design optimization and verification of a horizontal ... · in addition, the low airfoil of the...

Design Optimization and Verification of a Horizontal Stabilizer for the SeaStryder600 Wing-

In-Ground-Effect (WIG) Aircraft

by

Stephen Haley

A thesis submitted in conformity with the requirements for the degree of Master of Applied Science

Graduate Department of Mechanical and Industrial Engineering University of Toronto

© Copyright by Stephen Haley 2012

ii

Design Optimization of a Wing-In-Ground-Effect (WIG) Aircrafts

Horizontal Stabilizer

Stephen Haley

Master of Applied Science

Graduate Department of Mechanical and Industrial Engineering

University of Toronto

2012

Abstract

Aircraft manufacturer Aquavion Systems is currently designing and constructing prototypes for

its revolutionary new fleet of aircraft called the SeaStryder. During the prototyping phase, it was

discovered that the center-of-gravity of the SeaStryder600 was too far aft and outside of the

acceptable range. To solve this design issue, it was hypothesized that the weight of the horizontal

stabilizer may be reduced without compromising its structural integrity. The following document

analyzes this hypothesis and provides two alternative designs. Each design exceeds the design

requirements, meets additional requirements requested by industry, and provides a significant

degree of weight savings. The first design provides a 25% weight reduction. The second design

provides an 18% weight reduction as well as a 160% increase in loading capacity. The designs

proposed have both been verified through the use of Finite Element Analysis as well as by means

of experimentation where two prototype wings were constructed and tested to failure confirming

the analytical results.

iii

Acknowledgments

I would like to express my sincere gratitude towards the University of Toronto for providing me

with the opportunity to study at one of the most prestigious schools in the country. The following

research would not have been possible without the equipment, working facilities, and financial

support that was provided, and I am sincerely appreciative of this. Particularly, I would like to

thank my academic supervisor, Professor Cleghorn, for the guidance and support. The countless

hours spent reviewing and editing my research was greatly beneficial. Lastly, I would like to

thank my industry supervisor, Ray Richards, for providing the industry project with Aquavion

Systems Corp. The project was a pleasure to work on and taught me many invaluable skills,

particularly relating to the application and analysis of composite materials.

iv

Table of Contents

Table of Contents

Acknowledgments .......................................................................................................................... iii

Table of Contents ........................................................................................................................... iv

List of Tables ................................................................................................................................ vii

List of Figures .............................................................................................................................. viii

List of Appendices ........................................................................................................................ xii

1 Introduction ................................................................................................................................ 1

1.1 The SeaStryder .................................................................................................................... 1

1.2 Motivation ........................................................................................................................... 4

1.3 Objectives ........................................................................................................................... 4

1.4 Introduction to Wings and the Theory of Flight ................................................................. 5

1.4.1 Airfoil Definitions and Terminology ...................................................................... 5

1.4.2 Overview of aerodynamics ..................................................................................... 8

1.4.3 Boundary Layer Concept ...................................................................................... 11

1.4.4 Considerations for Finite Wings ........................................................................... 12

1.4.5 Planform Considerations: ...................................................................................... 13

1.4.6 Wing-in-Ground-Effect Aircraft ........................................................................... 14

2 Background Research ............................................................................................................... 15

2.1 Current Industry Design of Horizontal Stabilizer ............................................................. 15

2.1.1 Physical Configuration .......................................................................................... 15

2.1.2 Specifications Provided by Industry ..................................................................... 16

2.1.3 Current Design Issues ........................................................................................... 17

2.2 Future Design Considerations ........................................................................................... 18

v

2.2.1 Design Requirements ............................................................................................ 18

2.2.2 State-of-art ............................................................................................................ 20

3 Materials and Methods ............................................................................................................. 24

3.1 Design Concept ................................................................................................................. 24

3.1.1 Design Overview .................................................................................................. 24

3.1.2 Servo Housing and Elevator Control .................................................................... 25

3.1.3 Base Support Design ............................................................................................. 27

3.1.4 Composite Sandwich Structure ............................................................................. 28

3.1.5 Accelerometers ..................................................................................................... 29

3.2 Materials and Apparatus ................................................................................................... 29

3.2.1 Apparatus .............................................................................................................. 29

3.2.2 Materials Used ...................................................................................................... 30

3.3 Preliminary Experimentation ............................................................................................ 31

3.3.1 Base Support Experiment ..................................................................................... 31

3.3.2 Composite Sandwich Panel Experiment ............................................................... 35

3.4 Design Validation and Experimentation ........................................................................... 39

4 Results and Discussion ............................................................................................................. 44

4.1 Base Support Experiment ................................................................................................. 44

4.2 Composite Sandwich Panel Experiment ........................................................................... 45

4.2.1 Beam 1 .................................................................................................................. 45

4.2.2 Beam 2 .................................................................................................................. 46

4.2.3 Beam 3 .................................................................................................................. 47

4.2.4 Beam 4 .................................................................................................................. 49

4.2.5 Beam 5 .................................................................................................................. 50


4.3.1 Wing 1 ................................................................................................................... 53

vi

4.3.2 Wing 2 ................................................................................................................... 54

5 Conclusions and Recommendations ........................................................................................ 56

5.1 Base Support Experiment ................................................................................................. 56

5.2 Composite Sandwich Panel Experiment ........................................................................... 56


5.4 Scientific contributions ..................................................................................................... 58

5.5 Future Recommendations ................................................................................................. 58

Bibliography ................................................................................................................................. 59

vii

List of Tables

Table 1: Composite Sandwich Panel Structures ........................................................................... 29

Table 2: 4 Bar Linkage Specifications .......................................................................................... 90

Table 3: Beam Study Results ...................................................................................................... 101

Table 4: Wing Study Results ...................................................................................................... 104

Table 5: Deflection-Load Results ............................................................................................... 105

viii

List of Figures

Figure 1: Quarter Scale SeaStryder600 Hydroplaning ................................................................... 1

Figure 2: Double Dihedral Wing Configuration ............................................................................. 2

Figure 3: SeaStryder600 ................................................................................................................. 3

Figure 4: Airfoil Nomenclature ...................................................................................................... 6

Figure 5: Wing Sweep .................................................................................................................... 6

Figure 6: Control Surfaces .............................................................................................................. 7

Figure 7: Rib and Spar Configuration ............................................................................................. 7

Figure 8: Cartesian coordinate system used in the study of aerodynamics .................................... 8

Figure 9: Section Forces and Moments .......................................................................................... 9

Figure 10: Boundary Layer Flow over a Streamlined Body at a low angle of attack ................... 11

Figure 11: Example of Boundary Layer Separation ..................................................................... 11

Figure 12: Example of Wingtip Vortices ...................................................................................... 12

Figure 13: Airfoil Downwash ....................................................................................................... 13

Figure 14: Current Industry Design of the Quarter Scale Horizontal Stabilizer ........................... 15

Figure 15: Industry Testing of Support Spar ................................................................................ 16

Figure 16: Stryder600 Side View ................................................................................................. 17

Figure 17: Close up of Industry Design Showing Servo Assembly ............................................. 17

Figure 18: Servo Assembly Cover Plate ....................................................................................... 18

Figure 19: Configuration of R. Warsi Sullivans Design ............................................................... 21

ix

Figure 20: Failure of Design ......................................................................................................... 22

Figure 21: Z-pin reinforcement proposed by D. U. Long ............................................................. 23

Figure 22: Elevator Control Mechanism Model ........................................................................... 26

Figure 23: Elevator Control Mechanism Model Cross-Section .................................................... 26

Figure 24: Base Structure Model .................................................................................................. 27

Figure 25: Simple Base Structure Design ..................................................................................... 28

Figure 26: Base Structure with added Support Walls ................................................................... 28

Figure 27: Construction of First Support Base ............................................................................. 32

Figure 28: Construction of Second Support Base ......................................................................... 33

Figure 29: Addition of composite Sandwich Panel ...................................................................... 33

Figure 30: Base 1 - Unloaded ....................................................................................................... 34

Figure 31: Base 1 - Fully Loaded Prior to Failure ........................................................................ 34

Figure 32: Base 2 - Unloaded ....................................................................................................... 34

Figure 33: Base 2 - Fully Loaded Prior to Beam Failure .............................................................. 35

Figure 34: Composition of Beam 1 ............................................................................................... 35





Figure 39: Experimental Set-Up for Testing Elasticity ................................................................ 38

x

Figure 40: Vertical Stabilizer Initial Core Templates ................................................................... 39

Figure 41: Finishing the Vertical Stabilizer Core ......................................................................... 40

Figure 42: Finalizing the Vertical Stabilizer Replica ................................................................... 40

Figure 43: Wing Base Support ...................................................................................................... 41

Figure 44: Completing the Base Support ...................................................................................... 41

Figure 45: Completing the Wing .................................................................................................. 42

Figure 46: Divinycell Hotwire Cutting Failure ............................................................................. 43

Figure 47: Base Failure ................................................................................................................. 44

Figure 48: CAD Analysis of Beam 1 ............................................................................................ 46

Figure 49: Failure of Beam 2 ........................................................................................................ 46


Figure 51: Beam 2 - Unloaded (Top) and Fully Loaded (Bottom) ............................................... 47


Figure 53: Upper Surface Failure of beam 4 in Tension .............................................................. 49


Figure 55: Delamination Failure of Beam 5 ................................................................................ 51


Figure 57: Beam 5 - Unloaded (a) and Fully Loaded (b) ............................................................. 52

Figure 58: CAD Analysis of Wing 1 ............................................................................................ 54

Figure 59: CAD Analysis of Wing 2 ............................................................................................ 55

xi

Figure 60: Graphical Results of Beam Experiment ...................................................................... 56

Figure 61: 4 Bar Linkage Schematic ............................................................................................ 90

Figure 62: DS8711 Servo Torque (24.99 kg-cm)/Hinge Moment Characteristics ....................... 91

Figure 63: Beam 1 Loading Positions ........................................................................................... 92

Figure 64: Beams 2-5 Loading Positions ..................................................................................... 92

Figure 65: Wing Loading Positions ............................................................................................. 93

Figure 66: Accelerometer Input Vibration Example .................................................................... 94

Figure 67: Output after Separation of Bending and Torsional Vibration ..................................... 94

Figure 68: Beam 1 Cross-Section ................................................................................................. 95




Figure 72: Beam 3 Cross-Section ............................................................................................... 100

Figure 73: Deflection-Load Curve .............................................................................................. 105

Figure 74: CAD Deflection Study .............................................................................................. 105

Figure 75: Geometric Limit of Deflection ( 10 for illustrative purposes) ................................ 109

Figure 76: True Geometric Limit of Deflection ......................................................................... 109

xii

List of Appendices

Appendix I – Design Details and Documentation ........................................................................ 61

Appendix II – NACA 0009/0012 Profiles .................................................................................... 89

Appendix III – Hinge Moment Calculations and Characteristics ................................................. 90

Appendix IV – Beam Loading Positions ...................................................................................... 92

Appendix V – Example of Accelerometer Input / Output ............................................................ 94

Appendix VI – Experimental Results: Beam Bending Study ....................................................... 95

Appendix VII – Experimental Results: Wing Study ................................................................... 102

Appendix VIII –Experimental Results: Stiffness Test ................................................................ 105

Appendix IX – Material Data Sheets .......................................................................................... 106

Appendix X – Evaluation of Design Requirements .................................................................... 109

1

Chapter 1 Introduction

1 Introduction

1.1 The SeaStryder

Aquavion Systems Corp is a privately owned aircraft manufacturer located in Ajax, Ontario.

