A STUDY OF THE REVERSE DELTA WING USING PARTICLE IMAGE VELOCIMETRY (PIV)
BY
AFAQ ALTAF
A thesis submitted in fulfilment of the requirement for the degree of Master of Science (Mechanical Engineering)
Kulliyyah of Engineering International Islamic University
Malaysia
JANUARY 2011
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ABSTRACT
Particle Image Velocimetry was used in a low speed wind tunnel to investigate the vortex structures of a slender reverse delta wing at various angles of attack and roll.
This work investigates the characteristics of the vortices generated downstream in planes perpendicular to the free stream direction and their dependence on angles of attack and roll at chord-based Reynolds number of Rec=3.82×105. The peak tangential velocities show a trend similar to a delta wing. A six component force balance was used to obtain the aerodynamic coefficients for a reverse delta wing and a delta wing.
A qualitative comparison with existing computational results is executed for velocity vectors and surface pressure contours; whereby the computational results of surface pressure contours agree with the six component force balance results, revealing that a regular delta wing exhibits a higher CL than a reverse delta wing. It can be concluded that a regular delta wing vortex at a particular angle of attack exhibits a tangential velocity magnitude twice that of a reverse delta wing; exhibits higher circulation and vorticity than a reverse delta wing vortex.
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ﺺﺨﻠﻣ ﺚﺤﺒﻟا
ﺔﻋﺮﺴﻟا سﺎﻴﻘﻟ ﻰﺌﻳﺰﺠﻟاﺮﻳﻮﺼﺘﻟا ﺔﻴﻨﻘﺗ ﺖﻣﺪﺨﺘﺳا )
PIV ( ﻰﺋاﻮﻬﻟا ﻖﻔﻨﻟا ﻰﻓ
ﻟا نود سﻮﻜﻌﻤﻟا ﺎﺘﻟﺪﻟا حﺎﻨﺟ لﻮﺣ ﺔﻴﺋاﻮﻬﻟا تﺎﻣاوﺪﻟا ﺔﻴﻠﻜﻴه ﺔﺳارﺪﻟ ﻰﺗﻮﺼ
ﺔﻔﻠﺘﺨﻣ نراود و مﻮﺠه ﺎﻳاوز ﺪﻨﻋ .
ﺺﺋﺎﺼﺧ رﺎﺒﺘﺧا ﻢﺗ ﺔﺳارﺪﻟا ﻩﺬه ﻰﻓ و
ﻰﻠﻋ ىدﻮﻤﻌﻟا يﻮﺘﺴﻤﻟا ﻰﻓ ﺔﻴﺟرﺎﺨﻟا تﺎﻣاوﺪﻟا ىﺪﻣو مدﺎﻘﻟا ءاﻮﻬﻟا ﻩﺎﺠﺗا
ﻟازﺪﻴﻟﻮﻨﻳر ﻢﻗر نﻮﻜﻳ ﺎﻣﺪﻨﻋ ﺎﻬﻴﻠﻋ ناروﺪﻟا ﺔﻳوازو مﻮﺠﻬﻟا ﺔﻳواز ﺮﻴﺛﺎﺗ ﺪﻤﺘﻌﻤ
ىوﺎﺴﻳ حﺎﻨﺠﻟاﺮﺗو ﻰﻠﻋ 10
5× 3.82 . سﻮﻜﻌﻤﻟا ﺎﺘﻟﺪﻟا حﺎﻨﺟ ﻦﻴﺑ ﻞﺛﺎﻤﺗ ﺪﺟو
ﺔﻴﺿﺮﻌﻟا ﺔﻋﺮﺴﻟا ةورذ ﺔﻤﻴﻗ ﻲﻓ ىدﺎﻌﻟا ﺎﺘﻟﺪﻟا حﺎﻨﺟو .
مﺪﺨﺘﺳاو
ﺎﺘﻟد حﺎﻨﺠﻟ ﺔﻴﺋاﻮﻬﻟا ﺎﻜﻴﻣﺎﻨﻳﺪﻟا تﻼﻣﺎﻌﻣ سﺎﻴﻘﻟ ﺔﺘﺴﻟا لﺎﻤﺣﻼﻟ نزاﻮﺘﻟازﺎﻬﺟ ﺎﺘﻟد حﺎﻨﺟو سﻮﻜﻌﻣ .
