The Effect of Spanwise Location of an Active Boundary Layer Fence on Swept Wing

Performance

Thesis

Presented in Partial Fulfillment of the Requirements for the Degree Master of Science in

the Graduate School of The Ohio State University

By

Ali N. Hussain, B.S.

Graduate Program in Aeronautical and Astronautical Engineering

The Ohio State University

2019

Thesis Committee

Dr. Jeffrey P. Bons, Advisor

Dr. James W. Gregory, Committee Member

Copyrighted by

Ali N. Hussain

2019

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Abstract

Active flow control (AFC) in the form of a wall normal slot was investigated on a

NACA 643-618 laminar wing model. The wing model has a leading-edge sweep of (Λ =

30°) and tests were performed using a chordwise Reynolds number of 100,000 with specific focus on the stall characteristics and changes in performance of an AFC slot with spanwise location. The study included comparing a passive boundary layer fence (BLF) and an AFC slot at the same spanwise locations: 0.60z/b, 0.70z/b, and 0.80z/b.

Changing the location of the passive BLF resulted in increases in the maximum lift coefficient (퐶퐿푀푎푥) over the baseline ranging from 19.3% - 10.20%, with higher gains seen for fences closer to the root. The BLF at 0.60z/b experienced an unstable longitudinal static stability derivative (퐶푀훼) at an angle 13° higher than the baseline.

Moving the fence outboard to 0.80z/b resulted in delaying this unstable behavior by an additional 훼 = 6°. A gradual decline in stability was also seen when moving the fence to a higher spanwise location as the wing maintained a more positive coefficient of moment, indicating a greater ‘pitch-up’ tendency.

For the AFC tests, coefficient of momentum (퐶휇) was set at 3.98%, 7.07%, and

10.33%. Maximum lift performance increased monotonically as more momentum was introduced into the system. An AFC slot was found to improve upon aspects of the BLF performance at all spanwise locations tested for a 퐶휇 = 10.33%. The slot was found to

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improve 퐶퐿푀푎푥 by up to 23.4% at the lowest spanwise location tested (0.60z/b) and this decreased to a 10.6% benefit at 0.80z/b. In all cases, moving the slot further outboard by

0.10z/b increases the stall angle by an additional 4°; in the case of 0.80z/b the wing never experiences an unstable spike in longitudinal static stability derivate normally associated with stall. At all spanwise locations, the AFC outperformed the BLF in terms of delaying stall and an unstable 퐶푀훼.

Fluorescent tufts were used to visualize the surface flow for the baseline and AFC at all spanwise locations. The results corroborated the load cell findings and helped to visualize separation and spanwise flow at key angles of attack such as stall for the baseline wing. At higher angles of attack, evidence of the AFC slot could clearly be seen in the attached, streamwise flow directly outboard of the fence location. Evidence of a fence and tip vortices were also present for the AFC configurations, explaining the increased lift benefits seen.

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Acknowledgments

I would like to acknowledge the many people that have helped me during my studies and have had an impact on my time here at The Ohio State University. Firstly, I would like to thank my advisor, Dr. Bons, for the opportunity to perform research both as an undergraduate and graduate student. Through your guidance and encouragement, I have become a better engineer, researcher, and person as I move forward in my career.

I would like to thank my fellow students and colleagues at the Aerospace

Research Center. Thank you for your friendship, encouragement, and motivation over the past two years. I can’t wait to see all that you achieve in your careers and in life!

Lastly, and most importantly, I would like to thank my parents, Nafisa and

Nayyer Hussain, as well as my brother, Mustafa Hussain. Their passion for learning has shaped me as I undertook this journey. They have helped me succeed and grow, and they keep me motivated for the future challenges ahead. They have always supported me through my endeavors and this thesis was possible because of their encouragement.

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Vita

2014…………………………………………Ontario High School,

Ontario, OH

2018…………………………………………B.S. Aerospace Engineering,

The Ohio State University

2018 – Present……………………...... Graduate Teaching Associate, Dept. of

Mechanical and Aerospace Engineering,

The Ohio State University

Fields of Study

Major Field: Aeronautical and Astronautical Engineering

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Table of Contents

Abstract ...... iii Acknowledgments...... v Vita ...... vi List of Tables ...... ix List of Figures ...... x Nomenclature ...... xii Chapter 1. Introduction ...... 1 1.1 Motivation ...... 1 1.2 Background ...... 3 1.3 Objectives ...... 10 Chapter 2. Experimental Setup ...... 12 2.1 Low Speed Wind Tunnel ...... 12

2.2 NACA 643-618 Airfoil ...... 14 2.3 Flow Control Methods ...... 16 2.3.1 Passive Flow Control: Boundary Layer Fence ...... 16 2.3.2 Active Flow Control: Steady Slotted Blowing ...... 17 2.4 Data Acquisition ...... 23 2.4.1 Hot-Film Probe Velocity Characterization ...... 23 2.4.2 Global Wing Forces ...... 24 2.4.3 Surface Flow Visualization ...... 26 Chapter 3: Experimental Results ...... 28 3.1 Global Force Performance at Various Spanwise Locations ...... 29 3.1.1 Spanwise Location: 0.6z/b ...... 29 3.1.2 Spanwise Location: 0.7z/b ...... 36 3.1.3 Spanwise Location: 0.8z/b ...... 42 vii

3.2 Spanwise Dependence Performance Results ...... 49 3.2.2 Global Lift Performance ...... 50 3.2.3 Longitudinal Static Stability Performance ...... 53 3.3 Surface Flow Visualization: Fluorescent Tufts...... 55 3.4 AFC Slot Effect on Global Performance ...... 63 3.5 Comparison of Present Study to Previous Studies ...... 68 Chapter 4: Conclusion...... 72 Chapter 5: Future Work ...... 75

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List of Tables

Table 1: Performance Summary of Baseline, BLF, AFC at spanwise location 0.60z/b ... 36 Table 2: Performance Summary for Baseline, BLF, AFC at Spanwise Location of 0.70z/b ...... 42 Table 3: Performance Summary of Baseline, BLF, AFC at Spanwise Location of 0.80z/b ...... 49 Table 4: Longitudinal static stability derivative comparison for Baseline, BLF, and AFC (퐶휇 = 10.33%) ...... 54 Table 5: Aerodynamic forces for wing configurations at 훼 = 14° ...... 57 Table 6: Aerodynamic forces for wing configurations at 훼 = 19° ...... 58 Table 7: Aerodynamic forces for wing configurations at 훼 = 30° ...... 60 Table 8: Aerodynamic forces for wing configurations at 훼 = 36° ...... 62 Table 9: Aerodynamic forces for AFC: 0.60z/b at 훼 = 29° and 훼 = 31° ...... 63 Table 10: Performance Comparison to Historical Studies ...... 71

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List of Figures

Figure 1: Generic boundary layer fence (leading edge only) [3] ...... 2 Figure 2: Computational pressure distribution for BLF at 훼 = 15° [17] ...... 6 Figure 3: Pressure contours of simple fence (left) and extended fence (right) for a BLF on a T-38 wing [17] ...... 7 Figure 4: 퐶퐿 푣푠. 훼 from Salmi [15] ...... 8 Figure 5: LSWT at the Aerospace Research Center ...... 13 Figure 6: LSWT schematic at the Aerospace Research Center (ARC) ...... 14 Figure 7: NACA 643-618 airfoil profile...... 14 Figure 8: AFC and BLF wing mounting configurations (a) 0.60z/b; (b) 0.70z/b; (c) 0.80z/b ...... 16 Figure 9: Boundary Layer Fence ...... 17 Figure 10: AFC slot geometry CAD model ...... 18 Figure 11: Diagram of active flow control system ...... 19 Figure 12: AFC Slot exit velocity characterization for 퐶휇 = 7.07% ...... 20 Figure 13: Slot exit velocity distribution at 훼 = 25° for wind-off (top) and wind-on (bottom) for 퐶휇 = 9.92% [23] ...... 21 Figure 14: Hot film calibration setup; from top left clockwise: (a) Hot film jet calibrator (b) Close-up of hotwire probe over jet orifice (c) Hot film probe over wing AFC slot ... 24 Figure 15: Load Cell Shaft mounting ...... 25 Figure 16: Baseline wing with tufts (processed image) ...... 27 Figure 17: 퐶퐿 푣푠. 훼 (top) and Δ퐶퐿 푣푠. 훼 (bottom) for baseline, BLF, and AFC slot at spanwise location of 0.60z/b ...... 30 Figure 18: 퐶퐷 푣푠. 훼 (top) and Δ퐶퐷 푣푠. 훼 (bottom) for baseline, BLF, and AFC slot at spanwise location of 0.60z/b ...... 31 Figure 19: 퐶푀 푣푠. 훼 (top) and CMα 푣푠. 훼 (bottom) for baseline, BLF, and AFC slot at spanwise location of 0.60z/b ...... 32 Figure 20: 퐶푀훼 푣푠. 훼 (zoomed-in, top) and 퐿/퐷 푣푠. 훼 (bottom) for baseline, BLF, and AFC slot at spanwise location of 0.60z/b ...... 33 Figure 21: 퐶퐿 푣푠. 훼 (top) and Δ퐶퐿 푣푠. 훼 (bottom) for baseline, BLF, and AFC slot at spanwise location of 0.70z/b ...... 37 Figure 22: 퐶퐷 푣푠. 훼 (top) and Δ퐶퐷 푣푠. 훼 (bottom) for baseline, BLF, and AFC slot at spanwise location of 0.70z/b ...... 38 Figure 23: 퐶푀 푣푠. 훼 (top) and CMα 푣푠. 훼 (bottom) for baseline, BLF, and AFC slot at spanwise location of 0.70z/b ...... 39 Figure 24: 퐶푀훼 푣푠. 훼 (zoomed-in, top) and L/D 푣푠. 훼 (bottom) for baseline, BLF, and AFC slot at spanwise location of 0.70z/b ...... 40 Figure 25: 퐶퐿 푣푠. 훼 (top) and Δ퐶퐿 푣푠. 훼 (bottom) for baseline, BLF, and AFC slot at spanwise location of 0.80z/b ...... 44 Figure 26: 퐶퐷 푣푠. 훼 (top) and Δ퐶퐷 푣푠. 훼 (bottom) for baseline, BLF, and AFC slot at spanwise location of 0.80z/b ...... 45

