DEGREE PROJECT IN MATHEMATICS, SECOND CYCLE, 30 CREDITS STOCKHOLM, SWEDEN 2017

An Experimental Study of the High- Lift System and Wing-Body Junction Wake Flow Interference of the NASA Common Research Model

DESIRÉE BRUNDIN

KTH ROYAL INSTITUTE OF TECHNOLOGY SCHOOL OF ENGINEERING SCIENCES

An Experimental Study of the High- Lift System and Wing-Body Junction Wake Flow Interference of the NASA Common Research Model

DESIRÉE BRUNDIN

Degree Projects in Systems Engineering (30 ECTS credits) Degree Programme in, Aerospace Engineering, (120 credits) KTH Royal Institute of Technology year 2017 Supervisor at NASA Ames Research Center: Kurtis R. Long Supervisor at KTH: Per Engvist Examiner at KTH: Per Engvist

TRITA-MAT-E 2017:33 ISRN-KTH/MAT/E--17/33--SE

Royal Institute of Technology School of Engineering Sciences KTH SCI SE-100 44 Stockholm, Sweden URL: www.kth.se/sci

Abstract

This thesis investigates the turbulent flow in the wake of the wing-body junction of the NASA Common Research Model to further reveal its complex vortical structure and to contribute to the reference database used for Computational Fluid Dynam- ics validation activities. Compressible flows near two wall-boundary layers occurs not only at the wing-body junction but at every control surface of an airplane, therefore increased knowledge about this complex flow structure could potentially improve the estimates of drag performance and control surface efficiency, primarily for minimizing the environmental impact of commercial flight. The airplane model is modified by adding an inboard flap to investigate the influence from the deflection on the vorticity and velocity field. Future flap designs and settings are discussed from a performance improvement point of view, with the investigated flow influence in mind. The experimental measurements for this thesis were collected using a Cobra Probe, a dynamic multi-hole pressure probe, for Reynolds numbers close to one million based on the wing root chord. A pre-programmed three-dimensional grid was used to cover the most interesting parts of the junction flow. The facility used for the tests is a 120 cm by 80 cm indraft, subsonic wind tunnel at NASA Ames Research Center’s Fluid Mechanics Lab, which provides an on-set flow speed of around Mach 0.15, corresponding to approximately 48 m/s.

Keywords: Common research model, wind tunnel experiments, subsonic flow, wing-body junction, commercial aircraft, Cobra Probe measurements, high-lift con- trol systems, flow interference, drag estimate.

Abstrakt

Den här avhandlingen undersöker det turbulenta flödet runt övergången mellan flyg- planskropp och vinge på en NASA Common Research Model för att vidare utforska den komplexa, tredimensionella strukturen av flödet och bidra till NASA’s officiella databas för jämförelser med simulerade flöden. Kompressibla flöden nära tvåväggs- gränsskikt uppkommer inte bara vid övergången mellan flygplanskropp och vinge utan även vid varje kontrollyta på ett flygplan. Ökad kunskap om flödets beteende vid sådana områden kan därför bidra till en bättre uppskattning av prestanda och effektivitet av kontrollytorna och flygplanet i sin helhet, vilket kan bidra till minskad miljöpåverkan från kommersiell flygtrafik. Flygplansmodellen är modifierad genom montering av en vingklaff på den inre delen av vingen, detta för att undersöka hur olika vinklar på klaffarnas nedböjning påverkar flödets struktur och hastighetsfält. Framtida klaffdesigner och inställningar för ökad prestanda diskuteras även utifrån denna påverkan. Mätningarna i vindtunneln gjordes med en Cobra Probe, ett dynamisk tryckmät- ningsinstrument, speciellt designad för turbulenta och instabila flöden. Reynold- snumren som generades av den subsoniska, indrags-vindtunneln var ungefär en miljon baserad på vingrotens längd, vilket motsvarar knappt en tiondel av normala flygförhållanden för samma flygplansmodell.

Preface

This thesis is the result of my Master’s degree project at the Department of Ex- perimental Aero-Physics at NASA Ames Research Center during the spring 2017. The work concerns collection of three-dimensional wind tunnel data in the wake of the wing-body junction region of the NASA Common Research Model (CRM). The collected data is included in the official NASA CRM wind tunnel database, which primary purpose is to provide a reliable foundation for Computational Fluid Dynamics (CFD) validation studies. This thesis is divided into eight chapters. Chapter one includes a brief motivation to the project and an introduction to the CRM. Chapter two presents the history and design of the Boeing 777, the aircraft frame which the CRM is based on. Chapter three acts as an introduction to high-lift control systems in order to get familiar with some of the modern commercial aircraft flap designs and functions. Chapter four describes the turbulent flow structure from a theoretical point of view and also includes a short description of CFD methods and why they are not able to fully predict the flow behaviour. The experimental approach is covered in chapter five where the model, test facility and conditions are thoroughly explained. Chapter six includes the experimental results and chapter seven estimates some aerodynamic properties and optimize them with respect to flap settings. Chapter eight covers the final conclusions and mentions potential future work to be carried out on this topic.

Acknowledgments

I offer my sincerest gratitude to my supervisor Kurtis R. Long at the Fluid Me- chanics Lab at NASA Ames Research Center, for all the support, inspiration and encouragement. Kurt is by far the best mentor anyone could ever ask for and it has been a great honor to learn from him during this project. I want to thank Porsche Parker and everyone else at the internship office at NASA Ames for giving me the opportunity to come to a NASA center and learn from the very best. They have provided me with help and support in all aspects of this internship and always gone that extra mile to make everyone feel at home. A huge thank you to everyone at the Fluid Mechanics Lab who have helped me with small and big tasks during my project; Rabindra D. Mehta, Barry Porter, David Keil, Joseph Burces and Eduardo Ramirez. Without your contributions it would not have been possible to complete this thesis. I especially want to thank Witold Koning at NASA Ames Aeromechanics Branch for providing CFD support and interesting discussions throughout my project. Finally, I wish to thank my supervisor at KTH, Per Enqvist, for continuous feedback and inspiration. Thank you all.

Contents

1 Introduction 1 1.1 Nomenclature ...... 1 1.2 Abbreviations ...... 2 1.3 Objective ...... 2 1.4 NASA Common Research Model ...... 3 1.5 Experimental Method ...... 3

2 Boeing 777 Overview 4 2.1 Background ...... 4 2.2 High-Lift Design ...... 4 2.2.1 777 Leading-edge slats ...... 5 2.2.2 777 Trailing-edge flaps ...... 5 2.3 Future ...... 6

3 High-Lift Controls in General 9 3.1 Purpose ...... 9 3.2 Function ...... 10 3.3 Designs ...... 11 3.3.1 Trailing-edge devices ...... 12 3.4 Challenges ...... 14 3.5 Future Development ...... 15

4 Description of the Flow 16 4.1 Turbulent Flow Structure ...... 16 4.2 Simulating the Flow ...... 18 CONTENTS

5 Experimental Approach 20 5.1 Model Description ...... 20 5.2 Facility ...... 21 5.3 Data Collection ...... 22 5.4 Correction Methods ...... 25

