Master of Science Thesis

Passive Flow Separation Control on an Airfoil-Flap Model The Effect of Cylinders and Vortex Generators

D.P. Jansen

August 2012

Faculty of Aerospace Engineering · Delft University of Technology

Passive Flow Separation Control on an Airfoil-Flap Model The Effect of Cylinders and Vortex Generators

Master of Science Thesis

For obtaining the degree of Master of Science in Aerospace Engineering at Delft University of Technology

D.P. Jansen

August 2012

Faculty of Aerospace Engineering · Delft University of Technology Delft University of Technology

Copyright ¥c Aerospace Engineering, Delft University of Technology All rights reserved. Delft University Of Technology Department of Aerodynamics

The undersigned hereby certify that they have read and recommend to the Faculty of Aerospace En- gineering for acceptance a thesis entitled “Passive Flow Separation Control on an Airfoil-Flap Model” by D.P. Jansen in partial fulfillment of the requirements for the degree of Master of Science .

Dated: August 2012

Head of department: Prof.dr.ir.drs. H. Bijl

Supervisor: Dr.ir. L.L.M. Veldhuis

Reader: Dr.ir. B.W. van Oudheusden

Reader: Ir. W.A. Timmer

Acknowledgements

This report is written as a thesis for the degree of Master of Science at the Faculty of Aerospace Engi- neering, Department of Aerodynamics, at Delft University of Technology. I would like to take the opportunity to thank all the people who have helped me in accomplishing this graduation research. I would like to thank Dr. ir. L.L.M. Veldhuis, my supervisor, for his full support, enthusiasm and sharing of his in depth knowledge on experimental research and the field of flow separation control. I could not have wished for a better advisor for this project. I also would like to thank Dipl. ing. S. Bernardy and L. Molenwijk for their willingness in answering my questions and their assistance in performing the wind tunnel experiments. Then I would like to thank my parents for their motivation and advice in times most needed. Special thanks goes to R. Peacock for his sincere curiosity and help on the project. Finally I also want to mention Kim-Lan, my girlfriend, because without her love and unlimited support I definitely would not have achieved this result.

Delft, The Netherlands D.P. Jansen August 2012

v D.P. Jansen vi MSc. Thesis Summary

In the on-going search for improvements in high lift device performance, the field of flow separation control has seen explosive growth in recent years. This is due to a continuously improved understanding of fluid mechanics, the development in experimental and computational techniques and the variety of flow control mechanisms existing today. These multiple mechanisms can be differentiated into active and passive flow control techniques. In this research two passive flow separation control methods are tested on their ability to postpone flow separation on an airfoil with trailing edge flap model of the Extra EA-400. The two methods are the application of vortex generators on the airfoil and the positioning of vortex producing cylinders in the slot of the airfoil flap system. The 2D airfoil flap model was tested by experimental and numerical investigation. The flap deflection was ◦ ◦ ◦ set at δf = 45 , δf = 50 and δf = 55 to test the flap at moderate to highly separated flow conditions. Measurements in the experiment were done by pressure holes and probes and oil flow visualization. Numerical calculations were performed in FLUENT as part of the ANSYS 13.0 package, which was also used to create the grid. The model is a RANS analysis with the k − ω SST turbulence model and low Re near wall treatment. It was found in the experiments that both the vortex generators and the cylinders were able to delay flow separation and thereby increase the lift of the airfoil flap system. Both devices are able to create vortices that entrain flow into the boundary layer, which gets re-energized. The cylinders proved to be superior compared to the vortex generators in the ability to postpone the point of flow separation. With respect to size and positioning the cylinders also showed a larger bandwidth than the vortex generators at which the devices were effective. The results from the experiments indicated that the reduced frequency for the cylinder vortex shedding resulting in the largest delay in flow separation, matched the theoretical predictions found in literature. Especially at high flap deflections significant lift gains were obtained up to ◦ 6% for the vortex generators and 18% for the cylinders at a deflection of δf = 55 . At the flap deflection ◦ of δf = 45 the flow separation delay capabilities of the cylinders are less effective. The frequency tests, oil flow visualization and boundary layer measurements were used to understand the performance and working principles of the cylinders. The numerical simulations showed difficulties in reproducing the results for the cylinders obtained from the experiment. While for the attached airfoil flow the numerical simulations showed good agreement with the experiment, the flow separation on the flap at high deflections was not accurately modelled for both the baseline and cylinder configuration of the airfoil flap. The vortices created in the simulation showed a different frequency than found in the experiment and only minor effects of the cylinder were observed. Separation on the flap was very slightly delayed, although no significant mixing effects of the vortices were visible in the boundary layer.

vii D.P. Jansen viii MSc. Thesis Contents

Acknowledgements v

Summary vii

List of Figures xii

List of Tables xvii

Nomenclature xxi

1 Introduction 1 1.1 The need for high lift systems ...... 1 1.2 Limits in high lift performance ...... 3 1.3 Flow separation in high lift performance ...... 5 1.3.1 Plain airfoil characteristics ...... 5 1.3.2 Flapped airfoil characteristics ...... 6 1.4 Flow separation control ...... 8 1.4.1 Active flow separation control techniques ...... 8 1.4.2 Passive flow separation control techniques ...... 10 1.5 Opportunity of research ...... 13 1.6 Layout of the report ...... 13

2 Experimental model setup 15 2.1 Introduction ...... 15 2.2 The wind tunnel and model ...... 15 2.3 Passive flow separation methods ...... 18 2.3.1 Vortex generators ...... 19

ix Contents

2.3.2 Cylinders ...... 20 2.4 Measurement techniques ...... 23

2.4.1 Pressure measurements and calculation of Cl and Cd ...... 23 2.4.2 Frequency measurements ...... 25 2.4.3 Oil flow visualization ...... 26

3 Experimental results 27 3.1 Introduction ...... 27 3.2 Baseline model ...... 27 3.3 Vortex generators ...... 31 3.4 Cylinders ...... 34 ◦ 3.4.1 Flap deflection δf = 55 ...... 34 ◦ 3.4.2 Flap deflection δf = 50 ...... 42 ◦ 3.4.3 Flap deflection δf = 45 ...... 43

3.4.4 Comparison of δf effects on cylinder performance and further investigation . . . . 47 3.5 Conclusion experimental investigation ...... 52

4 Numerical investigation 55 4.1 Introduction ...... 55 4.2 General setup ...... 55 4.3 The numerical domain ...... 56 4.4 Model validation ...... 61 ◦ 4.4.1 Low flap deflection δf = 15 , no flow separation ...... 61 ◦ 4.4.2 High flap deflection δf = 45 ...... 62 ◦ 4.4.3 Very high flap deflection δf = 55 ...... 63 4.4.4 Conclusion on baseline model validation ...... 63 4.4.5 Influence of tunnel wall at large flap deflection ...... 63 4.4.6 Modelling laminar zones in CFD ...... 65 4.5 Results for baseline and cylinder configurations ...... 67 4.5.1 Baseline and cylinder velocity field ...... 68 4.5.2 Vortex development ...... 68 4.5.3 Vortex influence on the flap ...... 74 4.6 Conclusion on numerical investigation ...... 76

5 Conclusion and Recommendations 79 5.1 Conclusion ...... 79 5.2 Recommendations ...... 80

D.P. Jansen x MSc. Thesis Contents

Bibliography 81

A Test matrix for the wind tunnel experiments 83

B Additional pictures on the wind tunnel and model layout 85

C Output file example of pressure measurements from Profmeasure 89

MSc. Thesis xi D.P. Jansen Contents

D.P. Jansen xii MSc. Thesis List of Figures

1.1 General high lift system with single slotted trailing edge and leading edge flap...... 2 1.2 Lift polar showing the effects of a trailing edge and leading flap deflection on a plain airfoil. 2 1.3 Circular-arc mean lines A, B, C obtained from the Joukowski transformation. Figure from Smith [1975]...... 3 1.4 Flow around a cylinder with a very high circulation, such that the two stagnation points coincide. Figure from Smith [1975]...... 4 1.5 Visualization of flow separation on a flap. Figure from Little et al. [2009]...... 4 1.6 Visualization of a laminar separation bubble...... 5 1.7 Visualization of flow separation...... 6 1.8 Lift polar showing the general effect of increasing trailing edge flap deflection...... 7 1.9 Triple-slot flap system of the B727-200. Figure from Rudolph [1996]...... 8 1.10 PIV visualization of flow separation postponement on a wind turbine blade by applying synthetic jets at α = 16 ◦ and Re = 1 .6 · 10 5. Figure from Maldonado et al. [2010]. . . . . 9 1.11 Components of an electrodynamic synthetic jet. Figure from Cattafesta [2011]...... 9 1.12 Working principle of boundary layer suction. Figure from Boermans [2008]...... 10 1.13 Schematization of a piezoelectric flap. Figure from Veldhuis and van der Jagt [2010]. . . . 10 1.14 Schematization of a DBD plasma actuator. Figure from LeBeau [2007]...... 11 1.15 Vortex generators displayed in different setups. Figures from von Stillfried et al. [2010]. . 11 1.16 Effect of counter-rotating micro VG’s on RMS anemometer output from hot-film sensors ◦ on flap upper surface with α = 8 and M∞ = 0 .2. Figure from Lin [1999]...... 12 1.17 Lift-enhancing tabs as used by Lee et al. [2005]...... 12

2.1 The Low Speed Low Turbulence Wind Tunnel at Delft University of Technology...... 16 2.2 Cross section of the model and definition of flap parameters...... 16 2.3 The model positioned in the wind tunnel with the cylinder test setup...... 17 2.4 Boundary layer measurements on model lower side at x/c = 0 .67 ...... 18

xiii List of Figures

2.5 The model with VG’s applied (left) and cylinders (right)...... 19 2.6 Vortex generator geometrical definitions...... 20 2.7 Vortex generator geometry and construction method...... 20 2.8 Cylinder position and definition of coordinates. Hinge point is referenced to the airfoil nose with [x, y ] = [0 .72 c, 0.06 c]...... 21 2.9 Cylinder geometrical definitions...... 21 2.10 Reduced frequency dependency on cylinder diameter...... 23 2.11 Test setup of the wake rake...... 25

3.1 Lift polars for the baseline configurations at multiple flap deflections at Re = 2 .0 · 10 6. . . 29 ◦ ◦ 3.2 Pressure distributions for baseline configurations at multiple a.o.a for δf = 55 , δf = 50 ◦ and δf = 45 ...... 30 3.3 Effect of height, spacing and edge distance on the lift for multiple VG configurations at ◦ δf = 55 ...... 32 3.4 Flap and airfoil lift contributions for multiple VG configurations with h∗ = 0 .023 , d∗ = 0 ◦ and δf = 55 ...... 33 3.5 Pressure distribution for optimum VG configuration at multiple a.o.a. with λ/h = 4 .0, ∗ ∗ ◦ h = 0 .023 , d = 0 and δf = 55 ...... 33 3.6 Pressure distributions for optimum VG and baseline configuration at low and high a.o.a. ∗ ∗ with λ/h = 4 .0, h = 0 .023 , d = 0 and δf = 55 ...... 34 ◦ 3.7 Effect of size and positioning on the lift for multiple cylinder configurations at δf = 55 . . 35 ◦ 3.8 Effect of size and positioning on the lift for multiple cylinder configurations at δf = 55 . . 36 + 3.9 Lift gains ∆Cl dependency against reduced frequency F at high and low angle of attack ∗ ◦ ◦ with r = 0 .090 , θ = 37 and δf = 55 ...... 36 3.10 Flap and airfoil lift contributions for multiple cylinder configurations with D∗ = 0 .0167 ◦ and δf = 55 ...... 37 3.11 Pressure distributions for optimum cylinder and baseline configuration at low and high ∗ ∗ ◦ ◦ a.o.a. with D = 0 .0167 , r = 0 .090 , θ = 37 and δf = 55 ...... 38 3.12 Oil flow visualization on the flap for the baseline and optimum cylinder configuration at ◦ δf = 55 ...... 39 3.13 Effect of positioning on the lift for multiple cylinder configurations with D∗ = 0 .0167 and ◦ δf = 50 ...... 40 3.14 Flap and airfoil lift contributions for multiple cylinder configurations with D∗ = 0 .0167 ◦ and δf = 50 ...... 41 3.15 Pressure distributions for optimum cylinder and baseline configuration at low and high ∗ ∗ ◦ ◦ a.o.a. D = 0 .0167 , r = 0 .090 , θ = 30 , δf = 50 ...... 42 3.16 Oil flow visualization on the flap for the baseline and optimum cylinder configuration at ◦ δf = 50 ...... 44 3.17 Effect of positioning on the lift for multiple cylinder configurations with D∗ = 0 .0167 and ◦ δf = 45 ...... 45 3.18 Flap and airfoil lift contributions for multiple cylinder configurations with D∗ = 0 .0167 ◦ and δf = 45 ...... 46

D.P. Jansen xiv MSc. Thesis List of Figures

3.19 Pressure distributions for optimum cylinder and baseline configuration at low and high ∗ ∗ ◦ ◦ a.o.a. with D = 0 .0167 , r = 0 .090 , θ = 37 and δf = 45 ...... 47 3.20 Oil flow visualization on the flap for the baseline and optimum cylinder configuration at ◦ δf = 45 ...... 48 3.21 Frequency measurements at multiple positions behind the cylinder. Pressure fluctuations ∗ ∗ ◦ ◦ |∆P (f) | are plotted against frequency with D = 0 .0167 , r = 0 .090 , θ = 37 and δf = 45 . 50 ◦ 3.22 Frequency measurement locations at the flap at δf = 45 ...... 51 3.23 Total pressure measurements of the boundary layer at various locations on airfoil and flap for the baseline and optimum cylinder configuration with D∗ = 0 .0167 , r∗ = 0 .090 , θ = 37 ◦ ◦ and δf = 45 ...... 53

4.1 Visualization of the numerical domain...... 57 4.2 Mesh in the vicinity of the cylinder...... 58 + ◦ 4.3 ywall for the baseline model at δf = 45 ...... 58 4.4 Convergence of the steady state solution...... 59 4.5 Locations of the monitor points...... 60 4.6 Convergence of the velocity magnitude in the monitor points...... 61 ◦ 4.7 Pressure distributions for CFD and experiment at low flap deflection at δf = 15 . . . . . 62 ◦ 4.8 Pressure distributions for CFD and experiment at high flap deflection at δf = 45 . . . . . 63 ◦ 4.9 Pressure distributions for CFD and experiment at very high flap deflection at δf = 55 . . 64 4.10 Total domain with modelled tunnel walls on upper and lower side...... 64 4.11 Pressure distributions for the CFD simulation modelled with and without tunnel wall at ◦ δf = 45 ...... 65 ◦ 4.12 Designation of laminar and turbulent zones around the airfoil and flap at δf = 45 . . . . 66 4.13 Numerically calculated pressure distributions with/without laminar zones compared to ◦ experiment at δf = 45 ...... 66 4.14 Total pressure measurements in the boundary layer visualized for numerical simulations ◦ modelled with laminar zones and for fully turbulent simulations at δf = 45 . Measurements are performed at x/c = 0 .787 ...... 67 ◦ 4.15 Mean velocity field and streamlines for the baseline and cylinder configuration at δf = 55 . 69 4.16 Mean velocity field and streamlines in the slot for the baseline and cylinder configuration ◦ at δf = 55 ...... 70 ◦ 4.17 Mean pressure distribution for the baseline and cylinder configuration at δf = 55 . . . . 71 ◦ 4.18 Mean pressure vectorization for the baseline and cylinder configuration at δf = 55 . . . . 71 4.19 Mean skin friction coefficient for the baseline and cylinder configuration at the flap upper ◦ surface at δf = 55 ...... 72 ◦ 4.20 Vorticity magnitude at different time instances at very high flap angle. δf = 55 . . . . . 73 ◦ 4.21 Lift coefficient development in time for the baseline and cylinder configuration at δf = 55 . 74 ◦ 4.22 Velocity magnitude development in time at the monitor points at δf = 55 ...... 75 4.23 Visualizations of the line rakes at x = 0 .8965 c and at x = 0 .9365 c...... 75

MSc. Thesis xv D.P. Jansen List of Figures

◦ 4.24 Horizontal velocity component at x/c = 0 .8965 for very high flap angle at δf = 55 . . . . 76 ◦ 4.25 Horizontal velocity component at x/c = 0 .9365 for very high flap angle at δf = 55 . . . . 77

B.1 Exterior view of the wind tunnel test section...... 86 B.2 Interior view, looking down stream, of the wind tunnel test section. Clearly visible are the pitot tube and the wake rake that is positioned behind the model...... 86 B.3 Detailed view on the positioning of the cylinders. The mechanism supporting the cylinder is clearly visible as well as the zigzag tape that was used to remove the lower surface laminar separation bubble...... 87 B.4 Detailed view of the suction orifices at the wing/wall junction shown here around the airfoil upper surface...... 87 B.5 View of the zigzag positioning of the pressure holes on the airfoil and flap model surface. 88

◦ C.1 Raw data of pressure measurements for baseline run at δf = 55 ...... 92

D.P. Jansen xvi MSc. Thesis List of Tables

2.1 Selected cylinder sizes with corresponding F +...... 22

A.1 Test matrix for the vortex generators...... 83 A.2 Test matrix for the cylinders...... 84

xvii List of Tables

D.P. Jansen xviii MSc. Thesis Nomenclature

Abbreviations 2D 2-Dimensional 3D 3-Dimensional AC Alternating Current CCD Charge-Couple Device CFD Computational Fluid Dynamics CPU Central Processing Unit DBD Dielectric Barrier Discharge DES Detached Eddy Simulation EXP Experimental FCD Flow Control Device FFT Fast Fourier Transform LE Leading Edge LES Large Eddy Simulation NACA National Advisory Committee for Aeronautics PISO Pressure-Implicit with Splitting of Operators PIV Particle Image Velocimetry RAM Random-Access Memory RANS Reynolds Average Navier-Stokes SST Shear-Stress Transport TE Trailing Edge UV Ultraviolet

xix List of Tables

VG Vortex generator

Coefficients

Cd Drag coefficient

Cl Lift coefficient

Cl0 Lift coefficient at zero angle of attack

Clmax Maximum lift coefficient

Cn Normal coefficient cp Pressure coefficient

Greek Symbols α Angle of attack

αs Stall angle λ Spacing between two VG pairs µ Dynamic viscosity

σmon Convergence parameter

τw Wall shear stress β Angle between VG’s and flow direction β Curvature ω Specific dissipation rate of turbulent kinetic energy δ Boundary layer thickness

δf Flap deflection angle ∆ Change in variable

Latin Symbols c Airfoil chord D Cylinder diameter d Distance between FCD and separation point D∗ Cylinder diameter in units of c d∗ Distance between FCD and separation point in units of c f Frequency F + Reduced frequency h Height of a VG

D.P. Jansen xx MSc. Thesis List of Tables

h∗ Height of a vortex generator in units of c

L Sound pressure level

Lf Length of object over which flow separation is delayed

M∞ Free stream Mach number

P Pressure fluctuation in frequency domain p∞ Free stream pressure ps Pressure fluctuation pref Reference sound pressure pt Total pressure q∞ Free stream dynamic pressure r Cylinder distance from reference point r∗ Cylinder distance from reference point in units of c Re Reynolds number

Re y Turbulent Reynolds number S Distance between cylinder and wall s Spacing between two VG’s in a pair

St Strouhal number t Time t0 Reference time U Total velocity

U∞ Free stream velocity x Chordwise coordinate y Normal coordinate y+ Inner distance variable for a boundary layer

MSc. Thesis xxi D.P. Jansen List of Tables

D.P. Jansen xxii MSc. Thesis CHAPTER 1

Introduction

High lift aerodynamics has always been an important aspect of aircraft design. This is due to its major influence on aircraft performance in the essential flight phases of take-off and landing. Flap systems seen on aircraft today are primarily based on technology from the early 1940s, but studies in high lift design have gained a new interest. Although many of the currently used high lift devices consist of outdated complex multi-element flap systems, there is a growing trend in recent years for simpler, more cost effective designs. These mechanically simpler systems are still required to perform well in terms of aerodynamic performance. Increasing computational resources has given numerical approaches like CFD (Computational Fluid Dynamics) more possibilities to study the flow around high lift devices. This in combination with still developing experimental techniques in wind tunnel testing, like LDV (Laser Doppler Velocimetry) and PIV (Particle Image Velocimetry), has increased knowledge and performance of high lift aerodynamics to comply to the current demand of effective high lift systems.

1.1 The need for high lift systems

For several decades airplanes have been applied with high lift devices. The necessity of high lift systems lies in the fact that under certain circumstances the lift generated by the wings is not enough for the required aircraft performance. Although the wings produce enough lift for the aircraft to fly at cruise conditions, they do not suffice at take-off and landing. A lift increase is necessary in order to climb to the cruise altitude at take-off and to be able to fly at the necessary low manoeuvre speeds in case of landing. While a rise in lift for the wings is easily obtained by increasing the angle of attack, which is always done during take-off and landing, the necessary lift improvement to obtain maximum performance of the aircraft is unrealizable without the use of high lift systems. According to Meredith [1992] increasing the performance of high lift devices has the following possible benefits:

• A shorter take-off and landing field length is required.

• A higher take-off weight is possible.

• An increase in range can be accomplished.

1 Chapter 1. Introduction

Figure 1.1 – General high lift system with single slotted trailing edge and leading edge flap.

