose LandingGear Configurations. Ph.D. dissertation, University of Florida, 2013.

[48] M. Zdravkovich. Flow Around Circular Cylinders, Volume 1: Fundamentals. FlowAround Circular Cylinders: A Comprehensive Guide Through Flow Phenomena, Ex-periments, Applications, Mathematical Models, and Computer Simulations. OxfordUniversity Press, New York, New York, 1997.

[49] M. M. Zdravkovich. Flow Around Circular Cylinders, Volume 2: Applications. FlowAround Circular Cylinders: A Comprehensive Guide Through Flow Phenomena, Ex-periments, Applications, Mathematical Models, and Computer Simulations. OxfordUniversiy Press, New York, New York, 2003.

[50] W. Zhao. Microphone Calibration. Technical report, University of Toronto, FlowControl and Experimental Turbulence Laboratory, Toronto, Ontario, 2013.

A | Wind Tunnel Layout

60

Appendix A. Wind Tunnel Layout 61

0.4

82

.44

0.6

10

.91

0.2

82

.78

0.8

64

.57

0.9

11

.22

2.4

41

.40

2.2

51

.83

2.87

2.59 x 2.59

6.0

5

6.10

0.91

0.70

0.6

2 1.02

1.52

3.96

1.33

3.05

Co

ntr

ac

tio

n

Up

stre

am

Sil

en

cer

Do

wn

stre

am

Sil

en

cer

Fan

an

d M

oto

r

Ple

nu

m

An

ech

oic

Ch

am

be

r No

zzle

Di!

use

r

Jet

Co

lle

cto

r

FigureA.1:Schematic

oftheAcousticW

indTu

nnel

Facility.

Alldimension

sin

meters.

B | Surface Microphone Calibration

62

Appendix B. Surface Microphone Calibration 63

Fre

qu

en

cy

, k

Hz

Sensitivity, mV/Pa

12

34

57.5

7.6

7.7

7.8

7.98

8.1

8.2

8.3

8.4

8.5

Mic

rop

ho

ne

R1

Fre

qu

en

cy

, k

Hz

Sensitivity, mV/Pa

12

34

57.5

7.6

7.7

7.8

7.98

8.1

8.2

8.3

8.4

8.5

Mic

rop

ho

ne

R2

Fre

qu

en

cy

, k

Hz

Sensitivity, mV/Pa

12

34

5

7.6

7.88

8.2

8.4

Mic

rop

ho

ne

L2

Fre

qu

en

cy

, k

Hz

Sensitivity, mV/Pa

12

34

58.5

8.6

8.7

8.8

8.99

9.1

9.2

9.3

9.4

9.5

Mic

rop

ho

ne

L1

Fre

qu

en

cy

, k

Hz

Sensitivity, mV/Pa

12

34

58

8.1

8.2

8.3

8.4

8.5

8.6

8.7

8.8

8.99

Mic

rop

ho

ne

D

Fre

qu

en

cy

, k

Hz

Sensitivity, mV/Pa

12

34

57.5

7.6

7.7

7.8

7.98

8.1

8.2

8.3

8.4

8.5

Mic

rop

ho

ne

A

Fre

qu

en

cy

, k

Hz

Sensitivity, mV/Pa

12

34

58.5

8.6

8.7

8.8

8.99

9.1

9.2

9.3

9.4

9.5

Mic

rop

ho

ne

B

Fre

qu

en

cy

, k

Hz

Sensitivity, mV/Pa

12

34

57.5

7.6

7.7

7.8

7.98

8.1

8.2

8.3

8.4

8.5

Mic

rop

ho

ne

C

FigureB.1:Calibration

curveforallsurface

microph

onesensors.

Solid

lines

representthemeanan

dda

shed

lines

representtw

ostan

dard

deviations

from

themeanbe

tween20

0Hzan

d1kH

z,respectively.

C | Landing Gear Model Drawings

64

A A

75

3

2

1

SECTION A-A

4

6

ITEM NO. PART NUMBER / DWG. NUMBER DESCRIPTION QTY.1 UTIASAA-RLGM-010 MAIN STRUT ASSEMBLY 12 UTIASAA-RLGM-020 DRAG STRUT 13 UTIASAA-RLGM-030 WHEEL ASSEMBLY 24 B18.3.1M - 6 x 1.0 x 70 Hex SHCS -- 24NHX M6 x 1.0 x 70 SHCS 15 B18.3.1M - 4 x 0.7 x 20 Hex SHCS -- 20NHX M4 x 0.7 x 20 SHCS 16 B18.2.4.1M - Hex nut, Style 1, M6 x 1 --D-N M6 x 1HEX NUT 17 B18.2.4.1M - Hex nut, Style 1, M4 x 0.7 --D-N M4 x 0.7 HEX NUT 1

D

B

A

8 7 6 5 4 3 2

THE INFORMATION CONTAINED IN THIS DRAWING IS THE SOLE

PROPERTY OF THE UNIVESITY OF TORONTO. ANY REPRODUCTION IN

PART OR AS A WHOLE WITHOUT THE WRITTEN PERMISSION OF THE

UNIVERSITY OF TORONTO IS PROHIBITED.

