Aerodynamics of an Unmanned Aerial Vehicle

Submitted byCheah Ai Lin

Department ofMechanical Engineering

In partial fulfillment of the requirements for Degree of Bachelor of Engineering National University of Singapore

Session 2011/2012

SUMMARY

This thesis consists of the design and analysis of the aerodynamics of an unmanned aerial vehicle (UAV).

The aim of this UAV is for infantry men to deploy the UAV in an urban operations environment where there is space constraint. The design of the UAV has to be relatively small and portable so that infantry men are able to carry it during their operations. The wings of the aircraft are designed to be foldable so that it can be kept for storage and released when deployed. Thus, the restriction of the size of the aircraft would mean that the wingspan of the aircraft has to be relatively short. In addition, a launcher will be used to launch the UAV.

There are a few main factors which determine the span of the aircraft; namely the estimated weight of the aircraft, the aerofoil shape, aspect ratio and the lift required by the aircraft. There is a wide range of aerofoils to choose from and based on the aerofoil data, calculations were done to ensure that the minimum lift provided by the aerofoil was sufficient to lift the aircraft. The choices were then narrowed down before a 3-D computational fluid dynamics using Solidworks Flow Simulation was performed on the Solidworks model.

Aerofoil selection was done using the 2-dimensional (2D) analysis of aerofoil shapes to determine the lift characteristics. The X-foil which is a well-established aerofoil computational program is capable of generating the characteristics of each aerofoil.

Computation Fluid Dynamics (CFD) was then carried out on the 3-D model using Solidworks Flow Simulation. The main aim of using CFD was to ensure that the aircraft modelled in Solidworks is subjected to a real-life situation and that the lift provided by the aerofoil selection previously is capable of generating sufficient lift for the given speed. Based on the analysis and selection process of different aerofoils, NACA8414 was chosen as the wing aerofoil and SD 8020 was chosen as the tail aerofoil. Computational Fluid Dynamics analysis was also used to plot the Moment vs Angle of Attack (AOA) graph so as to ensure that the aircraft is longitudinally stable in trim conditions. In addition, the lift generated for each individual surface, namely the wing and the tail, can be determined from the Solidworks flow simulation. The individual lift surfaces for the wing and the tail are used to determine the position of the wing and tail and also to calculate the centre of gravity.

The fabrication and manufacturing of the components such as the wings, tails and fuselage were done in-house using the CNC Hot Wire Cutter machine.

Flight instruments and equipments were mounted on the UAV and flight data were collected. A powerless glide test was conducted to determine the Lift/Drag ratio. This is to validate the theoretical and practical data collected during the flight tests. The difference in theoretical and practical data were discussed and accounted for.

ACKNOWLEDGEMENT

The author wishes to express sincere appreciation and gratitude to the supervisor, Associate Professor Gerard Leng for his patience and guidance throughout this project. His role as a mentor was invaluable. Associate Professor Gerard Leng guided the team throughout the course of the project. Without his guidance, the team would not be able to meet the requirements and deadlines. Gratitude is also extended to the Staff and Technicians of the Dynamics Lab for their administrative, technical support and assistance throughout the project. Special thanks also go to Leong Jun Yi and Andrew Ong who provided advice to the manufacture and flying of the UAV. Last but not least, the author would also like to thank her teammates (Alphonsus, Long Qiang, Trixie and Shiao Loong) who had put in their best efforts for this project. It had been an enjoyable experience working with them.