Over the past few years Aquavion Systems has been developing and testing a fleet of aircraft

called the SeaStryder. As a modern-day seaplane, the SeaStryder has been developed to provide

customers with the freedom of taking off and landing on water without the limited performance

of traditional seaplanes. In addition, the low airfoil of the SeaStryder allows it to operate as a

wing-in-ground-effect (WIG) aircraft, providing significant fuel savings. The SeaStryder is a

revolutionary conceptual aircraft, featuring hydroplaning float technology (HFT). This patented

technology employs integrated technological features which together provide an extremely

efficient and functional seaplane. A primary component of this technology is the design of the

forward wing. This low, buoyant wing provides both the necessary aerodynamic support in

flight, as well as the floatation required while making way in water. This wing configuration also

provides a high degree of lateral roll stability, at rest and while making way. During takeoff, the

flat dihedral underside of the wing provides an extremely efficient hydroplaning surface, quickly

raising the aircraft up and into a true hydroplane condition as shown in Figure 1.

Figure 1: Quarter Scale SeaStryder600 Hydroplaning

Dihedral float wing

hydroplaning

2

Once hydroplaning, takeoff is possible in only a matter of seconds, which is remarkable in

comparison to traditional seaplane designs. While very low hydroplaning speeds are possible, at

higher speeds the trailing edge of the main wing, near the root section, serves as a high speed

hydroplaning surface which allows for increase aircraft mobility and breaks the water surface

suction to allow for a smooth liftoff. This integration of functions is the basis of Hydroplaning

Float Technology (HFT). Traditional seaplanes require the use of sponsons and often additional

floats to provide the necessary lateral roll stability. These features, while necessary on the water,

inherently produce significant drag when airborne. This results in increased fuel consumption

and inferior performance. HFT eliminates all of this by allowing the wing to serve many

functions without compromising performance.

The main wing also contains a double dihedral to improve the stability and maneuverability of

the aircraft. The inner portion of the main wing consists of a 10 degree dihedral, similar to that of

a deep-v Planning hull. The outer portion of the wing is inclined an additional 5 degrees as

shown in Figure 2.

Figure 2: Double Dihedral Wing Configuration

Together, these features serve a variety of purposes. The initial 10 degree dihedral allows for

smooth operation on the water, easily cutting waves, allowing for banked turns of up to 10

degrees (which traditional seaplanes cannot do), and quickly getting the vessel into a planing

condition when necessary. The additional 5 degree incline prevents the entire wing from being

submerged thus eliminating the possibility of a dangerous spinout on the water. As well, the

discontinuity minimizes damage due to sea spray by limiting the wetted surface to only the inner

dihedral. This second incline also naturally causes the aircraft to be self-leveling during flight.

Most aircraft, if left uncorrected, will gradually begin to bank to one side or the other until the

plane eventually enters into a spiral dive. To avoid this, continuous pilot input or an auto-pilot

Additional 5

Degree Incline

3

system is necessary. The double dihedral configuration inherently maintains level flight and even

after a banked turn, will automatically return the aircraft into a balanced level position.

The design of the horizontal stabilizers (or rear wings) also play a vital role in the HFT.

Traditional aircraft generally maintain a slight downward pitching moment about the main wing.

This leaves the horizontal stabilizers responsible for providing a slight down-lift to compensate

and keep the aircraft balanced. The SeaStryder operates in a tandem wing configuration whereby

both the main wing and the horizontal stabilizers provide significant uplift. This is due to the fact

that the aircraft must maintain a more aft-ward center of gravity to allow for smooth

hydroplaning when navigating as a seagoing vessel. The tandem wing configuration provides the

additional advantage of an increase ranged and endurance as there is no longer a need to

maintain the balancing moment known as trim-drag.

The SeaStryder series has a broad potential market ranging from general aviation to commercial

uses such as passenger and cargo transportation, military use and UAV (Unmanned Aerial

Vehicle) applications. The SeaStryder is also particularly well suited for high-speed ground

effect transportation. Any aircraft with a low lying primary wing such as the SeaStryder is highly

susceptible to experiencing the advantageous effects of ground effect when flying within one

wing-span of the ground. At this low altitude the air beneath the wing acts as a cushion beneath

the wing and dramatically increases the fuel efficiency though a “cushioning” effect. This

phenomenon is discussed in more detail in Section 1.4.6, Wing-in-Ground-Effect Aircraft.

The SeaStryder600 shown in Figure 3 is one of three varying configurations of the SeaStryder.

Figure 3: SeaStryder600

The smallest of the SeaStryder line up is the SeaStryder200, a two passenger, 773 [kg] aircraft

with a cabin width of 1.2 [m]. This is followed by the SeaStryder600, a 4 passenger, 2273 [kg]

aircraft with a cabin width of 2.0 [m]and standing headroom. Lastly, is the much larger

4

SeaStryder6000, a 100 passenger commercial aircraft capable of supporting a 4560 [kg] payload.

The SeaStryder600 is the design of interest in this research. A quarter scale prototype of the

aircraft has been constructed for analysis and preliminary flight testing. The following are the

specifications for the full scale SeaStryder600 aircraft proposed by Aquavion Systems:

Amphibious operation

Sleeps 4 people (2 fore & 2 aft)

Stand-up headroom (72 inches)

80-inch cabin width

Lav/Shower/Galley facilities

Door & Canopy entry

Hydroplaning floatwing for lowest drag in flight

Walkway or on-water swim access to engine/props

Fly-by-wire control operation

Glass cockpit

Fully composite airframe

Tricycle gear, water thrusters

5000 lb (2273 kg) maximum weight

Range: 1300 nm

Max: 228 mph (198 kts)

Cruise: 200 mph (174 kts)

59 mph (51 kts) stall with flaps

Single 600 hp turboprop or twin 300 hp piston engines

1.2 Motivation

Industry testing of the scale prototype model of the SeaStryder600 has revealed the need to

reduce weight in the tail end, specifically in the horizontal stabilizer fins. The following research

is premised on the belief that greater weight reduction is achievable by incorporating modern

design methods and materials. This must be done while maintaining the structural rigidity and

damping to avoid aeroelastic flutter.

1.3 Objectives

The primary objectives of the research are to design and verify a new horizontal stabilizer with

reduced weight, and acceptable strength and stiffness characteristics. The scope includes

identifying optimal characteristics of key design features and components, specifically, material

choices, panel thicknesses, internal geometry, and eventually the resulting analytically defined

properties and fabrication method.

5

1.4 Introduction to Wings and the Theory of Flight

1.4.1 Airfoil Definitions and Terminology

The following terms are used throughout the document and should be understood:

Leading Edge: The forward edge of a wing, see Figure 4.

Trailing Edge: The aft edge of a wing, see Figure 4.

Chord Length: Distance from the leading edge to the trailing edge, see Figure 4.

Wingspan: The distance measured from one wingtip to the opposing wingtip.

Aspect ratio: Ratio of wingspan to chord length.

Camber line: The line midway between the upper and lower surfaces of an airfoil section as

measured perpendicular to the camber line itself, see Figure 4.

Downwash: The downward velocity component of air flowing off the trailing edge.

Planform: The shape and layout of a fixed-wing aircraft's fuselage and wing structure.

Sweep: Angle between the wingspan and centerline of aircraft (0° sweep would be perpendicular

wings), see Figure 5.

Port/Starboard: Starboard (Stbd) refers to the right hand side of the aircraft when facing forward,

Port refers to the left.

Control Surface: A movable surface used to direct the flow of air off the aircraft, thus providing

a means of aircraft control, see Figure 6.

Aileron: A control surface mounted on the trailing edge of the main wing of an aircraft, see

Figure 6.

Elevator: A control surface mounted on the trailing edge of the horizontal stabilizer, see Figure

6.

Spar: Main wing support structure generally extending the length of the wingspan, see Figure 7.

6

Rib: A component extending from the spar providing a profile to form the airfoil shape, see

Figure 7.

Figure 4: Airfoil Nomenclature1

Figure 5: Wing Sweep2

1 Warren, F. (2009) Airfoil Nomenclature. [Print] Mechanics of Flight. 1st ed. Wiley

2 Chan and Song. (2009) Wing Sweep. [image online] Available at: http://ap-physics-b-

project.wikispaces.com/Aerodynamics+of+Airplanes+-+Chan+%26+Song [Accessed: May 2012]

7

Figure 6: Control Surfaces3

Figure 7: Rib and Spar Configuration4

3 Fear of Flying Help. (2008) Control Surfaces. [image online] Available at: http://www.flyingfear.net/articles/how-

an-aircraft-flies-control-surfaces.html [Accessed: May 2012]

4 Airbus Airplane Wing Structure and Terminology (2007) Airplane Wing Components. [image online] Available at:

http://www.nomenclaturo.com/tag/airplane-wing [Accessed: May 2012]

8

1.4.2 Overview of aerodynamics

Force and moment nomenclature:

These forces and moments are due to the net effects of pressure and shear stress distributions

integrated over the entire surface of the body where:

The notations follow the commonly used Cartesian coordinate system as seen in Figure 8.

Figure 8: Cartesian coordinate system used in the study of aerodynamics5

5 Warren, F. (2009) Cartesian Coordinate System. [Print] Mechanics of Flight. 1st ed. Wiley

9

In two dimensions, the aerodynamics forces acting on an airfoil section are completely defined in

terms of either axial and normal forces, or lift and drag. These two representations of

aerodynamics forces are related solely by the angle of attack, , which is the angle between

and the section cord line, as presented in Figure 9:

Figure 9: Section Forces and Moments6

In terms of dimension-less coefficients these become:

Eq. 1

Eq. 2

Eq. 3

Eq. 4

Eq. 5

6 Warren, F. (2009) Section forces and moments. [Print] Mechanics of Flight. 1st ed. Wiley

10

In relation to airfoils, and are the free stream density and velocity, respectively; is the

planform area, and is the mean chord length.

denotes the aerodynamic coefficient per unit span. Therefore, the relationship between lift and

drag, and, axial and normal forces, can be described as:

Eq. 6

Eq. 7

or,

Eq. 8

Eq. 9

Two important locations along the cord line are: the Center of Pressure, , defined as the point

about which the resultant moment is zero; and the Aerodynamic center, , defined as the point

about which the resultant moment is independent of .

In dimensionless terms, it can readily be shown that the equations derived for , and are:

Eq. 10

Eq. 11

For incompressible flow, the quarter cord is generally assumed to be the aerodynamic center7.

While this is only an approximation, it is quite accurate and often used for subsonic flight under

Mach number 0.3. Also, for symmetric airfoils, the quarter cord can also be shown to be the

center of pressure.

7F., Warren. Mechanics of Flight. 1st ed. Wiley, 2009. Print.

11

1.4.3 Boundary Layer Concept

For the high-Reynolds-number flows typically encountered in flight, the viscous shear forces are

small relative to the pressure forces. This allows analysis to be simplified to an inviscid flow

analysis outside of the boundary layer.

Boundary Layer separation, commonly referred to as a stall, occurs because of a positive

pressure gradient induced on the upper aft portion of the wing during high attack angles. This

positive pressure gradient eventually reaches a critical value where the flow cannot sustain itself

along the surface of the airfoil and causes the boundary layer to separate from the airfoil as

shown in Figure 10 and Figure 11.