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Figure 27: 퐶푀 푣푠. 훼 (top) and 퐶푀훼 푣푠. 훼 (bottom) for baseline, BLF, and AFC slot at spanwise location of 0.80z/b ...... 46 Figure 28: 퐶푀훼 푣푠. 훼 (zoomed-in, top) and 퐿/퐷 푣푠. 훼 (bottom) for baseline, BLF, and AFC slot at spanwise location of 0.80z/b ...... 47 Figure 29: 퐶퐿 푣푠. 훼 for Baseline, BLF, and AFC (퐶휇 = 10.33%) at all spanwise locations ...... 50 Figure 30: Normalized 퐶퐿푀푎푥 for all flow control configurations ...... 52 Figure 31: Longitudinal static stability derivatives for Baseline, BLF, and AFC (퐶휇 = 10.33%) ...... 53 Figure 32: Fluorescent tuft grid used in flow visualization ...... 55 Figure 33: Tufts visualization for baseline, AFC: 0.60z/b, AFC: 0.70z/b, AFC: 0.80z/b (퐶휇 = 10.33%) at 훼 = 14° ...... 57 Figure 34: Tufts visualization for baseline, AFC: 0.60z/b, AFC: 0.70z/b, AFC: 0.80z/b (퐶휇 = 10.33%) at 훼 = 19° ...... 58 Figure 35: Tufts visualization for baseline, AFC: 0.60z/b, AFC: 0.70z/b, AFC: 0.80z/b (퐶휇 = 10.33%) at 훼 = 30° ...... 60 Figure 36: Tufts visualization for baseline, AFC: 0.60z/b, AFC: 0.70z/b, AFC: 0.80z/b (퐶휇 = 10.33%) at 훼 = 36° ...... 61 Figure 37: Tufts visualization for AFC: 0.60z/b (퐶휇=10.33%) at 훼 = 29° & 훼 = 31° .. 63 Figure 38: Visualization of Forces Caused by AFC Slot ...... 64 Figure 39: 퐶퐿 푣푠. 훼 (left) and 퐶퐷 푣푠. 훼 (right) Forces for Wind off Tests ...... 65 Figure 40: Normalized 퐶퐿푀푎푥 for all flow control configurations with slot force corrections ...... 67 Figure 41: Spanwise dependence of 퐶퐿푣푠. 훼 푐omparison of Salmi [15] and present study (BLF)...... 69 Figure 42: Spanwise dependence of 퐶퐿푣푠. 훼 푐omparison of Salmi [15] and present study (AFC) ...... 70

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Nomenclature

α = angle of attack [°] AR = aspect ratio [4.3] 2 Aslot = Area of the slot exit [m ] b = semispan of wing [0.254 m] c = chord [0.1016 m] ceff = effective chord length [0.117 m] 퐷 CD = drag coefficient [-], 퐶퐷 = 2 2 .5휌푈∞푆∗cos (Λ) 퐿 CL = lift coefficient [-], 퐶퐿 = 2 2 .5휌푈∞푆∗cos (Λ) 푀 CM = moment coefficient [-], 퐶푀 = 2 2 .5휌푈∞푆푐푒푓푓∗cos (Λ) Δ퐶푀 CMα = longitudinal static stability derivative [-], 퐶 = 푀훼 Δ훼 2 푚̇ ( ) 2∗퐴푠푙표푡 휌∗퐴푠푙표푡 Cμ = momentum coefficient [%], 퐶휇 = ∗ 푆 푈∞ e = Oswald efficiency factor, [0.75] Rec = chordwise Reynolds number [-] S = wing planform area [m2] U = velocity [m/s] x = streamwise or chordwise direction y = direction normal to the streamwise flow z = corresponds to the spanwise direction Λ = sweep angle [°]

Subscripts edge = local boundary layer edge jet = slot exit ∞ = free stream conditions max = maximum value

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Chapter 1. Introduction

1.1 Motivation

Aircraft are regularly designed to perform very well in specific operating conditions and are usually designed for a distinct purpose. However, in today’s world, aircraft are frequently called upon to operate outside these optimal design conditions, and as a result, their performance can suffer. For small scale aircraft, such as unmanned aerial vehicles (UAVs), attempting to operate outside of their set design conditions can render them impractical. UAVs have become much more prevalent in use for both military and commercial applications. Due to their small size, these aircraft operate at lower speeds and in a low Reynolds number regime. At this flight regime, the aerodynamics are extremely important to consider and the impact on the aircraft’s operation must be well understood.

One benefit to operation at lower Reynolds numbers is the ability to maintain a laminar boundary layer over much of the wing’s surface. This type of boundary layer causes less drag than a turbulent boundary layer over the same aircraft. Certain airfoils, such as the NACA 643-618 airfoil, are geometrically designed to encourage laminar flow by pushing the thickest part of the airfoil further aft and creating a favorable pressure gradient to delay transition. Unfortunately, laminar boundary layers are characterized as having low momentum and thus are more susceptible to separation, particularly at high angle of attack (훼) conditions. When flow separates over a wing, also known as stall, the

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performance drastically decreases and can result in a loss in lift and control. To prevent and mitigate stall, flow control techniques, both active and passive in nature, have been developed.

Passive flow control represents geometric modifications to an object, including boundary layer trips [1] that preemptively transition the boundary layer from laminar to turbulent. Turbulent boundary layers have higher energy and mixing which help to keep flow attached to the surface and delay separation and stall. Another method for passive flow control is a boundary layer fence (BLF, also called wing fence). This flow control method first introduced by Wolfgang Liebe in 1938 and discussed by Nickel and

Wohlfahrt [2] is a physical fence built along the suction surface, sometimes extending to the pressure side, of the wing that helps to maintain lift and performance at higher 훼’s.

One downside to all passive methods is that although they provide important benefits in their optimal flight condition, they cannot be removed or ‘turned off’ in other conditions which may lead to higher weight or increased drag in certain situations.

Figure 1: Generic boundary layer fence (leading edge only) [3]

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Active flow control (AFC) describes methods that use energy or momentum to manipulate the flow field around an object. Some examples are pulsed jets [4, 5], plasma actuators [6, 7], and fluidic oscillators [8]. A main attraction to AFC methods is the ability to activate and deactivate the control depending on the flight condition, which optimizes any benefits while mitigating potential off-design negative consequences.

Although active flow control has been a dynamic area of research, not much work has been done specifically focusing on swept wings. Swept wings have complex flow conditions and stall characteristics that require different flow control techniques to be successful. Previous research focused on momentum injection at the hinge of flaps (near the separation line) using jets and fluidic oscillators [9, 10]. In addition to momentum addition, the reduction of spanwise crossflow was a driving factor in flow control.

1.2 Background

Swept wing flow control presents additional challenges when considering the three-dimensional nature of the flow. Due to the angle of the flow relative to the leading edge of the wing, flow is considered in both a chordwise and a spanwise direction (from root to tip). The spanwise flow does not directly contribute to lift and so the wing experiences a lower effective velocity than the freestream. This causes lower lift than a straight wing at the same angle of attack (훼), and particularly at low speeds and high angles of attack (훼), this can lead to earlier separation. The three-dimensional nature of the flow also leads to different stall characteristics. An infinite two-dimensional wing will stall relatively uniformly along the spanwise direction. In contrast, a swept wing will start

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to stall near the wing tip first and then move towards the root. This non-uniform stall pattern creates complex pitching moments about the aerodynamic center of the wing which can drastically impact the stability of the aircraft. To be successful, flow control on swept wings has historically targeted this spanwise flow to maintain performance and control of the wing.

Nickel and Wohlfahrt include a summary from Liebe’s work where he drew the conclusion that stopping the cross flow specifically in the boundary layer region, hence the name “boundary layer fence”, was the main mechanism of control [2]. Research has shown that a wing fence must be quite tall compared to the boundary layer in order to be effective. Further research from Zhidkosti et al, showed that the fence was actually responsible for creating a secondary vortex in addition to the wing tip vortex [11]. This second vortex appeared on the suction side of the wing attached to the inboard side of the fence. Just as the wing tip vortex forms as a result of a pressure differential on the suction and pressure sides of a wing, this fence vortex was created because of the pressure differential across the wing fence. Essentially, the fence acted as a vortex generator and helped to entrain high energy flow from the freestream into the boundary layer [12, 13].

The addition of this high energy flow helps to keep flow attached and delay stall.

Further research and testing provided details about the optimal design of the fence, including length and height. Queijo et al [14] tested numerous wing fence designs on a swept wing (Λ = 35°, 푅푒 = 1.1푀) and found that the optimal location for a fence was either 0.36z/b or 0.73z/b. In particular, a fence at 0.73z/b was found to have a reasonable compromise between pitching moment and lift performance characterizes. It

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was also discovered that a fence was more effective when it extended over and wrapped around the leading edge. It is hypothesized that the portion of the fence that wraps around the leading edge helps to form and keep the fence vortex attached as it moves around the wing and along the suction surface. This is more critical for leading edges of a larger radius. Queijo et al [14] found that sharper leading edges did not require a fence to overhang the leading edge, and in some cases this was detrimental. A reduction in fence length from 0.80x/c to 0.26 x/c proved to have little effect on performance, yet further reduction in length caused unstable pitching moments at lower angles of attack. The height of the fence was only significant for fences of a reduced chordwise extent, in which case taller fences proved to perform better. In 1952, Salmi (Λ = 45°, 푅푒 = 2푀) also found that a fence of height 0.15t (15% max thickness of the airfoil) was almost as effective as a fence of 0.60t (60% max thickness of airfoil) [15].

Williams et al and Solfelt [16, 17] both researched the effects of wing fences on a

T-38 aircraft using wind tunnel, flight test, and computational simulations. Wind tunnel testing revealed that placing the fence at 0.825z/b provided a maximum lift enhancement of 6.3% and flow visualization revealed that outboard of the fence were regions of increased attached flow and decreased spanwise cross-flow. Solfelt’s work [17] confirmed the existence and revealed the location of the two vortices found by Zhidkosti et al [11]. The counter-rotating nature of the vortex pair creates a high-pressure interaction that draws freestream flow down to the surface of the wing, seen in Figure 2.

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Figure 2: Computational pressure distribution for BLF at 훼 = 15° [17]

Both studies showed that extending the wing fence beyond the leading edge proved to be more effective. Through wind tunnel testing, Williams et al showed the effect of extending the fence around the leading edge. They found that the extended fence provided approximately 4-5% more lift over the baseline as compared to the non- extended fence for a given angle of attack and Reynolds number [16]. Through simulations, Solfelt [17] was able to reveal the reasoning behind this difference in performance. As seen in Figure 3, the extended fence (right) shows a much stronger fence vortex as compared to the simple fence (left). The stronger fence vortex is able to entrain more momentum to the surface of the wing, helping to delay separation and increase the benefits provided by the BLF.