6 Experimental Results 29 6.1 Velocity Field ...... 29 6.2 Downwash ...... 29

7 Aerodynamic Performance 35 7.1 Drag estimate ...... 35 7.2 Lift estimate ...... 37 7.3 Optimization ...... 38

8 Discussion 42 8.1 Future Work ...... 43

9 Summary 44 Chapter 1

Introduction

1.1 Nomenclature

CD – Drag coefficient CL – Lift coefficient M – Mach number α – Angle of attack (◦) ◦ δf – Flap deflection ( ) Re – Reynolds number ÆR – Wing aspect ratio 2 Sref – Wing reference area (m ) cref – Wing mean chord length (m) croot – Wing root chord length (m) b – Wing span (m) ρ – Air density (kg/m3) q – Dynamic pressure (Pa) V∞ – Mean free stream velocity (m/s) u – Downstream velocity (m/s) v – Lateral velocity (m/s) w – Vertical velocity (m/s) U – Downstream velocity ratio V – Lateral velocity ratio W – Vertical velocity ratio

1 2 CHAPTER 1. INTRODUCTION

1.2 Abbreviations

NASA – National Aeronautics and Space Administration ARC – Ames Research Center FML – Fluid Mechanics Lab BDAS – Basic Data Acquisition System CRM – Common Research Model CFD – Computational Fluid Dynamics AIAA – American Institute of Aeronautics and Astronautics TFI – Turbulent Flow Instrumentation TE – Trailing-Edge LE – Leading-Edge BL – Boundary Layer WBT – Wing/Body/Tail WBTF – Wing/Body/Tail/Flap

1.3 Objective

One of the oldest, unsolved interference problems in aeronautics is the interference flow between the wings and the fuselage of a fixed-wing airplane. Large regions of separated flow are developed, creating a flow structure that is highly anisotropic, three-dimensional and time dependent. The main reason behind its complexity is the two-wall boundary layer interaction, most turbulence models today assume an isotropic flow and are therefore not applicable even for a single boundary layer re- gion. Experimental data for the wing-body junction will increase the knowledge about this flow structure and thus, help provide a foundation for future CFD mod- els. This experimental study aims to contribute to the NASA CRM wind tunnel database for CFD validation by taking velocity measurements in the wake of the wing-body junction of the NASA CRM, a structural model of a Boeing 777. Flow affected by two-wall boundary layer interaction can also be found near the control surfaces on an airplane, this study will therefore investigate how the junction flow is influenced by the high-lift system. Inboard, trailing-edge devices are believed to significantly influence this flow, and since simulations cannot yet predict the flow at the junction, experimental data could bring new and valuable information about this influence. The most critical parts of a flight are the takeoff and landing seg- ments, during which the high-lift system is primarily utilized. Increased knowledge about how different settings change the lift and drag characteristics can potentially yield better estimates of performance and which settings that are optimal under what conditions. 1.4. NASA COMMON RESEARCH MODEL 3

1.4 NASA Common Research Model

Upon requests for a common, modern and public airframe for CFD validations, NASA initiated the CRM in 2007. It was believed to greatly simplify data compar- isons and collaboration between research institutes if a common aircraft model was used. The CRM geometry is based on the Boeing 777, a large transport aircraft that is widely used by airline companies all around the world, and still subject to new model releases such as the 777X family in 2020. The airframe was chosen based on three desired properties, the model should be based on a transonic transport air- liner designed for a cruise speed of Mach 0.85, a nominal lift coefficient of 0.5 and Reynolds number around 4 million per reference chord. Previous research involving the CRM includes all the AIAA drag prediction workshops since the model was first initiated, and a variety of other workshops, such as high-lift prediction. The CRM have also been subject of aerodynamic shape optimization, aerostuctural optimization and CFD simulations.

1.5 Experimental Method

The 3%, semi-span CRM was mounted in the 120 cm by 80 cm test section of the indraft, subsonic wind tunnel at the Fluid Mechanics Lab (FML) at NASA Ames Research Center (ARC). Measurements were taken at a three-dimensional grid in the wake of the wing-body junction for free stream wind speeds of around Mach 0.15, or 48 m/s. Two different model configurations were tested – WBT and WBTF, at zero angle of attack. The device used for measuring the velocity field was a Cobra Probe, a dynamic multi-hole pressure probe from Turbulent Flow Instrumentation (TFI), designed for measuring three-component velocities in unsteady, turbulent flows. The data was collected using the Basic Data Acquisition System (BDAS) and processed using Matlab and Tecplot 360. Chapter 2

Boeing 777 Overview

2.1 Background

The Boeing 777 program was the first all new commercial jet aircraft to be designed by Boeing since the 757 and 767 families in 1980. The design started as a derivative of the 767 with the purpose of filling the size gap between the 767 and 747. However, customer requirements influenced the design process and a completely new family of aircraft was developed instead. This new family entered service in 1995 and included several models, designed for flights ranging between 5300 and 8600 nautical miles. The two high-by pass ratio turbofan engines that power the 777 are the largest engines ever made, the nacelle has the same diameter as the fuselage of a Boeing 757. The engines can be manufactured by any of the world’s leading jet engine companies, such as General Electric (GE), Pratt and Whitney or Rolls-Royce [1]. The 777 is widely used by airline companies all over the world and an upgraded version, the 777X, is planned to enter service in 2020. Since the range of an average flight with a 777 is medium-to-long, improvements made for this airframe have the potential of reducing the environmental impact more than improvements on a short-ranged aircraft. This is one of many reasons why the 777 was chosen as the common research model.

2.2 High-Lift Design

The optimal high-lift system should provide minimum takeoff drag while producing sufficient lift to meet the approach speed requirement. The 777’s high-lift system features the best technologies used on the 757 and 767, modified to meet customer

4 2.2. HIGH-LIFT DESIGN 5 requirements [2]. Customer requirements include demands on payload capabilities, range, cruise speed, airport compatibility etc. Other inputs that need to be con- sidered when designing a high-lift system are for example design objectives and regulatory and producibility requirements. Generally, these demands lead to a ba- sic wing design that determines the target design parameters CL,max and L/D. The control settings are then determined so that the former is maximized during landing and the latter during takeoff. These parameters are used as input in the high-lift design process [1]. The high-lift wing flow structure was too complex for CFD alone around the time the 777 was designed. A combination of two- and three-dimensional computer codes were used together with experimental wind tunnel tests to decide which design that best satisfied the requirements. The need for an efficient and robust computer code for testing high-lift design set the direction for future CFD development at the time when Navier-Stokes and Euler codes were not mature enough for the complex configurations that needed to be tested [1].

2.2.1 777 Leading-edge slats The leading-edge device on the 777 is the same as on the 757, a three position, single-slotted slat system both inboard and outboard of the wing, with a sealed takeoff and a slotted landing position. The inboard slat is tapered between the side of body and the engine strut and the outboard slat system consists of six spanwise, constant-chord panels ranging from outboard the nacelle chine almost all the way to the wing tip. The chords varies from 9% of the local trapezoidal wing chord at the side of body up to 33% of the chord at the wing tip. The maximum deflection settings are around 35◦ inboard and 31.6◦ outboard [2]. The slat settings are extremely sensitive to positioning in order for the flow not to separate immediately at the leading-edge. When the 777 slats were designed, two-dimensional CFD codes were used for evaluating the shape of the device and the rough position for each setting. Once a basic design was determined, the three- dimensional optimization process for the exact positioning was carried out by wind tunnel tests for large Reynolds numbers.

