inverted delta wig effect aircraft 2

Aerodynamics and Propulsion for an Inverted Delta Wing-In-

Ground Effect Aircraft

Submitted by: Jiang Junde

Department of Mechanical Engineering

In partial fulfillment of the requirements for the Degree of Bachelor of Engineering

National University of Singapore

Session 2005 / 2006

Aerodynamics and Propulsion of an Inverted Delta Wing In Ground Effect Aircraft

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Summary The objective of this project is to design, fabricate and perform flight tests to

investigate the Wing-In-Ground (WIG) effect of an Inverted Delta Wing

configuration aircraft. It involves the study of aerodynamics and stability

characteristics on a model aircraft, using both software simulations and actual

flight tests to determine the characteristics.

This project can be broadly categories into 3 major milestones and they are:

1) Design and simulations using computational and 3-D engineering software

2) Fabrication and laboratory-based experimental testing

3) Actual flight testing under various conditions

Literature review was first done to find out more about the WIG aircrafts and their

characteristics. It was followed up with further studies on the aerodynamic

theories needed to design and compute the WIG model aircraft’s various

parameters. Gambit was used to create the CAD model of the aircraft and Fluent

was used to run the CFD simulations in order to determine the optimal

parameters of the aircraft. The model aircraft was then fabricated with the

parameters obtained from the simulations. The propulsion system was selected

based on the flying speed and the drag computed by the CFD simulations. A

control system was planned with the objective of attaining static and dynamic

flight stability, flight manoeuvrability and automatic height control. Experiments


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were done on the various control system components to determine their

respective characteristics.

After getting the optimal parameters, a model aircraft was built and the

performance of the aircraft was evaluated by doing test flights at various places

and conditions, with on board sensors to verify the results predicted by the CFD

simulations. Both land and water flight tests were carried out to verify its

amphibious capabilities. The aircraft had demonstrated great versatility in its

control and maneuvering during the numerous field tests conducted.

In the short 9 months, the project team had successfully attained the initial

objectives of designing, analyzing, fabricating and achieving a working WIG

aircraft. The stability theory and test flight results were also presented at the NUS

Centennial Open House 2006.

In conclusion, the project has successfully demonstrated the capabilities of an

Inverted Delta Wing configuration WIG aircraft and its immense potential in the

field of high speed marine transport, for both commercial and military usage.


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Acknowledgement

The author wishes to express his thanks and heartfelt gratitude to the following

persons for their various contributions and important assistances rendered during

the project.

• A/P Gerard Leng Siew Bing, Project Supervisor, for providing the necessary

guidance and invaluable advice throughout the course of the project.

• Mr Lee Qihui, Project Member and Friend, for his contribution and effort in the

stability and control aspect of the project..

• Mr Ahmad Bin Kasa, Ms Amy Chee, Ms Priscilla Lee and Mr Cheng Kok

Seng, Staff of the Dynamics & Vibration lab, for their help and support during

the project.

• Mr Kam Mun Loong, Mr Tan Han Yong, Mr Oi Tze Liang and Mr Teoh Wei Lit,

PHD and Masters Students of COSY lab, for their help and encouragement.

• Mr Tan Gee Boon, Nicholas and Mr Ng Kah Yong, Filming Crew and Friends,

for providing relentless assistances in filming the various test flights.

• Ms Lim Weiyee, Web page Designer and Friend, for her help in designing and

developing the website for this WIG project.


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Table of Contents Acknowledgement. ........................................................................................................... III

List of Figures and Tables................................................................................................. VI

List of Symbols ………………………………………………………………………...VII 1 Introduction............................................................................................................- 1 -

1.1 Project Objectives ...........................................................................................- 1 -

1.2 Structure of the Dissertation ...........................................................................- 2 -

2 Fundamentals of Ground Effect aerodynamics......................................................- 3 -

2.1 Chord Dominated Ground Effect (CDGE) .....................................................- 3 -

2.2 Span Dominated Ground Effect (SDGE)........................................................- 5 -

3 Preliminary CFD Analysis.....................................................................................- 7 -

3.1 CFD – Basic Background Information ...........................................................- 7 -

3.2 Pre-processing.................................................................................................- 8 -

3.3 Numerical Schemes ......................................................................................- 10 -

3.3.1 SIMPLE ................................................................................................- 11 - 3.3.2 Upwind Scheme .................................................................................- 11 -