Cl

Airfoil Airfoil + trailing edge flap Airfoil + leading edge flap 0 α

Figure 1.2 – Lift polar showing the effects of a trailing edge and leading flap deflection on a plain airfoil.

• Steeper climb or approach angles are possible, enhancing safety during landing. This steeper climb or approach also reduces noise emission at ground level. • A reduction in stall speed can be achieved, leading to improved slow dive performance for emergency landings.

The most widely used high lift system is the configuration of an airfoil with a leading edge flap (slat) and at least one, often more, trailing edge flaps as is seen in figure 1.1. There are different types of both devices that have specific effects on the wing performance. In essence their goal is to increase the camber of the wing and to enlarge the wing surface area. These characteristics are realized by a flap and slat extraction and deflection, which both contribute to the gain in total lift of the wing. The lift gaining effects can best be visualized in a lift polar, which shows the dependency of the lift coefficient on the angle of attack. Figure 1.2 shows the lift polar for a plain airfoil and for an airfoil with trailing/leading edge flap.

The plain airfoil shows the well-known lift development with a linear relation between Cl and α at lower angle of attack, followed by the stagnation of lift at the stall angle until it drops at higher angle of attack. It is seen that the trailing edge flap deployment delivers a direct increase in lift when compared to the plain airfoil over the whole range of angle of attack. Similar to the plain airfoil lift development the lift increases proportional with higher angle of attack until a maximum is reached. It is seen that the stall angle at this lift maximum for the trailing edge flap deflection is lower than for the plain airfoil. Leading edge flaps do not have an immediate effect on the lift at low angles of attack, but due to an increase of the stall angle, realize higher maximum lift values.

D.P. Jansen 2 MSc. Thesis 1.2. Limits in high lift performance

Figure 1.3 – Circular-arc mean lines A, B, C obtained from the Joukowski transformation. Figure from Smith [1975].

Thus leading and trailing edge flaps indeed are beneficial in obtaining a higher lift, clarifying their use on current aircraft at take-off and landing. It is also clear that certain limits exist in lift performance, as is shown in figure 1.2 at high angle of attack. The working principles of high lift devices need further explanation, but first it is necessary to seek clarification concerning the observed limits that apply in lift performance.

1.2 Limits in high lift performance

To better understand what may be attainable in high lift performance it is necessary to have knowledge of theoretical maximum possible lift limits. Smith [1975] considered the flow around circular-arc mean lines in potential flow conditions. The circular-arc mean lines are shown in figure 1.3. Stagnation points are present at both ends of the arcs. By applying Joukowski airfoil theory to these arcs Smith [1975] found the following definitions for obtainable lift:

◦ ◦ Cl = 2 π (sin (α + β)) /cosβ (0 ≤ β ≤ 45 ) (1.1)

◦ ◦ Cl = 4 πsinβsin (α + β) (45 ≤ β ≤ 90 ) (1.2)

In the case of β = 90 ◦ the two stagnation points are so close together that a single stagnation point is formed. Figure 1.4 visualizes the flow under these conditions. It is this flow situation where the highest possible lift in theory is attained for which equation 1.2 gives Clmax = 4 π. In the case of an angle of attack of α = 45 ◦ and β = 45 ◦, which resembles quite closely the curvature of modern day flapped airfoils, the theoretical maximum is determined by equation 1.1 as Clmax = 8 .89 . This still seems rather high, since even aerodynamically advanced triple slotted flap systems have lift coefficients in the range of Clmax = 3 .5 − 4.0 at maximum, as identified by Torenbeek [1976]. So what is the reason that the theoretical values for Clmax are far from being achieved in practice? The answer lies in the assumption of potential flow in determining the theoretical limits. In potential flow viscous effects are not taken into account and boundary layers around objects are not modelled. In reality these boundary layers and viscous effects play a very important role in the behaviour of the flow around high lift systems. In the real world the appearance of boundary layers incorporate flow phenomena like transition and flow separation. It turns out that especially the aspect of flow separation has a dominant impact on the performance of high lift systems. Figure 1.5 graphically illustrates this phenomenon of flow separation at a highly deflected flap. This picture is obtained from PIV measurements on an airfoil and flap at zero incidence ◦ on a δf = 30 flap deflection performed by Little et al. [2009]. It is seen that flow is separated over the complete length of the flap and a drastic loss in lift efficiency is encountered. The general explanation for flow separation to occur is that the momentum of the flow in the

MSc. Thesis 3 D.P. Jansen Chapter 1. Introduction

Figure 1.4 – Flow around a cylinder with a very high circulation, such that the two stagnation points coincide. Figure from Smith [1975].

Figure 1.5 – Visualization of flow separation on a flap. Figure from Little et al. [2009]. vicinity of the wall is too low for the boundary layer to overcome the occurring adverse pressure gradient. The flap deflection of 30 ◦ tested by Little et al. [2009] is representative for deflections at which flow separation generally occurs for single slotted flaps. While equations 1.1 and 1.2 suggest that curvature β and angle of attack α can easily be increased to values of 45 ◦ for maximum lift, practice shows that this is not possible without the onset of flow separation and thus much lower lift limits apply. The flow separation has its unfavourable impact on the system shown in figure 1.5 on a high flap deflection, but it is also the reason for the occurrence of stall at high incidence angles as was observed earlier in figure 1.2. Also at stall of (flapped) airfoils high adverse pressure gradients appear with the onset of flow separation as a result. While the theoretical limits from potential flow are not representative for real world high lift applications, they do indicate that there is room for improvement in performance of current high lift devices in case phenomena like flow separation are tackled. In this thesis an investigation will be performed in how flow separation can be eliminated or postponed on an airfoil flap system with high flap deflection. For this a further understanding of the working principles of high lift devices is required together with the way flow separation affects the performance.

D.P. Jansen 4 MSc. Thesis 1.3. Flow separation in high lift performance

Figure 1.6 – Visualization of a laminar separation bubble.

1.3 Flow separation in high lift performance

The concept of flow separation in high lift performance has been studied from the early pioneering years in aviation up until today. Wright et al. [1924] obtained a patent on the design of a split flap that included a good qualitative understanding of the flow processes including flow separation. Although improvements in mathematical analysis, experimental methods and numerical simulations have been made ever since, modern aerodynamics still has difficulty accurately predicting flow separation. Flow separation does have a simple definition in mathematical terms. The occurrence of flow separation is often expressed through the definition of the wall shear stress, which has the form:

∂U τw = µ (1.3) ∂y y=0

Flow separation occurs if the velocity gradient at the wall is zero, thus when the wall shear stress is zero. The problem in localizing flow separation in numerical simulations is to accurately calculate this wall shear stress. The following section will further describe in more detail the known process of flow separation in the case for plain and flapped airfoils.

1.3.1 Plain airfoil characteristics

As explained earlier, increasing the angle of attack for an airfoil towards the stall angle for high lift performance will eventually cause flow separation to occur on the airfoil. For a plain airfoil Torenbeek [1976] states that there are multiple flow separation types encountered for stall: thin airfoil stall , leading edge stall , trailing edge stall or a combination of leading and trailing edge stall . The type of stall is mainly determined by the Reynolds number and the curvature of the nose, which affects another phenomenon called laminar to turbulent flow transition. The way this flow transition takes place at higher angle of attack determines the type of stall that will occur. Leading edge stall and thin airfoil stall are quite similar. At lower incidence angles, the high adverse pressure results in a flow transition where the laminar boundary layer separates and quickly reattaches as a turbulent boundary layer. This is called a short separation bubble, see figure 1.6 for a graphical illustration. For thin airfoils operating at higher incidence angles, the reattachment point gradually moves aft producing a long separation bubble. Ultimately, the boundary layers fail to reattach on the airfoil and flow separation has occurred as visualized in figure 1.7. For leading edge stall the short separation bubble also transforms into a long separation bubble, but in this case the change is more abrupt. Trailing edge stall occurs at higher Reynolds number and for thicker airfoils where the laminar to turbulent flow transition is more focused in a point, instead of a short or long separation bubble. In this case

MSc. Thesis 5 D.P. Jansen Chapter 1. Introduction

Figure 1.7 – Visualization of flow separation. separation is determined by the boundary layers’ ability to cope with the general pressure rise at the trailing edge, which leads to trailing edge separation at a higher angle of attack. Combined stall occurs at intermediate Reynolds number and leading edge thickness where both leading edge and trailing edge stall phenomena occur.

1.3.2 Flapped airfoil characteristics

For flapped airfoil systems high lift performance and the phenomenon of flow separation at stall is more complicated than for single airfoil configurations. This is because the multiple elements, which have their own flow characteristics, also influence each other. This affects the possible onset of flow separation, which can occur on both the flap and the airfoil. Over all, more parameters are involved in determining the high lift performance and flow separation characteristics.

Multi-element flow effects

To better understand how the parameters determine the high lift and flow separation characteristics of the high lift system, it is necessary to more accurately identify the working effects of the multiple elements. Like said, the elements influence each other. These effects are primarily pressure influences, which are well described by Smith [1975]:

• Slat effect : The circulation on a forward element reduces velocities on the downstream element, which reduce negative pressure peaks on its nose.

• Circulation effect : In turn, the circulation on the downstream element induces velocities on the upstream element, particular near the trailing edge, which increases its circulation.

• Dumping effect : The increased velocity at the trailing edge of the forward element relieves the upper surface pressure recovery impressed on the boundary layer, so alleviating separation problems.

• Off-the-surface pressure recovery effect : The boundary layer from forward elements is dumped at velocities appreciably higher than free stream. The final deceleration to free stream velocity is done in an efficient way; without this effect the boundary layer is unable to overcome the entire pressure rise. The deceleration of the wakes occur out of contact with a wall; this is usually more effective than the best possible deceleration with a wall.

• Fresh boundary layer effect : Each new element starts out with a fresh boundary layer at its leading edge. Thin boundary layers can withstand stronger adverse pressure gradients than thick ones. Hence, breaking up a boundary layer into several thinner boundary layers is favourable to the delay of separation.

D.P. Jansen 6 MSc. Thesis 1.3. Flow separation in high lift performance

Cl

◦ δf = 0 ◦ δf = 20 ◦ δf = 30 ◦ δf = 40

0 α

Figure 1.8 – Lift polar showing the general effect of increasing trailing edge flap deflection.

The earlier mentioned camber effect of the slat and flap deployment is a more general explanation of the detailed described effects stated above. These effects also explain the working principles of the slat and flap on the lift performance of a plain airfoil earlier shown in figure 1.2. Torenbeek [1976] elucidates that in general the deflection of a trailing edge flap results in an increased circulation, which induces an upwash at the nose. The adverse pressures increase and airfoils liable to leading edge stall will show flow separation at an angle of attack, which is below that of the plain airfoil. The working effect of a slat counteracts this effect by reducing the negative pressure at the nose, thereby increasing the stall angle and maximum obtainable lift.

Flow separation on a flap

Similar to the initiation of stall at high angle of attack for a plain airfoil, the deflection for a flap shows a behaviour where at too large angles the flow can separate. The trailing edge flap hereby acts kind of as a plain airfoil where dependent on the flap shape, multiple types of stall can occur. Figure 1.8 shows the general effect of multiple flap deflections for a trailing edge flap in a lift polar. It shows that initially the lift polar shifts up for higher flap deflections at the cost of a lower stall angle. However there is a deflection at which a further increase of the flap deflection results in flow separation and the lift will stagnate and even drop at even higher deflections. At this point flow separation is occurring at the flap even on low angles of attack. Besides the flap deflection there are several other design parameters that determine the effects described by Smith [1975] and the possible onset of flow separation. Flap chord lengths, gap sizes, slot shapes all affect flap performance, although for trailing edge flaps the most direct and important parameter remains the flap deflection. Numerous studies have been performed by Recant [1940] and Cahill [1949], to accurately identify the performance of the multiple design parameters. It was also discovered that multiple slots resulting in several flap elements were favourable in postponing flow separation compared to single element counterparts. This led to the development of multi-slotted flap systems designed in the 1960’s - 1970’s, where performance increase was found through all the effects described by Smith [1975], although these concepts were not even fully understood at that time. The multi-slotted flap systems realized higher flap deflections and were able to tackle flow separation. Typical values for triple

MSc. Thesis 7 D.P. Jansen Chapter 1. Introduction

Figure 1.9 – Triple-slot flap system of the B727-200. Figure from Rudolph [1996].

◦ ◦ slotted flaps were around δf = 50 deg to generally δf = 30 for single slotted flaps. This led to very sophisticated triple slotted high lift device systems being applied by Boeing on the B727, see figure 1.9, up to the B737 and eventually also the B747, at which the triple slotted inboard and outboard trailing edge flaps found their complexity peak. While lift gains, identified by Torenbeek [1976], of 40% compared to the single slotted counterparts were no exception, the development of multi-slotted flap systems also showed downsides. The deployment of multi-slotted flaps requires a mechanically complex system. This affects the development time and costs and puts high pressure on maintenance management. In addition, an aerodynamic downside is that the external hinges and actuators necessary for these systems generate parasitic drag when stowed in the cruise configuration. Of course there is also the direct weight penalty on the aircraft for such bulky systems. This led to a change of view that nowadays simpler systems are preferable as identified by Ashby [1996] and Rudolph [1996]. More knowledge of high lift performance and flow separation is obtained over the past decades with the growing trend of more accurate numerical simulations and experimental methods. This helped the development of new techniques in controlling flow separation and some of them already find their real life applications today. Although these methods are new, in general the goal remains the same as with the complex multi-slotted systems. That is to find ways to delay flow separation and thereby increase the possible flap deflection and improve high lift performance.

1.4 Flow separation control

For a couple of decades other approaches, besides the multi-slot flap systems, have been studied in order to control flow separation. The idea is to return to simpler single or double slotted flaps and incorporate these new flow control techniques. These approaches can be divided into two types of flow separation control techniques: active flow separation control and passive flow separation control. Active flow separation control techniques are based on putting energy into the flow, while passive control techniques do not induce energy in the system. In practice, active flow separation control can lead to higher lift performance improvements compared to passive techniques, often at the cost of increased complexity of the system. Both flow control techniques are based on either directly increasing the momentum in the boundary layer or by creating vortices transporting higher momentum free stream flow to within the boundary layer. Increasing the momentum of a boundary layer will generally increase the ability to overcome the adverse pressure gradient.

1.4.1 Active flow separation control techniques

Cattafesta [2011] classifies the active flow control techniques into the following solutions: fluidic , moving object/surface , plasma and others . This section will briefly describe each of these control techniques.

D.P. Jansen 8 MSc. Thesis 1.4. Flow separation control

(a) Baseline situation. (b) Situation with sinusoidal actuation of syn- thetic jets.

Figure 1.10 – PIV visualization of flow separation postponement on a wind turbine blade by applying ◦ synthetic jets at α = 16 and Re = 1 .6 · 10 5. Figure from Maldonado et al. [2010].

Figure 1.11 – Components of an electrodynamic synthetic jet. Figure from Cattafesta [2011].

Fluidic actuators

Fluidic actuators use fluid injection or suction to obtain a certain amount of control on the flow. Although many subclasses exist, the two most commonly used fluidic actuators are synthetic jets and boundary layer suction/blowing . Figure 1.11 shows a schematic of an electrodynamic synthetic jet configuration. Synthetic jets are based on alternately ingesting and expelling fluid into the flow to create vortices and a higher momentum boundary layer. For the actuator shown in figure 1.11 this is done by a diaphragm, which will oscillate under influence of electrodynamic transduction. The magnet generates a magnetic field with a magnetic flux density, which under the influence of an alternating current (AC-current) in the wound coils, results in an alternating force induced on the coils. This causes the diaphragm to oscillate, which results in fluid flowing in and out the cavity through an orifice or slot leading to vortices in the boundary layer. Maldonado et al. [2010] performed PIV measurements on wind turbine blades equipped with synthetic jets at high angle of attack. Significant improvements in CLmax were found in the order of 12% at a delayed stall angle of 2◦ compared to the baseline configuration. The delay in flow separation on the turbine blade is illustrated in figure 1.10. Boundary layer suction/blowing is another way to increase the momentum in the boundary layer. With boundary layer suction this is done by removing the low momentum flow in the vicinity of the wall, where usually the fluid is expelled at another location. Boundary layer blowing directly increases the momentum of the boundary layer and can even be done oscillatory to add vortices to the flow as well. Figure 1.12 graphically explains the working principle of boundary layer suction.

MSc. Thesis 9 D.P. Jansen Chapter 1. Introduction

Figure 1.12 – Working principle of boundary layer suction. Figure from Boermans [2008].

Figure 1.13 – Schematization of a piezoelectric flap. Figure from Veldhuis and van der Jagt [2010].

Moving object/surface actuators

There are several types of moving object/surface actuators, but the most common is the piezoelectric flap. An AC-voltage across the piezoelectric device causes the flipping motion of the flap, which then interacts with the flow. In this way vortices can be created in different sizes and direction depending on the geometry and orientation with respect to the local free stream flow. A typical example of a cavity-type piezoelectric flap is shown in figure 1.13.

Plasma actuators

Plasma actuators come in different forms, having slightly different techniques to obtain the plasma. The most popular variant is the Dielectric Barrier Discharge (DBD) plasma actuator, which consists of two electrodes that are separated by a dielectric material. The air passing the electrodes becomes ionized by the voltage that is applied. The ionized air, now called plasma, produces forces in the air due to the attained electric field gradients by the electrodes. These forces can induce velocity components to the flow and thus can be used for effective flow control including delay of flow separation. Figure 1.14 shows the general setup of a DBD plasma actuator. The general advantages of plasma actuators with other active control techniques is that they are easily applicable and very compact, which gives a lot of design freedom. The main evident disadvantage is the amount of energy necessary to supply a high voltage to the electrodes.

1.4.2 Passive flow separation control techniques

There are various forms of passive flow separation control techniques. Unlike active flow control no direct energy is brought into the system, meaning that passive flow control is solely based on mixing

D.P. Jansen 10 MSc. Thesis 1.4. Flow separation control

Figure 1.14 – Schematization of a DBD plasma actuator. Figure from LeBeau [2007].

(a) Counter-rotating (b) Counter-rotating (c) Co-rotating. (d) Multiple-row. “common flow “common flow up”. down”.

Figure 1.15 – Vortex generators displayed in different setups. Figures from von Stillfried et al. [2010]. high momentum fluid to areas of low momentum, hence to boundary layers which are on the verge of separation.

Vortex generators

Vortex generators (VG’s) are the most commonly known passive control devices and are already used in different industries. Although VG’s come in various shapes and sizes, in general a vortex generator is build up as a small vertical plate positioned at an angle with respect to the local free stream flow. With appropriate dimensioning and positioning stream wise vortices can be created which can be used to control the flow. Vortex generators can be classified in different ways. First there exist co-rotating and counter-rotating types, depicted in figure 1.15, where in general better results are obtained with the latter. Then there also exists a segmentation in terms of scale, hence conventional boundary layer VG’s and more recently the development of sub-boundary layer VG’s , also called micro vortex generators. Research in sub-boundary layer VG’s as performed by Lin [1999] show significant improvements in reducing flow separation comparable to their larger conventional counterparts without the effects of increased drag. The counter-rotating micro VG’s led to lift increases up to 10% on a leading and trailing edge flapped airfoil at accompanying drag reductions of 50% . This was achieved by a delay in the separation location from approximately 45% flap chord to at least 85% flap chord. Figure 1.16 shows the results in flow separation delay by means of measured anemometer RMS values tested for two Reyonolds numbers.

In general the interest in micro VG’s and creation of small-scale perturbations is quite similar to the trend found in the field of active flow control techniques. Micro VG’s as other small-scale flow control solutions have a small bandwidth and the location of separation needs be somewhat fixed in order for the micro VG’s to be effective.

MSc. Thesis 11 D.P. Jansen Chapter 1. Introduction

Figure 1.16 – Effect of counter-rotating micro VG’s on RMS anemometer output from hot-film sensors on ◦ flap upper surface with α = 8 and M∞ = 0 .2. Figure from Lin [1999].

Figure 1.17 – Lift-enhancing tabs as used by Lee et al. [2005].

Lift-enhancing tabs

Lift-enhancing tabs or Gurney flaps are small ’plates’, which are located generally at trailing edges of lift generating devices. A typical Gurney flap is several boundary layers in height and is usually positioned perpendicular to the flow. Figure 1.17 shows the working principle of lift-enhancing tabs. Two counter rotating vortices are created aft of the Gurney flap, entraining the flow from the airfoil upper surface around the top recirculation region. Hence the flow stays attached over the airfoil flap surface and separation can be delayed. Ashby [1996] found promising results, where various tabs on both the main airfoil and trailing edge flap were tested. Lift increases up to 11% were found at some angle of attack, while the maximum lift coefficient gained 3% with respect to the baseline configuration. Even more active and passive flow separation control techniques exist and in general it can be said that very interesting results are obtained with various designs. In general active flow control techniques share the main similar disadvantage. Whether it is applying an AC-current or voltage as for the synthetic jets and plasma actuators, or fluid being blown into the system by a pump with boundary layer blowing, significant amount of energy needs to be put into the system. Often problems are encountered with the practical implementation of these mechanisms. Part of these problems are alleviated by a migration of the active flow control research field in terms of approach, as discussed by Cattafesta [2011]. Increased flow physics understanding changed the concept from inducing large vortices to the system to focus on creating smaller instabilities to the flow. This makes it feasible to reduce power, size and mass. However for high-speed applications these ’small-scale’ devices currently still lack bandwidth and often have control related issues. With passive flow control, designs are often easier applicable and design aspects like size,

D.P. Jansen 12 MSc. Thesis 1.5. Opportunity of research

mass and maintenance are concept wise less an issue. Also facets like costs and safety are considered to be less problematic for passive flow control solutions. This is the reason why passive techniques like vortex generators can already be found on various recent aircraft designs. This makes passive control very if not more interesting for the near future.

