Design of a Variable Span Wing for an Unmanned Aerial Vehicle Aerospace Engineering

Design of a Variable Span Wing for an Unmanned Aerial Vehicle

Marcelo António Borralho da Silveira

Thesis to obtain the Master of Science Degree in Aerospace Engineering

Supervisors: Prof. Filipe Szolnoky Ramos Pinto Cunha Prof. Luís Filipe Galrão dos Reis

Examination Committee Chairperson: Prof. Fernando José Parracho Lau Supervisor: Prof. Filipe Szolnoky Ramos Pinto Cunha Member of the Committee: Dr. Frederico José Prata Rente Reis Afonso

July 2019 ii Agradecimentos

Aos meus orientadores, Professor Lu´ıs Reis e Professor Filipe Cunha quero agradecer por todo o apoio, disponibilidade e pacienciaˆ para me orientar e ajudar durante todo o desenvolvimento deste trabalho. A toda minha fam´ılia. Aos meus pais por me apoiarem e suportarem incondicionalmente durante todos estes anos e por me encorajarem a ser melhor. E ao meu irmao˜ por toda a companhia e amizade que temos partilhado. A todos os meus amigos e colegas que me acompanharam durante o curso por tornarem esta viagem nao˜ so´ mais facil´ como inesquec´ıvel.

iii iv Resumo

As tecnologias para aeronaves adaptativas (Morphing Aircraft) fornecem uma estrategia´ para alterar e adaptar a configuraçao˜ da aeronave as` condiçoes˜ de voo. Esta abordagem ao design de aeronaves tenta encontrar melhorias de eficienciaˆ para aeronaves do futuro. A asa telescopica´ de envergadura variavel´ oferece as caracter´ısticas de uma asa com elevada envergadura e alongamento desejaveis´ para um voo eficiente a baixa velocidade, ao mesmo tempo que acomoda aquelas de uma asa menor com reduzido alongamento para um melhor desempenho em voo de alta velocidade. Este trabalho apresenta o projeto de uma asa telescopica´ com envergadura variavel´ entre 3,5m e 5m para um ve´ıculo aereo´ nao˜ tripulado (UAV) de 150kg de media´ autonomia para patrulha e vigilancia.ˆ Varios´ mecanismos de asa telescopica´ foram estudados e descritos. A metodologia para o calculo´ de cargas aerodinamicas,ˆ tensoes˜ estruturais maximas´ e dimensionamento dos elementos da longarina telescopica,´ com base em formulas´ anal´ıticas e calculos´ aerodinamicosˆ recorrendo a` teoria da linha sustentadora, e´ apresentada. A estrutura da aeronave e´ projetada para sustentar um fator de carga de 3,8g com um fator de carga maximo´ de 5,7g. O desempenho da aeronave equipada com a asa de envergadura variavel´ foi avaliado e comparado em termos de resistenciaˆ com duas aeronaves equipadas com uma asa fixa com 4,3m e 5m de envergadura, respectivamente. A asa adaptativa supera a asa fixa de 5m para velocidade de mergulho de 150kt com uma reduçao˜ de resistenciaˆ aerodinamicaˆ de 23% e pode permitir um aumento de 68% em massa ate´ que nenhum benef´ıcio de desempenho seja observado. Em comparaçao˜ com a asa fixa de 4,3 m, a resistenciaˆ aerodinamicaˆ e´ reduzida ate´ 16% a` velocidade de loiter de 70kt com uma penalidade de peso de 13%.

Palavras-chave: Asa Adaptativa, envergadura variavel,´ UAV, dimensionamento estrutural, desempenho.

v vi Abstract

Morphing aircraft technologies provide a strategy to change and adapt the aircraft shape to the flight conditions. This approach to aircraft design aspires to find efficiency improvements for future aircraft. The variable span telescopic wing offers the characteristics of a large span wing with high aspect ratio desirable for efficient low speed flight while accommodating those of a smaller span wing with low aspect ratio for better flight performance at high subsonic speed. This work presents the design of a telescopic wing with span ranging from 3.5m to 5m for a medium- endurance patrol and surveillance 150kg unmanned aerial vehicle (UAV). Different telescopic wing mechanisms were studied and described. Methodology for calculating aerodynamic loads, maximum structural stresses and sizing of the telescopic spar elements, based on analytical formulas and aerodynamic calculations using the lifting line theory, is present. The aircraft structure is designed to sustain a load factor of 3.8g with an ultimate load factor of 5.7g. The performance of the variable span wing aircraft was assessed and compared in terms of drag with two fixed wing aircraft with 4.3m and 5m wingspan respectively. The morphing wing outperforms the 5m fixed wing for 150kt dive speed with 23% drag reduction and can allow for a 68% increase in mass until no performance benefit is noticed. Compared with the the 4.3m fixed wing the drag reduction achieved is up to 16% at loiter speed of 70kt with weight penalty of 13%.

Keywords: Morphing wing, variable span, UAV, structural sizing, performance.

vii viii Contents

Agradecimentos...... iii Resumo...... v Abstract...... vii List of Tables...... xi List of Figures...... xiii Nomenclature...... xvii Glossary...... xix

1 Introduction 1 1.1 Motivation...... 1 1.2 Morphing Overview...... 2 1.3 Objectives...... 4 1.4 Thesis Outline...... 4

2 Morphing Background 7 2.1 Historical Perspective...... 7 2.1.1 Main Aircraft Morphing Research Programs...... 9 2.2 Wing Morphing...... 13 2.2.1 Twist...... 13 2.2.2 Sweep...... 15 2.2.3 Dihedral (Folding and Winglets)...... 16 2.2.4 Chord...... 18 2.2.5 Camber...... 20 2.2.6 Span...... 29

3 Telescopic Wing Design Strategy 35 3.1 Design Framework...... 35 3.1.1 UAV Characteristics...... 36 3.1.2 Mission Proﬁle...... 37 3.1.3 Proﬁle Selection...... 38 3.2 Telescopic Wing...... 40 3.2.1 Main Spar...... 43

ix 3.2.2 Ribs...... 44 3.3 Structural Strength and Sizing...... 45 3.3.1 Aerodynamic Loads...... 47 3.3.2 Moment and Shear Diagrams...... 47 3.3.3 Maximum Stress and Sizing...... 49

4 Results 51 4.1 Structural Strength and Sizing Results...... 51 4.1.1 Aerodynamic Loading...... 52 4.1.2 Bending Moment and Shear Diagrams...... 53 4.1.3 Maximum Stress...... 53 4.1.4 Remarks...... 57

5 Aerodynamic Performance Studies 59 5.1 Performance Assessment...... 59 5.2 Remarks...... 64

6 Conclusions 67 6.1 Conclusions and Achievements...... 67 6.2 Future Work...... 68

Bibliography 69

A Technical Datasheets 79 A.1 TEKEVER AR5 Datasheet...... 79

x List of Tables

1.1 Effects of wing geometric parameters on aircraft performance from [7]...... 3

3.1 TEKEVER AR5 speciﬁcations...... 37 3.2 Aircraft speciﬁcations...... 37 3.3 Aircraft desired performance...... 37 3.4 Low Reynolds airfoils characteristic at Re=1,000,000...... 39 3.5 Airfoil evaluation matrix...... 39 3.6 Concept evaluation matrix...... 42 3.7 Morphing aircraft optimal speeds...... 46

4.1 Final telescopic spar dimensions...... 51 4.2 Reaction loads: Case A – Stall Speed...... 53 4.3 Reaction loads: Case B – Dive Speed...... 53 4.4 Maximum section loading: Case A Stall Speed...... 56 4.5 Maximum section loading: Case B Dive Speed...... 56 4.6 Maximum normal and shear stress: Case A Stall Speed (5m span)...... 56 4.7 Maximum normal and shear stress: Case A Stall Speed (3.5m span)...... 56 4.8 Maximum normal and shear stress: Case B Dive Speed (5m span)...... 56 4.9 Maximum normal and shear stress: Case B Dive Speed (3.5m span)...... 57 4.10 Maximum deﬂection: Case A Stall Speed (5m span)...... 57

5.1 Optimal aircraft speeds...... 64 5.2 Wing drag for different speeds with a comparison with the 5m span wing...... 64 5.3 Optimal morphing wing weight penalties to match ﬁxed 4.3m and 5m span wing drag... 64

xi xii List of Figures

2.1 Clement Ader’s Eole – a (non-ﬂying) shape changer, [8]...... 7 2.2 Ivan Makhonine’s MAK-10 telescopic wing aircraft (1931), [9]...... 8 2.3 21st Century Aerospace Vehicle (artist rendering), [14]...... 9 2.4 (a) Morphing Aircraft developed by NextGen in the DARPA MAS 3 Program, [5]; (b) Lock- heed Martin baseline morphing concept, [2]...... 11 2.5 (a) SADE’s Smart Leading Edge concept, [17]; (b) SARISTU wing demonstrator with integrated leading, trailing edge smart devices and active winglet, [18]...... 12 2.6 Wing morphing concepts types, [6]...... 13 2.7 (a) Digital Morphing Wing and its discrete lattice building-block elements, [25]; (b) MAV wing with torque rod, [24]...... 14 2.8 (a) Grumman F-14 Tomcat, [29]; (b) Batwing concept conﬁgurations for highlift, climb, cruise, loiter and maneuver, [4]...... 15 2.9 MAV undergoing neutral (top) and positive (bottom) gull-wing morphing, [34]...... 17 2.10 (a) Experimental model as mounted in the wind tunnel, [36]; (b) HECS wing concept schematic, [40]...... 18 2.11 (a) Bakshaev LIG-7 schematic, [9]; (b) Sliding Rib concept developed by Cornstone Re- search Group, [42]...... 19 2.12 (a) Chord extension through the use of extendable trailing edge plate, [44]; (b) Chord extension through elastic deformation of structure, [45]...... 19 2.13 Cross section of an airfoil, [4]...... 20

2.14 Airfoil camber effect on: (a) Cl vs α curve; (b) Cd vs Cl curve, adapted from [46, 47].... 21 2.15 High lift devices configurations: (a) Cross section schematic, [47]; (b) Landing and take-off flaps, [10]...... 21 2.16 (a) Compliant trailing edge device with fairings (transition sections), [62]; (b) Flap with eccentric actuator device, [63]...... 24 2.17 FishBAC morphing airfoil wind tunnel at rest and deflected downwards, [68]...... 25 2.18 Spanwise shape variations of the smart trailing edge control surface, [72]...... 26 2.19 The morphing aircraft before flight number 6 on April 29, 2010, [74]...... 27 2.20 Shape morphing aero control surface’s strucure: airfoil at minimum camber (up); airfoil at maximum camber (bottom), [79]...... 28

xiii 2.21 Morphing rib structure, [80]...... 29 2.22 The three segment telescopic wing from [87]: (a) Actuators; (b) Wind tunnel test setup.. 31 2.23 General CAD view of the variable span wing from [92]...... 32 2.24 Morphing core fabricated section, [94]...... 33 2.25 Adaptive Aspect Ratio concept: (a) Isometric view (retracted); (b) Top view (extended), [67, 96]...... 34

3.1 Detailed design process...... 36 3.2 Mission profile...... 38 3.3 Seilig S4233 low Reynolds airfoil...... 38 3.4 Plant view of span morphing concept A (extended)...... 41 3.5 Isometric view of span morphing concept A (retracted)...... 41 3.6 Plant view of span morphing concept B (extended)...... 41 3.7 Isometric view of span morphing concept B (retracted)...... 41 3.8 Plant view of span morphing concept C (extended)...... 42 3.9 Isometric view of span morphing concept C (retracted)...... 42 3.10 Plant view of the telescopic spar (retracted)...... 43 3.11 Isometric view of the telescopic spar (retracted)...... 43 3.12 Isometric view of the sliding rib with bounding box dimensions...... 45 3.13 Velocity-load diagrams: (a) 5m wingspan configuration (maximum) (b) 3.5m wingspan configuration (minimum)...... 46 3.14 Telescopic spar loading schematic...... 47 3.15 Free body diagram of each telescopic spar element...... 48 3.16 Free body diagram of differential length of a beam, taken from [107]...... 48 3.17 (a) Rectangular hollow beam cross section (b) Normal stress cross sectional distribution, [109]...... 49

4.1 Final telescopic spar geometry schematic...... 51 4.2 Semi span lift distribution for maximum and minimum wingspan conﬁgurations (case A: n= 5.7; stall condition)...... 52 4.3 Semi span lift distribution for maximum and minimum wingspan conﬁgurations (case B: n= 5.7; V= dive speed)...... 52 4.4 Bending moment diagram of the maximum wingspan (5m) wing at stall condition and n= 5.7 53 4.5 Shear force diagram of the maximum wingspan (5m) wing at stall condition and n= 5.7.. 54 4.6 Bending moment diagram of the minimum wingspan (3.5m) wing at stall condition and n= 5.7...... 54 4.7 Shear force diagram of the minimum wingspan (3.5m) wing at stall condition and n= 5.7. 54 4.8 Bending moment diagram of the maximum wingspan (5m) wing at dive speed and n= 5.7 55 4.9 Shear force diagram of the maximum wingspan (5m) wing at dive speed and n= 5.7... 55 4.10 Bending moment diagram of the minimum wingspan (3.5m) wing at dive speed and n= 5.7 55

xiv 4.11 Shear force diagram of the minimum wingspan (3.5m) wing at dive speed and n= 5.7... 56

5.1 Morphing aircraft wing model: (a) 5,000mm wingspan (b) 3,500mm wingspan...... 59 5.2 Wing lift coefficient with angle of attack for different span wings...... 60 5.3 Polar curves for different span configurations...... 60 5.4 Airspeed for level flight with angle of attack for different span configurations...... 61

5.5 CL/CD with level ﬂight airspeed for different span conﬁgurations of telescopic wing.... 62

5.6 Optimal CL/CD curve with level flight airspeed for the telescopic wing...... 62 3/2 5.7 CL /CD with level flight airspeed for different span configurations of telescopic wing... 62 3/2 5.8 Optimal CL /CD curve with level flight airspeed for the telescopic wing...... 63

5.9 CL/CD curve comparison between optimal morphing wing and 4.3m ﬁxed span wing... 63 3/2 5.10 CL /CD curve comparison between optimal morphing wing and 4.3m ﬁxed span wing.. 63

xv xvi Nomenclature

Greek symbols

α Angle of attack.

µ Dynamic viscosity.

ν Kinematic viscosity.

ρ Density.

σxx Normal stress.

τxy Shear stress.

Roman symbols b Width. c Wing section chord.

CD Three-dimensional coefﬁcient of drag.

Cd Two-dimensional coefﬁcient of drag.

CL Three-dimensional coefﬁcient of lift.

Cl Two-dimensional coefﬁcient of lift.

CM Three-dimensional coefﬁcient of moment.

D Aircraft drag. h Beam cross section height.

Izz Moment of inertia.

L Aircraft lift.

L/D Lift to drag ratio.

M Reaction moment.

Mz Bending moment.

xvii n Load factor. py Distributed load.

Q First moment of area.

R Reaction force.

Re Reynolds number.

S Shear force.

T Aircraft thrust. t Beam cross section thickness. t/c Airfoil relative thickness.

V Velocity.

Vy Shear force.

W Aircraft weight. w Beam cross section width. y Perpendicular distance from the neutral axis to a point on the beam cross section.

Subscripts

A Point A.

AB Beam AB.

B Point B.

C Point C. c Cruise.

CD Beam CD. d Dive. l Loiter. m Maneuver. s Stall. to Take-off. x, y, z Cartesian components.

xviii xix Glossary

AAW Active Aeroelastic Wing. AdAR Adaptive Aspect Ratio wing concept. AFRL Air Force Research Laboratory (USA). AoA Angle of attack. CFD Computational Fluid Dynamics. DARPA Defense Advanced Research Projects Agency (USA). DLR Deutsches Zentrum fur¨ Luft- und Raumfahrt, German Aerospace Center. DMF Dynamic Modulus Foam. EADN Enhanced Adaptive Droop Nose. EAP Electro Active Polymer. EASA European Aviation Safety Agency. EU European Union. FE Finite Element. FishBAC Fish Bone Active Camber. HECS Hyper Elliptic Cambered Span. LLT Lifting Line Theory. MAS Morphing Aircraft Structures Program by DARPA. MAV Micro Air Vehicle. MDO Multi-Disciplinar Optimization. MIT Massachusetts Institute of Technology. MORPHLET MORPHing WingLET. NASA National Aeronautics and Space Administration (USA). NOVEMOR Novel Air Vehicle Conﬁgurations: From Flutter- ing Wings to Morphing Flight. PTFE Polytetraﬂuorethylene. RPV Remotely Piloted Vehicle.

xx SADE Smart High Lift Devices for Next-Generation Wings. SARISTU Smart Intelligent Aircraft Structures. SLE Smart Leading Edge. SMA Shape Memory Alloys. SMP Shape Memory Polymer. UAV Unmanned Aerial Vehicle. UHMWPE Ultra-High-Molecular-Weight-Polyethylene. V-n Velocity vs Load diagram. VSW Variable Span Wing.

xxi xxii Chapter 1

Introduction

1.1 Motivation

The aeronautical industry currently focuses on developing an environmentally efficient aviation with work and research being done on reduction of emissions, engines and alternative fuels, air traffic management, safety aspects of air transport, etc. Morphing technologies belong to the new fields and approaches to aircraft design that strive to find improvements in aircraft efficiency.