ا ﺔﻋﺮﺴﻟا راﺪﻘﻣ نا ﺔﺳارﺪﻟا ﻩﺬه ﻦﻣ ﺞﺘﻨﺘﺳاو ﺔﻴﺿﺮﻌﻟ
ﺔﻟﺎﺣ ﻲﻓ ﺎﻬﺘﻤﻴﻗ ﻦﻴﺗﺮﻣ لدﺎﻌﺗ ﺔﻨﻴﻌﻣ مﻮﺠه ﺔﻳواز ﺪﻨﻋ ﺎﺘﻟﺪﻟا حﺎﻨﺟ تﺎﻣاوﺪﻟ تﺎﻣاوﺪﻟا ﺮﻳﺪﺗو ناورﺪﻟا ﺔﻋﺮﺳ ﻢﻴﻗ ﻚﻟﺬآو سﻮﻜﻌﻤﻟا ﺎﺘﻟﺪﻟا حﺎﻨﺟ ماﺪﺨﺘﺳا ماﺪﺨﺘﺳﺎﺑ ﺎﻧرﺎﻘﻣ ﺎﺘﻟﺪﻟا حﺎﻨﺟ ماﺪﺨﺘﺳا ﺔﻟﺎﺣ ﻰﻓ ﻲﻠﻋا ﻊﻓﺮﻟا ةﻮﻗ تﻼﻣﺎﻌﻣو حﺎﻨﺟ سﻮﻜﻌﻣ ﺎﺘﻟد .
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APPROVAL PAGE
I certify that I have supervised and read this study and that in my opinion, it conforms to acceptable standards of scholarly presentation and is fully adequate, in scope and quality, as a thesis for the degree of Master of Science in Mechanical Engineering.
...……….
Ashraf Ali Omar Supervisor
...……….
Waqar Asrar Co-Supervisor
I certify that I have read this study and that in my opinion, it conforms to acceptable standards of scholarly presentation and is fully adequate, in scope and quality, as a thesis for the degree of Master of Science in Mechanical Engineering.
...……….
Ashraf Ali Omar
Supervisor
This thesis was submitted to the Department of Mechanical Engineering and is accepted as a fulfilment of the requirement for the degree of Master of Science in Mechanical Engineering.
...……….
Waqar Asrar
Head, Department of Mechanical Engineering
This thesis was submitted to the Kulliyyah of Engineering and is accepted as a fulfilment of the requirement for the degree of Master of Science in Mechanical Engineering.
...…..……….
Amir Akramin Shafie Dean, Kulliyyah of
Engineering
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DECLARATION
I hereby declare that this thesis is the result of my own investigations, except where otherwise stated. I also declare that it has not been previously or concurrently submitted as a whole for any other degrees at IIUM or other institutions.
Afaq Altaf
Signature ……….. Date ………...
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INTERNATIONAL ISLAMIC UNIVERSITY MALAYSIA
DECLARATION OF COPYRIGHT AND
AFFIRMATION OF FAIR USE OF UNPUBLISHED RESEARCH
Copyright © 2010 by Afaq Altaf. All rights reserved.
A STUDY OF THE REVERSE DELTA WING USING PARTICLE IMAGE VELOCIMETRY (PIV)
No part of this unpublished research may be reproduced, stored in a retrieval system, or transmitted, in any form or by any means, electronic, mechanical, photocopying, recording or otherwise without prior written permission of the copyright holder except as provided below.
1. Any material contained in or derived from this unpublished research may only be used by others in their writing with due acknowledgement.
2. IIUM or its library will have the right to make and transmit copies (print or electronic) for institutional and academic purposes.
3. The IIUM library will have the right to make, store in a retrieval system and supply copies of this unpublished research if requested by other universities and research libraries.
Affirmed by Afaq Altaf.
……….. ………
Signature Date
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ACKNOWLEDGEMENTS
I would like to thank International Islamic University Malaysia for supporting this work through the Research Management Centre under the research grant no. EDW B0905-294.
I would like to express my gratitude towards, Prof. Dr. Ashraf Ali Omar (supervisor) and Prof. Dr. Waqar Asrar (co-supervisor) for believing in me and providing me with an opportunity to be working under their supervision. I am thankful to them for providing me an opportunity to learn the technique of Particle Image Velocimetry, broadening my understanding of experimental aerodynamics and for assisting me throughout the research work.
