FYP FINAL

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Acad Year11/12

SOLAR UAV: Airframe And Landing Gear Design,

Fabrication, Testing And Evaluation (I)

PROJECT NO.

A186

SO

LA

R U

AV

: Airfram

e An

d L

and

ing G

ear Design

, Fab

rication, T

esting A

nd

Evalu

ation (I)

Chu Wei Xin

SCHOOL OF MECHANICAL AND AEROSPACE ENGINEERINGNANYANG TECHNOLOGICAL UNIVERSITY

Year 2011/2012

Title

Solar UAV - Airframe And Landing Gear Design, Fabrication, Testing And Evaluation (I)

SUBMITTED

BY

CHU WEI XIN

SCHOOL OF MECHANICAL AND AEROSPACE ENGINEERING

A final year project reportpresented to

Nanyang Technological Universityin partial fulfilment of the

requirements for theDegree of Bachelor of Engineering (Mechanical Engineering)

Nanyang Technological University

Year (2011/2012)

TABLE OF CONTENTS

Page

ABSTRACTI

ACKNOWLEDGEMENTii

LISTS OF FIGURESiii

LIST OF TABLESiv

CHAPTER ONE

INTRODUCTION1

CHAPTER TWO

Literature Review4

CHAPTER THREE

Conceptual Design

CHAPTER FOUR

ANSYS Analysis Results

CONCLUSIONS

FUTURE WORKS

REFERENCES

APPENDICES

ABSTRACTUnmanned Aerial Vehicles have seen an unprecedented growth in recent years in

both military as well as civilian application domain. This has increased the interest

and research in unmanned technology at academic level. This thesis shows the

conceptual design of a solar powered Unmanned Aerial Vehicle (UAV) that has a

wing span of 15 meters with a take-off mass of not more than 60 Kg. The conceptual

idea adopted is the concept of a powered glider. Gliders are designed to have the

minimum drag for any given amount of lift. The fuselage is a long and narrow

section, together with long and thin wings help to achieve minimum drag while

maintaining the lift. Any aircraft designed in this manner are able to have efficient

climbing rates and can glide long distances at a high speed with a minimum loss of

height in between.

AcknowledgmentsThe student would like to send out his heart-felt gratitude for the opportunity to work

on the Solar UAV - Airframe and Landing Gear Design, Fabrication, Testing And

Evaluation (I) project. The student would like to express heartfelt appreciation to his

supervisor A/P Liu Yong and Co-supervisor A/P Li Peifeng, for the guidance and

patience that they have shown in guiding the student, and their willingness to impart

their valuable experience, knowledge and skill to the student.

And also, a great thank you to the other professors who are in-charge of other

students, who painstakingly help to coordinate the work between the different teams.

List of FiguresFigure 1: High-Altitude Long Endurance (HALE) UAV, Zephyr

Figure 2: Solar Impulse

Figure 3: DARPHA Vulture Program

Figure 4: RQ-11 Raven

Figure 5: Predator

Figure 6: MQ-8B Fire Scout UAV

Figure 7: A160T HummingBird

Figure 8: Tethered Aerostat Radar System

Figure 9: Advanced Airship Flying Laboratory

Figure 10: Flight Control Components

Figure 11: Spars, booms and skin configuration

Figure 12: Spars, booms, skin and stringers configuration

Figure 13: Torsion box made of spars, stringers and skin

Figure 14: Lug/pin attachment for a fighter jet

Figure 15: Wing Root Triple Lug Joint

Figure 16: Various Configurations of Lug Joints

Figure 17: Double Shear Lug Design with Hollow Tube

Figure 18: Body Axis, Moments, Rates and Controls

Figure 19: Distributed versus concentrated forces

Figure 20: Aircraft Sizing Flow

Figure 21: Mission Profile

Figure 22: Proposed UAV Conceptual Configuration

Figure 23: NACA 23012 Profile

Figure 24: Ribs (Side View)

Figure 25: Ribs (Isometric View)

Figure 26: Spar (I-Beam – Side View)

Figure 27: Spar (I-Beam – Isometric View)

Figure 28: C-Channel (Side View)

Figure 29: C-Channel (Isometric View)

Figure 30: Leading Edge

Figure 31: Airfoil Skin with Stringers (Side View)

Figure 32: Airfoil Skin with Stringers (Isometric View)

Figure 33: Airfoil Assembly (Isometric View)

Figure 34: Airfoil Assembly (Internal Wing Structure)

Figure 35: Elliptical Loading

Figure 36: Equal flanged section and examples of sections with one axis of symmetry

Figure 32: Airfoil Skin Meshing

Figure 33: Airfoil Internal Structure Meshing

Figure 34: Airfoil Internal Structure Meshing (Zoom-In)

Figure 35: Stress Analysis

Figure 36: Deformation Analysis

Figure 37. I Beam Schematic

Figure 38 – 43. I Beam Construction

Figure 44. Test Section 1 Assembly

Figure 45. Test Section 1 with simulated solar panels

List of TablesTable 1: Targets / Milestones for Final Year Project

Table 2: Comparison of Joints

Table 3: Design Targets

Table 4: Proposed Configuration Parameters

Table 5: Dimensions of I-Beam

Table 6: Max Allowable Loading Values

Table 7: Derived Parameters of I-Beam

Table 8: Circular Spar Dimensions

Table 9: Volume and Stress Result Tabulation

Table 10: Single Spar Result Tabulation

Table 11: Front Spar Result Tabulation

Table 12: Rear Spar Result Tabulation

Table 13: Comparison of Results between I-Beam and Circular Spar

Table 14: Shear Stress Tabulation of Different Sections

Table 15: Mass Estimation of Airfoil

Chapter One Introduction

1.1 Background

An Unmanned Aerial Vehicle (UAV) is an aircraft without a human operator on

board. They are designed to carry out the various missions and come back so that it

can be reused. These aerial vehicles can be operated out of line sight and at altitudes

where they cannot be detected by people on the ground. They are being controlled by

an operator in a command and control center rather than on the aircraft itself.

The motivation behind this project is partially to design an UAV that can fulfill

mission requirements, and at the same time, be energy efficient. Commonly used fuel

for planes are jet fuel, causing a negative impact on the environment due to its

emissions. One of the methods is to eliminate the use of jet fuel and replacing the

energy source with solar energy. Solar powered aircraft will not only be successful,

but they are also a solution to the current environmental problem and will become a

solution for the future.

1.2 Objectives

The primary objective of this project to design a solar powered UAV that cruises at

8,000 meters while carrying out its mission requirements. It has to operate for a

continuous time frame of 24 hours and with a cruise speed of no more than 20km/h

Primary source of energy will be solar energy, this will be stored in batteries installed

in the airfoil structure. The wing is made up of a combination of several components,

mainly the rib-and-spar structure which is made of carbon-fiber reinforced

composites and covered by a solar panel embedded film with stringers running span

wise. All these individual components help to maintain the structural integrity of the

UAV throughout all the possible scenarios/loading conditions that the UAV may

encounter during the mission. The few main challenges in this project was mainly

how to integrate the different components together across different development

teams, while maintaining the simplicity of design and achieving the desired mission

requirements.

1.3 Scope

This report outlines the work in the design of a UAV. The project team is broken

down into several specialized teams (Structures, aerodynamics and propulsion, flight

control and avionics, solar energy). This report focuses on the structural design of the

UAV, mainly the airfoil structure, and has also been analyzed for its aerodynamics,

stresses, optimization, Computer Aided Drawing (CAD - SolidWorks) and Finite

Element Analysis (FEA).

The structures team comprises of 4 members. Melvin Chow Shun Jie will be dealing

primarily with the airframe configuration and design, Augustus Yip Bao Sheng will

be in-charge of the design of structural components, Chu Wei Xin main role is on the

SolidWorks design of the UAV and lastly Tan Shan Zhi main role is on the structural

testing of the individual components.

Table 1 in the next page highlights the main tasks that the student has to complete

and the datelines that he has set for himself.