1.5 Opportunity of research

In section 1.4 a variety of methods in flow separation control has been discussed including obtained results by other research. It is identified that for the near future passive flow control can give many solutions. One of the solutions not discussed up to this point is a passive method, researched by van der Steen [2009], where cylinders were used on a simplified airfoil flap system to prevent separation occurring over the flap at high flap deflections. Research shows that under specific conditions, the flow past cylinders can generate perpendicular vortices to the free stream, more generally known as the Von-Kármán vortex street. van der Steen [2009] has shown that these vortices can delay flow separation and can be more effective then conventional vortex generators. This thesis will continue this research by utilizing these passive flow separation techniques on an existing high lift device system being the airfoil flap sytem of the Extra EA-400. As was pointed out in section 1.3 the interest lies in applying the separation control on high flap deflections where flow separation is eminent. More specifically the goal of this thesis is: To investigate the performance improvements of cylinders and vortex generators on the Extra EA-400 airfoil flap system at high flap deflections by experiment and numerical simulations . The Extra EA-400 is a six-seat single-engine monoplane produced by Extra Flugzeugbau GmbH in collaboration with Delft University of Technology, which developed the aerodynamic design of the wing. For the airfoil flap system of this aircraft, Delft University of Technology has a 2D prismatic wind tunnel model. Hence, the study is a 2D investigation that will simplify the analysis compared to a 3D wing especially for the numerical simulation. For a full 3D wing analysis, lateral effects like wing tip vortices and wing fuselage interactions will significantly increase the complexity of the flow structures. These 3D effects are not studied in this thesis.

1.6 Layout of the report

The layout of the report will be as follows. In chapter 2 the setup of the experiment will be discussed which includes an overview of the wind tunnel, apparatus and the model. Further the passive flow control devices working principles and measurement techniques will also be described. Chapter 3 will give the results and discussion from the experiment. Then in chapter 4 the setup and results for the numerical simulation will be discussed. Finally the conclusion and recommendations for this research will be given in chapter 5.

MSc. Thesis 13 D.P. Jansen Chapter 1. Introduction

D.P. Jansen 14 MSc. Thesis CHAPTER 2

Experimental model setup

2.1 Introduction

This chapter will describe the experimental investigation performed on the airfoil flap model of the Extra EA-400. This model is based on the aerodynamic profile NLF-MOD22B, designed by the Delft University of Technology. In section 2.2 a brief description is given about the wind tunnel and the model. This is followed by an explanation and some preliminary calculations for the applied passive lift devices in section 2.3. Finally section 2.4 gives an overview of the applied experimental techniques and measurement systems including an explanation on the required data.

2.2 The wind tunnel and model

The wind tunnel

The wind tunnel used is the Low Speed Low Turbulence Wind Tunnel of Delft University of Technology. It is a closed vertical arranged circuit wind tunnel, operating at tunnel speeds of 0 − 120 m/s with turbulence levels below 0.05% . The test section has dimensions of 1.80 × 1.25 m and a length of 2.4 m. An electric motor drives the flow through the diffuser towards the settling chamber, after which the air is accelerated towards the test section. Rotatable side plates are installed in the test section to give freedom in adjusting the angle of attack. Figure 2.1 shows the complete layout of the wind tunnel.

The model

The NLF-MOD22B model used for the experimental and numerical research is an airfoil with a single slotted trailing edge flap system. The NLF-MOD22B airfoil flap system is schematically depicted in figure 2.2 for its retracted and deployed state respectively, the latter also showing the geometric details of the flap setting.

15 Chapter 2. Experimental model setup

Figure 2.1 – The Low Speed Low Turbulence Wind Tunnel at Delft University of Technology.

Figure 2.3 shows the composite wind tunnel model, which was installed in a vertical position between the two rotatable circular endplates. The model has a chord of 0.6m and span of 1.25 m. The 30% chord flap can be deployed through a translation and rotation of the turntables positioned at the flap endplates.

From earlier research on the NLF-MOD22B flap system performed by Boermans and Rutten [1995] it was found that artificial tripping of the boundary layer on the airfoil lower surface was necessary in order to eliminate a detrimental laminar separation bubble on the flap lower surface. This was also observed in preliminary wind tunnel tests of this research. For this reason zigzag tape with a thickness of 0.5 mm is placed at 58% chord length. The thickness of the zigzag tape was determined by the local height of the boundary layer. The boundary layer thickness was measured through a vertical pitot tube traverse. Figure 2.4 shows the performed measurements including the boundary layer state before and after the application of the zigzag tape.

Figure 2.2 – Cross section of the model and definition of flap parameters.

D.P. Jansen 16 MSc. Thesis 2.2. The wind tunnel and model

Figure 2.3 – The model positioned in the wind tunnel with the cylinder test setup.

MSc. Thesis 17 D.P. Jansen Chapter 2. Experimental model setup

6

5

4

3 [mm] y

2

◦ 1 α = 0 ◦ α = 12◦ α = 0 zigzag at x/c = 0 .58 ◦ α = 12 zigzag at x/c = 0 .58 0 0 300 600 900 1200 1500 ptot

Figure 2.4 – Boundary layer measurements on model lower side at x/c = 0 .67 .

It should be noted that the presence of the wind tunnel walls clearly affects the 2D state of the flow through the possibility of locally early separation. This was avoided by applying boundary layer suction at the rotatable endplates through 5 mm diameter holes, spaced 10 mm apart and from the main wing respectively, and through 4 mm diameter holes, spaced 8 mm apart and from the flap. These suction orifices are displayed in the picture of figure B.4 in appendix B. Tufts were applied to the model in preliminary experiments to test the necessary suction level to prevent the occurrence of flow separation at the wing/wall junction. In these preliminary tests suction level was increased up to the point where tufts near the wing/wall junction showed steady behaviour with an orientation in line with the general flow direction. The flap is positioned at a fixed overlap and gap size of respectively 0.0c and 0.035 c, which produces the highest lift according to measurements from Boermans and Rutten [1995]. In order to ensure flow separation on the flap to test the vortex generating devices on their performance, the flap needs to be positioned at a large deflection. This means that it is necessary to investigate at which deflection angle flow separation starts to occur. Boermans and Rutten [1995] performed studies on the pressure distributions for several flap deflections and concluded that flow separation starts to occur at the flap for a flap deflection of 40 ◦. From these results the flap deflections at which the model will be tested are set at 45 ◦, 50 ◦ and 55 ◦.

2.3 Passive flow separation methods

The airfoil flap configuration model as discussed in section 2.2 will serve as the benchmark setup for the wind tunnel investigation. This section will describe the passive separation control devices tested in this study.

D.P. Jansen 18 MSc. Thesis 2.3. Passive flow separation methods

Figure 2.5 – The model with VG’s applied (left) and cylinders (right).

2.3.1 Vortex generators

The vortex generators were tested at the lower side of the airfoil near the local trailing edge, as depicted in figure 2.5. While VG’s are more often found on upper surfaces of wings, downstream of the leading edge, the idea here is to position the devices upstream of the flap leading edge. As discussed later in this section, vortices need a specific distance to develop and it takes time for the flow to mix with the entrained high-energy flow. There is a certain distance from the separation point at which vortex generators become effective. If positioned on the upper side of the airfoil it becomes difficult to reach the flap boundary layer with vortices. Positioning on the flap itself is difficult because of the short distance with respect to the point of separation. From the lower side of the airfoil this can be done more effectively since the vortices have enough time develop. This makes the positioning of VG’s at the end of the lower side of the airfoil an interesting study case. Figure 2.5 depicts the conceptual idea that streamlines that come from the lower side of the airfoil and continue close over the flap upper surface are entrained with a vortex shedding to delay separation over the flap. Correct dimensioning and positioning of the VG’s should give an optimal configuration, at which the necessary streamlines are entrained with vortex shedding. As explained in section 2.3.1 different types of vortex generators exist. For this research only the counter rotating VG’s were tested at several sizes. Figure 2.6 shows the geometrical definitions of VG’s, which include:

• λ for the mutual spacing of the vortex generators. Here λ is defined as the distance between the VG pairs.

• h to define the height of the vortex generator.

• d for the distance between the VG’s and the bottom trailing edge of the NLF-MOD22B model.

The vortex generators were all tested on different values for these parameters. For the remaining param- eters in figure 2.6, fixed values were chosen of β = 18 ◦ and s = h respectively, for which van der Steen [2009] found the best results. The height of the vortex generators was difficult to determine. In van der Steen [2009] VG heights ranging 1 from h = 3 δ (micro VG’s) to h = δ proved to be successful, however these devices were applied for quite a different flow situation, an inclined flat plate. Furthermore micro VG’s should be positioned close to the point of separation to be effective, which is unrealizable in this case when the VG’s are positioned at the lower end of the airfoil. It was expected that slightly larger VG’s were necessary in order for the vortices to remain their strength through the slot and be effective over the flap. Figure 2.4 shows that the boundary layer height at location of the VG’s is 6 mm after the transition tape is applied. From this

MSc. Thesis 19 D.P. Jansen Chapter 2. Experimental model setup

Figure 2.6 – Vortex generator geometrical definitions.

Figure 2.7 – Vortex generator geometry and construction method. result VG heights were chosen of 6 mm , 14 mm and 28 mm , which were tested at multiple spacing’s λ. The complete test matrix for the VG’s is given in table A.1 in appendix A. Note that in the test matrix and in shown results in following sections the non-dimensionalized parameters h∗ and d∗ are used. Here h∗ and d∗ are defined as h∗ = h/c and d∗ = d/c . The VG’s were produced from a 0.1 mm aluminium sheet. A construction method was applied similar to van der Steen [2009]. Small rectangular plates were bend over their diagonal until the two created planes are perpendicular, creating the delta shaped vortex generator shown in figure 2.7. The devices were connected to the model through double sided tape, which made it easy to remove and test for different configurations.

2.3.2 Cylinders

The cylinders were tested in the slot just in front of the trailing edge flap. Similar to the vortex generators, the concept is to entrain the streamlines close over the flap surface with a vortex shedding developing in the slot. This is depicted in figure 2.5. As well as for the vortex generators positioning and size are of key importance for the cylinder performance. Figures 2.8 and 2.9 show the parameter definition for the positioning and size of the cylinders. The positioning is expressed in the cylindrical coordinates r and θ. The diameter is specified as D. The origin is chosen to be at the hinge of the supports, at [x, y ] = [0 .72 c, 0.06 c] measured from the airfoil nose; The connection system will be explained later in this section. The area of interest is also depicted in figure 2.8, which resulted in multiple selections of r, θ and D to test the cylinders. The complete test matrix for the cylinders is given in table A.2 in appendix A. Note that in the test matrix and in shown results in following sections the non-dimensionalized parameters r∗ and D∗ are used. Here r∗ and D∗ are defined as r∗ = r/c and D∗ = D/c . It must be noted that the circulation of the Von Kármán vortex street is lateral to the flow direction instead of lengthwise, which is the case for the VG devices. Unlike for the vortex generators some theoretical basic equations exist in order to estimate the correct cylinder size beforehand. The shedding frequency of a vortex street behind a circular object can be determined with the following equation:

D.P. Jansen 20 MSc. Thesis 2.3. Passive flow separation methods

Figure 2.8 – Cylinder position and definition of coordinates. Hinge point is referenced to the airfoil nose with [x, y ] = [0 .72 c, 0.06 c].

Figure 2.9 – Cylinder geometrical definitions.

MSc. Thesis 21 D.P. Jansen Chapter 2. Experimental model setup

D [mm] F + Cylinder I 10 3.6 Cylinder II 15 2.4 Cylinder III 20 1.8

Table 2.1 – Selected cylinder sizes with corresponding F +.

St · U∞ f = (2.1) D

Here St is the Strouhal number, a dimensionless number that describes oscillating flow mechanisms, and U∞ is the free stream velocity of the fluid. The Strouhal number for cylinders is 0.2 for Reynolds numbers from 100 to 10 5 as given by White [2006]. It should be noted that equation 2.1 applies for free stream flow. In the vicinity of a wall, research shows that the shedding of vortices by the cylinder is affected by the distance from the wall S, which is clarified in figure 2.9. This was investigated by Wang and Tan [2007], which tested a cylinder wake close to a fully developed turbulent boundary layer with thickness δ = 0 .4D. The research showed that vortex shedding takes place as in free stream when S/D > 0.8. For S/D < 0.8 the vortex shedding is asymmetrical with stretched and underdeveloped vortices from the near wall side. At ratios of S/D < 0.3 the vortex shedding can even be completely suppressed. The experiment further showed the Strouhal number remains constant at about 0.2, but that the state of the turbulent boundary layer could significantly affect flow characteristics at different S/D . Besides taking into account the influence of the wall, or in this specific investigation the influence of the slot boundaries and flap, the question remains at which shedding frequency the momentum exchange in the boundary layer is optimal. Nishri and Wygnanski [1998] and Greenblatt and Wygnanski [2000] performed studies on flow separation control through periodic excitation and identified that the vortex frequency in a different form is more useful. They introduced the reduced frequency F +:

fL F + = f (2.2) U∞

This expression is a dimensionless form of the vortex frequency f. In equation 2.2 Lf is by definition the length of the object over which flow separation must be delayed. Combining equation 2.2 with 2.1 leads to the interesting result:

L F + = St f (2.3) D

This means that the reduced frequency is independent of velocity, considering St is independent of Reynolds number between 100 and 10 5. Greenblatt and Wygnanski [2000] found an optimum reduced frequency for forcing a separated flow to reattach at F + = 1 . For maintaining attached flow the forcing frequency should be between 3 < F + < 4. These results were used as a basis for estimating the cylinder sizes that have been tested. When for Lf the length of the flap is chosen, the reduced frequency can be plotted as a function of the cylinder diameter, see figure 2.10. From a theoretical point of view it is preferable to test a cylinder size corresponding to F + = 1 , but this would result in cylinder sizes of 40 mm and is considered impractical given the limited size of the slot. Cylinder diameters of 10 mm , 15 mm and 20 mm were selected, which have reduced frequencies in the range of 1.8 < F + < 3.6. The diameters and corresponding reduced frequencies are summarized in table 2.1. The cylinders were mounted at the required position through four fixed supports with adjustable can- tilevers. The cylinders crossed the total span of the model so that the concept of 2D flow is maintained.

D.P. Jansen 22 MSc. Thesis 2.4. Measurement techniques

5

4

3 + F

2

1

0 0 10 20 30 40 50 D

Figure 2.10 – Reduced frequency dependency on cylinder diameter.

Positioning of the construction was done with care. The brackets need to be oriented far enough from the pressure orifices to avoid flow interference effects, keeping in mind that a too large support spacing may cause insufficient bending stiffness for the construction and possibly undesirable vibrations of the cylinders. The cantilevers were able to rotate and, through the multiple positioning holes, have freedom of movement in horizontal and vertical direction. In appendix B pictures are provided to see the mechanism in more detail.

2.4 Measurement techniques

There are several experimental methods to determine airfoil flap configuration performance and to analyse the effect of vortex generating devices, of which the latter is particularly important for this study. For this research pressure measurements, oil flow visualization and frequency measurements were used. The next section will shortly explain and discuss these techniques.

2.4.1 Pressure measurements and calculation of Cl and Cd

Multiple pressure measurement methods were applied which can give direct insight in local flow properties, like flow transition and flow separation. The pressure measurements can also be used to calculate certain performance parameters. These parameters are the dimensionless coefficients Cl and Cd, which represent the lift and drag. All the obtained pressure quantities were measured by tubing pressure orifices to a digital multi-manometer from Pressure Systems, which is able to directly transform the pressure offsets to discrete values. On the upper and bottom side of both the airfoil and flap a total of 85 pressure orifices were located, where for each pressure hole measurements are performed at a frequency of 330 Hz in an interval of 4.0 s. The pressure holes are located on the model in a zigzag position that will prevent the pressure holes to locally disturb each other, see figure B.5 in appendix B. The pressure measurements can be used to accurately calculate the lift in the following way. First the measured static pressures at the pressure orifices are translated to dimensionless pressure coefficients through:

MSc. Thesis 23 D.P. Jansen Chapter 2. Experimental model setup

p − p∞ Cp = (2.4) q∞

Here p is the local static pressure at the orifice and p∞ and q∞ are respectively the far field total and dynamic pressure. The latter two are obtained by measurements with a pitot static tube. This tube is positioned in the plane of the mid-span section on the tunnel sidewall opposite to the model lower surface and about two chord lengths in front of the leading edge. This is visualized in the picture of figure B.2. The pressure distribution is integrated through the trapezium rule to obtain a normal coefficient for both the airfoil and flap with:

T E airfoil x C = C − C d (2.5) nairfoil ˆ plower pupper c LE airfoil ! "

T E flap x C = C − C d (2.6) nflap ˆ plower pupper c LE flap ! "

This gives a total normal coefficient of:

Cn = Cnairfoil + Cnflap (2.7)

This can be used to calculate the lift using:

Cl = Cn/cos (α) − Cd · tan (α) (2.8)

To calculate the lift coefficient from equation 2.8, a drag coefficient is necessary. For the drag calculation another pressure measurement approach is used, the so-called wake-traverse method. A total pressure and static pressure wake rake is positioned behind the model. The dimensions of the wake rake are shown in figure 2.11. The positioning of the wake rake in the tunnel is seen in the picture of figure B.2 in appendix B. The pitot tubes are positioned horizontally, although ideally they should be aligned with the local flow behind the airfoil flap configuration that is deflected downwards by the downwash of the wing. However it was very unpractical to adjust the orientation of the wake rake for each observation. Still the total pressure tubes are allowed to have an angle misalignment of 40 ◦ which was assumed to be sufficient for the measurements. The wake rake pressure measurements are used to determine the profile drag of the model using the method proposed by Jones [1936]. The equation used in this method to determine the drag is given by:

pt − p pt − p∞ y Cd = 2 1 − d (2.9) ˆ ò q∞ 3 ò q∞ 4 1 c 2

This gives the necessary information to calculate the lift coefficient. An important issue remains in the computation of the previously mentioned performance parameters. Since the wind tunnel experiment needs to provide 2D results, some corrections need to be applied for the presence of the tunnel wall. Apart from the local influence on the boundary layer flow separation as

D.P. Jansen 24 MSc. Thesis 2.4. Measurement techniques

(a) Front view.

(b) Top view.

Figure 2.11 – Test setup of the wake rake. discussed in section 2.2, the wall also has two other effects, namely solid blockage and wake blockage. The solid blockage effect is caused by an increase in the free stream velocity attributed to the volume distribution of the body. The wake blockage effect is a serious constraint on the growth of the wake. For streamlined bodies at small angles these effects are small, but especially for the wake blockage become of significant performance for bluff bodies. This includes airfoil flap configurations at high flap deflections and high angle of attack. Allen and Vincenti [1944] proposed a correction model for this situation, which is based on a wall representation of image doublets and sources. It is applied to the pressure distribution computation and also for the lift and drag coefficient calculation in order to get the required 2D results.

All measured (un)corrected pressures and tunnel conditions were collected by Labview from which the aerodynamic coefficients are calculated in Profmeasure. Appendix C shows the output file from Profmea- sure for such a single measurement, which is then read into MATLAB for further postprocessing.

2.4.2 Frequency measurements

A mapping of the frequency was created of the produced vortices aft of the cylinder. A listening tube was connected to an external storage system and gave the possibility to analyse results on screen in Labview. This gave a recorded sampled signal of the pressure fluctuation ∆ps against time in the vicinity of the cylinder. This however is a noisy signal, which is difficult to analyse for measured frequencies. Therefore the signal was transformed through a Fast Fourier Transform (FFT) in MATLAB to the frequency domain. The FFT transformation applied here is given by the following equation:

N −1 2πi (k−1) j ∆P (k) = ∆ps (j) e N (2.10) Øj=1

The measurements were performed at multiple locations to gain insight on the development of a possible vortex street. These locations are visualized in figure 3.22.

MSc. Thesis 25 D.P. Jansen Chapter 2. Experimental model setup

2.4.3 Oil flow visualization

Oil visualization gives direct qualitative insight in the characteristics of the flow. It can confirm observa- tions made from analyzing data of quantitative measurements. When applied properly it can be utilized for detecting both transition as well as flow separation. For the visualization fluid a mixture of oil with a fluorescent dye was prepared. For the respectively low tunnel speed, a thinner was added to make sure that the fluid viscosity was not too high and that the oil would settle quickly for each measurement. A UV light was used for enhanced visibility with photography. The images were recorded with a Nikon D80 CCD camera at high aperture to ensure sufficient light exposure.