TABLE OF CONTENTS

SUMMARYii

ACKNOWLEDGEMENTiv

TABLE OF CONTENTSv

LIST OF FIGURESvii

LIST OF TABLESix

LIST OF SYMBOLSx

1. INTRODUCTION1

2. OBJECTIVES2

3. LITERATURE REVIEW5

3.1 Aerofoils5

3.2 Planforms7

3.3 Centre of Gravity and Aerodynamic Centre8

3.4 Dihedral10

3.5 Drag11

4. DESIGN11

4.1 Selection of Aerofoil11

4.2 Design of UAV15

4.3 Wing Aerofoil and Wing Size15

4.4 Tail Aerofoil and Tail Size19

4.5 Downwash Angle22

5. COMPUTATIONAL FLUID DYNAMICS23

6. DRAG26

7. STABILITY28

7.1 Longitudinal Stability28

7.2 Lateral Stability29

9. RESULTS AND EVALUATION30

9.1 Verifying Lift30

9.2 Lift-to-Drag Ratio30

10. RECOMMENDATIONS FOR FUTURE WORK33

11. CONCLUSION34

REFERENCES35

Appendices36

LIST OF FIGURES

Figure 1: Project Flow Chart3

Figure 2: Geometric Parameters of Aerofoil [7]5

Figure 3: Aerodynamic Centre [9]9

Figure 4: Balancing moments about the CG [3]9

Figure 5: Front view of Aircraft having dihedral of 10

Figure 6 : Aircraft rolls to the right due to a right sideslip10

Figure 7: Lift characteristics of selected aerofoils14

Figure 8: Cl vs Cd comparisons for the selected three aerofoils using Xfoil16

Figure 9: Flow analysis across the Solidworks model17

Figure 10: Tail Configuration [4]20

Figure 11: Horizontal Tail effective angle of attack [11]23

Figure 12: Basic Mesh of Computational Domain and Refinement of Mesh24

Figure 13: Lift vs AOA25

Figure 14: Moments vs AOA graph from Solidworks Flow Simulation28

Figure 15: Rolling stability due to dihedral [1]29

Figure 16: Powerless Glide Test31

Figure 17: 2D NACA 8414 Aerofoil Characteristics generated using Xfoil36

Figure 18: Downwash angle for no sweepback37

Figure 19: Typical values for horizontal and vertical tail volume coefficients38

Figure 20: Coefficient of Drag for 4-digit aerofoils at low Reynolds number39

Figure 21: Coefficient of Drag of fuselage vs Fineness ratio39

Figure 22: Example of Drag Breakdown for Piper Cherokee 18040

Figure 23: Refinement of cells and curvature41

Figure 24: Altitude vs Line Number from ArdumegaPilot Mission Planner42

Figure 25: Latitude, Longitude and Altitude of the UAV during Glide Test42

Figure 26: Latitude and Longitude conversion to distance (m)42

LIST OF TABLES

Table 1: Shape of Aerofoils6

Table 2: Comparison of the characteristics of various planforms [5]7

Table 3: Wing Aerofoil characteristics16

Table 4: Main Wing Specifications18

Table 5: Dimensions for parameters stated in Figure 1021

Table 6: Wing and Tail specifications25

Table 7: Drag components26

Table 8: Sample results obtained from Solidworks Flow Simulation41

Table 9: Wing Parameters43

Table 10: Horizontal Tail Parameters43

Table 11: Vertical Tail Parameters44

LIST OF SYMBOLS

ARAspect ratio

b Span

c Chord

CL3D Lift Coefficient

Cl2D Lift Coefficient

ClSlope of sectional lift curve, deg1

CmSlope of pitching moment coefficient, deg1

DDrag, N

eOswalds efficiency factor

HVertical distance, m

iAngle of incidence, deg

lLength of fuselage, m

L Lift, N

M Moment per unit length Nm m1

Density, kg/m3

S Wing Surface Area, m2

SwetWetted Surface Area, m2

VFree stream velocity, ms1

VHHorizontal Tail Volume Ratio

W Weight, N

X Horizontal distance, m

Angle of attack, deg

Diameter of fuselage, m

1. INTRODUCTION

The research on an Unmanned Aerial Vehicle (UAV) has been ongoing as UAV proved to be essential and useful in helping infantry men in carrying out their missions. In todays modern battlefield, more and more UAVs are being developed to further aid people in carrying out their operations. One of the most basic functions of the UAV is for surveillance purposes. With a camera mounted on the UAV, infantry men are able to survey the grounds without endangering their lives. In addition, an autopilot system will also be mounted on the UAV to stabilise it; thus it does not require infantry men to control the plane at all times. Currently, the L3 cutlass tube-launched Unmanned Aircraft System (UAS) and the Switchblade serve the purpose of launching the UAS into the sky and has a camera to monitor its flight and surroundings.

This project aims to develop a portable UAV which is capable of launching vertically into the air in an urban operations environment and deployment of the wings as the UAV is launched into the sky. A camera system is placed on board to provide a live feed back to the infantry men. This thesis will cover the aerodynamics of the UAV: from the design and selection process of the aerofoils for the wings and tails, the lift and flow analysis of the aircraft as well as the drag of the aircraft.

In the process of designing the UAV, there are some constraints which need to be addressed so as to ensure that the final UAV model meets the mission requirements.

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1. The UAV has to be portable such that an infantry man can easily bring it along during his operations. The estimated size of the UAV is about the size of the fieldpack. This constraint limits the wingspan and total length of the UAV.