Phase 1 : Initial Research and Fact Finding (7 weeks)Item Duration Comments

Understanding of mission requirements and formation of structure with Solar UAV Team

1 week (22 Aug – 28 Aug)

Completed

Literature Research on the structures of the UAV

3 weeks(29 Aug – 18 Sep)

Completed

1. To obtain the airfoil profile data from Aerodynamics Group

2. With the obtained data, Structures Group will come out with the internal structure of airfoil

3 weeks(19 Sep – 9 Oct)

Completed

Phase 2 : Airfoil Modelling (4 weeks)SolidWorks 3D Modelling (Airfoil) 4 weeks

(10 Oct – 6 Nov)Completed

Exam Preparation : 7 Nov – 1 DecPhase 3 : Airframe Modelling and Integration (8 weeks)

Consolidation of team data and modification of requirements (if any)Consolidation of interim report

1 week(2 Dec – 11 Dec)

Completed

Interim Report Submission 12 Dec 2011Design of airfoil section joints 1 weeks

(12 Dec 19 Dec)Completed

Building of the test sections 5 weeks (19 Dec – 29 Jan)

Completed

Phase 4 : Design of Assembly Jig (3 weeks)Design of test rig 3 weeks

(30 Jan – 20 Feb)Completed

Phase 5 : Conclusion and ImprovementsItem Duration Comments

Collation of results and drafting of final report

3 weeks(7 Feb – 18 Mar) – Concurrent with Design of Assembly Jig

Completed

Final Report Draft Submission 19 Mar 2012Improvements to final report with conclusions and suggested activities for follow on project

3 weeks(19 Mar – 8 Apr)

Completed

Final Report (unbound) 9 Apr 2012Exam Preparation : 10 Apr – 4 May

Preparation for oral presentation 5 May – 9 MayOral Presentation 7 – 9 May 2012Final Improvements to Final Report 3 weeks

(10 May – 30 May)Final Report Submission 31 May 2012

Table 1. Targets/Milestones for Final Year Project

Chapter Two Literature Review

2 Literature Review

This section discusses previous researches and developments on the topic of solar

powered UAV.

2.1 History of Solar Energy

Solar energy is not something new. In 7th century B.C., humans used a piece of glass

to focus the sun’s energy on a small area, resulting in a fire. The Greeks and Romans

used solar energy to light torches for religious purposes. They were so serious about

solar energy that they built glass houses to create the right conditions to grow plants.

Humans benefitted from solar power. But it wasn’t until 1776 that the first solar

collector was built. It was built by Horace de Saussare and his collector was shaped

like a cone that would boil ammonia that would then perform like refrigeration and

locomotion.

In 1861, Auguste Mouchout created a steam engine that was powered by solar

energy. However at that time, the technology was very costly and it could not be

possible to reproduce it or even maintained. This did not stop the efforts of many

scientist who knew that solar power can be harnessed and used in many different

ways. In 1880’s, the first light converting photovoltaic cells are built.

Since then, solar energy technologies have been harnessed to provide heating,

photovoltaic cells and thermal electricity. Using these technologies, humans are able

to convert solar energy into electrical energy to power our appliances.

2.2 First Solar Powered Flight and Subsequent Developments

The first solar powered flight was accomplished by Sunrise I, which was designed by

Astro Flight Inc., has a service ceiling of 6,100 meters. The historic flight was

powered only by sunlight, took place in 1974 at Fort Irwin, California. More than

one thousand solar cells are embedded on the wing structure was the sole source of

energy, producing around 450 watts of power. Sunrise I had a wingspan of 10 meters

and a gross weight of 10 kilograms. The structure of the Sunrise I was built mainly

out of spruce, balsa and maple. However Sunrise I was damaged during a windstorm.

A second model, Sunrise II was then developed. It weighed less and was able to have

more output power in comparison to Sunrise I. It has a total of 4,480 solar cells with

an output of 600 watts. It has a climb rate of 91 meters per minute and a service

ceiling of 23,000 meters.

Over the years, interest in solar powered flights increased. Competition also became

much more intense. In 1996, the Solar Solitude, built by aircraft enthusiast Dave

Beck, flew a record distance of 38.84 kilometers at an altitude of 1,283 meters.

As technology advances, there was also much interest in long endurance AUVs. An

example is the High-Altitude Long Endurance (HALE) UAV, Zephyr. It is

developed by a British Company, QinetiQ. It has a construction of carbon fiber

(CFRP), and uses the sunlight to charge the lithium-sulphur batteries during the day,

which powers the aircraft at night. The 2008 version of the Zephyr, with a wingspan

of 18 meters, weighed 30 kilograms. Whereas the 2010 version, with a wingspan of

22.5 meters, weighed 50 kilograms. The Zephyr has the record of an AUV with a

flight that lasted 336 hours and 22 minutes and results have been verified by the

Fédération Aéronautique Internationale (FAI). The altitude of flight was at 21,562

meters.

Figure 1. High-Altitude Long Endurance (HALE) UAV, Zephyr

Previous researches and developments were all for Unmanned Aerial Vehicles. In

2003, the first solar powered manned flight was completed by Solar Impulse. It flew

for duration of 26 hours only on solar energy, the first of its kind. The solar impulse

has a wingspan of 64 meters and weighs 1,588 kilograms. The aircraft was powered

by 11,000 solar cells, which powered four 7 kilowatt electric motors. It has a cruise

speed of 111 kilometers per hour.

http://en.wikipedia.org/wiki/F%C3%A9d%C3%A9ration_A%C3%A9ronautique_Internationale

Figure 2. Solar Impulse

Currently, the United States Defence Advance Research Project Agency (DARPA)

Vulture program, aims to develop an AUV that can enable a High-Altitude

Unmanned Aerial Vehicle to operate continuously for a period of five years. The

main idea of the Vulture program is to combine benefits of an aircraft and a satellite

into one system. The aircraft has highly efficient electric motors and propellers.

Together with a high aspect ratio, a 400 foot wing for increased solar power and

aerodynamic performance. DARPA will continue to work with Boeing to advance

this technology further and a functional prototype will be built.

Figure 3. DARPHA Vulture Program

2.3 Types of Unmanned Aerial Vehicles (AUV)

There are 3 main categories of UAVs. They are namely fixed-wing, rotary-wing, a

balloon or an airship. Hence one of the ways to classify different forms of AUV is by

their configuration.

1. Fixed-Wing UAV

Figure 4. RQ-11 Raven Figure 5. Predator

The most commonly used configuration is the fixed-wing UAV. It can be employed

for a vast range of applications. The reason why it is the most selected configuration

is that these are quite stable and does not require complicated control systems as

compared to other systems.

2. Rotary-Wing UAVs

Figure 6. MQ-8B Fire Scout Figure 7. A160T HummingBird

The advantages of Rotary-Wing UAVs allowed it to have excellent manoeuvrability

and its ability to hover. However, rotary-wing UAVs are relatively unstable and it

poses more challenges for the controller.

3. Airship UAVs

Figure 8. Tethered Aerostat Figure 9. Advanced Airship Radar System Flying Laboratory

2.4 Components of an UAV System

The UAV system consists of 3 main components. They include the UAV platform,

the payload as well as the ground command and control centre.

2.4.1 UAV Platform

The UAV platform consists of the following sub-components.

1. Airframe

The airframe is a very vital component of the UAV. It is the primary load bearing

structure and carries payloads that are necessary for the mission. Hence, the airframe

must be structurally reliable to sustain the various loads that is being applied to it

while on the ground or in flight. At the same time, the airframe must have the

required aerodynamic properties and be light enough.

2. Propulsion System

The UAV propulsion system is dependent on the mission requirements. An UAV

with a mission requirement of quick insertion and extraction from a warzone, it must

have sufficient speed and hence a more powerful propulsion system is required.

3. Flight Control

Figure 10. Flight Control Components

Since the UAV has no human operator onboard, it is really vital that the UAV is

equipped with reliable Global Positioning Systems, sensors and avoidance systems.

The complexity of the flight control system is again dependent on the mission

requirements.

2.4.2 Payload

The types of payload being carried on the UAV are dependent on the mission

requirements. An UAV with the mission requirement of doing spy missions would be

equipped with cameras, various sensors, radars and sensors.

2.4.3 Ground Command and Control Centre

The ground command and control centre is used to monitor the systems of the UAV

as well as to control the UAV. The ground command and control centre will have

avionics flight displays, navigation systems, position mapping systems as well as

system health monitoring systems. The communication is not just from the ground

control centre, but also from the UAV back to the control centre. Information such as

video, pictures can be transmitted back to the ground control centre.