D.P. Jansen 26 MSc. Thesis CHAPTER 3

Experimental results

3.1 Introduction

This chapter will give an overview of the results obtained from the experimental investigation. Initially the NLF-MOD22B airfoil flap system was tested for its baseline configuration. From the experiments performed by Boermans and Rutten [1995] it was identified that flow separation occurs at the flap at a deflection of 40 ◦. Based on this result, the choice was made to test the flow separation delay capabilities of the passive lift devices on the model for flap angles of respectively 45 ◦, 50 ◦ and 55 ◦. All tests in the experimental investigation are performed on a tunnel speed of 50 m/s corresponding to a Reynolds number of Re = 2 .0 · 10 6. An overview of the results from the baseline run tests is presented in section 3.2. Then section 3.3 and 3.4 will respectively show the effects of the performance of the passive high lift devices with respect to the baseline run results at the selected flap deflections. Finally in section 3.5 conclusions will be drawn based on the experimental investigation.

3.2 Baseline model

The baseline model is defined as being solely the airfoil and flap with one of the three deflections 45 ◦, 50 ◦ and 55 ◦. As discussed in section 2.4 the wind tunnel experiments consisted of pressure measurements, oil flow visualization and frequency measurements. In this section the pressure measurements and resulting lift polars of the baseline models will be discussed to understand the effects of the angle of attack and the three flap deflections on the model performance. Discussion of the oil flow visualization and frequency measurements for the baseline models is done in section 3.4 where the baseline is directly compared to the similar tests done for the cylinder configuration.

Baseline lift polars

In figure 3.1 the lift polars are shown for the three tested flap deflections. In general the lift coefficient shows similar behaviour for all three cases. For all cases the angle of attack is increased from −6◦ to

27 Chapter 3. Experimental results

13 ◦. Figure 3.1 shows that the lift increases with the angle of attack until the stall angle is reached and ◦ abruptly drops. It is clear that the lift curve for δf = 45 shows the largest obtainable lift at all angles ◦ of attack with a maximum of Cl = 3 .23 at α = 12 . Although at lower angles of attack the flap with ◦ δf = 50 can show comparable lift performance, it is noticeable that the lift starts to stagnate in this case at α = 8 ◦. This is an unexpected result for which a supplementary run was performed. However, no differences in the lift development between the subsequent runs were measured. At a flap deflection of ◦ δf = 55 it is clear that the lift is significantly lower over the complete range of angle of attack. Further ◦ results will show that at flap angles larger than δf = 45 the lift becomes highly affected by the onset of flow separation. A remark should be made on the tests performed for the flap deflection of 55 ◦. Early wind tunnel tests revealed that the a.o.a. had to be lowered first to negative angles in order to have a semi-attached flow on the flap. This meant that if the model was directly set at α = 0 ◦ at the start of the run a completely separated flow on the flap was observed. It is a rather peculiar phenomenon and is also reported by Baragona [2004] and Biber [1995] on other single slotted airfoil flap researches. This insinuates the presence of a stall hysteresis caused by laminar bubble bursting on the flap. The observed stall hysteresis was kept in mind, but no further research was done on the hysteresis by means of stall recovery tests. The laminar bubble bursting may also be the cause the stagnation in lift found at high angle of attack for the flap deflection of 50 ◦. In fact Baragona [2004] discovered in the same setup at which hysteresis for α = 0 ◦ was observed, changes in the slot geometry could cause laminar bubble bursting and accompanying lift stagnations at high angle of attack. The tests performed by Baragona [2004] and also Biber [1995] seem to resemble the early observations in the lift polars of this wind tunnel research for the NLF-MOD22B airfoil flap. The next section will further elaborate on the observations made from the lift polars by discussing the accompanying pressure distributions.

Baseline pressure distributions

Figure 3.2 shows the pressure distributions for the baseline configuration at flap deflections 45 ◦, 50 ◦ and 55 ◦ for multiple angles of attack. Overall similar behaviour can be observed in the pressure distribution development. For all deflections a large negative pressure peak develops on the leading edge of the main airfoil, while the pressure distributions on the flap show more variety in behaviour. Focusing on the flap pressures it is clear that flow separation on the flap upper surface is visible at all deflections. Especially at 55 ◦ the sudden increase in pressure about halfway the flap chord and the high steady pressure toward the trailing edge, clearly indicates the onset of flow separation. The flow separation is caused by the very high adverse pressure gradient the flow needs to overcome, due to the very low peak pressures at the flap leading edge. Similar behaviour is found at 50 ◦ and 45 ◦ though the adverse pressure gradients are less severe and the point of separation shifts more towards the trailing edge. For the dependency on the angle of attack the following observations can be made. With increasing angle of attack, the stagnation point on the airfoil moves rearward along the lower surface and the suction peak at the nose of the airfoils upper surface develops significantly resulting in a lower pressure at the upper surface and a higher lift Clairfoil for the main element. From figure 3.1 it is visible that stall is reached ◦ at an angle of αs = 13 for all the tested flap deflections. Figures 3.2a - 3.2c show that at stall, flow separation is now found on the airfoil due to the severe adverse pressure gradient at this high angle of attack. The stall on the airfoil is of a leading edge type, due to its sudden appearance and the onset of separation close to the leading edge. In general the pressure distributions on the flap show less severe changes when the angle of attack is increased. Still, a tendency can be observed that for higher incidence angles the pressure on the upper

flap surface in fact increases and smaller lift coefficients Clflap are created. This effect is more often found for airfoil flap configurations for example by Boermans [1998] on an experimental research for a NACA 63-415 airfoil with slotted flap. At higher angle of attack the pressure distribution on the flap

D.P. Jansen 28 MSc. Thesis 3.2. Baseline model

3.5 ◦ Baseline, δf = 55 ◦ Baseline, δf = 50 ◦ Baseline, δf = 45 3

2.5 l C

2

1.5

1 −10 −5 0 5 10 15 ◦ α [ ] (a) The baseline lift polars.

0.1

0

y/c −0.1

−0.2

0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 x/c (b) The tested flap deflections.

Figure 3.1 – Lift polars for the baseline configurations at multiple flap deflections at Re = 2 .0 · 10 6.

MSc. Thesis 29 D.P. Jansen Chapter 3. Experimental results

−8 ◦ Baseline, α = 0 ◦ Baseline, α = 6 −7 ◦ Baseline, α = 12 ◦ Baseline, α = 13 −6

−5

−4 p C −3

−2

−1

0

1 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 1.1 1.2 x/c ◦ (a) δf = 55 .

−8 −8 ◦ ◦ Baseline, α = 0 Baseline, α = 0 ◦ ◦ Baseline, α = 6 Baseline, α = 6 −7 ◦ −7 ◦ Baseline, α = 12 Baseline, α = 12 ◦ ◦ Baseline, α = 13 Baseline, α = 13 −6 −6

−5 −5

−4 −4 p p C C −3 −3

−2 −2

−1 −1

0 0

1 1 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 1.1 1.2 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 1.1 1.2 x/c x/c ◦ ◦ (b) δf = 50 (c) δf = 45 .

◦ ◦ Figure 3.2 – Pressure distributions for baseline configurations at multiple a.o.a for δf = 55 , δf = 50 and ◦ δf = 45 . is relieved which also favours the delay of flow separation. This is due to the fact that the upper flap pressure distribution is suppressed by the displacement effect of the main airfoil large wake above the flap. This means that it may be possible that if the airfoil flow separation could be delayed and a higher stall angle could be reached, for example by applying a leading edge flap, the separation on the flap at these high incidence angles can be removed. In fact figures 3.2a - 3.2c show that at stall, the flow on the flap is completely attached, indicating the aforementioned potential. Focusing on figure 3.2b partly explains the lift drop at high a.o.a. observed in figure 3.1. It reveals that ◦ at α = 12 the flow over the flap is largely separated resulting in a low Clflap . This also has its effect on the airfoil pressure distribution especially at its trailing edge, which is probably caused by a smaller circulation effect of the flap on the main element. This at least explains that the root cause of the lift stagnation at high angle of attack is found in the flap performance. As hypothesized in the previous section it could be that a laminar bubble burst on the flap has caused this. Section 3.4.2 will provide additional insight through the performed oil flow visualizations.

D.P. Jansen 30 MSc. Thesis 3.3. Vortex generators

3.3 Vortex generators

This section will describe the wind tunnel results for the airfoil flap model tested with vortex generators. The vortex generators were only tested for a flap deflection of 55 ◦. The performance will be addressed by observing the lift polars and comparing them to the baseline situation. This is followed by a closer inspection on the pressure distributions. The complete test matrix for the vortex generators is presented in appendix A.

VG lift polars

Figure 3.3 shows the lift performance of two sets of multiple VG configurations presented in several lift polars. Equivalent to the baseline tests the angle of attack is increased from −6◦ to 13 ◦ at which the lift was measured. This time the model did not necessarily needed to be lowered first to negative angles to ◦ ◦ have semi-attached flow on the flap at α = 0 , as was the case for the baseline model at δf = 55 . As is covered in section 2.3.1 several VG configurations have been selected and tested. It can be seen from the figures that the VG configurations all show an increase in lift at lower angles of attack but generally have a decrease in lift near the stall angle. It should be noted that the main area of interest for ◦ a lift increase is more around α = 0 and not at Clmax . This is because the model is tested with a trailing edge flap only, hence not in combination with a leading edge slat which gives more performance at higher angle of attack. From figure 3.3 it can be seen that VG’s with smaller and larger heights ( h∗ = 0 .010 and h∗ = 0 .070 ) show only minor effect in lift improvements. The smaller height, h∗ = 0 .010 , results vortices that are too small and therefore insufficient in mixing enough high-energy flow to the boundary layer at the flap. The larger VG’s, h∗ = 0 .070 , are probably causing adverse vortex interaction since they lead to very fluctuating lift values also visible in the lift polar. The spacing also shows an effect where values should not either be too large or too small for an optimal vortex performance. VG’s at a spacing of λ/h = 6 .5 and λ/h = 4 .5 do not result in a significant lift improvement. This is probably due to a spacing that is too large and where an insufficient amount of healthy air is pumped into the boundary layer. Only the VG configuration with spacing λ/h = 4 .0 and height h∗ = 0 .023 shows a significant performance gain over the whole range of α, this setup is found as the optimum VG configuration. In ◦ this case a gain is found at α = 0 with lift increases up to ∆Clo = 0 .11 which is an improvement of 6.0% compared to the baseline. At high angles improvements are obtained of ∆Clmax = 0 .064 an improvement of 2.2% . It is clear for all configurations that stall is not delayed, although this is not surprising since the stall for this airfoil flap model is mainly determined by the airfoil stall characteristics and not the performance of the flap.

The total lift can be subdivided into the contributions of both the flap and airfoil, Clflap and Clairfoil . Figure 3.4a shows that the lift gain at low incidence angles is initially both on the flap and airfoil. However as the angle of attack increases the lift gain for the airfoil decreases while the contribution from the flap remains fairly constant. Another visualization can be made of the airfoil and flap lift contribution by plotting the ratio Clairfoil /C lflap versus the a.o.a. Figure 3.4b indicates that compared to the baseline configuration, the lift at higher angles of attack has a larger contribution from the flap lift component.

VG pressure distributions

Figure 3.6 shows the pressure distributions for the found optimum VG configuration at multiple angles of attack. A very comparable pressure distribution development to figure 3.2 is seen for the different angles of attack. Comparison of the flap pressure distributions of the baseline and VG configurations shows that the tendency of a higher flap pressure at larger angle of attack for the baseline model is not found for the airfoil flap model with VG’s. In figure 3.6 it is visible that the flap pressure remains fairly constant for increasing angle of attack. This also explains the result found from figure 3.4b where at higher angle of attack a larger contribution to the total lift is found from the flap lift component.

MSc. Thesis 31 D.P. Jansen Chapter 3. Experimental results

3.5 3.5

Baseline Baseline ∗ ∗ VG, λ/h = 6 .5 VG, λ/h = 6 .5, h = 0 .010, d = 0 ∗ ∗ VG, λ/h = 4 .5 VG, λ/h = 4 .5, h = 0 .010, d = 0 ∗ ∗ 3 VG, λ/h = 4 .0 3 VG, λ/h = 4 .0, h = 0 .070, d = 0 .13

2.5 2.5 l l C C

2 2

1.5 1.5

1 1 −10 −5 0 5 10 15 −10 −5 0 5 10 15 ◦ ◦ α [ ] α [ ] (a) VG’s set 1 with h∗ = 0 .023 and d∗ = 0 . (b) VG’s set 2.

0.1 0.1

0 0

y/c −0.1 y/c −0.1

−0.2 −0.2

0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 x/c x/c (c) VG positions set 1. (d) VG positions set 2.

Figure 3.3 – Effect of height, spacing and edge distance on the lift for multiple VG configurations at δf = ◦ 55 .

D.P. Jansen 32 MSc. Thesis 3.3. Vortex generators

3.5 Baseline VG, λ/h = 6 .5 VG, λ/h = 4 .5 3 VG, λ/h = 4 .0

2.5 15 Baseline VG, λ/h = 6 .5 2 VG, λ/h = 4 .5 VG, λ/h = 4 .0

airfoil 12 l C , flap

l 1.5

C 9 f lap l /C

1 airf oil l 6 C

0.5 3

0 −10 −5 0 5 10 15 0 ◦ α[ ] −10 −5 0 5 10 15 ◦ α[ ]

(a) Clflap and Clairfoil versus α. (b) Clairfoil and Clflap versus α.

∗ ∗ Figure 3.4 – Flap and airfoil lift contributions for multiple VG configurations with h = 0 .023 , d = 0 and ◦ δf = 55 .

−8 ◦ VG, α = 0 ◦ VG, α = 6 −7 ◦ VG, α = 12 ◦ VG, α = 13 −6

−5

−4 p C −3

−2

−1

0

1 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 1.1 1.2 x/c

∗ Figure 3.5 – Pressure distribution for optimum VG configuration at multiple a.o.a. with λ/h = 4 .0, h = ∗ ◦ 0.023 , d = 0 and δf = 55 .

MSc. Thesis 33 D.P. Jansen Chapter 3. Experimental results

−8 −8 Baseline Baseline VG, = 4 0, ∗ = 0 023, ∗ = 0 VG, = 4 0, ∗ = 0 023, ∗ = 0 −7 λ/h . h . d −7 λ/h . h . d

−6 −6

−5 −5

−4 −4 p p C C −3 −3

−2 −2

−1 −1

0 0

1 1 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 1.1 1.2 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 1.1 1.2 x/c x/c (a) α = 0 ◦. (b) α = 12 ◦.

Figure 3.6 – Pressure distributions for optimum VG and baseline configuration at low and high a.o.a. with ∗ ∗ λ/h = 4 .0, h = 0 .023 , d = 0 and δf = 55 .

Figure 3.6 shows the pressure distribution of the optimum VG configuration compared with the baseline model at α = 0 ◦ and α = 12 ◦. It can be seen for both situations that the optimum VG configuration has a slightly different pressure distribution on the flap compared to the baseline model. The pressure at the nose is slightly higher, although the pressure over the rest of the flap chord is lower. This is because the sudden pressure rise is shifted more towards the trailing edge, which indicates that flow separation seems to be delayed. At zero angle of attack the upper surface pressure distribution of the airfoil is lowered, possibly due to an increased circulation effect. This effect is more visible at lower angle of attack than at α = 12 ◦.

3.4 Cylinders

This section will describe the wind tunnel results for the airfoil flap model tested with vortex generating cylinders. The cylinders were tested for flap deflections of 55 ◦, 50 ◦ and 45 ◦. The performance of the cylinders will be addressed by observing the lift polars and comparing them to the baseline situation. This is followed by a closer inspection on the pressure distributions and performed oil flow visualizations. To better understand the working principles of the cylinder vortices, microphone frequency measurements and transverse total pressure measurements were executed at a flap deflection of 45 ◦. These results will also be presented. The complete test matrix for the vortex generators is given in appendix A.

◦ 3.4.1 Flap deflection δf = 55

◦ Cylinder lift polars for δf = 55

Figure 3.7 and 3.8 show the lift performance of several cylinder configurations in a lift polar at a flap ◦ deflection of δf = 55 . Similar to the VG tests the model did not need to be lowered first to negative angles to have semi-attached flow on the flap at α = 0 ◦. It can be seen that multiple cylinder configurations show significant improvement over the whole range of α in lift. Cylinders with varying diameters all show a performance gain when positioned properly, hence in the range of r∗ = 0 .06 ↔ 0.09 and θ = 30 ◦ ↔ 37 ◦. The optimum configuration is found at a D∗ = 0 .0167 , r∗ = 0 .090 and θ = 37 ◦. An interesting graph is illustrated in figure 3.9 where the lift gains ∆Cl are visualized against corresponding reduced frequencies

D.P. Jansen 34 MSc. Thesis 3.4. Cylinders

3.5 3.5

Baseline ∗ ◦ Baseline ∗ ∗ ◦ Cylinder, r = 0 .060, θ = 37 Cylinder, D = 0 .0167, r = 0 .060, θ = 55 ∗ ◦ ∗ ∗ ◦ Cylinder, r = 0 .090, θ = 30 Cylinder, D = 0 .0167, r = 0 .060, θ = 18 ∗ ◦ ∗ ∗ ◦ Cylinder, r = 0 .090, θ = 37 Cylinder, D = 0 .0333, r = 0 .090, θ = 37 3 3

2.5 2.5 l l C C

2 2

1.5 1.5

1 1 −10 −5 0 5 10 15 −10 −5 0 5 10 15 ◦ ◦ α [ ] α [ ] (a) Cylinders set 1 with D∗ = 0 .0167 . (b) Cylinders set 2.

0.1 0.1

0 0

y/c −0.1 y/c −0.1

−0.2 −0.2

0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 x/c x/c (c) Cylinder positions set 1. (d) Cylinder positions set 2.

◦ Figure 3.7 – Effect of size and positioning on the lift for multiple cylinder configurations at δf = 55 .

F +. The reduced frequency with the highest lift gain is determined in the experiment as F + = 3 .6, which is some what higher than the optimum value of F + = 1 found by Greenblatt and Wygnanski [2000]. The frequency corresponds to an optimum cylinder diameter of D∗ = 0 .0167 . For this configuration the largest lift gain is found near zero angle of attack, the main area of interest, with lift increases up to ∆Cl0 = 0 .32 which is an improvement of 18% compared to the baseline. At high angles improvements are obtained of

∆Clmax = 0 .24 . A trend is visible in figure 3.9 that as frequency is reduced (larger cylinders) lift gains become smaller.

The total lift can be subdivided into the contributions of both the flap and airfoil, Clflap and Clairfoil . In figure 3.10a the lift contributions of three configurations including the optimum, D∗ = 0 .0167 , r∗ = 0 .090 and θ = 37 ◦, are plotted. It shows comparable trends to the earlier found vortex generator airfoil and flap lift contributions. For increasing angle of attack the cylinder lift gain for the airfoil decreases. A similar effect is observed for the flap lift, which at this point differs slightly with the optimum VG configuration that showed a reasonably independent flap lift to the angle of attack. As for the optimum VG configuration, it can be seen from figure 3.10b that the contribution of the flap to the total lift performance is increased at higher angle of attack when compared to the baseline configuration. Similar to the case of the vortex generators the cylinders do not delay the stall of the airfoil flap model.

◦ Cylinder pressure distributions for δf = 55

Figure 3.11 shows the pressure distribution for the optimum cylinder configuration compared to the baseline model at α = 0 ◦ and α = 12 ◦. It shows at both angles of attack that the pressure peak at

MSc. Thesis 35 D.P. Jansen Chapter 3. Experimental results

3.5

Baseline ∗ ∗ ◦ Cylinder, D = 0 .025, r = 0 .090, θ = 30 ∗ ∗ ◦ Cylinder, D = 0 .025, r = 0 .105, θ = 30 ∗ ∗ ◦ Cylinder, D = 0 .025, r = 0 .045, θ = 42 3

2.5 l C

2

1.5

1 −10 −5 0 5 10 15 ◦ α [ ] (a) Cylinders set 3 with D∗ = 0 .025 .

0.1

0

y/c −0.1

−0.2

0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 x/c (b) Cylinder positions set 3.

◦ Figure 3.8 – Effect of size and positioning on the lift for multiple cylinder configurations at δf = 55 .

0.35

0.3

0.25

0.2 l C ∆ 0.15

0.1

0.05 α = 0 ◦ α = 12 ◦ 0 0 1 2 3 4 5 F+

+ Figure 3.9 – Lift gains ∆Cl dependency against reduced frequency F at high and low angle of attack ∗ ◦ ◦ with r = 0 .090 , θ = 37 and δf = 55 .

D.P. Jansen 36 MSc. Thesis 3.4. Cylinders

3.5

Baseline ∗ ◦ Cylinder, r = 0 .060, θ = 37 ∗ ◦ Cylinder, r = 0 .090, θ = 30 ∗ ◦ Cylinder, r = 0 .090, θ = 37 3

2.5 15

Baseline ∗ ◦ Cylinder, r = 0 .060, θ = 37 ∗ ◦ Cylinder, r = 0 .090, θ = 30 2 ∗ ◦ Cylinder, r = 0 .090, θ = 37

airfoil 12 l C , flap

l 1.5

C 9 f lap l /C

1 airf oil l 6 C

0.5 3

0 −10 −5 0 5 10 15 0 ◦ α[ ] −10 −5 0 5 10 15 ◦ α[ ]

(a) Clairfoil and Clflap versus α. (b) Clairfoil /C lflap versus α.

0.1 0.1

0 0

y/c −0.1 y/c −0.1

−0.2 −0.2

0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 x/c x/c (c) Cylinder positions. (d) Cylinder positions.