3.4 Accuracy of CFD simulations results ...........................................................- 11 -

3.5 Variables in the CFD simulations .................................................................- 13 -

4 Design of the Prototype .......................................................................................- 15 -

4.1 First Weight Estimation ................................................................................- 16 -

4.2 Fuselage Design ............................................................................................- 16 -

4.3 Wings Design................................................................................................- 17 -

4.3.1 Determining the Optimal Anhedral Angle .......................................- 18 - 4.4 Angle of Incidence of the Wings ..................................................................- 19 -

4.4.1 Cruising Height/Chord Ratio.............................................................- 19 - 4.4.2 Determining the Angle of Incidence ................................................- 20 -

4.5 Horizontal Stabilizer .....................................................................................- 21 -

4.6 Control Surfaces............................................................................................- 23 -

4.7 Aerodynamic Characteristics ........................................................................- 24 -

5 Propulsion System ...............................................................................................- 28 -

5.1 Determining Maximum Drag........................................................................- 28 -

5.2 Motors Selection ...........................................................................................- 30 -


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5.3 Experimental Verification of Motor’s Technical Specifications ..................- 32 -

5.4 Determining the Thrust of the Motor............................................................- 33 -

6 Fabrication and Integration of the Prototype .......................................................- 34 -

6.1 Integration of the Prototype ..........................................................................- 35 -

7 Results and Analysis of the Flight Test Results...................................................- 37 -

7.1 Indoor Flight Test .........................................................................................- 38 -

7.2 Outdoor Flight Test.......................................................................................- 41 -

8 Conclusion and Recommendations......................................................................- 44 -

8.1 Recommendations.........................................................................................- 46 -

References ……………………………………………………………………………- 48 -

Appendix A: Historical Development in WIG………………………………………..- 51 -

Appendix B: Fundamentals of Fluid Mechanics ……………………………………..- 53 -

Appendix C: Pressure Correction Method..…………………………………………..- 56 -

Appendix D: AXI 2814/12 Motor Specifications.........................................................- 58 -

Appendix E: Motor Thrust Experiment ........................................................................- 58 -

Appendix F: Graphs of Height Readings (cm) versus Time (s) ……………………..- 60 -

Appendix G: Detail Mass Breakdown of Components……………………………... - 61 -

Appendix H: Determination of Experimental Angle of Attack ..………...…………. - 62 -

Appendix I: Tabulation and Graphs of CFD Simulations ………………………….. - 63 -


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List of Figures

Figure 2.1: Contour plot of static pressure on an airfoil …………………………- 4 -

Figure 2.2: Vortex strength of an aircraft in flight…………………………………- 6 -

Figure 3.1: Geometry and Mesh for Overall Flow Domain………………………- 9 -

Figure 3.2: Mesh of a wing-fuselage-tail model…………………………………- 10 -

Figure 3.3: CL and CD vs. number of iterations when TOL is 10-5 ..…………. - 12 -

Figure 3.4: Graph of CL versus Degree of Refinement ...…………………….. - 13 -

Figure 4.1: Graph of Lift/Drag ratio versus Anhedral Angle ..…………………- 19 -

Figure 4.2: Graph of Lift/Drag Ratio versus Height/Chord Ratio………………- 20 -

Figure 4.3: Graph of Lift/Drag ratio versus Angle of Attack....…………………- 21 -

Figure 4.4: Graph of Cm versus Angle of Attack..…………...…………………- 22 -

Figure 4.5: Graph of Moment versus Deflection Angle for Rudder…………...- 23 -

Figure 4.6: Graph of Moment versus Deflection Angle for Elevator…………..- 24 -

Figure 4.7: Graph of Coefficient of Lift versus Angle of Attack………………..- 25 -

Figure 4.8: Graph of Coefficient of Lift at α=0º versus Height/Chord Ratio….- 25 -

Figure 4.9: Graph of Coefficient of Lift Gradient versus Height/Chord Ratio...- 26 -