The extend to which morphing technologies can improve the performance of an aircraft is closely related to its role and whether it is frequently changed throughout the aircraft lifetime and missions. A multi-role aircraft can beneﬁt greatly from morphing technologies that can adapt it to the situation at hand, while an aircraft that remains in the same ﬂight conditions throughout the majority of its service time has less to gain by adapting to the secondary conditions.

In the past decades, unmanned aerial vehicles (UAV) have seen a surge in interest and development for both military and civil applications. These aircraft provide aerial platforms with lower risk and lower operational costs. They are used for varied purposes, involving reconnaissance, surveillance and attack missions in the military camp and patrol, land management, science research and delivery tasks in the civilian applications. The different purposes of a UAV might be used as well as the necessary ﬂexibility to alter objectives mid-mission makes them good candidates to test the beneﬁts of morphing technologies.

This work approaches these technologies as a solution to a specific design problem. Although many morphing concepts and devices have been studied and researched comparatively ”very little has been reported about the benefits and penalties of such technologies, particularly as systems integrated in the whole aircraft environment.” The subject of the evaluation of the benefits and penalties of the morphing technologies is quite complex and can only be made by comparing a baseline non-morphing aircraft to the morphing aircraft. The choice of fixed wing aircraft to compare with the morphing aircraft is vital and depends on the missions and the role of the aircraft: a low subsonic speed morphing aircraft will outperform a high subsonic speed fixed wing aircraft at a mission dominated by low speed flight conditions, [1].

1 1.2 Morphing Overview

Morphing aircraft technology is an expression that tries to synthesize shape and/or structure trans- formation technology in the specific case of aircraft. However, there is no consensus about its exact definition. Etymologically morphing derives from the Greek –morphos, means shape. Morphing, as defined by Rodriguez [2], means to transform, to become “one of various distinct forms of an organism.” This association to biology is pertinent, since a lot of morphing aircraft technologies consist of biomimicry approaches that seek to emulate biological structures and functions, such as the flight of birds, insects, etc. [1,2]. In aeronautics, morphing is a research field that looks to make aircraft more efficient and able to operate under a wide range of varying flight conditions. Commonly it is referred as Bowman et al. [3] described: ”a set of technologies that increase a vehicle’s performance by manipulating certain characteristics to better match the vehicle state to the environment and task at hand.” Using this definition, the Wright Flyer is the first morphing aircraft since it used wing warping as a form of bank angle control. In addition, technologies such as flaps, slats, retractable landing gears and others associated with conventional aircraft are considered morphing technologies. This contradicts the connotation of radical changes in shape or shape changes that require futuristic technologies and materials that morphing carries, [1–3]. For that reason, these devices present among conventional aircraft, flaps, slats, retractable landing gears, etc. are associated to fixed wing technologies and based on traditional materials and mechanical components, such as hinged surfaces, and in opposition morphing technologies are associated with novel materials and actuators with promising new interesting behavior, [1,4]. Even if the expression fixed wing may misleadingly imply that conventional aircraft with conventional wings do not have moving parts, its use is helpful to clarify the concept of morphing aircraft technology since this technology intends to provide capabilities beyond the conventional fixed wing aircraft. More- over, it is helpful as a baseline comparison in the assessment of benefits and penalties of a morphing technology, [1,4]. Regardless of the type of technology and its extent, materials, actuators and structures it may use the objectives and aimed capabilities of morphing technologies are clear. As Mcgowan et al. [5] identified they aim to increase “the adaptability of the vehicle to enable optimized performance at more than one point in the flight envelope.” From a design standpoint, this means a morphing aircraft can be viewed as if it were multiple different vehicles. Whereas conventional aircraft commonly are optimized for just one design point, [5–7]. Conventional aircraft are generally optimized towards a single design point therefore this characteristic of morphing aircraft enables them to perform the missions of several conventional aircraft. Addition- ally, this ability of having a design optimized for more than one flight phase allows a morphing aircraft to perform more missions efficiently or even to perform a mission that would require more than one conventional aircraft, [5]. The table 1.1 summarizes the effect of geometric wing parameters have on the aircraft performance. As an example, high swept wings are more adequate for transonic and supersonic

2 flight while low swept wings are capable of greatly efficient low speed flight. This possibilities have been motivating engineers to go further on the development of ambitious morphing concepts, [5–7].

Table 1.1: Effects of wing geometric parameters on aircraft performance from [7]

Wing Parameter Effects of Variability – all other parameters unchanged

Plan Area + Increased lift, load factor capability - Decreased parasitic drag Aspect Ratio + Increased L/D, loiter time, cruise distance, turn rates Decreased: engine requirements - Increased maximum speed; Decreased parasitic drag Dihedral + Increased Rolling moment capability, lateral stability - Increased maximum speed Sweep + Increased critical Mach no., dihedral effect Decreased high-speed drag

- Increased CLmax Taper Ratio Wing efﬁciency (spanwise lift distribution); Induced drag Twist Distribution Prevents tip stall behavior; Spanwise lift distribution Airfoil Camber Zero-lift angle of attack, airfoil efﬁciency, separation behavior Airfoil Thickness/Chord Ratio + Improved low-speed airfoil performance - Improved high-speed airfoil performance Leading Edge Radius + Improved low-speed airfoil performance - Improved high-speed airfoil performance Airfoil Thickness Distribution Airfoil characteristics, laminar/turbulent transition

Morphing concepts can provide fantastic features to aircraft, and in the future, according to Rodriguez [2], morphing aircraft technologies will be used to:

1. ”Improve aircraft performance to expand its ﬂight envelope.”

2. ”Replace conventional control surfaces for ﬂight control to improve performance and stealth.”

3. ”Reduce drag to improve range.”

4. ”Reduce vibration or control ﬂutter.”

The design of morphing mechanisms up to their implementation faces the difficult task of not only providing aerodynamic efficiency improvements but also keeping the aircraft structurally strong, lightweight and controllable. The necessity to adapt shape to the flight regime means the internal structure of morphing aircraft should be able to change in an efficient and reversible manner. The skin should be able to accommodate the change of internal structure by either sliding or being flexible which renders it a virtually non-load-carrying element and thus degrades the overall structural stiffness. However, the skin still needs to be strong enough to transmit the aerodynamic loads. Additionally, surface continuity and smoothness and guaranteeing that the gaps and seams are minimal in all aircraft configurations is deeply related to how aerodynamically efficient the aircraft can be. These complexities lead to a heavier structure and ”demand new effective ways of changing structural shape with less actuation force, along with materials with good structural properties: high strain with no plastic deformation, good fatigue resistance, low environmental impact, etc”, [4,7]. To change the configuration of aircraft in-flight are required higher actuation forces to cope with aerodynamic and friction forces. This asks for stronger and heavier actuators, which increase power requirements. For that reason, it is vital to search new clever ways of actuation that may use aerodynamic

3 forces for supplemental actuation. Morphing aircraft control faces the dilemma of using conventional surfaces or using asymmetrical morphing. Highly complex internal structures reduce space for conventional control systems. Moreover, the effectiveness of control surfaces has to be secured for all configurations without compromising the efficiency of the lifting surfaces. Algorithms can be developed to use morphing mechanisms asymmetrically as a means of control. However, this requires a power efficient actuation system that can perform reliably at the high frequencies demanded. Additionally, it is vital that the control of the morphing systems should assure the internal coordination of the moving substructures and actuators in order to minimize the human input in the underlying process. Ultimately, the integration of all these components should not compromise field operations and reliability of the aircraft nor the ability for maintenance and repairs, [4,7]. Since morphing aircraft are likely to be heavier, during the design, development and implementation of these devices the benefits of these technologies should be weighted against its drawbacks. As de- mystified in [5] morphing aircraft technologies are entirely case specific: ”results can only be compared in context with the specific mission specifications (speed, Reynold’s number, etc.) and the trade-offs of each specific technical approach (weight, power, etc.).” Thus, the benefits and trade-offs of morphing technologies cannot be generalized to a single answer. Also stating that in a aircraft design context the concern about ”how” morphing is done, either using futuristic technologies and material or more conventional approaches, is not the critical aspect and that ”potentially enables artificial constraints that reduce innovation and limit the use of more optimal applications.” Ultimately, morphing technologies and subsystems should be evaluated by their effect on the overall aircraft mission performance, [5]. Aircraft morphing concepts are differentiated into three different categories: wing, fuselage/tail and engine morphing. From now on, this work will highlight only the wing morphing, [7].

1.3 Objectives

The objectives of this work are:

• Design a telescopic wing for a medium-altitude and medium-endurance patrol and surveillance UAV.

• Study morphing technologies capable of providing span change to a unmanned aerial vehicle.

• Design a structure capable of supporting the aircraft certiﬁcation loads.

• Assess the value of the morphing solution found when compared to a conventional ﬁxed wing conﬁguration as a whole aircraft package.

1.4 Thesis Outline

This work is organized into 6 different chapters:

1. Introduction – presents an overview on the topic that will be discussed, Morphing.

4 2. Morphing Background – a historical perspective and the current state-of-art is presented.

3. Telescopic Wing Design Strategy – the initial problem and project speciﬁcations are given. The telescopic wing mechanism and subsequent project aspects are explored.

4. Results – the results of the telescopic wing design are presented.

5. Aerodynamic Performance Studies – the morphing wing performance is compared with a ﬁxed wing counterpart.

6. Conclusion and Future Work.

5 6 Chapter 2

Morphing Background

This chapter presents a bibliographical review and the theoretical background on morphing: from a historical perspective examples of morphing aircraft, the more relevant projects and their area of investigation and the recent state of the art with the current technical difﬁculties.

2.1 Historical Perspective

One might feel astounded by morphing aircraft futuristic designs and concepts but morphing’s basic ideas predate even the first flight of the Wright brothers. This comes from the biomimmicry aspect of human technological advance, our desire to recreate the solutions nature presents. Early aviation enthusiasts observed birds fly and how they change their wing shape to instinctively adapt to the desired flight mode (soar, dive, etc), [8–11]. In 1890, Clement´ Ader proposed with a great deal of vision a morphing wing design, that can be seen in figure 2.1, and developed ideas for the future of aviation warfare later published in 1909. His description of the general military aircraft and of reconnaissance aircraft: ”Whatever category airplanes might belong to, they must satisfy the following general conditions: their wings must be articulated in all their parts and must be able to fold up completely. When advances in aircraft design and construction permit, the frames will fold and the membranes will be elastic in order to diminish or increase the bearing surfaces at the wish of the pilot.”

Figure 2.1: Clement Ader’s Eole – a (non-ﬂying) shape changer, [8]

7 “The wings will be extendable in flight and their surface will be increased or decreased at will. These airplanes will be characterized by their agility but will also be of solid construction.” He also stated that their ”wings will be bat-type or preferably bird type, long and narrow” and ”adjustable, so that in flight they can be reduced by a half or a third or even less”, [8, 12]. For the Wright brothers aircraft control was an essential aspect to powered flight and they used wing warping as a mean of aircraft roll control in their 1903 flyer. The twisting of the wing structure resulted in the ability to increase the angle of attack of the outboard sections of the wing and thus achieving an asymmetrical lift distribution, [9]. Historically, there have been two different ways to design and build structures capable of achieving the required in-flight wing shape change for aircraft control and operation, [9]:

1. Use internal actuators to bend and twist wing structure to create the appropriate deformations on compliant surfaces, much like birds control their wings.

2. Deploy small discrete wing components composed by hinged surfaces and structures to reshape the aircraft’s wing in-ﬂight, such as ﬂaps, slats, ailerons, etc. These are present in many of today’s conventional aircraft.

Ivan Makhonine was a pioneer in large in-flight wing planform change to reduce stall speed for better take-off and landing operations while keeping the possibility of a small wing for high-speed flight. He developed the telescopic wing aircraft MAK-10 that first flew in August 1931, figure 2.2. Makhonine’s telescopic wing was capable of an increase in span of 60% (from 13 to 21 meters) and in area of 57% (from 21 to 33 square meters), [9, 10].

Figure 2.2: Ivan Makhonine’s MAK-10 telescopic wing aircraft (1931), [9].

Recently, with the developments in smart technologies, materials, sensors, actuators and hardware, new morphing approaches and old concepts as well have been researched. In fact, in the last few decades, there were several signiﬁcant and expansive research programs supported by the National Aeronautics and Space Administration (NASA) and the Defense Advanced Research Projects Agency (DARPA) as well as collaborative projects funded by the European Union, [9].

8 2.1.1 Main Aircraft Morphing Research Programs

NASA’s Morphing Program

One of the earlier programs dedicated to morphing aircraft technologies is NASA’s Aircraft Morphing program. Created in the mid-1990’s at Langley Research Center and Dryden Flight research Center it included fundamental multidisciplinary research efforts in smart materials, adaptive structures, micro flow control, optimization and controls. The main motivation behind this project was never to develop complete morphing vehicles but to develop and assess morphing concepts and technologies that would enable multi-point efficiency, [2,5,9, 13]. The object of study was centered around adaptability and not just limited to large shape changes. The morphing aim was further defined as ”efficient, multi-point adaptability”. This definition includes very small to large vehicle shape changes achieved by structural or aerodynamic approaches. It also synthesises that morphing technologies strive to be lighter weight and/or more energy efficient than conventional technologies, capable of performing disparate mission requirements and have versatility and resilience to varying conditions or problems, [5]. In 2001, the project released the figure 2.3 providing an inspiring vision of the future of aeronautics to engineers, students and countless airplane enthusiasts. However, it further added to the impression that morphing means only radical shapes and changes. NASA’s ”Aircraft Morphing Program” paved the way forward with innumerable new morphing concepts and technologies. ”Most importantly, this effort led to the education and development of young researchers in this area. At its height there were 80–100 NASA researchers involved with university and industry partners”, [9].

Figure 2.3: 21st Century Aerospace Vehicle (artist rendering), [14].

DARPA’s Morphing Program

DARPA is the primary U.S.A. Department of Defense research agency and its mission is to identify and advance new technologies to improve military capabilities through very fundamental research and

9 the development of complex systems including aircraft and spacecraft. During the 1990’s, DARPA developed several smart structure projects such as Smart Materials and Structures Demonstration Program, Smart Rotor Program and the Smart Wing Program. These programs displayed the value of smart material based actuation and control systems applied to helicopter rotor blades, transonic aircraft wings and undersea vehicles and engine inlets, [9]. The DARPA Morphing Aircraft Structures program (DARPA/MAS) used a Hunter-Killer unmanned aerial vehicle (UAV) concept as an example of a morphing system capabilities. In this case the aircraft combined features of a reconnaissance aircraft with an attack aircraft. Its objectives were defined as “the design and fabrication of effective combinations of integrated wing skins, actuators and mechanisms, structures, and flight controls to achieve the anticipated diverse, conflicting vehicle mission capabilities via wing shape change”, [9]. DARPA/MAS program began in late 2002 and ended in early 2007. It had three phases:

• Phase 1: Concept proposal.

• Phase 2: Design/Build with a wind-tunnel structural test.

• Phase 3: Flight test demonstration.