2.2.2 777 Trailing-edge flaps The design of the trailing-edge device was strongly influenced by landing require- ments. The approach speed limits put demands on the maximum lift required and the attitude was constrained in both directions. Too high attitude upon touchdown would cause the aft of the body to hit the ground, and too low attitude would instead cause the plane to land on the nose wheel. 6 CHAPTER 2. BOEING 777 OVERVIEW

The challenges faced when designing the trailing-edge high-lift system concerned mainly the strict dimension limits posed by the wing. The percentage of the wing that can carry control surfaces is a lot smaller than on most aircraft. The airfoil is also designed with an unusual high cusp in order to provide good cruise performance, which puts strict limits on the maximum allowed depth of the flaps making the leading-edge design a challenge. The inboard aft flap is deflected on hooked tracks below the aft end of the main flap, which improves the takeoff L/D thanks to the provided Fowler motion at low flap angles. Wind tunnel experiments were conducted in order to create a baseline to which CFD could be applied for variations in the configuration. A wide range of devices were tested both inboard and outboard, ranging from the simplest single-slotted flaps to the most complex double-slotted main/aft flaps. The plan was originally to use single-slotted flaps both inboard and outboard due to simplicity, but changes in the mission goals required a high-speed aileron to be added and thus, more lift needed to be provided by the high-lift system. The final design therefore consists of double-slotted flaps inboard and single-slotted flaps outboard. The double-slotted flaps provide the required extra lift during approach and landing but without the extra pitch trim difficulties that they would have imposed if they were added outboard instead of inboard. The outboard flaps are single-slotted for simplicity but can be deflected up to 37◦ while still generating a lot of lift, although wind tunnel tests showed some separation at these high angles of deflection. The maximum deflection angle setting is 43◦ for the main inboard flap and 67◦ for the aft inboard flap [2][1]. There are two control surfaces on the trailing-edge of the wing besides the high- lift system, which primary purpose is not to generate lift. However, the flaperon, the inboard high-speed lateral control surface, is deflected to match the inboard flap when deployed, in order to fill the gap in the spanwise lift distribution and improve lift and drag characteristics. The aileron, the outboard control surface, is primarily used for low speed flight control but is drooped to match the outboard flaps to provide lift and reduce induced drag during takeoff.

2.3 Future

The new Boeing 777X family was released in 2013 and is estimated to enter service in 2020. The new planes will be powered by an even larger engine manufactured by GE, with a world record fan diameter of 3.35 meters. This engine will not only be the largest one ever manufactured but also the most fuel-efficient and quiet engine on the market. There have been some complaints regarding the fact that the engine can only be manufactured by one company, since this will remove the 2.3. FUTURE 7

Figure 2.1: A Boeing 777 in flight with deflected double-slotted trailing-edge flaps. competition part of the business and airline companies could end up having to pay more for the engines than before. However, an increase in the technology and optimization process behind an airplane design makes it very difficult and maybe most importantly, very expensive, to design an airliner that delivers optimal performance for a range of engines. The Boeing Company states that designing the 777X models with this extra option would add millions of dollars to the retail price, and limit the optimization process. The engine is no longer an external part of the plane that can be added in a final step, it is integrated in the design process from the start and changes in the engine design could affect many other aspects, such as control surfaces or even wing design. The most ground breaking feature of the new models are probably the composite, folding wings which span 64.8 meters on the ground and extend to 71.8 meters in flight, see Figure 2.2 and 2.3. The new wing design lets the plane fit at any airport while increasing the efficiency in air, reducing the fuel consumption even further. With the 95 billion USD order for 259 new aircraft, the new 777X family is already the largest product launch by dollar value in the history of commercial aviation, and most major airline companies has shown a great interest for the new models. Thus, NASA selecting the Boing 777 as the common, structural shape for the CRM seems to be a very tactical move since it should dominate long range flights for a while to come. 8 CHAPTER 2. BOEING 777 OVERVIEW

Figure 2.2: A future Boeing 777X on the ground with folded wings.

Figure 2.3: A future Boeing 777X in flight with full wing span. Chapter 3

High-Lift Controls in General

This chapter aims to provide a basic understanding of the purpose of high-lift control systems and how the most common designs work. A brief discussion about some challenges faced when analyzing the performance of high-lift systems will follow together with some ideas for future development.

3.1 Purpose

During the critical takeoff and landing segments of a flight, the aircraft will travel with a large incidence. Most large, commercial transport aircraft are optimized to fly with a small incidence at an altitude close to 11000 meters. Therefore, while climbing to or descending from this optimal altitude, the aircraft is operating ineffi- ciently and as a result consuming a lot of fuel. Furthermore, flying at low altitudes implies flying through denser atmosphere where the aircraft will experience more drag, which will further add to the increased fuel consumption. Flying at low alti- tudes also cause higher noise levels for the people on the ground, adding to already high stress levels and other issues. These are the most important reasons for why a short climb and descent is desired, and to operate the aircraft at the large incidence that these segments of the flight implies, there is a need for extra lift production in order to avoid stalling. The high-lift control system serves as a mean for the aircraft to operate at higher angles of attack without stalling by producing extra lift. The extra lift is especially needed during takeoff and landing in order to get to and from the optimal cruise altitude as efficiently as possible. Basically, the leading-edge device produces lift by keeping the flow attached further across the wing and the trailing-edge device

9 10 CHAPTER 3. HIGH-LIFT CONTROLS IN GENERAL produces lift by changing the effective camber of the wing. The details regarding how the extra lift is produced is given in Section 3.2.

3.2 Function

Each element of the wing, including the high-lift system, experiences a circulation effect from surrounding elements. The trailing-edge flap causes a change in cir- culation upon deflection and help the wing itself produce more lift by causing an upwash effect on the flow at the trailing-edge. The circulation also produces lift by causing a local increase in the flow velocity which yields higher negative pressure on the upper side of the wing. The lift produced by the flap airfoil itself is rather small compared to the indirect lift production that the flap causes by influencing the flow structure experienced by other elements. The high-lift system produces wakes similar to the main wake produced by the wing. If the wakes from different elements interfere with each other there is a higher risk of flow separation due to a thicker boundary layer being created, which cannot withstand larger pressure gradients. When the flow separates there will be no contribution to the lift from that part of the wing and, if the separation is significant enough, the aircraft will stall. Probably the most important feature of the high-lift system is the so called fresh boundary layer effect, which is the main reason it can produce such a significant amount of lift. The fresh boundary layer effect is present when the wake from the different elements does not coincide and each element instead creates a new boundary layer. Since this boundary layer is thinner than for a single airfoil of equivalent length, the boundary layer is resistant to larger pressure gradients and thus, generates more lift. The function of trailing-edge high-lift devices, simply known as flaps, is to change the effective incidence and camber of the wing section by deflecting down from the trailing-edge of the wing. The deflection increases the lift curve with almost a constant value for a wide range of angles of attack, this constant is known as the lift coefficient increment, ∆CL, and is defined as the value by which the lift coefficient is increased by, at an incidence of 10◦ above the zero-lift angle, see Figure 3.1. This represents a normal landing incidence for a large transport aircraft, and it is convenient to define it at this incidence since extra lift production is especially important during approach and landing, in order to avoid stalling at low speed close to the ground. The lift coefficient increment is more or less independent of test conditions and angle of attack, but will vary with wing aspect ratio, as the lift curve slope itself varies with aspect ratio [3]. The drag coefficient increment, ∆CD0, has a higher dependency on test condi- 3.3. DESIGNS 11

Figure 3.1: Sketched graph for illustrating the lift coefficient increment for varying angles of attack [3]. tions and incidence than the lift coefficient increment. However, it is still rather small for a wide range of angle of attack and it follows that the drag coefficient increment is assumed to be constant as well. The main difference between the two is the angle of attack for which they are defined. Since drag should be minimized during takeoff, the drag coefficient increment is taken to be the change in drag co- efficient at 6◦ above the zero-lift angle, corresponding to a normal takeoff incidence for a large transport aircraft.