Figure 5.1: Submerged portion of the fuselage at rest…………………………- 30 -

Figure 5.2: Graph of Thrust (N) versus Throttle (%)……………………………- 33 -

Figure 6.1: Photographs of the prototype’s skeleton structure………………..- 34 -

Figure 6.2: Schematic layout of the prototype in plan view……………………- 35 -

Figure 6.3: Photograph of the FS8 and the prototype side by side...…………- 36 -

Figure 7.1: Screenshots of the prototype flying in the MPSH 2……………….- 39 -

Figure 7.2: Graph of CL versus Angle of Attack………………………………....- 39 -

Figure 7.3: Graph of CD versus Angle of Attack…………………………………- 40 -

Figure 7.4: Screenshots of the prototype flying over water……………………- 42 -

List of Tables Table 4.1: First Estimation of mass breakdown of components ……………………..- 16 -

Table 5.1: Comparison of power generated with different propellers………………. - 30 -

Table 7.1: Velocity of the Prototype in Air (land takeoff)………………………….. - 38 -

Table 7.2: Velocity of the Prototype in Air (water takeoff)………………………… - 41 -


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LIST OF SYMBOLS c Chord Length

h Height

h/c Height to Chord Ratio

CL Coefficient of Lift

CD Coefficient of Drag

CM Coefficient of Moment

α Angle of Attack

θ Pitch Angle

δE Angle of Deflection of Elevator

δR Angle of Deflection of Rudder

δA Anhedral Angle

xcp Center of Pressure

xa/c Aerodynamic center

xθ Aerodynamic center of Pitch

xh Aerodynamic center of Height

XCG X-Coordinate of C.G, with origin at Point A

YCG Y-Coordinate of C.G, with origin at Point A

ZCG Z-Coordinate of C.G, with origin at Point A

b Wing Span

AR Aspect Ratio

Re Reynolds Number

U Reference Velocity / Free Stream Velocity

L Reference Length

ν Kinematics Viscosity

ρ Density

CMα Derivative of the Coefficient of Moment With Respect To Angle of Attack

CM0 Y intercept of the Coefficient of Moment

CLα Derivative of the Coefficient of Lift With Respect To Angle of Attack

CMh Derivative of the Coefficient of Moment With Respect To Height


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CLh Derivative of the Coefficient of Lift with Respect To Height

CMwf Characteristic Curve of the Coefficient of Moment for Wing-Fuselage

Combination

CMt Characteristic Curve of the Coefficient of Moment for Tail

CMwft Characteristic Curve of the Coefficient of Moment for Wing-Fuselage-Tail

Combination

VH Tail Volume Ratio

CMαt Derivative of the Coefficient of Moment of the Tail With Respect To Angle

of Attack

C M0t Y Intercept of the Characteristic Curve of the Coefficient of Moment of the

Tail

CLαt Derivative of the Coefficient of Lift of the Tail With Respect To Angle of

Attack

ε Downwash Angle

iw Wing Angle of Incidence

it Tail Angle of Incidence

lt Distance between the C.G and a/c of tail

St Area of tail

η Tail efficiency

τ Control Surface effectiveness parameter

Mp Maximum Overshoot

ts Settling Time

tr Rise Time

ξ Damping Ratio

ωn Natural Frequency


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1 Introduction Ground Effect is a phenomenon when a lift generating device, like a wing, moves

very close to the ground surface which increases the lift-to-drag ratio. This

phenomenon which results in an increase in the aerodynamic efficiency of the

vehicles was first exploited by the Russians, whom designed and build the first

WIG craft during the cold war to transport their troops and supplies. However

these WIG crafts were huge planes that were built for military purposes and it is

only in recent years that there is focus on small scale commercial WIG crafts.

The amount of written literature regarding such small scale crafts is limited, and

thus this project was initiated. This project aims to design, analyze, fabricate and

test fly a small scale wing in ground model aircraft to investigate and

demonstrate the effects of ground effect.

This assignment is a joint effort of 2 final year project students covering the areas

of aerodynamics, propulsion, stability and control. In this thesis, the

aerodynamics and propulsion aspects of the WIG aircraft will be discussed and

presented.

1.1 Project Objectives The following objectives are to be achieved:

• Literature review and theoretical study of WIG aircraft’s aerodynamics

characteristics.



• Determining the optimal design and operating parameters for the WIG

model aircraft through the use of CAD and CFD simulation software.

• Determine the propulsion system.

• Fabrication of the WIG model aircraft.

• Test flight of the fabricated prototype.

• Verification of theoretical results against actual flight performance

1.2 Structure of the Dissertation This thesis is divided into 8 Chapters and they are organized as follows:

Chapter 2 – Fundamentals of Ground Effect aerodynamics.