Figure 10: Boundary Layer Flow over a Streamlined Body at a low angle of attack8

Figure 11: Example of Boundary Layer Separation9

8 Warren, F. (2009) Boundary layer flow over a streamlined body at a low angle of attack. [Print] Mechanics of

Flight. 1st ed. Wiley

9 Warren, F. (2009) Examples of boundary layer separation. [Print] Mechanics of Flight. 1st ed. Wiley

12

1.4.4 Considerations for Finite Wings

Wing tip vortices

For an infinite wing, it is correctly assumed that the flow of air only takes place in the plane

consisting of the foil cross section. However, for a finite wing, this is not entirely true,

particularly near the wing tips. At the tip of a wing, air is able to flow from the high pressure

regions below the wing to the low pressure regions above the wing, thus creating a vortex

moving in the span wise direction as seen in Figure 12. This vortex is the primary component of

drag, commonly known as “induced drag”, experienced by an aircraft in subsonic flight.

Attempts to reduce this induced drag include wing tapering and the inclusion of winglets to

prevent vortex formation.

Figure 12: Example of Wingtip Vortices10

Downwash

Downwash is a physical phenomenon predicted by both thin airfoil theory as well as the vortex

panel method. The downwash, located directly below the aircraft, is equally balanced by upwash

10 Langley Research Center (1990) Airplane Vortex. [image online] Available at: http://archive.org/details/NIX-EL-

1996-00130 [Accessed: Dec 2010]

13

beyond the tip vortices, shown in Figure 13. As the size of the tip decreases compared to the

overall size of the wing, the magnitude of the induced drag is reduced due to the reduced

effective downwash.

Figure 13: Airfoil Downwash11

1.4.5 Planform Considerations:

Most airfoils are subjected to induced drag caused by air flowing around the tip of the wing in

the span-wise direction. By decreasing the size of the tip of the wing, or by adding “winglets”,

this induced drag can be greatly reduced. Also, by increasing the aspect ratio of the airfoil, the

wing becomes more effective in that it has a lower amount of induced drag relative to the lift

generated. These methods of reducing drag are commonly seen on lower speed aircraft in the

form of various types of winglets or raked wingtips, as well as elliptical, and more commonly

tapered, wing profiles. A recent study on the use of winglets, as well as varying degrees of taper

at sub-sonic speeds, demonstrates the benefit of using winglet technology as well as the

importance of wing tapering12

. The SeaStryder currently incorporates winglets, raked wingtips,

as well as a tapering wing profile.

11 Scott, J. (2005) Trailing Vortex. [image online] Available at:

http://www.aerospaceweb.org/question/nature/q0237.shtml [Accessed: Dec 2010]

12 Azlin, M.A., C.F. Mat Taib, S. Kasolang, and F.H. Muhammad. "CFD analysis of winglets at low subsonic flow."

World Congress on Engineering. (2011): 87-5.

14

High-speed aircraft are influenced by a phenomenon known as wave drag. During high speed

flight, wave drag is significantly more powerful then induced drag, and therefore, airfoils are

designed solely around reducing this type of drag. Common wing designs of high speed aircraft

include highly swept wings with a very thin profile, changing slowly over a larger cord length.

This design allows for reduced drag during high-speed flight, however, poor low speed

performance. The SeaStryder is not designed for high-speed flight where wave drag becomes a

problem and is resultantly characterized as a low-speed (Sub-Sonic) aircraft.

1.4.6 Wing-in-Ground-Effect Aircraft

A Wing-In-Ground effect (WIG) aircraft such as the SeaStryder is a specialized type of aircraft

designed to take advantage of a well-known phenomenon in aviation known as ground effect.

When in flight within approximately one wingspan of the ground, an aircraft experiences a

floating sensation caused by the increased pressure beneath the airfoil. This “cushion effect”

reduces the required angle of attack by providing additional lift, and as well, provides a reduction

in the induced drag. The induced drag, primarily caused by wingtip vortices is reduced due to the

inability of the vortices to form freely on the wing tips. Theoretically, a wing that does not

produce any wingtip vortices has virtually no induced drag. This effect is often seen in wind

tunnel tests where the wing spans the entire width of the wind tunnel. As an aircraft approaches

the ground, these effects are magnified, particularly on aircraft with low airfoils such as the

SeaStryder.

15

Chapter 2 Background Research

2 Background Research

2.1 Current Industry Design of Horizontal Stabilizer

2.1.1 Physical Configuration

The current industry design consists of a rib and spar configuration followed by a thin fiberglass

outer shell as shown in Figure 14.

Figure 14: Current Industry Design of the Quarter Scale Horizontal Stabilizer

The support spar is a hollow tapering channel beam constructed from carbon fiber while the ribs

were constructed from balsa wood. A female mold of the wing was constructed and used to form

the fiberglass surface into the desired NACA profile. The profile chosen for the wing by industry

was a NACA 0012 (root) to NACA 0009 (tip) tapering profile. These profiles can be seen in

Appendix II – NACA 0009/0012 Profiles. The wing connected to the fuselage by sliding the

hollow support spar on to an additional support member which supported both the left and right

wing spars through the vertical tail section. Two threaded rods extended from the wing base and

were used to securely fasten each wing to its opposing wing, creating a clamping action across

the vertical stabilizer. The elevator hinge was constructed from a Kevlar strip that ran the entire

span of the wing, joining the elevator to the wing but leaving a 1.59 [mm] separation gap

between the two. This provided a simple, non-mechanical hinge with excellent longevity and

durability and is desired in future designs. The elevator made up for 119 [mm] of the trailing

Wing Root

NACA 0012

Wing Tip

NACA 0009 Servo Housings

Elevator Servo Assembly Support Spar

Threaded Rods

Elevator Hinge

16

edge of the wing. The core section of the wing, however, varied in chord length from root to tip.

The root NACA profile, located on the centerline of the fuselage, was scaled to provide a 474

[mm] root chord length, and the tip profile, located 1219 [mm] from the centerline, was scaled to

provide a 320 [mm] tip chord length. The elevator servos were mounted within the rib and spar

space frame. Control rods extended through the fiberglass surface and were connected to the

lower edge of the elevator. Due to this design, additional cover plates and flexible gasket

couplings were necessary to protect the control rod and seal off the servo from water and outside

debris. The construction methods and materials involved in the manufacturing of this design

were limited to commonly available tools, equipment and materials and all layup procedures

were performed by hand.

2.1.2 Specifications Provided by Industry

The load capacity of the complete wing assembly was never tested, however since the design is

internally supported by a single support spar, the spar was individually tested prior to assembly

as shown in Figure 15.

Figure 15: Industry Testing of Support Spar

The spar was deemed capable of supporting a load and moment of 240.2 [N] and 113.2 [Nm],

the expected load due to a 6G dive of a 35.38 [kg] aircraft. The stiffness characteristics (Bending

and Torsion) of the spar or wing assembly were not measured. The combined weight of the left

and right vertical stabilizers was 4.49 [kg], therefore, each stabilizer weighed approximately

2.245[kg].

Horizontal Stabilizer

Support Spar

Weights used for testing

Vertical Stabilizer

17

2.1.3 Current Design Issues

Any aircraft’s center of gravity (CG) must be positioned within a limited range in order to

maintain a certain level of stability during flight. Because of this, the consequence of any

unnecessary weight aft of the CG is amplified due to the fact that it must be compensated for

forward of the CG to maintain stability. In the case of the Stryder600, this compensating weight

is added in the nose cone of the aircraft to provide the most benefit. It has been determined by

industry that the aft weight of the aircraft is too great and is causing unnecessary amounts of

weight to be added to the nose cone as compensation. It is believed that the current horizontal

stabilizers are one of the primary reasons for this excess aft weight and that an optimization of

the current design is needed. The horizontal stabilizers are located near the most aft section of

the aircraft as seen in Figure 16 and therefore any unnecessary weight in this component is

highly unfavorable.

Figure 16: Stryder600 Side View

An additional minor issue with the horizontal stabilizer is the servo control rod mechanism

shown previously in Figure 14. The control rods currently extend past the fiberglass surface as

shown in Figure 17, exposing the servo and control mechanism.

Figure 17: Close up of Industry Design Showing Servo Assembly

Servo control

rod

Horizontal

stabilizer

position

Elevator

18

This exposure causes the mechanism to become susceptible to water and/or physical damage,

and disrupts the airflow around the wing. Because of this, the design requires an additional cover

plate and flexible gasket coupling to help protect the servo and control rod as shown in Figure

18, however these additional external components are undesirable.

Figure 18: Servo Assembly Cover Plate

2.2 Future Design Considerations

2.2.1 Design Requirements

The following is a list of the requirements that the new design must meet for it to be a suitable

replacement for the current design; these requirements have been requested by Aquavion

Systems:

The weight of each horizontal stabilizer must be less than the current weight of 2.25 [kg].

Ideally, a weight reduction of at least 10% from the current design would be favorable.

The horizontal stabilizer must be able to safely withstand a 6G dive. As well, an

assumption is to be made that the newly designed aircraft will increase the previous

loading on the wings by 25% due to its additional weight. This equates to a load of 450.5

[N] and 224 [Nm] on each horizontal stabilizer calculated in Appendix X – Evaluation of

Design Requirements: Minimum Load Requirement, while maintaining a minimum

factor of safety of 1.5

This load must not cause a deflection so severe that the hinge can no longer pivot. This

equates to a maximum allowable deflection across the horizontal stabilizer hinge line of

Servo cover

plates

19

12.7 [mm] as shown in Appendix X – Evaluation of Design Requirements: Maximum

Deflection Requirement. Any deflection beyond this will not allow the elevator to pivot

and cause a loss of control.

The elevator range of motion should meet or exceed the current design of a +/- 25 degree

incline/decline of the leading control surface.

The elevator hinge moment provided by the servo control arms must exceed the

maximum expected moment caused throughout all operational conditions. As well, the

new design should employ the use of at least three servos along the hinge line similar to

the current design.

These servo motors must be serviceable, as well as protected from weather.

The horizontal stabilizers should utilize the proven Kevlar hinge concept similar to the

current design.

The horizontal stabilizers should contain an accelerometer(s) for measuring bending and

torsional vibration.

The horizontal stabilizers must be easily removable and connect to the vertical stabilizer

in the same fashion as the current industry design at a 3.5 degree positive rigging angle.

The horizontal stabilizers must follow the same NACA 0012/0009 profile as seen in

Appendix II – NACA 0009/0012 Profiles.

The materials and manufacturing methods used must be practical and cost effective.

Minimizing the cost of manufacturing is a key objective. As well, the cost of prototyping

must be kept to an absolute minimum as no further funding is available.

The horizontal stabilizers must be able to support the addition of winglets.

A structural prototype should be constructed to confirm all results.

20

2.2.2 State-of-art

The internal configuration of current leading edge designs with similar design requirements

generally consist of a primary spar, a cellular foam core material, and a load bearing skin

material as the outer shell. A recent comparable study performed by C. Kong on the structural

design of a horizontal tail wing demonstrates this highly effective design13

. Utilizing the skin as

a load bearing surface is inherently the optimal way to improve the rigidity as the second

moment of area, also known as the moment of inertia, is exponentially related to the distance

from the cross-sectional plane to the load bearing material. This can be seen in Equation 12 for

the Second Moment of Area.

Eq. 12

It is clear that the most beneficial configuration to maximize the Second Moment of Area is to

use the most distant area possible, the wing surface, as the load bearing structure. In this

configuration carbon fiber is commonly used for high-end aircraft as it is commercially available

and provides a high level of strength and rigidity relative to its weight, much higher than the

common alternative material, fiberglass. The core materials used in this type of composite

sandwich panel structure vary from once design to another. The analysis performed by C. Kong

utilized a carbon fiber shell which surrounded Divinycell, an aerospace grade structural foam.