3.3 Designs

The high-lift control system used on Boeing 777 were described in the previous chapter and an overview of different devices will be presented here. There are ad- 12 CHAPTER 3. HIGH-LIFT CONTROLS IN GENERAL vantages and disadvantages for each design and there are several factors to take into consideration when deciding on a high-lift design. One design will not be optimal for every airplane configuration, some designs will generate too much lift which will induce unnecessary drag and cause a severe reduction in overall performance. If the device does not generate enough lift it will make takeoff difficult, if not impos- sible, on shorter runways or in bad weather, and be dangerous during the critical approach and landing phase of the flight. High-lift leading-edge devices pushes the separation aft across the wing chord in order to avoid stalling at high angles of attack. The design and setting of the device needs to be very precise in order to obtain desired effect and not cause the flow to separate right at the leading-edge. The most common leading-edge devices are the Krueger flaps and slats. They come in different configurations and in general they are more complex, both in design and operation, the more aerodynamically optimized they are [2].

3.3.1 Trailing-edge devices Since this experimental study investigates the influence on the flow from a trailing- edge high-lift device, different flap designs will be discussed in more detail here. The plain flap, see Figure 3.2, is the most basic flap design and also the general design of most attitude control surfaces, however, as a high-lift device it is rarely used due to its poor lift production ability for subsonic flight.

Figure 3.2: Illustration of a plain flap [3].

The split flap, also known as single-slotted flap, normally refers to a flap that splits the trailing-edge of the wing chordwise and deflects down by rotating about a hinge, see Figure 3.3. The term can also refer to any hinged flap that does not change the upper side of the wing as it deflects, i.e. the flap is not a part of the basic wing when it is not deployed. The slotted flap can be combined with Fowler motion, which considerably increases the area when the flap is fully deployed, see Figure 3.4, [3]. A more efficient flap design is the double-slotted flap, see Figure 3.5, which con- sist of a second slotted flap added to the trailing-edge of a main split flap. The effective camber increase is smoother than for the split flap and thus, it allows 3.3. DESIGNS 13

Figure 3.3: Illustration of a single-slotted flap [3].

Figure 3.4: Single-slotted flap with Fowler motion [3]. for further extension, yielding a larger lifting area. The slots themselves also pro- vide beneficial effects by allowing high energetic flow from below the wing to pass through to the upper side and produce more lift. The double-slotted flap can also be combined with Fowler motion.

Figure 3.5: Illustration of a double-slotted flap [3].

Most large transport aircraft today use double-slotted flaps inboard and single- slotted flaps outboard. The reason behind it is that single-slotted flaps are signifi- cantly simpler to design and use, and weigh less than double-slotted flaps. However, for a large transport aircraft, single-slotted flaps are not able to produce the required lift. As mentioned earlier, it is possible to generate too much lift, thus double-slotted flaps are most likely not needed for the whole wing and they are generally placed on the inboard section due to the wing being more rigid closer to the body. Since the 14 CHAPTER 3. HIGH-LIFT CONTROLS IN GENERAL inboard section of the wing also lie closer to the aerodynamic centre it will require less pitch trimming for the system once in flight [3].

3.4 Challenges

Easy analysis of the lift, drag and pitching moment coefficient increments are pos- sible when only considering a rectangular wing with a flap system attached. Chal- lenges arise when the wing considered is swept, an increase in sweep angle has a tendency to reduce the maximum lift coefficient, see Figure 3.6. However, previous experimental studies show that the lift coefficient increases slightly for moderate aspect ratios and a sweep angle below 20◦ [3].

Figure 3.6: Effect from wing sweep on the lift coefficient increment [3]. The CRM has a sweep angle of 35◦, close to the value illustrated here.

Previously mentioned theories about the characteristics of high-lift devices are 3.5. FUTURE DEVELOPMENT 15 based on a single wing. Considering the effects from the fuselage greatly complicates the theory and even has the ability of changing some of the characteristics of the device. For example, the expectations on the interference effects from the wing-body junction causes the split flap to seem more favourable compared to the double- slotted flap. However, considering all effects shows that the slotted flap induces less drag than the split flap, despite the unfavourable interference effects from the junction.

3.5 Future Development

As material technologies evolve and new lightweight, high-strength composite mate- rials become available to the aeronautical industry, thinner airfoils can be made for increased performance and reduced weight. As the wing airfoil size decreases, the space left for supporting the wing control surfaces become very limited. Designing trailing-edge flap systems that can fit into this limited space is a challenge and it seems as if the only sensible option will be to use single-slotted flaps. The cost in terms of weight and complexity that a double-slotted flap imposes is simply not worth the extra lift it produces. The trailing-edge flap performance would be greatly improved if the spanwise discontinuities could be eliminated. The discontinuities create vortices and cause higher drag and noise. They are a result from needing gaps for thrust-gates for the engine jet and for attitude control surfaces, such as flaperons and ailerons that are used for roll- and pitch control. In order to design a continuous trailing-edge flap system, the control surfaces must either be moved or integrated with the high-lift system. The engine must also be mounted low enough for the jet to clear at least simple-slotted flaps at full deflection. These new challenges aside, a continuous trailing-edge flap system is probably the only way to achieve a simple yet cost- effective and performance increasing high-lift system for future aircraft. It is also, most likely, the only feasible solution to noise reduction, something that is highly prioritized in modern aircraft design. Chapter 4

Description of the Flow

This chapter explains the turbulent junction flow structure from a theoretical point of view. The flow can be investigated using computational methods, however, these have failed for the junction region and the reason why will be discussed below.

4.1 Turbulent Flow Structure

The flow around a wing-body junction is a complex interaction between a turbu- lent boundary layer, vortex legs and a viscous wing wake, see Figure 4.1. The flow behaves differently depending on what main physical feature that drives the inter- action, and this will change multiple times along the intersection between the wing and the fuselage. The flow is divided into three parts after what main physical feature that dominates the flow at that streamwise location, see Figure 4.2.

1. The first part includes the flow upstream from the wing leading-edge and at the leading-edge itself. There will be an onset turbulent boundary layer and a three-dimensional separation that initiates at the leading-edge. Strong pressure gradients cause a primary root vortex and possibly also a secondary vortex to be created in this part of the junction flow.

2. The boundary layer between the first and the second part of the flow is arbi- trary but yet clearly characterized by the absence of strong pressure gradients in the second part. This part is instead driven by Reynolds stresses and the flow is therefore changing very slowly in the spanwise direction – the flow is in some sense in equilibrium during this section. The primary vortex created in the first region, commonly referred to as the horseshoe vortex, experiences

16 4.1. TURBULENT FLOW STRUCTURE 17

Figure 4.1: Illustration of the flow around an airfoil with a single-slotted fowler- flap. The orange represent the boundary layer, the blue is the separation and the surrounding flow is the inviscid flow.

some viscous diffusion in the second part, and the pressure gradients start to increase again as the flow gets closer to the trailing-edge of the wing, where part three takes over. 3. The third part of the flow is driven by viscous wake interactions at the trailing- edge and further downstream from the wing. The flow has a downward move- ment spanwise along the wing, towards the body, and this behaviour causes an imminent flow separation. A new vortex is initiated from the wing-body junction at the trailing-edge of the wing and the primary horseshoe vortex legs, that have been travelling on either side of the wing, are now free to interact. This interaction is believed to contribute significantly to the sep- aration as they carry a lot of streamwise momentum. As a result from the pressure difference above and below the wing, the vortex legs are not equal in size once they reach the trailing-edge, the vortex leg that travelled under the wing will have grown spatially whilst the other one contracted [4].