Three contractors (Lockheed Martin, Hypercomp/NextGen and Raytheon Missile Systems) were challenged to produce concepts tied to missions that required controlled, in-flight increases in wingspan, wing planform area and shape, [10]. Lockheed Martin’s model changes its configuration to efficiently fly at low and high speeds. Its morphing wing folds to reduce planform and wetted area and for this reason this design requires substantial use of advanced materials to guarantee surface smoothness and continuity, fig.2.4. HyperComp/NextGen’s aircraft model is capable of considerable in-plane shape changes and reduction of surface area. The change to the wing planform affected the span, chord and sweep; Five different configurations were possible. And the model presented by Raytheon Missile Systems comprises a telescopic wing design that enables an increase in 50% of wingspan and provides up to 75% more loiter time. The limited volume for structural elements and actuators and the large wing loading are the main design challenges. Despite all three designs had satisfied DARPA’s technical goals, DARPA’s management decided to only continue to phase two with the concepts of Lockheed Martin and HyperComp/NextGen. This phase had the objective of demonstrating the concepts’ ability to perform under realistic aerodynamic loads and to validate analytical prediction methods, [15]. DARPA approved the phase 3 of the MAS program to test morphing flight control with three distinct goals: to ”demonstrate the ability to morph in flight in operational useful time frames and flight conditions while maintaining full flight control(flight test 1); Demonstrate improved morphing aircraft maneuverability (flight test 2); Demonstrate acceptable landing in a variety of conditions simulating a damaged or malfunctioning wing system (flight test 3)”, [9]. The program’s third phase focused on flight-testing a low-speed 200 lb Remotely Piloted Vehicle (RPV) version of both concepts. HyperComp/NextGen ”designed and flight-tested a RPV with 40%

10 (a) (b)

Figure 2.4: (a) Morphing Aircraft developed by NextGen in the DARPA MAS 3 Program, [5]; (b) Lockheed Martin baseline morphing concept, [2]

change in wing area, 73% change in span and a 177% change in Aspect Ratio”, presented in fig 2.4. This feat is one of the largest and reversible change in wetted area displayed in-flight. The scaling of the flight test result to the full-scale aircraft is complex as a result of the differing flight regimes of the RPV (low subsonic speeds and low altitude) and the full scale concept, [5].

The MAS program showed varied morphing wing solutions based on entirely different concepts. The greatest challenge faced is the integration of all aircraft systems (power, materials, controllers and actuation) into a confined aerodynamic shape that provides a fully operational aircraft. Weisshaar [9] stated that the principle obstacle hindering the production of full scale flight demonstrator based on the wind tunnel testing of small scale models was ”the inability to define a challenging military mission that captured the imagination of funding agencies.” The lack of a defined role, the requirements and objective for the aircraft cannot be fully defined to support the next technical steps of the concept design, [9].

Even though the MAS program was completed without the use of exotic materials, one of the recog- nised challenges and limiting technologies for morphing aircraft and devices is the demand for material that can be easily deformed but can still carry substantial loading, [9].

SADE

In 2008, started the Smart High Lift Devices for Next Generation Wings (SADE) project, a European collaborative effort co-funded by the European Union (EU) and coordinated by the German Aerospace Center (DLR) with 12 other partners. According to EU’s Transport Research and Innovation Monitoring and Information System website ”SADE aimed at a major step forward in the development and evaluation of the potential of morphing airframe technologies and contributed to the research work calling for the reduction of carbon dioxide and nitrogen oxide emissions through new intelligent low-weight structures”, [16].

The research focused on seamless smart high lift devices to enable the laminarisation of slim high- aspect-ratio wings and thus reduce commercial aircraft parasitic drag and fuel consumption. This project ﬁnished in 2012. It included the development and extensive wing tunnel testing of Smart Leading Edge (SLE) concept, a cross section of the device is displayed in ﬁgure 2.5,[16, 17].

11 (a) (b)

Figure 2.5: (a) SADE’s Smart Leading Edge concept, [17]; (b) SARISTU wing demonstrator with integrated leading, trailing edge smart devices and active winglet, [18]

SARISTU

Smart Intelligent Aircraft Structures (SARISTU) is the other main EU funded morphing project. The initiative coordinated by Airbus brings together 64 partners from 16 European countries. Both SARISTU and SADE are part of EU’s 7th Framework Programme under FP7-TRANSPORT - Speciﬁc Programme ”Cooperation”: Transport (including Aeronautics) section, whose main research objective is ”to develop safer, ”greener” and ”smarter” European transport systems for the beneﬁt of all citizens”, [18–20].

This multidisciplinary project ranged from 2011 to 2015 and its research included several technology streams: morphing structures, integrated sensing and multifunctional materials, with extensive development, ground and wing tunnel testing and validation and integration of the technologies on a fully functional aircraft level, [18, 20].

SARISTU aimed to improve aircraft performance through drag reduction possible with a laminar wing with the integration of different conformal morphing concepts: Enhanced Adaptive Droop Nose (EADN), adaptive wing trailing edge, active winglet morphing and elastomer-based skins for kink-less and gap-less wing surface. In figure 2.5 is presented the scaled demonstrator of the wing with the several morphing devices integrated. The integration of these technologies allows for aircraft control and operation under diverse and distinct conditions while benefiting from the best capabilities of a laminar wing for cruise flight, [18].

Other technological breakthroughs included: ”developments on health monitoring analysis methods; Improvements in the ability to upscale nanocomposites from the basic resin all the way to industrial relevant laminates with complex and large geometries, including conductive nanocomposites and investigation of possible co-bonding with metallic materials”; Lastly, the project demonstrated method of a validation and integration of concepts from materials up to sub-components and components with an incremental perspective aimed at the performance of the ﬁnal model, [4, 18].

12 2.2 Wing Morphing

Morphing aircraft technologies have been developed and studied in an array of levels and approaches, from the ”design of actuators based on conventional or novel materials which exhibit particular responses to different stimuli” to ”morphing concept definitions base on aerodynamic studies involving changes in the aircraft geometrical parameters”, [1]. Wing morphing technologies are described and divided according to the morphing concept which they are related to, and these in order are usually associated with the change of a particular wing geometrical parameter. Three major categories as defined by Barbarino et al. [6]: planform change, out-of-plane and airfoil (see fig 2.6), [1,6, 21].

Figure 2.6: Wing morphing concepts types, [6]

2.2.1 Twist

A wing is said to have twist if the angles of attack of the spanwise sections of the wing are not equal. Twist affects the spanwise lift/load distribution by managing the effective angle of attack of the wing’s spanwise airfoil sections perform under, [22]. Regularly, twist is used to delay wing tip stall when it has already happened in the wing root as to allow the pilot to keep roll control of the aircraft. To achieve this much appreciated stall delay the wing sections at the tip are rotated downwards relatively to the root so as the effective angle of attack of the airfoil at the wing tip is lower. This is referred as ”wash-out”. It can also be used to alleviate bending moment at wing root and decrease induce drag at high lift configurations. Additionally, twist changes the pitching moment on swept wings so to reduce the necessity for actuation of the horizontal stabilizer which helps to reduce tail trim drag. This effect is fundamental in tail-less aircraft configurations, [1, 22]. Twist morphing can be used as an effective form of roll control by asymmetrically altering the produced lift. In 1903 the Wright brothers obtained active roll control by warping the wings of their flyer. NASA’s Active Aeroelastic Wing (AAW) program, active from 1996 through 2005, procured multidisciplinary technology that integrates ”air vehicle aerodynamics, active controls, and structural aeroelastic behavior to maximize air vehicle performance”, [23]. The result was the experimental aircraft X-53, adapted from an F/A-18 fighter aircraft equipped with active wing twist roll control. It used a more flex-

13 ible wing, active leading and trailing edge devices which managed to use the surrounding airﬂow to induce twist on the wing structure with little internal actuation. The concept is viewed as a return to the idea pioneered by the Wright brothers, [1,6,9, 23].

In recent times twist morphing has also been applied to a different class of vehicles: the micro aircraft vehicles (MAV). Typically with a size scale similar to biological systems of flying beings, they use membrane wings where the lack of internal structure makes the implementation of ailerons non-viable. In 2005 Abdulrahim et al. [24] implemented torque rods as lateral control devices, fig 2.7. Testing reported that the ”membrane wings on these vehicles can be morphed with little power but with significant benefits.” The greater authority in vehicle control is achieved by decoupling lateral control from longitudinal control as MAV usually are solely controlled by ruder and elevator, [24].

(a) (b)

Figure 2.7: (a) Digital Morphing Wing and its discrete lattice building-block elements, [25]; (b) MAV wing with torque rod, [24]

This type of device has also been tested in larger scale models, such as a morphing swept wing tailless aircraft. The twist morphing capability eliminates the necessity of winglets and washout to coun- teract the adverse yaw flying wing with elevons usually suffer during maneuvering. Thus taking full advantage of the efficiency of the flying wing. Wind tunnel testing suggested 7% to 28% drag reduction during maneuvering and 15% lift to drag ratio (L/D) increase in cruise flight when compared to an elevon equipped wing, [26].

Other twist morphing concepts were developed and tested. In 2004, spars capable of changing chord wise position to control the position of the ﬂexural axis and allow the wing to twist under aerodynamic loading. In 2010, a torque rod was use to warp the skin of a rotor wing and induce twist. The skin was disconnected at the trailing edge so as to reduce actuation energy. The maximum peak-to-peak angle of twist measured was 27◦ at the wing tip and wind tunnel tests showed that the device was capable of a change up to 0.7 of the CL value in a 0.68m span wing, [1, 27, 28].

Recently, a new twist morphing wing concept, the Digital Morphing Wing, was developed by engineers at the Massachusetts Institute of Technology (MIT) and NASA. The wing internal structure is constructed from discrete lattice elements join with a carbon ﬁber torque tube for twisting the whole structure. The structure is lightweight and simple to fabricate and repair, ﬁg 2.7,[25].

14 2.2.2 Sweep

There are three main reasons to use swept wings in aircraft design, [1,9, 22]:

1. To avoid compressiliblity effects on drag. By decreasing the airstream component perpendicular to the wing it self the wave drag at transonic speeds can be delayed. And by keeping the wing inside the mach cone at supersonic flight with wings designed for subsonic flight instead of dedicated wings which are very inefficient at low speeds.

2. To provide pitch stability for tailless (ﬂying) wings and lateral stability in both roll and yaw for any kind of ﬁxed wing aircraft.

3. To improve longitudinal static stability by reducing the distance between the aircraft center of gravity and the wing aerodynamic center;

The variable sweep wing (swing wing) was created from the necessity of combining efficient low speed flight of low sweep wings with high speed flight of highly swept military aircraft. Pivoting of the wing has been the most used mechanism to change sweep. The first aircraft to feature a swing wing was the 1951 Bell X-5. Its wing also translated forward simultaneously to control the aerodynamic center position. During the 60’s and 70’s this technology was used in several military aircraft, famously the Grumman F-14 Tomcat, shown in figure 2.8, and the Rockwell B-1 Lancer. Due to its heavier weight, complex maintenance and the development of aircraft carrier take-off and landing systems the swing wing technology ceased to be the best solution, [1,6, 10].

(a) (b)

Figure 2.8: (a) Grumman F-14 Tomcat, [29]; (b) Batwing concept conﬁgurations for highlift, climb, cruise, loiter and maneuver, [4]

During DARPA’s Morphing Aircraft Structures program Hypercomp/NextGen developed the batwing concept, figure 2.8, a morphing aircraft concept capable of in-flight sweep angle change. It was equipped with an endoskeleton wing box covered with a deformable polymer skin and actuated via electromotor allowing 5 different sweep configurations. ”This aircraft design moves between five different wing plan- forms and creates wing planform changes in area, span, chord, and sweep that vary by 51%, 36%,

15 110%, and 30 deg, respectively”, [9]. In 2007, a UAV model was flight tested and demonstrated 40% area change 73% span change and 117% aspect ratio change, [1,9]. Variable sweep wing configurations have been researched to a certain degree in the field of Micro Air Vehicles (MAV). A MAV equipped with a wing capable of a sweep change between 15 and 60 degrees to reduce drag by lowering the frontal profile size showed a drag reduction of 60% at a speed of 20 m/s, [30]. A bio inspired design with multi-joint wing simulating a bird’s wing bone structure capable of achieving different wing and sweep setups. The concept provided sweep raging from -30 and 30 degrees both symmetrically and asymmetrically. Results show evidence of better turning radius for swept configurations and with asymmetrical setups sensor pointing in crosswinds up to 44 degrees a big improvement is noticed because ”the common approach to sensor pointing despite crosswinds is turning into the wind and crabbing downrange to periodically point the sensor; however, such an approach is certainly not optimal due to the lack of continuous coverage by the sensor along the desired line of sight”, [31].

2.2.3 Dihedral (Folding and Winglets)

In aircraft design dihedral angle is used primarily for roll stability reasons but can also be a measure to increase the Oswald efficiency coefficient and reduce induced drag, [1,6, 22]. A wing with a positive dihedral angle has improved roll stability. Any small disturbance in bank angle tends to initiate an aircraft fall to the side of the lower wing, creating a wind velocity component that in a wing with positive dihedral angle provides an increase in the angle of attack of that side of the wing and decrease the angle of attack of the higher wing. The wing will produce a differential lift that results in a roll moment that returns the aircraft to null bank angle. Negative dihedral is used to increased aircraft maneuverability by creating a wing configuration that is naturally unstable to small disruptions, [1, 22]. Recently variable dihedral/gull wings have been researched because of their ability to enhance aircraft performance and flight control. Besides controlling stability behaviour and reducing the induced drag (by changing the vorticity distribution) a variable dihedral wing can: control the aerodynamic span; replace conventional control surfaces; and improve the stall characteristics, [6]. The first dihedral morphing aircraft was the IS-1 fighter designed in 1932 by Nikitin-Shevchenko. It was a bi-plane capable of changing into a monoplane to operate at higher speeds. The experimental aircraft XB-70 Valkyrie also used a form of out of plane morphing. Its wing was hinged and the outer panel was able to be downwards rotated to control the lift-drag ratio both at subsonic and supersonic speeds, [6]. The Lockheed Martin concept developed during the DARPA/MAS program had a wing (Z-wing) with multiple articulated section capable of reducing wing span and area, increasing wing sweep and lowering the aircraft wetted area, fig.2.4. The wing folds in to allow for high dash speeds and unfolds for increased endurance, [6,9, 32]. The model went through wind tunnel testing. Besides the wing folding mechanism the model included several technologies including seamless skins and thermopolymer based actuation. Several

16 studies performed computational fluid dynamics (CFD) and finite elements (FE) calculations to evaluate aeroelasticity effects of the wing. Some unexpected behaviors were encountered, though none were found to be adverse, [1,9, 15, 33]. Gull configurations were studied on biologically inspired concepts with avian morphology. The Uni- versity of Florida performed flight tests on micro air vehicles with articulated wing that allow for independent inboard and outboard sections of the wing capable of varying their dihedral angle. The results and aircraft behavior were replicated through computational fluid dynamics (CFD) analyses. The gull wing configuration can be optimized for several mission objectives such as ”long range cruise, endurance, loitering, minimum radius turns, steep descents and sensor-pointing.” The last one can be achieved with asymmetrical dihedral configurations. The optimized configurations converged, in several cases, to the biological solutions, [11, 34, 35].

Figure 2.9: MAV undergoing neutral (top) and positive (bottom) gull-wing morphing, [34].

Morphing winglet structures have been developed either with the objective of aircraft control or to improve wing efficiency at different flight phases. Studies on variable cant winglets reported in [36] showed that in swept wings and with asymmetrical configurations, winglets were able to achieve roll control with proverse yaw. The winglets generated control moments about multiple axis and the longitudinal and lateral control dynamics were coupled. ”A single pair of adaptive winglets cannot replace all the conventional control surfaces and an elevator was included to trim the aircraft, especially during level turns at arbitrary bank angles”, [6]. Another pair of active winglets constituted the solution for the highly coupled control dynamics, fig 2.10. This control system was shown to be ”more effective than a more conventional one at moderate and high lift coefficients, so it could be applied effectively to low speed or high-altitude morphing aircraft”, [1, 36]. In 2007, Ursache et al. [37] developed the morphing winglet (MORPHLET) concept to improve flight performance of narrow body aircraft. A low-fidelity mulitdisciplinary design optimization (MDO) was performed to optimize the cant, twist, and span of each end wing partition to maximize the vehicle performance at three different flight phases: start of initial cruise, start of final cruise, and end of descent. Performance studies predict a 4-5% improvement over fixed winglets, [6].

17 Other morphing winglet concepts shown performance gains in several aspects of ﬂight: reducing stall speed and drag at cruise and maximum speeds, enhancing range, endurance and providing the potential for increased payload, [1,6, 38, 39].

(a) (b)

Figure 2.10: (a) Experimental model as mounted in the wind tunnel, [36]; (b) HECS wing concept schematic, [40]

It was shown by NASA researchers that a hyper-elliptically swept planform wing with a cambered span (HECS), with a separate hyper-elliptical span wise profile, has aerodynamic improvements over flat wings. Several designs attempt to achieve such configurations through morphing have been research. Manzo and Garcia [40] investigated a system comprised of multiple discrete articulated wing sections. Two actuation approaches were developed: a tendon based spool system powered by a DC motor and a shape memory alloys (SMA) based system. Mechanisms that target a smooth and continuous bending of the wing have also been investigated, [6, 40, 41].