∗ Figure 3.10 – Flap and airfoil lift contributions for multiple cylinder configurations with D = 0 .0167 and ◦ δf = 55 .

MSc. Thesis 37 D.P. Jansen Chapter 3. Experimental results

−8 −8 Baseline Baseline Cylinder, ∗ = 0 0167, ∗ = 0 090, = 37 ◦ Cylinder, ∗ = 0 0167, ∗ = 0 090, = 37 ◦ −7 D . r . θ −7 D . r . θ

−6 −6

−5 −5

−4 −4 p p C C −3 −3

−2 −2

−1 −1

0 0

1 1 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 1.1 1.2 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 1.1 1.2 x/c x/c (a) α = 0 ◦. (b) α = 12 ◦.

Figure 3.11 – Pressure distributions for optimum cylinder and baseline configuration at low and high a.o.a. ∗ ∗ ◦ ◦ with D = 0 .0167 , r = 0 .090 , θ = 37 and δf = 55 . the leading edge of the upper flap surface is significantly lower. The sudden pressure rise is also found further towards the flap trailing edge, which suggests that flow separation is postponed. It further seems that the pressure recovery is smoother for the cylinder configuration. The enhanced flap lift also reduces pressures at the upper airfoil, which is probably achieved through an increased circulation effect of the flap on the main element.

◦ Cylinder oil flow visualization δf = 55

Figure 3.12 shows oil visualization pictures of the flap for the baseline and cylinder configuration at α = 0 ◦ and α = 10 ◦ respectively. As explained in section 2.4.3 oil flow visualization gives direct visual insight in the flow characteristics. The visible flow characteristics will now be discussed. Observation of picture 3.12a shows that the fluorescent oil is streaked downstream from left to right up till the dividing vertical line. This is the point where laminar to turbulent transition takes place. At this point, inside the bubble, there is very little flow and oil is clumped up resulting in the clear dividing line. After the bubble the flow reattaches and becomes turbulent. In this area all the oil streaks along with the uniform flow leaving a clean surface. This is because the turbulent flow has a high shear stress that scours away the oil. Further downstream the turbulent flow is unable to cope with the adverse pressure gradient and separates. The separation point is clearly seen by a second dividing line where the flow has come to rest and is clumped up again. After this point the flow is non-uniform and circulates and oil is streaked in different directions. These oil flow patterns are seen in all the pictures, although clear differences are also visible. Most noticeable is the clear shift downstream of the second vertical line for the cylinder configuration model with respect to the baseline configuration. This confirms the observations from the pressure distributions in section 3.4.1 that flow separation has indeed been postponed. This is visible at both α = 0 ◦ and α = 10 ◦. This is a major indication that the cylinder vortices indeed have a positive effect on delaying flow separation. Comparing the oil patterns between the baseline and cylinder configurations also shows that the vortex- entrained flap has a less evident transition line than the baseline run. This indicates two things. First, the boundary layer growing from the flap leading edge does have a laminar zone despite the experienced vortices entrained from the cylinder. Second, the cylinder configuration has a softer laminar to turbulent transition, probably having a shorter bubble.

D.P. Jansen 38 MSc. Thesis 3.4. Cylinders

(a) Baseline configuration at α = 0 ◦. (b) Baseline configuration at α = 10 ◦.

(c) Optimum cylinder configuration at α = 0 ◦. (d) Optimum cylinder configuration at α = 10 ◦.

Figure 3.12 – Oil flow visualization on the flap for the baseline and optimum cylinder configuration at ◦ δf = 55 .

MSc. Thesis 39 D.P. Jansen Chapter 3. Experimental results

3.5

Baseline ∗ ◦ Cylinder, r = 0 .090, θ = 18 ∗ ◦ Cylinder, r = 0 .105, θ = 30 ∗ ◦ Cylinder, r = 0 .090, θ = 30 3

2.5 l C

2

1.5

1 −10 −5 0 5 10 15 ◦ α [ ] (a) Cylinders set 1.

0.1

0 y/c −0.1

−0.2

0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 x/c (b) Cylinder positions set 1.

∗ Figure 3.13 – Effect of positioning on the lift for multiple cylinder configurations with D = 0 .0167 and ◦ δf = 50 .

D.P. Jansen 40 MSc. Thesis 3.4. Cylinders

3.5

Baseline ∗ ◦ Cylinder, r = 0 .090, θ = 18 ∗ ◦ Cylinder, r = 0 .105, θ = 30 ∗ ◦ Cylinder, r = 0 .090, θ = 30 3

2.5 15

Baseline ∗ ◦ Cylinder, r = 0 .090, θ = 18 ∗ ◦ Cylinder, r = 0 .105, θ = 30 2 ∗ ◦ Cylinder, r = 0 .090, θ = 30

airfoil 12 l C , flap

l 1.5

C 9 f lap l /C

1 airf oil l 6 C

0.5 3

0 −10 −5 0 5 10 15 0 ◦ α[ ] −10 −5 0 5 10 15 ◦ α[ ]

(a) Clairfoil and Clflap versus α. (b) Clairfoil /C lflap versus α.

0.1 0.1

0 0 y/c y/c −0.1 −0.1

−0.2 −0.2

0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 x/c x/c (c) Cylinder positions. (d) Cylinder positions.

∗ Figure 3.14 – Flap and airfoil lift contributions for multiple cylinder configurations with D = 0 .0167 and ◦ δf = 50 .

MSc. Thesis 41 D.P. Jansen Chapter 3. Experimental results

−8 −8 Baseline Baseline Cylinder, ∗ = 0 0167, ∗ = 0 090, = 30 ◦ Cylinder, ∗ = 0 0167, ∗ = 0 090, = 30 ◦ −7 D . r . θ −7 D . r . θ

−6 −6

−5 −5

−4 −4 p p C C −3 −3

−2 −2

−1 −1

0 0

1 1 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 1.1 1.2 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 1.1 1.2 x/c x/c (a) α = 0 ◦. (b) α = 12 ◦.

Figure 3.15 – Pressure distributions for optimum cylinder and baseline configuration at low and high a.o.a. ∗ ∗ ◦ ◦ D = 0 .0167 , r = 0 .090 , θ = 30 , δf = 50 .

◦ 3.4.2 Flap deflection δf = 50

◦ Cylinder lift polars for δf = 50

Figure 3.13 shows the lift performance of several cylinder configurations in a lift polar at a flap deflection ◦ of δf = 50 . It can be seen that again a lift increase with a cylinder configuration is possible over the ◦ whole range of α. This improvement however is less evident when compared to the results of the δf = 55 case. The optimum configuration, found at a D∗ = 0 .0167 , r∗ = 0 .090 and θ = 30 ◦, shows a lift increase ◦ of in ∆Cl0 = 0 .14 and 0.32 in Clmax . At first glance this is opposite to the δf = 55 case, where the lift increase is mainly found at low angles of attack. However as explained earlier and can be seen in figure ◦ 3.13, the baseline lift polar for δf = 50 at high angle of attack shows a lift stagnation. At least it is clear from figure 3.13 the cylinder configuration near Clmax does not suffer from this lift stagnation and therefore results in a large ∆Clmax . It is easy to see from figure 3.13 that if the stagnation would not ◦ occur for the baseline configuration, a comparable trend in lift increase to the δf = 55 case would be apparent. With regard to the optimum location and size of the cylinders, very comparable results are ◦ + found to the δf = 55 case. Again reduced frequencies of the cylinder vortices with F = 3 .6 are proven to be most efficient.

The contributions of both the flap and airfoil lift performance, Clflap and Clairfoil are visualized in figure 3.14a. It shows that in this case the performance improvement mainly lies in an increase of the airfoil lift. In fact the flap lift does not increase at all compared to the baseline configuration until a higher angle of attack is reached. It can be seen that as the flap lift is increased from α = 10 ◦ and onwards compared to the no-cylinder situation, the lift on the airfoil shows a larger ∆Cl than at lower angles of attack. This indicates an increased circulation effect of the flap on the airfoil at high incidence angles. This also suggests that the lift increase found at lower α is not due to an increased circulation effect. At zero angle of attack the baseline flap lift is even slightly higher with Clflap = 0 .546 compared to Clflap = 0 .544 for the cylinder variant. Finally figure 3.14b shows that for the optimum cylinder configuration the contribution of the flap to the total lift performance is increased at higher angle of attack when compared ◦ to the baseline configuration, a characteristic also observed at δf = 55 for the VG and cylinder tests.

D.P. Jansen 42 MSc. Thesis 3.4. Cylinders

◦ Cylinder pressure distributions for δf = 50

Figure 3.15 indicates that at α = 0 ◦ there are just small differences in flap pressure distributions between the original configuration and the best cylinder configuration. At the flap the negative pressure peak is slightly reduced but longer maintained, however a general gain in negative pressure at the flap upper surface is minimal. At α = 12 ◦ the pressure distributions show large differences. It is clear that these differences mainly lie in the fact that for the baseline model separation the separation point on the flap has evidently shifted towards the leading edge. This results in a significant increased and a longer maintained negative pressure peak for the cylinder configuration at this angle of attack. Again the flap upper surface ◦ ◦ ◦ pressure has a smoother recovery at both α = 0 and α = 12 . Overall comparison with the δf = 55 ◦ case concludes that the lift increase for the flap deflection of δf = 50 is mainly obtained through an improved airfoil pressure distribution.

◦ Cylinder oil flow visualization for δf = 50

◦ The oil flow pictures in figure 3.16 show comparable phenomena as seen earlier for the δf = 55 visual- izations. Clear transition and separation lines are visible accompanied between the laminar, turbulent and flow separated zones respectively. Again the cylinder configurations at both α = 0 ◦ and α = 12 ◦ see a shift in the separation line with respect to the original configurations. Where for the α = 0 ◦ case the effect is only minor, for the α = 12 ◦ the situation changes from a large separation zone for the baseline to an almost completely attached flow situation for the cylinder configuration. This indicates that the cylinders are capable of entraining the flow to the extend of even completely removing the flow separa- tion. It is at this angle where earlier the lift polars show a lift stagnation for the baseline model possibly indicating the presence of a laminar separation bubble burst. Figure 3.16b shows some interesting details of the encountered phenomenon. It is seen that the transition bubble is almost at the nose of the flap. The flow then shortly separates after the bubble. This indicates that the flow separation initiated close to the leading edge on the flap is indeed caused by a laminar bubble burst at high angle of attack. It remains unclear why the bubble burst only occurs at this specific condition, hence short range of high angle of attack at this flap deflection. The possible effect of bubble bursting is not found for the cylinder configuration. It is clear that the cylinder vortices besides having direct positive lift effects on the system are also capable of removing laminar bubble bursting, which results in the major improvements in delay of flow separation seen in figure 3.16b and 3.16d.

◦ 3.4.3 Flap deflection δf = 45

◦ Cylinder lift polars for δf = 45

Figure 3.17 shows the lift performance of several cylinder configurations in a lift polar at a flap deflection ◦ of δf = 45 . It can be seen that only minor lift improvements are obtained over the whole range of angle of attack when cylinders are used to entrain the flow with vortices. The optimum configuration, ∗ ∗ ◦ found at a D = 0 .0167 , r = 0 .090 and θ = 37 , showed a lift increase of 0.042 in Cl0 and 0.063 in ◦ Clmax . The optimum position is found to be similar to the other two cases with flap deflections δf = 55 ◦ + and δf = 50 . Again reduced frequencies of the cylinder vortices with F = 3 .6 are proven to be most efficient.

The contributions of both the flap and airfoil lift performance, Clflap and Clairfoil are visualized in figure 3.18. It shows that the performance lies in an increase of the airfoil lift for the whole range of α, accompanied by a flap lift increase at higher angle of attack. This behaviour is also seen in the ◦ δf = 50 case, although now the performance improvements are much less evident.

MSc. Thesis 43 D.P. Jansen Chapter 3. Experimental results

(a) Baseline configuration at α = 0 ◦. (b) Baseline configuration at α = 12 ◦.

(c) Optimum cylinder configuration at α = 0 ◦. (d) Optimum cylinder configuration at α = 12 ◦.

Figure 3.16 – Oil flow visualization on the flap for the baseline and optimum cylinder configuration at ◦ δf = 50 .

D.P. Jansen 44 MSc. Thesis 3.4. Cylinders

3.5

Baseline ∗ ◦ Cylinder, r = 0 .090, θ = 30 ∗ ◦ Cylinder, r = 0 .090, θ = 37 3

2.5 l C

2

1.5

1 −10 −5 0 5 10 15 ◦ α [ ] (a) Cylinder set 1.

0.1

0 y/c −0.1

−0.2

0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 x/c (b) Cylinder positions

∗ Figure 3.17 – Effect of positioning on the lift for multiple cylinder configurations with D = 0 .0167 and ◦ δf = 45 .

Figure 3.18b shows that for the optimum cylinder configuration the contribution of the flap to the total lift performance is increased at higher angle of attack when compared to the baseline configuration.

◦ Cylinder pressure distributions for δf = 45

Figure 3.19a reveals that at α = 0 ◦ there are small differences in flap pressure distributions between the original configuration and the best cylinder configuration. Though the negative pressure is longer maintained at the flap upper surface, the negative pressure peak is reduced. This explains that overall the flap lift is not increased, although flow separation may be slightly delayed. At α = 12 ◦ the differences in the pressure distributions of the baseline and best cylinder configurations are more evident. Especially the pressure over the flap upper surface pressure shows a smoother recovery.

◦ Cylinder oil flow visualization for δf = 45

The oil flow pictures in figure 3.20 again show clear transition and separation lines between the laminar, turbulent and flow separated zones respectively. As well as for the other flap deflections the cylinder configurations at both α = 0 ◦ and α = 12 ◦ see a shift in the separation line with respect to the original configurations. However these shifts are less distinct which support the observations made in the pressure distributions where also small differences were noticed. Noticeable at α = 12 ◦ is the baseline configuration suffering from a small area of separated flow just before the trailing edge, while the cylinder configuration shows completely attached flow over the whole flap.

MSc. Thesis 45 D.P. Jansen Chapter 3. Experimental results

3.5

Baseline ∗ ◦ Cylinder, r = 0 .090, θ = 30 ∗ ◦ Cylinder, r = 0 .090, θ = 37 3

2.5

15

Baseline ∗ ◦ 2 Cylinder, r = 0 .090, θ = 30 ∗ ◦ Cylinder, r = 0 .090, θ = 37 airfoil l

C 12 , flap

l 1.5 C 9 f lap l /C 1 airf oil l 6 C

0.5

3

0 −10 −5 0 5 10 15 0 ◦ α[ ] −10 −5 0 5 10 15 ◦ α[ ]

(a) Clairfoil and Clflap versus α . (b) Clairfoil /C lflap versus α.

0.1 0.1

0 0 y/c y/c −0.1 −0.1

−0.2 −0.2

0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 x/c x/c (c) Cylinder positions (d) Cylinder positions

∗ Figure 3.18 – Flap and airfoil lift contributions for multiple cylinder configurations with D = 0 .0167 and ◦ δf = 45 .

D.P. Jansen 46 MSc. Thesis 3.4. Cylinders

−8 −8 Baseline Baseline Cylinder, ∗ = 0 0167, ∗ = 0 090, = 37 ◦ Cylinder, ∗ = 0 0167, ∗ = 0 090, = 37 ◦ −7 D . r . θ −7 D . r . θ

−6 −6

−5 −5

−4 −4 p p C C −3 −3

−2 −2

−1 −1

0 0

1 1 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 1.1 1.2 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 1.1 1.2 x/c x/c (a) α = 0 ◦. (b) α = 12 ◦.

Figure 3.19 – Pressure distributions for optimum cylinder and baseline configuration at low and high a.o.a. ∗ ∗ ◦ ◦ with D = 0 .0167 , r = 0 .090 , θ = 37 and δf = 45 .

Very clear are the transition lines in the baseline oil visualizations, which are hardly visible for the cylinder configurations. This is similar to the behaviour found at the other flap angles. Since this was encountered at all flap angles it was decided to further examine this observation. The softer laminar to turbulent transitions gave rise to the idea that influencing the transition itself could delay the flow separation. It is commonly known for example that turbulent boundaries are more resistant to adverse pressure gradients and can indeed have a positive effect on delaying flow separation. It was decided to perform several tests with applying zigzag tape on the flap to investigate the influence of transition on ◦ the flow separation. This was only done for an angle of δf = 45 . It was found however that zigzag tapes on different locations resulting in various turbulent flow situations on the flap did not result in significant influences on the delay of flow separation. From this could be concluded that a different transition is not sufficient to positively affect the location of flow separation.

3.4.4 Comparison of δf effects on cylinder performance and further investi- gation

In section 3.4.1 to 3.4.3 the pressure measurements, lift polars and oil flow visualizations have been shown and discussed for the cylinder configuration studies. The results concerning the working effect of ◦ the cylinders are not straightforward. It is seen that at high flap deflection of δf = 55 the cylinders indeed have a positive effect on flow separation delay and total lift with an increase up to 18% in Cl0 with ◦ ◦ ∆Cl0 = 0 .32 compared to the baseline. However, as the flap angle decreases to δf = 50 and δf = 45 , lift increases are significantly less with ∆Cl0 = 0 .14 and ∆Cl0 = 0 .042 respectively. At higher angle of attack a similar but slightly different trend is observed. In general lift gains are less ◦ ◦ at high angle of attack, ∆Clmax = 0 .24 for δf = 55 and ∆Clmax = 0 .063 for δf = 45 , excluding the ◦ exceptional case of the measurements performed at δf = 50 where a laminar separation bubble very ◦ negatively affected the baseline case. The difference is that at a high flap deflection of δf = 55 the lift ◦ gain decreased with increasing angle of attack, while at δf = 45 the cylinders perform slightly better around ∆Clmax .

MSc. Thesis 47 D.P. Jansen Chapter 3. Experimental results

(a) Baseline configuration at α = 0 ◦. (b) Baseline configuration at α = 12 ◦.

(c) Optimum cylinder configuration at α = 0 ◦. (d) Optimum cylinder configuration at α = 12 ◦.

Figure 3.20 – Oil flow visualization on the flap for the baseline and optimum cylinder configuration at ◦ δf = 45 .

D.P. Jansen 48 MSc. Thesis 3.4. Cylinders

What is certain is that there is an optimum position found for the cylinders at r∗ = 0 .090 and θ = 37 ◦ for all flap deflections. It seems higher θ result in cylinders being too close to the airfoil wall in the slot, reducing the capability of the cylinder to develop an effective vortex street as seen by Wang and Tan [2007]. Too low values of θ possibly result in vortices not passing the flap upper but lower surface. The parameter r∗ mostly affects the position of the vortex creation with respect to the point of separation. From the experiments it is likely that for values of r∗ < 0.090 the vortex strength at the point of separation reduces being less effective. For values of r∗ > 0.090 the vortex source moves closer to the separation point which may be good from a view of increased vortex strength, but is probably negatively affected by the increasingly narrower slot. These effects are similar for all the tested flap deflections. Also the optimum size seems to be the same for all flap cases, that is D∗ = 0 .0167 . It should be noted that beforehand it was expected that larger cylinders would perform better due to the fact that vortex strength increases with larger cylinder sizes and the reduced frequency is closer to F + = 1 , which is the theoretical optimum for reattaching separated flows. However as mentioned earlier, the cylinder vortices are affected very much by wall effects, which especially with the larger cylinders can be significantly influenced. It may be that the larger cylinders require a larger gap size or different overlap, because they may be too close to the airfoil or flap surface to produce an effective vortex street. The larger cylinders can also have a direct blocking effect through the slot negatively influencing the flow. Larger gaps and overlaps are not tested in this experimental investigation. Up to this point some phenomena are reasonably understood and some conclusions can be drawn. Still it remains unclear why there is a loss in effectivity for the cylinders at lower flap deflections. For this reason additional tests, that is frequency measurements and boundary layer measurements, are performed at a ◦ flap deflection of δf = 45 to gain more insight in this behaviour.

Frequency measurements

◦ At the flap angle of δf = 45 frequency measurements are performed at various locations near the flap for an angle of attack of α = 0 ◦. The experimental test setup is explained in section 2.4.2. The FFT transformed signal measured by the listening tube results in the graphs depicted in figure 3.21, where the single sided amplitude of the pressure fluctuation, |∆P (f) |, is plotted in the frequency domain. The measurements are performed for both the cylinder with D = 10 mm and the baseline configuration. The locations at which the measurements are performed are shown in figure 3.22. The locations are chosen such that the measurements will visualize a development of the vortices along the flap. Focusing on the signal measured at location 1, directly in the wake of the cylinder, there is clear peak visible at a frequency of 770 Hz for the cylinder measurements. This peak indicates the vortices shed behind the cylinder, which have a shedding frequency of 770 Hz . This is very close to the theoretical 0.2·40 St ·U∞ 1 estimation of f = 0.010 = D = 800 Hz , based on a local slot flow velocity of V = 40 m/s . This is a clear proof that vortices are indeed shed from the cylinder at the slot at the expected frequency. More important is that from this it can be concluded that the vortices are shed at a reduced frequency of F + = 3 .6 at which the cylinders capability of reattaching separated flow is close to optimal by theory, according to Greenblatt and Wygnanski [2000]. More information is given by the frequency measurements. Also in figures 3.21b and 3.21f, thus at location 2 and 3, it is seen that peaks are present at 770 Hz for the cylinder frequency measurements. These are smaller when compared to the clear peak in figure 3.21a. In free stream conditions at location 6 a clear top is observed at 770 Hz as well, although small. The peaks at the cylinder vortex frequency are not seen at location 4 and 5. These observations indicate the following. First the reduction in pressure fluctuations at 770 Hz from location 1 to 3 reveals that the vortex strength reduces substantially over this distance. At position 3 the vortex strength from the cylinders is at a similar level of intensity as smaller circulation structures that start to develop due to the onset of flow separation. It seems that at location

1The local slot flow velocity V = 40 m/s is based on data from initial numerical simulations.