I would like to express my gratitude towards, Omer A. Elsayed, for teaching me the technique of Particle Image Velocimetry and helping me in analyzing the data.
I would like to express my gratitude towards Hani Bin Ludin @ Jamaluddin for simulating the flow for the research model.
I would like to thank the technical staff at the automotive laboratory of IIUM for helping me throughout my research work. I would personally like to thank Mohd.
Norhafiz Bin Adnan, Firdaus Hakeem Bin Ma’arof, Meor Jalalullail Bin Meu Othman, Azliza Binti Embong, Ramlee Ariffin and Azweeda Binti Dahalan. Wind tunnel testing, fabrication of the experimental model and experimental setups would not have been possible without their generous help.
I would also like to express my gratitude towards the entire staff of mechanical engineering department for helping me directly or indirectly in achieving my set goal.
Finally, I would like to thank my family members and friends for the great support they have showered towards me. I am thankful to them for showing patience towards me and providing me with valuable advice.
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TABLE OF CONTENTS
Abstract ...iii
Absract in Arabic ...iii
Approval Page ...iv
Declaration Page ...v
Copyright Page...vi
Acknowledgements ...vii
List of Tables ...x
List of Figures ...xi
List of Abbreviations ...xvi
List of Symbols ...xvii
CHAPTER ONE: INTRODUCTION ...1
1.1 Overview ...1
1.2 Overview of Previous Research Work ...3
1.3 Problem Statement and its Significance ...5
1.4 Research Philosophy ...5
1.5 Research Objectives ...6
1.6 Research Methodology ...6
1.7 Scope of Research ...7
1.8 Thesis Organization ...7
CHAPTER TWO: LITERATURE REVIEW ...9
2.1 Introduction ...9
2.2 Delta Wing ...12
2.3 Reverse Delta Wing ...16
2.4 Methods of Minimizing Wake Vortex Encounters ...18
2.4.1 Detection and Avoidance ...18
2.4.2 Wake Vortex Alleviation ...20
2.4.3 Vortex Dissipation ...22
2.5 Summary ...23
CHAPTER THREE: EXPERIMENTAL APPARATUS AND METHODOLOGY ...25
3.1 Introduction ...25
3.2 Experimental Apparatus...25
3.2.1 The Wind Tunnel ...25
3.2.2 The Model ...26
3.2.3 The Measurement Technique ...27
3.3 Experimental Procedure ...30
3.4 Summary ...33
CHAPTER FOUR: PIV ANALYSIS METHODS ...34
4.1 Introduction ...34
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4.2 Vector Statistics ...34
4.3 Scalar Map ...36
4.4 Vorticity ...37
4.5 Streamlines ...38
4.6 Summary ...39
CHAPTER FIVE: THREE DIMENSIONAL FLOW BOUNDARY CORRECTIONS ...40
5.1 Introduction ...40
5.2 Buoyancy ...40
5.3 Solid Blockage ...41
5.4 Wake Blockage ...43
5.4.1 Approximate Blockage Corrections ...43
5.5 Streamline Curvature ...44
5.6 General Downwash Corrections ...46
5.7 Summary ...47
CHAPTER SIX: RESULTS AND DISCUSSIONS ...48
6.1 Introduction ...48
6.2 Velocity Vectors ...48
6.3 Vorticity Contours ...53
6.4 Tangential Velocity ...54
6.5 Circulation...66
6.6 Simulation ...78
6.7 Aerodynamic Coefficients ...80
6.8 Summary ...87
CHAPTER SEVEN: UNCERTAINTY ANALYSIS ...89
7.1 Introduction ...89
7.2 Statistical Uncertainty ...90
7.3 Uncertainty in the Wing Angle of Attack ...90
7.4 Uncertainty in Displacement ...91
7.5 Total Uncertainty ...92
7.6 Summary ...92
CHAPTER EIGHT: CONCLUSIONS AND RECOMMENDATIONS ...93
8.1 Conclusions ...93
8.2 General Recommendations ...94
8.3 Recommendations for Further Studies ...95
BIBLIOGRAPHY ...97
LIST OF PUBLICATIONS ...100
APPENDIX: CALIBRATION PROCESS, FIELD OF VIEW MEASUREMENT AND CALCULATION OF THE LASER SEPARATION TIME ...101 1
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LIST OF TABLES
Table No. Page No.