There is a complex flow structure when the body boundary layer approaches the wing leading-edge. The boundary layer separates due to the adverse pressure gradient created by the wing leading-edge, resulting in a vortex that wraps around the wing continuing downstream in a horseshoe shape. The cross-section bluntness at the leading-edge and the wing sweep both have a strong effect of the location and strength of the primary horseshoe vortex. Methods for controlling junction flow, actively or passively, with the goal of reducing or eliminating the horseshoe vortex 18 CHAPTER 4. DESCRIPTION OF THE FLOW

Figure 4.2: Wing-body junction flow structure divided into three parts. The blue lines represent the wing, the red lines shows where the different parts separate, and the grey spirals illustrates the horseshoe vortex initiating at the leading-edge. and/or its inherent unsteadiness and thereby reducing drag, vibration and/or noise, have all more or less failed [5]. In order to satisfy fluid dynamics conservation laws, the vortex theorems, stream- wise legs of the upstream vortex will stretch around the wing in a horseshoe shape with each leg having opposite vorticity. There is a region of separation around the leading-edge of the wing and a strong flow acceleration between the leading-edge and where the wing has its maximum thickness [6].

4.2 Simulating the Flow

Computational Fluid Dynamics (CFD) is one method used for analyzing the wing- body interaction, besides experimental studies. Low order inviscid codes or panel methods are likely able to show first order effects of this interaction when set up cor- rectly. A three-dimensional fully viscous solution with boundary layer-interaction of the main wing, fuselage, and separate high lift devices will be very challeng- ing even for the most successful CFD codes today and despite recent advances in supercomputing capabilities. When only focusing on the lifting capability of an aircraft, the boundary layer can be excluded in many cases. Inviscid codes will likely yield accurate results for the lifting profiles since the contribution from the boundary layer is negligible when its thin enough, which is the case for flows with high Reynolds number, i.e. regular flight conditions. However, the drag cannot be properly estimated without 4.2. SIMULATING THE FLOW 19 including the boundary layer in the analysis, and since the focus of this study is the flow at the wing-body junction, where a significant amount of drag is developed, CFD codes are not sufficient for a complete investigation of the interaction. While CFD is able to make good predictions, it is likely that experimental studies are always necessary to tune or correct the CFD results. Not only does the two-wall boundary layer have a very complicated interference but there will also be a suction peak forming at the junction, potentially introducing local supersonic flow. This phenomenon has the ability to increase the drag on the aircraft by introducing local shock waves. The adverse interference of the two surfaces is usually reduced by introducing a fairing between them, which also reduces the boundary layer build up and thus, the drag. When control surfaces are added to the investigation the problem becomes even more complicated. Adding more elements does not only imply more individual boundary layers to be solved for but also a completely new behaviour of the resulting boundary layer due to interference. One major reason to why the high-lift system can generate so much lift is the generation of a new boundary layer on the isolated flap or control surface, as mentioned in Section 3.2. Most methods used for analyzing this feature are based on a single, two-dimensional wing. In other words, the three- dimensionality of the wing is not considered and, even less likely, the interaction with the fuselage. Chapter 5

Experimental Approach

There is no junction data available for the CRM to this date and this experimental study aims to contribute to the CRM database while also investigating the influ- ence on this junction flow from the high-lift system. This chapter will explain the experimental set-up in detail and how the wind tunnel tests were executed.

5.1 Model Description

The model used for the wind tunnel tests was a 3% scale, semi-span CRM in two different configurations; Wing/Body/Tail (WBT) and Wing/Body/Tail/Flap (WBTF). Once data was collected for the first configuration, an inboard flap was added to the trailing-edge of the wing to create the WBTF configuration. Since this study focus on the flow at the wing-body junction, only the inboard flap was considered, since it is believed to influence this flow significantly more than the outboard flap. See Figure 5.1 – 5.2 for CAD model images and Table 5.1 for a list of design parameters. The true inboard trailing-edge flap on Boeing 777 is double-slotted but, for simplicity, the flap used for this wind tunnel experiment was approximated to a flat plate without slots, deflecting straight down from the wing trailing-edge. The flap was made of aluminium sheet with a constant thickness of 0.08 cm, mounted to the wing using aluminium foil tape with a thickness ten times smaller than the flap. The dimensions of the deflected part of flap were 17.5 cm by 6.88 cm, and to simplify the mounting there was an overlap of 2.5 cm for the wing and the flap. Since the wing root chord is 36 cm this means that the flap chord is approximately 20% of the root chord, cflap = 0.2croot.

20 5.2. FACILITY 21

(a) WBT

(b) WBTF

Figure 5.1: The CRM, as seen from the test section window, in two different con- figurations.

Four different flap deflections were tested for comparison, the exact angles were 11.2◦, 20.5◦, 30.5◦ and 44.5◦, but will from now on be referred to as 10◦, 20◦, 30◦ and 40◦ in the text. The two smaller angles are in the range for normal takeoff deflections and the largest angles are in the range for normal landing deflections.

5.2 Facility

The test facility where the experiments were performed was the indraft, subsonic wind tunnel at FML, NASA ARC. The wind tunnel can generate a maximum wind- speed of Mach 0.15, or approximately 48 m/s. The test section in which the model was mounted measures 120 cm by 80 cm with a ceiling clearance of 80 cm. The model was mounted with the top of the fuselage 30 cm from the test section ceiling and the bottom 30 cm from the test section floor. The wind tunnel mount wall is interchangeable to accommodate angles of attack within ±6◦, however, only zero angle of attack were tested here due to time limitations. See Figure 5.3 for CAD model images of the test facility and the placement of the model. For each test, the wind tunnel was operated close to full capacity, which yielded 22 CHAPTER 5. EXPERIMENTAL APPROACH

Table 5.1: Values of the CRM properties.

Property Symbol Value Wing aspect ratio ÆR 9.0 ◦ Sweep angle (leading-edge) Λl.e. 35 Wing span (semi-span) b 0.89 m 2 Wing reference area Sref 0.388 m 2 Flap reference area Sflap 0.012 m Wing root chord croot 0.36 m Flap root chord cflap 0.0688 m a Reynolds number of almost one million based on the root chord.

5.3 Data Collection

The velocity data was collected using a Cobra Probe from TFI, a dynamic multi- hole pressure probe designed for unsteady, turbulent flows. A list of other collected data and the sensors used can be seen in Table 5.2, and from the measured variables other parameters were deduced, see Table 5.3. The Cobra Probe has four pressure taps on a faceted head and measures three- component velocities and static pressures within a range of ±45◦, with a frequency of 2000 Hz. For flow speeds above 2 m/s and turbulence intensities below 30%, the cobra probe has a measurement accuracy of ±0.5 m/s for the velocities and ±1◦ for the pitch and yaw angles. The Cobra Probe replaces the commonly used hot-wire and other types of anemometers. The probe can withstand moderate knocks and contaminated flow and requires no calibration. It is 160 mm long with a maximum diameter of 14 mm and the measurement head measures only 2.6 mm. For the wind tunnel experiments conducted here, the probe was pre-programmed to take measurements at certain points in a three-dimensional grid. To assure that the probe did not hit the model or any other sensors in the wind tunnel, the grid was verified by letting the probe travel to all points without wind on. If a point was out of bound, or caused the probe to hit the model, it was easier to notice when the wind was off and the test section window was open. Which points to include in the grid was decided partly by looking at previous wind tunnel experiments, but also by conducting a water channel experiment where a similar wing-body junction was investigated using fluorescent dye and UV light to highlight interesting areas. Especially the junction wake has a very interesting flow structure due to separation and generated vortices from the wing-body junction. Therefore, the grid covered 5.3. DATA COLLECTION 23

Table 5.2: Measured variables by different sensors, during each wind tunnel run.