2.5 Applications

Listed below are some of the applications of UAVs taken from the UAVS Website.

1. Aerial Policeman and Crowd Monitoring

2. Aerial Reconnaissance

3. Aerial Traffic and Security Watch

4. Air to Air Missiles, Air to Ground Missiles, Anti-Tank Missiles

5. Battlefield Management

6. Crop Dusting, Crop Management

7. Disaster damage estimation, Disaster effects management

8. Fire Fighting

9. Fishery Protection, Forestry

10. Geophysical surveys

11. Guided Shells

12. Life raft Deployment

13. Litter on beaches and in parks

14. Maritime and Mountain Search and Rescue

15. Mineral exploration

16. Oil and Gas Exploration and Production

17. Oil and gas pipeline

18. Pollution Control and Air Sampling

19. Search and Rescue

20. Telecoms relay and signal coverage survey

21. Waterways and shipping

22. Wide Area Munitions Deployments

2.6 Wing Structural Design

In 1930s, the structural design (Figure 10) of the airfoil had two spars running span-

wise throughout the wing. These spars would withstand the bending and shearing

stresses and at the same time, to attach the airfoil skin. The spar is the primary load

bearing structure of the airfoil.

Figure 11. Spars, booms and skin configuration

Developments (1950s) along the way had the airfoil skin as a load bearing member

of the structure. Stringers were also introduced to reduce / prevent buckling of the

airfoil skin caused by torsion loads. This is shown in Figure 11.

Figure 12. Spars, booms, skin and stringers configuration

By the 1980s, airfoil design evolved into a “torsion box” design (Figure 12). The

torsion box uses properties of thin surfaces to carry the imposed loads primarily

through tension while the close proximity of the enclosed core material compensates

for the tendency of the opposite side to buckle under compression. The torsion box

runs span wise along the length, providing the torsional stiffness and longitudinal

stiffness required in the wing. The torsion box consists of the airfoil skin, spars at

both ends and the stringers.

Figure 13. Torsion box made of spars, stringers and skin

2.7 Joints

Wing joint design is one of the most critical areas in aircraft structures, in particular

for fatigue consideration of long life structure. There are basically two types of wing

joint design, (i) fixed joint and (ii) rotary joint. Rotary joints are not discussed in this

report since they are beyond the scope of this report. The advantages and

disadvantages of different fixed joints are summarized in Table 2 below:

Joints Advantages DisadvantagesSpliced plates Widely used due to its light weight and

more reliable and inherent fail-safe featureSlightly higher cost, manufactural fitness required

Tension bolts Less manufactural fitness required, easy to assemble or remove. More economic for military fighter with thin airfoil

Heavy weight penalty

Lug/pin Less manufactural fitness required, easy to assemble or remove. More economical for military fighter with thin airfoil


Combination of spliced plates & tension bolts

Reliable and inherent fail-safe feature, and less manufactural fitness required


Table 2. Comparisons of Joints

The best fatigue design, of course, is one with no joints or splices. This is

accomplished on the modern transports which have no joints across the load path

except at the side of the fuselage. Wing sweep plus dihedral and manufacturing joints

requirements make the joint at the side of fuselage necessary It is important to keep

the joint short. A long joint tends to pull load in from adjoining areas.

Figure 13 below illustrates a lug/pin attachment for a fighter jet. This design

configuration is a highly loaded wing root joint. The lug/pin attachment as shown in

Figure 13 is widely adopted in aircraft wing designs because of good load transfers

without excessive stress concentration. This characteristic contributes in ensuring

fatigue life. The high structural efficiency of the lug/pin attachment is another reason

to attract designers attention during wing root design.

Most of the light loaded wings for general aviation aircraft adapt a single main front

spar and an auxiliary rear spar construction. Therefore, the wing root join usually is a

triple point lug joint as illustrated in Figure 14. The upper and lower lugs at the front

spar pickup wing bending loads, vertical shear loads and wing torque, the single luge

at the auxiliary rear spar take the wing vertical shear loads and torque only.

Figure 14. Lug/pin attachment for a fighter jet

Figure 15. Wing Root Triple Lug Joint

Figure 16. Various configurations of lug joints

Also, an additional feature to consider is the bolt within a hollow tube. Figure 16

illustrates a fail-safe design feature of “bolt within a hollow tube” which is used in

lug/pin arrangements. In case the hollow tube fails the bolt will take the load.

Figure 17. Doubler shear lug design with hollow tube

2.8 Dynamics of Flight

2.8.1 Introduction

This section introduces the basic background information about flight dynamics as it

is necessary to keep the aircraft stable and under control.

2.8.2 Six Degrees of Freedom (6 DOF)

An aircraft has six degrees of freedom. It can move upwards/downwards,

forward/backward, left/right and rotate about 3 axis (Pitch, Roll, Yaw). All these

contribute to the stability of an aircraft. When an aircraft is in a non-equilibrium

state, where it is disturbed in any of its axis of freedom, it must return to its original

state of equilibrium.

Figure 18. Body Axis, Moments, Rates and Controls

The deflections in control surfaces change the curvature in wing/tail surfaces and

changes the moments in the corresponding body axis. For instance, the ailerons will

control the roll of the aircraft by having one aileron being deflected upwards and the

other deflected downwards. Deflection of the rudders will cause a moment about the

Z axis (Yaw) and deflection of elevators will cause a moment about the Y axis

(Pitch).

2.8.3 Equations of Motion

Flight mechanics is the analysis of an aircraft motion using Newton’s laws. Most

airframe structures are flexible to some extent, but in this context, the airframe is

assumed to be rigid body. Newton’s laws are valid when written relative to an

inertial reference frame, which simply means that the reference frame is not rotating

or accelerating. If the equations of motion are being derived relative to an intertial

reference frame and if approximations characteristics of aircraft motion are

introduced into these equations, the resulting equations are those for flight over a non

rotating flat earth. Hence, for aircraft motion, the earth is taken as an approximate

inertial reference frae, and it is called the flat earth model. This would little to a small

error in most analyses. A general derivation of the equations of motion involved the

use of material system involving both solid and fluid particles. The end result is a set

of equations giving the motion of the solid part of the aircraft subject to

aerodynamic, propulsive and gravitational forces. To simplify the derivation of

equations, the correct equations for the forces are assumed to be known. Then the

equations describing the motion of the solid part of aircraft are derived. The aircraft

is assumed to have a right-left plane of symmetry with forces acting on the center of

gravity and the moments acting about the center of gravity. Forces acting on an

aircraft in flight are due to distributed surface forces and body forces. The surface

forces come from the air moving over the airplane and through the propulsion

system, while the body forces are due to the gravitational effects. Any distributed

force can be replaced by a concentrated force along a specific line of action. To have

all the forces acting through the same point, the concentrated force can be replaced

by the same force acting at the point of interest plus a moment about that point to

offset the effect of moving the force. The point usually chosen for this purpose is the

center of mass, or the center of gravity, because the equations of motion are the

simplest.

Figure 19. Distributed versus concentrated forces

The equations governing the translational and rotational motion of an aircraft are the

following.

1. Kinematic equations giving the translational position and rotational position

relative to the earth reference frame.

2. Dynamic equations relating forces to translational acceleration and moments

to rotational acceleration.

3. Equations defining the variable-mass of aircraft (Center of gravity, mass and

moments of inertia) versus time.

4. Equations giving the position of control surfaces and other movable parts of

the aircraft versus time.

2.8.4 Stability and Control

Stability and control studies focuses on the motion of the center of gravity relative to

the ground and motion of the aircraft about the center of gravity. Stability and control

studies include the use of the six degrees of freedom equation of motion. They are

being divided into two major categories (1) Static stability and control and (2)

Dynamic stability and control.

1. Static Stability and Control

It is concerned with the static stability, center of gravity effects. Given a disturbance

in a steady flight condition, static stability investigates the tendency of an aircraft to

reduce the disturbances. This can be achieved by focusing on the signs of the forces

and moments. There are also limits as to how far forward or rearward the center of

gravity can shift, away from the position of the center of gravity when the aircraft is

in equilibrium.

2. Dynamic Stability and Control

Dynamic stability and control is concerned with the motion of an aircraft in a

disturbance such as wind gust which changes the speed, angle of attack and/or the

sideslip angle. These can be performed using computer simulations, however it is

difficult to determine cause and effects.