2. The wings have to be folded for ease of storage and for portability. This restricts the dimensions of the wingspan and its chord.

3. Telemetry and equipment weight about 1kg and the wings must be capable of sustaining at least 10N of lift.

2. OBJECTIVES

This project consists of mainly three phases: Selection of aerofoils, Computational Flow Dynamics (CFD) of the selected aerofoils and Flight Test Analysis. Each phase contributes to the final objectives and thus completing the individual phases on time means that the other group members can move on with their parts using the data collected.

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Figure 1: Project Flow Chart

The main objective for each phase is to:

1. Selection of Aerofoil

Select an aerofoil for the main wing and the tail respectively. The dimensions of the wing and tail namely the span, aspect ratio and chord length are determined by the optimisation program. Theoretical calculations would be performed to ensure that the aerofoils selected are able to provide enough lift and ensure longitudinal stability of the aircraft during trim condition.

2. Verified by Solidworks Flow Simulation

The final model which is modelled using Solidworks will be analysed using the Solidworks Flow Analysis. Computational Fluid Dynamics (CFD) would be performed on the model and the respective lift values can be obtained from the

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wing and tail surfaces. Pitching moment of the UAV can be obtained from the Solidworks flow simulation.

3. Flight Test

Performing numerous flight tests on the final model further validates the theoretical calculations and simulation results. Data can be collected during flight using the telemetry on board and the glide test ratio can be determined.

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3. LITERATURE REVIEW

3.1 Aerofoils

Aerofoils refer to the cross-sectional shape of the wings of an aircraft. An aerofoil consists of leading edge, trailing edge, maximum thickness, maximum camber, camber line and chord. The general aerodynamic characteristics of the wing such as lift, drag and stall characteristics are affected by aerofoil geometric parameters. The geometric parameters are summarized below:

Figure 2: Geometric Parameters of Aerofoil [7]

The maximum thickness and thickness distribution contributes to the amount of lift generated by the aerofoil. With a thicker aerofoil, the lift generated is higher but form drag also increases due to flow separation. The aerofoil generates lift due to the difference in pressure on the upper and lower surfaces. According to

Bernoullis equation, as flow passes the upper surface which has a larger curvature compared to the lower surface, the velocity of flow speeds up and this result in a lower pressure. On the lower surface, the velocity of flow is not as high as compared to the upper surface, thus the pressure at the lower surface is higher

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than the pressure on the upper surface. The difference in pressure between the upper and lower surfaces generates a net upward force known as lift.

Cambered aerofoils provide greater lift due to greater curvature and they have a nose down or negative pitching moments. This negative pitching moment must be offset by the down ward force from the horizontal tail.

Table 1: Shape of Aerofoils

Types of AerofoilProperties

Heavily cambered1.High lift

2. Low Cm

Moderately cambered Flat bottom1.Moderate lift

2.Moderate Cm

Symmetrical no cambered1.Zero lift at zero

AOA

2.Virtually no

pitching moment

Moderately cambered aerofoil is highly recommended for the main wings so that it can provide a higher suction force and hence higher lift force generated at low speeds. As for the tail, symmetrical aerofoil would be chosen as the main purpose of having the tail is to balance the upward force generated by the wing. Thus,

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when given a negative angle of incidence of the tail, the tail generates a downward force. For symmetrical aerofoil, the lift generated is zero when angle of attack is zero.

3.2 Planforms

Different types of planforms changes the lift distribution of the wings. Some of the more common planforms are summarised below.

Table 2: Comparison of the characteristics of various planforms [5]

ProfileAdvantagesDisadvantages

Elliptical1.Generates highest lift1.Difficult to fabricate

coefficients2.Gives little warning

2. Lowest induced dragprior to complete stall,

pilot has little time to

react due to poor aileron

effectiveness

Rectangular1.Easy to fabricate1.Higher induced drag

2.Provides adequate2.Trailing vortices at

stall warning andwing tips increases drag

aileron effectiveness

and thus increase

stability

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Tapered1. Lower induced drag1. Tip stall will lead to a

than rectangular wingsloss of aileron control

(more effective only at

high speeds)

The figure above summarizes the three common planforms. Although the elliptical planform is most capable of generating high lift and reducing drag, the shape of the planform is too complex and thus makes it difficult to fabricate. Tapered wings also proves to be more efficient in reduce drag, however it is only effective at high speeds. Eventually, rectangular wings are chosen as they are simple to manufacture and drag can be reduced by ensuring a smoother surface on the wings and also the aerodynamic shape of the fuselage.