MSc. Thesis 49 D.P. Jansen Chapter 3. Experimental results

0.05 0.05

Baseline ∗ ∗ ◦ Baseline ∗ ∗ ◦ Cylinder, D = 0 .0167, r = 0 .090, θ = 37 Cylinder, D = 0 .0167, r = 0 .090, θ = 37

0.04 0.04

0.03 0.03 [dB] [dB] | |

∆P(f) 0.02 ∆P(f) 0.02 | |

0.01 0.01

0 0 0 200 400 600 800 1000 1200 1400 1600 1800 2000 0 200 400 600 800 1000 1200 1400 1600 1800 2000 Frequency [Hz] Frequency [Hz] (a) Location 1. (b) Location 2.

0.05 0.05

Baseline ∗ ∗ ◦ Baseline ∗ ∗ ◦ Cylinder, D = 0 .0167, r = 0 .090, θ = 37 Cylinder, D = 0 .0167, r = 0 .090, θ = 37

0.04 0.04

0.03 0.03 [dB] [dB] | |

∆P(f) 0.02 ∆P(f) 0.02 | |

0.01 0.01

0 0 0 200 400 600 800 1000 1200 1400 1600 1800 2000 0 200 400 600 800 1000 1200 1400 1600 1800 2000 Frequency [Hz] Frequency [Hz] (c) Location 3. (d) Location 4.

0.05 0.05

Baseline ∗ ∗ ◦ Baseline ∗ ∗ ◦ Cylinder, D = 0 .0167, r = 0 .090, θ = 37 Cylinder, D = 0 .0167, r = 0 .090, θ = 37

0.04 0.04

0.03 0.03 [dB] [dB] | |

∆P(f) 0.02 ∆P(f) 0.02 | |

0.01 0.01

0 0 0 200 400 600 800 1000 1200 1400 1600 1800 2000 0 200 400 600 800 1000 1200 1400 1600 1800 2000 Frequency [Hz] Frequency [Hz] (e) Location 5. (f) Location 6.

Figure 3.21 – Frequency measurements at multiple positions behind the cylinder. Pressure fluctuations ∗ ∗ ◦ ◦ |∆P (f) | are plotted against frequency with D = 0 .0167 , r = 0 .090 , θ = 37 and δf = 45 .

D.P. Jansen 50 MSc. Thesis 3.4. Cylinders

0.4 (6)

0.3

0.2

y/c 0.1

(2) (1) (3) 0 (4)

(5) −0.1

−0.2 0.7 0.8 0.9 1 1.1 1.2 x/c

◦ Figure 3.22 – Frequency measurement locations at the flap at δf = 45 .

3 the vortices are already too weakened to have a positive effect on flow mixing in the boundary layer. It follows that the vortices can only be effective before this point, between 0.845 < x/c < 0.95 . This could ◦ ◦ be the clarifying answer to why the cylinders are so effective at δf = 55 and not at δf = 45 . Observing ◦ the pressure distributions in figures 3.11, 3.15 and 3.19 shows that for δf = 55 the separation occurs ◦ ◦ before x/c = 0 .95 , while for δf = 50 and δf = 45 the separation point is slightly behind x/c = 0 .95 . At ◦ the large flap deflection of δf = 55 the strength of the vortices at the point of separation is sufficient to ◦ energize it, while at δf = 45 this is not the case.

It may be that the quick decay of vortices is not the only effect for a loss in mixing efficiency. The measurements at location 6 at which lower pressure fluctuations are measured than directly behind the cylinder clarifies the following. It shows the expected behaviour of these pressure fluctuations to decrease with increasing distance from their point of origin. The spectrum measured at location 6 is a propagated signal originated from the vortex street passing from location 1 to 2. This means that if not directly measured in a vortex the measured fluctuations can be lower than the actual value for the closely passing vortex. It could be that the measurements of |∆P (f) | along the flap could also indicate that the vortex gradually deviate from the streamline close to the wall, at which the measurements are performed, leading to a loss in effectivity of re-energizing the boundary layer. This effect would be more critical for the lower flap deflections for which the vortices need to pass a larger distance along the flap surface. The exact effect of this is not known, since the path of the vortices is not determined.

In section 3.4.4 the differences in separation delay at high a.o.a. is addressed. A point can now also be ◦ made on the effect of the angle of attack, which at δf = 55 had a larger negative influence on cylinder performance, with regard to the vortex sustainability just discussed. Observing the pressure distribution ◦ at figure 3.11 for the baseline shows that for the high flap deflection δf = 55 the pressure distribution on the flap remains quite the same compared to the case at zero angle of attack. For smaller flap deflections, ◦ i.e. δf = 45 , there is a clear trend that at higher angle of attack the pressure increases on the flap, at the leading edge as well closely after. This indicates that the encountered adverse pressure gradient by the flow is lower. This could have a positive effect on the conservation of the vortices, which as explained in this section has a large influence. This would then explain why at lower flap deflections the cylinder effect is more significant at high incidence angles with respect to low angles.

MSc. Thesis 51 D.P. Jansen Chapter 3. Experimental results

Boundary layer measurements

Additional to the frequency analysis tests, transverse total pressure measurements are performed for the ◦ flap angle of δf = 45 at the trailing edge of the airfoil and several locations on the flap. Figure 3.23 shows the results for these measurements at low and high angle of attack, α = 0 ◦ and α = 12 ◦ respectively. Normally a boundary layer shape is expressed in velocity U/U e. However with the reasonable assumption that within the boundary layer the static pressure is independent in the direction normal to the surface, the total pressure can indeed give an indication of the boundary layer development. The total pressure ptot here is referenced with the ambient pressure pamb at the wind tunnel. Observing figure 3.23a at position 1, the trailing edge of the main element, the total pressure measure- ments reveal that at α = 0 ◦ a very standard boundary layer profile is seen and that there is no difference in boundary layer development between the baseline and cylinder configuration. At position 2 there are some differences between both configurations. First it is seen that there is a deviation in the total pressure very close to the wall for the baseline configuration. This is likely due to the transition point, which is close at this measurement point. As explained from the oil flow visualizations the laminar to turbulent transition is bubble is quite significant for the baseline case and that shows here in the total pressure deviation. Further from the wall the boundary layer developments are quite similar with an increase in total pressure (increasing velocity) and comparable boundary layer heights. Outside the boundary layer the total pressure for the baseline configuration is much higher. This is because the wake of the cylinder directly has a negative effect on the total pressure. Going further upwards a loss in total pressure is observed for both configurations as is caused by the wake of the airfoil. Finally at the free stream flow similar total pressure levels are found. Overall the total pressure developments do not show specific indications that the boundary layer for the cylinder case is entrained by vortices and gets energized. This may agree with the observed performance at α = 0 ◦ for this flap angle. Here only a minor total lift gain was measured, and as explained in the previous section, is not expected to have much vortex interaction at this position of x/c = 0 .937 , because of the low intensity of the vortices at this flap deflection of ◦ δf = 45 . Still some improvements would be expected. This suggests that the total pressure itself is not sufficient enough in this case to give a good indication of the boundary layer profile. At position 3 for the same angle of attack no improvements are seen in the boundary layer development expressed by the total pressure. Only from figure 3.23b which are the same measurements performed at α = 12 ◦ at location 3 there is a small sign of velocity entrainment in the boundary layer. The increase in total pressure at boundary layer height could indicate an improved velocity profile at this position, which according to the oil flow visualizations should be completely attached. However this again does not fully agree with the total pressure measurements. To conclude the performed total pressure measurements show logical and clear phenomena outside the boundary layer, but are insufficient to accurately describe the behaviour of the flow within the boundary layer. To improve local static pressures need to be known as well to obtain true boundary layer velocity profiles.

3.5 Conclusion experimental investigation

Vortex generators and cylinders have been tested in different configurations to test the ability of these devices to add momentum to the boundary layer. The hypothesis is that the shed vortices mix the boundary layer with high momentum outside flow and should result in a delay of flow separation. The experiment shows that flow separation indeed can be delayed by the effects of the passive lift devices and that lift of the total system is improved. First the baseline configuration was tested for all flap deflections. It is found that the clean model at the tested gap, overlap and flap deflections is suspected of possible bubble bursting and stall hysteresis. The VG and cylinder configurations do not show indications for this, perhaps improving the situation with their effects on the flow. Baseline oil flow visualization shows signs of laminar bubble bursting at

D.P. Jansen 52 MSc. Thesis 3.5. Conclusion experimental investigation

0.25 Ptot Baseline Ptot Optimum Cylinder

(1)

(2) 0.1

(3) y/c

−0.05

−0.2 0.7 0.8 0.9 1 1.1 1.2 x/c (a) α = 0 ◦.

0.1 Ptot Baseline (1) Ptot Optimum Cylinder

−0.05 (2)

(3) y/c

−0.2

−0.35 0.7 0.8 0.9 1 1.1 1.2 x/c (b) α = 12 ◦.

Figure 3.23 – Total pressure measurements of the boundary layer at various locations on airfoil and flap ∗ ∗ ◦ for the baseline and optimum cylinder configuration with D = 0 .0167 , r = 0 .090 , θ = 37 ◦ and δf = 45 .

MSc. Thesis 53 D.P. Jansen Chapter 3. Experimental results

high angle of attack. The stall of the multi-element model is of a leading type found on the airfoil for all configurations. The VG’s and cylinders do not increase the stall angle. The baseline configurations at all tested flap deflections show large areas of flow separation at the flap. For the cylinder configurations the separation is significantly delayed in most cases when compared to the baseline cases. This is seen in the various surface pressure measurements and oil flow visualizations. Accompanied to the effects of delays in flow separation are measurements of significant lift increases. It is ◦ found that cylinders performance is best at high flap deflections, hence at δf = 55 . The VG’s are only ◦ tested at high flap deflection of δf = 55 and show a significant performance increase in only some of the tests. It can be concluded that the cylinders capabilities to delay flow separation and increase the lift are superior to the vortex generators. This fact lies, besides in smaller noticed effects in flow separation delay, also in the narrower bandwidth at which VG’s are proven to be effective. While for the cylinders multiple positions and sizes are efficient in creating flow separation delaying vortices, the VG’s show a performance gain at only a single configuration. This is at a height of 0.023 c and spacing λ/h = 4 , where lift gains of ∆Clo = 0 .11 are measured, an improvement of 6% . Larger spacing’s result in an insufficient amount of healthy air to be pumped into the boundary layer, while a configuration with smaller VG spacing’s possibly suffer from vortex interfering effects. The cylinder configurations show performance ◦ ◦ ◦ gains of 18% with ∆Cl0 = 0 .32 at α = 0 and δf = 55 while effects were less at δf = 45 with an ∗ ◦ increase of only ∆Cl0 = 0 .042 . The optimum position is found at r = 0 .090 and θ = 37 at a size of D∗ = 0 .0167 and these settings proved to be best for all flap settings. Larger cylinder sizes are tested ◦ only at the flap deflection of δf = 55 and they prove to be less effective, although reduced frequency values are closer to the preferable F + = 1 . In these cases the cylinders could block and negatively affect the flow through the gap. Higher values of θ result in vortices that come too close to the airfoil surface, while low values of θ lead to vortices passing the flap on the lower side instead of the upper side. r∗ can have a large influence as well where values of r∗ > 0.090 result in the cylinders getting too close to the airfoil and flap surface in the slot. It is found through the frequency measurements that the vortex decay along the flap upper surface is significant and result in a vortex strength that is too small toward the flap ◦ trailing edge to be efficient at flap deflections of δf = 45 . This explains the much smaller lift gains at the lower flap deflections. In these cases a larger cylinder diameter having stronger vortices could have been more effective, although this has not been tested. In case larger cylinders are tested or when the cylinders are positioned closer to the point of separation on the flap (thus increasing r∗) should probably be done in combination with a larger gap size and overlap setting to deny negative wall effects for the vortex production. Changing the gap and overlap also directly affects the performance of the slot and thereby the lift characteristics of the complete model. ◦ The effect of the angle of attack is not completely understood. At a flap deflection of δf = 55 the lift ◦ gains are higher at low incidence angles while at δf = 45 increases are larger at high incidence angles. This could be due to a less severe adverse pressure gradient developing on the flap at lower flap angles that could have a positive effect on the vorticity preservation.

D.P. Jansen 54 MSc. Thesis CHAPTER 4

Numerical investigation

4.1 Introduction

In section 3.4 it has been proved that the cylinders can have a positive effect on the delay of flow separation and hence increase the lift. This has been quantified and visualized by the pressure measurements, frequency measurements and oil flow visualizations. However, more insight is necessary in order to determine the exact working principle of the cylinders. In addition, a direct influence of the cylinder on the local flow is investigated. To increase our understanding on these aspects a CFD (Computational Fluid Dynamics) investigation is performed. The goal in this numerical investigation is to develop a model which can give comparable results to the experiment and provide additional information for the cylinder working principles. Section 4.2 will explain the general build-up of the CFD model. Followed by section 4.3 which will discuss the numerical domain of the model. In section 4.4 several validations on the CFD model are presented. The results of the numerical investigation are discussed in section 4.5 followed by the conclusion.

4.2 General setup

The numerical analysis was performed by the commercially available FLUENT package, which is part of the ANSYS 13.0 simulation software. FLUENT is able to solve the equations of mass, momentum and energy conservation. For this model the energy equation is not solved, because the flow is assumed to be incompressible. The mass and momentum equation were solved as unsteady to accurately capture the vortex generation. However some initial steady computations were also performed as will be discussed in section 4.3. To simulate the turbulence FLUENT can apply three types of methods: Reynolds Averaging Navier Stokes (RANS), Large Eddy Simulation (LES) or Detached Eddy Simulation (DES). For this simulation RANS is selected as LES and DES simulations require much more computational effort. RANS models the turbulence through the definition of a mean and fluctuating component of all the flow variables that generates additional equations. These equations are closed by a selected turbulence model which in this

55 Chapter 4. Numerical investigation

case is done with the k − ω SST model. The model is known for its good behaviour for separated flows. This is indicated by several numerical investigations like Catalano and Amato [2003] who performed a turbulence model study for multi-element high lift configurations and praised the capabilities of the k −ω SST model in such cases. The model also gives the opportunity to serve as a low-Reynolds turbulence model where the flow is directly modelled all the way through the viscous sub-layer instead of applying the commonly used semi-empirical wall functions to capture the boundary layer. Although this low-Reynolds behaviour requires a higher density mesh and thus computational effort, it is known to be more accurate in capturing flow separation.

4.3 The numerical domain

For the grid generation a two dimensional C-type structured mesh is chosen, which is illustrated in figure 4.1. The grid generation is done using the Ansys Meshing tool. The most forward point and both the upper and lower side of the total domain are at a distance of 10 c from the airfoil flap configuration, while the rear boundary is at a distance of 15 c. Multiple zones are modelled in order to create the structured grid and to allow more control on specifying higher mesh densities for crucial areas. A Near-field zone is defined as a high-density region close to the airfoil flap model. Also a Near-wake zone is modelled as a high-density region. Both regions are indicated in figure 4.1b. The structured mesh leads to a grid of approximately 600000 elements. The choice of a multi-zone structured grid is based on the following. Although it takes more effort to create such a grid for complex shaped geometries than an unstructured grid, it can have the following benefits. First a structured grid gives more control on specifying the mesh and can prevent the mesh to have cells with bad shape, hence a high skewness. Second the numerical diffusion is kept to a minimum when the flow is aligned with the direction of the grid, which is not realizable with an unstructured grid. This increases the accuracy and can be used to decrease the number of cells leading to shorter computation times. Of course in the areas of severe flow separation and back flow, in the region behind the flap in this case, the structured grid suffers from similar numerical diffusion errors as a structured grid and more cells are necessary. Murayama et al. [2006] investigated the influence of an unstructured and structured mesh for the numerical and experimental research of a 3-dimensional wing flap model. For the low flap deflection both meshes resulted in numerical comparable results to experimental investigations. Still slightly better results were found for the structured mesh in computing the lift coefficient, with deviations in general being ≤ 1% against ≤ 1 − 3% for the unstructured mesh. As explained in the previous section the boundary layer is treated by a low-Reynolds turbulence model. For the high-density boundary layer mesh the ANSYS [2009] manual suggests a minimum of 10 cells within the viscosity affected near-wall region ( Re y < 200 ), which in this model is set at 15 cells or more at all locations. Another constraint with the low-Reynolds turbulence model is a very small height of the first cell often expressed through the definition of y+. The ANSYS [2009] manual explains that these y+values should be in the order of 1 to a maximum of 4 − 5. Figure 4.3 shows the y+ values found for the baseline configuration at a flap deflection of 45 ◦. It shows that the desired y+values are met for most of the domain.

Boundary conditions and initial conditions

The far-field boundary conditions are specified with a velocity inlet of 50 m/s and a pressure outlet at ambient pressure conditions of 101325 pa . The inlet and outlet turbulent properties are set at a turbulent intensity of 0.1% , which corresponds to the turbulence characteristics of the Low Speed Low Turbulence Wind Tunnel used in the experiment. Further a turbulence viscosity ratio of 1 is applied, which is common for exterior flows as specified in the ANSYS [2009] manual. The contours of the airfoil, flap and cylinder are given wall boundary conditions with zero slip.

D.P. Jansen 56 MSc. Thesis 4.3. The numerical domain

(a) Mesh of the total domain. The red contours in the center of the grid visualize the geometry of the airfoil flap configuration.

(b) Mesh of the airfoil flap domain. Also visible is the Near-field zone around the airfoil and flap and the Near-wake region behind the flap.

Figure 4.1 – Visualization of the numerical domain.

MSc. Thesis 57 D.P. Jansen Chapter 4. Numerical investigation

Figure 4.2 – Mesh in the vicinity of the cylinder.

10 Airfoil Lower 9 Airfoil Upper Flap Lower Flap Upper 8

7

6

5 + wall y 4

3

2

1

0 0 0.2 0.4 0.6 0.8 1 1.2 x/c

+ ◦ Figure 4.3 – ywall for the baseline model at δf = 45 .

D.P. Jansen 58 MSc. Thesis 4.3. The numerical domain

3

2.5

2

l 1.5 C

1

0.5

0 0 2000 4000 6000 8000 10000 12000 14000 16000 iteration

Figure 4.4 – Convergence of the steady state solution.

Steady solution

The unsteady simulations were preceded by steady simulations. The steady state solutions provide accurate initial conditions for the complete flow field and speed up the computations performed with the unsteady simulations. Figure 4.4 shows the convergence of the lift coefficient for a steady state solution. It can be observed that convergence is obtained after 16000 iterations.

Unsteady solution

The unsteady simulations are calculated by the PISO solver, which is known to be efficient for unsteady simulations as stated by ANSYS [2009]. Pressure, momentum and turbulent properties are calculated with 2nd order spatial discretization accuracy. The transient formulation is through a 1st order implicit method. For the unsteady simulations setting the correct time step is of primary importance. The time step was set to accurately capture the vortex shedding of the cylinder. From the experiment and theory a vortex shedding frequency is found of 770 Hz and 800 Hz respectively for a cylinder with D = 10 mm at a Strouhal number of 0.2. For this numerical simulation the experimentally found result is considered to be a better prediction and a frequency is estimated of 770 Hz . This corresponds with a time period of 1.130 · 10 −3 s for each vortex shedding. It is important to capture the shedding with a sufficient amount of time steps. Lam and Wei [2006] suggest a minimum of 20 time steps is necessary to capture a cylinder vortex street. Using this and including a small margin results in an applied time step of 5.0 · 10 −5 s for the numerical simulations. After each iteration in time the solution needs to converge to steady state. For the unsteady simulations of the configurations with cylinder, convergence is analysed by monitoring the velocity magnitude in multiple points displayed in figure 4.5. A parameter is defined to quantify the convergence in the monitor points: Ui+1 − Ui σmon = (4.1) Ui

Figure 4.6 shows σmon between each time step. In general it is found that after 25 iterations the velocities −5 in monitor points 1 and 2 are converged to σmon < 1 · 10 . However, after this amount of iterations the

MSc. Thesis 59 D.P. Jansen Chapter 4. Numerical investigation

Figure 4.5 – Locations of the monitor points.

D.P. Jansen 60 MSc. Thesis 4.4. Model validation

x 10 −3 1.5 monitor point 1 monitor point 2 monitor point 3 1

0.5

0 mon σ

−0.5

−1

−1.5 4000 4050 4100 4150 4200 iteration

Figure 4.6 – Convergence of the velocity magnitude in the monitor points.

−4 velocity at monitor 3 mostly shows values of σmon > 1·10 and does not seem to be fully converged yet. For this reason, it was decided to fix the amount of iterations between each time step at 50 iterations. It is possible that less iterations may prove to be sufficient as well.