5.1 Corrected values of Aerodynamic coefficients 47
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LIST OF FIGURES
Figure No. Page No.
1.1 Aircraft wake vortex encounter 2
1.2 Wake vortex created by an aircraft 2
1.3 Schematic diagram of trailing vortices near a runway 4 2.1 Schematic of the subsonic flow field over the top of a delta wing
at an angle of attack 12
2.2 Leading edge vortices over the top surface of a delta wing at an
angle of attack. 13
2.3 The flow field in the crossflow plane above a delta wing at an
angle of attack, showing the two primary leading-edge vortices 14 2.4 Flow over a forward swept wing and a swept-back wing 18
2.5 The Green Wake Project Concept 20
3.1 Reverse Delta Wing Geometry 26
3.2 Inclined Plane Measurement Scale 27
3.3 Principle layout of a PIV system 28
3.4 Double cavity Nd:YAG PIV laser 29
3.5 Experimental Setup 31
3.6 Experimental setup as seen in the wind tunnel test section 31 3.7 PIV image map of the reverse delta wing tip vortex (left wing tip)
at α=30º, φ=0º 32
4.1 Velocity vector maps 35
4.2 Tangential velocity maps 36
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4.3 Tangential velocity contours 37
4.4 Vorticity maps 38
4.5 Vorticity contours 38
4.6 Streamlines 39
5.1 Values of K1 and K3 for a number of bodies 42 5.2 Values of τ1 for a number of tunnel types 42 5.3 Values of δ for a wing with uniform loading in a closed
rectangular tunnel 45
5.4 Values of τ2 for open and closed jets 45 6.1 Reverse delta wing vortex at α=5º, φ=0º and Rec=3.82×105 50 6.2 Regular delta wing vortex at α=5º, φ=0º and Rec=3.82×105 50 6.3 Reverse delta wing vortex at α=10º, φ=0º and Rec=3.82×105 51 6.4 Regular delta wing vortex at α=10º, φ=0º and Rec=3.82×105 51 6.5 Reverse delta wing vortex at α=15º, φ=0º and Rec=3.82×105 51 6.6 Regular delta wing vortex at α=15º, φ=0º and Rec=3.82×105 52 6.7 Reverse delta wing vortex at α=20º, φ=0º and Rec=3.82×105 52
6.8 Regular delta wing vortex at α=20º, φ=0º and Rec=3.82×105 52 6.9 Velocity vectors at α=20°, φ=5° at a location of x/c=3.418 for a
reverse delta wing and (b) a regular delta wing with Vorticity (1/s)
as its background 54
6.10 Vorticity contour at α=20°, φ=5° at a location of x/c=3.418 for a
reverse delta wing and a regular delta wing 54 6.11 Tangential velocity distributions of a reverse delta wing at α=5º,
x/c=1.359 and α=5º, x/c=3.418 56 6.12 Tangential velocity distributions of a regular delta wing at α=5º,
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x/c=1.359 and α=5º, x/c=3.418 56 6.13 Tangential velocity distributions of a reverse delta wing at α=10º,
x/c=1.359 and α=10º, x/c=3.418 57
6.14 Tangential velocity distributions of a regular delta wing at α=10º,
x/c=1.359 and α=10º, x/c=3.418 57
6.15 Tangential velocity distributions of a reverse delta wing at α=15º,
x/c=1.359 and α=15º, x/c=3.418 58
6.16 Tangential velocity distributions of a regular delta wing at α=15º,
x/c=1.359 and α=15º, x/c=3.418 58
6.17 Tangential velocity distributions of a reverse delta wing at α=20º,
x/c=1.359 and α=20º, x/c=3.418 59
6.18 Tangential velocity distributions of a regular delta wing at α=20º,
x/c=1.359 and α=20º, x/c=3.418 59
6.19 Tangential velocity distributions of a reverse delta wing at φ=0º,
x/c=1.359 and φ=0º, x/c=3.418 60 6.20 Tangential velocity distributions of a regular delta wing at φ=0º,
x/c=1.359 and φ=0º, x/c=3.418 60 6.21 Tangential velocity distributions of a reverse delta wing at φ=5º,
x/c=1.359 and φ=5º, x/c=3.418 61 6.22 Tangential velocity distributions of a regular delta wing at φ=5º,
x/c=1.359 and φ=5º, x/c=3.418 61 6.23 Tangential velocity distributions of a reverse delta wing at φ=-5º,
x/c=1.359 and φ=-5º, x/c=3.418 62
6.24 Tangential velocity distributions of a regular delta wing at φ=-5º,
x/c=1.359 and φ=-5º, x/c=3.418 62
6.25 Tangential velocity distributions of a reverse delta wing at φ=10º,