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Figure 3: Pressure contours of simple fence (left) and extended fence (right) for a BLF on a T-38 wing [17]

The spanwise location of a passive BLF has been studied extensively along with the number of fences used [14, 15, 18, 19]. Trends showed that moving fences outboard

(closer to the tip) resulted in a lower maximum lift coefficient but delayed the onset of an unstable pitching moment to higher angles of attack. Figure 4 presents data from Salmi

[15] which shows that moving the BLF outward from 0.575z/b to 0.8z/b yields a noted 2° delay in the onset of stall. However, the wing produces a lower maximum lift; a 6.5% improvement over the baseline is seen at the lower spanwise location while the lift improvement is only 4.6% for the fence at 0.80z/b.

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Figure 4: 퐶퐿 푣푠. 훼 from Salmi [15]

As more of the wing surface is outboard of the fence, there is a greater region to take advantage of the attached flow and produce lift. Queijo et al showed that moving a fence outboard from 0.65z/b to 0.76z/b resulted in a gradual decline in the longitudinal stability characteristics of the wing before stall [14]. Specifically, the coefficient of moment (퐶푀) stayed more positive (pitch-up tendency) for the fences at a higher spanwise location. Pratt and Shields also found a greater improvement in stability over the baseline for a fence at 0.575z/b as compared to one located at 0.80z/b [20]. However, having a larger wing area outboard of the fence causes it to remain susceptible to earlier onsets of stall and a large unstable pitching moment. Having a smaller area outboard of the fence decreased the maximum lift coefficient, but also increased the angle at which an unstable pitching moment spike and static stability derivative increase occurred. When operating at moderately high angles of attack, and when lift generation is key (i.e. takeoff 8

or landing) a fence at a lower spanwise position (closer to root) may be more beneficial as the aircraft will generally not be operating close to the extended stall angle and unstable regions of performance. Conversely, when stability is critical, or when maneuvering is necessary at high angles of attack, a fence closer to the tip would ensure that the flow over the control surfaces (ailerons, flaps) remains attached so that they are effective. In general, when deciding on the spanwise location of a wing fence, it would be prudent to consider the positioning of control surfaces and ensure that the wing fence provides performance enhancement in that immediate region.

To further improve effectiveness, multiple fences can be used at various spanwise locations to target and augment the effectiveness of ailerons and trailing edge flaps [15,

19, 20]. With leading and trailing edge flaps deployed, Salmi [15] found that having two wing fences was more than twice as effective as a single fence at reducing the instability of the wing. Using multiple fences can be especially beneficial by having one fence more inboard to provide a larger lift benefit and having another fence closer to the tip which would augment the effectiveness of the control surfaces in the direct vicinity of the fence.

Although passive BLFs has been proven and used on actual aircraft, there are some detrimental effects that must be considered. Since the fence is a passive flow control, it cannot be removed during flight which causes an unavoidable drag and weight penalty at all times. While it will provide lift and stability performance enhancements in specific regimes, when used in off-design conditions, the performance can suffer and negatively impact the aircraft. Another consequence of a BLF is present if an aircraft experiences sideslip. If the fence is at a significant angle relative to the oncoming

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freestream, the flow can hit the surface of the BLF and cause drastic moments and instability for the aircraft.

Recently, a study was done by Walker and Bons [21] to study the physical mechanisms responsible for the effectiveness of a passive BLF along with a replication of these mechanisms using active flow control (AFC) through a steady blowing slot. Using a

NACA 643-618 airfoil (Λ = 30°, 푅푒푐 = 100,000), they showed that AFC could improve on the performance benefits of a BLF by delaying stall and an unstable pitching moment by an additional 훼 = 7°. In addition, the active flow control avoided a drag penalty caused by the BLF from 훼 = 14° − 19°, and increased the maximum lift coefficient (CL) by 12.8% over the baseline. Optimizing the location and characteristics of the wall- normal slot geometry through further research can lead to greater benefits while reducing the momentum flux needed.

1.3 Objectives

The current investigation is tasked with determining an optimal spanwise location for a wall normal steady slotted blowing jet. Using a NACA 643-618 airfoil, both passive

BLF and AFC will be tested at 3 spanwise locations ranging from 0.60z/b to 0.80z/b.

Testing both types of flow control will allow for a direct comparison and investigation into the differences in lift and stability enhancement on a swept wing. To characterize performance, global force measurements will be taken with a six-axis load cell for all configurations. This will aid in determining the impact of moving flow control to different spanwise locations. In addition, fluorescent tufts will be used for surface flow

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visualization. Along with the load cell, this technique will provide a better understanding of the flow field that is occurring at key moments, such as stall, and provide an opportunity to visualize the spanwise flow and separated flow over the wing. First, a baseline wing with no flow control was tested. Following this, passive flow control via the boundary layer fence was added to the wing at the 3 spanwise locations and tested using the techniques described above. Finally, the active flow control via steady slot blowing is implemented at the same spanwise locations, and then the results are compared both regarding how they relate to the passive control, but also focusing on the differences noted with spanwise location.

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Chapter 2. Experimental Setup

2.1 Low Speed Wind Tunnel

The low speed wind tunnel (LSWT) at the Ohio State University Aerospace

Research Center (ARC), shown in Figure 6, was used for all of the experimental investigations performed in the current study. This open-loop wind tunnel, powered by a centrifugal blower, has two distinct test sections; the section used for this study can be seen on the right side of Figure 5. The air then passes through a large plenum which houses a baffle plate, honeycomb straightener, and fine mesh screens. Downstream of this, a splitter plate and vacuum are used to remove the growing boundary layer before the test section. To measure the freestream velocity, a retractable pitot-static probe and a

GE Druck 0-3” H2O differential pressure transducer are used. A schematic of the wind tunnel is shown in Figure 6. The usable test section has a cross sectional area of 0.133m2 and dimensions of 0.381m x 0.349m. Clear acrylic walls allow for easy optical access and flow visualization techniques such as tufts or particle image velocimetry.

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Figure 5: LSWT at the Aerospace Research Center

The LSWT is capable of speeds up to 30 m/s with a flow uniformity at the inlet of

±2% and a freestream turbulence level of 0.5%. For the current study, data is collected at a chordwise 푅푒푐 = 100,000 which uses the true chord length (0.1016m) and the component of velocity that is normal to the leading edge. The nominal freestream velocity used for testing is ~17.6m/s. Small variations in daily conditions cause this to change to maintain the selected Reynolds number.

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Figure 6: LSWT schematic at the Aerospace Research Center (ARC)

2.2 NACA 643-618 Airfoil

The airfoil used for this study is the NACA 643-618 laminar airfoil. To encourage laminar flow over the surface, the maximum thickness is designed as far aft as possible.

The airfoil profile can be seen in Figure 7. The maximum thickness is 18% located at x/c

= 37.1% and the maximum camber is 3.31% at a location of x/c = 54.1%. For the current study, the wing model has a leading-edge sweep of Λ = 30°. The semispan of the wing is 0.254m, or approximately 73% of the test section height. This space allows for the formation of wing tip vortices which is crucial to studying swept wing aerodynamics

[22]. The geometric chord of the wing is 0.1016m. The sweep angle of 30° gives the model a streamwise chord (푐푒푓푓) of 0.117m and an aspect ratio (AR) of 4.3.

Figure 7: NACA 643-618 airfoil profile

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For each test, the wing was rotated from 0° ≤ 훼 ≤ 40°. The blockage ratio of the wing at 훼 = 40° is 18%. Wind tunnel blockage corrections are available [22], and this high blockage should be considered when comparing data outside the current study.

However, this does not pose a problem for the current scope of work when comparing flow control against the baseline experimental performance.

The wing model was designed in SolidWorks and then 3D printed via stereolithography. The wing model was printed in multiple sections including a 0.1016m span root section, 0.0508m span baseline section, 0.0508m span active flow control

(AFC) section, a 0.0508m tip section, and several 0.0254m span sections. The multiple pieces of the wing allow for more flexibility in setting the location of the BLF/AFC without the need to reprint the entire wing. The location of the BLF/AFC slot was varied from 0.60z/b to 0.80z/b and can be seen in Figure 8. A metal rod was threaded through the middle of the wing pieces and tightened with nuts at the root and tip. Multiple metal dowel pins were used between each adjacent section to prevent shifting or movement of the wing pieces. A water-based putty was then used to fill in the hollowed-out section containing the rod/nut at the tip. The wing model was painted black using Ultra Cover

Flat Black Paint + Primer spray paint before load cell and tuft testing were completed.

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(a) (b) (c)

Figure 8: AFC and BLF wing mounting configurations (a) 0.60z/b; (b) 0.70z/b; (c) 0.80z/b

2.3 Flow Control Methods

2.3.1 Passive Flow Control: Boundary Layer Fence

Passive flow control was implemented in the form of a boundary layer fence

(BLF). The design was influenced by previous studies on the characteristics of an effective BLF, which include wrapping around the leading edge, and being tall relative to the boundary layer height [15, 20] This BLF, shown in Figure 9, is made from a 1mm thick steel plate. It is angled in the streamwise direction along the effective chord of the airfoil. The height is 0.6t (60% thickness of airfoil) and the fence starts from the trailing

푥 edge on the suction side of the wing and wraps around the leading edge to = 0.25 on 푐푒푓푓 the pressure side of the wing. It is painted in the same manner as the wing.

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Figure 9: Boundary Layer Fence

For this study, various locations of the BLF were tested to directly compare the performance to that of the active flow control configuration. The locations chosen for the fence are 0.6z/b, 0.7z/b, and 0.8z/b to match the AFC slot locations.

2.3.2 Active Flow Control: Steady Slotted Blowing

Active flow control (AFC) was implemented on the wing in the form of a chordwise slot, seen in Figure 10 , which uses steady blowing to create a fluidic boundary

푥 layer fence. The slot extends from = 0.25 on the pressure side, wraps around the 푐푒푓푓

푥 leading edge, and ends at = 0.75 on the suction side and has dimensions of 0.33mm 푐푒푓푓 wide and 133mm long. The slot has two connectors near the leading edge which help to maintain a consistent slot width throughout as well as provide some structural rigidity.

Steady blowing is implemented using an Alicat Scientific MCR-3000SLPM mass flow

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controller. This controller is capable of flows up to 3000 SLPM at an accuracy of

±0.8% + 0.2% of full-scale.

Figure 10: AFC slot geometry CAD model

A diagram of the flow control system can be seen in Figure 11. The dotted line represents a copper pipe that is embedded into the wing and guides the air to the slot opening. As seen in Figure 10, the air enters the AFC section near the leading edge and is directed into a large internal volume which acts as a plenum to help the air exiting the slot have a more uniform velocity distribution.