Name Symbol Sensor Range Accuracy Unit Barometric pres- pbaro Absolute pres- [0,133.3] 0.05% kPa sure sure x-ducer Differential pres- pctr Bidirectional dif- [-2.7,2.7] 0.05% kPa sure across con- ferential pressure traction x-ducer ◦ ◦ Facility total Tfac 100 ohm plat- [-17.8,82.2] 1.8 C C temperature inum RTD tem- perature sensor Relative humid- Hrel Solid state hu- [3,95]% ±2% % ity midity sensor

planes at 3 cm, 8 cm, 13 cm, 18 cm and 23 cm downstream from the wing trailing- edge/body junction for the WBT configuration and 8 cm, 13 cm, 18 cm and 23 cm downstream for the WBTF configuration, the flap extending backwards caused the plane 3 cm downwind to be out of bounds. Each plane consists of 11 × 21 = 231 points for lateral (spanwise) values between -3.6 and -1 cm inboard and vertical values between 5 cm below and 5 cm above the wing. See Figure 5.4 for a plot of the grid in the lateral/vertical-plane. The data was measured by the Basic Data Acquisition System (BDAS) and processed by the TFI device control software to get the calculated parameters. The probe took measurements at 2000 Hz for five seconds at each point of the grid. A time-averaged value was then calculated by the software and saved as a text file.

The variables measured by other sensors than the Cobra Probe during the wind tunnel run are listed in Table 5.2 and Equation 5.2 – 5.4 show how some parameters are calculated from measured variables. 24 CHAPTER 5. EXPERIMENTAL APPROACH

Table 5.3: Calculated parameters deduced from measured data.

Name Symbol Dependency Formula Unit Atmospheric pres- ptsc pbaro,Dtsc 144pbaro − kPa sure at tunnel 0.07285Dtsc centerline Absolute total ptot ptc, pcorr ptc + pcorr kPa pressure at tunnel centerline Static pressure at pstat ptot, q ptot − q kPa tunnel centerline q 2/7 Mach number in Mts ptot, pstat 5 (ptot/pstat) − 1 1 test section ◦ Static tempera- Tstat Tfac,Mtun See Eq. (5.1) C ture in test sec- tion p Speed of sound c Tstat,R 49.0223 Tstat,R ft/s Test section dry pvap,B ptc,Tstat See Eq. (5.2) psf bulb vapor pres- sure Test section dew pvap,P Hrel, pvap,B See Eq. (5.3) psf point vapor pres- sure Static density ρstat pstat, pvap,P,Tstat,R See Eq. (5.4) slug/ft 2 Test section dy- q pstat,Mtc 0.7pstatMtun psf namic pressure 5.4. CORRECTION METHODS 25

1.8Tfac − 273.15 Tstat = 2 (5.1) 1 + 0.2Mtun ( ) −3
31.5Tstat − 0.064 pvap,B = 12.77 + 0.442 · 10 ptc · exp (5.2) 433.5 + 1.8Tstat H p p = rel vap,B (5.3) vap,P 100   −6 20.89pstat − 0.38pvap,P ρstat = 582.56 · 10 (5.4) 1.8Tstat − 273.15 5.4 Correction Methods

The velocity measurements were corrected for sensor mounting imperfection by taking measurements far away from the model, where the flow was assumed to be laminar and completely streamwise. The measured lateral and vertical velocities at the free stream were then subtracted from the velocities measured at the grid. The true downstream velocity was larger than the measured value, which was corrected for by dividing the values measured at the grid by the values measured at the free stream. The free stream measurements were taken at each downstream location at a lateral distance of 30.5 cm and vertical distance of 25.4 cm from the trailing-edge of the wing at the junction.

ui(y, z) uˆi(y, z) = , vˆi(y, z) = vi(y, z) − V∞ , wˆi(y, z) = wi(y, z) − W∞ U∞ where i = 1, 2, 3, 4 corresponds to each downwind location. 26 CHAPTER 5. EXPERIMENTAL APPROACH

(a) WBT

(b) WBTF

Figure 5.2: The CRM from above in two different configurations. 5.4. CORRECTION METHODS 27

(a) Side view of wind tunnel.

(b) Top view of wind tunnel.

Figure 5.3: Wind tunnel CAD images with the model mounted in the test section. The text section dimensions are 120 cm by 80 cm by 80 cm (not including the plenum). 28 CHAPTER 5. EXPERIMENTAL APPROACH

Figure 5.4: A plot of the grid representing the points at which measurements where taken. The grid is shown in the yz-plane and was repeated for 4-5 downstream locations (x). The free stream measurements in the beginning and at the end of each run is not shown in this grid plot. The measurements started at the top right corner and travelled down vertically before moving inboard (left). Chapter 6

Experimental Results

6.1 Velocity Field

The test conditions are listed in Table 6.1 as averaged values measured during each wind tunnel test. The velocity data is processed using Matlab for two-dimensional contour plots of the velocity field, see Figure 6.1 and 6.2, and Tecplot 360 is used for three- dimensional contour plots with a CAD model included for reference, see Figure 6.3. The velocity field yields information about properties such as vorticity, separation, downwash, drag and lift. The downstream velocity ratio is plotted as an intensity where zero correspond to no downstream components of the velocity field and one correspond to free stream. A wake is expected downwind from the wing which should cause a region of slower flow right behind the wing. The velocity compo- nents in the lateral/vertical-plane, i.e. any flow components not in the downstream direction, is plotted to investigate how the wing/flap affects the flow more specifi- cally. Trajectory lines are included to visualize the inplane movement of the flow, which can show vortex formations and give an idea of where they arise from.

6.2 Downwash

The deflected flap yields a downwash on the flow field and therefore an upwash on the wing which corresponds to extra lift. The amount of downwash a certain flap deflection yields could bring information about what settings are effective and when they reach the maximum lift production. The downwash can be seen in the flow trajectory that is plotted over the contour of the mean velocity ratio for different

29 30 CHAPTER 6. EXPERIMENTAL RESULTS

Table 6.1: Averaged values for measured variables for each wind tunnel run.