x/c=1.359 and φ=10º, x/c=3.418 63
6.26 Tangential velocity distributions of a regular delta wing at φ=10º,
x/c=1.359 and φ=10º, x/c=3.418 63
6.27 Tangential velocity distributions of a reverse delta wing at φ=-10º,
x/c=1.359 and φ=-10º, x/c=3.418 64
6.28 Tangential velocity distributions of a regular delta wing at φ=-10º,
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x/c=1.359 and φ=-10º, x/c=3.418 64
6.29 Circulation distributions of a reverse delta wing at α=5º,
x/c=1.359 and α=5º, x/c=3.418 66 6.30 Circulation distributions of a regular delta wing at α=5º,
x/c=1.359 and α=5º, x/c=3.418 67 6.31 Circulation distributions of a reverse delta wing at α=10º,
x/c=1.359 and α=10º, x/c=3.418 67
6.32 Circulation distributions of a regular delta wing at α=10º,
x/c=1.359 and α=10º, x/c=3.418 68
6.33 Circulation distributions of a reverse delta wing at α=15º,
x/c=1.359 and α=15º, x/c=3.418 68
6.34 Circulation distributions of a regular delta wing at α=15º,
x/c=1.359 and α=15º, x/c=3.418 69
6.35 Circulation distributions of a reverse delta wing at α=20º,
x/c=1.359 and α=20º, x/c=3.418 69
6.36 Circulation distributions of a regular delta wing at α=20º,
x/c=1.359 and α=20º, x/c=3.418 70
6.37 Circulation distributions of a reverse delta wing at φ=0°,
x/c=1.359 and φ=0°, x/c=3.418 72 6.38 Circulation distributions of a regular delta wing at φ=0°,
x/c=1.359 and φ=0°, x/c=3.418 72 6.39 Circulation distributions of a reverse delta wing at φ=5°,
x/c=1.359 and φ=5°, x/c=3.418 73 6.40 Circulation distributions of a regular delta wing at φ=5°,
x/c=1.359 and φ=5°, x/c=3.418 73 6.41 Circulation distributions of a reverse delta wing at φ=-5°,
x/c=1.359 and φ=-5°, x/c=3.418 74
6.42 Circulation distributions of a regular delta wing at φ=-5°,
x/c=1.359 and φ=-5°, x/c=3.418 74
6.43 Circulation distributions of a reverse delta wing at φ=10°,
x/c=1.359 and φ=10°, x/c=3.418 75
6.44 Circulation distributions of a regular delta wing at φ=10°,
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x/c=1.359 and φ=10°, x/c=3.418 75
6.45 Circulation distributions of a reverse delta wing at φ=-10°,
x/c=1.359 and φ=-10°, x/c=3.418 76
6.46 Circulation distributions of a regular delta wing at φ=-10°,
x/c=1.359 and φ=-10°, x/c=3.418 76
6.47 Pressure contours over the lower surface and upper surface at
α=30°, φ=0° for a reverse delta wing and a regular delta wing 79 6.48 Streamlines over a reverse delta wing and a regular delta wing, 80
at α=30°, φ=0°
6.49 A comparison of 65° delta wing lift coefficient based on theory,
experiments and CFD Simulation 81
6.50 Aerodynamic coefficients versus angle of attack (α) for a
reverse delta wing and a regular delta wing at φ=0° 82 6.51 Aerodynamic coefficients versus angle of attack (α) for a
reverse delta wing and a regular delta wing at φ=10° 83 6.52 Aerodynamic coefficients versus angle of attack (α) for a
reverse delta wing and a regular delta wing at φ=-10° 84 6.53 Aerodynamic coefficients versus angle of attack (α) for a
reverse delta wing and a regular delta wing at φ=20° 85 6.54 Aerodynamic coefficients versus angle of attack (α) for a
reverse delta wing and a regular delta wing at φ=-20° 86
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LIST OF ABBREVIATIONS
CCD Charge Coupled Device
DW Delta Wing
et al. (et alia): and others
FAA Federal Aviation Administration
FLAME Future Laser Atmospheric Measurement Equipment LIDAR Laser Doppler Radar
MFLAME Multifunction Future Laser Atmospheric Measurement Equipment NASA National Aeronautics and Space Administration
Nd:YAG Neodymium Yttrium Aluminium Garnet PIV Particle Image Velocimetry
RDW Reverse Delta Wing
UV LIDAR Ultra Violet Laser Doppler Radar
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LIST OF SYMBOLS
c Chord length (m) k Number of points (m) r Radius (m)
rc Core radius (m)
Rec Reynolds number based on chord V∞ Free stream velocity (m/s)