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Figure 11: Diagram of active flow control system

For the current investigation, the effect of the spanwise location of the slot was studied. By taking advantage of the modular setup, the slot location was varied from

0.6z/b – 0.8z/b in 0.10z/b span increments. Numerous studies have investigated a passive

BLF in this range of locations [14, 15, 20, 21]. By varying the location, an optimal location for the slot could be determined. Moving forward, this information could be used to better introduce AFC into the design process.

Using a hot-film probe, the flow exiting the slot was characterized. For a closer comparison to a passive fence, the flow would mimic the shape and distribution of the

BLF seen in Figure 9. However, the internal geometry of the chamber changed the distribution as can be seen in Figure 12. The rod running through the middle of the wing creates an obstruction for the flow. Also noted is the large radius of the leading edge and the small connectors spanning the slot and reducing the local jet velocity in those location. In the 3D printing process, part of the slot became wider, measuring 0.381mm

푥 aft the rod on the suction side (from 0.35 < < 0.60) due to material shrinkage. This 푐푒푓푓 19

is compared to the nominal slot width of 0.33mm for the rest of the slot. However, the connecting rods near the leading edge prevented the shrinkage in that area, which helps to explain the change in velocity profile along the slot for the location aft of the rod.

Figure 12: AFC Slot exit velocity characterization for 퐶휇 = 7.07%

It is important to consider that the slot exit velocity characterization was performed in a benchtop setting, outside of the wind tunnel. During flight, a wing experiences a non-uniform pressure distribution across the surface of the wing including a large low-pressure region on the suction surface near the leading edge, known as the suction peak. This pressure distribution changes with angle of attack and would have an impact on the slot exit velocity presented above, when the wing is in the wind tunnel.

Walker showed through CFD simulations the changes experienced by the slot exit velocity distribution when exposed to a freestream velocity [23]. Although the slot used in that study was slightly different (1.0mm width compared to 0.33mm width for the current study) and there were no connecters spanning the leading edge of the slot, the 20

simulation showed the relative changes of the velocity profile between the wind-off and wind-on cases. The slot exit velocity shown in Figure 13 (from Walker [23]) was simulated at an 훼 = 25° and a spanwise location of 0.70z/b, however this distribution will likely change with angle of attack and location due to the difference in pressure distribution around the airfoil. The wind-off distribution is much more uniform since the internal rod geometry is not modeled in this case. For the wind-on simulation, there is a lower velocity magnitude seen near the slot exit and a tendency for the velocity to exit closer to the ends of the slot on both the suction and pressure sides.

Figure 13: Slot exit velocity distribution at 훼 = 25° for wind-off (top) and wind-on (bottom) for 퐶휇 = 9.92% [23]

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For testing, three different coefficients of momentum were chosen at each spanwise location; 퐶휇 = 3.98%, 7.07%, 10.33%. Coefficient of momentum is calculated using Equation (1) and describes the momentum flux required to achieve a level of control. Specifically, it expresses the ratio of the jet momentum flux used by the AFC to the freestream dynamic pressure multiplied by the reference area (wing area for this study). One assumption used in this equation is that the freestream density is the same as the density of the air coming out of the AFC slot. The 퐶휇 values used in the current study are higher than some traditional methods of active flow control. However, for replicating a passive BLF using wall normal steady blowing, Walker and Bons [21] showed that much higher coefficients of momentum were necessary to achieve appreciable gains in performance.

푚̇ 2 ( ) (1) 2퐴푠푙표푡 휌 ∗ 퐴푠푙표푡 퐶휇 = ∗ 2 푏 ∗ 푐푒푓푓 푈∞

For example, AFC methods that are implemented though pulsed blowing to target flow instabilities only require small momentum flux to achieve control (퐶휇 ~ 0.03%) [5].

Vortex generator jets use more momentum flux, 퐶휇 ~ 0.30%, to synthetically create streamwise vortices on a surface [24, 25]. AFC to replicate a boundary layer fence, on the other hand, targets the interruption of the spanwise flow through steady momentum injection along a chordwise slot. Thus, it requires a much larger 퐶휇 to be effective.

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2.4 Data Acquisition

Both quantitative and qualitative experimental techniques are utilized in the present study. First, a hot-film probe is used to characterize the flow exiting the slot for the AFC configurations. A load cell measures global wing forces and moments, while fluorescent mini-tufts are used to help visualize the surface flow over the wing.

2.4.1 Hot-Film Probe Velocity Characterization

To characterize the active flow control, a hot-film probe was used to measure the slot exit velocity. The TSI hot-film probe functions as a constant temperature anemometer. The 20휇푚 diameter sensor is maintained at a constant temperature as different velocities are applied. The probe was calibrated with a TSI 1127 jet calibrator and a 10mm isentropic circular nozzle. This was done to ensure that the entire probe can be fully immersed in the flow. Using this setup, the probe was exposed to different velocities to establish a relationship between Nusselt number and Reynolds number.

Thus, when testing the slot using the probe, local velocities could be determined.

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Figure 14: Hot film calibration setup; from top left clockwise: (a) Hot film jet calibrator (b) Close-up of hotwire probe over jet orifice (c) Hot film probe over wing AFC slot

A standard U-tube manometer was used to measure the differential pressure. The hot film is calibrated from 0 – 100 m/s. Using a secured stand and a traverse, the hot film was held 2.5mm away from the nozzle and the slot during calibration and testing, respectively. When testing the AFC slot, the probe was moved along the width of the slot until the center was found which corresponds to the maximum velocity at that point.

2.4.2 Global Wing Forces

To obtain quantitative performance data, an ATI Industrial Automation Gamma six-component force/torque cell was used to determine global lift, drag, and pitching

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moment. This load cell can measure forces up to 130N in the radial direction with a resolution of 0.025N. Data were sampled at 1,024 Hz for 32s resulting in 32,768 individual readings that are then averaged to obtain a single force measurement. All configurations tested for this study have data collected for the range of 훼 = 0° − 40°.

A shaft is used to mount the airfoil to the load cell and passes through an opening in the wind tunnel, shown in Figure 15. Wooden blocks are used to cover the open area without interfering with the load cell or wing to ensure that the measurements are not impacted.

Figure 15: Load Cell Shaft mounting

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The force and moment coefficients that will be presented have had swept wing transformations applied [26, 27]. Utilizing the flow normal to the leading edge of the wing in the corresponding calculations allows for a clearer comparison between swept and unswept wing performance and experimental data. This transformation is solely

1 based on the sweep angle (Λ). The factor can be applied to the lift, drag, and cos2(Λ) moment coefficients for this purpose, and these transformations have been performed for the data in this study.

2.4.3 Surface Flow Visualization

Fluorescent tufts were applied to the wing to aid in surface flow visualization.

Tufts are a common tool to qualitatively or quantitatively measure the flow on the surface of an object. They are typically small segments of thread that move and can indicate the direction of the flow. The length of tufts is based on the scale of the object and must be long enough to visualize the flow, yet short enough to not disturb the flow and impact the results. For the present study, the tufts are used only for flow visualization and are not present during quantitative load cell data collection. The method of tufting used for the wing was adapted from Walker [23]. Tufts were only applied to select configurations.

Tufts were made from #9271dd Neon Yellow polyester thread. The tufts measure

19mm long and are oriented in the streamwise direction. To minimize interference between individual tufts, they were spaced 25mm apart in the streamwise direction (+X) and 18mm apart in the span wise direction (+Z). The resulting grid pattern included alternating rows of 15 and 14 tufts and a total of 116 on the wing suction surface. This

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pattern can be seen in Figure 16. Instead of glue, the tufts were applied using #4768 Flat

Black Model Master Acrylic Paint to ensure there was no fluorescence or reflectivity during the test when the blacklight is turned on.

Figure 16: Baseline wing with tufts (processed image)

The blacklight used was an OPPSK black light with 18 LED black lights. To capture the images, the “Slow Shutter Cam” iPhone app was used on an iPhone 6s. The

‘light trail’ setting was used with a 8s shutter speed and full light sensitivity. The images were then post-processed using Zoner Photo Studio X and Gimp 2.10.8.

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Chapter 3: Experimental Results

The first step in this study was to establish the baseline performance of the current wing model. In total, 13 different configurations were tested when considering the spanwise location and varying coefficients of momentum for the active flow control cases. Load cell measurements are collected for all cases, however fluorescent tufts are only applied to the baseline and active flow control configurations.

Load cell results for each location include 퐶퐿, 퐶퐷, Δ퐶퐿, 푎푛푑 Δ퐶퐷; the latter two help to clearly show the differences in performance and regions of benefit vs. detriment, as compared to the baseline for each of the BLF and AFC configurations. In addition, stability performance is presented through 퐶푀 푎푛푑 퐶푀훼, or the pitching moment coefficient and longitudinal static stability derivative, respectively. At stall, large positive pitching moments are created from the sudden loss of lift aft of the aerodynamic center of the wing. Delaying this unstable moment can increase the performance and controllable range of aircraft. The longitudinal static stability derivative is the slope of the 퐶푀 푣푠. 훼 curve and is used to find a trim condition where the moment about an aircraft is zero.

When this derivative is negative, the aircraft tends to return to static equilibrium when perturbed by a disturbance. For swept wings, the complex three-dimensional nature of the flow can lead to problems with stability when flow begins to separate.

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3.1 Global Force Performance at Various Spanwise Locations

3.1.1 Spanwise Location: 0.6z/b

Before any flow control was applied, the baseline wing performance was characterized, and the results are shown in Figure 17 - Figure 19. The maximum 퐶퐿 for the baseline NACA 643-618 wing tested was 퐶퐿푚푎푥 = 1.28 at 훼 = 28°. Additionally, there is a clear reduction in the lift curve slope beginning at 훼 = 12°. This is also where the coefficient of moment begins to increase, eventually leading to an unstable static stability derivative. As more lift is lost aft of the aerodynamic center, the positive pitch- up tendency of the wing is more pronounced. The baseline configuration reaches a maximum unstable static stability derivative at 훼 = 14° (퐶푀훼 = 0.02325/°). The lift curve slope then decreases even more significantly at 훼 = 19°, indicating a more drastic stall and loss of lift across the entire wing. This also corresponds to a positive static stability derivative. As seen in Figure 18, there is a noted increase in CD from 훼 = 18° −

20° where the drag increases by 56% over that range. The baseline wing experiences a maximum L/D of 8.41 at 훼 = 11°.