Chapter 3 – Preliminary CFD analysis.

Chapter 4 – Design of the prototype.

Chapter 5 –Propulsion system.

Chapter 6 – Fabrication and Integration of the Prototype.

Chapter 7 – Results and analysis of the flight test data.

Chapter 8 – Project Conclusion and Recommendations for further study

Stability and Control of an Inverted Delta Wing In Ground Effect Aircraft


2 Fundamentals of Ground Effect aerodynamics

When a wing approaches the ground, an increase in lift as well as a reduction

in drag is observed which results in an overall increase in the lift-to-drag ratio.

The cause of the increase in lift is normally referred to as chord dominated

ground effect (CDGE) or the ram effect. Meanwhile, the span dominated

ground effect (SDGE) is responsible for the reduction in drag. The

combination of both CDGE and SDGE will lead to an increase in the L/D ratio

hence efficiency increases.

2.1 Chord Dominated Ground Effect (CDGE)

In the study of CDGE, one of the main parameters which one considers is the

height-to-chord (h/c) ratio. The term height here refers to the clearance

between the ground surface and the airfoil or the wing. The increased in lift is

mainly because the increased static pressure creates an air cushion when the

height decreases. This result in a ramming effect whereby the static pressure

on the bottom surface of the wing is increased, leading to higher lift. Fig. 2.1

shows the difference between an airfoil without ground effect (a) and with

ground effect (b). Theoretically, as the height approaches 0, the air will

become stagnant hence resulting in the highest possible static pressure with a

unity value of coefficient of pressure.


Figure 2.1: Contour plot of static pressure on an airfoil; a) out of ground effect, b) in ground

effect.

Following the convention of the study of aerodynamics, the solutions of the

aerodynamic forces, Lift (L) and Drag (D), and moment (M) are normally

presented in a form of dimensionless coefficient which are define as the

following:

Where ρ∞ is density of air, S is projected area on ground plane, V is free

stream velocity and c is the chord length. Rozhdestvensky[1] has predicted for

a case a flat plate with infinite span in the presence of extreme ground effect

(h/c < 10%), a closed form solution for CL and CM can be obtained by a

modification to the thin airfoil theory and the solutions are given as:



In equation 2.5, the coefficient of moment is taken with respect to the leading

edge. By taking the moment at the leading edge, the center of pressure, xp is:

Hence unlike the case of a symmetrical airfoil out of ground effect, the center

of pressure is at one-third of the cord instead of one-forth. Coincidentally, for a

symmetrical airfoil, the center of pressure coincides with the aerodynamic

center. This is however not true for a cambered airfoil.

2.2 Span Dominated Ground Effect (SDGE)

On the other hand, the study of SDGE consists of another parameter known

as the height-to-span (h/b) ratio. The total drag force is the sum of two

contributions” profile drag and induced drag. The profile drag is due to the

skin friction and flow separation. Secondly, the induced drag occurs in finite

wings when there is a ‘leakage’ at the wing tip which creates the vortices that

decreases the efficiency of the wing. In SDGE, the induced drag actually

decreases as the strength of the vortex is now bounded by the ground. As the

strength of the vortex decreases, the wing now seems to have a higher

effective aspect ratio as compared to its geometric aspect ratio (b2/S),

resulting in a reduction in induced drag.



Figure 2.2: Vortex strength of an aircraft in flight; Left: Out of ground effect. Right: In ground effect

From Prandtl’s lifting line theory [2], the induced drag can be calculated by

Where e is known as the span efficiency and AR is the aspect ratio. In the

presence of ground effect, Rozhdestvensky [3] shows that e α 1/h, hence from

equation 2.7,

From Equation 2.8, it can be shown that the induced drag will decrease

linearly with height.




3 Preliminary CFD Analysis

In order to fabricate the prototype, most of the parameters of the model have

to be obtained either through computational or experimental methods. By

using a computational method like Computational Fluid Dynamics (CFD)

analysis, the process of determining these optimal parameters will be faster

as well as cheaper due to the ease of adjusting the CAD models and

simulation parameters versus actual fabrication of experimental models.