Honeycomb and aluminum core materials are often used as well, as described in a recent

sandwich panel study performed by Jong Woong Lee14

. This study tested sandwich panels

commonly used as wing skins and fuselage components consisting of graphite/epoxy facesheets

surrounding aluminum honeycomb core materials. While extremely light, a leading disadvantage

noted in this study was the unpredictability of a buckling failure. The buckling failures occurred

at stresses consistently significantly lower (nearly half) than predictions made by FEM analysis.

The primary cause of failure was delamination of the facesheets due to the limited bonding area

13 Kong, C., H. Park, Y. Kim, and K. Kang. "Structural Design on Wing of a Small Scale WIG Vehicle with

Carbon/Epoxy and Foam Sandwich Composite Structure." 16th International Conference on Composite Materials.

(2007).

14 Lee, Jong Woong, Cheol Won Kong, Se Won Eun, Jae Sung Park, Young Soon Jang, Yeong Moo Yi, and

Gwang-Rae Cho. "Compression test of composte sandwich panel." Key Engineering Materials. 348-349. (2007):

605-4

21

of the honeycomb structure, which eventually lead to global buckling. Because of this, only

small-cell or closed cell core materials are generally used when buckling is a potential failure

mode. In a study performed by R. Warsi Sullivan testing carbon-composite Unmanned Air

Vehicle (UAV) wings, where buckling was known to be a potential failure mode, Divinycell was

again the material chosen as the primary core material15

. This study, along with a similar recent

study performed by Myoungkeon Lee on the composite wing structure of an UAV16

, are of

particular relevance, as weight reduction in an UAV is critical to its operational limitations.

Myoungkeon Lee concluded that when utilizing a load bearing carbon fiber skin, the buckling

load was the limiting factor and efforts to improve buckling strength were substantially more

critical than improving material strength. As well, this study noted the importance of orienting

the carbon fiber plies at ±45° to the buckling axis to improve the buckling strength. In the design

proposed by R. Warsi Sullivan, a rib and spar structure supported an upper and lower sandwich

panel skin as shown in Figure 19.

Figure 19: Configuration of R. Warsi Sullivans Design17

15 Sullivan, R. Warsi, Y. Hwang, M. Rais-Rohani, and T. Lacy. "Structural Analysis and Testing of and Ultralight

Unmanned-Aerial-Vehicle Carbon-Composite Wing." Journal of Aircraft. 46. no. 3 (2009): 814-7.

16 Lee, Myoungkeon, Changmin Cho, and Seyong Jang. "HALE UAV Composite Wing Structure Design."

Advanced Materials Research. 123-125. (2010): 105-4.

17 Sullivan, R. (2009) Panel Configuration. [Journal Article] Structural Analysis and Testing of and Ultralight

Unmanned-Aerial-Vehicle Carbon-Composite Wing

Rib

Spars

Upper/Lower

Surfaces

22

The upper and lower skins consisted of carbon fiber plies with a low density Divinycell H60

foam core. Figure 20 shows the failure of this design due to buckling at the aileron cut-out.

Figure 20: Failure of Design18

Looking more closely at the buckling failure of sandwich structures, a study performed by

Baoxiang Shan noted that the critical bucking load of simple sandwich structures can be

approximated by employing the use of Equation 13, Euler’s Buckling load equation19

.

Eq. 13

Here, is the variable of interest as it is the thickness of the panel. It is clear that an increase in

thickness has a dramatic effect on the critical buckling strength as the thickness is exponentially

related to the critical buckling load.

All of the designs discussed thus far contain a rigid rib and spar type internal configuration

which provide the shape and transfer the skin load to a single spar. A conceptual design for a

18 Sullivan, R. (2009) Panel Failure. [Journal Article] Structural Analysis and Testing of and Ultralight Unmanned-

Aerial-Vehicle Carbon-Composite Wing

19 Shan, B, and A Pelegri. "Apporoximate Analysis of the Buckling Behavior of Composites with Delamination."

Journal of Composite Materials. 37. no. 8 (2003): 673-13.

23

spar-less sandwich panel wing design was proposed by D. U. Long20

. With this design the core

shear stresses normally supported by a spar are supported by the core as well as the addition of a

novel Z-pin reinforcement technique shown in Figure 21.

Figure 21: Z-pin reinforcement proposed by D. U. Long21

As shown, the cross-member reinforcements provide lateral stability, transferring the shear loads

to the facesheets. The transfer of these loads to the fuselage was not discussed in this paper as it

was purely a conceptual design proposal. An additional study performed by Zhang Ji-Kui22

also

concluded that by stitching the laminate sandwich panel, the buckling strength was improved,

and in addition, the torsional stiffness was improved, which is critical in flutter prevention.

20 Long, D.U. "Mechanical Properties and Structural Design of a Novel Composite Box for High Aspect Ratio

Wing." Journal of THERMOPLASTIC COMPOSITE MATERIALS. 24. no. 11 (2011): 819-17.

21 Long, D.U. (2011) Z-Pin reinforcement. [Journal Article] Mechanical Properties and Structural Design of a Novel

Composite Box for High Aspect Ratio Wing

22 Ji-Kui, Zhang, Li Zheng-neng, and Kou Chang-he. "Structural Design of High Aspect Ratio Composite Material

Wing." Institute of Aircraft Design, Beijing. (2004).

Core Material

Z-pin

reinforcement

Upper/Lower

Surfaces

24

Chapter 3 Materials and Methods

3 Materials and Methods

3.1 Design Concept

The following section describes the details of the proposed design. This design is based on state-

of-art techniques, cost and manufacturing constraints as well as available materials.

3.1.1 Design Overview

Two primary considerations when designing the internal configuration of the wing were strength

and stiffness. Strength limits the amount of lift the wind can provide, particularly relevant during

a dive recovery maneuver. From the design requirements, the maximum lift required equates to a

load of approximately 450 [N]23

. This occurs during a 6G dive recovery operation, where the

aircraft is experiencing six times the normal gravitational force. Wing Stiffness is of equal

importance. If the wing does not have adequate stiffness, then it may be prone to flutter, a

catastrophic condition whereby the wing begins to vibrate at its natural dynamic frequency. This

critical flutter frequency is a function of stiffness, damping, and wing speed. As well, if the

deflection of the wing is too great, the elevators can no longer pivot, as the hinge line becomes

an arc. Based on the wing requirements, additional weight reduction requirements, and cost

constraints, the proposed design utilizes a load-bearing skin, fully surrounding an ultra-light

core, to form a single composite sandwich structure similar to the design proposed by D. U.

Long. An obvious disadvantage of this design is the inability to utilize the space inside the

wings, however, this is not relevant for stabilizers as they are too far aft. As well, additional

loading would cause poor weight distribution and would defeat the purpose of the weight

reduction. This design can be manufactured cost effectively both in terms of material and

manufacturing methods. By using the wing skin as the support frame, the maximum possible

stiffness to weight ratio is achieved as the moment of inertia is exponentially related to the

distance from the mid-plane to the load bearing material. However, this design is highly

susceptible to surface buckling in compression as seen in the study performed by R. Warsi

23 Appendix X – Evaluation of Design Requirements, Minimum Load Requirement

25

Sullican24

. Therefore, methods to increase the buckling strength are critical. As mentioned in the

study performed by Myoungkeon Lee, one method of improving the buckling strength is to

orient the carbon fiber plies at ±45° to the buckling axis25

. In addition, by using the entire

thickness of the wing as a sandwich panel, as opposed to separate upper and lower sandwich

panel as developed by C. Kong26

, the panel thickness is maximized, a critical factor in

determining buckling strength. An additional complication of this design arises due to the load

transfer from the wing to the fuselage. Traditional rib and spar configurations transfer the wing

load to the fuselage through a main support spar. With the solid core skin support frame concept,

no such spar exists, and therefore an alternative method must be used. An innovative solution to

this problem involves the use of a specialized base support discussed in this section. The

materials selected for this design consisted of carbon fiber as the structural material, and

Divinycell-H80 and polystyrene for the core materials. Carbon fiber was selected for the

structural material because it provides the highest level of stiffness and strength relative to its

weight, compared to materials that are commonly available. Divinycell H80 was selected as the

base core material as it is an aerospace grade structural foam which provides a cost effective

lightweight solution. It is inherently extremely stiff, and is commonly used in conjunction with

carbon fiber and in sandwich structures to prevent buckling. Finally, polystyrene was selected as

a primary core material due to its excellent formability and extremely low density.

3.1.2 Servo Housing and Elevator Control

The two primary elevator control issues with the current design are the service hatch locations

and the protrusion of the elevator linkage arms. The proposed design addresses both of these

issues through the use of an extremely light and compact servo housing assembly, designed to

allow the servo motor to be accessed and serviced from the trailing edge of the wing, between

the elevator and wing as shown in the Figure 22. The servo housing assembly has been designed

24 Sullivan, R. Warsi, Y. Hwang, M. Rais-Rohani, and T. Lacy. "Structural Analysis and Testing of and Ultralight

Unmanned-Aerial-Vehicle Carbon-Composite Wing." Journal of Aircraft. 46. no. 3 (2009): 814-7.

25 Lee, Myoungkeon, Changmin Cho, and Seyong Jang. "HALE UAV Composite Wing Structure Design."

Advanced Materials Research. 123-125. (2010): 105-4.

26 Kong, C., H. Park, Y. Kim, and K. Kang. "Structural Design on Wing of a Small Scale WIG Vehicle with

Carbon/Epoxy and Foam Sandwich Composite Structure." 16th International Conference on Composite Materials.

(2007).

26

to enable the linkage arms to retract into the housing as the servo is controlled, illustrated in

Figure 23.

Figure 22: Elevator Control Mechanism Model

Figure 23: Elevator Control Mechanism Model Cross-Section

This allows for a more compact design which does not protrude from the wing or cause any

aerodynamic disturbances on the wing surface. The housing design also meets the requirement to

protect the servo and electrical conduits from extreme weather conditions.

The torque and range of motion of the elevator flap was analyzed using a standard four-bar

linkage approach. An estimate of the pressure generated on the flap, throughout all elevator

positions during maximum speed flight, provided a basis from which to calculate the minimum

required servo motor torque. Based on the findings, a specific servo was selected, the JR DS8711

Digital Ultra Torque Metal Gear Servo, commonly used in aviation. The resultant moment

provided by the servo, transmitted through a variable four-bar linkage, was calculated for each

Fully extended

elevator

Fully retracted

elevator

Servo housing

location

27

possible linkage scenario at all flap angles and is provided in Appendix III – Hinge Moment

Calculations and Characteristics. As specified, this design also uses a non-mechanical Kevlar

hinge and is designed to provide a +/- 25 degree flap angle.

3.1.3 Base Support Design

The base support is a critical component of the load bearing skin concept. It provides a method

of transferring the load from the skin to a single spar where it connects to the vertical stabilizer.

Figure 24 illustrates the relationship between the vertical stabilizer and the wing base support.

Figure 24: Base Structure Model

The structure makes up for the initial 7.6% of the wingspan. It is designed to alleviate stress

concentrations caused by the load transfer and provides a solid uniform base from which the

remainder of the wing extends. The base support design allows the entire wing to be removed as

required. Given the structural importance of this component (since its failure would be

catastrophic) two designs were developed, constructed, and tested in the first round of

experimentation. The first design consisted of a structural foam core (Divinycell-H80)

surrounded by four layers of 5.7oz carbon fiber. A carbon fiber pipe made from six layers of

carbon fiber extended entirely through this structure and was the support channel for the vertical

stabilizer connecting rod. A cross-section of this design is shown in Figure 25.