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Figure 17: 퐶퐿 푣푠. 훼 (top) and Δ퐶퐿 푣푠. 훼 (bottom) for baseline, BLF, and AFC slot at spanwise location of 0.60z/b

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Figure 18: 퐶퐷 푣푠. 훼 (top) and Δ퐶퐷 푣푠. 훼 (bottom) for baseline, BLF, and AFC slot at spanwise location of 0.60z/b

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Figure 19: 퐶푀 푣푠. 훼 (top) and CMα 푣푠. 훼 (bottom) for baseline, BLF, and AFC slot at spanwise location of 0.60z/b

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Figure 20: 퐶푀훼 푣푠. 훼 (zoomed-in, top) and 퐿/퐷 푣푠. 훼 (bottom) for baseline, BLF, and AFC slot at spanwise location of 0.60z/b

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A BLF at a spanwise location of 0.60z/b is responsible for increasing the maximum lift to 퐶퐿푀푎푥 = 1.527, which is an increase of 19.3% over the baseline. It is also responsible for lift benefits over the baseline from 훼 = 13° − 26°. The BLF delays stall and an unstable longitudinal static stability derivative until 훼 = 27° (increase of 13° over the baseline). It should be noted that the passive BLF reaches a 퐶푀훼 = 0.154 /°, which is a much larger instability than in the baseline case. There is also a region of a drag penalty from 훼 = 13° − 26° which directly corresponds to the region of improved lift, and the BLF maintains slightly higher drag throughout the range of 훼′푠 tested.

Herein lies one of the main drawbacks of the passive fence; although it can provide significant benefit, it must be considered how the performance suffers in off-design flow conditions since it cannot be removed or ‘turned off’. This drag penalty is avoided with the AFC and the active control even shows lower drag than the baseline in the same range of 훼’s as mentioned above. The passive boundary layer fence shows a benefit to efficiency (L/D) at lower angles of attack as compared to the baseline with 퐿/퐷푀퐴푥 =

10.05 at 훼 = 9°. It remains higher than the baseline from 0° < 훼 < 10°, after which it becomes lower than the baseline due to the higher drag penalty.

The AFC slot improves upon some of the benefits of the BLF at this spanwise location. With a 퐶휇 = 3.98%, the 퐶퐿푀푎푥 is increased to 1.427 which is a 11.5% increase over the baseline and this flow control provides lift benefits starting from 훼 = 20° − 32°.

This configuration also delays an unstable longitudinal static stability derivative until

훼 = 31°, with a maximum 퐶푀훼 = 0.04058 /° at 훼 = 32°. This is an improvement over the BLF not only in terms of a higher stall angle, but also in that the instability at stall is

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much smaller. Although not a main focus of this study, the AFC slot shows an improvement over the baseline and the BLF in L/D for all AFC configurations for 훼 >

11°. By increasing the 퐶휇 to 7.07%, the wing almost matches the maximum lift of the passive BLF at 퐶퐿푀푎푥 = 1.52. However as seen in Figure 17, the clear difference is that the AFC maintains increased lift to a much higher angle of attack of 훼 = 32°. In addition, at stall, the instability in the static stability derivative is still smaller than that of the BLF.

Lastly, at the highest 퐶휇 tested, the lift performance shows a non-linearity in its behavior near 퐶퐿푀푎푥. Until 훼 = 29°, the lift performance matches that of the 퐶휇 = 3.98% case and underperforms as compared to the middle coefficient of momentum value (7.07%). Then, the CL suddenly spikes to a 퐶퐿푀푎푥= 1.58 until finally stalling at 훼 = 33°, the same as the other AFC configurations. Although this behavior represents a much larger performance gain with regards to lift, the consequences with regards to stability must be considered.

The two rapid changes in lift (sudden increase and then stall) create very undesirable changes in the moment about the airfoil and could cause serious problems for the stability and control of an aircraft. Flow visualization using tufts helps to shed some light on the physical mechanism occurring that cause this non-linear behavior.

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Table 1: Performance Summary of Baseline, BLF, AFC at spanwise location 0.60z/b

Wing Configuration Baseline BLF AFC Slot AFC Slot AFC Slot 푪흁 - - 3.98% 7.07% 10.33% CLMax 1.28 1.527 1.427 1.52 1.58 휶풔풕풂풍풍° 19° 27° 33° 33° 33° 횫CLMax % - 19.3% 11.5% 18.75% 23.44% Lift Benefit Range - 13°-26° 20°-33° 20°-35° 20°-37° Delay Unstable 푪푴휶 - 13° 18° 18° 18° 푳/푫풎풂풙 8.42 10.05 9.24 8.83 8.48

3.1.2 Spanwise Location: 0.7z/b

The results for the load cell performance data are shown below for a spanwise location of 0.70z/b. The 퐶퐿 results in Figure 21 show that the passive BLF is responsible for a 17.2% increase over the baseline with a 퐶퐿푚푎푥 = 1.5 at an 훼 = 29°. The passive fence is also able to delay stall and the onset of a large, unstable pitching moment until

훼 = 30° , an increase of 16° over the baseline. However, the large spike in pitching moment results in an instability in the longitudinal static stability derivative having a positive value of 0.1307/° at stall. This large ‘pitch-up’ moment is indicative of a loss of lift aft of the aerodynamic center, and its magnitude is much larger than that of the baseline case. Figure 22 shows that the passive fence produces an undesirable increase in

CD over the baseline, notably from 훼 = 14° − 18°. Additionally, the BLF reaches a

L/DMax = 8.7 at 훼 = 10°. Similarly to the fence at 0.60z/b, it provides increased L/D at lower angles of attack, but then falls below the baseline for 훼 > 13° in this scenario.

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Figure 21: 퐶퐿 푣푠. 훼 (top) and Δ퐶퐿 푣푠. 훼 (bottom) for baseline, BLF, and AFC slot at spanwise location of 0.70z/b

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Figure 22: 퐶퐷 푣푠. 훼 (top) and Δ퐶퐷 푣푠. 훼 (bottom) for baseline, BLF, and AFC slot at spanwise location of 0.70z/b

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Figure 23: 퐶푀 푣푠. 훼 (top) and CMα 푣푠. 훼 (bottom) for baseline, BLF, and AFC slot at spanwise location of 0.70z/b

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Figure 24: 퐶푀훼 푣푠. 훼 (zoomed-in, top) and L/D 푣푠. 훼 (bottom) for baseline, BLF, and AFC slot at spanwise location of 0.70z/b

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The AFC slot performs similarly to the BLF with higher 퐶휇 values even exceeding the performance gains. The lift and longitudinal static stability performance are shown to increase over the baseline as 퐶휇 is increased. The lowest 퐶휇 =

3.98% achieves a maximum 퐶퐿 = 1.39 at an 훼 = 34° (increase of 8.6% over the baseline) while delaying an unstable pitching moment increase until 훼 = 34°. Lift benefit is seen over the range of 훼 = 20° − 34°. In addition, the longitudinal static stability derivative increases to a positive 퐶푀훼 = 0.04208/° at this angle, however this is a smaller magnitude than in the passive BLF test, indicating a small instability. A 퐶휇 =

7.07% is responsible for a lift increase of 14.8% over the baseline (퐶퐿푚푎푥 = 1.47 at 훼 =

25°) and delaying stall until 훼 = 36°, an increase of 18° over the baseline case and 6° further than the passive BLF. Lastly, the highest 퐶휇 = 10.33 is responsible for a 18.8% lift increase over the baseline with a 퐶퐿푚푎푥 = 1.52 at 훼 = 25°. The two highest 퐶휇 levels show significant lift enhancement over the baseline from 훼 = 19° − 36°. Increasing the

퐶휇 by 3% provided relatively small additional lift and drag enhancement yet did not improve the stability by an appreciable amount. In fact, the unstable longitudinal static stability derivative has a higher magnitude for the 퐶휇 = 10.33% case (퐶푀훼 = 0.06657/°) as compared to the 퐶휇 = 7.07% case (퐶푀훼 = 0.04096/°). Also, as shown in Figure 22, all of the AFC slot configurations are able to avoid a large increase in drag (drag penalty) associated with the passive BLF. This is one of the critical benefits to any type of active flow control; it can be turned on/off as needed and only for the specific flight regimes that it provides benefits for. The AFC slot at this location does not show an increased L/D

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over the BLF until 훼 = 13°, which aligns with the lower drag seen by the ‘fluidic fence’ as compared to the passive fence. Additionally, the AFC slot shows a similar L/D trend to the baseline until 훼 = 19°. This coincides with the stall of the baseline wing which explains the drastic drop in L/D, while the AFC configurations remain higher until 훼 =

23°.

Table 2: Performance Summary for Baseline, BLF, AFC at Spanwise Location of 0.70z/b

Wing Configuration Baseline BLF AFC Slot AFC Slot AFC Slot

푪흁 - - 3.98% 7.07% 10.33% CLMax 1.28 1.5 1.39 1.47 1.52

휶풔풕풂풍풍° 19° 30° 35° 37° 37° 횫CLMax % - 17.20% 8.60% 14.80% 18.80% Lift Benefit Range - 17°-29° 19°-34° 19°-36° 19°-36°

Delay Unstable 푪푴휶 - 16° 21° 23° 23° 푳/푫풎풂풙 8.42 8.68 7.62 7.82 8.07

3.1.3 Spanwise Location: 0.8z/b

At a spanwise location of 0.80z/b, a passive BLF is responsible for an increased

퐶퐿푚푎푥 = 1.41 at 훼 = 32°, which is an increase of 10.2% over the baseline. As noted, this is smaller than the lift increases noted by the BLF at previous spanwise locations. Since there is less wing area outboard of the fence, there is a smaller area for which lift to be recovered and thus, a smaller maximum lift benefit. However, this smaller reference area is less susceptible to total wing flow separation and the entire region is in very close proximity to the fence, and therefore directly in the path of the fence and tip vortices. Due to this, a passive BLF at 0.80z/b is able to delay the onset of a spike in pitching moment

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until 훼 = 33°, which results in an unstable longitudinal static stability derivative of

퐶푀훼 = 0.0765 /°. This passive BLF fence location provides lift benefit for an additional

13° and delays an unstable pitching moment for 훼 = 19° as compared to the baseline. As is the case with the previous BLF locations, there is also a drag penalty from 훼 = 15° −

19°. A clear trend can be seen by referencing the changing performance of a passive BLF as the spanwise location moves closer to the tip; the maximum lift benefit decreases but the stall angle is increased. Following the pattern seen at the previous two locations, the

L/D is increased at lower angles, and then becomes lower than the baseline after 훼 =

13°.