Variable Name Symbol No Flap 10◦ 20◦ 30◦ 40◦ Unit Wind speed Vwind 47.9 47.9 47.9 47.9 47.9 m/s Mach number Mts 0.14 0.14 0.14 0.14 0.14 1 Reynolds number Re 0.99 0.97 0.96 0.98 0.95 1 ◦ Static tempera- Tstat 14.4 16.1 17.2 15.0 20.0 C ture ◦ Total tempera- Ttot 15.6 17.2 18.3 16.1 21.7 C ture Relative humid- Hrel 74 62 40 83 40 % ity 3 Static density ρstat 1.21 1.20 1.19 1.19 1.18 kg/m Static pressure pstat 100.6 99.6 99.1 99.1 99.6 kPa Dynamic pres- q 1.34 1.34 1.34 1.39 1.34 kPa sure Total pressure ptot 101.9 100.9 100.5 100.5 100.9 kPa Dynamic viscos- ν 1786 1796 1800 1791 1815 Pa s ity configurations and flap settings, see Figure 6.4. This downwash can be calculated by looking at how much of the flow is moving vertically compared to downstream,

W (x, z) tan θ = (6.1) U(x, z) where W (x, z) is the vertical velocity ratio, and U(x, z) is the downstream velocity ratio in the downstream/vertical-plane. See Table 6.2 for the maximum downwash angles at each streamwise location. The downwash angles from the flow trajectories obtained for the WBTF configurations are normalized using the results from the WBT configuration, since it is of interest to know how the flap changes the flow compared to the basic wing. The maximum angles are found below the wing, which can be seen in the figures, this corresponds well to what is expected since the flap deflects down from the leading-edge and should therefore have a larger impact on the flow below the wing than above it. 6.2. DOWNWASH 31

(a) 8 cm downwind

(b) 23 cm downwind

Figure 6.1: Downstream velocity ratios for each configuration and flap setting, at two different downstream locations.

Table 6.2: Maximum downwash angles, θ, for different flap deflections, δf , at four downstream locations. The flap downwash angle is normalized with the downwash angle obtained from the measurements without flap on. All angles are given in degrees.

δf θx=8cm θx=13cm θx=18cm θx=23cm 10 9.56 7.46 5.49 3.74 20 9.01 7.00 5.07 3.39 30 19.3 23.3 25.1 27.8 40 7.15 5.13 3.27 1.51 32 CHAPTER 6. EXPERIMENTAL RESULTS

(a) 8 cm downwind

(b) 23 cm downwind

Figure 6.2: Inplane velocity ratios (lateral/vertical) for each configuration and flap setting, at two different downstream locations. Trajectory lines are included to show the movement of the flow, especially vortex formation. The outline of the wing/flap is included as dashed, black lines. The inboard, aft tip of the flap is marked with a triangle in the bottom right corner. 6.2. DOWNWASH 33

Figure 6.3: Three-dimensional plot of the downstream velocity ratio and the CRM WBT CAD model to show where data was collected. This data corresponds to the configuration with a 30◦ flap deflection. Only three out of four planes are plotted here. 34 CHAPTER 6. EXPERIMENTAL RESULTS

(a) WBT

(b) WBTF

Figure 6.4: Contour plot of the mean velocity field in the downstream/vertical plane closest to the body (1 cm outboard), for each configuration. Chapter 7

Aerodynamic Performance

The drag and lift can be estimated using a control volume analysis of the momentum deficit in the wake of the wing-body junction. By analyzing the measured data, an estimate of how the flap deflection affect the drag and lift of the wing section close to the junction can be obtained. Worth noting is that the collected data only covers a small section of the wing and therefore is not able to give an accurate, absolute value of the drag and lift produced by the plane, but rather an estimate of how the inboard flap affects the flow and what angle settings may be preferred during certain segments of the flight.

7.1 Drag estimate

The drag can be estimated by considering a control volume that encloses the three- dimensional region covered by the data collection. The optimal control volume for estimating drag accurately is one that confines the whole airplane, however, that analysis would require data to be collected at every point surrounding the airplane. Since this study focus on the wing-body junction wake, the analysis is constrained to this region. The theoretical drag is given by Z D = ρUdA (7.1) where ρ is the air density, U is the downstream component of the velocity field and dA is the surface segment affected by the drag. The drag estimate is based on the change in downstream velocity, ∆U, due to the presence of the wing and, for the WBTF configuration, the inboard flap at

35 36 CHAPTER 7. AERODYNAMIC PERFORMANCE different angles. The collected data consist of a discrete number of points and the integral in Equation (7.1) is therefore approximated by a summation of all measured values in each plane that confines the control volume. For one plane the drag is calculated as X D = ρA∆U

Since the lateral component of the velocity field is significantly smaller than the ver- tical component or the downstream component at every point, the downstream/ver- tical planes of the control volume will not be included in the analysis. The part of the drag through the lateral/vertical plane is calculated at a down- stream distance of 23 cm. This plane is chosen since it is the one furthest away from the wing trailing-edge and will therefore have the steadiest flow that describes the influence from the wing and flap the best. The on-set flow on the wing is assumed to be completely laminar and streamwise, the reduction in the downstream velocity ratio can therefore be calculated by subtracting the measured downstream velocity ratio from the free stream value.

∆Ujk = U∞ − Ujk = 1 − Ujk where Ujk correspond to the measured downstream velocity ratio at point (yj, zk). The drag at x = 23 cm is

11 21 11 21 X X X X
Dyz x=23cm = ρAyz∆Ujk = ρAyz 1 − Ujk x=23cm j=1 k=1 j=1 k=1 where Ayz is the area spanned by y ∈ [−3.6, −1] cm and z ∈ [−5, 5] ⇒ Ayz = 26 cm2. The drag through the downstream/lateral planes of the control volume is calcu- lated the same way but now the plane is spanned by x ∈ [8, 23] cm and y ∈ [−3.6, −1] cm at z = −5 cm and z = 5 cm. The drag through the top plane is

4 11 X X
Dxy z=5cm = ρAxy 1 − Uij z=5cm i=1 j=1 and similar for the bottom plane,

4 11 X X   Dxy z=−5cm = ρAxy 1 − Uij z=−5cm i=1 j=1 7.2. LIFT ESTIMATE 37

2 The areas are the same for both planes, Ayz = 39 cm . The total drag is then

Dtot = Dxy z=−5cm + Dxy z=5cm + Dyz x=23cm

See Table 7.1 for calculated values. The drag coefficient is usually what is compared and optimized, it can be calculated from the drag formula

1 2 D D = ρU CDSref = qCDSref ⇒ CD = 2 qSref where q is the dynamic pressure which is measured separately during each wind tunnel run. The reference area Sref is taken to be the wing root chord length multiplied by the width of the grid for the WBT configuration, and the wing root chord length plus the flap chord length multiplied by the grid width for the WBTF configuration. Other reference areas could be considered, this area is chosen in order to get an estimate of the partial drag coefficient from this section of the wing. The drag coefficients for the different configurations are then

Dtot Dtot CD = = WBT Sref,w 0.05croot Dtot Dtot Dtot CD = = = WBTF Sref,wf 0.05(croot + cflap) 0.06croot

Where cflap = 0.2croot is used to simplify the expression. See Table 7.1 for calculated values of the estimated drag and partial drag coefficient, for each flap setting.

7.2 Lift estimate

The lift is estimated in the same manner as the drag, the only difference being that the vertical component of the velocity field is considered instead of the down- stream component. Let Wij be the vertical velocity ratio in the downstream/lateral plane and Wjk the vertical velocity ratio in the lateral-vertical plane. The vertical free stream value W∞ is zero since the free stream is assumed to be completely streamwise, this gives

∆Wij = W∞ − Wij = 0 − Wij = −Wij

∆Wjk = W∞ − Wjk = 0 − Wjk = −Wjk 38 CHAPTER 7. AERODYNAMIC PERFORMANCE

The lift is then

4 11 X X Lxy z=−5cm = −ρAxy Wij z=−5cm i=1 j=1 4 11 X X Lxy z=5cm = −ρAxy Wij z=5cm i=1 j=1 11 21 X X Lyz x=23cm = −ρAyz Wjk x=23 cm j=1 k=1 and the total lift is

Ltot = Lxy z=−5cm + Lxy z=5cm + Lyz x=23cm The partial lift coefficient is obtained from the lift formula,

1 2 L L = ρV CLSref = qCLSref ⇒ CL = (7.2) 2 qSref where the dynamic pressure and reference area are the same as in the drag case.