Another advantage of using CFD is its ability to perform flow visualization. Air

being invisible, under normal circumstances, the human’s naked eye is unable

to see how the air behaves. Typically, flow visualization is being carried out

either in a smoke tunnel or water tunnel. But with CFD, flow can be visualize

by analyzing the velocity vector plots and injecting tracking the particles being

injected into the simulation and by observing the flow pattern will enable a

better understanding of the physics of the flow.

3.1 CFD – Basic Background Information

The essence behind CFD is to solve the governing equations for fluid (the

Navier-Stoke’s equations) which normally take the form of integral or partial

differential equations using numerical methods. The non-dimensional form of

the incompressible Navier-Stoke’s equation can be written as (See Appendix

B for derivation):


In general, analytical solutions to the highly non-linear Navier-Stokes equation

are difficult to obtain, CFD is therefore needed to obtain a set of numerical

solutions and this was done using Fluent, a commercial CFD code based on

the Finite Volume Method.

3.2 Pre-processing Before the CFD simulations can be done in Fluent, the CAD models of the

prototype will have to be modelled out in a CAD program called GAMBIT. The

program is also used for mesh generation and implementing boundary

conditions. Due to the complex shape of the prototype, unstructured mesh is

used due to its adaptability. Mesh density control is also apply in order to save

computational power and time by having coarser grids at the boundaries of

the domain and finer grids near area of interests and where the geometries

are more complex. In addition, to avoid generating any highly skew mesh,

mesh control is also needed to ensure that the transition from fine to coarse

mesh is smooth.



Figure 3.1: Geometry and Mesh for Overall Flow Domain

Although WIG aircrafts are by nature “sea planes”, the physics behind the

interaction between the craft and air-water interface is very complex to model.

With over a hundred simulations needed to determine the optimal parameters,

it will not be feasible to run the time consuming two phase flow simulations,

especially when the undulating surface effect is actually negligible [5]

according to the literature review that was done. Thus in order to cut down the

computational effort, the boundary condition of the ground is assume to be a

hard moving wall as shown in Fig. 3.1.

Being a subsonic flow, due to the elliptic nature of the governing equation, the

propagation of disturbances can be felt throughout the domain. To reduce any

numerical error from being introduced, the outer boundaries are place far

away from the model. In addition, in order to compensate for the large domain



and to reduce the computational effort, symmetry boundary condition will be

use on the plane of symmetry of the model for the case of a 3D flow analyses.

Figure 3.2 shows an example of the mesh across the wing-fuselage-tail

combination of the craft.

Figure 3.2: Mesh of a wing-fuselage-tail model

3.3 Numerical Schemes The numerical scheme chosen to discretize the pressure equation 3.1 and the

momentum equation 3.2 are the semi-implicit method for pressure-linked

equations (SIMPLE) and the second order upwind scheme respectively. The

reasons are given as follow:




3.3.1 SIMPLE Equation 3.2 is the transport equation for the velocity components. However,

unlike compressible flow, there is evidently no transport equation for pressure

as the pressure terms only appears in the momentum equations 3.2 but not

3.1. Therefore when equation 3.2 is solved to obtain the solutions for velocity,

these solutions will not satisfy the continuity equation 3.1. The SIMPLE

scheme [6], which is an iterative process, is developed to correct the pressure

field so as to obtain the correct velocity field which will satisfy the continuity

equation.

3.3.2 Upwind Scheme Another problem faced during the process of solving incompressible flow

equation is that if an oscillating pressure field is present in the fluid, the

application of standard central difference scheme on the pressure derivatives

will cause these fluctuating or zigzag effects to be not reflected in the

momentum equation. One proposed solution to take care of the fluctuation is

to use a staggered mesh. However, this technique can only be used on

structured mesh therefore the alternative solution to this is to use the upwind

scheme (See Appendix C for more details).

3.4 Accuracy of CFD simulations results To ensure proper convergence of the solutions, a few arbitrary simulations

was done to find out the minimum required tolerance value needed for

convergence criteria. Since the lift and drag are the two most important

parameters needed, the solutions of the two parameters are observed with

different tolerance value. When the fluctuation of the lift and drag are


sufficiently small in the next successive steps of iterations, the solutions are

said to have converged sufficiently.

From the study shown in Fig. 3.3, it is found that the default tolerance value of

Fluent, 10-3, is insufficient. To ensure a more accurate solution is obtained,

the tolerance must be set at around 10-5.