2.8.5 Aircraft Sizing

The first step is the conceptual design of the aircraft. The configuration of the aircraft

is being selected after which the engines and wings are being sized according to the

mission requirements. The sizing of an aircraft is an iterative process. The sizing is

determined by pre-determined take-off gross weight, the engine size and the wing

size. The end result of sizing process is an aircraft with the dimensions. Next, the

geometry of an aircraft is determined by assuming that the center of gravity is

located at the wing aerodynamic center, such that the aircraft is in a state of static

equilibrium. Once the dimensions are obtained, the various instruments/components

are then installed at various positions that maintain the static equilibrium of the

aircraft. Statistical formulas are then used to estimate the weight of the individual

components and hence the gross take-off weight can be obtained. If the gross take-

off weight is not close enough to the initial pre-determined take-off weight, the

whole process is repeated again. Figure 12 below shows the sizing flowchart.

Figure 20. Aircraft Sizing Flow

Chapter 3 Conceptual Design Phase

3 Conceptual Design Phase

3.1 Introduction

The airfoil is a vital structure that is used to sustain the air loads that can be

experienced both in flight or when the aircraft is on the ground. The wing consists of

two essential parts, internal wing structure and external wing structure. The internal

wing structure consists of spars, ribs and stringers, whereas the external wing

structure, consist of the airfoil skin. Both the internal and external wing structure

must be able to sustain the bending moments, torsion and shear stress that may be

encountered while the aircraft is on the ground or in flight.

In this section, the purpose and conceptual design of the various wing components

will be explained. After which the design will analysed for its ability to maintain its

structural integrity. SolidWorks and ANSYS will be used for the design and analysis

work.

3.8.1 Airfoil Structure

3.8.1.1 Internal Structure

1. Ribs

Ribs are essential to maintain the shape of the wing section, to avoid buckling of the

airfoil skin. The ribs are required to support the wing-panels, to achieve the desired

aerodynamic shape and to maintain it. It has to effectively transfer large forces, add

strength and prevent buckling. There are various forms of ribs in the industry.

(a) Formed Ribs

Formed ribs are made from a sheet of metal that is being bent into shape.

(b) Plate-type Ribs

Consists of sheet metal, which has upturned edges, holes are being cut in the ribs to

reduce the weight. These ribs are commonly used in conditions of light to medium

loading.

(c) Truss Ribs

Consist of airfoil profiles that are joined together. They are suitable for a wide range

of loading types.

(d) Closed Ribs

Close ribs are constructed from profiles and sheet metal, they are suitable for closing

off sections of the wing. It is also suitable for a range of loading conditions.

(e) Forged Ribs

Forged ribs are manufactured using heavy-press machinery and are used for sections

where very high loads apply.

(f) Milled Ribs

Milled ribs are solid structures, they are being manufactured by milling excess

material from a solid of metal and they are also used in regions where very high

loads apply.

2. Stringers

Stringers are structural components that are attached to the skin panels that run the

length of the wing, and they are being attached to the ribs. There are several methods

to attach the ribs to the stringers, and then to the skin. Firstly, the stringers and ribs

can both be uninterrupted. The stringers will run over the rib, this would result in a

gap between the rib and skin. Rib and skin are indirectly connected; this would result

in a bad transfer of air loads from the skin to the rib. Another method is that the

stringers can be interrupted at the rib, interrupting the stringer in this way will result

in a weakened structure, and this would usually require additional strengthening

structure, known as a doubler. When holes are cut in the ribs to allow for the

attachment of the stringers, this also results in weakening of the ribs.

3. Spars

Spars are employed to support the ribs. Spars can take on the shape of an I-Beam.

The spar is the primary load bearing structure of the wing. The spar carries much

more load at the root as compared to the wing tips. It is important to consider the

loads that are experienced in flight, forces such as bending and shearing. At the same

time, forces in flight will cause the twisting of the wing. In order to overcome this

problem, a second spar / structure is usually included to form a torsion box structure.

In this case, the skin will serve as a spar-cap to resist bending, as part of the torsion

box to resist torsion and for the effective transfer of aerodynamics forces.

3.1.1.2 External Structure

1. Skin

The purpose of the airfoil skin is to give it the aerodynamic shape, to carry a fair

share of air loads, to carry torsional loads. Having the airfoil skin to carry a portion

of the load is called a stressed skin. The airfoil skin can be attached to the inner

structure through the use of bonding or riveting.

3.2 Design Targets and Mission Specifications

The design targets set forth for the design team is shown in Table 3 below. This shall

be used as a baseline for the team to base their calculations and designs upon. With

this pre-determined input information, the design team is broken down into 4

different teams(aerodynamics and propulsion, flight control and avionics, solar

energy and structures). Each team will be in-charge of their tasks but will have close

collaboration with every other team.

Take-off mass 60 Kg (Maximum)L/D = 20CL,Cruise ~ 1CD0 ~ 0.025Maximum climb angle = 5 deg (Minimum)Propeller efficiency = 80%Motor efficiency = 85%Max motor power = 2 x 1 KW (With reserves)Wing area for solar power = 15 m2

Power for 24hrs mission = 18 KWH (With reserves)

Table 3. Design Targets

3.2.1 Payload

The maximum allowed takeoff weight, inclusive of all payloads on board, is to not

exceed 60kg. The max design payload has yet to be determined.

3.2.2 Crew members required

The mission would require at least 4 members to be on site to operate the UAV. Two

members are to launch the UAV by hand, one will be in charge of piloting the UAV

and the other for flight tracking and data analysis.

3.2.3 Endurance

The UAV is required to have an operational period of around 24 hours.

3.2.4 Cruise Speed and Altitude

The targeted cruise speed would be 20km/h at an altitude of 8,000 meters.

3.2.5 Take off and Landing

The main consideration for this UAV was that it was to be as light weight as

possible, there will be no landing gear installed. This would allow for more room for

solar cells, batteries and payload. Resulting in higher power output. The UAV will be

hand launched during takeoff and skid land softly in a designated area to reduce the

amount of damage to the UAV.

3.3 Mission Profile and Critical Mission Requirements

Figure 21. Mission Profile

Phase 1: The UAV will be hand launched.

Phase 2: The proposed UAV should have a climb rate that allows it to climb to

8,000 meters in 2 hours and 30 minutes approx.

Phase 3-5: The UAV will cruise at an altitude of 8,000 meters for a maximum of

24 hours.

Phase 4: The descent will have duration of 2 hours and 30 minutes from 8,000

meters.

The critical mission requirements are that:

1. Total mass is not to exceed 60kg.

2. Endurance time to be no less than 24 hours.

3.4 Market Analysis

The main objective of this solar powered UAV is to encourage less fuel emissions,

allowing for greener aviation. Over the last few years, major companies have started

to invest in greener technology. Two major milestones were the design of the Solar

Impulse and the Zephyr that increased the popularity of solar powered UAVs. UAVs

can be used for a wide range of applications and hence there is a large market for this

technology.

3.5 Proposed UAV conceptual configuration

The configuration of the solar UAV is illustrated in the Figure 21. This will serve as

a guideline to the various teams, especially the structures team. The structures team

will be in charge of the design of the various components as well as integrating the

different structures / components. For the simplicity of calculations, the chord is set

at 1 meter with a span of 15 meters. This will enable us to have a total wing area of

approximately 15 m2 for the solar panels to generate sufficient power to stay in flight.

Figure 22. Proposed UAV Conceptual Configuration

NACA Profile 23012Wing Span 15 meters (Tip-to-tip)Chord 1 meterDihedral 8°-10°Anhedral 10°-15°

Table 4. Proposed Configuration Parameters

3.6 Airfoil Selection

Figure 23. NACA 23012 Profile

NACA 23012 has been selected by the aerodynamics team to be used on the Solar

UAV due to its low drag and high stall angle. So based on this NACA profile, the

structures team will know the scale of the wing based on a 1 meter chord length

which will enable them to work within the boundaries.

3.7 SolidWorks Modelling (Conceptual Design)

This section shows the conceptual design of the various components that upon

assembly, produces the complete wing structure (Internal and External). Engineering

drawings for the individual components are included in the APPENDIX section. The

assembled airfoil will also be illustrated below. All calculations included in this

conceptual design stage are based on a wing load of 2.5g. The estimated lift force

generated will be around 1471N and assumed to be uniformly distributed load

instead of elliptical loading.