3.3 Centre of Gravity and Aerodynamic Centre

The position of centre of gravity (CG) of the UAV has an impact on the longitudinal stability and selection of horizontal tails angle of incidence. The general rule of thumb for the CG is to be placed slightly in front of the aerodynamic centre (AC) of the wings i.e. 25% of the chord.

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Figure 3: Aerodynamic Centre [9]

The AC is the point where all the lift acts and where the pitching moment coefficient is not affected by the change in angle of attack. By placing the CG slightly in front of the AC, the anticlockwise moment produced by the wings due to the distance between the AC and CG have to be balanced by a clockwise moment produced by the horizontal tail. Thus, the horizontal tail has to exert a downward force to counteract the moment produced by the wings. And this amount of downward force required will determine the angle of incidence of the tail.

Figure 4: Balancing moments about the CG [3]

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3.4 Dihedral

Dihedral helps to bring about lateral stability, also known as roll stability. Dihedral angle is the angle between the plane perpendicular to the root chord and the plane which passes through the upper and lower surfaces of the wing. In the case of dihedral, the wing tip is higher than the wing roots. Having a dihedral angle will improve the lateral stability of the plane.

Figure 5: Front view of Aircraft having dihedral of

Figure 6 : Aircraft rolls to the right due to a right sideslip

As the aircraft sideslips to the right as depicted in Figure 6, the right wing experience a higher angle of attack and this increases the lift experienced by the right wing. The left wing, on the other hand, experiences a much lower angle of attack and thus the lift decreases. This cause a restoring moment to the left which brings the aircraft back to equilibrium [3].

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3.5 Drag

Drag is often divided into two main parts; the induced drag and the zero-lift drag. The induced drag is associated with the generation of lift and it is dependent on the AOA of the UAV. The result of induced drag comes from the downwash effect, where the trailing vortices generated at the wing tips induces a small downward velocity component in the downwards direction of the wing. The zero-lift drag is contributed by the various parts of the UAV which is not dependant on production of lift such as the horizontal tail, fuselage, wings, and vertical tail and so on.

4. DESIGN

4.1 Selection of Aerofoil

The aerofoil shape of the wings has to be selected based on the criteria that the main wings must be capable of providing enough lift to support the weight of the entire UAV. A conventional wing and tail configuration will be adopted. The

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aerofoil shape of the tail has to be selected such that it provides a downward force which balances the lift force produced by the main wings.

The selection of the wing aerofoil was done using a 2D computational program known as the Xfoil. The Xfoil consists of a wide database for NACA aerofoils and the Cl, Cd and Cm of the aerofoil can be easily determined. The credibility and accuracy of Xfoil has been verified and was deemed to be accurate. Thus, the use of Xfoil sped up the process of selecting and narrowing down the suitable aerofoils for the UAV. Given a large database to choose from, 5 aerofoils were selected, namely the SPICA, NACA 4415. NACA 8414, NACA 0012 and SD 8020 [10]. The information of the respectively aerofoils selected were plotted using Xfoil and the graphs are as shown below.

SPICA

1.5

1

Cl

0.5

0

-100AOA 1020

-0.5

12

Cl

-10

NACA 44151.8

1.6

1.4

1.2

1

0.8

0.6

0.4

0.2

0

-0.2 01020

-0.4AOA

NACA 8414

2.5

2

1.5

Cl

1

0.5

0 -10 0 10 20

AOA

13

SD 8020

1.2

1

0.8

0.6

Cl

0.4

0.2

0

05101520

-0.2AOA

NACA 0012

1.4

1.2

1

0.8

Cl0.6

0.4

0.2

0

-0.205101520

AOA

Figure 7: Lift characteristics of selected aerofoils

The Cl obtained from the graph plotted above has to be adjusted for aspect ratio and converted to the estimated 3D CL using the Lifting Line theory:

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The Cl can be obtained by finding the gradient of the graph plotted in Figure 7. A high aspect ratio is preferred for the wings as it provides better 3D lift characteristics. It produces less drag at lower speeds.

4.2 Design of UAV

A conventional design will be adopted for the UAV; main wings and tail. The main wings serve to provide lift for the entire UAV and the tail will be used to balance the aircraft.

To calculate the lift of the UAV provided by the wings, the equation shown below can be used.

In order to find the lift of the wings, the chord length, wing span, aspect ratio are estimated and after which the dimensions of the wings are determined by the optimisation program. The known values are then factored into the lift equation to obtain the total lift generated by the wings.