4.4 Model validation

Before the model was used to simulate airfoil flap configurations with cylinders, it was tested for some ◦ easier case studies. This involved baseline cases where three different flap deflections of δf = 15 , ◦ ◦ δf = 45 and δf = 55 were tested. The first case study is the baseline airfoil flap configuration at a low flap deflection of 15 ◦ at which flow separation does not occur.

◦ 4.4.1 Low flap deflection δf = 15 , no flow separation

First a low flap deflection is chosen at which flow separation is not expected on the flap. Pressure data for this flap setting is obtained from Boermans and Rutten [1995], since this flap deflection was not tested in the experiment of this research. The flap for this case is positioned with a gap size and overlap of 0.03 c and 0.08 c respectively. Figure 4.7 shows in one plot the pressure distributions of the airfoil and flap for the reference case and the numerical simulation. It can be seen that at low flap angle there are some differences in the pressure distributions, but in general they seem to agree well with each other. Firstly, it must be noted that a small error exists for the experimental computation at the main airfoil trailing edge. This seems to be a problem in the post-processing Profmeasure Fortran code, where an incorrect method is applied to connect the upper and lower pressure distribution at the last pressure hole. A similar though less evident effect can be seen at the flap trailing edge pressure distribution. It can be stated that the numerically obtained pressure distributions are better representations at these locations. Figure 4.7 further shows that flow separation does not occur at the flap upper surface in both the experimentally and numerically obtained pressure distribution. A small area of separated flow is observed at the bottom of the airfoil at x/c = 0 .7, which is captured by both methods. However, the numerically calculated pressure seems to be lower when compared to experiments.

MSc. Thesis 61 D.P. Jansen Chapter 4. Numerical investigation

−8 Exp Baseline −7 CFD Baseline

−6

−5

−4 p C −3

−2

−1

0

1 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 1.1 1.2 x/c

◦ Figure 4.7 – Pressure distributions for CFD and experiment at low flap deflection at δf = 15 .

Another observation shows that the peak pressure at the flap nose is lower for the numerically calculated distribution, which also results in a higher Clflap compared to the experiment. As explained earlier in section 3.1, Smith [1975] says that the main element can be very much affected by the circulation effect of its downstream element. The higher Clflap could explain the slightly lower pressure distribution over the airfoil due to an increased circulation effect. The reason for the lower peak pressure at the flap could be a result of the earlier mentioned flow separation occurring in the slot at the bottom of the airfoil, which the CFD model is not able to capture very accurately.

◦ 4.4.2 High flap deflection δf = 45

◦ Now a flap deflection of δf = 45 is chosen at which flow separation definitely will occur on the flap. Figure 4.8 shows the pressure distribution of the airfoil and flap for the experiment and the numerical simulation. Again a small error is observed in the experimentally obtained pressure distribution at the airfoil trailing edge. Further it is clear that the numerical simulation does not capture the flow separation on the flap correctly. The point of separation is in front of the experimentally obtained location and the accompanying pressure values are significantly different. Similar to the low flap angle test case the peak pressure at the flap is lower, but for this case also the pressure near the flap trailing edge. Although the exact separation point is not correct, the CFD simulation does show the capability to accurately model the pressure recovery. Evaluating the pressure gradient of the flap at the recovery show similar values for CFD and experiment. Further calculating the surface integral from both flap pressure distributions also show that very similar lift coefficients Clflap are obtained. The pressure distribution on the main airfoil is very comparable to the experimental result. This can be expected, since no flow separation occurs on the main airfoil, but it may also be a result of the comparable

Clflap flap lift contributions. Due to the comparable flap lift contributions the net circulation effect of the numerically and experimentally obtained flap pressure distributions can be considered similar and hence do not result in differences on the main airfoil pressure distributions.

D.P. Jansen 62 MSc. Thesis 4.4. Model validation

−8 Exp Baseline −7 CFD Baseline

−6

−5

−4 p C −3

−2

−1

0

1 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 1.1 1.2 x/c

◦ Figure 4.8 – Pressure distributions for CFD and experiment at high flap deflection at δf = 45 .

◦ 4.4.3 Very high flap deflection δf = 55

◦ Finally a very high flap angle of δf = 55 is simulated. Also for this flap deflection the numerical simulations do not agree perfectly with the experimental results. Again the flow separation on the flap is not modelled correctly and also here lower pressures are found at the flap leading and trailing edge than in the experiment. A significantly higher Clflap from the CFD simulation can be observed and explains the over prediction of the pressure on the main airfoil again due to its circulation effect.

4.4.4 Conclusion on baseline model validation

From the previous model validations at low and high flap angle it can be concluded that the model especially has trouble with accurately capturing the pressure distribution on the flap. At low angles this is already visible in an over predicted negative pressure peak at the flap nose, while at higher angles the pressure distribution over the complete flap chord shows discrepancies with the experiment. In general the pressure distribution on the airfoil is well captured. This means that at this point the numerical model is quantitatively questionable. This must be noted knowing that numerical studies incorporating cylinders will involve even more complex flow phenomena around the flow separation. However, it must be said that there are trends observed for the baseline model for increasing flap deflection, which imply that the model can be appropriate for indicating trends with investigating cylinder configurations. At this point it was decided to further investigate possibilities to improve the performance of the numerical model before cylinder configurations are tested. The over predicted pressure at the flap nose for low and high flap deflections presumed that it was necessary to test the influence of laminar zones in the CFD model. For the large differences in modelling flow separation at high flap angles, it was decided to examine the possible influence of the tunnel wall and blockage effects.

4.4.5 Influence of tunnel wall at large flap deflection

The experimental results were corrected for the influence of the wind tunnel wall by the method of Allen and Vincenti [1944]. These corrections made are due to blockage effects of the tunnel walls on the model.

MSc. Thesis 63 D.P. Jansen Chapter 4. Numerical investigation

−8 Exp Baseline −7 CFD Baseline

−6

−5

−4 p C −3

−2

−1

0

1 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 1.1 1.2 x/c

◦ Figure 4.9 – Pressure distributions for CFD and experiment at very high flap deflection at δf = 55 .

Figure 4.10 – Total domain with modelled tunnel walls on upper and lower side.

These however are only corrections on the angle of attack and the performance parameters through a source and sink method. These corrections do not take into account possible interference effects on the separation point, the region of transition, and the structure of the near-wake. This makes it uncertain if the Allen & Vincenti corrections are sufficient to model the complete tunnel wall interaction with the airfoil flap model. It may be that the actual tunnel wall, especially at high flap deflections, influences the location of flow separation. Therefore a test is performed for the numerical simulation where instead of far field boundaries a nearby tunnel wall is modelled. The adjusted domain is shown in figure 4.10. It should be noted that the tunnel walls are modelled as symmetry planes to avoid the development of boundary layers. This is done to avoid the growth of very thick boundary layers, which would otherwise grow along the semi-infinitely long tunnel walls and result in a significantly smaller effective tunnel throat. It must be noted that the influence of the tunnel walls was tested on an initial mesh resulting in dissimilar computed pressure distributions when compared to the numerical baseline distribution shown in figure 4.8. This initial mesh still is suitable to test the effect of the tunnel walls.

Figure 4.11 shows that the wind tunnel walls result in a lower pressure distribution on the upper airfoil surface. The velocity over the airfoil is increased since the effective area the flow can move through is

D.P. Jansen 64 MSc. Thesis 4.4. Model validation

−8 Exp Baseline CFD Baseline −7 CFD Baseline with tunnel wall

−6

−5

−4 p C −3

−2

−1

0

1 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 1.1 1.2 x/c

Figure 4.11 – Pressure distributions for the CFD simulation modelled with and without tunnel wall at ◦ δf = 45 . decreased. This blockage effect decreases the pressure over the airfoil, although this effect is only small. Further effects like a shift in the location of the separation point on the flap are not visible. This suggests that the blockage effect does not have a significant influence on the position of separation on the flap. For this reason, it was decided not to incorporate the wind tunnel walls in further numerical simulations.

4.4.6 Modelling laminar zones in CFD

Effects on the pressure distribution

Up till this point the flow in the numerical simulation was modelled as fully turbulent. This is very common for numerical simulations, although it is not an exact model of reality. As discussed in section 3.1 the laminar to turbulent transition can have a large influence on the development of flow separation. In general, laminar flow is less resistant to an adverse pressure gradient than its turbulent counterpart and will therefore separate earlier. However the effect of laminar zones for this complex flow situation is not exactly straightforward. A test is performed with the influence of laminar zones in CFD, where laminar to turbulent transition locations are known from the experimental data. To test the improvement of the model a plot is made of the pressure distribution and compared to the fully turbulent zone model and the pressure distribution obtained from the experiment. This is shown in figure 4.13. It must be noted that the computational pressure distributions are from calculations performed on an initial mesh. This mesh does not have a proper meshed Near-wake behind the flap and there are too few cells. The computation can still be used to show a trend for the incorporation of laminar zones. The laminar and turbulent zone designation is depicted in figure 4.12. Figure 4.13 shows that modelling laminar zones has a small but noticeable effect on the flow situation compared to the fully turbulent numerical model. The pressure distribution on the airfoil for the laminar zones clearly shows the transition point occurring at 0.35 c. The location of the separation point at the flap is slightly shifted backwards and the pressure is lower just behind the flap leading edge. Also visible is the higher pressure at the flap trailing edge for the laminar zone simulation, which is matching better with the distribution from the experiment.

MSc. Thesis 65 D.P. Jansen Chapter 4. Numerical investigation

◦ Figure 4.12 – Designation of laminar and turbulent zones around the airfoil and flap at δf = 45 .

−8 ◦ EXP Baseline, δf = 45 ◦ −7 CFD Baseline, δf = 45 ◦ CFD Baseline, δf = 45 , with laminar zones −6

−5

−4 p C −3

−2

−1

0

1 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 1.1 1.2 x/c

Figure 4.13 – Numerically calculated pressure distributions with/without laminar zones compared to ex- ◦ periment at δf = 45 .

D.P. Jansen 66 MSc. Thesis 4.5. Results for baseline and cylinder configurations

Effects on the boundary layer

Defining laminar zones also affects the development of the boundary layer over the airfoil and flap. To determine this influence the total pressures of the fully turbulent and laminar zone computations are plotted along a vertical line from the airfoil surface at the trailing edge. These can be compared directly with the baseline total pressure measurements performed in the experiment.

0.2 0.2 EXP Ptot Baseline EXP Ptot Baseline CFD Ptot Baseline CFD Ptot Baseline, with laminar zones y/c y/c

0.8 0.8 x/c x/c

(a) Ptot measured for the numerical simulation mod- (b) Ptot measured for the numerical simulation with elled fully turbulent. modelled laminar zones.

Figure 4.14 – Total pressure measurements in the boundary layer visualized for numerical simulations mod- ◦ elled with laminar zones and for fully turbulent simulations at δf = 45 . Measurements are performed at x/c = 0 .787 .

Figure 4.14 shows that the value of the total pressure for both numerical cases just outside the boundary layer coincides very well with the experimentally measured value. The main difference lies in the height of the boundary layer. For the fully turbulent flow model the boundary layer is thicker than measured in the experiment while for laminar zone numerical simulation it is thinner. It is easy to understand that the latter has a smaller height result since fully developed turbulent boundary layers along the same geometry will always be thicker than ones where also a laminar and transition region is apparent, as explained by White [2006]. The fact that the boundary layer is thinner than observed in the experiment suggests that the transition location could be shifted more towards the airfoil leading edge. It should be said that the slope ∂U/∂y is modelled accurately by the laminar zone model. The effects seen in the pressure distributions and within the boundary layer indicate that the situation is improved compared to the fully turbulent modelling, hence the distributions seen in previous sections 4.4.2 and 4.4.3 could be adjusted to incorporate laminar zones and see improvements. Still the effects are only small. The exact incorporation of transition itself into a numerical simulation needs more justifications and research as is done here. For these reasons, the incorporation of laminar zones is not included in further simulations.

4.5 Results for baseline and cylinder configurations

The numerical model has been assessed by several validation tests. It is clear that determining the correct position of flow separation with the RANS model remains a troublesome task and does not exactly correspond to the experimental results, especially at very high flap angle. The incorporation of tunnel walls and laminar to turbulent transition did not significantly increase performance of the model to capture the flow separation more accurately. Quantitatively the CFD model is too inaccurate to

MSc. Thesis 67 D.P. Jansen Chapter 4. Numerical investigation

directly compare with experimental result. However, trends and certain flow phenomena could definitely be captured. This section will address the results found from the cylinder configuration in CFD and the baseline numerical simulations. At some points comparisons are made between results found from the experiments. It must be noted that only one cylinder configuration is tested. This is the optimum cylinder configuration found in the experiment with a diameter of 10 mm or D∗ = 0 .0167 at the position ∗ ◦ ◦ ◦ of r = 0 .090 and θ = 37 . The numerical model is developed for δf = 45 and δf = 55 and is tested only for a zero angle of attack.

4.5.1 Baseline and cylinder velocity field

◦ ◦ The computed mean pressure distributions for the baseline configurations at δf = 45 and δf = 55 are already shown in section 4.4. Figure 4.15a and 4.16a show the accompanying mean velocity fields ◦ of the baseline model at a flap deflection of δf = 55 , also supplied with the mean stream traces. The velocity fields indeed show the large area of separation on the flap as well as the flow-separated area near the airfoil lower side trailing edge. Figure 4.15b and 4.16b visualize velocity fields for the cylinder configuration, which in general shows a very similar result to the baseline configuration with comparable zones of highly separated flow. From this picture no clear delay of flow separation can be detected. Also from figure 4.17 showing the pressure distributions for the baseline and cylinder configuration a shift in the separation point is hard to detect. Although some minor effects are visible, neither a clear shift in separation nor an increase in lift is shown. This does not correspond to the experimental investigation at which a significant shift in the separation point was detected accompanied by an increase in lift of 18% . This gives the suggestion that the numerical model is not able to capture the flow mixing capabilities of the vortices as captured in the experiment. Further investigation of the numerical results will clarify this. To better observe the minor pressure effects on the airfoil and flap, the pressures are translated to local pressure vectors perpendicular to the surfaces in figure 4.18. Focusing on the flap reveals that the mean negative pressure is lower on the flap nose for the cylinder model. This is caused by a decreased flow velocity, also visible in figure 4.16b. This decreased flow velocity may be caused by the cylinder vortex generation utilizing total energy from the flow, leading to a decreased mean velocity over the flap nose. The net effect of this decreased velocity and higher pressure effect is that the flap upper surface pressure recovery is relieved and in fact flow separation is very slightly delayed. This observed phenomenon is not due to a mixing effect but may be a direct result of the presence of the cylinder in the flow. Figure 4.19 shows the mean skin friction coefficient on the flap upper surface. As indicated in section 3.1, the mean friction coefficient is an efficient parameter to indicate the location of flow separation. Also from this plot only a slight improvement can be seen for the cylinder configuration in the delay of flow separation. In the next section further investigations are performed on the cylinder vortex shedding development and an explanation is sought why the flow mixing does not seem to be captured by CFD.

4.5.2 Vortex development

An important aspect of actual flow mixing in the boundary layer is a correct simulation of the vortex development behind the cylinder. Therefore, the vortex generation and development is closer studied in this section. For the experiment it is measured that vortices are created behind the cylinder and that the vortex frequency agrees very well with the theoretically determined value based on a Strouhal number of St = 0 .2. It is interesting to see if the numerical simulation also shows a clear Von-Kármánn vortex street at similar vortex frequencies. Figure 4.20 visualizes the vorticity magnitude in the vicinity of the cylinder at different time instances. It is clear that vortices are shed from the top and bottom of the cylinder. A close inspection of the shed vortices, see figure 4.20d - 4.20f, reveals that from the lower surface a very clear vortex is created which

D.P. Jansen 68 MSc. Thesis 4.5. Results for baseline and cylinder configurations

(a) Baseline model.

(b) Cylinder model.

◦ Figure 4.15 – Mean velocity field and streamlines for the baseline and cylinder configuration at δf = 55 .

MSc. Thesis 69 D.P. Jansen Chapter 4. Numerical investigation

(a) Baseline model.

(b) Cylinder model.

Figure 4.16 – Mean velocity field and streamlines in the slot for the baseline and cylinder configuration at ◦ δf = 55 .

D.P. Jansen 70 MSc. Thesis 4.5. Results for baseline and cylinder configurations

−8 CFD Baseline CFD Cylinder −7

−6

−5

−4 p C −3

−2

−1

0

1 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 1.1 1.2 x/c

◦ Figure 4.17 – Mean pressure distribution for the baseline and cylinder configuration at δf = 55 .

Baseline Baseline Cylinder Cylinder

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 1.1 x/c x/c (a) Airfoil pressure vectors. (b) Flap pressure vectors.

Baseline Baseline Cylinder Cylinder

0.7 0.8 0.9 0.7 0.8 0.9 x/c x/c (c) Airfoil pressure vectors in the slot. (d) Flap pressure vectors in the slot.

◦ Figure 4.18 – Mean pressure vectorization for the baseline and cylinder configuration at δf = 55 .

MSc. Thesis 71 D.P. Jansen Chapter 4. Numerical investigation

0.04

τwmean Baseline

τwmean Cylinder

0.03

w 0.02 τ

0.01

0 0.85 0.9 0.95 1 1.05 1.1 x/c

Figure 4.19 – Mean skin friction coefficient for the baseline and cylinder configuration at the flap upper ◦ surface at δf = 55 . develops in time. The vortex created from the cylinder upper surface however develops in a stretched way and quickly dissolves as it passes the flap leading edge, which is an unexpected result. A closer look reveals that the boundary layer generated and shed from the airfoil lower side could explain the upper vortex behaviour. At all displayed time instances the shed airfoil boundary layer generates significant circulation that passes the cylinder upper side and clearly distorts the cylinder upper surface vortex. It looks like this boundary layer vorticity directs the flow and cylinder vortices vertically towards the lower airfoil surface, which forces the upper cylinder vortex to develop close to the wall. As discussed earlier in section 3.4, research of Wang and Tan [2007] showed that cylinders which are positioned to close the wall, S/D ∗ < 0.8, can show asymmetric vortex generation with stretched close wall vortices. It seems that this is also the case here for the numerical simulation. It may be, that the numerical modelling of the off- surface boundary layer is not done accurately enough and thus causes the aforementioned problem. The development of a separated boundary layer can be significantly affected by the local meshing, however no time was available to solve or further investigate this problem. The end result of the asymmetric vortex generation is that the upper surface vortex does not show a circulation effect over the flap. This will be further explained in section 4.5.3. The stretched vortex from the cylinder upper surface also seems to cause a second problem. The frequency at which the vortices are simulated does not match with the frequency found in the experiment. This frequency was calculated by observing the velocity fluctuations in monitor point 1 and 2, which are plotted in figure 4.22. It shows that the velocity change due to the upper surface cylinder vortex is hardly captured at both the monitor points but its effect is noticeable. From this figure a vortex shedding cycle of 0.00315 s can be determined, hence the time for the development of two vortices. This corresponds to a vortex frequency of 635 Hz , where experiment showed a frequency of 770 Hz , a difference of 17% . Simulations with finer meshes around the cylinder, smaller time steps or increased number of iterations between each time step have been tested but did not show significant improvements on the found vortex frequency. It may be that the asymmetrical vortex generation is an influence on the observed difference in the experimentally and numerically found cylinder vortex frequency. Remarks can also be made on the comparison of a numerical 2-D model with a 3-D occurred experimental flow phenomenon. In the 3-dimensional space of the experiment the larger and smaller flow structures are allowed to have energy cascades in all directions, while for the numerical simulation these degradations are only allowed in two directions. Besides 3D effects the chosen turbulence model also affects the vortex development. Extensive studies on simple 2D free stream cylinder configurations as performed by Rahman et al. [2007] show that vortex frequencies can already deviate for 7 − 20% depending on the chosen turbulence model. This indicates that the way of how small-scale eddies cascade is resolved can already have a large influence on the vortex frequencies of larger flow structures.

D.P. Jansen 72 MSc. Thesis 4.5. Results for baseline and cylinder configurations

(a) t = t0 (b) t = t0 + 0 .0005 s

(c) t = t0 + 0 .00095 s (d) t = t0 + 0 .00175 s

(e) t = t0 + 0 .00255 s (f) t = t0 + 0 .00315 s

◦ Figure 4.20 – Vorticity magnitude at different time instances at very high flap angle. δf = 55

MSc. Thesis 73 D.P. Jansen Chapter 4. Numerical investigation

3

2.5

2

l 1.5 C

1

0.5

Baseline Cylinder 0 0 0.005 0.01 0.015 0.02 0.025 0.03 0.035 0.04 t[s]

◦ Figure 4.21 – Lift coefficient development in time for the baseline and cylinder configuration at δf = 55 .

It is time to evaluate current obtained results. It is seen in section 4.5.1 that flow separation is not ◦ significantly delayed as was observed at δf = 55 for the experimental investigation. It is suggested that, although a direct influence of the cylinder has a minor positive effect, the numerical model is not able to capture the flow mixing effect of energizing the boundary layer. This section shows that the vortices development necessary for a flow mixing effect is not as expected. First there is an asymmetrical shedding with stretched vortices at the cylinder upper side. Second, and this could well be caused by the first observation, the vortex shedding is at a much lower frequency. These two observations could be the direct cause of the inability to have a mixing effect. In the next section the vortex influence on the flap is researched and discussed.