1. Ribs

The rib is designed such that it is being divided into 3 portions. The 3 sections are

mainly the leading edge section, the middle section and the trailing edge section.

Slots are also cut into rib to make way for the stringers that span across the entire

wing span. This is illustrated in Figure 23 and Figure 24.

Figure 24. Ribs (Side View)

Figure 25. Ribs (Isometric View)

2. Spar – Comparison between an I Beam and a Circular Spar

The key parameters to determine when designing a spar is the ability of the spar to

sustain bending stress, torsion stress and shearing stress. The spars are placed at the

max thickness of the airfoil, which is 25% of chord length from the leading edge.

The reason why it is being located there is because the maximum wing loads are

encountered there. In this portion regarding spars, the team will be comparing the

advantages of single circular spar configuration, double circular spar configuration

and I-Beam configuration. And from the tabulated bending stress and shearing stress

Leading Edge Section

Middle Section

Middle Section

Adjoining Portions

Slots for Stringers

results, we will be able to determine which is the ideal spar profile based on weight,

strength to resist bending and twisting loads.

(a) I-Beam In the following page, Table 5 is the dimensions for the I-Beam that has been

decided upon. Table 6 shows the maximum loading conditions based on design

requirements. Table 7 shows list of parameters that are derived for the I-Beam based

on Table 5 and Table 6. The vital parameters are being highlighted in blue. These

values (Bending Stress, Shear Stress, Deflection Angle, Deflection), will be

compared against those for circular spar (Single Spar, Double Spar) in the later

portion in this section.

Table 5. Dimensions of I-BeamTerm Symbol Calculated Value

Volume (Total) V (m3) 1.800E-03Bending Moment (Total) M (Nm) 2341.965

Shear Force (Total) V (N) 735.75Table 6 . Max Allowable Loading Values

2nd moment of inertia (1 Spar only) I (m4) 4.475E-07

Volume (1 Spar only) V (m3) 1.800E-03Bending Moment (1 Spar only) M (Nm) 2341.965Bending Stress (1 Spar only) σ (MPa) 272.13

Top half Cross sectional Area (1 Spar only) A (m2) 1.200E-04Top half Centroid distance from N.A. (1 Spar only) y bar (m) 4.017E-02

First moment of inertia (1 Spar only) Q (m3) 4.820-06Thickness (1 Spar only) t (m) 1.000E-03Shear Force (1 Spar only) V (N) 7.358E+02Shear Stress (1 Spar only) VQ/IT τ (MPa) 7.92

Terms Meters Centimetersh (m) 0.104 10.4

h1 (m) 0.1 10t (m) = (h - h1)/2 0.002 0.2

b (m) 0.035 3.5tw (m) 0.001 0.1

y (m) Centroid to top 0.052 5.2

Deflection angle (1 Spar only) dw/dx (deg) 4.79Deflection (1 Spar only) w (m) 0.49

Table 7. Derived Parameters of I-Beam

(b) Circular Spar

Circular spar is one of the selected designs for sustaining the loads. Table 8 shows

the dimensions (Outer diameter, Inner diameter and Length). Table 9 will show the

max loading forces acting on the wing structure. For Table 10, Table 11 and Table

12, the results for the Bending Stress, Shear Stress, Deflection Angle, Deflection are

obtained based on the equations that are mentioned earlier on. In the case of the

circular spar, two conditions are being compared, Single Spar condition and Double

Spar condition.

Dimensions Front Spar Rear Spar One SparOuter radius (m) 0.05 0.036 0.05Inner radius (m) 0.049 0.035 0.0488

Length (m) 7.5

Table 8. Circular Spar Dimensions

Term Symbol Calculated Value

Volume (Total) V (m3) 4.006E-03Bending Moment (Total) M (Nm) 2341.965Shear Force (Total) V (N) 735.75

Table 9. Volume and Stress Result Tabulation

Single Spar Condition



Cross sectional Area (1 Spar only) A (m2) 3.725E-04Centroid distance (1 Spar only) y bar (m) 4.940E-02

First moment of inertia (1 Spar only) Q (m3) 1.840E-05Thickness (1 Spar only) t (m) 1.200E-03Shear Force (1 Spar only) V (N) 7.358E+02Shear Stress (1 Spar only) VQ/IT τ (MPa) 24.82Deflection angle (1 Spar only) dw/dx (deg) 4.72

Deflection (1 Spar only) w (m) 0.48Table 10. Single Spar Result Tabulation

Double Spar Condition

2nd moment of inertia (Front Spar) I (m4) 3.811E-07

Volume (Front Spar) V (m3) 2.333E-03Bending Moment (Front) M (Nm) 1873.572Bending Stress (Front) σ (MPa) 245.83

Cross sectional Area (Front) A (m2) 3.110E-04

Centroid distance (Front) y bar (m) 4.950E-02

First moment of inertia (Front) Q (m3) 1.540E-05Thickness (Front) t (m) 1.000E-03Shear Force (Front) V (N) 5.886E+02Shear Stress (Front) VQ/IT τ (MPa) 23.78Deflection angle (Front) dw/dx (deg) 4.50Deflection (Front) w (m) 0.46

Table 11. Front Spar Result Tabulation

2nd moment Inertia (Rear Spar) I (m4) 1.406E-07

Volume (Rear Spar) V (m3) 1.673E-03Bending Moment (Rear) M (Nm) 702.5895Bending Stess (Rear) σ (MPa) 179.92

Cross sectional Area (Rear) A (m2) 2.231E-04Centroid distance (Rear) y bar (m) 3.550E-02

First moment of inertia (Rear) Q (m3) 7.918E-06Thickness (Rear) t (m) 1.000E-03Shear Force (Rear) V (N) 1.902E+01Shear Stress (Rear) VQ/IT τ (MPa) 1.07Deflection angle (Rear) dw/dx (deg) 4.58Deflection (Rear) w (m) 0.46

Table 12. Rear Spar Result Tabulation

Comparison of I-Beam versus Circular Spar

In Table 13 below shows the comparison between 3 different configurations. They

are namely the Double Spar Condition, Single Spar Condition and the I-Beam. As

seen from the results, it is noted that the I-Beam has the lowest weight of 2.804 kg, a

deflection angle of 4.792 degrees and deflection of 0.486 m in comparison with the

other configurations. Although the I-Beam has the highest deflection angle as

compared to the rest, it was well within the design limits and offered the lowest

weight. The I-Beam can also withstand the high tensile and compressive stresses of

bending moment. Hence the Single I-Beam configuration will be selected for wing

structure due to its rigidity and weight saving characteristics. Together with an

additional C-channel at the trailing edge to form a torsion box. The problem with

circular spars is that it allows for the spars to ‘turn’ within the ribs, which reduces its

ability to transfer torsion loads. So an I-Beam has the advantage of being able to

transfer air loads effectively.

Carbon Fiber

Spar Front Rear1 Spar Only I beam

Allowable Tensile Strength (MPa) 2295.9552295.95

5 2295.955 2295.955Safety Factor 9.340 12.761 8.912 8.437

Allowable Shear Strength (MPa) 2295.9552295.95

5 2295.955 2295.955

Safety Factor 96.5522142.67

9 92.507 289.733Young's Modulus (GPa) 137.9 137.9 137.9 137.9Density (kg/m3) 1578 1578 1578 1578Mass (kg) 3.681 2.640 4.408 2.804Deflection angle (deg) 4.502 4.576 4.718 4.792Deflection (m) 0.456 0.464 0.478 0.486

Table 13. Comparison of Results between I-Beam and Circular Spar

Hence, based on the above calculations and selected I-Beam profile, the I-Beam is

designed to conform to the shape of the airfoil, as illustrated on the following page.

This is with the purpose that the top and bottom surface of the I-Beam will be flush

with the surface of the rib when it is assembled.

Figure 26. Spar (I Beam – Side View)

Figure 27. Spar (I Beam – Isometric View)

3. C-Channel (Rear Spar)

Similar to the I-Beam, the C-Channel is designed to conform to the shape of the

airfoil. This with the purpose that the top and bottom surface of the Channel will be

flush with the surface of the rib when it is assembled. The C-Channel (Rear Spar)

will connect the middle section rib and the trailing edge section of rib together. C-

Channel is used to create a torsion box within the wing structure.