Vθ Tangential velocity (m/s) x Streamwise coordinate (m) y Spanwise coordinate (m) z Transverse coordinate (m)
x/c Ratio of streamwise coordinate and chord length α Angle of attack (deg)
φ Roll angle (deg) Γ Circulation (m2/s) a Wing lift curve slope CL Lift coefficient CD DragCoefficient Cm Moment Coefficient L/D Lift to drag ratio
N Number of measurement samples
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σ Variance
Vi Mean velocity (m/s) K3 Body shape factor
τ1 Ratio of tunnel test section shape to (model span/tunnel width) τ2 Downwash correction factor
δ Boundary correction factor
λ Ratio of tunnel height to tunnel width k Ratio of wing span to tunnel width C Tunnel Test section area (m2) S Surface area (m2)
b Span (m)
t Thickness (m)
B Test section width (m) H Test section height (m) εsb Solid blockage εwb Wake blockage εt Total blockage
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CHAPTER ONE INTRODUCTION
1.1 OVERVIEW
Vortices created by aircrafts are an inevitable consequence of the creation of lift.
Vortices persist for many miles and wake vortex encounters pose a grave hazard to trailing aircrafts that fly in close proximity especially during takeoff and landing because the wing tip vortex circulation is at a maximum (Arndt et al., 1991). The severity of the interaction that can occur when a plane encounters one of these tip vortices is very grave. A trailing plane travelling at a speed of 60 m/s, which flies axially into the tip vortex generated by a B757, experiences huge changes in incident flow angle which can both stall the wing and generate large rolling moments that the aircraft ailerons cannot control (Page et al., 1991). The potential for additional disasters as a result of wake vortex interactions is obviously still great.
Vortices produced by small aircrafts are almost negligible, but vortices created by larger and heavier aircrafts can be extremely dangerous for a distance of many miles to a trailing aircraft. Wake vortex and turbulence generated by large aircrafts can cause instability, uncontrollable rolls, and sudden loss of altitude to a trailing aircraft. There have been incidents, especially at lower altitudes during landing approaches, when wake turbulence has resulted in fatal accidents because of
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insufficient time and altitude for pilots to regain full control of their aircraft after being buffeted violently by the powerful vortices.
Figure 1.1 illustrates various types of wake encounters and Figure 1.2 is an image of a wake vortex created by a climbing aircraft.
Figure 1.1: Aircraft wake vortex encounter [Brian and Robert, 2004].
Figure 1.2: Wake vortex created by an aircraft [www.airliners.net, retrieved 2010].
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Tip vortices limit the spacing between aircrafts within the takeoff and landing corridors at busy airports and hence, increase the time intervals between consecutive landings and takeoffs.
Brian and Robert (2004) state that air traffic control authorities have adopted a set of operational regulations that have provided commercial aviation with a safe solution to the problem of aircraft wake vortex encounters. The regulations include aircraft segregation by size, controlling flight paths during takeoff and landing, and maintaining a fixed separation between aircrafts based on the size of the lead aircraft.
Due to an uncertainty in knowing where the vortices are located relative to the flight path, the separation distances between aircrafts have to be selected on an overly conservative basis. These procedures limit the traffic capacity at many major airports since a smaller number of aircrafts can be entertained daily than could be handled if a permanent solution to wake vortex encounters becomes available.