The main wings of the UAV will be placed on the same level as that of the tail. This would induce a downwash effect which will decrease the effective angle of attack of the tail.

4.3 Wing Aerofoil and Wing Size

The selected wing aerofoils are namely SPICA, NACA 4415 and NACA 8414. The characteristics of each aerofoil are summarised below:

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120Cl/Cd vs AOA

100

80

/CdNACA

608414

ClNACA

404415

SPICA

20

0

05101520

AOA

Figure 8: Cl vs Cd comparisons for the selected three aerofoils using Xfoil

Table 3: Wing Aerofoil characteristics

StallCl at MaxShape of

AerofoilMax ClStall AngleCl/Cd at

Pattern= 6oaerofoil

6o

SPICA1.2913oGradual1.0764.8Flat Bottom

15oRelatively

NACA 44151.54Sharp1.11103.6Flat Bottom

NACA 84141.9513oGradual1.49109.2Moderately

Cambered

Initially, SPICA and NACA 4415 were chosen as the wing aerofoils due to the shape of the aerofoil and the ease of fabrication. However, having done the Solidworks Flow Simulation on these two aerofoils, it did not provide sufficient lift for the UAV.

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Figure 9: Flow analysis across the Solidworks model

NACA 8414 was eventually chosen as the main wings aerofoil. This can be seen from Table 3, the aerofoil characteristics of NACA 8414 best fits the criteria for the UAV.

The wingspan and wing chord of the UAV are limited by size. A high aspect ratio is preferred as it reduces induced drag and increases the 3D CL value. In order to support a weight of 1kg (10N), the wing span was set at its maximum, 100cm wingspan, wing chord to be 13cm and the AOA of the wing will be place at 6o so as to maximise the L/D ratio.

Theoretical calculations were performed on the chosen wing aerofoil, NACA 8414. The lift obtained would be verified using Solidworks Flow Simulation, which will be explained in later parts of the thesis.

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Sample calculation of lift provided by the main wings:

Table 4: Main Wing Specifications

ParametersDimensions

Wing Chord, c13 cm

Wing Span, b100cm

Wing Area, S0.13m2

Aspect Ratio7.69

Coefficient of lift curve slope,0.0945 /o = 5.414 /rad

Angle of incidence,6o = 0.1047 rad

Density, 1.225kg/m3

Velocity, V20m/s

Substituting the above into the 3D CL equation,

Thus, based on the above sample calculation for NACA 8414 3D wing, the calculated lift value provides 40% more lift, providing a safety factor of 1.4. With that, the NACA 8414 was implemented as the main wings aerofoil and the finalised Solidworks model of the UAV will be simulated under the Solidworks Flow Simulation.

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It must be noted that the calculated value of lift comes entirely from the main wings and that the downward force exerted by the tail is not considered. Thus, the total lift of the aircraft is in fact:

4.4 Tail Aerofoil and Tail Size

The selected tail aerofoils are namely SD8020 and NACA 0012. Both aerofoils chosen are symmetrical in shape and its zero lift angle of attack occurs at 0o and thus, it is suitable to act as a tail aerofoil so that when the elevator is deflected downwards, the UAV will have a pitching up moment and vice versa. In addition, a slight negative AOA of the tail aerofoil will be able to provide a balance condition for the aircraft. SD8020 was eventually chosen for the tail aerofoil as it has a lower drag coefficient and the thickness is smaller than NACA 0012.

In order to ensure that UAV is statically stable, the tail configuration has to be such that it balances the lift force provided by the main wings.

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Figure 10: Tail Configuration [4]

The position of CG was determined by placing the CG slightly in front of the 25% wing chord. The distance lt was set at at initial value of 35cm (this was limited by the length of the rod which we used to support the entire UAV). The value of lt can be adjusted based on the anticlockwise moment required by the tail to balance the UAV. The distance lt can be placed further away from the CG to obtain a larger moment arm exerted by the tail downward force.

By taking moments about the CG, the downwards lift generated by the tail can be obtained.

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Sample Calculations:

Table 5: Dimensions for parameters stated in Figure 10

ParametersDimensions

Lift, Lw14.3N

Distance from wing AC to CG, lw1.5cm

Tail coefficient of lift curve slope

(SD8020) ,

Distance from tail AC to CG, lt35cm

Having decided that SD8020 as the tail aerofoil, the downward force required by the tail can be found using the above equation. With that, the CL value can be obtained using the lift equation. By substituting the known values into the equation, the CL can then be used to find the angle of incidence of the tail using the 3D CL equation.