Figure 28. C-Channel (Side View)

Figure 29. C-Channel (Isometric View)

4. Leading and Trailing Edge

The leading and trailing edge will be manufactured out of low density foam. The

purpose of this component is to maintain the shape and structural integrity of the

leading edge.

Figure 30. Leading Edge

Isometric View

Side View

Isometric View

Side View

5. Airfoil Skin and Stringers

The airfoil has stringers being designed into it. This is to prevent buckling of the

airfoil skin and to support the weight of the skin. The airfoil skin is made up of a

thin film that has the solar panel embedded in it. This is used to form an envelope

around the internal structure of the wing. This is important as the airfoil skin

must be properly attached to the internal structure to allow for effective transfer

of air loads to the spars. There will be no need for the use of screws / nuts for

attachment.

Figure 31. Airfoil Skin with Stringers (Side View)

Figure 32. Airfoil Skin with Stringers (Isometric View)

Airfoil Assembly

Figure 32 and Figure 33 shows the fully assembled airfoil. The primary means of

attachment is through the use of adhesives such as epoxy. Figure 33 shows the

internal structure of the airfoil. As explained earlier on, the rib is being divided into 3

sections, and connected through the use of the front spar and C-Channel (Rear Spar).

The area between the front spar and the C-Channel will form the torsion box. The

torsion box uses the properties of thin surfaces to carry the imposed loads primarily

through tension and compensate for the tendency of the opposite side to buckle under

compression.

Figure 33. Airfoil Assembly Conceptual Design (Isometric View)

Figure 34. Airfoil Assembly Conceptual Design (Internal Wing Structure)

Isometric View

Isometric View

Airfoil Skin

C-Channel

Leading Edge Foam

Ribs

Front Spar

Trailing Edge foam

Chapter 4 Detailed Design Phase

4 Detailed Design Phase

4.1 Detailed Calculations

In order to size the components correctly such that it can withstand the different

loads, detailed calculations have to be done to determine the various stress values.

Based on the material that we have selected and the geometry of conceptual ideas

that we have developed, the following stress values for bending and shearing could

be obtained.

Calculations for the spar are most vital in this stage as it is the primary load bearing

structure in the airfoil.

4.1.1 Bending Calculation

In the previous chapter it was highlighted that the team would have to make a

selection between the circular spar (single and double spar configuration) and the I-

Beam configuration. To compare the bending stress, the following equation is being

utilized.

σ=MyI x

σ is the bending stress;

M - the moment about the neutral;

y - the perpendicular distance to the neutral axis;

Ix - the second moment of area about the neutral axis x.

Shearing Stress

τ=VQ¿

V – Total shear force at the location;

Q – Statical moment of area;

t – Thickness in the material perpendicular to the shear;

I – Moment of Intertia of the entire cross sectional area.

Figure 35. Elliptical Loading

A half wing section of 7.5 meters is utilized for the calculations. The wing loading

decreases as it moves nearer to the tip of the wing, conforming to that of an

elliptically loaded condition. The fixed end represents the fuselage where the wing is

attached to and the wing tip is the end of the cantilever beam.

4.1.1.1 I-Beam

In the following page, Table 5 is the dimensions for the I-Beam that has been

decided upon. Table 6 shows the maximum loading conditions based on design

requirements. Table 7 shows list of parameters that are derived for the I-Beam based

on Table 5 and Table 6. The vital parameters are being highlighted in blue. These

values (Bending Stress, Shear Stress, Deflection Angle, Deflection), will be

compared against those for circular spar (Single Spar, Double Spar) in the later

portion in this section.

Table 5. Dimensions of I-Beam


Volume (Total) V (m3) 1.800E-03Bending Moment (Total) M (Nm) 2341.965

Shear Force (Total) V (N) 735.75Table 6. Max Allowable Loading Values



Top half Cross sectional Area (1 Spar only) A (m2) 1.200E-04Top half Centroid distance from N.A. (1 Spar only) y bar (m) 4.017E-02

First moment of inertia (1 Spar only) Q (m3) 4.820-06Thickness (1 Spar only) t (m) 1.000E-03Shear Force (1 Spar only) V (N) 7.358E+02Shear Stress (1 Spar only) VQ/IT τ (MPa) 7.92Deflection angle (1 Spar only) dw/dx (deg) 4.79Deflection (1 Spar only) w (m) 0.49

Table 7. Derived Parameters of I-Beam

Terms Meters Centimetersh (m) 0.104 10.4

h1 (m) 0.1 10t (m) = (h - h1)/2 0.002 0.2

b (m) 0.035 3.5tw (m) 0.001 0.1

y (m) Centroid to top 0.052 5.2

4.1.1.2 Circular Spar

Circular spar is one of the selected designs for sustaining the loads. Table 8 shows

the dimensions (Outer diameter, Inner diameter and Length). Table 9 will show the

max loading forces acting on the wing structure. For Table 10, Table 11 and Table

12, the results for the Bending Stress, Shear Stress, Deflection Angle, Deflection are

obtained based on the equations that are mentioned earlier on. In the case of the

circular spar, two conditions are being compared, Single Spar condition and Double

Spar condition.

Dimensions Front Spar Rear Spar One SparOuter radius (m) 0.05 0.036 0.05Inner radius (m) 0.049 0.035 0.0488

Length (m) 7.5

Table 8. Circular Spar Dimensions


Volume (Total) V (m3) 4.006E-03Bending Moment (Total) M (Nm) 2341.965Shear Force (Total) V (N) 735.75

Table 9. Volume and Stress Result Tabulation

Single Spar Condition



Cross sectional Area (1 Spar only) A (m2) 3.725E-04Centroid distance (1 Spar only) y bar (m) 4.940E-02

First moment of inertia (1 Spar only) Q (m3) 1.840E-05Thickness (1 Spar only) t (m) 1.200E-03Shear Force (1 Spar only) V (N) 7.358E+02Shear Stress (1 Spar only) VQ/IT τ (MPa) 24.82Deflection angle (1 Spar only) dw/dx (deg) 4.72Deflection (1 Spar only) w (m) 0.48

Table 10. Single Spar Result Tabulation

Double Spar Condition

2nd moment of inertia (Front Spar) I (m4) 3.811E-07

Volume (Front Spar) V (m3) 2.333E-03Bending Moment (Front) M (Nm) 1873.572Bending Stress (Front) σ (MPa) 245.83

Cross sectional Area (Front) A (m2) 3.110E-04

Centroid distance (Front) y bar (m) 4.950E-02

First moment of inertia (Front) Q (m3) 1.540E-05Thickness (Front) t (m) 1.000E-03Shear Force (Front) V (N) 5.886E+02Shear Stress (Front) VQ/IT τ (MPa) 23.78Deflection angle (Front) dw/dx (deg) 4.50Deflection (Front) w (m) 0.46

Table 11. Front Spar Result Tabulation

2nd moment Inertia (Rear Spar) I (m4) 1.406E-07

Volume (Rear Spar) V (m3) 1.673E-03Bending Moment (Rear) M (Nm) 702.5895Bending Stess (Rear) σ (MPa) 179.92

Cross sectional Area (Rear) A (m2) 2.231E-04Centroid distance (Rear) y bar (m) 3.550E-02

First moment of inertia (Rear) Q (m3) 7.918E-06Thickness (Rear) t (m) 1.000E-03Shear Force (Rear) V (N) 1.902E+01Shear Stress (Rear) VQ/IT τ (MPa) 1.07Deflection angle (Rear) dw/dx (deg) 4.58Deflection (Rear) w (m) 0.46

Table 12. Rear Spar Result Tabulation

4.1.1.3 Comparison of I-Beam versus Circular Spar

In Table 13 below shows the comparison between 3 different configurations. They

are namely the Double Spar Condition, Single Spar Condition and the I-Beam. As

seen from the results, it is noted that the I-Beam has the lowest weight of 2.804 kg

for a 7.5 meter section, a deflection angle of 4.792 degrees and deflection of 0.486 m

in comparison with the other configurations. Although the I-Beam has the highest

deflection angle as compared to the rest, it was well within the design limits and

offered the lowest weight. The I-Beam can also withstand the high tensile and

compressive stresses of bending moment. Hence the Single I-Beam configuration

will be selected for wing structure due to its rigidity and weight saving

characteristics. Together with an additional C-channel at the trailing edge to form a

torsion box. The problem with circular spars is that it allows for the spars to ‘turn /

rotate’ within the ribs, which reduces its ability to transfer torsion loads. So an I-

Beam has the advantage of being able to transfer air loads effectively.