1.2 OVERVIEW OF PREVIOUS RESEARCH WORK
The first series of NASA (National Aeronautics and Space Administration) wake turbulence research flights were conducted by the NASA Dryden Flight Research Centre in 1969-70, when large wide-bodied jet transports were being introduced into commercial and military service and a mix of large, medium, and small aircraft was being seen at nearly every major airport. These first tests measured the strength of vortices created by large “jumbo” aircraft and measured their effects on smaller trailing aircraft. Between the years 1972 and 1974, three additional series of wake turbulence tests were flown at NASA Dryden with emphasis on the use of wing flaps, speed brakes, and spoilers to alter the formation, strength, and behaviour of vortices generated by large aircraft.
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Data from the research flights aided the FAA (Federal Aviation Administration) in establishing the current separation distances that require that during landings a small aircraft must remain six miles behind a large wide-bodied aircraft, and two large wide-body jets cannot be closer than four miles. During takeoffs, three-minute intervals are required between small and large aircraft.
Separation distances of up to five miles are also applied between various sizes of aircraft while cruising at or near the same altitude.
Wake turbulence generated by an aircraft may be a hazard for the following aircraft. Figure 1.3 shows a schematic diagram of trailing vortices near a runway. To protect against this risk, regulations specify minimum separation between aircraft during both approach and landing. Work is now underway to overcome this restriction, or at least to reduce the required separation under certain conditions, which would increase traffic capacity during peak periods and reduce saturation at heavily used airports. In turn, this would enable postponing or avoiding, the construction of new airports or runways, thereby considerably decreasing costs and environmental impact.
Figure 1.3: Schematic diagram of trailing vortices near a runway.
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Worldwide research has been focused on increasing airport capacity by minimizing the wake vortex hazard. Hinton et al. (1999) focus their research work on developing a vortex detection system that can be used to safely reduce aircraft spacing by enabling aircrafts not to cross the paths of vortices.
1.3 PROBLEM STATEMENT AND ITS SIGNIFICANCE
Aircraft wake vortex flow is turbulent flow that is formed behind an aircraft as it passes through air. The problem faced by the aviation industry is how to alleviate the wake vortex generated by aircrafts so as to avoid trailing aircrafts to be bathed in turbulent vortex flows which can cause a high degree of instability and loss of control.
The purpose of this thesis is to investigate the flow over a reverse delta wing.
Reverse delta wings may be used in vortex alleviation. They may also be utilised in forward swept winged aircraft. PIV technique, a minimally intrusive method, is used in this work to study the vortex characteristics of a reverse delta wing at various angles of attack and roll at a chord-based Reynolds number of Rec=3.82×105. Vortex characteristics such as maximum tangential velocity, vorticity and circulation are studied. The six component force balance is used to study the aerodynamic coefficients of the reverse delta wing and the delta wing. Existing simulation results are studied and compared with this research work.
1.4 RESEARCH PHILOSOPHY
Although many proposed ideas of wake vortex alleviation have been adopted and practiced, missing links still exist which do not allow for sufficient wake vortex alleviation. Europeans have recently suggested an add-on device which resembles an inclined delta-type wing (reverse delta wing) to alleviate wake vortices by causing
6
instability within the vortices, but they have thus far withheld all such potential information.
This study shall explore characteristics of the reverse delta wing vortices and discuss the results in detail, so as to find out if the reverse delta wing as an add-on device has the capability to alleviate wake vortices.
1.5 RESEARCH OBJECTIVES
The research objectives are:
1. To visualize the flow over a reverse delta wing.
2. To investigate the characteristics of the vortices using Particle Image Velocimetry.
3. To experimentally obtain the aerodynamic coefficients of a regular delta wing and a reverse delta wing using the Six Component Force Balance and to compare them with each other.
1.6 RESEARCH METHODOLOGY
The research was a laboratory based experimental work. The main objective of the research was to investigate the characteristics of the vortices using Particle Image Velocimetry.
1. Advanced PIV technique was used to measure the wake vortex behind a regular and a reverse delta wing.
2. Angle of attack and roll angle were varied.
3. Measurements were conducted at two planes downstream of the wing.