The horizontal tail area can be estimated by the following:

Horizontal Tail Volume Ratio,

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For glider, VH is estimated to be 0.6 (refer to Figure 19).

Sample calculations for angle of incidence for the horizontal tail:

By substituting into the lift equation,

From the optimisation program, tail AR = 3.11. Using the 3D CL to find the angle of incidence of tail,

(downwards)

4.5 Downwash Angle

Downwash is generated due to the flow over the main wings, generating a small velocity component in the downwards direction due to the wing trailing vortices. The downwash generated by the main wings results in a decrease in the angle of incidence of the tail. The effect of downwash causes a deflection of the flow over the tail in the downwards direction.

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Figure 11: Horizontal Tail effective angle of attack [11]

Where

Based on Figure 17,

5. COMPUTATIONAL FLUID DYNAMICS

Computational Fluid Dynamics (CFD) was performed using Solidworks flow simulation. Flow simulation was used to determine the global goals and surface goals, namely the total lift of the UAV, lift generated by the wing and tail surfaces, drag of the UAV as well as the moment characteristics of the UAV. Refinement of mesh was done on the aircraft to ensure that the curvature and partial cells are properly meshed and laminar and turbulent flow were selected to ensure that the

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flow simulation done in Solidworks was as close to the real life situation as possible. An animation of flow over the UAV was also performed to analyse the flow patterns as it passes the UAV. CFD model from flow simulation is based on time dependent Navier-Stokes equations for reiterative calculations to generate the required data. Due to time constraint and the level of refinement of the cells, CFD was performed on the final Solidworks model of the UAV and used to determine the lift data at different angles of attack.

Figure 12: Basic Mesh of Computational Domain and Refinement of Mesh

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Lift

18

16

14

12

Lift (N)10

8

6

4

2

0

0123456

AOA (degrees)

Figure 13: Lift vs AOA

CFD results shown in Figure 13 indicate that at an AOA of 1o, the lift generated is about 13N at 20m/s. The stalling angle of the UAV is about 5o. The design of the UAV is capable of achieving sufficient lift at a speed of 20m/s, a 30% margin before stalling occurs at 5o.

Table 6: Wing and Tail specifications

Angle of

WingAerofoil ShapeWingspan (cm)Chord (cm)Aspect Ratioincidence

Main WingsNACA 8414100137.696o

Horizontal TailSD 8020289.03.111.22o

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6. ESTIMATED DRAG

The estimated drag of the UAV can be calculated using the formula which includes the induced drag and the zero lift drag coefficients:

The estimated drag of the UAV is made up of the main components such as the fuselage, main wings, horizontal tail and rudder.

e = 0.85, Tail AR = 3.11, Wing AR = 7.69

= 0.0108

The CD0 values were obtained by calculating the fineness ratio for each major component and referenced with the graphs shown in Appendix 4 and 5.

Table 7: Drag components

ComponentsReference Area, AFineness ratioCD

Zero Lift Drag---

Main Wings0.130.140.0065 X 1.5 =

(NACA 8414)0.00975

Horizontal Tail0.0260.100.0058 X 1.5 =

(SD 8020)0.0087

Vertical Tail (SD0.009750.100.0058 X 1.5 =

8020)0.0087

Fuselage0.030X 1.5=

(, l =0.0059451.954

0.045

0.156m)

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With reference to values obtained in Table 7:

From the data logger, ground speed recorded was 17m/s. Assuming linear aerodynamics, the CFD lift value which was obtained at 20m/s can be scaled down to 17m/s using the formula below:

Thus,

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7. STABILITY

7.1 Longitudinal Stability

Static longitudinal stability of an aircraft is crucial as it determines the ability of the aircraft to return to its original position when displaced by a gust of wind or there is a sudden change in the AOA. To ensure that the aircraft is longitudinally stable,

Moments vs AOA0.5

0.4

(Nm)0.3

0.2

Moments

0.1

0

0123456

-0.1

-0.2AOA

Figure 14: Moments vs AOA graph from Solidworks Flow Simulation

Solidworks Flow Simulation is used to determine the CM of the entire UAV. Based on the graph generated above, it can be seen that it is a negative downward sloping graph which cuts the positive y-axis (Cm) and x-axis (AOA). The graph depicts that as AOA increases, the negative pitching moment (nose down) will bring the aircraft back to trim position and as AOA decreases, the positive

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pitching moment (nose up) will pitch up the aircraft to trim position. Trim condition is when the summation of forces is zero (thrust equals drag) and the summation of all moments is zero. Thus, the trim position is when the slope cuts the x-axis (Cm = 0). By generating the Moments vs AOA graph, the position of the Centre of Gravity of the UAV can be determined to ensure trim conditions.