Ltot Ltot CL = = WBT Sref,w 0.05croot Ltot Ltot CL = = WBTF Sref,wf 0.06croot

See Table 7.1 for calculated values of the estimated lift and partial lift coefficient, for each flap setting.

7.3 Optimization

As can be seen in Table 7.1, the lift and drag are both increasing up to a flap deflection of 30◦ and then decreasing again for 40◦. This trend implies that there should be an optimal flap setting somewhere in between these deflection angles. For the case with no flap and flap angle 40◦, the drag is larger than the lift, which means that the airplane would not be able to fly with these settings if this was the absolute values for drag and lift. However, only a small section of the wing is considered and there is therefore a possibility of the remaining part of the wing supplying sufficient lift to account for this. 7.3. OPTIMIZATION 39

Table 7.1: Estimated values for the partial lift and drag coefficients obtained with control volume analysis of the momentum deficit. The values give an indication of trends in lift and drag for different flap deflections, rather than an absolut measure.

◦ 2 δf ( ) Sref (m ) D (N) L (N) CD CL L/D 0 0.018 49.71 40.97 1.703 1.404 0.8242 10 0.022 81.06 91.67 2.777 3.141 1.131 20 0.022 179.5 215.3 6.150 7.377 1.200 30 0.022 255.1 296.9 8.741 10.17 1.164 40 0.022 216.3 213.1 7.410 7.301 0.9855

During landing, high lift is desired and the induced drag is not considered in the optimization process. The drag is actually a positive consequence from the high lift production during approach and landing since the speed need to be reduced significantly before touchdown. Therefore, a simplified optimization problem for flap deflection during landing can be written as 
max CL = CL max δf

For takeoff, however, the maximum lift is desired while at the same time minimizing the drag. This gives an optimization problem on the form n o

max CL/CD = CL/CD max = L/D max δf

Both optimization problems are solved by interpolating the calculated values with a third degree polynomial, namely

−3 3 2 CL(δf ) = −0.68 · 10 δf + 0.033δf − 0.093δf + 1.4
−5 3 2 L/D (δf ) = 0.79 · 10 δf − 0.0012δf + 0.041δf + 0.83 Maximizing these functions with respect to flap angle graphically gives the optimal settings for landing and takeoff, respectively.

◦ CL max = 10.33 , for δf = 33 Landing
◦ L/D max = 1.2 , for δf = 22 Takeoff See Figure 7.1 and 7.3 for plots of the calculated points and the cubic fit. Figure 7.2 shows a plot of the estimated partial drag coefficient and the corresponding cubic fit, which is not used in the optimization but included for reference. 40 CHAPTER 7. AERODYNAMIC PERFORMANCE

Figure 7.1

Figure 7.2 7.3. OPTIMIZATION 41

Figure 7.3 Chapter 8

Discussion

The downstream velocity ratio plotted in Figure 6.1 shows a flow separation ini- tiating somewhere between 30◦ and 40◦ flap deflection. The flow for the 40◦ flap setting is clearly separated and the 30◦ flap setting shows that some separation may have initiated although still not completely. The separation imply that there would be little or no lift produced by this section of the wing at these flap angles. Modern airplane wings are twisted to avoid stalling (i.e. the flow separating) at the wing tips, and it is therefore likely that the whole wing would not stall for this setting. However, since the flaps are used to increase the lift, it is probably never desired to put them at an angle that makes them stall. From this result it is now known that, for this flap design and these test conditions, the maximum flap angle should be around 30◦, larger angles will not be beneficial, neither in terms of lift nor drag. The trajectory lines in the inplane velocity ratio in Figure 6.2, shows that the flow in the lateral/vertical plane is dominated by vertical flow, with a vortex forming in the region of the flap tip, closest to the body. The inplane contours shows how much the model affects the flow, since it should be completely streamwise when there is no model present. The influence the model has on the flow structure seem to increase with flap angle, this agrees well with what is expected. The flow ratio in the lateral direction is very small which implies that there may be very little streamwise motion along the inboard section of the wing. The velocity field in the downstream/vertical plane shows the expected wake behind the wing for the case with no flap and 10◦ flap deflection, see Figure 6.4. For larger deflections, the influence from the flap become so significant that the wing wake is not visible in the contour plot anymore. The pink area in the top right corner corresponds to the body-boundary layer, and this boundary layer gets thicker with flap deflection up to 30◦, for 40◦ the boundary layer has reduced again.

42 8.1. FUTURE WORK 43

The wake from the flap interacts heavily with the body-boundary layer, causing a lot of extra drag. This phenomenon is one example of a result that CFD would not be able to predict. Unfortunately, the maximum boundary layer interference seems to occur for a flap angle close to the one generating the maximum lift, which makes it difficult to obtain good lifting behaviour without inducing drag. Although this result is already widely known, hence the term lift induced drag, which makes up a large part of the total drag of an airplane. The downwash angles for the largest flap deflection are smaller than for the 30◦ flap, implying that the optimal setting for lift production should be somewhere between 20◦ and 40◦. The boundary layer interference is largest for this flap deflec- tion, which is not unexpected since lift will induce drag and thus, the flap setting that produces the most lift should also produce the most drag.

8.1 Future Work

The CRM wind tunnel data collection is an on-going project that NASA will con- tinue to contribute to. Due to time limitations, only one angle of attack was con- sidered in this study, although NASA are planning on continuing the tests with start in mid June 2017, covering at least three angles of attack. There are plans on collecting data in the on-set body-boundary layer at the nose of the fuselage, and above the wing at the wing-body junction, all for varying angles of attack. As mentioned earlier, the future high-lift systems will most likely be simpler and lighter than the ones used today due to lightweight materials that cannot support heavy, complex control mechanisms. If the high-lift system is to be simpler, it has to be optimized and designed in such a way that it will still generate a sufficient amount of lift for all aspects of the flight. In order to optimize the design efficiently, more has to be known about the flow structure it give rise to, and simulating this using CFD would greatly simplify the analysis. Hopefully this data can contribute to CFD validations, but more data for this region is still needed. Chapter 9

Summary

Three dimensional wind tunnel data for a 3% semi-span CRM was collected using a Cobra Probe. Data was first collected without flap and then flaps with four different angles were mounted and tested for, at zero angle of attack. The indraft, subsonic wind tunnel was operated at full wind speed capacity of Mach 0.15, which generated Reynolds numbers close to one million based on the wing root chord. The collected data was then processed using Matlab and Tecplot 360 to investigate how the flow structure changed with flap settings. From contour plots of the velocity field it could be assumed that the flow was fully separated at the largest flap angle and that the flap clearly pushes the flow down. The maximum downwash effect was obtained for a flap angle of 30◦ and this is also the angle that made the wake from the wing interact with the body-boundary layer the most, which is probably the reason why the drag was maximum at this flap angle as well. The purpose of the experiment was to collect wing-body junction data for the CRM in order to contribute to the CFD reference database used for simulation validations. The collected data in this study will be made public shortly after this thesis is completed. The high-lift system interference close to the body of the aircraft was interesting to investigate since CFD analysis leaves out this part as it cannot solve for the boundary layer. CFD models may be able to predict lift profiles fairly well but drag estimates are more difficult, especially the drag that arises from the boundary layer interactions.

44 Bibliography

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