Carbon Fiber

Spar Front Rear1 Spar Only I beam

Allowable Tensile Strength (MPa) 2295.9552295.95

5 2295.955 2295.955Safety Factor 9.340 12.761 8.912 8.437

Allowable Shear Strength (MPa) 2295.9552295.95

5 2295.955 2295.955

Safety Factor 96.5522142.67

9 92.507 289.733Young's Modulus (GPa) 137.9 137.9 137.9 137.9Density (kg/m3) 1578 1578 1578 1578Mass (kg) 3.681 2.640 4.408 2.804Deflection angle (deg) 4.502 4.576 4.718 4.792Deflection (m) 0.456 0.464 0.478 0.486

Table 13. Comparison of Results between I-Beam and Circular Spar

Hence, based on the above calculations and comparisons, the I-Beam profile is

selected.

4.1.2 Structural Calculation

Upon obtaining the structural rigidity and the feasibility of the wing section,

assumptions have to be made to calculate the shear flow.

1) Flanges and stringers are assumed to carry only axial stress.

2) Skin and Webs are assumed to carry only shear stress.

3) The axial stress in the stringers and shear stress in the sheet are assumed to be

constant throughout their respective thickness

4) Shear stress is constant in the web between stringers.

4.1.2.1 Shear Flow

With regard to an I-beam spar configuration, the areas are idealized to the points 1, 2,

3 and 4. Also, for further simplification, the airfoil is assumed to be symmetric.

Hence this segment of the report only aims to obtain a rough estimated value of the

shear stresses occurring in the airfoil.

SectionThickness (mm)

Shear Flow (N/mm)

Max Shear Stress (MPa)

1-2 (front spar flange) 1.000 5.6586 5.6586

arc 1-2 (leading edge) 0.025 0.1734 6.9360

1-3 (top skin) 0.025 0.1726 6.9040

2-4 (bottom skin) 0.025 0.1726 6.9040

3-4 (rear spar flange) 0.500 0.6008 1.2016

arc 3-4 (trailing edge) 0.025 0.0476 1.9040

Table 14. Shear stress tabulation of different sections

As seen in Table 14, the maximum allowable shear stress amounts to 6.936MPa,

which is acting on the Teflon backing film of the solar panels.

4.1.2.2 Shear Center

The shear centre is that point through which the loads must act if there is to be no

twisting, or torsion. The shear centre is always located on the axis of symmetry;

therefore, if a member has two axes of symmetry, the shear centre will be the

intersection of the two axes.

Figure 36. Equal flanged section and examples of sections with one axis of symmetry

In an airfoil, likewise, the shear centre is the location where by if the wing load acts

upon that point, there is no twisting or torsion. After calculation of shear stress, using

the shear flow distribution in the wing, the shear centre of the wing section is

obtained. Measuring from the leading edge, the shear centre is 37.38mm behind the

front spar.

4.2 Detailed Design

The detailed design phase aims to address the issues highlighted in the conceptual

design phase as well as to size the various components to the sizes that is capable of

withstanding the various loadings experienced in flight. Changes are made to the

design in order to allow for the various equipment to be installed on the wing section.

For instance, the solar team requires that the battery pack should be installed within

the leading edge airfoil. The battery pack has dimensions of L 270mm X B 45mm X

90mm and requires a hole of that dimension to be cut out within the leading edge

airfoil.

Figure. Leading and Trailing edge airfoil

Each rib, if based on the design in conceptual design phase, has a mass of 98.94g.

That is quite a significant weight as a meter of wing section has 3 ribs (approx 300g),

which makes up one-third of the overall structural mass. Since there are much

unused areas in a rib, and based on conventional airfoil design, holes are being cut

out in the ribs to reduce the overall mass.

Figure. Ribs

Figure . Airfoil Assembly (Detailed Design)

4.3 ANSYS

Simulation on ANSYS is carried out to ensure that the various components are sized

correctly to withstand the various loadings in flight.

4.3.1 Meshing Conditions

The surface of the airfoil must be meshed to a fine finite element due to the fact that

it is a curve surface and it is vital for the team to analyze the stress / deformation as

accurately as possible. The mesh sizes for the various components will vary

according to the geometry and how vital the component is in sustaining a load. In

Figure 32 the leading edge has to be finely meshed due to the reason it has a curved

surface as compared to the coarsely meshed spar, for more accurate results.

Figure 32. Airfoil Skin Meshing

Figure 33. Airfoil Internal Structure Meshing

Figure 34. Airfoil Internal Structure Meshing (Zoom-In)

4.3.2 Stress Analysis

Figure 35. Stress Analysis

Based on the results from ANSYS shown in Figure 35, the stress experienced by the

airfoil internal structures are within acceptable range. The lowest possible stress is

reflected by the dark blue areas ( ), with a stress of 740.51 Pa. The highest stress

that is encountered in the structure is reflected by areas in yellow ( ), with a stress of

1.2045e^7 Pa. This is still within acceptable range. The bending stress that has the

most significant impact on the spar is towards the root of the wing and gradually

decreases as it moves towards the tip, this is due to the fact that the root is the only

region that is connected to the fuselage of the aircraft that forms a cantilever beam.

This max stress encountered in the structure is still within tolerable range as there are

no failure states shown in the analysis results.

4.3.3 Deformation Analysis

Figure 36. Deformation Analysis

From the deformation result shown in Figure 36, it is noted that the least deformation

is located at the region where the airfoil is connected to the fuselage, represented in

blue ( ). The location with the highest deformation is at the wing tips represented in

yellow ( ). The deformation of the structure ranges from 0m to 0.00015m. This can

be experienced both in flight and when on the ground due to the fact that the wing tip

is not being supported. The deformation is still within acceptable range as no failure

in the structure is reflected in the analysis results.

Further tests have to be conducted which will also include the wing section joints.

Chapter 5 Building of Test Section

5 Building of Test Section

The building of an airfoil test section has been carried out in order to verify the

results from ANSYS. The airfoil test section will be tested for its maximum possible

loading that it can sustain, the max deflection possible, to test the joints between the

various airfoil sections as well as to have a visual layout of where each and every

component will go. This requires the team to work closely with the solar team, as the

solar team will install the solar panels, batteries and maximum power point tracking

(MPPT) in or within the airfoil. The materials that we are working with are pre-

impregnated carbon fiber, high and low density foam. Pre-impregnated carbon fiber

is carbon fiber that already has epoxy applied on the carbon fiber sheets.

The weights for the individual components are then tabulated so that the team can

estimate the overall airfoil weight.

This is how we have constructed the various components, taking the I-Beam as an

example. Each layer of carbon fiber is 0.25mm thick, in order for the structure to be

sufficiently rigid to withstand the various loads, we will need to increase the number

of plies used. The Figure 37 below shows the I-Beam, each line will represent a layer

of carbon fiber. First the C-Section (highlighted in Red) is being constructed by

laying pre-preg carbon fiber over a mould as shown in Figure 38. After which the

two C-Sections are placed back to back as seen in the Figure 39. To further

strengthen the top and bottom sections of the I-Beam, 6 more layers of carbon fiber

are overlaid to increase thickness and ability to hold the I section together as seen in

Figure 40.

These components are then put into a vacuum bag as seen in Figure 41 and Figure

42, where air is being extracted prior to putting into the autoclave machine. The

process is carried out at elevated temperature and elevated pressure. Elevated

pressure allow for a high fiber volume fraction and low void content for maximum

structural efficiency.

Figure 37. I-Beam Schematics

Figure 38. Figure 39. Figure 40.

Figure 41. Figure 42. Figure 43.

As for the Styrofoam components, they are made fabricated using hot wire cutters.

Epoxy was utilized to bond all the various components together. The various

components fabricated and their ‘building block’ dimensions are as follows:

1. I-Beam Spar

Figure

2. C-Channel (Rear Spar)

CFRP Section

LayersDimensions

(mm)Specifications

Strips6 (top),

6 (bottom)

30 x 1000 N.A.