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Figure 25: 퐶퐿 푣푠. 훼 (top) and Δ퐶퐿 푣푠. 훼 (bottom) for baseline, BLF, and AFC slot at spanwise location of 0.80z/b

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Figure 26: 퐶퐷 푣푠. 훼 (top) and Δ퐶퐷 푣푠. 훼 (bottom) for baseline, BLF, and AFC slot at spanwise location of 0.80z/b

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Figure 27: 퐶푀 푣푠. 훼 (top) and 퐶푀훼 푣푠. 훼 (bottom) for baseline, BLF, and AFC slot at spanwise location of 0.80z/b

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Figure 28: 퐶푀훼 푣푠. 훼 (zoomed-in, top) and 퐿/퐷 푣푠. 훼 (bottom) for baseline, BLF, and AFC slot at spanwise location of 0.80z/b

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For the AFC slot at a spanwise location of 0.80z/b, a 퐶휇 = 3.98% yields a

퐶퐿푚푎푥 = 1.33 or an increase of 3.52% over the baseline. Interestingly, in a pattern displayed by all three 퐶휇 tested at this location, the AFC configurations do not show either an unstable longitudinal static stability derivative or a drastic loss of lift normally associated with stall, over the range of 훼’s tested. Instead, there is a much more gradual decline in the performance at especially high angles. This drastic delay in stall angle is thought to be due to the close proximity of the fence and tip vortices due to the placement of the slot. With the two vortices in close proximity to each other, the interaction is amplified which helps to keep flow attached to much higher angles of attack. In addition, flow entrainment caused by the slot is also likely a factor in the benefits seen. Using flow visualization, this can be seen in section 3.3 Surface Flow Visualization: Fluorescent

Tufts.

As is the trend with the previous spanwise locations, increasing the 퐶휇 yields higher lift benefits. The middle coefficient of momentum can reach a 퐶퐿푀푎푥 = 1.38, representing a 7.8% increase over the baseline. Looking at Figure 25, this AFC configuration provides lift enhancement over the baseline at the same angle as the BLF, however it provides a greater enhancement sooner, and more consistent lift throughout the region in which it is beneficial. Since the loss of lift is more gradual, there is no large pitching moment spike or unstable static stability derivative to be concerned with. Lastly, the highest 퐶휇 = 10.33% just surpasses the performance of the BLF and reaches a

퐶퐿푀푎푥 = 1.416. As is the case with 퐶휇 = 7.07%, this AFC configuration produces the maximum lift at 훼 = 24°, which is notably lower than the BLF. This means that to take 48

full advantage of this flow control, a lower angle can be used without concern for stall or unstable pitching moments.

Having flow control at this high of a spanwise location in isolation does not seem to properly utilize the benefits available. Boundary layer fences that are placed near the tip routinely are used in combination with other fences more inboard on the wing. The two fences working together amplify the effectiveness of each fence individually. Similar techniques could be explored with AFC, using multiple slots and even different 퐶휇’s for each spanwise location to optimize momentum input with performance.

Table 3: Performance Summary of Baseline, BLF, AFC at Spanwise Location of 0.80z/b

Wing Configuration Baseline BLF AFC Slot AFC Slot AFC Slot

푪흁 - - 3.98% 7.07% 10.33%

CLMax 1.28 1.41 1.33 1.38 1.416

휶풔풕풂풍풍° 19° 33° >40° >40° >40°

횫CLMax % - 10.20% 3.52% 7.81% 10.60% Lift Benefit Range - 19°-33° 20°-40° 20°-40° 20°-40°

Delay Unstable 푪푴휶 - 19° 26° 26° 26° 푳/푫풎풂풙 8.42 10.94 8.84 9.283 11.70

3.2 Spanwise Dependence Performance Results

In the previous sections, each spanwise location was tested using a passive BLF or AFC slot and the results are discussed in detail. The following section looks to compare the spanwise dependence of the performance of AFC to a passive BLF. This comparison is important because applying flow control at specific spanwise locations can be used to target different performance changes. It is also important to understand if there

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is an optimal location for a single fence when considering the balance between lift performance and stability.

3.2.2 Global Lift Performance

Shown in Figure 29 is a 퐶퐿 푣푠. 훼 plot of all three BLF locations tested along with three AFC slot locations at 퐶휇 = 10.33%. This was chosen for the comparison as it achieved the highest lift performance benefits across all locations tested. It is clear from the results that the AFC slot replicates the trends seen by the BLF; namely the decrease in maximum lift and increase in stall angle as the flow control location is moved closer to the tip. This aligns well with the past studies referenced earlier [15, 20]

Figure 29: 퐶퐿 푣푠. 훼 for Baseline, BLF, and AFC (퐶휇 = 10.33%) at all spanwise locations

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Interestingly, while the BLF provides 7% greater lift benefit at a location of

0.70z/b as compared to 0.80z/b, moving the fence inboard to 0.60z/b only increases the lift benefit by an additional 2%, even though the wing experiences stall at an 훼 that is 3° lower. However, the AFC slot provides the same relative change in lift enhancement between all three spanwise locations, providing a more consistent spanwise performance.

At all locations, the AFC slot with the highest 퐶휇 provides a greater 퐶퐿푀푎푥 than the BLF, and in the case of a spanwise location of 0.60z/b, the difference is over 5% more lift.

Additionally, all three 퐶휇 tested follow a more consistent linear trend in lift benefit with spanwise location than the BLF. This means that within the spanwise locations tested, moving the AFC slot to another location yields a predictable result that follows a clear trend, rather than dropping off suddenly as is the case for the BLF. Figure 30 shows the ratio of 퐶퐿푀푎푥 for the flow control configurations normalized by the 퐶퐿푀푎푥 for the baseline.

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Figure 30: Normalized 퐶퐿푀푎푥 for all flow control configurations

One of the clear signs of the effectiveness of AFC can be seen when comparing the performance of AFC and a BLF at 0.60z/b. The AFC slot not only out performs the passive fence (23.4% lift benefit compared to 19.3%), it also delays stall considerably further. In fact, the AFC slot at this location delays the stall angle as far as the best-case scenario for the BLF (0.80z/b). By utilizing an AFC configuration, both the lift benefit and increased stall angle can be achieved while keeping the slot further inboard.

As previously mentioned, for an AFC slot at 0.60z/b and 퐶휇 = 10.33%, the lift performance is maximized over a very specific range of 훼’s. This must be recognized when considering the complete picture of the performance of the AFC slot at this spanwise location. The rapid change in lift also has implications for stability which is discussed in the following section. 52

3.2.3 Longitudinal Static Stability Performance

Additional consideration should be given to the stability characteristics of the flow control techniques at different spanwise locations, and especially the tradeoff between lift and stability performance. As previously mentioned, the baseline wing experiences a maximum longitudinal static instability at 훼 = 14° where 퐶푀훼 =

0.02325 /°. This coincides with a loss of lift aft of the aerodynamic center and causes an unstable pitch up moment on the wing. This instability is delayed considerably for all flow control configurations; however, the magnitudes are much larger once stall eventually occurs.

Figure 31: Longitudinal static stability derivatives for Baseline, BLF, and AFC (퐶휇 = 10.33%)

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Clearly, each of the AFC configurations presented are not only able to delay the

unstable static stability derivative to a higher angle than a passive BLF at the same

location, the instability itself is also smaller (Figure 31). This is critical when operating

near these limits. Having a smaller instability at a higher angle of attack lends itself to a

wider range of control for an aircraft. In addition, while the instabilities present in the

BLF cases do decrease with higher spanwise location, the rate at which they decrease is

considerably less than the instabilities seen in the AFC cases. The maximum 퐶푀훼

decreases by about 15% when a BLF is moved from 0.60z/b to 0.70z/b; for an AFC slot,

this reduction is 49%, indicating a much larger benefit to stability for the AFC as

compared to the BLF (Table 4). Specifically, for the AFC slot at 0.80z/b, there is no large

positive spike in 퐶푀훼 as is the case with other control configurations and its maximum

instability is essentially zero at a relatively low angle of attack. This means that although

the lift benefits are much lower than other locations, having a AFC slot at 0.80z/b would

provide the highest 훼 without stall, which could be a desirable attribute for some flight

operations.

Table 4: Longitudinal static stability derivative comparison for Baseline, BLF, and AFC (퐶휇 = 10.33%)

Wing Configuration Baseline BLF AFC 푪흁 - - - - 10.33% 10.33% 10.33% Spanwise Location: z/b - 0.60 0.70 0.80 0.60 0.70 0.80 Max 푪푴휶 0.02325 0.154 0.1307 0.07656 0.131 0.06657 0.009545 휶° 14° 26° 29° 32° 32° 36° 16°

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3.3 Surface Flow Visualization: Fluorescent Tufts

To better visualize the flow control mechanisms at play, fluorescent tufts were applied to the suction surface of the wing. These tufts allowed for surface flow visualization of the different AFC configurations and a comparison of the stall characteristics present. The grid of tufts used for the current study is shown in Figure 32.

Figure 32: Fluorescent tuft grid used in flow visualization

Tests were run on the baseline wing along with the AFC configurations at each spanwise location from 0° ≤ 훼 ≤ 40° with the tufts applied. The results are consistent with the load cell findings in that the baseline configuration begins to lose lift (evident from reduction of lift curve slope) and experiences more spanwise flow starting at α =

14°, as shown in Figure 33. This coincides with the large positive spike in moment for the baseline wing. For the AFC wing at 0.60z/b, as expected there is some unsteady flow and entrainment near the slot, however the wing shows fewer signs of spanwise flow 55

except for at the trailing edge near the root. A similar flow pattern is seen for the AFC wing at 0.70z/b. For the AFC slot at 0.80z/b, there appears to be slightly more spanwise flow in the regions inboard of the fence. However, all of the AFC configurations show a clear pattern; the flow outboard of the slot remains streamwise with no spanwise components. At this angle of attack, the baseline wing is actually producing a higher 퐶퐿 than the AFC wings even though the flow looks very similar (Table 5). This is due to the negative lift force created by the air coming out of the slot and is discussed further in section 3.4 AFC Slot Effect on Global Performance.

Next, at 훼 = 19°, the baseline wing is shown to have some flow separation for the region 0.50z/b - 1.0z/b and the flow near the root is almost entirely spanwise (Figure 34).

This corresponds with stall, shown in the load cell findings, in which some of the flow across the wing is no longer attached. The AFC wings at 0.60z/b and 0.70z/b, on the other hand, do not show any flow separation, although the spanwise flow has increased inboard of the slot and near the trailing edge (where the slot does not extend). On the

0.80z/b wing, there is slightly more spanwise flow present, however it has still not separated. By looking at the load cell findings in Table 6, the difference is clear; while the baseline wing stalls and the lift curve essentially becomes flat, the AFC wings continue to produce lift. This is also the angle at which major lift performance gains begin to occur.