7.2 Lateral Stability

Dihedral effect on the UAV was due to the wing position. The wings of the UAV are positioned at the top of the fuselage, i.e. high wings. The high wings enable greater stability as it guides the position of cross flow around the fuselage in a sideslip, changing the angle of attacks at the wing root and wing tip, resulting in a net lift component which brings the UAV back to stability.

Figure 15: Rolling stability due to dihedral [1]

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9. RESULTS AND EVALUATION

9.1 Verifying Lift

As the UAV is capable of flying, it goes to show that sufficient lift is provided by the main wings. Given that the weight of the UAV is approximately 1kg, the CL value can be verified from the straight and level flight at ground speed of 17m/s.

The theoretical CL calculated was 0.45 and the slight reduction in CL could be due to wind conditions, the accuracy of the shape of the aerofoil and also interference from the aircraft body due to presence of fuselage, wing spars and so on.

9.2 Lift-to-Drag Ratio

One of the most important criteria when building the UAV is the lift-to-drag ratio. Maximum lift occurs near the stall angle and minimum drag occurs at low angles of attack for cambered aerofoil. Thus, to get the maximum lift-to-drag ratio, the lower lift value has to be compensated to achieve a lower drag value. The UAVs required lift is determined by the total weight of the UAV. In this case, a minimum of 10N of lift is required from the main wings to support the UAV. Lower drag is obtained by ensuring a smooth surface of the wings such that flow separation is delayed. Lower drag leads to better climb performance and better glide ratio. In addition, it also reduces the thrust of the motor needed to overcome the drag.

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Theoretical lift-to-drag ratio can be determined through CFD analysis, where the UAV is assumed to be operating at maximum CL.

The theoretical value of L/D can be compared with that of the powerless glide which was conducted during one of the flight tests. The latitude, longitude and altitude of the UAV were recorded during the test, where the UAV was travelling at 17m/s.

Figure 16: Powerless Glide Test

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The difference between the actual and theoretical L/D could be due to the increase in drag caused by the presence of the folding wing mechanism which was placed on the top of the fuselage. The rough surfaces of the mechanism, protruding wires, camera and drag from exposed wing spars at the joints could have increased the total drag of the UAV significantly and these values of drag were not accounted for in the estimated drag calculation due to its complexity in shape. Interference drag, caused by the connection between the folding wing mechanism and the wings were not taken into account during calculation.

In addition, as there is no angle of attack indicated mounted on the UAV, it is not possible to determine the AOA the plane is flying at. The ground speed data collected by the GPS is subjected to a 3m error and this contributes to the error in calculating range and altitude (see Appendix 7).

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10. RECOMMENDATIONS FOR FUTURE WORK

With more time and resources, the Solidworks Flow Simulation can be further enhanced by increasing the mesh and refinement number. This would provide a more accurate data.

The lift of the UAV provided by the wings can be increased by having a canard configuration. Given more time, the canard configuration can be explored and the use of canard may be implemented in our design. With the addition of a canard, it may allow for a shorter main wingspan and hence easier portability of the UAV. In addition, in terms of the material used to manufacture the UAV, the foam can be of smoother surface so as to reduce drag and lighter in order to reduce weight.

Drag can also be further reduced by streamlining the fuselage and ensure that the exposed surfaces such as the wing joints and spars are covered with a smooth layer.

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11. CONCLUSION

In conclusion, the aerodynamics aspects of the UAV were met. The entire process from the selection of aerofoils to the implementation of the wings and tails on the actual UAV were discussed in the thesis. Several prototypes were built to improve on the design of the UAV. In addition, flight tests that were conducted proved the airworthiness of the UAV and also ascertain the stability of the plane. With the equipment placed on board, the wings and tails were capable of providing lift and balance respectively during flying.

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REFERENCES

1. Flight Emergency & Advanced Maneuvers Training. (2009).

Lateral/Directional Stability. Flightlab.