4.5.3 Vortex influence on the flap

The influence of the shed vortices on the flow over the flap can be displayed in several ways. Figure 4.24 and 4.25 show line rakes of the horizontal velocity component at two positions above the flap for the baseline and cylinder configuration at different time instances. Each figure also illustrates a zoomed in situation for the boundary layer. The locations of these line rakes are indicated in figure 4.23. It must be noted that in the complete time interval a single (lower cylinder side) vortex has passed over the flap. Both figures show that the line rakes for the cylinder configuration have more variations in the horizontal velocity than for the baseline case mostly outside the boundary layer, due to the passing of a vortex. It is also clear from figures 4.24a, 4.24b, 4.25a and 4.25b that overall a velocity loss exists in the wake of the cylinder. This is due to a total pressure loss, which is also observed in the experiment and shown in figure 3.23, section 3.4.4. Figure 4.25c reveals that in the boundary layer the average negative velocity component is slightly smaller at the point of separation indicating a minor separation delay, which could be the result from the earlier explained decreased pressure difference with the leading edge. Figure 4.24c shows that at the time instance of t = t0 + 0 .0170 the boundary layer is given a velocity increase at the near wall regime of the boundary layer. This could be a very small sign of a flow mixing effect though it is observed at a very small time interval. It does show that the numerically computed boundary layer is capable to react on a vortex entrainment from outside the boundary layer. Still the effect is minimal and is not sufficient to have a positive effect on delay of flow separation. It should be noted that it is here at the boundary layer where the RANS model most likely has the biggest difficulty accurately capturing the

D.P. Jansen 74 MSc. Thesis 4.5. Results for baseline and cylinder configurations

120 monitor point 1 monitor point 2 monitor point 3

90 ]

60 m/s [ V

30

0 0 0.005 0.01 0.015 t[s]

◦ Figure 4.22 – Velocity magnitude development in time at the monitor points at δf = 55 .

flow. As was seen in the previous section, where the model has difficulty determining the exact location of the separation point, it is the near wall treatment where these models start to show deviations with the experiments. It may be that the boundary layer does not correctly cope with the entrained vortex. For this different near wall treatments could be tested in combination with various turbulence models. It may be that averaging nature of the RANS models itself is not capable of capturing the desired effects.

Figure 4.23 – Visualizations of the line rakes at x = 0 .8965 c and at x = 0 .9365 c.

MSc. Thesis 75 D.P. Jansen Chapter 4. Numerical investigation

0.1 0.1 t = t0 t = t0 t = t0 + 0 .0050 t = t0 + 0 .0070 t = t0 + 0 .0105 t = t0 + 0 .0140 t = t0 + 0 .0170 t = t0 + 0 .0210 0.08 0.08

0.06 0.06 /c /c wall wall y y 0.04 0.04

0.02 0.02

0 0 0 10 20 30 40 50 60 70 80 90 100 110 0 10 20 30 40 50 60 70 80 90 100 110 Vx[m/s ] Vx[m/s ] (a) Cylinder configuration. (b) Baseline configuration.

0.02 0.02 t = t0 t = t0 t = t0 + 0 .0050 t = t0 + 0 .0070 t = t0 + 0 .0105 t = t0 + 0 .0140 t = t0 + 0 .0170 t = t0 + 0 .0210 0.016 0.016

0.012 0.012 /c /c wall wall y y 0.008 0.008

0.004 0.004

0 0 0 10 20 30 40 50 60 70 80 90 100 110 0 10 20 30 40 50 60 70 80 90 100 110 Vx[m/s ] Vx[m/s ] (c) Cylinder configuration, zoomed in on boundary (d) Baseline configuration, zoomed in on boundary layer. layer.

◦ Figure 4.24 – Horizontal velocity component at x/c = 0 .8965 for very high flap angle at δf = 55 .

4.6 Conclusion on numerical investigation

The goal of the numerical investigation is to develop a model that gives comparable results to the experiment and provide additional information on the working principles of the cylinders. The simulations indicate that both items remain a troublesome task for the numerical model to perform. ◦ ◦ A 2D model is created of the NLF-Mod22B airfoil at a flap setting of δf = 45 and δf = 55 . While in general the flow around the airfoil is calculated correctly, an extensive model validation shows that it is not possible to accurately capture the flow separation occurring on the flap. While at low flap angles the ◦ RANS model performs quite well it loses accuracy at higher flap angles δf > 55 . Further investigations on tunnel wall effects and incorporation of laminar zones, do not show significant improvements. It seems that the application of a complex and resource demanding near wall treatment in combination with the k − ω SST turbulence model known for its favourable behaviour in high separation flows, is not sufficient to correctly model the boundary layer in the strong adverse pressure gradient that arises on the flap. This means that the baseline configuration already shows deviations with the experimentally obtained results. Still simulations were performed on the cylinder configurations to observe possible trends measured by

D.P. Jansen 76 MSc. Thesis 4.6. Conclusion on numerical investigation

0.1 0.1 t = t0 t = t0 t = t0 + 0 .0050 t = t0 + 0 .0070 t = t0 + 0 .0105 t = t0 + 0 .0140 t = t0 + 0 .0170 t = t0 + 0 .0210 0.08 0.08

0.06 0.06 /c /c wall wall y y 0.04 0.04

0.02 0.02

0 0 −20 −10 0 10 20 30 40 50 60 70 80 90 100 110 −20 −10 0 10 20 30 40 50 60 70 80 90 100 110 Vx[m/s ] Vx[m/s ] (a) Cylinder configuration. (b) Baseline configuration.

0.02 0.02 t = t0 t = t0 t = t0 + 0 .0050 t = t0 + 0 .0070 t = t0 + 0 .0105 t = t0 + 0 .0140 t = t0 + 0 .0170 t = t0 + 0 .0210 0.016 0.016

0.012 0.012 /c /c wall wall y y 0.008 0.008

0.004 0.004

0 0 −20 −10 0 10 20 30 40 50 60 70 80 90 100 110 −20 −10 0 10 20 30 40 50 60 70 80 90 100 110 Vx[m/s ] Vx[m/s ] (c) Cylinder configuration, zoomed in on boundary (d) Baseline configuration, zoomed in on boundary layer. layer.

◦ Figure 4.25 – Horizontal velocity component at x/c = 0 .9365 for very high flap angle at δf = 55 .

MSc. Thesis 77 D.P. Jansen Chapter 4. Numerical investigation

the computational model. The cylinder tested has the settings of the optimum configuration found in the experiment with a diameter of 10 mm or D∗ = 0 .0167 at the position of r∗ = 0 .090 and θ = 37 ◦. The simulations with cylinders show difficulties concerning the vortex development and an interaction with the modelled boundary layer. The vortices created behind the cylinder are asymmetrical and the frequency does not correspond to measurements performed in the wind tunnel. Vortices calculated in the simulation are at a shedding frequency of 635 Hz where the experiment shows a frequency of 770 Hz . An undesired interference effect with the airfoil lower surface boundary layer and 3D effects are possible explanations for these observations. It is seen that the boundary layer on the flap hardly responds to this vortex development and a flow mixing effect is not detected. Results may be better if the vortex generation is simulated better, however questions remain on the capability of the modelled boundary layer to cope with the entrained vortices from the outside flow.

D.P. Jansen 78 MSc. Thesis CHAPTER 5

Conclusion and Recommendations

5.1 Conclusion

The goal of this thesis was to investigate the performance improvements of cylinder vortex streets on the Extra EA-400 airfoil flap system in a 2D situation . In the experiments cylinders with diameters in ◦ ◦ the range of 10 − 20 mm were tested at flap deflections of δf = 45 − 55 and compared to the baseline configurations. At a flap deflection of 55 ◦ a comparison is also made with the behaviour of regular VG’s with heights h∗ = 0 .023 . The model and applicable devices were tested at an inlet velocity of 50 m/s in a closed loop wind tunnel. A fixed gap and overlap was used for the slot parameters at 0.035 c and 0.0c respectively. The static pressure measurements and oil flow visualizations indicate that both the cylinders and VG’s are capable of delaying flow separation on the flap and increasing the lift. It is found that the performance of cylinders are superior to the regular VG’s in feasible separation delay as well as in a larger bandwidth ◦ at which cylinders are effective. At the flap deflection of 55 an increase in lift is seen of ∆Cl0 = 0 .32 for the cylinders and 0.11 for the VG’s. The VG’s positioning and sizing should be done with precision. Large spacing’s result in an insufficient amount of healthy air to be pumped into the boundary layer, while a configuration with small VG spacing’s suffer from vortex interfering effects. Low VG heights prove to be less effective as the created vortices are small and are decayed before reaching the area of flow separation. VG’s with a height of 0.023 c and spacing λ/h = 4 are found to be optimal. ◦ The tested cylinders become less effective as the flap deflection lowers to δf = 45 . Frequency tests showed that the vortex strength reduces rapidly along the flap surface, losing the ability to energize the boundary layer. At lower flap angles the vortices need be effective further toward the trailing edge, since ◦ the location of flow separation shifts aft with decreasing flap deflection. At δf = 45 a lift increase of

∆Cl0 = 0 .042 is measured with the application of cylinders. The optimum configuration is found to be the same for all flap deflections at r∗ = 0 .090 and θ = 37 ◦ with a cylinder size of D∗ = 0 .0167 (10 mm ). The reduced frequency F + at the optimum configurations is determined to be in the range of 3−4, which is close to the most efficient frequency according to literature. The cylinders require attentive placement within the airfoil flap slot. Positioning the cylinders too close to the airfoil or flap surface results in a distorted vortex street, while cylinders at a large distance from the separation location rapidly lose efficiency.

79 Chapter 5. Conclusion and Recommendations

A numerical investigation has been performed with ANSYS Fluent to provide additional information with respect to the experimental research. A RANS model is chosen to reduce computation resource demands and modelling complexity. To cope with the large areas of flow separation a k − ω SST turbulence model with low Re near wall treatment is selected. Validations of the 2D model at several flap deflections ◦ give varying results. At a flap deflection of δf = 15 the numerical model performs well and pressure ◦ ◦ distributions are comparable to the experiment. At a flap setting of δf = 45 and δf = 55 the numerical model proves to be poor in calculating an accurate flow situation around the flap. This results in an incorrect prediction of the position of flow separation. Simulations for the cylinder configurations show difficulties in calculating a realistic vortex development. A shedding frequency is measured of 635 Hz where the experiment shows a frequency of 770 Hz . In the numerical simulation the vortices are also affected too much by the airfoil surface in the slot, which leads to a stretched upper vortex and as a result an asymmetrical vortex street. There is little interaction between the vortices and the boundary layer, and no significant energizing effect of the boundary layer is observed. It is concluded that the numerical model is not able to capture this interaction. A small positive effect is noticed of the direct influence of the cylinder on the flow in the slot and separation is very slightly delayed. The results of the CFD investigation indicate the following. Although not captured in the numerical simulation, the separation delays and positive lift increments found in the experiment are primarily due to velocity entraining effects of the vortices inside the boundary layer and not only due to a direct pressure effect of the cylinder.

5.2 Recommendations

The investigation on passive high lift control applied on a practical airfoil flap model has shown interesting capabilities of vortex generating devices such as VG’s and cylinders. While the research in this thesis led to some important answers to certain questions, it also gave rise to new questions and research possibilities. Based on the findings from this investigation the following recommendations can be made for future studies both by experiment and numerical simulations:

• The cylinders should be tested on an airfoil flap system with a leading edge slat. The cylinders in this research showed capabilities of flow separation at high angle of attack. It would be interesting

to see if the cylinders are capable of further increasing Clmax in combination with the stall delaying effects of a leading edge slat. The model can be further improved to practical situations when a full wing is applied with cylinders. In this case lateral effects on the model can be further studied. • Larger cylinders could be applied together with a larger gap size/overlap to create larger vortex structures in order to improve the effectiveness of the cylinders at lower flap angles. • Flow visualization techniques like PIV (Particle Image Velocimetry) could give additional infor- mation on the behaviour of the flow in the slot and at the flap. PIV will visualize the vortex development for the different cylinder positions in the slot and give further insight in the vortex decay observed by the frequency measurements. Velocity profiles within the boundary layer should ◦ be measured for the δf = 55 configuration to better understand the flow mixing effects of the cylinders close to the wall. • It is recommended to perform a Detached Eddy Simulation (DES) or Large Eddy Simulation (LES) for further numerical investigations. DES and LES would model small-scale eddies and resolve large eddies directly resulting in a more accurate representation of the large vortex structures behind the cylinder and in the wake of the flap. The DES could still have problems in the boundary layer where it models the flow with unsteady RANS models. This could remain troublesome for the boundary layer vortex interaction and LES may prove to be more accurate. The up scaling from RANS to DES would require much higher demands on computation resource, with LES demanding even more from CPU and RAM-memory capacity.

D.P. Jansen 80 MSc. Thesis Bibliography

H.J. Allen and W.G. Vincenti. Wall interference in a two-dimensional flow wind tunnel with consideration of the effect of compressibility. Technical report, NACA, 1944. ANSYS. ANSYS FLUENT 12.0 User’s Guide . Inc., 2009. D. L. Ashby. Experimental and computational investigation of lift-enhancing tabs on a multi-element airfoil. Technical report, Stanford University Department of Aeronautics and Astronautics, 1996. M. Baragona. Unsteady Characteristics of Laminar Separation Bubbles an Experimental and Numerical Investigation . PhD thesis, TU Delft, 2004. K. Biber. Physical aspects of stall hysteresis of an airfoil with slotted flap. In AIAA Paper 1995-0440 , 1995. L.M.M. Boermans. Aerodynamic design of aircraft and advanced transportation systems . 1998. L.M.M. Boermans. Practical implementations of boundary layer suction for drag reduction and lift enhancement at low speed. In Presentation at KATnet II Workshop, Ascot UK , 2008. L.M.M. Boermans and P.B. Rutten. Two-dimensional aerodynamic characteristics of airfoil NLF-MOD22 with fowler flap. Technical report, TU Delft, 1995. J. F. Cahill. Summary of section data on trailing-edge high-lift devices. Technical report, NACA, 1949. P. Catalano and M. Amato. An evaluation of RANS turbulence modelling for aerodynamic applications. Aerospace Science and Technology , 7:493–509, 2003. L. N. Cattafesta. Actuators for active flow control. Annual Reviews , 43:247–272, 2011. D. Greenblatt and I. Wygnanski. The control of flow separation by periodic excitation. Progress in Aerospace Sciences , 36:487–545, 2000. B.M. Jones. Measurements of profile drag by the pitot-traverse method. In British Aero. Res. Council , 1936. K.M. Lam and C.T. Wei. Characteristics of vortices shed from a circular cylinder and an inclined flat plate. In Computational Wind Engineering , 2006. P. LeBeau. Flow control using plasma actuators, 2007. URL http://www.pa.uky.edu/ . B. Lee, K. Yee, W. Joo, and D. Lee. Passive control of dynamic stall via nose droop with gurney flap. In AIAA Journal , 2005.

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J. C. Lin. Control of turbulent boundary-layer separation using micro-vortex generators. AIAA Journal , 3404:99, 1999.

J. Little, M. Nishihara, I. Adamovich, and M. Samimiy. Separation control from the flap of a high-lift airfoil using DBD plasma actuators. In AIAA Paper 2009-145 , 2009. V. Maldonado, J. Farnsworth, W. Gressick, and M. Amitay. Active control of flow separation and structural vibrations of wind turbine blades. Wind Energy , 13:221–237, 2010. P.T. Meredith. Viscous phenomena affecting high-lift systems and suggestions for future CFD develop- ment. In AGARD CP-515 , 1992.

M. Murayama, T. Imamura, K. Yamamoto, and K. Kobayashi. Comparison of RANS simulations of multi-element high-lift configurations. In AIAA Paper 2006-1396 , 2006.

B. Nishri and I. Wygnanski. Effects of periodic excitation on turbulent flow separation from a flap. AIAA Journal , 36(4):547–556, 1998. Mahbubar Rahman, Mashud Karim, and Abdul Alim. Numerical investigation of unsteady flow past a cir- cular cylinder using 2-D finite volume method. Journal of Naval Architecture and Marine Engineering , 4:27–42, 2007.

I. G. Recant. Wind-tunnel investigation of a NACA 23030 airfoil with various arrangements of slotted flaps. Technical report, NACA, 1940.

K. C. Rudolph. High-lift systems on commercial subsonic airliners. Technical report, NASA, 1996.

A.M.O. Smith. High-lift aerodynamics. Journal of Aircraft , 12:518–523, 1975.

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M.V. van der Steen. Passive off-surface flow separation control methods on a simplified flapped configu- ration. Master’s thesis, TU Delft, 2009.

L.L.M. Veldhuis and M. van der Jagt. Separation postponement by means of periodic surface excitation. In ICAS , 2010.

F. von Stillfried, S. Wallin, and A. V. Johansson. An improved passive vortex generator model for flow separation control. In AIAA Journal , 2010.

X. Wang and S. Tan. Near-wake flow characteristics of a circular cylinder close to a wall. In Journal of Fluids and Structures , 2007. F.M. White. Viscous Fluid Flow . McGraw-Hill, 2006.

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D.P. Jansen 82 MSc. Thesis APPENDIX A

Test matrix for the wind tunnel experiments

In this appendix the test matrix for the wind tunnel experiments are presented. The test matrix is two-fold, one for the tested vortex generators and one for the tested cylinders.

Table A.1 – Test matrix for the vortex generators.

∗ ∗ Run δf λ/h h d Baseline 55 - - - VG 1 55 4.5 0.010 0 VG 2 55 6.5 0.010 0 VG 3 55 4.5 0.023 0 VG 4 55 4.5 ∗ 0.023 0 VG 5 55 9 0.023 0 VG 6 55 13.5 0.023 0 VG 7 55 6.5 0.023 0 VG 8 55 4 0.023 0 VG 9 55 4∗ 0.023 0 VG 10 55 4 0.07 0.13 VG 11 55 8 0.07 0.13 ∗ = shifted λ/ 2h w.r.t. previous run.

83 Appendix A. Test matrix for the wind tunnel experiments

Table A.2 – Test matrix for the cylinders.

∗ ∗ Run δf D r θ Baseline 55 - - - Cyl 1 55 0.017 0.060 37 ◦ Cyl 2 55 0.017 0.060 18 ◦ Cyl 3 55 0.017 0.060 55 ◦ Cyl 4 55 0.017 0.060 0◦ Cyl 5 55 0.017 0.060 30 ◦ Cyl 6 55 0.017 0.090 30 ◦ Cyl 7 55 0.017 0.090 37 ◦ Cyl 8 55 0.033 0.090 37 ◦ Cyl 9 55 0.033 0.090 10 ◦ Cyl 10 55 0.033 0.105 -165 ◦ Cyl 11 55 0.025 0.090 30 ◦ Cyl 12 55 0.025 0.105 30 ◦ Cyl 13 55 0.033 0.105 37 ◦ Cyl 14 55 0.25 0.045 30 ◦ Cyl 15 55 0.25 0.045 42 ◦ Baseline 50 - - - Cyl 16 50 0.017 0.090 30 ◦ Cyl 17 50 0.017 0.090 18 ◦ Cyl 18 50 0.017 0.090 54 ◦ Cyl 19 50 0.017 0.105 30 ◦ Cyl 20 50 0.017 0.075 30 ◦ Baseline 45 - - - Cyl 22 45 0.017 0.090 30 ◦ Cyl 23 45 0.017 0.090 37 ◦

D.P. Jansen 84 MSc. Thesis APPENDIX B

Additional pictures on the wind tunnel and model layout

This appendix will show some additional pictures of the Low Speed Low Turbulence Wind Tunnel of Delft University of Technology. Also some detailed pictures of the model are presented.

85 Appendix B. Additional pictures on the wind tunnel and model layout

Figure B.1 – Exterior view of the wind tunnel test section.

Figure B.2 – Interior view, looking down stream, of the wind tunnel test section. Clearly visible are the pitot tube and the wake rake that is positioned behind the model.

D.P. Jansen 86 MSc. Thesis Figure B.3 – Detailed view on the positioning of the cylinders. The mechanism supporting the cylinder is clearly visible as well as the zigzag tape that was used to remove the lower surface laminar separation bubble.

Figure B.4 – Detailed view of the suction orifices at the wing/wall junction shown here around the airfoil upper surface.

MSc. Thesis 87 D.P. Jansen Appendix B. Additional pictures on the wind tunnel and model layout

Figure B.5 – View of the zigzag positioning of the pressure holes on the airfoil and flap model surface.

D.P. Jansen 88 MSc. Thesis APPENDIX C

Output file example of pressure measurements from Profmeasure

89 Appendix C. Output file example of pressure measurements from Profmeasure

D.P. Jansen 90 MSc. Thesis MSc. Thesis 91 D.P. Jansen Appendix C. Output file example of pressure measurements from Profmeasure

◦ Figure C.1 – Raw data of pressure measurements for baseline run at δf = 55 .

In this appendix an example of an output file for a single pressure measurement is given. This output is produced by Profmeasure, which presents both the uncorrected pressure data and the pressure data corrected by the method of Allen & Vincenti. The input for Profmeasure is the measured pressures and tunnel conditions gathered by Labview. Also input is given in Profmeasure for the Allen & Vincenti correction (in this case the cx400_d30 input file is used for all measurements). The output data was further post-processed in MATLAB to visualize the results, which are shown in this thesis.

D.P. Jansen 92 MSc. Thesis