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Figure 33: Tufts visualization for baseline, AFC: 0.60z/b, AFC: 0.70z/b, AFC: 0.80z/b

(퐶휇 = 10.33%) at 훼 = 14°

Table 5: Aerodynamic forces for wing configurations at 훼 = 14°

Wing Configuration Baseline AFC Slot AFC Slot AFC Slot Spanwise Location - 0.60z/b 0.70z/b 0.80z/b

푪흁 - 10.33% 푪푳 1.101 1.016 1.047 1.054 푪푫 0.1494 0.1147 0.1389 0.1149 푪푴 -0.0404 -0.0936 -0.0230 -0.0332 푪푴휶 0.0162 -0.0016 0.0053 0.0109

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Figure 34: Tufts visualization for baseline, AFC: 0.60z/b, AFC: 0.70z/b, AFC: 0.80z/b (퐶휇 = 10.33%) at 훼 = 19°

Table 6: Aerodynamic forces for wing configurations at 훼 = 19°

Wing Configuration Baseline AFC Slot AFC Slot AFC Slot Spanwise Location - 0.60z/b 0.70z/b 0.80z/b

푪흁 - 10.33% 푪푳 1.235 1.241 1.263 1.26 푪푫 0.3399 0.2008 0.2307 0.2659 푪푴 -0.0199 -0.0910 -0.0098 -0.0189 푪푴휶 -0.0308 -0.0008 -0.0044 0.0009

Looking at both configurations at much higher angles of attack, there are still clear differences in the flow over the wings in terms of separation and flow direction. At an 훼 = 30°, the flow over the baseline wing is essentially entirely separated as evidenced by the ‘fanned’ out tufts across the entire wing (Figure 35). Contrastingly, the AFC wing at 0.60z/b shows far less separation; only a small portion near the trailing edge that is aft of the AFC slot. This is likely because the AFC slot does not extend all the way to the

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trailing edge. Also, there is a small region of well aligned, attached flow near the leading edge, directly outboard of the slot. This flow has been entrained and attached to the surface due to the counter rotating vortex pair formed by the fence and tip vortices. In addition, while the flow is heavily spanwise inboard of the ‘fence’, there is not much separation present. The AFC slot at 0.70z/b displays similar characteristics as the fence at 0.60z/b, with a slightly larger region of streamwise and attached flow present.

However, there seems to be slightly more separated flow inboard of the slot and near the trailing edge behind the slot. Referring to the load cell findings (Table 7), the slot at

0.70z/b does produce less lift than a slot at 0.60z/b (for 훼 = 30°). This region of flow is likely indicative of that. Also, there is less flow that is separated near the trailing edge in the vicinity of the slot for this AFC slot location. The AFC slot at 0.80z/b shows some spanwise flow and separation directly inboard from ~0.60 < z/b < 0.80. Further inboard of that region shows flow to be more spanwise. However, the flow is very clearly oriented in the streamwise direction for the entire region of the wing that is outboard of the slot. This is likely due to the proximity of the fence and tip vortices as well as the flow entrainment of the slot (spanwise flow directly inboard of slot). Since the fence and tip are so close, the vortex interaction is stronger which helps to maintain attached flow in the region between them to higher angles of attack.

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Figure 35: Tufts visualization for baseline, AFC: 0.60z/b, AFC: 0.70z/b, AFC: 0.80z/b (퐶휇 = 10.33%) at 훼 = 30°

Table 7: Aerodynamic forces for wing configurations at 훼 = 30°

Wing Configuration Baseline AFC Slot AFC Slot AFC Slot Spanwise Location - 0.60z/b 0.70z/b 0.80z/b 푪흁 - 10.33% 푪푳 1.258 1.58 1.473 1.387 푪푫 0.7527 0.7659 0.7516 0.733 푪푴 -0.0147 -0.2337 -0.0829 -0.0933 푪푴휶 -0.0080 -0.1007 -0.0039 -0.0069

When an even higher angle of attack, 훼 = 36° (Figure 36), is considered, the baseline wing shows that much of the flow is separated. This is consistent with the pattern of the wing at 훼 = 30°. The AFC slot at 0.60z/b has already stalled and experienced an unstable pitching moment spike prior to this angle; however, the flow has not separated as much as the baseline case. Most of the flow that is not separated is heavily spanwise, and at this angle, the wing is not producing considerably more lift than

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the baseline case, however the effects of the AFC slot are still present (Table 8). Near the slot, on both sides, flow entrainment is visible as tufts are drawn towards the slot in the middle. At this 훼, the AFC configuration with the slot at 0.70z/b has not stalled yet; this is evident by the region of attached flow directly outboard of the slot at the leading edge.

Inboard of this, the flow over the wing shows signs of separation from ~0.35z/b –

0.70z/b. The large region of separated and spanwise flow is predictive of the upcoming stall at 훼 = 37°. The AFC slot at 0.80z/b shows an even greater region of attached, streamwise flow. The entire 20% span outboard of the ‘fence’ shows attachment even at this high angle of attack. This is likely due to the interaction and proximity of the fence and tip vortices as well as flow entrainment due to the slot. Inboard of the slot, there is spanwise flow from 0.60 < z/b < 0.80 and closer to the root from 0.30 < z/b < 0.60 the flow is more clearly separated.

Figure 36: Tufts visualization for baseline, AFC: 0.60z/b, AFC: 0.70z/b, AFC: 0.80z/b (퐶휇 = 10.33%) at 훼 = 36°

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Table 8: Aerodynamic forces for wing configurations at 훼 = 36°

Wing Configuration Baseline AFC Slot AFC Slot AFC Slot Spanwise Location - 0.60z/b 0.70z/b 0.80z/b

푪흁 - 10.33% 푪푳 1.28 1.369 1.436 1.372 푪푫 0.9657 0.9294 0.9422 0.8819 푪푴 -0.0823 -0.1675 -0.1314 -0.1497 푪푴휶 -0.0100 -0.0136 -0.0069 -0.0161

One of the unique findings of the AFC slot performance at a spanwise location of

0.60z/b was the non-linear lift response seen at 훼 = 30° − 32°. As explained previously in Figure 17, the lift curve slope drastically increases at these angles and provides an additional lift benefit of 10% over the baseline. Shown in Figure 37 are two flow visualizations from the AFC slot at 0.60z/b; one image is taken before the lift spike and the other is taken in the region of increased lift. There is a clear difference in the flow directly outboard of the slot that corroborates the load cell findings (Table 9). While the wing on the left (훼 = 29°) experiences some separation near the leading edge and spanwise flow closer to the trailing edge, the wing on the right (훼 = 31°) shows a region of attached and streamlined flow at the leading edge. In addition, the flow near the trailing edge is much less spanwise than at the lower angle of attack. This difference in flow field can be attributed to the interaction between the fence and the tip vortex. At

훼 = 31° the flow tufts show distinct patterns that indicate the presence of these two vortices. First, the flow directly outboard of the slot is attached at the leading edge. As the vortex moves towards the trailing edge the tufts fan out slightly indicating the

‘spreading out’ of the vortex. Near the tip, the tufts show flow in the spanwise direction

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and the pattern indicates a cone-like vortex forming at the fence and spreading towards the tip. This agrees with the simulations conducted by Solfelt [17].

Figure 37: Tufts visualization for AFC: 0.60z/b (퐶휇=10.33%) at 훼 = 29° & 훼 = 31°

Table 9: Aerodynamic forces for AFC: 0.60z/b at 훼 = 29° and 훼 = 31°

Wing Configuration AFC Slot Spanwise Location 0.60z/b 휶 29° 31°

푪흁 10.33% 푪푳 1.424 1.567 푪푫 0.6541 0.7659 푪푴 -0.1330 -0.2439 푪푴휶 -0.1007 -0.0104

3.4 AFC Slot Effect on Global Performance

For this form of active flow control, the effect of the slot itself must also be considered when looking at the complete picture regarding the performance benefits. In this case, at higher 퐶휇, the slot imparts a significant force on the airfoil just from the air 63

exiting the slot. These forces must be considered as a part of this AFC method currently, until future methods and optimizations can either limit their effects or investigate another way to utilize these forces.

It helps to first visualize why this is occurring. Shown in Figure 38 is a diagram of how air exits the slot at a given spanwise location. As clearly seen, most of the air exits from the suction side, and thus it imparts a force in the negative lift direction.

Figure 38: Visualization of Forces Caused by AFC Slot

This effect is much more pronounced at lower angles of attack. However, the advantage of AFC is that since it does not provide benefits in this condition, it can simply be turned off with no negative consequences. Shown in Figure 39 are the results of a

‘wind-off’ test to measure the 퐶퐿 and 퐶퐷 caused by the slot. It should be noted that with a freestream velocity, the exit flow out of the slot will have a different distribution and likely create slightly different 퐶퐿 and 퐶퐷. To calculate these coefficients, at each 퐶휇, the forces in the X and Y direction were recorded at 훼 = 0° with the AFC activated and no free-stream velocity. These forces were then mathematically rotated to find the 퐶퐿 and 퐶퐷 at all other angles 0° < 훼 < 40°. As seen in Figure 39, there is considerable negative lift

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being generated, even at higher angles of attack. For example, the highest 퐶휇 = 10.33% has a negative 퐶퐿 ~ − 0.08 at angles of attack of ~ 35°, which is the angle of the greatest benefit over the baseline (for AFC slot at 0.70z/b).

Figure 39: 퐶퐿 푣푠. 훼 (left) and 퐶퐷 푣푠. 훼 (right) Forces for Wind off Tests

The slot also imparts a small amount of negative drag (or thrust) on the airfoil.

This is due to the slot blowing normal to the wall and the aft parts of the airfoil slot having a small curvature which causes a small component of the air to be in the positive thrust direction. This benefit is mainly seen at higher angles of attack since as the airfoil rotates, the slot is pointing in a direction in which the negative lift produced at lower angles of attack is now negative drag (positive thrust). Figure 39 shows that this benefit varies from ~퐶퐷 = 0 to a maximum benefit of 퐶퐷 = −0.07 depending on the coefficient of momentum.

Looking back on the performance of the AFC slot, when considering the negative impact of the slot on the lift performance, the results are even more drastic. At all

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locations, the 퐶휇 = 10.33% AFC slot improves the lift performance over the baseline by varying degrees. However, that result (Figure 30) essentially described the net performance benefit. For example, at 0.70z/b, the AFC slot is able to provide a net benefit of Δ퐶퐿푀푎푥 = 0.24 over the baseline while the BLF shows a Δ퐶퐿푀푎푥 = 0.22 over the baseline. While this performance is comparable, considering the gross effect of the slot shows a much larger improvement for the AFC slot. As seen in Figure 39, at 훼 =

25°, the slot creates 퐶퐿 = −0.084 for a 퐶휇 = 10.33%. If this is added to the previous

퐶퐿푀푎푥 for the AFC slot, this increases the benefit to Δ퐶퐿푀푎푥 = 0.324 or an increase of

25.3% over the baseline. Shown below is Figure 40 detailing the normalized 퐶퐿푀푎푥 for each type of flow control which now includes both the net and gross benefits of AFC for comparison. When accounting for the forces out of the slot, both the 퐶휇 = 7.07% and

퐶휇 = 10.33% are shown to improve upon the BLF performance over the baseline. Across all configurations and locations, the gross lift benefit of AFC is shown to be 3-5% higher than the net benefit presented earlier in this report.