2.2.4 Chord

Chord dimension is one of the driving dimensions of the wing area, and therefore has a direct effect on lift, drag and the wing aspect ratio. Spanwise chord variation changes taper ratio and lift distribution of the wing and can make the planform shape of the wing more suitable for flight in different flight situations, [1, 22]. Chord morphing may enable reducing induced drag either by approximating the planform shape of the wing to its most efficient elliptical shape or by increasing the wing’s aspect ratio. From an opposite perspective increasing chord and wing area might enable higher maximum lift values to expand flight envelope’s limits, altough this is not a common approach, [1,6]. The 1937 Bakshaev LIG-7, fig 2.11, was probably the first aircraft capable of chord variation in flight. The USSR aircraft was equipped with telescopic wing sections that extended during take-off and landing controlled manually by tensioned steel wires, achieving a 44% increase of the wing area, [9]. The conventional changing in chord length is done through leading/trailing edge-flaps, usually actuated by screw system. However, since fixed wings possess components such as spars, fuel tanks and

18 (a) (b)

Figure 2.11: (a) Bakshaev LIG-7 schematic, [9]; (b) Sliding Rib concept developed by Cornstone Re- search Group, [42] other additional components that make it structurally complex it is difficult to accommodate the necessary mechanisms for chord morphing without serious weight penalty or structural drawbacks, [4,6]. Cornstone Research Group developed an interpenetrating rib mechanism to change the chord length by means of miniature DC motors and lead screws, fig 2.11. The airfoil shape was maintained by a flexible honeycomb material that transferred aerodynamic loads. The use of shape memory polymers (SMP) heated by thin wires was suggested for the skin, [6, 42]. Perkins et al. [43] also experimented with dynamic modulus foams (DMF) as means of both actuation and airfoil shape. ”Although the prototype wing section was able to extend the chord upon heating, it could not return to its original shape upon cooling because of the low recovery stress of shape memory foams”, [6]. Besides the added structural weight and complexity, the smooth operation of the mechanism under aerodynamic load and maintaining chord-wise bending are significant challenges, [1, 43]

(a) (b)

Figure 2.12: (a) Chord extension through the use of extendable trailing edge plate, [44]; (b) Chord extension through elastic deformation of structure, [45]

Most reported morphing chord technologies are applications to rotary wing aircraft. The structure of rotor blades is characterized by a single D-spar and a honeycomb ﬁller. Coupled with a simple structure

19 the interior space of rotor blades is not usually used for other function components, such as fuel tanks and control surfaces, this makes them adequate for the implementation of chord morphing, [6]. Several studies increased the chord by extending a flat end plate through a slit at the trailing edge over a section of the rotor blade. Reported maximum altitude, gross weight and speed increase along with a reduction on main rotor power at the boundaries of the flight envelope. Another concept employs a morphing zero Poisson’s ratio cellular structure design coupled with flexible skins to allow for chord extensions up to 30%, [1, 44, 45]. Both can be observed in the fig 2.12.

2.2.5 Camber

In aerodynamics, camber is a parameter of the effective curvature of an airfoil. The mean camber line, illustrated in ﬁg 2.13, is a line equidistant from the upper and lower surfaces of an airfoil; and the chord line corresponds to a straight line which joins the leading and trailing edge points. The maximum camber, usually referred just as camber, is the maximum distance between the camber and chord lines, measured perpendicularly to the chord line, generally presented as a percentage of the chord length, [22, 46].

Figure 2.13: Cross section of an airfoil, [4].

The camber of an airfoil affects its performance in terms of lift and drag as well. Generally an increase in camber has a positive effect in the lift produce by the airfoil. Relative to a symmetrical airfoil a positive cambered airfoil can produce higher lift coefﬁcient at the same angle of attack, [1, 46, 47].

The typical effect of camber is a translation of airfoil Cl vs AoA curve upwards, providing a higher maximum lift coefficient, fig 2.14. Stall occurs at a lower geometrical angle of attack but at a higher absolute angle of attack (relative to the null lift angle of attack). The null lift angle of attack is negative and its absolute value increases with camber. The drag coefficient increases with the lift coefficient; however, for the same lift coefficient a cambered airfoil produces less drag than its symmetrical counterpart, [1, 46, 47]. For laminar flow airfoils which possess a range of lift coefficient values for which the corresponding drag coefficient are considerably lower, the so called ”drag bucket”. The Cd vs Cl curve suffers a translation due to camber, fig 2.14, which can be used to place the drag bucket on positive lift coefficient values and minimize drag for the more commonly used lifts, [46]. Common hinged control surfaces such as ailerons, rudder and elevator are devices which change the airfoil camber of the spanwise section they are inserted into by deflecting a part of the trailing edge generating a desired variation of the lift produced, [6, 48].

20 (a) (b)

Figure 2.14: Airfoil camber effect on: (a) Cl vs α curve; (b) Cd vs Cl curve, adapted from [46, 47]

Flaps and slats are other traditional devices widely used that change the airfoil camber of the wing sections they are located. In figure 2.15 are shown different conventional high lift devices configurations. These are known as high lift devices which enable wing configurations capable of generating the high lift coefficients that allow the aircraft to maintain level flight at low speeds essential in takeoff and landing operations. Often chord is simultaneously increased with camber. Wing configurations with flaps possess discontinuities in the airfoil contour and its upper surface is highly curved which can lead to flow separation. Low speed flight with flow characterized by lower Reynolds numbers and high angles of attack can aggravate this condition, increasing drag and drastically decreasing the lift generated. To delay flow separation and mitigate its effects slots are used to energize the airfoil boundary layers. This increases the number of elements and complexity of these devices, [1, 47].

(a) (b)

Figure 2.15: High lift devices conﬁgurations: (a) Cross section schematic, [47]; (b) Landing and take-off ﬂaps, [10]

Camber morphing concepts are largely aimed towards performing the same functions as high lift devices and hinged control surfaces but with increased efficiency and a wider range of flight phases, mainly possible by maintaining smooth and gapless airfoil surfaces that allows for a more laminar air flow in cruise flight conditions, [1,4, 18].

21 Besides lateral control and high lift configurations one preeminent aspect of camber morphing is the ability to achieve higher lift to drag ratios in all flight phases by changing airfoil shape to the optimal configuration for each flight phase. The optimal airfoil for a certain flight phase will perform better in those conditions than a generic airfoil as it can achieve higher lift over drag coefficients and thus be more efficient. Optimal airfoil shapes at most operating conditions can be achieved by varying the camber and the leading edge thickness. The more possible intermediate configurations and thus the more adaptable to the optimum configurations, the greater are the aerodynamic benefits, [6, 49–51]. Other applications include: improved periodic cruise with camber, thrust and elevator control; reducing transonic drag rise; managing lift distribution in certain maneuvers to allow for lighter structural designs; increase maneuvering performance; etc, [6, 49, 52, 53].

• Dynamic and cyclic cruise flight with camber, throttle and pitch control able to perform better than static cruise configurations. Alternating between a high thrust and low drag phase and a high lift configuration glide phase was shown to be more efficient cruise condition than a static fixed airfoil option, [52].

• Reducing the curvature of the upper airfoil surface provided a delay and reduction of the transonic drag rise, [53].

• Changing lift distribution over the span to reduce wing tip lift and concentrate a higher portion of the wing’s lift near the root as a means to reduce the root bending moment during certain maneuvers to enabling lighter structural designs, [49, 54, 55].

• Increase maneuvering performance. Conﬁgurations that provide marked increases in lift at high subsonic speeds but with an increase in proﬁle drag are not optimal in most situations however they can improve maneuvering performance of combat aircraft in such situation where ”low response times are essential and the drag rise” ”may not be so critical as that in a steady cruise”, [49, 54, 55].

• optimization to unsteady viscous ﬂow, [56].

Camber morphing is the most researched approach to aircraft shape morphing. Studies on camber morphing concepts applied to fixed wing aircraft as well as to rotorcraft are numerous. Camber morphing can be done by changing the whole underlying structure (ribs) or by the employment of leading or trailing edge devices placed on a fixed wing box much like slats and flaps. One of the first examples of a variable camber wing the 1920 Parker [57] wing is capable of changing the camber of the whole profile for adapting to high and low speed flight. Slats and leading edge devices’ main effort is not to increase the lift produced but to allow for higher stall angle of attack by preventing flow separation in the front of the airfoil, [1,4, 47, 57]. The major differences between camber morphing concepts are in the means of actuation procured. Smart materials increase the spectrum of actuation solutions. Actuators based on piezoelectric materials, shape memory alloys (SMA), electro active polymers (EAP) as well as conventional actuators for camber morphing were investigated so far, [1,6, 58].

22 Conventional actuation

Usually composed of internal mechanisms, conventional actuation is based on classical elements such as threaded or geared components moved by electromagnetic, hydraulic, pneumatic or engine power. There are several studies on airfoil and camber optimization to different flight conditions and maxi- mization of the lift to drag ratio both in commercial aircraft and commercial applications, [49–52, 54–56]. Monner [59] developed an approach capable of changing airfoil camber starting after 90% of the wing chord of a commercial aircraft to better cruise performance. The concept modified the trailing 50% of the Fowler flap to accommodate the varying geometry as on civil transport aircraft the ailerons and Fowler flaps are positioned in this region, [59]. In [60] it is proposed a variable flap for improved efficiency composed by conventional materials and actuators. Tests showed an average drag reduction of 2.7% at low speeds when compared to the clean airfoil and considerable actuation energy reduction (40%) compared to a plain flap system. Under the EU’s 7th Framework Programme several projects have developed morphing leading and trailing edge devices with a focus on commercial aircraft. ”The SADE project developed several concepts of internal mechanisms to be used in leading edge devices”, its final design the kinematic chain is illustrated in fig. 2.5. The SARISTU project developed and integrated in the same wing box both leading and trailing edge devices. Recently the european project Novel Air Vehicle Configurations: From Fluttering Wings to Morphing Flight (NOVEMOR) developed a droop nose demonstrator with compliant based actuation. ”The device featured a fiberglass composite skin with optimized three-dimensional thickness distribution, which was supported by topology-optimized superelastic nickel titanium and aluminum internal compliant mechanisms topology-optimized superelastic nickel titanium and aluminum internal compliant mechanisms”, [17, 18, 61]. In 2016, Aguiar [4] studied a smart leading edge device based on compliant magnesium alloys. The smart droop nose design used provides a variable camber wing with the absence of gaps, seams and steps making it a viable option for the future generation of high lift devices that must reduce aircraft drag and noise. The airfoil shape change was managed by a compliant mechanism and a stringer connection with the skin. A magnesium alloy, being a low density material with a high strength to weight ratio, was tested for its potential aeronautical applications. A structural optimization of the compliant internal structure was performed based on the alloy properties and the desired airfoil shape and change, [4]. FlexFoil is a variable geometry control surface technology developed by Flexsys. A compliant morphing trailing edge wing to improve range and endurance of a high altitude aircraft was designed, fabricated and flight tested, [64]. The aircraft wing can be seen in figure 2.16. The used compliant structures exploit the natural elasticity of aviation grade materials, providing large control deformations (from -9◦ to 40◦ ) and still retain a smooth and gradual surface as to maximize the laminar bondary layer extension over the wing and prevent flow separation. Flight testing revealed that laminar flow was maintained over 60% of the airfoil chord and that range improvement was over 15%. Later the technology FlexFoil was flight tested in Gulfstream GIII Business Jet showing fuel savings up to 12% and reducing the noise in 40%, [62, 64].

23 (a) (b)

Figure 2.16: (a) Compliant trailing edge device with fairings (transition sections), [62]; (b) Flap with eccentric actuator device, [63]

Di Matteo et al. [63] designed and analyzed a morphing trailing edge flap for a large aircraft high lift wing. An eccentric beam actuation mechanism was used to smoothly deflect the flexible rear part of the flap with an open sliding trailing edge, fig. 2.16. ”Nonlinear structural modeling was used for stress-strain and structure behavior under actuation and aerodynamic loading analysis”, [1, 63, 65]. Yokozeki et al. [66] developed a seamless flap like concept based on a flexible corrugated structure. The corrugated structure that composes the trailing part of the wing is equipped on a fixed wing box and actuated via a pulley and wire system connected to the trailing edge region. ”Nonlinear finite element analysis was used to estimate the morphing deformation and the actuation force, and the feasibility of the morphing system was confirmed.” The wing concept was tested in the wing tunnel in Reynolds number ranging from 500,000 to 1,500,000 and deflection angles up to 30◦ . The tested prototype was manufactured using a carbon fiber reinforced polymer material for the corrugated structure and the upper thin skin. The morphing concept presented has better lift properties than plain flap configuration and linearity between the deflection angle and lift coefficient, [66]. The biologically inspired concept named Fish Bone Active Camber (FishBAC) developed in [67] is ”composed of four main elements: a compliant skeletal core, a pre-tensioned elastomeric matrix composite skin, an antagonistic pair of tendons coupled to a nonbackdriveable spooling pulley as the driving mechanism and a non-morphing main spar”, fig. 2.17. The compliant core structure consists of a central bending beam with stringers branching off to support the skin. It is designed to have high anisotropy, with low chordwise bending stiffness but high spanwise bending stiffness. The actuation system is symmetrical and allows for bidirectional changes in camber. Its upper and lower tendons are positioned equidistant from the bending spine so to be under equivalent strains and simplify the actuation control, [67, 68]. A FishBAC wing tunnel model was tested and compared to a traditional discrete trailing edge flap model. The flapped and FishBAC airfoils display similar performance in terms of the lift coefficients generated over the range of camber changes induced; however, the FishBAC airfoil generates considerably less drag for equivalent levels of lift. The maximum obtainable lift-to-drag ratio was shown to be 20% to 25% higher and additionally it is sensitive to the angle of attack, [67, 68].

24 Figure 2.17: FishBAC morphing airfoil wind tunnel at rest and deﬂected downwards, [68]

Recently, a study on morphing elastically lofted transition for active camber control surfaces aim to address the gap present at the spanwise ends of the control surfaces and reduce the significant amounts of vorticity, noise, and drag generated. Computational fluid dynamics analysis of the desired transition shape indicates both an increase in lift and a decrease in drag produce by the wing. The complex three-dimensional shape change required is supported by carefully tailored ribs that have the necessary amount of bend/twist coupling to achieve the desired distribution of trailing edge angle along the span. The concept has no internal actuation system an no moving parts to facilitate its integration on new wing designs. An initial prototype demonstrator capable of large deflections and smooth transition surfaces was fabricated, [68]. Besides fixed wing applications, variable camber rotor blades have also been explored to address retreating rotor blade dynamic stall helicopters encounter near the limits of the flight envelope. Using a leading edge compliant structure actuated by a linear electromagnetic actuator to morph into the airfoil optimized to delay stall at high angles of attack. The device is actuated one time per rotor revolution de- laying blade stall and thereby providing substantial gains in forward speed, maneuverability, and payload capacity. A bench-top model was fabricated which demonstrated 6Hz flap response rate, [69]. Recently leading edge control surfaces were researched for possible application for control straight level flight of micro air vehicles in turbulent flow. ”Higher actuation rates produced dominant leading edge vortices and hence a transient lift enhancement over the airfoil. Lift spikes from high rate actuations could be exploited to compensate for the high frequency perturbations from gusts”, [70].

Piezoelectric actuation

Piezoelectricity is the electric charge that accumulates in the material in response to applied mechanical stress. The piezoelectric effect is a reversible process and piezoelectric materials react to input voltage with high forces. Although the linear displacements are usually very small (µm), these materials have a wide bandwidth and high frequency responsiveness. This initially caused interest in terms of vibration control and applications in aeroelasticity for vibration control in wings and rotor blades have

25 been studied, [1,4,6]. Testing of a remote piloted aircraft that uses an adaptive wing with piezoelectric actuators is reported in [71]. The aim of the study was to evaluate the ”feasibility of applying piezoelectric actuators to suppress aeroelastic vibrations in a flexible aircraft so that the aeroelastic response can be tailored to comply with specified dynamic performance characteristics”. The wing was equipped with conventional aerodynamic control surfaces (flaperons) for performance comparison. Closed-loop buffet attenuation, gust response alleviation and flutter suppression tests indicated considerably greater vibration reduction than that attained using conventional aerodynamic control surfaces [71]. DARPA/AFRL/NASA Smart Wing program, led by Northrop Grumman Corporation under the DARPA Smart Materials and Structures initiative with the intend to improve military aircraft aerodynamic performance in lift to drag ratio, maneuver capabilities and aeroelastic response through the use of smart materials and structures. The project consisted of two phases and hinge-less, smoothly contoured control surfaces with chord-wise and span-wise shape variabilities were developed. Phase 1 focused on static wing-shaping concepts and used low bandwidth actuation based on shape memory alloys. Phase 2 investigated high frequency response actuation involving piezoelectric actuation. The final wind tunnel model comprised 10 eccentric-driven segments connected together by a continuous outer skin and a flexible hinge pin at the trailing edge tip. The smart trailing edge control surfaces were deployed up to 20o in less than 0.33 s in various spanwise uniform and nonuniform shapes, including linear and quadratic ramps, sine, and cosine waves, fig. 2.18,[72].