Base

Support

Structure

28

Figure 25: Simple Base Structure Design

While this design has been developed to withstand the expected loads, it is inherently susceptible

to buckling on either edge of the support channel when overloaded. During the initial stages of

experimentation, it was concluded that this type of failure posed a significant risk which lead to

an alternate design with added support walls. This addition does not significantly add to the

weight and dramatically increases the loading capabilities of the base structure. However, it did

add a significant level of complexity to the manufacturing process. The added support walls can

be seen in the cross-section, Figure 26:

Figure 26: Base Structure with added Support Walls

3.1.4 Composite Sandwich Structure

The internal configuration of the wing, beyond the base support, is a composite sandwich panel

structure consisting of an ultra light foam core surrounded by a carbon fiber (CF) shell. The core

material carries both the shear forces and helps to prevent buckling of the thin-walled outer shell.

The primary failure mode of the composite sandwich panel structure is local surface buckling.

Therefore, the goal was to maximize the local buckling strength while minimizing the overall

weight. As well, the stiffness of the wing must be adequate to prevent flutter and allow for full

elevator movement during maximum load flight. Based on industry recommendation, available

materials and manufacturing capabilities, five varying configurations of composite sandwich

panel structures were considered. The configurations of these panels are given in Table 1:

Added

support

walls

Support

Channel

29

Configurations of Composite Sandwich Panel Structures under Consideration

Configuration 1 Configuration 2 Configuration 3 Configuration 4 Configuration 5

Layer 1 3 Layers of CF 3 Layers of CF 2 Layers of CF 2 Layers of CF 2 Layers of CF

Layer 2 0.25in Divinycell Polystyrene Polystyrene 0.125in Divinycell 0.25in Divinycell

Layer 3 Polystyrene 3 Layers of CF 2 Layers of CF Polystyrene Polystyrene

Layer 4 0.25in Divinycell 0.125in Divinycell 0.25in Divinycell

Layer 5 3 Layers of CF 2 Layers of CF 2 Layers of CF

Table 1: Composite Sandwich Panel Structures

3.1.5 Accelerometers

One of the design requirements is the ability to detect and measure potentially catastrophic

vibrational disturbances in the wing. This is accomplished by means of 2, 3-axis accelerometers

placed at tip of the wing, one at the leading edge, and one at the trailing edge. The effects of

torsional and bending vibration on the accelerometers can be each isolated by making use of the

superposition principle as each effect acts independently. Appendix V – Example of

Accelerometer Input / Output - depicts an example of a likely input to the accelerometer, and the

output after employing the superpositioning principle. This example demonstrates clearly how

the accelerometers can be used together to obtain the torsional and bending vibration in real time.

3.2 Materials and Apparatus

3.2.1 Apparatus

For the purpose of conducting experiments, a 38.1 [mm] diameter stainless steel bar was rigidly

supported to a fixed steel slat table using a cast-iron bench vise. This arrangement formed a

stable support member representing the support spar used on the vertical stabilizer. A checker

pattern backdrop was placed behind the set-up to help record deflection characteristics, where

each square in the pattern measured 6.35 [mm]. The loading of the beams was captured in high

definition at a rate of 60 frames per second. For the loading procedure, 2.27kg weights were

used and added as indicated in Appendix IV – Beam Loading Positions. While this limited the

resolution of the experiment, it was adequate in providing the experimental data needed to make

30

strength estimates, design decisions and ensure design requirements were being met. To make

many of the composite parts, a vacuum pump and necessary accessories were provided by the

University of Toronto. A precision scale was used throughout the construction process to record

all material masses, as well as to properly mix the epoxy resin system to the correct ratio. A

digital caliper with a 0.01[mm] resolution was used for taking measurements, as well as for

measuring the deflection during a three-point bend test. Finally, a simple hot wire cutter was

constructed for cutting and shaping the polystyrene parts.

3.2.2 Materials Used

1) BGF 5.7oz plain weave Carbon Fiber 3K-70-P (Composites Canada Product CA057-50)

2) PR2032 Aeropoxy Resin with PH3665 Hardener27

Combined Carbon Fiber/Epoxy Properties

Density [g/mm^3] (2 Layer): 0.00165 (measured)

Density [g/mm^3] (3 Layer): 0.00215 (measured)

Tensile Strength [MPa]: 310

Tensile Modulus [GPa]: 70

3) 3.175 and 6.35 [mm] Divinycell H80 Structural Foam Sheet28

Density [g/mm^3]: 8.00E-05

Tensile Strength [MPa]: 2.5

Compressive Strength [MPa]: 1.4

Shear Strength [MPa]: 1.15

Tensile Modulus [MPa]: 95

Shear Modulus [MPa]: 27

4) 609.6x2438.4x50.8[mm] Closed-Cell Extruded polystyrene (STYROFOAM LT)29

27 Provided by AEROPOXY - See

Appendix IX – Material Data Sheets, AEROPOXY Data Sheet

28 Provided by DIAB Group - See

Appendix IX – Material Data Sheets, DIAB Group Data Sheet

29 Provided by DOW Industries - See

31

Density [g/mm^3]: 2.80E-05 ( 2.78E-05 measured)

Tensile Strength [kPa]: 450

Compressive Strength [kPa] 250

Shear Strength [kPa]: 250

Shear Modulus [MPa]: 7.0

Tensile Modulus [MPa]: 12-15

Maximum use Temperature [°C]: 74

5) Airtech Econostitch Peel Ply (Composites Canada Product VAPP-ES)

6) Airtech Airweave N4 Breather Material (Composites Canada Product VABR-N4)

7) Airtech Vacuumed Bagging Film (Composites Canada Product VABF-WL6400-60)

8) Airtech AT-200Y Sealant Tape (Composites Canada Product VATP-AT200Y)

9) Gorilla glue (Polystyrene adhesive)

3.3 Preliminary Experimentation

3.3.1 Base Support Experiment

The objective of the base support experiment was to study the loading capacities and failure

modes of the two composite beam base structures and test the validity of the composite base

structure design concept. The critical applied load was defined as the load which first caused

severe deformation or crack propagation when the beam was loaded in a cantilever beam

orientation. It should be noted that in this experiment the loading capacity of the beam itself was

not relevant as only the base structure was of interest to the end result.

Both of the proposed base structures were designed to carry the expected maximum load

(including a safety factor of 1.50) of 450[N] evenly distributed across the span of the beam. The

second design has additional support members to help prevent the occurrence of internal

buckling of the main load bearing support walls. Both composite base structures were expected

to fail due to buckling under extreme loading.

Appendix IX – Material Data Sheets, DOW Industries Data Sheet

32

3.3.1.1 Procedure

The process used to create the prototype base support is described below. For simplicity, a

rectangular cross section was used as opposed to an airfoil for this stage of the experimentation.

1) The initial step was to construct the core of the base support as shown in Figure 27 (a-e).

(a) A sheet of 6.35 [mm] Divinycell H80 was cut and epoxied together to form a

76.2 49.38 203.2 [mm] solid block of Divinycell. (b) A 40.54 [mm] diameter hole was

then drilled into the block for the addition of the support spar channel. (c) The support

spar channel, measuring 38.1 [mm] inner diameter and 40.54 [mm] outer diameter, was

made from six layers of carbon fiber and inserted into the Divinycell core. Finally, four

layers of carbon fiber were added surrounding the Divinycell core and carbon fiber pipe.

(d) At this stage, a vacuum bagging process was used to ensure the carbon fiber was fully

impregnated and completely bonded to the core. (e) This completed the base structure.

(a) (b) (c)

(d) (e)

Figure 27: Construction of First Support Base

The second base support was made in a similar fashion, however, during the first step,

four layers of carbon fiber were added into three locations along the length of the base

support as shown in Figure 28 (a) and (b), before and after the finishing cuts and insertion of

the carbon fiber support spar channel.

33

(a) (b)

Figure 28: Construction of Second Support Base

2) Next the base supports were adhered to composite sandwich panels consisting of two

layers of 6.35 [mm] Divinycell H80 with a polystyrene foam core as shown in Figure 29

(a-c). Care was taken to avoid any imperfections along this joint that would later cause

stress concentrations in the carbon fiber shell. These structures were then surrounded by

three layers of carbon fiber and vacuum bagged once again.

(a) (b)

(c)

Figure 29: Addition of composite Sandwich Panel

3) Once fully cured, the bases were each tested to failure in a cantilever beam type

configuration. This was performed by connecting the beams to the fixed steel bar

representing the support spar used on the vertical stabilizer. The beams were then

uniformly loaded until failure, as shown in Figures 30-33 The failure modes were then

Support

walls

Support

walls

34

analyzed and the failure stresses were calculated. The results from this experiment were

then used to select a base support design.

3.3.1.2 Base 1

Figure 30: Base 1 - Unloaded

Figure 31: Base 1 - Fully Loaded Prior to Failure

3.3.1.3 Base 2

Figure 32: Base 2 - Unloaded

Beam loaded in cantilever type configuration

Fixed steel bar representing the support spar

Internal Failure of the Base Structure



35

Figure 33: Base 2 - Fully Loaded Prior to Beam Failure

3.3.2 Composite Sandwich Panel Experiment

The primary purpose of the Composite Sandwich Panel experiment was to determine the critical

strength of five varying designs of cantilevered composite sandwich panel beams. The stiffness,

maximum deflection before failure, and particular failure modes were also documented. Each

beam was made to the same overall dimensions and loaded in a similar fashion. The critical

applied load was defined as: the load which first caused severe local or global buckling of the

structure when loaded with a uniformly distributed load in a cantilever beam orientation. The

composition of each of the composite sandwich panels are depicted in Figures 34- 38.

Figure 34: Composition of Beam 1

No signs of failure

36




37


3.3.2.1 Procedure

The Procedure for making each of the beams followed essentially the same process. Again, a

rectangular cross-section was used at this stage of the experimentation for simplicity. By using a

rectangular cross-section, the area of stress concentration was easily predictable and straight-

forward to calculate. As well, the beams were simpler to construct and still provided the buckling

strength of the panel. The following procedure was used to construct the beams:

1) The first step was to cut the polystyrene core to the required dimensions. This was done

using a hotwire foam cutting tool. Each beam was cut to approximately 203.2 [mm] wide

by 1105 [mm] long and was cut to the required thickness such that the end result would

measure 50.8 [mm] in thickness30

.

2) If the sandwich panel core contained of layers of either 3.18 [mm] or 6.35 [mm] thick

Divinycell core material, the Divinycell was cut to the same length and width of the beam

and, adhered using a polystyrene adhesive.

30 Final measured dimensions can be found in Appendix VI – Experimental Results: Beam Bending Study,

Summary of Experimental Results

38

3) The sandwich panel core was then adhered to the previously constructed base structure.

Similar to the previous experiment, care was taken to ensure a smooth transition from the

core to the base structure to avoid stress concentrations in the carbon fiber shell.

4) The required numbers of carbon fiber layers were then added to the upper and lower

surfaces of the core material, and to the base structure. The fiber ply orientation was

0/90° along the axis of the beam. These carbon fiber layers were then impregnated with

an epoxy resin system and vacuum bagged until fully cured.

5) The modulus of elasticity of each composite sandwich panel was approximated via a

three-point bend test and compared with analytical values. This test involved simply

supporting each beam across a 787.4 [mm] span, loading the center of the beam with

weights and measuring the deflection of the beam using a digital caliper as shown in

Figure 39. The results of this study can be found in Appendix VI – Experimental Results:

Beam Bending Study.