Figure 3.3: CL and CD vs. number of iterations when TOL is 10-5

It is also important to ensure that the number of meshes used is sufficient to

get accurate solutions. By using a standard domain and mesh distribution, the

degree of mesh refinement was increased gradually in Fluent until the solution

for lift and drag has sufficiently small fluctuation. It was observed that after 6

degrees of refinement from fig. 3.4, ie each original mesh was divided into 7

smaller meshes, the solution has negligible fluctuation. Thus all the future

simulations were done with the same domain size, mesh distribution and at 6

degree of refinement.



Figure 3.4: Graph of CL versus Degree of Refinement

Coefficient of Lift vs Degree of Refinement(Anhedral Angle = 6 , Wing angle of incidence = 0, Angle of Attack = 0))

0.08

0.085

0.09

0.095

0.1

0.105

0.11

0.115

0.12

0 1 2 3 4 5 6 7 8 9

Degree of Refinement

Coef

ficie

nt o

f Lift

Series1Poly. (Series1)

3.5 Variables in the CFD simulations After validating the scheme needed for accurate CFD simulations solutions, it

can now be applied to investigate the characteristics and parameters of the

prototype. However, the three important parameters, Lift, Drag and Moment

are dependent on a number of variables:

L = f (ρ, V, S, ν, α, h, c, δA) - (3.3a)

D = f (ρ, V, S, ν, α, h, c, δA) - (3.3b)

M = f (ρ, V, S, ν, α, h, c, δA) - (3.3c)

As it will be time consuming and inefficient to run the CFD simulations based

on all the variables above, dimensional analysis has been applied to reduce

the number of variables to a few dimensionless parameters. The set of

dimensionless parameters can be obtained using the Buckingham pi’s

theorem [7] and the above equations will be reduced to:




CL = f (Re, α, h/c, δA) - (3.4a)

CD = f (Re, α, h/c, δA) - (3.4b)

CM = f (Re, α, h/c, δA) - (3.4c)

Therefore instead of seven variables, only four variables are needed for the

computation to obtain the characteristic of the WIG craft. However, the

operating range of the Reynolds number is expected to be small as the

prototype is assumed to be operating within 10 m/s to 15m/s. The range of Re

is thus given by:

1.3×105 < Re < 4×105

Since the Re range is within the same order of magnitude, variables can now

be further cut down to three, height/chord ratio, angle of attack and anhedral

angle.

CL = f (α, h/c, δA) - (3.6a)

CD = f (α, h/c, δA) - (3.6b)

CM = f (α, h/c, δA) - (3.6c)

The simulations will begin with analysis on a wing section to determine the

anhedral angle, followed by the entire craft for different angle of attack and

height to chord ratio to obtain the characteristics of the WIG. The results will

be presented and discussed in the next chapter.



4 Design of the Prototype

The design objectives of the prototype must first be defined before the exact

details and procedures can be filled in. The design objectives are as follows:

1. Carry a minimum payload of electronics equipment, power supply and

onboard instrumentation.

2. Able to fly in ground effect mode across both land and water surfaces

3. Maintain a straight and level flight.

4. Speed limit of not more than 15 m/s

5. Modular design for ease of any repair or modification

6. Environmental friendly

To satisfy the last requirement, electric motor is selected over IC Engine as it

does not produce any harmful emissive which pollutes the environment. It

also makes it easier to locate test sites as electric motors are less disruptive

in terms of smell, sound and safety level.

With the objectives in mind, certain design parameters of the model have to

be decided before the CFD simulations can be done to obtain the rest of the

optimal design parameters and the aerodynamics characteristics of the

model. To reduce the cost and fabrication time, off the shelves components

like servos, electric motors and propellers are used. Therefore the size of the

craft is also limited by the availability and the constraints of these

components.



4.1 First Weight Estimation

In order for a plane to fly, the lift force generated by the wings must at least be

equal to the weight of the plane. Thus a rough estimation of the weight of the

plane is done in order to set the minimum lifting force required. The weight of

the plane will directly affect the cruising velocity of the plane, thus from Table

4.1, the simulations will be based on a craft capable of lifting off with a

maximum take off weight of at least 2 kg.