PROPRIETARY AND CONFIDENTIAL

NEXT ASSY USED ON

UNLESS OTHERWISE SPECIFIED:DIMENSIONS ARE IN INCHES

TOLERANCES:

.XX .XXX FRACTIONS

.02 .005 1 1/16

REMOVE ALL BURRS AND SHARP EDGES.

DRAWN

CHECKED

TITLE:

SIZE

BDWG. NO. REV

WT: 987.84SCALE: 1:3

RUDIMENTARY DASH-8 MAIN LANDINGGEAR WIND TUNNEL MODEL ASSEMBLY

SHEET 1 OF 2

UTIASAA-RLGM-000

DO NOT SCALE DRAWING

1

12345678

CC

A

B

D

INSTITUTE FOR AEROSPACE STUDIESUNIVERSITY OF TORONTO4925 DUFFERIN STREETTORONTO, ONTARIOM3H 5T6

NAME

05/03/2013RC

DATE

THIS DRAWING WAS PRODUCED USING:

SOFTWARE: SOLIDWORKS VERSION: 2012

FILENAME: UTIASAA-RLGM-000

Appen

dix

C.

Landin

gG

ear

Model

Draw

ings

65

3.071

14.

565

12.

795

0.709

2.1

26

3.543 1.063

9.252

0.374

0.551 D

B

A

8 7 6 5 4 3 2






NEXT ASSY USED ON


TOLERANCES:

.XX .XXX FRACTIONS

.02 .005 1 1/16


DRAWN

CHECKED

TITLE:

SIZE

BDWG. NO. REV

WT: 987.84SCALE: 1:3

RUDIMENTARY DASH-8 MAIN LANDINGGEAR WIND TUNNEL MODEL ASSEMBLY

SHEET 2 OF 2

UTIASAA-RLGM-000


1

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CC

A

B

D


NAME

05/03/2013RC

DATE




Appen

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C.

Landin

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Model

Draw

ings

66

1.2

60

0.55

1

0.630

13.110

1.0

63

2.126

2

4

3

1

ITEM NO. PART NUMBER/DWG. NUMBER DESCRIPTION QTY.

1 UTIASAA-RLGM-011 MAIN STRUT 12 UTIASAA-RLGM-012 MAIN STRUT MICROPHONE INSERT 13 EK-26899-P03 SURFACE MICROPHONE (KNOWLES) 4

4 B18.6.7M - M2.5 x 0.45 x 4 Type I Cross Recessed FHMS --4N M2.5 X 0.45 X 4 FHMS 2

D

B

A

8 7 6 5 4 3 2






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TOLERANCES:

.XX .XXX FRACTIONS

.02 .005 1 1/16


DRAWN

CHECKED

TITLE:

SIZE

BDWG. NO. REV

WT: SCALE: 1:2

RUDIMENTARY LANDING GEAR WIND TUNNEL MODEL MAIN STRUT ASSEMBLY

SHEET 1 OF 1

UTIASAA-RLGM-010


1

12345678

CC

A

B

D


NAME

05/03/2013RC

DATE




Appen

dix

C.

Landin

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Model

Draw

ings

67

4.173

2 X M2.5x0.45 TAP

0.

500

0.

551

1.100

5.295

2.000

A

1/2-20 Machine Threads

12.795

2.000 3.118 4.528

1.100

1.0

63

2.126

1.282 1.717 1.260

0.118 TYP.

0.

630

0.138

0.315

R0.118 TYP.

0.3

15

DETAIL A SCALE 1 : 1

11.795 0.197

D

B

A

8 7 6 5 4 3 2






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TOLERANCES:

.XX .XXX FRACTIONS

.02 .005 1 1/16


DRAWN

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TITLE:

SIZE

BDWG. NO. REV

WT: SCALE: 1:2

RUDIMENTARY LANDING GEAR MAIN STRUT

SHEET 1 OF 2

UTIASAA-RLGM-011


1

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CC

A

B

D


NAME

05/03/2013RC

DATE




Appen

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Landin

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Model

Draw

ings

68

B B

1.149

1.282

1.717

2.500

M5 TAP X 2 0.315

D 1.292

R0.315

0.177 THRU

1.3

78 60°

1.0

63

0.315

DETAIL D SCALE 1 : 1

3.118 2.000

C

SECTION B-B

0.0

59 0.315

3.898 0.315

0.1

18 T

YP.

DETAIL C SCALE 1 : 1

D

B

A

8 7 6 5 4 3 2






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TOLERANCES:

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.02 .005 1 1/16


DRAWN

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TITLE:

SIZE

BDWG. NO. REV

WT: SCALE: 1:2

RUDIMENTARY LANDING GEAR MAIN STRUT

SHEET 2 OF 2

UTIASAA-RLGM-011


1

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CC

A

B

D


NAME

05/03/2013RC

DATE




Appen

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C.

Landin

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Model

Draw

ings

69

4.528

0.3

15

2 X M2.5 CSK 0.177 TYP.

A A

C

C

R0.276

0.128

0.0

59 T

YP.