Figure 39: Experimental Set-Up for Testing Elasticity

6) The composite sandwich panels were then each tested to failure in a cantilever beam type

configuration. This was performed by connecting the beams to the fixed steel bar

representing the support spar, similar to the base structure testing. The composite

sandwich panels were then uniformly loaded until failure. The failure modes were

analyzed and the failure stresses were calculated. The results from this experiment were

then used to select a composite sandwich panel for use in the wing design.

Vernier calipers to

measure deflection

when loaded

39

3.4 Design Validation and Experimentation

From the previous experimentation, the configuration of the support base and composite

sandwich panel was determined. A prototype of the structural components of the wings were

constructed and tested based on these results to experimentally verify the design. Complete

design details and specifications are provided in Appendix I – Design Details and

Documentation.

3.4.1.1 Procedure

The first step in constructing the prototype was to create a replica of the current vertical tail

section of the SeaStryder600. This was constructed using rigid foam insulation to form the basic

shape. A fibreglass matt outer layer was then applied in order to provide the required structural

rigidity. The base structures were then constructed to precisely match the contours of the vertical

tail section and the remaining wing sections were completed and tested. The procedure is

outlined below.

1) The original computer aided design (CAD) data of the vertical tail section was used to

create cross-sectional templates that were then used to cut out 1.5 inch layers of the

section, to build up the foam core mold of the vertical tail section as shown in Figure 40

(a-b).

(a) (b)

Figure 40: Vertical Stabilizer Initial Core Templates

2) This mold was then cut using a hotwire foam cutting tool, sanded to a smooth finish, and

clearance cuts were made. This process can be seen in Figure 41 (a-d). (a) The cross-

sectional layers were precisely cut to form a near finished airfoil shape. (b) A template of

the span-wise clearance holes was made and used to cut out the segments which later

40

form the support bar channel, fastening locations and electrical conduits. (c-d) Lastly, the

1.5 inch sections were permanently adhered together using the polystyrene adhesive.

(a) (b)

(c) (d)

Figure 41: Finishing the Vertical Stabilizer Core

3) Once cured, the entire structure was surrounded by six layers of fibreglass matt and

vacuum bagged as shown in Figure 42 (a)-(b). (c) A composite tubular support channel

was then added and the whole structure was again sanded to a smooth finish. (d) Finally,

a white gel coat was added as the top layer, completing the vertical tail section replica.

(a) (b)

(c) (d)

Figure 42: Finalizing the Vertical Stabilizer Replica

41

4) The base structures were the next components to be built, as shown in Figure 43 (a-c). (a)

Aluminum templates of the airfoil structure adjacent to the base structures were made and

used to form the foam core structure of the base using a foam hotwire cutting tool. (b)

While the designed called for the use of Divinycell as the base core material, the

manufacturing methods available were unable to cut the Divinycell, and therefore

polystyrene had to be used in its place. (c) Four layers of carbon fiber were added as

required into four locations along the length of the beam providing the additional support

walls and one finished edge.

(a) (b) (c)

Figure 43: Wing Base Support

5) Once the layers were added the remaining base structure was completed as depicted in

Figure 44 (a-c). (a) The support bar channel, fastening hardware cut-out and electrical

conduits holes were drilled. To form the complex freeform surface adjacent to the vertical

stabilizer, the base structures were connected to the vertical stabilizer and material was

gradually cut off in 3.18 [mm] incremental cuts using a hotwire. (b) The hotwire was

offset 3.18 [mm] from the surface of the vertical stabilizer and slid along the support base

joint, creating a very close fit between the surface of the vertical stabilizer and the base

support. (c) The remaining unfinished foam sides were then surrounded by four

additional layers of carbon fiber and the parts were vacuum bagged until cured.

(a) (b) (c)

Figure 44: Completing the Base Support

42

6) Next, using a similar procedure used to create the base cores, the polystyrene foam was

cut to form the core material and the wing structures were completed. The process is

depicted in Figure 45 (a-e). (a) Again, aluminum templates were made and used to guide

the hotwire cutting tool through the polystyrene. (b) The cores were then bonded together

using a polystyrene adhesive and the whole structure was sanded to a smooth finish,

particularly across the bonded joints. (c-d) Finally, two layers of carbon fiber were added

to the entire structure to form the load bearing shell. The first layer was oriented at 0/90°

to the vertical stabilizer centerline, the second at ±45°. (e) This was again vacuum bagged

until fully cure.

(a) (b) (c)

(d) (e)

Figure 45: Completing the Wing

7) Each wing was then tested to failure, again in a cantilever beam type configuration

similar to the composite sandwich panel testing. The wings were uniformly loaded until

failure. The failure modes were then analyzed and the failure stresses were calculated.

3.4.1.2 Construction Complications

During the construction processes, some complications arose. The most unfavorable

complication arose while the Divinycell was being cut for the base structures. When the base

structure was constructed for the beams, the Divinycell was cut to a rectangular shape using a

band saw. For the wing base structure, the Divinycell needed to be cut into a specific airfoil

43

shape, and therefore a hotwire was the only method available to create such a free-form shape.

When this was attempted the hotwire was not able to cut the Divinycell fast enough due to a lack

of power, and caused the surface to burn and become very irregularly shaped as seen in Figure

46. Polystyrene was used in place of the Divinycell as it is easily cut using a hotwire as seen

earlier in Figure 43.

Figure 46: Divinycell Hotwire Cutting Failure

However, polystyrene does not provide the structural stiffness necessary to prevent internal

buckling of the base structure. Therefore, as expected, during the wing experimentation the base

structure began to fail before the catastrophic failure of the wing structure beyond the base. This

initial failure did not affect the final results relating to the failure of the composite sandwich

panel beyond the base structure.

Another complication to the base structure arose when the support bar channels were drilled into

the polystyrene cores. Unintentionally, the holes were cut oversize by 4.22 [mm]. The design

called for a 40.23 [mm] diameter hole, however, 44.45 [mm] diameter holes were cut. To correct

this issue, the carbon fiber pipe shown in part drawing 003-000 (designed to be inserted into this

hole) was modified from a six layer carbon fiber pipe to a 18 layer pipe to compensate for the

overcut.

Divinycell Irregular

Surface Burning

44

Chapter 4 Results and Discussion

4 Results and Discussion

4.1 Base Support Experiment

Two base support experiments were preformed. The first experiment tested the simple base

support design, a Divinycell core wrapped with four layers of carbon fiber. The weight of this

structure was 155g. While it was capable of supporting the required loading condition (a load

and moment of 450 [N] and 224 [Nm], respectively), before a complete structural failure, at a

load of 423 [N], the structure began to creak and shift indicating that an internal failure was

beginning. This creaking continued and eventually signs of internal failure could be seen as the

base shifted on the support spar. Catastrophic failure eventually occurred at a load and moment

of 778 [N] and 428 [Nm], respectively. While unloading the beam, it was noted that even after

catastrophic failure the beam was still able to carry a load and moment of 379 [N] and 199 [Nm]

sustainably. After sectioning the beam it was apparent that the failure was caused by the

separation of the support channel from the thin carbon fiber base wall. Figure 47 shows this

cross-sectional cut revealing the failure mode.

Figure 47: Base Failure

The second experiment tested the base support that contained the additional support walls. This

structure weighed 184g. It was fully capable of supporting the required loads, and no signs of

internal failure were ever apparent. The maximum loading condition placed on this base support

Separation of

Support channel Delamination of

Divinycell

Composite

Sandwich Panel

Spar Support

Channel

45

was a load of 1157[N] and a moment of 595[Nm]. This structure was never brought to failure as

no beam was ever made that could exceed this loading condition without failing itself.

4.2 Composite Sandwich Panel Experiment

During this experiment, five beams were brought to failure in a cantilever beam type

configuration. This was to determine failure modes, material properties and most relevantly,

maximum buckling strength. The beams were constructed such that the maximum load required

to be supported by the wing, 450[N] and 224[Nm], was equivalent to a maximum loading in the

rectangular beam of 365[N] and 188[N] due to the reduced rectangular beam cross section.

Therefore, this was the minimum desired loading capability of each beam. A summary of the

results can be found in Appendix VI – Experimental Results: Beam Bending Study.

4.2.1 Beam 1

The first composite panel constructed and tested consisted of a polystyrene core, followed by an

upper and lower layer of ¼” Divinycell H80, plus three layers of carbon fiber on the upper and

lower surfaces. This beam contained the initial base support and therefore was only loaded until

the catastrophic failure of the base structure had occurred. The maximum load this beam ever

sustained was immediately prior to the failure of the base support at a load and moment of 778

[N] and 428 [Nm], respectively. The local spanwise deflection, which would result in the

curvature of the hinge line, was difficult to measure due to the gradual base failure. While the

base support did allow the entire assembly to deflect, a close inspection the frames captured

during the experiment did not reveal any significant signs of spanwise curvature/deflection.

However, a finite element analysis estimated that the maximum deflection of the beam over the

span of 1.27 [m] should be approximately 13.7 [mm]. This beam configuration was not tested

again as it had already significantly exceeded the loading requirement of 365 [N] and 188 [Nm]

at the base end. The maximum facing stress calculated was 71.5 [MPa] based on Sandwich Beam

Theory. This value agreed closely with the finite element analysis results which predicted a

stress of 72.1 [MPa] under the same loading conditions where the base meets the composite

sandwich structure as shown in Figure 48.

The mass per unit length of this beam, excluding the base structure, was measured to be 811

[grams/meter].

46

Figure 48: CAD Analysis of Beam 1

4.2.2 Beam 2

The second composite panel constructed and tested consisted of a polystyrene core with three

layers of carbon fiber on the upper and lower surfaces. This beam, and all subsequent beams

tested, contained the base with added support walls. This beam was able to support a maximum

load and moment of 979 [N] and 504 [Nm], respectively, and the failure was due to local

buckling adjacent to the base joint as expected. This sudden failure is shown in Figure 49.

Figure 49: Failure of Beam 2

Start of

Buckling

Failure


Fixed steel bar Base Support Composite

Sandwich Panel

47

The maximum facing stress calculated at this point was 87.7 [MPa] which again agreed with the

finite element analysis preformed on this beam which predicted a stress of 91.7 [MPa] as shown

in Figure 50.


A deflection of 14.7 [mm] was also predicted through finite element analysis and was again

virtually unnoticeable experimentally. Figure 51 shows the unloaded (top image) and loaded

(bottom image) beam during the experiment. Is it clear from this figure how little deflection

actually occurred. The measured mass per unit length of this beam was 614 [grams/meter],

making it the second lightest beam constructed but notably one of the strongest.

Figure 51: Beam 2 - Unloaded (Top) and Fully Loaded (Bottom)

4.2.3 Beam 3

The third composite panel constructed and tested consisted of a polystyrene core followed by

two layers of carbon fiber on the upper and lower surfaces. This beam failed due to localized

buckling on the bottom surface. Since the upper surface was still free of defects, the beam was

able to be tested again simply by turning it over. During the first test the beam was able to

Nearly unobservable deflection

48

support a maximum load and moment of 356 [N] and 183 [Nm], and during the second, a load

and moment of 445[N] and 229 [Nm]. The failure of this beam was again consistently due to

local buckling adjacent to the base joint as expected. The maximum facing stress calculated at

this point was 50.2 [MPa] for the first test, and 62.8 [MPa] for the second test, which again

agreed with the finite element analysis preformed on this beam which predicted stresses of 49.8

[MPa] and 62.2 [MPa] as shown in Figure 52. The maximum deflection during these loads was

estimated to be 5.9 [mm] and 7.4 [mm], again found only computationally. The measured mass

per unit length of this beam was 500 [grams/meter], making it the lightest of all the beams tested.