Figure 2.18: Spanwise shape variations of the smart trailing edge control surface, [72]

Vos et al. [73] reported increased roll control authority of a 1.4m span unmanned air vehicle using morphing ailerons based on piezoelectric actuators while reducing weight, part-count, and power consumption when compared to conventional servo-actuators. An elastic skin covering the outside of the wing generated the axial precompression in the piezoelectric elements to magnify control deflections and forces simultaneously. The control surfaces were capable of deflections up to ± 3.1◦ at actuation frequencies up to 34Hz, [73]. The Virginia Tech Morphing Wing Design Team designed, fabricated and tested an unmanned air vehicle equipped with variable camber morphing control surfaces. All conventional control surfaces (ailerons, elevator and rudder) were actuated via piezoelectric macro fiber composites. The first flight was on April 2010, making this vehicle was the first fully solid-state piezoelectric controlled, free flight

26 aircraft to be flight-tested, fig. 2.19. Its control surfaces employed no multi-piece parts or mechanisms, nor any means mechanical amplification for the piezoelectric actuators. All systems were powered by a single Lithium polymer battery, including the electric motor-driven propulsion system, [74].

Figure 2.19: The morphing aircraft before ﬂight number 6 on April 29, 2010, [74]

The high bandwidth response piezoelectric materials are capable of making them particularly apt to rotor blade applications. In 2008 Grohmann et al. [75] demonstrated the active trailing edge and active twist concepts applied to helicopter blades. To obtain the desired wing weight and pitching moment the size and placement of piezocomposite patches were optimized. The piezo-composite actuator developed possessed a relatively high strain and force output. Recently, Bilgen [76] presented the ongoing research and development of a solid-state piezo-composite rotor design for use in rotary systems. The concept was applied to an quad-copter and ﬂight tests were made. Three piezoelectric macro ﬁber composites actuators per rotor blade change its shape and camber. Despite a reduction in rotor speed due to electromagnetic drag the thrust produced still increases, [75, 76].

Shape Memory Alloy actuation

Shape Memory Alloys (SMA) are materials whose mechanical properties, namely Young modulus, change with temperature. These alloys can be deformed at low temperatures and will return to its original shape when heated. The strains achievable are around 4-5%. As SMA usually are electric conductors, a simple method to change their temperature is through the Joule effect. According to Barbarino et al. [6] by directly ”implementing these materials within the structural elements, an actuation capability is integrated within the structure, with significant benefits in terms of: (1) weight, (2) reliability and maintenance, and (3) structural and aerodynamic efficiencies.” Due to heating and cooling cycles, SMA actuators can only be applied to low frequency shape changing of the order of magnitude of 1 Hz, [1,6]. In 1996 Roglin and Hanagud [77] presented an adaptive rotor blade that used embedded SMA wires to actuate a trailing edge flap to morph the airfoil camber for a remotely piloted helicopter. This camber morphing system was then utilized to supersede the collective pitch control of the helicopter. The change in rotor thrust due to the SMA actuation was linearized by a controller which applies the heating required

27 to deform the SMA. Due to the low-frequency response of the actuation system, only steady collective control was attempted, i.e., all rotor blades receive the same input and maintain shape through out the full revolution, [77]. As mentioned earlier, the phase 1 of the DARPA/AFRL/NASA Smart Wing program studied static wing camber morphing concepts that used low bandwidth actuation based on shape memory alloys. A smart wing 16% scale wind tunnel model, representative of an advanced military aircraft wing, was built and tested at the NASA Langley Research Center. The smart wing model smooth hingeless flap and aileron designs were actuated using built-in SMA tendons and deflections of up to 10◦ were obtained. ”Under steady-state conditions, performance improvements of 8-12% in comparison to a conventional design incorporating hinged control surfaces, over a broad range of wind tunnel and model test conditions, were established”, [78]. Other SMA actuated camber morphing concept was proposed in [79]. The bio-inspired wing sections were made of an antagonistic shape morphing structure composed of a cellular flexible core between two SMA face sheets in sandwich configuration, fig. 2.20. The core had a vertebrate structure allowing for the modular elements to rotate relative to one another. Hence, the actuator system could return its natural without any mechanism to provide the elastic restoring force, [79].

Figure 2.20: Shape morphing aero control surface’s strucure: airfoil at minimum camber (up); airfoil at maximum camber (bottom), [79]

Campanile [80] described a method for airfoil camber variation of a compliant structure without move- able parts. The rigid rib in an airfoil is to be replaced with a compliant belt-rib with stiff in-plane spokes attached by solid-state hinges. The spokes are used to provide the airfoil shape and shape change. ”Assuming the spokes are tension-compression only elements, the shape behavior of the belt-rib is determined by the belt’s bending flexibility and the spokes configuration,” [1]. A prototype can be view in fig. 2.21. Static bench tests were performed using mechanical wires and SMA actuation, [80]. In 2009 the feasibility of an adaptive wing for a small UAV entirely actuated by shape memory alloy devices was evaluated in [81]. The wing design comprising a sandwich box sub-structure with laminated faces, flexible ribs and a flexible skin employed SMA torsion tubes for wing camber control, while levers powered by SMA wires are utilized for local shape control which allows for a small change in the airfoil

28 Figure 2.21: Morphing rib structure, [80]. thickness distribution. Finite element simulations assessed the capability of the wing to support aerodynamic loads, the actuation power requirements and their force and torque during ﬂight. To allow a complete shape recovery by SMA actuators, deformation was limited to 4%. The wing appears capable to smoothly deform, with small stresses enabling a variation in camber from 5◦ to 15.5◦ with SMA torsion tubes delivering up to 200Nm and drawing around 1.2kW of power, [81].

Electro Active Polymer actuation

Electro Active Polymers (EAP) are materials that exhibit change in shape or size when under electrical field stimulation. For some EAP the maximum strains achievable are up to 380%, [1]. Aa aero-structural optimization study of morphing airfoils for adptive wings was performed in [82]. Instead of a sequential approach to optimization, the mission profile was represented as a sequence of spanwise lift distributions and used them to set the optimization goals. ”This allows to specify goals based directly on aerodynamic performances instead of prescribing fixed geometrical shapes.” The airfoil outer shape and its mechanical properties were parameterized to form combined aero-structural optimization problems. This method was applied to a morphing concept using dielectric elastomers, a specific type of EAP, achieving favorable aerodynamic and structural morphing performances, [82]. A smart wing design for a micro air vehicle with coupled EAP and piezoelectric actuation is presented in [83]. The proposed smart wing structure consisted of a composite spar and ailerons that had bimorph active ribs consisting of piezoelectric macro fiber composites actuators. Actuation is enhanced by the use of an EAP skin to preload the piezoceramic fiber actuators with a compressive axial load. Analytical and FE models demonstrated deflections up to 30◦ in aerodynamic. The computational models were validated with experimental work, [83].

2.2.6 Span

The three main parameters that affect the wing planform are span, chord and sweep. Both chord and span have a directly proportional effect on wing planform area. As span increases so does the wing area. But opposite to the chord’s effect, span increase also raises the aspect ratio of the wing, a parameter that alters the lift to drag ratio. A higher aspect ratio will result in increased wing efﬁciency and both increased range and endurance. Aircraft’s inertial properties also change with a larger span.

29 At low speed induced drag is the dominant drag component; however, at high flight speeds the lift coefficient decreases and the friction drag component surpasses the induced drag component. Thus, aircraft with large wing span and high aspect ratio will generally have good range and fuel efficiency, but lack maneuverability and have relatively low cruise speeds. On the other hand low aspect ratio aircraft are faster and highly maneuverable, but lose out on aerodynamic efficiency, [1,6, 22, 46, 47]. A variable-span wing can potentially integrate into a single aircraft the advantages of both a high span configuration and a short span one. This morphing technology has aroused interest specially for military UAVs, which must loiter for extended periods of time during surveillance and be able to switch into high- speed dash mode to move reconnaissance area or to attack a target. An increase in span accompanied by the respective increase in aspect ratio and wing area results in a decrease in the lift distribution for the same flight condition, i.e., for the same overall lift. Nevertheless, bending moment at the wing root can considerably increase leading to one major drawback. The need of a heavier wing structure to support it. ”Thus, both the aerodynamic and the aeroelastic characteristics should be investigated in the design of variable-span morphing wings”, [1,6]. Although, sweep morphing can alter the overall span of the wing, the span morphing concepts reviewed in this topic will necessarily include actual variation of the length of the wing measured from root to tip. Many of span morphing concepts are based on a telescopic mechanism, following the ideas first pioneered by Ivan Makhonine, figure 2.2, where the wing outer panel telescoped inside the inner panel to enable span and wing area changes. The German FS-29 glider built in 1972 was designed to research the telescopic variable span wing. The wing extension and retraction was humanly powered and the aspect ration of the wing changed between 20 and 28, [1,6,9]. Neal et al. [84] described the design and construction of a fully adaptive aircraft configuration capable of large scale changes in sweep, span and twist with the purpose to investigate morphing for multi- mission UAV. The span mechanism was actuated by pneumatic actuators, achieved a wing area increase of up to 31% and variation of the span up to 113%. There are five different planform configurations for the prototype as well as independent twist control for each wing. Wind tunnel tests shown that only three planform configurations were needed to maintain minimum drag over a large range of flight conditions, [84]. Static aerodynamic and aeroelastic studies on a variable span wing of a long range cruise missile were performed and reported in [85]. At full extension, the wing reached a span of 1.5m which represented a 50% increase in span, while wing area increased about 38%. Compared to the conventional wing configuration, i.e., the base wing with 0% span extension, the fully extended morphing wing configuration achieved an approximately 30% increase in range with an approximately 25% decrease in drag. Roll control comparison analysis between differential span actuation and conventional aileron based roll control showed improved roll control authority for the morphing wing solution. Aeroelastic behavior be- comes worse as the wingspan increased. Static aeroelastic considerations showed that a variable-span wing requires increased bending stiffness, [86]. In 2007 Samuel and Pines [87] presented the design and testing of a segmented telescopic wing for an UAV. Hollow fiberglass shells were used to preserve spanwise airfoil geometry of the wing segments

30 and are connected to ribs coupled to the telescopic spar mechanism, fig. 2.22. To prevent wing twist and fluttering, two identical spars were utilized. Each comprised of 3 segments of aluminium tubes of decreasing diameter and increasing length, connected by ceramic linear bearings that are extended and retracted using input pressure. The wing was able to undergo a change of 230% change in aspect ratio. Seam heights were reduced to a minimum and seam covers were applied so as to minimize the parasitic drag resulting from those discontinuities. In a previous work a more considerable seam height discrepancy of a similar concept caused a increase in parasitic drag that showed an increase of 25% in total drag relatively to a corresponding fixed wing. Spanwise friction tapes were also applied to the wing to facilitate the smooth deployment of the telescopic wing. In wing tunnel tests, the telescopic wing achieved lift-to-drag ratios as high as 16, which was similar to its solid fixed wing counterpart, [1, 87].

(a) (b)

Figure 2.22: The three segment telescopic wing from [87]: (a) Actuators; (b) Wind tunnel test setup

Scissor like mechanisms are an alternative approach to change the shape of the wingbox. Com- prised of planar unit cells, they are adequate when the desired morphing is two-dimensional in nature. Depending on the application diamond or hexagonal shaped cells might be used. The overall wingbox shape change depends on the cell shape along with the cell arrangement. The work presented in [88] explored a scissor-mechanism to alter the aircraft span and sweep with a design based on the TSCh wing (NextGen Aeronautics Inc.). A scale prototype of the internal wing structure was fabicated. When morphed it reduced the overall span by 55% and increased the sweep angle in 44◦ . A reeling mechanism joined with a cable system is presented for actuation, [88].

A design for a wing section capable of independent span and chord morphing by the use of extendable ribs and spars was developed in [89]. A model was fabricated with a threaded shaft attached to the spar root mechanism with ribs sliding along the span that was able to provide an increase of 50% in span and of 50% in chord with the possibility to control taper of the wing. A scissors-like mechanism was employed with the aim to keep the ribs equally spaced along the wing span as it changed. The built prototype version was not functional as the system blocked due to friction. The wing skin was made of a composite elastomeric material able to satisfy the shape change. Results were compared between three different conﬁgurations: the morphing wing with a perfect skin capable of maintaining the optimum

31 shape, the morphing wing with the actual skin model which deformed and the corresponding fixed wing. ”The ideal skin model outperformed the fixed wing in all speed with exception of the speeds around the fixed wing’s optimum design speed. The actual skin model showed increased drag compared to the fixed wing at all speeds due to the skin deformation,” [1, 89, 90]. The work presented in [91] comprises the development and ground validation a variable span wing (VSW) equipped on a small UAV prototype. An in-house aerodynamic optimization code was used for drag minimization of the telescopic wing for operation in the speed range of 11m/s to 40m/s. At 30m/s the drag reduction compared with the fixed original wing (with the same span as the maximum span of the morphing wing) was about 10%, near the upper flight speeds, a 28% drag reduction was achieved. A full scale prototype was built in composite materials and is made of two segments. The inboard segment is fixed to the fuselage and uses a monocoque structure construction, a composite material in the shape of the airfoil profile reinforced with a carbon fiber spar. The outboard segment uses a typical wing structure and slides inside the inboard segment to change the span of the wing, fig. 2.23. A rack and pinion system actuated by an electromotor realizes the span change. Different bench tests performed under loading revealed that the wing was suitable to be installed on a small UAV, [91]. In a subsequent work, the concept was flight tested. ”UAV fitted with the variable span wing demonstrated full flight capability and confirmed numerical estimates of performance and showed improvements produced by the VSW over a conventional fixed wing”, [91, 92].

Figure 2.23: General CAD view of the variable span wing from [92]

Another study using complaint cellular structures for span morphing was performed, [93]. A 2D compliant cellular truss structure designed to change span and thus planform area and aspect ratio of the wing. This in flight variation enables the shifting of the aircraft drag and power curves with velocity in diverse flight conditions, improving the efficiency of the aircraft. The study included a scale analysis of the application of the concept to different gross weight aircraft. 85%, 74% and 68% are the span decrease capability values for a 5kg, a 50kg and 500kg aircraft, respectively; while 2.9%, 7.4% and 8.9% are the structural weight fractions of the wing relative to a 5kg, 50kg and 500kg aircraft, respectively, not accounting for actuation weight. It was concluded that while the structural capability to span morph decreased with gross weight, the aerodynamic benefits gained by morphing increased with gross weight. Thus, suggesting that there is a gross weight for which this morphing design is most advantageous, [93].

32 Figure 2.24: Morphing core fabricated section, [94]

Vocke et al. [94] developed a continuous span morphing wing with focus to reduced discrete moving parts or abrupt changes in the airfoil profile. The concept contains two main components: a zero- Poisson ratio morphing core and fiber-reinforced elastomeric matrix composite skin with a nearly zero- Poisson ratio in-plane, fig. 2.24. Fabrication techniques for the morphing core and complex elastomeric composite skin were investigated. A prototype wing suitable for UAV integration was built with a 61cm default span and capable of a 100% increase in both span and area. The wing was tested under dynamic pressures corresponding to the maximum UAV speed of 130km/h and was able to maintain its shape, with the maximum out-of-plane deflection being inferior to 2.54mm. A recent study investigated a similar design and the morphing core was 3d printed and assembled from different materials providing different directional stiffness, [94, 95].

In 2012 do Vale [1] performed both computational and experimental studies on span and camber morphing wings. Coupled FE and CFD analysis was used to evaluate the benefits of camber and span morphing. Extended comparison analysis between the morphing wing aircraft and fixed wing of an UAV 100N of weight that was optimized for a cruise speed of 30m/s demonstrated that the morphing wing aircraft usually out performs the fixed wing aircraft, only showing losses in climb rate and a slight 4% drag penalty at 30m/s. Performance gains include fuel savings of 46% and 19% minimum and maximum fixed wing speed, respectively. Three morphing wing prototypes were fabricated and tested: two telescopic wing prototypes and one conjoined telescopic and camber morphing wing prototype. The concepts follow a modular design comprising in a hollow inboard wing from which a smaller outboard wing slides out supported by plates on the leading and trailing edge. The wing parts are produced out of balsa wood and carbon fiber shells. Span increase of around 65-70% with the area increase around 50% is achieved with the span morphing concept. Span change actuation comes from a wire and pulley system. The prototypes were submitted to ground tests performed with the intent to evaluate loading limits, actuation power, time and energy and mass penalties, [1].