0.315 TYP.

0.677

1.661

2.646

3.630

B

SECTION A-A

0.220 TYP.

0.0

89 T

YP.

DETAIL B SCALE 4 : 1

0.0

30 T

YP.

SECTION C-C

D

B

A

8 7 6 5 4 3 2






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.02 .005 1 1/16


DRAWN

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TITLE:

SIZE

BDWG. NO. REV

WT: SCALE: 2:1

RUDIMENTARY LANDING GEAR MAIN STRUT MICROPHONE INSERT

SHEET 1 OF 1

UTIASAA-RLGM-012


1

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CC

A

B

D


NAME

05/03/2013RC

DATE




Appen

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C.

Landin

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Model

Draw

ings

70

0.197 TYP.

2 X M4 THRU

0.2

95 T

YP.

0.591 TYP.

0.787 TYP.

9.646

0.

374

D

B

A

8 7 6 5 4 3 2






NEXT ASSY USED ON


TOLERANCES:

.XX .XXX FRACTIONS

.02 .005 1 1/16


DRAWN

CHECKED

TITLE:

SIZE

BDWG. NO. REV

WT: SCALE: 1:1

RUDIMENTARY LANDING GEAR DRAG STRUT

SHEET 1 OF 1

UTIASAA-RLGM-020


1

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CC

A

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NAME

05/03/2013RC

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Model

Draw

ings

71

3.543

1.181 R0.354 TYP.

1

32

4

5

ITEM NO. PART NUMBER / DWG. NUMBER DESCRIPTION QTY.1 UTIASAA-RLGM-031 RUDIMENTARY LANDING GEAR WHEEL TREAD 12 UTIASAA-RLGM-032 RUDIMENTARY LANDING GEAR OUTER WHEEL CAP 13 UTIASAA-RLGM-033 RUDIMENTARY LANDING GEAR INNER WHEEL CAP 14 B18.3.1M - 3 x 0.5 x 20 Hex SHCS -- 20NHX M3 X 0.5 X 20 SHCS 35 EK-26899-P03 SURFACE MICROPHONES (KNOWLES) 5

D

B

A

8 7 6 5 4 3 2






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TOLERANCES:

.XX .XXX FRACTIONS

.02 .005 1 1/16


DRAWN

CHECKED

TITLE:

SIZE

BDWG. NO. REV

WT: SCALE: 1:1

RUDIMENTARY LANDING GEAR WHEEL ASSEMBLY

SHEET 1 OF 1

UTIASAA-RLGM-030


1

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CC

A

B

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NAME

05/03/2013RC

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Appen

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C.

Landin

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Model

Draw

ings

72

3 X M3 THRU

120°

120°

0.307

0.220

2.362

3.543

45° TYP.

0.0

89

0.0

89

A

0.177 TYP.

0.0

89 T

YP.

0.030 TYP.

DETAIL A SCALE 2 : 1

D

B

A

8 7 6 5 4 3 2






NEXT ASSY USED ON


TOLERANCES:

.XX .XXX FRACTIONS

.02 .005 1 1/16


DRAWN

CHECKED

TITLE:

SIZE

BDWG. NO. REV

WT: SCALE: 1:1

RUDIMENTARY LANDING GEAR WHEEL TREAD

SHEET 1 OF 1

UTIASAA-RLGM-031


1

12345678

CC

A

B

D


NAME

05/03/2013RC

DATE




Appen

dix

C.

Landin

gG

ear

Model

Draw

ings

73

3.543

3 x M3x0.5x0.325 TAP

120°

120°

0.394

R0.354

0.276

0.

630

0.

260

D

B

A

8 7 6 5 4 3 2






NEXT ASSY USED ON


TOLERANCES:

.XX .XXX FRACTIONS

.02 .005 1 1/16


DRAWN

CHECKED

TITLE:

SIZE

BDWG. NO. REV

WT: SCALE: 1:1

RUDIMENTARY LANDING GEAR OUTER WHEEL CAP

SHEET 1 OF 1

UTIASAA-RLGM-032


1

12345678

CC

A

B

D


NAME

05/03/2013RC

DATE




Appen

dix

C.

Landin

gG

ear

Model

Draw

ings

74

3.543

R1.417

0.197 THRU

0.591

120°

120°

0.480

R0.354

0.

256

TYP.

0.

134

TYP.

0.197 TYP.

0.

630

0.

260

D

B

A

8 7 6 5 4 3 2






NEXT ASSY USED ON


TOLERANCES:

.XX .XXX FRACTIONS

.02 .005 1 1/16


DRAWN

CHECKED

TITLE:

SIZE

BDWG. NO. REV

WT: SCALE: 1:1

RUDIMENTARY LANDING GEAR INNER WHEEL CAP

SHEET 1 OF 1

UTIASAA-RLGM-033


1

12345678

CC

A

B

D


NAME

05/03/2013RC

DATE




Appen

dix

C.

Landin

gG

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Model

Draw

ings

75

chekiri rafik 201406 mas thesis

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