49

4.2.4 Beam 4

The fourth composite panel constructed and tested consisted of a polystyrene core, followed by

an upper and lower layer of 1/8” Divinycell H80 and two layers of carbon fiber on the upper and

lower surfaces. This beam also failed due to localized buckling on the bottom surface, and

therefore was also able to be tested twice; however, the beam also contained a critical flaw.

During the construction of the beam, a small step developed between base support and the

composite panel. This flaw caused the beam to fail prematurely, however, the results were still

analyzed. During the first test, the beam was able to support a maximum load and moment of

275 [N] and 534 [Nm], and failed directly on the step as expected. During the second test, the

beam was able to support a load and moment of 623[N] and 320 [Nm]. During the second test

the beam failed in tension, where it had previously failed in compression on the step flaw as

shown in Figure 53.

Figure 53: Upper Surface Failure of beam 4 in Tension

The maximum facing stresses calculated at the failure point were 70.6 [MPa] for the first test,

and 82.3 [MPa] for the second test, which was again consistent with the finite element analysis

shown in Figure 54. Both of these values are noticeably low relative to the beam configuration

and are a direct result of the initial step flaw. The measured mass per unit length of this beam

was 653 [grams/meter].

Failure in Tension


Fixed steel bar Base Support

Composite

Sandwich Panel

50


4.2.5 Beam 5

The fifth and final composite panel tested consisted of a polystyrene core, followed by an upper

and lower layer of ¼” Divinycell H80 and two layers of carbon fiber on the upper and lower

surfaces. This beam was also tested twice; however it failed on both sides due to severe

delamination of the Divinycell under extreme loading as shown in Figure 55.

51

Figure 55: Delamination Failure of Beam 5

During the first test the beam was able to support a maximum load and moment of 801 [N] and

412 [Nm] and during the second a load and moment of 979[N] and 504 [Nm]. The maximum

facing stresses calculated at this point were 96.9 [MPa] for the first test, and 118.4 [MPa] for the

second test, which again agreed with the finite element analysis preformed on this beam which

predicted stresses of 97.4 [MPa] and 119.1 [MPa], respectively, as shown in Figure 56.

Delamination of Divinycell

from polystyrene core

52


The maximum deflection during these loads was estimated to be 12.6 [mm] and 15.4 [mm]. The

experimental results were again unable to measure this small cantilever tip deflection. The

images in Figure 57 show the unloaded (top image) and loaded (bottom image) beam during this

experiment. Lastly, the measured mass per unit length of this beam was 721 [grams/meter].

(a)

(b)

Figure 57: Beam 5 - Unloaded (a) and Fully Loaded (b)


At this stage, two configurations of the design have been fully developed for industry use and are

provided in Appendix I – Design Details and Documentation. Two structural prototypes were

constructed from these plans and tested to further verify the design.

53

4.3.1 Wing 1

Two identical wings of the first configuration were constructed and tested, each consisting of the

second base support and the third sandwich panel core structure. The completed wings each

weighed approximately 890 grams, resulting in an estimated complete wing assembly weight of

1694 grams, a 25% reduction from the current industry design. This structure primarily consists

of a polystyrene core surrounded by two layers of carbon fiber. Complete detailed specifications,

including a weight breakdown of the complete wing assembly can be found in Appendix I –

Design Details and Documentation, Configuration 1. Similar to the beam experiments, each wing

was able to be tested twice, which provided a total of four experimental test results. Based on the

previous results of the sandwich panel core experiment, a finite element analysis of the wing

predicted the failure of the wing to occur at a stress of 62.8 [MPa], similar to the third beam

which contained the same structure. This equates to a uniform loading of 549 [N]. Three of the

wing failures occurred at a uniform loading of 556 [N], and the fourth at 534 [N]. The finite

element analysis shown in Figure 58 was performed to evaluate the stresses generated prior to

failure due to the uniform loading of 556 [N].

54

Figure 58: CAD Analysis of Wing 1

The result, 63.2 [MPa], was consistent with the predicted value of 62.8 [MPa]. These results

confirmed repeatedly that the design requirements have been met as the minimum uniform load

requirement was 450 [N]. Lastly, the maximum local spanwise deflection in the wing due to

loading was 2.0 [mm], well within the allowable range of 12.7 [mm] to not impede the elevator

control.

4.3.2 Wing 2

Based on the accuracy and consistency of the prior finite element analysis, there was a high

degree of confidence in the procedure used to generate the finite element analysis results.

Therefore, in the interest of cost saving, the second configuration was not constructed, but

simply modeled based on the results of the sandwich panel core experiment similar to the

previous wing validation experiment. The estimated weight of this wing was 1854 grams, an

18% reduction from the current industry design. The complete detailed specifications and weight

breakdown of this wing assembly can also be found in Appendix I – Design Details and

55

Documentation, Configuration 2. The finite element analysis of this wing configuration, shown

in Figure 59, predicted the failure of the wing to occur at a stress of 87.7 [MPa], similar to the

second beam, which equates to a uniform loading of 1165 [N] on the wing.

Figure 59: CAD Analysis of Wing 2

The maximum local spanwise deflection estimated at this load was 2.6 [mm], also still well

within the allowable range. Given that the minimum load to be uniformly supported was 450

[N], this design exceeds the design requirement by 159% and continues to provide a significant

level of weight reduction. Although the flutter characteristics of this wing were not calculated,

the increased stiffness of this configuration may also be beneficial as it much less prone to this

phenomenon.

56

Chapter 5 Conclusions and Recommendations

5 Conclusions and Recommendations

5.1 Base Support Experiment

The first support base design would have been the preferred design, had it been reasonably close

in strength to the second, because it is much less complex to manufacture. However, the addition

of support walls made a significant increase in the base strength. Due to the improved rigidity

and strength of the second base structure, it was determined that similar support walls were

necessary and would be incorporated into the design of the wing base structure as well.

5.2 Composite Sandwich Panel Experiment

The inclusion of a Divinycell separation layer between the carbon fiber and the polystyrene

dramatically increased the buckling strength of the carbon fiber, however, it also significantly

added to the overall weight. It was found that by adding an additional layer of carbon fiber

instead, a larger increase in the buckling strength could be achieved with a much lower increase

in weight. The additional layer of carbon fiber also increased the stiffness and toughness of the

structure, whereas the Divinycell layer did not. This comparison is graphically illustrated in

Figure 60 by comparing beams 2 and 5.

Figure 60: Graphical Results of Beam Experiment

Beam 1

Beam 2

Beam 3

Beam 4

Beam 5

400

500

600

700

800

900

1000

300 400 500 600 700 800 900 1000 1100

Mas

s [g

ram

s/m

ete

r]

Average Sustainable Load [N]

Beam Strength vs Beam Mass

57

It is clear that beam 2, which contained three layers of carbon fiber, was both stronger and lighter

then beam 5, which contained two layers of carbon fiber along with a Divinycell separation

layer. One reason for this increased weight is the open cell structure of the Divinycell foam.

During the manufacturing process, the Divinycell foam absorbs excess resin from the carbon

fiber and cannot be removed. Even if this absorption could be avoided, the weight of the

Divinycell alone is still greater than the weight of an additional layer of carbon fiber.

While all the beams constructed exceeded the loading requirements, the most advantageous

structural properties were found in either beam 2 or beam 3. The third beam exceeded the

loading requirements by 9.8% on average, and was the lightest beam weighing only 500

[grams/meter]. The second beam weighed 614 [grams/meter], making it the second lightest.

However, it was the strongest and most rigid beam constructed, exceeding the loading

requirements by 168%. Due to the increased loading capacity, both core arrangements are being

proposed as varying configurations, each with unique advantages. The first arrangement provides

the lightest possible configuration which still meets the loading requirements. The second

arrangement still provides a significant weight reduction, while simultaneously providing a

dramatic increase in stiffness and loading capabilities, and therefore an increased safety factor in

the final design.


Based on the results of the prototype experimentation, two configurations of the design have

been developed and proposed, the first consisting of two layers of carbon fiber, and the second

consisting of three layers, both directly surrounding a polystyrene core. The first configuration

meets all the necessary design requirements with a weight reduction of 25% and a Factor of

Safety (FOS) of 1.85. The second configuration provides a weight reduction of 18.0%, however,

it also offers improved rigidity and an increased loading capacity of 159%. The Factor of Safety

(FOS) for this design becomes 3.88. Both designs incorporate the improved servo housing

assembly and provide sufficient rigidity to allow for complete motion of the elevator assembly

throughout all operational conditions. Both designs are fully capable of supporting an additional

winglet if necessary. There is high degree of confidence in these results as both wing

configurations were verified experimentally as well as by means of finite element analysis. The

beam and wing failures were both highly consistent with the finite element analysis estimates,

58

and as well, the failure mechanisms and maximum loads were highly consistent throughout the

experimentation, particularly notable given that they are composite structures.

5.4 Scientific contributions

The design developed and proven in this study provides the foundation for the development of a

truly spar-less wing design. This allows for a significant reduction in weight, cost, and

manufacturing time by eliminating all traditional rib and spar components. By utilizing current

methods of reducing surface delamination and component buckling, a single sandwich panel is

able to support the loads required for this component. A critical aspect of this design incorporates

the use an extremely rigid lightweight base structure located at the wing root. This base structure

replaces the spar in traditional rib and spar configuration and transfers the loads supported by the

wing skin to a single support channel where it fastens to the fuselage as a traditional wing would.

5.5 Future Recommendations

As a continuation of this research, it is recommended that further testing be performed to ensure

a safe and effective design. Primarily, a wind tunnel test of the wing structure should be

performed throughout all operation conditions and elevator positions. This should provide

conclusive results with regards to the structural integrity, flap control mechanism, and as well,

test for any possible signs of flutter. Even though composites generally have excellent fatigue

resistance properties, a fatigue test should be performed to ensure the design is fit for the extreme

cyclical loading it is expected to encounter. Lastly, it should be noted that the failure of a single

servo, particularly in a limit lockout position, could be catastrophic to the elevator control. This

scenario should be reviewed and the division of the elevator into independently controlled

partitions should be considered.