The Adaptive Aspect Ratio wing morphing concept has been developed at Swansea University. The span morphing concept based on a compliant skin material and a internal structural shape changing mechanism was presented in [67]. Comprising four key elements: the elastomeric matrix composite

33 skin, the telescopic spar, the sliding ribs and the strap drive mechanism, ﬁg. 2.25. The compliant skin is the most demanding element of the concept, a ﬁne balance between in-plane actuation force and out-of- plane stiffness under aerodynamic loads is required. An initial skin design optimization was performed based on analytical models of the eleastomeric material, [96].

(a) (b)

Figure 2.25: Adaptive Aspect Ratio concept: (a) Isometric view (retracted); (b) Top view (extended), [67, 96]

Recently, Beaverstock et al. [97] studied the effect of span morphing coupled with camber morphing on the mission performance of a 25kg UAV. The morphing wing was modeled after the the adaptive aspect ratio (AdAR) span morphing concept and the fish bone active camber (FishBAC) morphing concept. Calculations showed the span morphing wing had a 25% improvement in the aerodynamic efficiency over its fixed wing counterpart, for an allowable 50% retraction with a velocity range of 50–115 kph. Reducing the allowed span morphing range reduced the aerodynamic benefits. Improvements in the efficiency achieved through camber morphing were found to be more situational dependent and more sensitive to the velocity range in the mission, where span morphing appeared more robust for an increase in velocity range beyond the optimum. However, the camber change required for optimum performance is considerably smaller than the span change to the planform. ”Span morphing, at the optimal mission velocity range, with 25% allowable retraction, can allow up to a 12% increase in mass before no performance advantage is observed compared to a similar fixed wing aircraft, where the camber morphing only allows up to 3%”, [97].

34 Chapter 3

Telescopic Wing Design Strategy

The aim of this work is to design a telescopic wing for a surveillance UAV and evaluate its feasibility and performance. In this chapter, the design methodology of the telescopic aircraft wing is described. The design process is done in three major steps: First, for a starting point the initial sizing of an aircraft platform and a typical mission are deﬁned. Then, concepts for the morphing mechanism are developed, selected and described. Finally, a structural analysis, based on analytical formulas and aerodynamic calculations using the lifting line theory, is approached.

3.1 Design Framework

A design process was established, and the approach adopted in the current span morphing wing design is illustrated in figure 3.1. An unmanned aerial vehicle is an aircraft that does not require an on-board human crew in order to fly. UAVs may have the ability of fully autonomous flight or they can be remotely controlled by a human pilot. This makes them the preferred solutions for missions involving a high level of risk. Along side that, the pilot endurance is not a limiting factor in the mission duration. The size of UAVs can range from micro air vehicles, with size comparable to small birds, up to large surveillance ones comparable to commercial airliners. Long endurance surveillance UAVs aerodynamic design focuses on maximizing range and endurance. Designed with high operational resistance with low-weight high-strength materials, structurally able to sustain impact on landings. Besides the lower costs and risks, the complex and dynamic flight mission profiles of UAVs is the major factor that makes them particularly well suited to be morphing wing technologies test platforms. On the other hand, the typical commercial airliner can spend more than 90% of its flight mission in cruise. Therefore, their fixed wings can be optimized for this flight condition and the plane is able to achieve high lift to drag ratios while cruising. Even if in the rest of the flight mission profile the wings are to some extent inefficient, the overall mission efficiency will not suffer a great penalty. Most UAVs have mission profiles that require them to cycle between loitering, cruising, dashing, fast ascents and fast descents, [98].

35 Design Speciﬁcations

Mission Proﬁle

Concept Development

Design Matrix and Selection

Concept Description

Loading Calculation

Structural Analysis

Performance Analysis

Figure 3.1: Detailed design process

To assess the span morphing wing technology its performance is compared to the baseline performance of a fixed wing of similar size. This analysis is presented in chapter5. To that avail, the initial sizing of an aircraft platform capable of equipping both a fixed wing and the span morphing wing is performed. It is expected that the span variable wing will expand the aircraft flight envelope and increase wing efficiency in both loiter and cruise at variable speeds and variable aircraft weight, when the fuel carried is consumed. Furthermore, an enhancement in dash, climb and take-off and landing performance is anticipated. In this study, both structural and aerodynamics analysis were performed. XFLR5 software is used to calculate the aerodynamic forces at different flight phases and estimate the aerodynamic performance. A analytical model was used to calculate the structural stress under aerodynamic loading.

3.1.1 UAV Characteristics

As a design starting point, the initial sizing of a medium endurance and medium altitude UAV platform is determined. The fixed wing UAV model AR5 from TEKEVER is used as reference both in terms of dimensions and performance. Table 3.1 shows the main characteristics of this aircraft, further specifications can be found in appendix A.1. As the information gathered about this aircraft is limited further market research of similar size and performance aircraft was realized to find adequate wing loading ratio.

The aircraft main speciﬁcations are displayed on table 3.2. The key characteristic of the aircraft

36 Table 3.1: TEKEVER AR5 speciﬁcations

Takeoff mass [kg] 150 Length [m] 3 Span [m] 4.3 Tail conﬁguration H-tail Endurance [h] 12 is the variable span. The wing is set up to vary almost continuously its span between 5m and 3.5m, respectively the maximum and minimum span conﬁgurations.

Table 3.2: Aircraft speciﬁcations

Takeoff mass [kg] 150 Length [m] 3.5 Span [m] 3.5-5 Chord [m] 0.75 Wing area [m2] 3.75 Aspect ratio 4.7-6.7 Tail conﬁguration Inverted V-tail Tail span [m] 1.715 Tail chord [m] 0.429

Some considerations were made about the desired performance of the UAV and are detailed in table 3.3. As the AR5 UAV, thrust is assumed to be provided by a propeller powered by a two stroke engine, expected to generate 20kW of power. The aircraft is capable of taking-off and landing from small paved runways.

Table 3.3: Aircraft desired performance

Cruise speed [kts] 95 Cruise Altitude [m] 3,000 Loiter Speed [kts] 70 Dive Speed [kts] 150 Stall Speed [kts] 50 Maximum climb rate [m/s] 3 Lift-to-drag ratio 20 Take-off roll distance [m] 150

3.1.2 Mission Proﬁle

A representing mission profile was defined based on the specifications and performance of the aircraft. To better equate to a typical mission of an UAV designed for surveillance and patrol activities a large loiter flight phase was included. The mission includes the common flight phases in most aircraft. Such as take-off, climb, cruise and

37 Figure 3.2: Mission profile descent. However, the major flight phase is an 8 hour loiter and a high speed dash to test the fast deployment of the platform is also notable. The dash and cruise phases provide the aircraft its 150km of range. A schematic of the mission profile can be seen, figure 3.2.

3.1.3 Proﬁle Selection

Figure 3.3: Seilig S4233 low Reynolds airfoil

The airfoil profile is one of the most important aspects of the wing, and has direct impact on aerodynamic performance. As the aircraft it self, the ideal airfoil performs exceptionally well in the dominant flight phases, loiter and cruise, and does not limit aircraft performance. For a wing to achieve high lift to drag ratios, an airfoil that will produce the necessary lift and still have a minimal drag penalty is needed. Low drag and high lift to drag coefficient in the range of lift coefficients needed at cruise speeds to maintain level flight are the defining characteristics of the desired airfoil. Airfoil types that can reach very low drag coefficient usually have what is called a laminar drag bucket. It is important that this region encompasses the lift coefficients the wing sections will be subject to. Several low Reynolds (Re) airfoils were considered and some relevant data is displayed in table 3.4. For structural reasons: to ensure enough internal space for structural elements, airfoils with maximum thickness (t/cmax) inferior to 10% of the chord were not considered, [46]. The software XFLR5 was used to evaluate airfoil performance. This software provides several tools to analyze and study both airfoils and wings. The airfoil analysis tool set is a version of the Xfoil software originally developed by Mark Drela compiled with C++ and C programming languages. Xfoil accounts for forced or free laminar-turbulent transition, transitional separation bubble(s) and limited trailling edge separation, [99, 100]. The Reynolds number (Re) is a dimensionless parameter that is commonly used to characterize the

38 flow of a fluid. The Reynolds number is dependent on fluid properties and a characteristic length, that in the case of an airfoil equates to the length of the chord, and ”measures the ratio of inertial forces to viscous forces and describes the degree of laminar or turbulent flow”, [101]. Setups with the same Reynolds number will have the same flow characteristics even if the fluid, speed and characteristic length are different. Thus it is very useful for scaling and comparing wind tunnel tests of smaller models. The equation 3.1 calculates the Reynolds number.

ρV c V c Re = = (3.1) µ υ

Where c is the chord length of the airfoil and V , ρ, µ and υ are, respectively, the velocity, density, dynamic viscosity and kinematic viscosity of the ﬂuid.

Table 3.4: Low Reynolds airfoils characteristic at Re=1,000,000

Airfoil t/cmax [%] Cd min Cl max Cl/Cd max NACA 2412 12 0.00547 1.582 101.38 NACA 63-412 12 0.00559 1.4179 115.69 S4233 13.6 0.00605 1.3896 136.02 SD7062 14 0.00684 1.7442 121.55 FX 63-137 13.7 0.0073 1.8523 132.52

Airfoil geometry data was obtained from the databases available at the Airfoil Tools website, [102].

Several performance evaluation parameters were determined and the software XFLR5 was used to calculate airfoil lift and drag polars at 1,000,000 and 500,000 Re. These parameters include the range of operational lift coefficient and angle of attack, as well as the drag coefficient over the same range. Also maximum lift coefficient and maximum lift coefficient to drag coefficient ratio. Another constraint is the maximum thickness of the airfoil; for structural reasons the maximum thickness of a small chord airfoil cannot be under 10% of the chord.

Each parameter was evaluated in a range between the best performing and worst performing of a set of low Reynolds airfoils. A evaluation matrix with the grades normalized into a 0 to 5 scale is presented in table 3.5. The Seiling S4233 low Reynolds airfoil, ﬁgure 3.3, out performed the considered airfoils and for that reason it was selected.

Table 3.5: Airfoil evaluation matrix

Airfoil t/cmax Cd min Cd avg Cl range AoArange Cd cruise Cl/Cd max Cl max Total NACA 2412 2.5 4.5 3.5 2 2.5 4 2.5 2.5 24.9 NACA 63-412 2.5 4 4 1.5 2 4 3.5 1.5 24.6 S4233 4 3 3 3.5 3 3 4.5 2 26.5 SD7062 4 2.5 2 4.5 4 2 4 4 25.8 FX 63-137 4 2 1 4.5 4.5 0.5 5 5 24.1

39 3.2 Telescopic Wing

The objective of the span morphing concept is to achieve an efficient wing that can present the desired span to better suit the flight condition and to expand the aircraft flight envelope. The telescopic wing is intended to provide a maximum wing span of 5m to the aircraft and a minimum span of 3.5m. The maximum wingspan configuration has a slightly larger wingspan than the 4.3m of the AR5 UAV to compensate for the lower efficiency of a fully rectangular wing without winglets. The morphing wing nature also allows for this configuration to be completely dedicated to low speed flight whilst not compromising the aircraft higher speed performance. The minimum (3.5m) span configuration reduces the span by 30% in line with several span morphing concepts, [1, 91, 92, 96, 97], also has similar wing loading value to UAV of similar size and cruise speeds closer to the desired dive speed. The chord length is to be constant with 0.75m and the profile in the shape of the Seilig S4233 airfoil. That results in a semi span length varying between 1.75m and 2.5m. Although the comparison between the maximum and minimum wing span configuration shows the more drastic differences in performance and flight characteristics there are further benefits to be explored from the intermediate wing configurations. Therefore, the ability to provide any wingspan length between the maximum and minimum configurations is desired. Due to the high complexity and inherent necessity to be deformed, the morphing wing is more likely to deform and deviate from the desired aerodynamic shape than a non-morphing wing; thus losing aerodynamic performance and efficiency, [7, 87, 96]. The actuation rate can also be a design limitation especially when contemplating asymmetrical span extension as a means for roll control. In that case a low deployment and retraction speed can drastically reduce the banking authority of the aircraft. Some wing concepts were considered: Concept A is based on the AdAR design developed in works [96, 97]. The design span morphing capability is carried out by a telescopic spar. While the structure of the inboard part of the wing is fixed and of conventional characteristics, the outboard part of the wing includes sliding ribs and a compliant morphing skin made from a elastomeric matrix composite, see figures 3.4 and 3.5. Strained elastic materials experience the Poisson effect; Thus carbon fiber reinforcements are applied to the skin chord wise to reduce the Poisson ratio of the material and the deviations from the desired aerodynamic shape the wing might suffer. Additionally, a zero Poisson mesh made from polymers can be used to provide airfoil shape support to the skin at different wingspans. Span extension is actuated by a strap drive system powered with and electrical engine. Due to the compression forces introduced by the elastomeric skin the span retraction dispenses an actuation force oriented with the movement of the outboard spar. The mechanism responsible for varying the span of concept B is similar to the previous concept; However, instead of using a compliant skin structure, a telescopic rigid skin is adopted, figures 3.6 and 3.7. The structural loading are carried by a single telescoping spar and the shape of the airfoil surface is maintained with sliding ribs throughout the inboard part of the wing. The contact between the sliding ribs and the skin is intermediated with a low friction coefficient material to facilitate smooth movement.

40 Figure 3.4: Plant view of span morphing concept A (extended)

Figure 3.5: Isometric view of span morphing concept A (retracted)

The outboard part of the wing that is connected to the moving beam of the main spar has conventional ﬁxed structure and its skin slides from inside the inboard telescoping skin.

Figure 3.6: Plant view of span morphing concept B (extended)

Figure 3.7: Isometric view of span morphing concept B (retracted)

41 Design concept C is based on a telescopic wing design developed in works [1, 91, 92]. The telescopic wing consists of a hollow outer wing (inboard) from which the inner wing (outboard) slides out, figures 3.8 and 3.9. The hollow inboard part of the wing is made from a sandwich made from composite materials that is reinforced in the spanwise direction with carbon fiber beams. This structure has to provide both the airfoil surface and the structural stiffness to support the aerodynamic forces applied to the wing. The inner moving wing is structurally conventional and the actuation options to provide its movement and determine its position are varied. The outboard wing moves as a whole. The chord discrepancy between the fixed wing and the moving wing can be considerable and this design has the more significant difficulty mitigating that as the minimum thickness achievable of the skin of the fixed wing is limited by the structural necessities of that element.

Figure 3.8: Plant view of span morphing concept C (extended)

Figure 3.9: Isometric view of span morphing concept C (retracted)

Table 3.6 presents an assessment of the design concepts considered based on several evaluated parameters. It is worth noting that the table was constructed based on knowledge of these concepts and their advantages and drawbacks developed on previous studies and implementation efforts, that were reviewed on chapter2. A evaluation scheme ranging from 0 to 5, 5 being the best rating, was used.

Table 3.6: Concept evaluation matrix

Concept Aerodynamics Weight Manufacture Actuation System Actuation Energy Total A (AdAR) 5 4 3 5 2.5 19.5 B 4 4.5 5 3 5 21.5 C 2.5 5 3 4 5 19.5

42 In aerospace applications there is an innate necessity to ensure reduced weight. A morphing wing technology can be compared in terms the weight penalty the aircraft would need to carry to nullify the aerodynamic benefits over a mission profile when compared to a fixed wing of similar size. In terms of aerodynamic efficiency of the wing it is important to note that spanwise discontinuities on the wing surface increase drag and therefore reduce the lift to drag ratio. As demonstrated by Samuel and Pines, residual discontinuities have manageable effects. However, larger discontinuities may increase wing drag considerably in comparison with a fixed wing of similar shape. Also deformations in the airfoil skin can affect the lift and drag produced by the wing, [1, 87]. The fabrication costs of the design, effectiveness of its actuation system and the actuation energy drained comprise the remaining parameters. Concept B emerges as the best overall option.

3.2.1 Main Spar

The telescoping spar design utilizes only two beams, one ﬁxed that starts from the root of the wing and the second one ends at the tip of the wing. To achieve a large relative span extension the two beams are designed with the same length.