Table 4.1: First Estimation of mass breakdown of components

Components Mass (Kg)

Propulsion System (Propellers, motor(s),

speed controller(s), batteries) 0.750

Fuselage 0.250

Wings & Tail 0.450

Electronics (Wires, servos, receiver) 0.150

Automatic Height Control System 0.400

Total Mass 2.000

4.2 Fuselage Design

A batch of NUS mechanical engineering students had previously done a

similar WIG aircraft project with a different wing configuration design. Since

the focus of this project was only to investigate the aerodynamic

characteristics of the inverted delta wing in ground effect mode, thus the

fuselage design adopted was decided to be based on the one used in earlier

project in order to cut down on the time needed for fabrication.



The original fuselage was designed by Mr Toh Boon Whye [17], who had taken

care to minimize the hydrodynamic drag of the fuselage through careful

application of the naval architecture principles. The fuselage was also

designed to be as streamlined as possible according to the physics of low

speed aerodynamics in order to minimize the aerodynamic drag. Similar to a

low speed aerofoil, the nose of the fuselage was made as round as possible

and the trailing edge as thin as possible to allow air to flow around it smoothly

without much abruption. For details on the experiments conducted on the

fuselage design, please refer to Mr. Toh’s thesis.

4.3 Wings Design

Since the prototype is based on a commercial WIG model FS8, some of the

important geometric parameters like airfoil data, taper ratio and aspect ratio of

the wings are already known. The geometric parameter that needs to be

determined through simulations is the anhedral angle. With the fuselage

based on an earlier design, thus the overall size of the prototype is limited by

the size of the fuselage. It was arbitrarily determined that a wing dimension

suitable for the fuselage size would be 12:1 ratio scaled down from the actual

wings size. At this scale, the average chord length was calculated from the

following equation:

cavg = S/b - (4.1)

Thus the average chord length used for all subsequent calculations and

simulations are 0.534 m.



4.3.1 Determining the Optimal Anhedral Angle

Anhedral angle is defined as the angle of the wings below the horizontal.

Having an anhedral angle in the wing will usually cause the plane to be more

unstable laterally in the air, while giving it better lift/drag ratio, and thus

efficiency, in ground effect mode. This phenomenon is not really well

documented in the literature review done, and thus CFD simulations were

done over a range of anhedral angles in order to determine its effect on the

lift/drag ratio.

A total of 20 CFD simulations were done, where the anhedral angle was

varied from 6 to 10 degrees at an interval of 1 degree each, at 4 different

height/chord ratio. The angle of attack of the wing was set at an arbitrary

value of 0 degrees. It was observed that the highest lift/drag ratio occurs

between an anhedral angle of 8 and 9 degrees, as shown in fig 5.1 below,

except for the height/chord ratio of 0.01. It is also noted that as the

height/chord ratio increases, the effect of different anhedral angles is less

significant, given that the curve at the height/chord ratio of 0.025 is relatively

flat.

This implies that when the plane is very low to the ground, having a high

anhedral angle will give a high lift/drag ratio, but as the plane goes slightly

higher, the highest lift/drag ratio falls between anhedral angle of 8 to 9 degree,

and as the plane goes even higher, the effect of anhedral angle on the plane’s

lift/drag ratio becomes less significant. Thus the optimal anhedral angle based

on these simulations lies between 8 to 9 degree. An anhedral angle of 8


degree is chosen as the difference in lift/drag ratio is small, and this will allow

the wingtips to have a higher clearance from the ground.

Figure 4.1: Graph of Lift/Drag ratio versus Anhedral Angle

Graph 4: Lift/Drag Ratio vs. Anhedral Angle - In Extreme Ground Effect(For various h/c, Wing Angle of Incidence = 0 , Angle of Attack = 0)

0

0.5

1

1.5

2

2.5

3

3.5

4

4.5

5 6 7 8 9 10 11

Anhedral Angle

Lift/

Drag

Rat

io

h/c = 0.01h/c = 0.015h/c = 0.02h/c = 0.025h/c = 0.01h/c = 0.015h/c = 0.02h/c=0.025

4.4 Angle of Incidence of the Wings

Wings are usually attached to the fuselage of airplanes at a predetermined

angle of attack, and this angle is known as the angle of incidence. The angle

of incidence of the wing is typically the angle of attack which gives the best

lift/drag ratio, and this will allow for minimum fuselage drag while traveling in

cruising mode. For WIG aircraft, due to the fact that the lift/drag ratio is

strongly dependent on the flying height, thus it is important to decide on the

cruising height/chord ratio before the angle of incidence can be determined.