2. Johnson, P. K. (2006). Airfield Models. Retrieved 30 August , 2011, from http://www.airfieldmodels.com/information_source/index.htm

3. Lancaster, R. (1983). Principles of Glider Flight. Alexander Schleicher GmbH & Co.

4. Leng, G. (2012). Static Longitudinal Stability. National University of Singapore.

5. Lennon, A. (1996). In A. Lennon, The Basics of R/C Model Aircraft Design. Wilton: Air Age Inc.

6. Leong, J. Y. (2010/2011). Unmanned Aerial Vehicle Design and Manufacture. Singapore: National University of Singapore.

7. Luo, S. C. (2012). Aerodynamics and Propulsion. National University of Singapore.

8. McCormick, B. (1995). Aerodynamics, aeronautics and flight mechanics.

John Wiley and Sons, Inc.

9. Model Aircraft. (14 April, 2011). Retrieved 5 January, 2012, from Aerodynamics, Beginners' Guide : http://adamone.rchomepage.com/index.html

10. N.A. (25 February, 2012). Airfoil Investigation Database. Retrieved 12 August, 2011, from Airfoil Investigation Database: http://worldofkrauss.com/

11. N.A. (n.d.). Principles of Helicopter Flight. Retrieved 4 January, 2012, from http://www.cavalrypilot.com/fm1-514/Ch1.htm

12. NASA. (25 March, 2010). National Aeronautics and Space Administration. Retrieved 28 February, 2012, from Lift to Drag Ratio: http://wright.nasa.gov/airplane/ldrat.html

13. Sadraey, M. Tail Design. Daniel Webster College.

14. Sadraey, M. Drag and its Coefficient. Daniel Webster College.

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Appendix 1

Figure 17: 2D NACA 8414 Aerofoil Characteristics generated using Xfoil

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Appendix 2

Figure 18: Downwash angle for no sweepback

Based on Figure 18,

X-axis of graph =

Y-axis of graph == 0.58, where = 0 and A = Aspect Ratio (for this case)

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Appendix 3

Figure 19: Typical values for horizontal and vertical tail volume coefficients

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Appendix 4

Figure 20: Coefficient of Drag for 4-digit aerofoils at low Reynolds number

Figure 21: Coefficient of Drag of fuselage vs Fineness ratio

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Appendix 5

Figure 22: Example of Drag Breakdown for Piper Cherokee 180

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Appendix 6

Figure 23: Refinement of cells and curvature

Table 8: Sample results obtained from Solidworks Flow Simulation

AveragedMinimumMaximum

Goal NameUnitValueValueValueValue

GG Y - Component

of Force 1[N]12.0038987712.3998340811.8019926213.04796569

Wing[N]11.6330128912.0289647611.4425748412.65949167

----

Tail[N]0.0047834190.0185136930.0306122740.003170438

GG Z - Component

of Force 1[N]3.4981350793.540378953.4948358523.590902445

GG X - Component

of Torque 1[N*m]2.2407236172.3130952972.2053409532.427614326

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Appendix 7

Figure 24: Altitude vs Line Number from ArdumegaPilot Mission Planner

Figure 25: Latitude, Longitude and Altitude of the UAV during Glide Test

Figure 26: Latitude and Longitude conversion to distance (m)

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Appendix 8

Table 9: Wing Parameters

Wing Chord,0.13 m

Wingspan,1.00 m

Wing Area,0.13 m2

Wing Aspect Ratio,7.692

Wing Angle of Incidence,0.105 rad

Wing Angle of Attack,0.174 rad

2D Wing lift coefficient,1.49

3D Wing lift coefficient,0.45

3D Lift coefficient at 0 AoA,0.425

2D Wing Lift curve slope,5.414

2D Wing moment coefficient,1.936

Distance from Wing leading edge to CG,0.030 m

Distance of ac from Leading edge,0.0325 m

Table 10: Horizontal Tail Parameters

Horizontal Tail Chord,0.09 m

Horizontal Tail Span,0.28 m

Horizontal Tail Area,0.0252 m2

Horizontal Tail Aspect Ratio,3.11

Horizontal Tail Angle of Incidence,0 rad

Distance of Tail ac from cg,0.35 m

2D Tail lift coefficient,0.7292

3D Tail lift coefficient,0.09262

2D Tail lift curve slope,6.607

Downwash angle,0.037 rad

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Rate of change of downwash angle,0.448

Horizontal Tail Volume Ratio,0.522

Table 11: Vertical Tail Parameters

Vertical Tail mean chord0.075 m

Vertical Tail Area,m2

Vertical Tail Span,0.13 m

Distance of Tail ac from cg,0.35 m

Vertical Tail Volume Ratio,

-END OF THESIS-

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