Figure 3.10: Plant view of the telescopic spar (retracted)

Figure 3.11: Isometric view of the telescopic spar (retracted)

Taking into consideration other telescopic wing mechanisms of similar size and similar number of telescoping elements, a 10% overlap between the beams at the most extended conﬁguration was de- ﬁned. Resulting in two 1.375m beams, 1.250 from half of the semi span length and 0.125m extra to result in the 10% overlap. The computed shear and bending moments presented in chapter4 revealed

43 that the 10% overlap margin was enough margin to keep the section from being a critical section in terms of internal beam stress. Multiple spar wing structure designs can deal with wing torsion without each singular beam having to support a large amount of torsional loading. Conventional wing structures also use the skin and ribs to carry torsion originated from shear stresses. Being that the design only contains one spar and that the skin is disconnected from certain ribs, it is necessary to design a spar capable of supporting not only the bending but also the torsion generated by the aerodynamic loading, [7, 46]. From an aircraft design perspective, structural weight is something desired to be as low as possible; Thus, the beam cross section shape chosen must be apt to sustain torsional and bending loads while containing the minimum amount of material. While the hollow circular cross section is the most material efficient under torsion, similar shaped closed loop hollow cross sections perform adequately under torsion. The hollow rectangular cross section shape was chosen allows for a simple overall setup, fabrication and actuation, while also being a closely related shape to the wing profile. Following the design objectives expressed above, a aerospace grade aluminium alloy was chosen for its high weight to strength ratio. The Aluminium 2024 T3 alloy is mainly composed by aluminium and copper; Its components and properties can be consulted in [103]; of interest, the maximum yield strength of 345MPa and also the shear strength of 283MPa and the Young modulus of 73.1GPa. Pre-fabricated standards of hollow rectangular beams made from this alloy are available in the market. Another aspect of the main telescopic spar design is the movement between the two beams. To prevent jamming and allow for a smooth and manageable actuation, it is essential to keep low levels of friction between the beams. The bearing surfaces along which the sliding motion of the spar occurs are located at the end of each beam, i.e., one at the outboard end of the fixed beam and one at the inboard end of the moving beam. Sets of discrete bearing surfaces made from a low friction polymer placed on the exterior surface of the inner beam and the interior surface of the outer beam to facilitate the movement. PTFE (Polytetrafluorethylene) and UHMWPE (Ultra-High-Molecular-Weight-Polyethylene) polymer sheets have a low friction coefficient. Comparing the two materials UHMWPE has lower density and friction coefficient in a contact surface bearing surface-alumina. PTFE has higher compression strength; However, compression pressure’s order of magnitude was calculated to be inferior to the order of magnitude of the compression strength, based on the static structural results presented on chapter4. Thus, UHMWPE will provide the superior bearing surfaces, [96].

3.2.2 Ribs

The wing rib is a plate like structure situated in the plane of the two-dimensional airfoil profile which is critical to the structure and structural integrity of the wing. The rib provides the airfoil shape to the surface of the wing and its main function is to transmit the aerodynamic loads from the skin to the internal load bearing structures. With 6 sliding rib in the inboard section of the wing and 8 fixed ribs on the outboard section in the maximum wingspan configuration the sliding ribs are at 281.3mm from each other, the maximum

44 distance; In the minimum wingspan conﬁguration the sliding rib spacing is 93.8mm. While the distance between the ﬁxed ribs is constant at 458.3mm.

Figure 3.12: Isometric view of the sliding rib with bounding box dimensions

The sliding ribs are simply in contact with both the the main spar and the airfoil skin. While the the other ribs are fixed to both and can transfer effectively moments and forces, the sliding ribs need an additional structure to provide the leverage to deal with out of plane forces that can be amplified with the slightest deviation between the rib plane and the plane perpendicular to the spanwise direction. In figure 3.12 can be seen a box like structure that involves the spar through a length considerably larger than the thickness of the rib is used to provide the alignment stability necessary. A similar material to the bearing surfaces used in the telescopic spar is to be adopted to facilitate the movement between the ribs and both the spar and the airfoil skin.

3.3 Structural Strength and Sizing

The telescopic wing is designed to be able to structurally sustain a limit load factor n of 3.8. The major load bearing element of the wing is the telescopic spar and it is assumed to sustain the full aerodynamic loading. The structural flight worthiness of the aircraft is performed according to design specification of a normal category aeroplane provided by the European Aviation Safety Agency (EASA), under the norm: Certification Specifications for Normal, Utility, Aerobatic, and Commuter Category Aeroplanes CS-23, [104]. ”Strength requirements are specified in terms of limit loads (the maximum loads to be expected in service) and ultimate loads (limit loads multiplied by prescribed factors of safety)”, [104]. The norm specifies that a normal category aeroplane should be design for a limit load factor of 3.8g. The general aviation factor of security of 1.5, also specified in the CS-23 norm, should be applied to the limit load factor; Thus, the ultimate load factor is set to 5.7g. This is the critical load that the structure has to sustain without failing. The CS-23 norm prescribes a negative limit load of -1.52g that is 0.4 times the positive load factor, and a following negative ultimate load factor, with the general factor of safety, of -2.28g.

45 In ﬁgure 3.13 it is presented the aircraft Velocity-load (V-n) diagrams of two wing conﬁgurations with 5m and 3.5m wingspan and 150kg of mass. The speeds of interest and stall line were calculated in chapter5. The air properties of the international standard atmosphere at sea level were considered.

The following table 3.7 shows the stall (VS) and cruising (VC) speeds for level ﬂight, along with the dive speed (VD) and the design maneuvering speed (VM). According to the CS-23 speciﬁcations, the dive speed corresponds to 1.5 times the cruising speed and the design maneuvering speed corresponds to the stall speed at the limit load factor.

(a) (b)

Figure 3.13: Velocity-load diagrams: (a) 5m wingspan conﬁguration (maximum) (b) 3.5m wingspan conﬁguration (minimum)

Table 3.7: Morphing aircraft optimal speeds

Wingspan [m] VS [kt] VC [kt] VM [kt] VD [kt] (a) 5 47.658 83.409 92.927 125.114 (b) 3.5 59.367 111.439 115.665 167.159

In both V-n diagrams the critical flight situations in terms of structure integrity for both wing configurations are identified by points A and B. Point A being the wing stall condition at a 5.7 load factor and point B is a 5.7 loading at dive speed. The major load bearing element of the wing is the telescopic spar and it is assumed to sustain the full aerodynamic loading. Wing structure weight is not considered. Due to the elevated number of degrees of movement between the skin, ribs and spar, the former has to sustain a larger portion of the loads than a more conventional internal structure where the wingbox can better sustain the bending and torsional loads. Tools were setup in a modular manner to allow to calculate the maximum stress showed by the beam from a specific aerodynamic loading.

46 3.3.1 Aerodynamic Loads

For the sake of identifying the critical flight phase, i.e., the flight phase in which the telescopic spar will show the greater values of internal stress it is required to identify the loading of the wing and its distribution; The software XFLR5 was used which provides traditional methods and several tools to analyse airfoils and wings. The non-linear lifting line theory method is used to assess the airflow characteristics around the wing and to estimate the distribution of the aerodynamic forces over the wing span, [99, 100, 105, 106]. The lifting line theory provides a prediction of the spanwise lift distribution over a finite wing and subsequent determination of the aerodynamic characteristics of the wing from two-dimensional airfoil data. The three-dimensional flow is approximated and several two-dimensional aerodynamic concepts are extrapolated to the tree-dimensional case. Analytical results present high conformity with experimental results for medium to high aspect ratio wings with little sweep and dihedral over small angles of attack, [99, 106]. The spanwise lift distributions are calculated for the possible critical flight conditions. Namely, the stall line condition at ultimate load factor of 5.7g and flight at dive speed at the same load factor. These points are identified in the aircraft V-n diagrams by points A and B, see figure 3.13.

3.3.2 Moment and Shear Diagrams

A script was developed using MATLAB R that computes the shear and moment diagrams of the telescopic spar for a given load distribution. It takes as inputs the beam dimensions and its positions, i.e., how extended is the telescoping mechanism and a semi-span load distribution data ﬁle that is gathered from XFLR5. Moment and Shear Diagrams were calculated considering the aerodynamic loads were totally carried by the main telescopic spar.

Figure 3.14: Telescopic spar loading schematic

The loading is assumed to be the distributed aerodynamic load and due to symmetry only the semi- span needs to be considered. A schematic of the setup can be observed in ﬁgure 3.14; The AB beam

47 is the inner ﬁxed beam-element of the spar, and the outer moving bean-element is characterized by the CD beam. Some considerations and simpliﬁcations were taken:

• Beams are considered rigid bodies;

• The interaction between the two elements of the telescopic spar was simpliﬁed: points A and C are considered the only contact points between the beams and the reaction forces are discrete and simply vertical.

The free-body schematic of each beam-element and its loading can be seen in ﬁgure 3.15.

Figure 3.15: Free body diagram of each telescopic spar element

The shear and bending moments need to be computed for both beam-elements; Through these diagrams the beam cross section subject to the most extreme loading can be identiﬁed; The important values are the maximum bending moment and the maximum shear force.

Figure 3.16: Free body diagram of differential length of a beam, taken from [107]

From the free body of differential length of a beam, represented in ﬁgure 3.16, the formulas for shear and bending moment can be deduced, 3.2 and 3.3 respectively.

dV y = −p (3.2) dx y

where Vy(x) is the shear distribution and py(x) is the load distribution.

dM z = −V (3.3) dx y

where Mz(x) is the moment distribution. The integration of these formulas, equations 3.2 and 3.3, will provide the same result as using the free body diagram method: considering the beam virtually split at section x; Then, Vy(x) and Mz(x) will

48 respectively be equal to the summation of forces and moments applied from section x to either end of the beam, so as to maintain conservation of linear and angular momentum.

3.3.3 Maximum Stress and Sizing

A further MATLAB R script was created capable of calculating the maximum normal and shear stresses in a section of a hollow rectangular beam. The script takes into consideration the geometry of the section, i.e., height, width and thickness of the hollow rectangular section. Shear and bending moment are additional inputs. The calculations are made following the equations 3.4 for normal stress and 3.5 for shear stress.

Mzy σxx = − (3.4) Izz

where y is the distance from the neutral surface, in the case of a symmetrical cross section the neutral surface coincides with the x axis, and where Izz is the centroidal moment of inertia of the section, [108].

VyQ τxy = (3.5) Izzb

where Q is the ﬁrst area moment, [108].

The hollow rectangular beam cross section is represented in ﬁgure 3.17; Its dimensions are identiﬁed.

(a) (b)

Figure 3.17: (a) Rectangular hollow beam cross section (b) Normal stress cross sectional distribution, [109]

The geometry parameters needed to compute the normal and shear stresses of a hollow rectangular beam cross section contingent to its dimensions are given by the following equations: 3.6-3.8.

wh3 − (w − 2t)(h − 2t)3 I = (3.6) zz 12

49 wt(h − t) t(2h − 4t)(h − 2t) Q = + (3.7) 2 4

b = 2t (3.8)

where w, h and t are respectively the width, height and thickness of the hollow rectangular cross section as displayed on figure 3.17. Normal stress increases with the distance from the neutral axis of the cross section. In the hollow rectangular beam case, maximizing the distance y will provide the maximum normal stress - the value of interest. Shear stress is inversely proportional to the thickness of the beam. So to calculate the maximum shear it should be taken the part of the beam cross section where its thickness is minimum; Resulting in the equation 3.8. Since there are benefits in producing the lightest possible aircraft, another set of scripts was developed to find the dimensions of the cross section with the least area still capable of supporting the load applied to the beam and present a maximum normal stress below the maximum yield strength of the material of the beam. Additionally to the former presented tools, a MATLAB R was produced to assess the internal space of the airfoil profile and find the maximum external dimensions of the beam cross section. Final sizing process:

• Calculate the shear and moment loading applied to the beam for a given condition ﬂight condition.

• Identify the maximum spar/beam dimensions to ﬁt inside the wing proﬁle.

• Compute the critical stress on the maximum loaded cross section.

• Find the smallest beam cross section dimensions capable of supporting the loading.

To check the deﬂection of the wing under the critical loading conditions, a MATLAB R script was developed to calculate the maximum deﬂection of each beam for the situation analysed, based on the area-moment method/theorem.

50 Chapter 4

Results

4.1 Structural Strength and Sizing Results

After identifying the maximum bending moment and shear force applied to the section of each beam, a final beam geometry capable of sustaining the loading and with the smallest area was found. The critical structural case, i.e., the flight condition that generated the largest cross sectional loading corresponds to the point A of V-n diagram of the 5 meter span wing, figure 3.13 (a). This corresponds to a stall condition of the maximum wingspan with a 5.7 load factor. The minimum area geometry found is presented below. In fig. 4.1 the cross sections of both the inboard beam and the outboard beam are represented inside the airfoil profile. Thickness is not represented. The full dimensions are presented in table 4.1. The dimensions are as represented in figure 3.17.

Figure 4.1: Final telescopic spar geometry schematic

Table 4.1: Final telescopic spar dimensions

Beam Width w [mm] Height h [mm] Thickness t [mm] Area [mm2] AB 75.6 158.0 1.03 476.9 CD 79.7 164.0 0.15 73.0

51 4.1.1 Aerodynamic Loading

The ﬁgure 4.2 and 4.3 show the lift distribution over the semi span of the wing for the maximum and minimum wingspan achieved by the morphing wing. The area below each curve corresponds to half the lift generated by the wing, that in this case will be 5.7 times the weight of the aircraft. The lift distributions shown in ﬁgures 4.2 and 4.3 correspond to the lift distributions present at stall condition (case A) and at dive speed (case B).

Figure 4.2: Semi span lift distribution for maximum and minimum wingspan conﬁgurations (case A: n= 5.7; stall condition)

Figure 4.3: Semi span lift distribution for maximum and minimum wingspan conﬁgurations (case B: n= 5.7; V= dive speed)

Although it is difﬁcult to note, the dive speed case has a slightly heavier lift distribution in the inboard zones compared to the stall case. This results in a smaller root bending moment as corroborated by the calculated reactions presented in tables 4.2 and 4.3.

52 4.1.2 Bending Moment and Shear Diagrams

The results of the free body diagram of the telescopic spar, presented in ﬁgure 3.15, are represented in tables 4.2 and 4.3. The reaction forces the beams and the support are dependent on the lift distribution and the extension of the telescopic spar.

Table 4.2: Reaction loads: Case A – Stall Speed

Wingspan [m] MA [kN.m] RA [kN] RB [kN] RC [kN] (a) 5 4.748 4.177 5.001 2.944 (b) 3.5 3.272 4.177 1.904 -1.224

Table 4.3: Reaction loads: Case B – Dive Speed

Wingspan [m] MA [kN.m] RA [kN] RB [kN] RC [kN] (a) 5 4.728 4.177 4.954 2.909 (b) 3.5 3.259 4.176 1.893 -1.228

In the next set of ﬁgures 4.4 to 4.11, the bending moment and shear force through the wing semi span are illustrated for each wing (maximum 5m span and minimum 3.5m span) and for both stall and dive speed at the ultimate load factor of n= 5.7.

Figure 4.4: Bending moment diagram of the maximum wingspan (5m) wing at stall condition and n= 5.7

4.1.3 Maximum Stress

Tables 4.4 and 4.5 show the maximum sectional loading of each beam for flight case A and B respectively. The maximum normal stress and shear stress present in the cross section of the telescopic spar beams are exhibited in tables 4.6 through 4.9 under the maximum loading for each of the four configurations considered. The non-critical configurations are compared with the maximum normal and shear stress of the critical case and the result is also given as a percentage of the maximum design value.

53 Figure 4.5: Shear force diagram of the maximum wingspan (5m) wing at stall condition and n= 5.7

Figure 4.6: Bending moment diagram of the minimum wingspan (3.5m) wing at stall condition and n= 5.7

Figure 4.7: Shear force diagram of the minimum wingspan (3.5m) wing at stall condition and n= 5.7

54 Figure 4.8: Bending moment diagram of the maximum wingspan (5m) wing at dive speed and n= 5.7

Figure 4.9: Shear force diagram of the maximum wingspan (5m) wing at dive speed and n= 5.7

Figure 4.10: Bending moment diagram of the minimum wingspan (3.5m) wing at dive speed and n= 5.7

55 Figure 4.11: Shear force diagram of the minimum wingspan (3.5m) wing at dive speed and n= 5.7

Table 4.4: Maximum section loading: Case A Stall Speed

Wingspan [m] Max MAB [kN.m] Max SAB [kN] Max MCD [kN.m] Max SCD [kN] (a) 5 4.748 5.001 0.782 3.367 (b) 3.5 3.272 4.177 0.278 1.311

Table 4.5: Maximum section loading: Case B Dive Speed

Wingspan [m] Max MAB [kN.m] Max SAB [kN] Max MCD [kN.m] Max SCD [kN] (a) 5 4.728 4.954 0.774 3.332 (b) 3.5 3.259 4.176 0.279 1.306

Table 4.6: Maximum normal and shear stress: Case A Stall Speed (5m span)

Beam Max Normal Stress σxx [MPa] Max Shear Stress τxy [MPa] AB 344.85 34.80 CD 344.79 150.84

Table 4.7: Maximum normal and shear stress: Case A Stall Speed (3.5m span)

Beam Max Normal Stress σxx Max Shear Stress τxy [MPa] ∆ [%] [MPa] ∆ [%] AB 237.62 68.9% 29.06 83.5% CD 122.69 35.6% 58.72 38.9%

Table 4.8: Maximum normal and shear stress: Case B Dive Speed (5m span)

Beam Max Normal Stress σxx Max Shear Stress τxy [MPa] ∆ [%] [MPa] ∆ [%] AB 343.40 99.6% 34.47 99.1% CD 341.00 98.9% 149.29 99.0%

56 Table 4.9: Maximum normal and shear stress: Case B Dive Speed (3.5m span)

Beam Max Normal Stress σxx Max Shear Stress τxy [MPa] ∆ [%] [MPa] ∆ [%] AB 236.71 68.6% 29.06 83.5% CD 122.91 35.6% 58.52 38.8%

Table 4.10 presents the maximum deflection results of each beam of the telescopic spar, as well as a conservative estimate of the maximum wing deflection. Additionally, it is presented the angle defined by the three beam/wing points: tip, deflected tip and root, the latter being the vertex.