C-Section2 (left), 2 (right)

134 x 100017mm (top & bot) 100mm( flange)

CFRP Section

LayersDimensions

(mm)Specifications

C-Section

2 92 x 1000 20mm (top & bot), 52mm (flange)

3. Stringers

CFRP Section

Layers Dimensions (mm) Specifications

L-shaped 2 30 x 1000 10mm (top), 20mm (flange)

4. NACA 23012 Ribs

Ribs are fabricated out of blue high density foam. They are cut out using hot wire

cutters that is available for use at the machine shop. The initial plan was to fabricate

the ribs out of carbon fiber, but there were a few issues with carbon fiber as ribs.

Carbon fiber ribs do not have sufficient contact area with the film that goes around

the entire airfoil. It is always possible that we increase the thickness of the carbon

fiber rib, however that was not done as blue high density foam ribs would be much

lighter as compared to a carbon fiber rib with the same thickness as the blue high

density foam ribs. And also due to the ease of fabrication using blue high density

foam, blue high density foam was chosen as the material for the ribs.

5.1 Prototype Test Section 1

Test section 1 is developed to have a visual representation of where each and every

component will go. It also allows the team to assess the design and make changes to

it based on the equipment that will be installed within the wing section. For instance,

the batteries will be attached to the I-Beam, hence the ribs and leading edge foam

will require for an opening to allow the batteries to fit in. During the design phase

and the building phase, there was close and frequent communication between the

various design teams. This test section also allows the team to understand the

difficulties faced in fabricating the airfoil structure with just in-house tools.

Figure 44. Test Section 1 Assembly

Figure 45. Test section 1 with simulated solar panels

I-Beam

High-density foam ribs

Stringers

C-Section

Solar panel embedded film

Center rib

From the test section, it was most ideal that the solar panels have sufficient support at

the edges. The solar panels are being placed in modules of 7 by 3 panels, which is

being represented by the green boxes in Figure 45. The solar team require that the

middle of the mid solar panel be aligned to the middle of the center rib. Necessary

changes to the design are noted as follows:

1. Distance between ribs has to be changed from 395 mm to 410 mm to allow

for each module to be centralised with reference to the center rib.

2. Holes will be cut in the ribs to reduce the weight of the overall structure. This

would allow for 40% reduction in weight of each rib.

3. The stringers have to be strategically placed to allow the edges of each solar

panel to rest on for maximum support.

4. The I-Beam will be flushed to the top and bottom most surface of the airfoil

for maximum support.

5. Leading and Trailing edge foam will be added for maximum support for solar

panel embedded film.

6. Openings have to be cut in the rib and leading edge foam to allow for

mounting of batteries which will be attached to the I-Beam.

5.2 Prototype Test Section 2

Figure 46. Test section 2 with the design changes included

The necessary changes that was highlighted in the previous section was designed and

built into test section 2. To reduce the weight of each rib, holes were cut in strategic

locations, preferably well away from the cuts for the stringers so that the structure

will not be weakened significantly. These cuts in the rib would reduce the weight of

the ribs to 60% of the full rib. A small section of the leading edge foam is also cut

out to allow for the battery to be contained within, as highlighted in Figure 46. The

above two are the two major changes to the design.

The actual solar panel embedded film will be installed on this test section 2, for load

testing as well as to allow the solar energy team to conduct their necessary tests. A

test rig is designed and built in order to allow the team to attach the test section for

the various tests to be conducted. Figure __ shows the solar panel embedded film.

I Beam

C-Channel

Trailing Edge Foam

Stringers

Hole for lithium batteries

Ribs with holes for weight reduction

Leading edge foam with holes dug out for lithium battery

Figure . Solar Panel Embedded Film

Figure . Battery pack

ADD IN FINAL PIC of ENTIRE ASM

5.3 Mass Estimation

Component Unit Unit Densit Theoretic Qty Test Test

Mass (g)

Volume (m3)

y (kg/m3)

al DensitySection 1 Mass (g)

Section 2 Mass (g)

Front Spar290.

10.0002

41208.7

5 1578 1 290.1 290.1

Rear Spar 55.4 4.7E-051178.7

2 1578 1 55.4 55.4

Stringers 10.1 7.5E-061346.6

7 1578 8 80.8 80.8

Ribs104.

7 2.35

3246.3529

4 0

Ribs w Holes61.3

6 144.3764

7Leading Edge Foam

112.9 2 225.8 112.9

Trailing Edge Foam 46.7 2 93.4 65.38Skin (FEP Film) 50 2150 1 50 50Epoxy 200 1 200 200

Total Mass without Joints1241.852

9998.9564

7

Total Mass with Joints (est) 1198.747

8Table 15. Mass estimation of airfoil

Mass estimation is essential to ensure that the airfoil structure stay within designated

limits. Mass is by itself an important value as it can be used to estimate the

operational boundaries as well as the center of mass which gives the static / dynamic

stability of the UAV. The first test section differs from the second test section by not

having any holes cut out in the ribs. As noticed in test section 1, the ribs without

holes would have a mass of 246.35g as compared to the rib from the test section with

a mass of 144.37g. Another major difference in test section 1 and test section 2 is

that the leading edge foam of test section 2 has a hole cut in it to allow for the

batteries to be contained within the leading edge foam.

As reflected in the Table 15 above, test section 1 has a mass of 1241.85g and test

section 2 with a mass of 998g. All these values exclude the mass of batteries /

equipment that are installed within the airfoil structure.

Chapter 6 Load Testing Results

TO BE ADDED IN FOR FINAL REPORT DUE ON 9 TH APRIL, TESTS ARE CURRENTLY

BEING CONDUCTED

ConclusionOver the two semesters, the students from the structures team have developed a

finalized design for the airfoil. The students have looked into different configurations

and the team has decided upon an I-Beam wing structure, connected by ribs. The

finite element verification has also been conducted to ensure that the stresses and

deformation fall within design limits. In order to the complete CAD model, finite

element verification will be extended to the full CAD model, which will be inclusive

of the various joints, since it is most accurate prediction of the wing’ behaviour apart

from direct experimental testing. Control surfaces are not included in this design and

will be addressed in the future.

The students also built two test sections. The first test section was for the team to

visualize where the various equipments will be installed within the wing. Based on

the required changes highlighted in test section 1, the team then built a second test

section with the solar cells and battery-cells installed in place. The joints between the

various wing sections are also designed into the second test section. The second test

section will be used for load testing as well as to allow for the solar team to conduct

their tests.

Many aspects of the wing design have been considered in this report, there are still

much to be done before releasing for production. During the analysis, many

assumptions were made which may lead to the lack of accuracy.

Important Lessons Learnt Conduct more thorough design reviews.

Using literature or benchmark data from conventional aircraft design may not

always be useful.

Material selection and structural analysis play an important role in aircraft

design.

Manufacturing the design make design flaws more noticeable as compared to

having it on paper.

Future WorksLoad testing will have to be carried out on the test section to derive the point of

failure of the airfoil; this is to ensure that the stress, deformation and failure state fall

within the design limits. Once the design verification for the airfoil has been

complete through the use of finite element method and structural testing, the students

can proceed on into the design phase for the fuselage and the tail control surfaces.

Interfaces to integrate the airfoil-fuselage, airfoil-motors will have to be designed in

the upcoming semesters. The students will also have to look into means to simplify

the assembly process of the airfoil, be it through the use of an assembly jig or

purchasing tools to make the fabrication process more simplified.

References[1] The History of Solar Energy, US Department of Energy, Energy Efficiency and

Renewable Energy, http://www1.eere.energy.gov/solar/pdfs/solar_timeline.pdf

[2] White, Frank M., Fluid Mechanics, Sixth Edition, McGraw-Hill Companies, Inc.,

New York, NY, 2008, pp. 818

[3] Megson, T.H.G., Aircraft Structures for Engineering Students, Fourth Edition,

Elsevier Ltd., Oxford, UK, 2007

[4] Sun, C. T., Mechanics of Aircraft Structures, John Wiley & Sons, Inc., Hoboken,

NJ, 2006

[5] Second moment of area, Wikipedia,

http://en.wikipedia.org/wiki/Second_moment_of_area, Accessed May 8, 2009

Appendix

FYP FINAL

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