59

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60

Plantema, F, J., 1966, Sandwich Construction: The Bending and Buckling of Sandwich

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Shan, B, and A Pelegri. Approximate Analysis of the Buckling Behavior of Composites

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61

Appendix I – Design Details and Documentation

89

Appendix II – NACA 0009/0012 Profiles

NACA Profiles from Data Files

Equation Driven NACA Profiles (Used in CAD Model)

Cord Position

NACA 0012 Data File

NACA 0009 Data File

Equation Based NACA 0012*


1.009 0.000 0.000 0.000 0.000 1.000 0.000 0.000 0.001 0.001 0.989 0.002 0.001 0.003 0.002 0.957 0.006 0.004 0.007 0.005 0.905 0.013 0.010 0.014 0.010 0.835 0.021 0.016 0.022 0.017 0.750 0.031 0.023 0.032 0.024 0.655 0.040 0.030 0.041 0.031 0.552 0.049 0.036 0.049 0.037 0.448 0.055 0.042 0.056 0.042 0.345 0.059 0.044 0.060 0.045 0.250 0.059 0.044 0.059 0.045 0.165 0.055 0.041 0.055 0.041 0.095 0.046 0.034 0.046 0.035 0.043 0.033 0.025 0.033 0.025 0.011 0.018 0.013 0.018 0.013 0.000 0.000 0.000 0.000 0.000 0.011 -0.018 -0.013 -0.018 -0.013 0.043 -0.033 -0.025 -0.033 -0.025 0.095 -0.046 -0.034 -0.046 -0.035 0.165 -0.055 -0.041 -0.055 -0.041 0.250 -0.059 -0.044 -0.059 -0.045 0.345 -0.059 -0.044 -0.060 -0.045 0.448 -0.055 -0.042 -0.056 -0.042 0.552 -0.049 -0.036 -0.049 -0.037 0.655 -0.040 -0.030 -0.041 -0.031 0.750 -0.031 -0.023 -0.032 -0.024 0.835 -0.021 -0.016 -0.022 -0.017 0.905 -0.013 -0.010 -0.014 -0.010 0.957 -0.006 -0.004 -0.007 -0.005 0.989 -0.002 -0.001 -0.003 -0.002 1.000 0.000 0.000 -0.001 -0.001 1.009 0.000 0.000 0.000 0.000 *NACA 0012 = (0.12/0.2)*L*(0.2969*(x/L)^0.5-0.126*(x/L)^1-0.3516*(x/L)^2+0.2843*(x/L)^3-0.1015*(x/L)^4)

*NACA 0009 = (0.09/0.2)*L*(0.2969*(x/L)^0.5-0.126*(x/L)^1-0.3516*(x/L)^2+0.2843*(x/L)^3-0.1015*(x/L)^4)

where L = 1 (Cord Length)

0.00

0.10

0.20

0.30

0.40

0.50

0.60

0.70

0.80

0.90

1.00

-0.10 0.00 0.10

Plot of Profiles

NACA 0012 Data File

NACA 0009 Data File



90

Appendix III – Hinge Moment Calculations and Characteristics

Geometry S1/E1 S1/E2 S1/E3 S2/E1 S2/E2 S2/E3 Item Description

L1 [in] 0.236 0.236 0.236 0.354 0.354 0.354 Servo Link

L2 [in] 1.624 1.614 1.692 1.545 1.604 1.673 Linkage Rod

L3 [in] 0.486 0.605 0.728 0.486 0.605 0.728 Control Horn Link

L4 [in] 1.624 1.614 1.692 1.545 1.604 1.673 Hinge/Servo

Theta2 [Deg] 17.38 17.38 17.38 17.38 17.38 17.38 Hinge/Servo

Table 2: 4 Bar Linkage Specifications

Figure 61: 4 Bar Linkage Schematic

Variable Servo Links

S1 and S2

Variable Control

Horn Links E1,

E2 and E3

Linkage Rod

Hinge/Servo

91

Figure 62: DS8711 Servo Torque (24.99 kg-cm)/Hinge Moment Characteristics

-20

0

20

40

60

80

100

120

-40 -30 -20 -10 0 10 20 30 40

Torq

ue

[kg

-cm

]

Flap Angle [Deg]

Est. Magnitude of Wind Torque (54[m/s])

Config: S1/E1

Config: S1/E2

Config: S1/E3

Config: S2/E1

Config: S2/E2

Config: S2/E3

92

Appendix IV – Beam Loading Positions

Figure 63: Beam 1 Loading Positions

Figure 64: Beams 2-5 Loading Positions

Beam

Beam

Centroid of each mass stack

(placed in center of beam at

this spanwise position).

Start of Composite Sandwich Panel

93

Figure 65: Wing Loading Positions

Wing

94

Appendix V – Example of Accelerometer Input / Output

Figure 66: Accelerometer Input Vibration Example

Figure 67: Output after Separation of Bending and Torsional Vibration

-0.15

-0.1

-0.05

0

0.05

0.1

0.15

0 0.25 0.5 0.75 1 1.25 1.5

Dis

pla

cem

en

t [m

]

Time [s]

Accelerometer 1

Accelerometer 2

-5.00

-4.00

-3.00

-2.00

-1.00

0.00

1.00

2.00

3.00

4.00

5.00

-0.15

-0.10

-0.05

0.00

0.05

0.10

0.15

0 0.25 0.5 0.75 1 1.25 1.5

An

gle

[ra

d]

Dis

pla

cem

en

t [m

]

Time [s]

Bending (Est)

Bending (Real)

Torsion (Est)

Torsion (Real)

95

Appendix VI – Experimental Results: Beam Bending Study

Beam 1 Experimental Results

Trial 1 – Unloaded Beam

Total Beam Mass: 1150g

Mass/Length: 811 [g/m]

Trial 1 – Maximum Loading Before Failure

Maximum Load: 778 [N]

Maximum Moment: 423 [N-m]

Figure 68: Beam 1 Cross-Section


Beam loaded in Cantilever type configuration

96


Trial 1 – Unloaded







97



Total Beam Mass: 767.4g






Beam Mass: 767.4g


Not Captured


Max Load: 445 [N]


Not Captured


98














99




Beam Mass: 1039g





100


Beam Mass: 1039g






10

1


Beam Specifications

Beam Dimensions [mm]

Layer Thicknesses [mm]

Beam Description Failure Mode

Width Length Height

CF Div Foam Div CF Total

1 3 Layer CF + 6.7mm Div. Base Support Failure

203 1270 49.0

0.61 6.7 34.3 6.7 0.61 49.0

2 3 Layer CF CF Buckling (Perfectly)

203 1181 51.7

0.56 0.0 50.6 0.0 0.56 51.7

3 2 Layer CF all round CF Buckling (Perfectly)

203 1181 51.2

0.36 0.0 50.4 0.0 0.36 51.2

3b* 2 Layer CF all round CF Buckling (Perfectly)

203 1181 51.2

0.36 0.0 50.4 0.0 0.36 51.2

4 2 Layer CF + 3.2mm Div. CF Buckling (On Joint Flaw)

203 1181 48.0

0.41 6.7 37.2 3.2 0.41 48.0

4b* 2 Layer CF + 6.7mm Div. In tension (Due to Trial 1)

203 1181 48.0

0.41 6.7 37.2 3.2 0.41 48.0

5 2 Layer CF + 6.7mm Div. Core Buckling (Delamination)

203 1181 52.3

0.41 6.7 38.0 6.7 0.41 52.3

5b* 2 Layer CF + 6.7mm Div. Core Buckling (Delamination)

203 1181 52.3

0.41 6.7 38.0 6.7 0.41 52.3

*recycled from previous test

Masses [g]

Sectional Properties

Stress Analysis (Max Values)

Base Block Foam Div CF Resin Glue Total

Mass

[g] Length [mm]

Failure

[kgs]

Moment

[Nm]

Force

[N]

Facing Stress

[MPa]

Core Shear

[KPa]

155 231 265 332 156 16 1155

783 965

79.4 423 778 71.5 160.4

184 316 0 305 75 0 879

624 1016

99.8 504 979 87.7 190.5

184 315 0 194 74 0 767

508 1016

36.3 183 356 50.2 69.4

184 315 0 194 74 0 767

508 1016

45.4 229 445 62.8 86.8

220 232 181 207 107 25 971

663 1016

54.4 275 534 70.5 111.4

220 232 181 207 107 25 971

663 1016

63.5 320 623 82.3 130.0

227 237 245 203 110 16 1039

733 1016

81.6 412 801 96.8 153.0

227 237 245 203 110 16 1039

733 1016

99.8 504 979 118.4 187.0

CAD Analysis (Max Values)

Disp. [mm] Facing Stress [MPa]

13.7 72.1

14.74 91.7

5.914 49.8

7.389 62.2

9.324 72.3

10.88 84.3

12.58 97.4

15.38 119.1

Table 3: Beam Study Results

102

Appendix VII – Experimental Results: Wing Study Wing 1 Experimental Results (Port Wing)


Beam Mass: 889g





Beam Mass: 889g




103

Wing 2 Experimental Results (Starboard Wing)


Beam Mass: 893g





Beam Mass: 893g




10

4


Beam Specifications

Dimensions [mm]

Layer Thicknesses [mm]

Beam Description Failure Mode

Width Length Height

CF Div Foam Div CF Total

Wing 1 2 Layer CF Buckling+Base

333 1092 N/A

0.4 0.0 36.8 0.0 0.4 N/A

Wing 1b 2 Layer CF Buckling+Base

333 1092 N/A

0.4 0.0 36.8 0.0 0.4 N/A

Wind 2 3 Layer CF CAD Only

333 1092 N/A

0.6 0.0 36.8 0.0 0.6 N/A

Masses [g]

Sectional Properties

Stress Analysis (Max Values)

Base Block Foam Div CF Resin Glue Total

Mass

[g] Length [mm]

Failure

[kgs]

Moment

[Nm]

Force

[N]

Facing Stress

[MPa]

Core Shear

[KPa]

210 276 0 285 116 2 889 N/A N/A

56.7 277 556 63.5 90.7

217 276 0 283 115 2 893 N/A N/A

56.7 277 556 63.5 90.7

217 276 0 372 115 2 982 N/A N/A

118.8 580 1165 84.7 190.1

CAD Analysis (Max Values)

Disp [mm] Stress [Mpa]

1.996 63.2

1.996 63.2

2.667 88

Table 4: Wing Study Results

105

105

Appendix VIII –Experimental Results: Stiffness Test

Mass Added [kgs]

Beam 1 Deflection

[mm]

Beam 2 Deflection

[mm]

Beam 3 Deflection

[mm]

Beam 4 Deflection

[mm]

Beam 4b Deflection

[mm]

Predicted CAD Deflection [mm]

0.0 0.00 0.00 0.00 0.00 0.00 0.00

11.3 0.41 0.46 0.48 0.56 0.58 0.43

34.0 1.14 1.52 1.52 1.60 1.83 1.30

56.7 1.78 2.46 2.29 2.41 2.87 2.17

79.4 2.41 2.97 3.02 3.20 3.45 3.04

Table 5: Deflection-Load Results

Figure 73: Deflection-Load Curve

Figure 74: CAD Deflection Study

0.00

0.50

1.00

1.50

2.00

2.50

3.00

3.50

0.0 10.0 20.0 30.0 40.0 50.0 60.0 70.0 80.0

De

fle

ctio

n [

mm

]

Load [kg]

Beam 1 Deflection [mm]




Beam 4b Deflection [mm]

Predicted CAD Deflection [mm]

106

106

Appendix IX – Material Data Sheets

107

107

108

108

109

109

Appendix X – Evaluation of Design Requirements

Minimum Load Requirement:

The previously calculated 6G load on the wing, provided by industry, resulted in a load and

moment of 240 [N] and 113 [Nm]. The new weight of the aircraft is expected to cause an

increased loading of 25%. As well, a safety factor of 1.5 has been requested. This resulted in an

increased minimum load requirement of 450 [N] and 224 [Nm] on each horizontal stabilizer.

Maximum Deflection Requirement:

The maximum allowable deflection is limited by a geometric condition whereby the flexible

hinge can no longer reach the deflected wing from the elevator. This is due to the fact that the

elevator, when in its limit position (+/- 25 degrees), has a significantly higher spanwise vertical

stiffness than the wing, and for all practical purposes, does not deflect, though the wing does.

This geometric condition is limited by the spacing between the elevator and wing, which

contains the flexible hinge. The previous and current design spacing is required to be 1.5875

[mm]. This arrangement can be seen in Figure 75 and Figure 76. Figure 75 assumes a separation

of 15.875 [mm], ten times the requirement, and is for illustrative purposes only, Figure 76 is of

the true 1.5875 [mm] separation and shows the maximum allowable deflection is 12.6996 [mm].

Figure 75: Geometric Limit of Deflection ( 10 for illustrative purposes)

Figure 76: True Geometric Limit of Deflection

Unbent Elevator Bent Wing

Wing surface before Bending

design optimization and verification of a horizontal ... · in addition, the low airfoil of the...

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