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Figure 40: Normalized 퐶퐿푀푎푥 for all flow control configurations with slot force corrections

In addition to the considerations as relating to the present study, the forces imparted by the slot could also be manipulated as a means to target certain aerodynamic characteristics. Changing lift or drag on an aircraft in flight is presently done with mechanical control surfaces that change the shape or curvature of the wings, elevators, and rudders. This is what allows aircraft to maneuver; e.g. by increasing lift on one wing, one can induce a bank turn, or by deploying flaps on the wings or elevators, the lift produced can be considerably increased at slower speeds such as takeoff or landing. By activating the AFC slot on one wing only, this effectively creates an imbalance in the lift produced between the two wings – the same as an aileron, except without the need for any mechanical parts. These mechanical control surfaces can be large and add enormous

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amounts of weight to the aircraft. In some cases, the size of the control surface is set by failure cases rather than designed for optimal usage. For a dual engine aircraft, the vertical tail is designed so that the tail/rudder is able to counteract the moment of an engine failure, yet when not in use creates a large amount of drag and is an excess weight to carry. Using AFC as a means to generate aerodynamic forces on the wings would help greatly reduce the need of these large control surfaces. Another possibility is to change the slot shape/direction of blowing in such a way that specific aerodynamic performance can be targeted. By changing the shape and orientation of this form of active flow control, control surfaces could be designed to be much smaller in size, or even eliminated entirely.

3.5 Comparison of Present Study to Previous Studies

It is necessary to look at the present study in terms of previous work done regarding the spanwise location of passive BLF’s. This section includes a comparison to a previous study from Salmi [15]. In 1952, Salmi studied the effect of a BLF on a NACA

631-A012 airfoil with twist and camber. The wing model had a leading-edge sweep Λ =

45°, a taper ratio 휆 = 0.45, and was tested at a 푅푒푐 = 2M. Although the study was done with a very different wing and flow regime, a comparison of the trends is possible. Both studies used a passive BLF with a height of 0.60t (60% max thickness of the airfoil).

The data from the previous study has had the same swept wing transformations applied to it as the present work [26, 27]. This was done to more directly compare the data to the current study. In addition, all the data sets were adjusted for the induced angle

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of attack and plotted against 훼 − 훼푖. This is a means by which to account for the different aspect ratios of the wings.

퐶 (3) α = 퐿 i 휋 푒 퐴푅

While the trends seen with spanwise location are similar, the performance gains for the current study are notably higher. One explanation for this is that the current study uses a rectangular wing with a constant chord as opposed to a tapered wing used in the historical study. Outboard of the fence, there is a larger local wing area able to generate lift and by keeping flow attached in that area, the lift performance benefits are amplified as compared to a tapered wing.

Figure 41: Spanwise dependence of 퐶퐿푣푠. 훼 푐omparison of Salmi [15] and present study (BLF)

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Figure 42: Spanwise dependence of 퐶퐿푣푠. 훼 푐omparison of Salmi [15] and present study (AFC)

The trends from all three studies match the expectations for both the BLF and AFC flow control configurations. As explained earlier, the noted difference in the magnitude of performance gains can be explained due to the wing planform type (tapered vs. rectangular) for the different studies. However, in the case of Salmi, the BLF at a spanwise location of 0.575z/b provides no increase in the stall angle over the baseline but does increase the maximum lift coefficient by 6.5%. The fence at this location provides considerable lift benefits from 훼 − 훼푖 = 11° − 27°. Moving the fence outboard to

0.80z/b extends the stall angle by 2°, but only provides lift benefits from 훼 − 훼푖 = 14° −

27° and a maximum lift benefit of 4.6% over the baseline.

In the current study, similar trends are seen with the passive BLF. Lift benefits are apparent for 훼 − 훼푖 = 7° − 20° for a BLF at 0.60z/b and the stall angle is increased 70

by 훼 − 훼푖 = 8°. Moving the fence outboard to 0.70z/b decreases the maximum lift

benefit from 19.3% to 17.2% but increases the stall angle by an additional 훼 − 훼푖 = 3°.

Moving the BLF further to 0.80z/b causes a large drop in maximum lift benefit to 10.2%

over the baseline. As discussed in the context of previous studies on BLF [14, 15, 20],

there is a noted compromise between lift performance and stall angle or pitching moment.

In agreement with those studies, the current work shows that the optimal location for a

passive BLF is close to a spanwise location of 0.70z/b when considering both lift and

stability performance. For the AFC Slot, the maximum lift benefit follows a more linear

trend as presented in Figure 30. This form of flow control matches the trends of the

historical and present studies of a passive BLF. However, based on the performance,

there is not one spanwise location that presents itself as an ‘optimal location’. All of the

locations follow a consistent trend in the tradeoff between lift and stability performance.

It is necessary then to consider the specific usage cases and performance objectives in

order to select the best location for the AFC.

Table 10: Performance Comparison to Historical Studies

Salmi (1952) Present Study Present Study

Wing Configuration Passive BLF Passive BLF AFC (퐶휇 - 10.33%) Spanwise Location: z/b 0.575 0.8 0.6 0.7 0.8 0.6 0.7 0.8

횫푪푳푴풂풙% 6.50% 4.60% 19.30% 17.20% 10.20% 23.44% 18.75% 11.5% Baseline Stall Angle 23° 19° 19° Flow Control Stall Angle 23° 25° 27° 30° 33° 33° 37° >40°

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Chapter 4: Conclusion

The current study investigated the effect of changing slot configuration on the performance of active flow control at low speeds for a swept wing. This study was primarily concerned with studying post-stall performance and the ability to provide lift enhancement and delay stall to a higher angle, while reducing momentum input. A

NACA 643-618 airfoil was tested at a chordwise Reynolds number of 100,000 from 0° ≤

훼 ≤ 40°. Boundary layer fence (BLF) and active flow control (AFC) slot configurations varied the spanwise location of flow control to determine the impact on stability characteristics and lift. Load cell testing and surface flow visualization were used to determine the best configuration as well as to better understand the mechanisms of the flow.

Passive BLF testing revealed findings consistent with historical trends. Moving the flow control location outboard from 0.60z/b to 0.80z/b decreased the 퐶퐿푀푎푥 over the baseline from 19.3% to 10.2%. The occurrence of an unstable static stability derivative was delayed by 13° for the fence closest to the root and increased to 19° for the fence near the tip. There is a considerable drop in performance gains when moving the fence from 0.70z/b to 0.60z/b; lift is only improved by 2% yet stall occurs 3° earlier. Taking this finding into consideration, 0.70z/b seems to be a location for what can be regarded as an optimal fence location. Drag penalties were seen in all cases, especially during regions

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of increased lift. This reiterates the negative consequences of a passive BLF; it cannot be de-activated when in an off-design or non-optimal flight condition. This method of flow control carries an obvious weight and drag penalty at all times. In addition, if an aircraft is in side-slip, the oncoming freestream air would directly hit the surface of the fence and could create dangerous and complex roll and yaw moments and other instabilities for the aircraft

AFC through a wall-normal slot was found to follow the same general trends as the BLF but exceeds the lift performance at the highest 퐶휇 = 10.33 by increasing the

퐶퐿푀푎푥 as high as 23.4%. As spanwise location was increased (closer to the tip), the lift benefits did drop, but followed an almost linear trend. Unlike the BLF, there was no single location that showed a significant relative drop in performance when moving the flow control. This is important as it shows there is no clear optimal location, and the considerations that dictate the positioning of the flow control depend more on the performance advantages desired. The stall angle and onset of an unstable pitching moment and static stability derivative were also delayed considerably as compared to the

BLF. The dependence on coefficient of momentum (퐶휇) was most prevalent when looking at lift and drag performance as there was a monotonic increase with lift as 퐶휇 was increased from 3.98% to 7.07% to 10.33%. Considerations should be made for the high

퐶휇 needed for optimal benefits and the effect the slot itself has on performance.

Flow visualization revealed the reasoning behind some of the unique findings for the AFC slot at 0.60z/b. This configuration showed a non-linear response at 퐶휇 =

10.33% which caused a drastic increase in lift, then almost immediately stalling. Tufts 73

showed that the fence vortex, in combination with the tip vortex, was able to entrain flow to the surface and maintain attached streamwise flow in the direct vicinity of the slot.

This is a critical finding the helps to explain and corroborate the physical mechanisms thought to play a role in both BLF and AFC control; suspending the spanwise flow component and the counter rotating vortex pair (fence vortex and tip vortex) that helps to entrain flow to the surface.

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Chapter 5: Future Work

Although this study helps to further the understanding of an active flow control replication of a boundary layer fence by looking at the performance changes with spanwise location, further research should be done to investigate more about the physical mechanisms behind the flow control, as well as testing this flow control method on other types of wings and in other flight regimes. This will allow for a much more complete picture of the capabilities of this active flow control method and provide greater insight into how this technology may be used in the design process.

1. In combination with an optimal spanwise slot location as investigated

presently, different slot configurations, or other methods of active flow control

should be studied to investigate potentially more effective flow control

methods. This can include segmented slots, angled slots (as opposed to wall-

normal blowing), discrete jets, or even fluidic oscillators. These options may

provide other mechanics of control not seen by using a wall normal

continuous slot.

2. Investigate pulsed blowing to reduce the energy needed for the performance

benefits. Steady blowing generally requires much higher 퐶휇 than other types

of AFC. Pulsed blowing could also target natural flow instabilities to enhance

certain aspects of the flow control as well as reducing the momentum input.

3. Test the effect of the number of AFC slots on the performance of flow control.

Past studies using BLF have shown that using multiple fences can be even

more effective than a single fence. By recognizing the different benefits of 75

AFC slots at different spanwise locations, combining numerous slots could

provide an even larger overall performance benefit for a given scenario. This

could include different size slots at different locations or only activating a

single slot vs. multiple slots when necessary.

4. Investigate the feasibility for an AFC slot to be used as a means to reduce the

size or replace mechanical control surfaces as the primary way to maneuver an

aircraft. This would take much more additional testing and reliability

assurances, but by eliminating the extra weight and drag caused by large

mechanical flaps, slats, and ailerons, aircraft could be designed to be more

efficient and maneuverable.

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