Table 4.10: Maximum deﬂection: Case A Stall Speed (5m span)

Beam Tip deﬂection [mm] angle [◦ ] AB 13.5 0.6◦ CD 2.6 0.1◦ AD (both) 53.4 1.2◦

4.1.4 Remarks

Analyzing the previous results, it is possible to infer the following observations:

• The smallest distance between the spar and the airfoil proﬁle is 4.95mm. This corresponds to the distance from the forward upper edge of the outboard beam (CD) to the proﬁle. The minimum distance measured from the inboard beam is 7.43mm.

• The space between the beam elements of the telescopic spar where they overlap is different for the vertical sides to the horizontal sides of the beams. Between the horizontal wall it measures 3.88mm and between the vertical walls it measures 5.70mm.

• Using the aluminium alloy 2420 T3 the beam AB has a mass of 1822.9g and the beam CD a mass equal to 279.2g.

• The distinction between the lift distributions of the maximum and minimum wingspan conﬁgurations is quite clear. As of the result of both generating the same overall amount of lift (the area below each line is equal to half the lift produced by the wing), the 5m wing has a lower intensity lift distribution over a larger span. Oppositely, the 3.5m wing’s lift distribution has higher values over a smaller span.

• Reaction force A is constant throughout all configurations, as it is dependent on the total lift produced by the wing which is defined by the constant load factor of 5.7. However, the moment reaction at the root changes with all configurations as it is dependent of how the lift is distributed along the semi span.

57 • The reaction force C has a positive value for the maximum wingspan conﬁguration but a negative value for the minimum wingspan conﬁguration. This change indicates that the resulting force of the lift distribution over the beam CD is situated to the outboard of point B (positive) or to the inboard of point B (negative).

• From the bending moment diagrams we can see that its maximum value is located at the root of the wing, which is the expected behavior of a cantilever beam.

• The maximum bending moment the CD beam experiences is located at point B for the 5m wingspan conﬁguration.

• The stall situation for the 5m wingspan conﬁguration clearly stands out as the critical case showing the largest values of bending moment and shear force for both beam AB and CD.

• The difference between the loadings is much more considerable when comparing the 5m wingspan with the 3.5m wingspan conﬁguration than when comparing the stall ﬂight condition with the dive speed condition for the same wingspan. The same effect can be observed when comparing the maximum stresses present in the beams.

• Beam CD has considerably larger maximum shear stress than beam AB due to a much lower thickness.

58 Chapter 5

Aerodynamic Performance Studies

5.1 Performance Assessment

In this chapter, the aerodynamic performance of the morphing wing is studied in varied aspects using diverse analysis in XFLR5. The morphing wing is compared to two fixed wings, one with the same span as the maximum span configuration of the morphing wing (5m) and one with the same span as the TEKEVER AR5 (4.3m). They are compared in terms of drag and the weight penalties necessary for nullifying the morphing wing benefits are provided. First, the lift and drag polars of the airfoil profile were computed for a large range of Reynolds number so as to provide the ability to assess the wing over a likewise large range of flight speeds. The Reynolds numbers analyzed ranged from 60,000 up to 6,000,000. XFLR5 non-linear lifting line theory method (LLT) provides the wing aerodynamic characteristics, including aerodynamic coefficient over the whole wing: CL,CD,CM, etc., as well as the spanwise distribution of the local aerodynamic coefficients and other flow parameters. XFLR5 has the possibility to compute sequential analyses for varying values of angle of attack while either fixing air velocity or the lift produced by the wing. For a fixed angle of attack, a varying speed analysis can be performed.

(a) (b)

Figure 5.1: Morphing aircraft wing model: (a) 5,000mm wingspan (b) 3,500mm wingspan

An aerodynamic analysis varying speed and angle of attack was performed for the wing for ﬁve different span conﬁgurations: 3.5m, 4m, 4.5m and 5m to characterise the morphing wing and 4.3m to

59 represent the fixed wing. The analysis is performed at sea level considering a standard ISO atmosphere. Wing polars and the aircraft level flight graph are displayed in figures 5.2–5.4. The wing lift and drag coefficients are calculated using the area of the maximum wingspan configuration.

Figure 5.2: Wing lift coefﬁcient with angle of attack for different span wings

Figure 5.3: Polar curves for different span conﬁgurations

The cruise and loiter performance of the morphing wing and the fixed wings was studied. Both cruise and loiter flight conditions are considered steady and level flight conditions, i.e., where the altitude and speed are kept constant. From the balance of the forces applied to the aircraft we have the equations 5.1 and 5.2 defining respectively the steady and level flight conditions for small angles of attack.

T − D = 0 (5.1)

where T and D are the aircraft thrust and drag forces.

L − W = 0 (5.2)

60 Figure 5.4: Airspeed for level ﬂight with angle of attack for different span conﬁgurations

where L is the aircraft lift and W its weight. In cruise flight the objective is to maximize the distance possible to be travelled, by producing the necessary lift to level flight with the least drag possible. It is the condition that requires minimum thrust to maintain level flight and for a constant altitude the cruise speed is the one that provides the maximum value of the ratio CL/CD,[22]. The loiter flight phase is used to maximize the endurance of the aircraft in that flight condition. For level flight it is the condition that requires the least amount of power to maintain it. Loiter speed is the 3 2 flight speed that will have the maximum value of CL /CD ,[22].

Figure 5.5 shows the curve of the CL/CD with speed for the four different wingspan configurations of the morphing wing. As expected, depending on the flight speed some configurations are more efficient than others, i.e., have a higher lift to drag ratio. The lift coefficient necessary for level flight decreases as the airspeed increases and the wing configurations with more area will generate more drag; Thus from a certain point the most efficient wing configuration changes. Therefore, it is possible to draw an approximation of the optimum CL/CD curve with speed for the morphing wing using the highest available points. That graph is displayed in figure Figure 5.6. Figure 5.7 shows the equivalent curve for the loiter parameter for the varied wingspan configurations of the telescopic wing. Similarly to the cruise conditions the optimal curve to maximize available flight endurance can be approximated, figure 5.8.

3/2 In figures 5.9 and 5.10, the CL/CD and CL /CD curves with level flight speed of the 4.3m and 5m span fixed wings are compared with the corresponding approximated optimal curves of the morphing wing. Table 5.1 lists the optimal aerodynamic speeds of the morphing aircraft for both maximum and minimum wingspan configurations and likewise the speed of the 4.3m span fixed wing aircraft. To note that the optimal aircraft speeds of the fixed 5m span wing are equal to those of the maximum configuration span configuration of the telescopic wing. The drag produced by the wings was calculated for loiter, cruise and dive speeds. Table 5.2 presents

61 Figure 5.5: CL/CD with level ﬂight airspeed for different span conﬁgurations of telescopic wing

Figure 5.6: Optimal CL/CD curve with level ﬂight airspeed for the telescopic wing

3/2 Figure 5.7: CL /CD with level ﬂight airspeed for different span conﬁgurations of telescopic wing

62 3/2 Figure 5.8: Optimal CL /CD curve with level ﬂight airspeed for the telescopic wing

Figure 5.9: CL/CD curve comparison between optimal morphing wing and 4.3m ﬁxed span wing

3/2 Figure 5.10: CL /CD curve comparison between optimal morphing wing and 4.3m ﬁxed span wing

63 Table 5.1: Optimal aircraft speeds

Wingspan Va [kt] Vto [kt] Vl [kt] Vc [kt] Vd [kt] 3.5m 59.4 71.2 86.0 111.4 167.2 5m 47.7 57.2 68.6 83.4 125.1 4.3m 52.2 62.7 75.3 97.7 146.5 these results for all wing configurations and both fixed wings. Furthermore, the comparison between the drag of each configuration and the drag of the optimal morphing wing is provided in percentage of the optimal drag for the corresponding speed.

Table 5.2: Wing drag for different speeds with a comparison with the 5m span wing

Wingspan Drag (Vl=70 kt) Drag (Vc=90 kt) Drag (Vd=150 kt) [N] ∆ [%] [N] ∆ [%] [N] ∆ [%] 3.5m 84.8 +62.2 61.5 +29.6 72.8 0 optimal 52.3 0 47.5 0 72.8 0 5m 52.3 0 47.5 0 94.4 +29.7 4.3m 62.6 +19.8 51.1 +7.6 83.2 +14.3

It is noted that the smaller drag levels are found at the cruise speed. In table 5.3 are the estimated weight penalties of the morphing wing relative to the 4.3m and 5m span fixed wing. This measure corresponds to the necessary extra weight the optimal telescopic wing aircraft would have to carry to nullify the aerodynamic benefits of the morphing wing regarding the fixed wing. An analysis with the aid of XFLR5 was performed to find the total weight for each considered speed which would require the morphing wing aircraft to generated the same amount of drag as the corresponding fixed wing. This parameter is valuable in assessing a morphing application as it provides the additional weight limit the morphing device must not cross to provide aerodynamic benefits.

Table 5.3: Optimal morphing wing weight penalties to match ﬁxed 4.3m and 5m span wing drag

Weight Penalty [%] Vl=70 kt Vc=90 kt Vd=150 kt 4.3m (ﬁxed) 13.2% 9.5% 39.1% 5m (ﬁxed) 0% 0% 68.1%

5.2 Remarks

Analysing the wing polars, it is found the 5m wing produces higher lift coefficient values than the 3.5m wing. The consequence of this fact, as corroborated by figure 5.4 and table 5.1, is that for the aircraft to achieve level flight with the smaller wing configuration it needs to fly at higher speeds than the configurations with larger areas. Figure 5.3 shows that the wing configurations with lower span can produce less drag for the same level of lift when the lift coefficient is low enough. Inversely, for the higher lift coefficient levels the drag

64 cost for the 3.5m wing is heavier. Consequently, from figures 5.5 and 5.6 it can be concluded that the larger wing is more adequate for lower air speeds and the smaller wing to higher air speeds. 3/2 The CL /CD curves, in figures 5.5 and 5.6, show similar behavior to the CL/CD curves; however, this graph is more significant to finding the optimal loiter speed, as loiter far from the optimal velocity is not advisable. Comparing the optimal morphing wing aircraft to the fixed wings aircraft, results clearly show that the optimal telescopic wing outperforms the 5m span fixed wing at air speeds raging from about 100kt to the maximum air speed in terms of efficiency. Comparing the optimal morphing wing aircraft to the 4.3m span fixed wing we cannot see such a significant benefit at the higher speed range. Nonetheless, the morphing wing also outperforms the 4.3m wing at the lower air speed range, which is the more important region for surveillance missions. As flight envelope performance increase example, the morphing can maintain a small optimal loiter speed of 68.6kt for the surveillance and patrol part of missions while having a maximum speed increase of 33.6% in relation to the 5m fixed wing.

65 66 Chapter 6

Conclusions

6.1 Conclusions and Achievements

The main purpose of this work was to design a telescopic wing for a medium endurance patrol and surveillance UAV. The morphing telescopic wing concept capable of changing wingspan in fight was designed and its structure and aerodynamic performance were studied. As a starting point, the aircraft platform as well as the desired wing aerodynamic shape were defined based on the TEKEVER AR5. A market research of UAVs with similar size and mission role allowed to complement the information gathered and helped to characterize mission profile and performance objectives. As the assessment of a morphing technology is entirely case dependent, this characterization of the aircraft platform was used to compare the morphing wing performance to a conventional fixed wing solution. Several concepts and mechanisms to attain the desired in flight span variation and wing shape were investigated in section 3.2. The chosen concept was defined and its critical elements were sized. The aircraft has a design maximum load factor of 3.8 and an ultimate load factor of 5.7 as prescribed in the EASA CS-23 norm. The velocity-load diagrams of the morphing wing aircraft were analysed and a methodology to find the critical point structure wise and thereupon provide a structure capable of supporting the ultimate load factor was developed and implemented. The aerodynamic assessment of the telescopic wing performance was performed using the lifting line theory method through the XFLR5 software. Comparison between the morphing wing performance and both a fixed wing optimized for loiter and a fixed wing that represented a compromise solution was carried out. From the arrived results in section 5.1, it can be concluded the morphing wing will outperform significantly the optimized fixed wing in the extremes of the operating range. If the comparison is against the compromise solution for the fixed wing then the morphing wing will marginally or moderately outperform the fixed wing in the vast majority of the flight envelope. The morphing wing displays a drag reduction compared to the 5m wing optimized for the same loiter speed of about 23% for an air speed equal to 150kt; Compared with the 4.3m wing the morphing wing reaches a drag reduction of about 16%, 7% and 12% for air speeds of 70kt, 90kt and 150kt respectively.

67 These efficiency gains allow for a weight penalty of 68% when comparing the optimal morphing wing with the 5m wingspan fixed wing at 150kt. Comparing against the 4.3m wingspan fixed wing the weight penalties are 13%, 9% and 39% for air speeds of 70kt, 90kt and 150kt respectively.

6.2 Future Work

• Aerodynamic analysis with higher ﬁdelity CFD methods that can account for the morphing wing discontinuities will provide better insights into the performance of the telescopic wing and what beneﬁts it can give.

• Finite element structural analysis that studies all structural elements can help to trim down structure weight and assure structural integrity. With structure and its weight deﬁned, a stability analysis can be done.

• The design decision between the morphing wing or an alternative fixed wing should be assessed over the entire extension of the aircraft’s typical missions. The actuation system, power and propulsion should be determined and studied so the energy expended by the morphing aircraft over the mission profile can be compared to that expended by the fixed wing aircraft.

• Other morphing devices and its conjoined integration may be investigated to provide additional morphing capabilities such as sweep or camber morphing to further expand the ﬂight envelope.

• Perform experimental tests to assure the feasibility of the telescopic mechanism and corroborate the computational calculations.

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77 78 Appendix A

Technical Datasheets

A.1 TEKEVER AR5 Datasheet

79 AR5 MISSION: PATROL We’re always one step ahead to the unknown. Meet Europe's first UAS-based maritime surveillance system.

Wide Area Surveillance Maritime Surveillance Communications Relay

EMSA & ESA EUROPEAN MARITIME PATROLLER ITAR FREE SATCOM ENABLED

LOWEST TOTAL COST OWNERSHIP

Our System is your tool From software, to electronics, to airframe, we master all the details. We adapt. [email protected] www.tekever.com 2 Venture Rd, Chilworth We’re TEKEVER Hampshire We deliver your promise SO16 7NP, United Kingdom AR5 LIFE RAY Ready to deploy. The AR5 LIFE RAY Evolution is a medium-altitude and medium-endurance fixed wing UAS. Search & Rescue, maritime surveillance and patrol missions benefit from the increased endurance, reduced operating costs and lower risk to life offered by the AR5.

Multiple platforms, one system AR5 Tactical UAS ranging from 150 to 500Kg MTOW Dimensions 4.3 x 3 m Runway take off & landing. BLOS satellite communications High precision video, imagery and Cruise sensor data in real-time Speed 140 km/h

Flexible architecture, supporting Comms multiple types of payloads and datalinks Range unlimited

Highest production standards, and MTOW prepared for certification 150 kg

Used in multiple collaborative Payload projects for testing and validation 150 kg Capacity 50 kg ITAR Free Satcom enabled. Endurance Fully managed LoS and BLoS 12 h datalink handover Short unpaved runways for take-off Recovery and landing runway Automatic take-off and landing Launch (ATOL) runway

PAYLOAD OPTIONS THE EUROPEAN Multi sensor 3 axis gyro stabilized MARITIME PATROLLER Flexible payload options. gimbal AIS Mini SAR AR5 LIFE RAY EVOLUTION is the LIDAR European Maritime Patroller, LWIR selected by the European Space LRF Agency (ESA) and in partnership with the European Maritime Safety Agency (EMSA) to test and demonstrate European-wide UAS-based maritime surveillance activities Low maintenance.