4.4.1 Cruising Height/Chord Ratio

As mentioned in section 2.1, the lift/drag ratio increases as the height/chord

ratio decreases. This relationship is shown to be accurate by doing a set of 5 National University of Singapore

- 19 -


simulations in fig. 4.2, with varying height/chord ratio whilst other variables are

kept constant. It is thus intuitive that the ideal cruising height should be as low

as possible, however due to certain physical limitations, ie the wings are

already mounted at a certain height above the ground on the fuselage, and

thus the minimum cruising height/chord ratio possible is 0.1. The absolute

flying height of the prototype will thus be 5.34 cm based on an average chord

length of 0.534 m.

Figure 4.2: Graph of Lift/Drag Ratio versus h/c ratio

Graph 5: Lift/Drag Ratio vs. Dimensionless Height (h/c)(For Anhedral Angle = 8, Wing angle of incidence = 3)

8

9

10

11

12

13

14

15

16

17

0 0.05 0.1 0.15 0.2 0.25 0.3Dimensionless Height (h/c)

Lift/

Drag

Rat

io

Wing alone: AA = 8, AoA=3Wing alone: AA = 8, AoA=3

4.4.2 Determining the Angle of Incidence

With the cruising height/chord ratio fixed at 0.1, a set of 4 simulations was

done with varying of the angle of attack. It can be observed in fig. 5.3 that the

highest lift/drag ratio lies between the angle of attack of 4 to 7 degrees. Since

the difference in the lift/drag ratio between 4 and 7 degrees is minimum, an

angle of incidence of 4 degrees was chosen, as it is ideal to keep the angle of

attack as low as possible to prevent flow separation on the wings. National University of Singapore

- 20 -


Figure 4.3: Graph of Lift/Drag ratio versus Angle of Attack

Graph 1: Lift/Drag Ratio vs. Angle of Attack(For Anhedral Angle = 8, h/c = 0.1 Wing angle of incidence = 3)

y = -0.0365x2 + 0.366x + 13.71

y = -0.0214x2 + 0.3458x + 11.062

8

9

10

11

12

13

14

15

16

17

-6 -5 -4 -3 -2 -1 0 1 2 3 4 5 6 7 8 9

Angle of Attack

Lift

/Dra

g R

atio

Wing alone: AA = 8, h/c = 0.1

Wing + hull: AA=8, h/c=0.1

Wing alone: AA = 8, h/c =0.1


4.5 Horizontal Stabilizer

The horizontal stabilizer is needed balance an aircraft longitudinally in order to

achieve stability. It can be either mounted aft of the main wing, which is the

conventional method, or fore of the main wing, which is then known as a

canard. Although a canard design is more efficient, however since the FS8

WIG aircraft utilizes a conventional tail design, thus the prototype will be using

aft horizontal stabilizer in order to be as accurate as possible. In addition to

longitudinal stability, a WIG requires height stability. In order to achieve height

stability, the horizontal stabilizer is normally mounted high above the ground,

out of ground effect.

The detailed analysis of longitudinal and height stability is carried out by Mr.

Lee Qihui. Since the horizontal stabilizer is like a secondary pair of wings



mounted on the tail and is mounted out of ground effect, hence the horizontal

stabilizer will be simulated as a wing in free stream velocity. By varying the

angle of attack and keeping the rest of the parameters constant for CAD

models of wings + fuselage, tail and wings + fuselage + tail, graphs of Cm vs α

can be plotted, which will allow the dimensions, the tail moment arm and

angle of incidence of the tail to be determined. Based on the fig. 5.4 below,

the parameters of the tail can be calculated to give:

b = 0.55 m

c = 0.25 m

AR = 2.2

Tail moment arm = 0.465 m from CG of the prototype

Angle of incidence = 0 degrees

For details regarding the derivation of these parameters, please refer to Mr

Lee’s thesis.

Figure 4.4: Graph of Cm versus Angle of Attack

Coefficient of Moment (CM) vs. Angle of Attack(For Anhedral Angle = 8, h/c = 0.1 Wing angle of incidence = 3)

y = 0.0082x - 0.0178

y = 0.0118x - 0.0229

y = -0.0273x + 0.0846

y = -0.0156x + 0.063

-0.6

-0.5

-0.4

-0.3

-0.2

-0.1

0

0.1

0.2

0.3

0.4

0.5

-8 -6 -4 -2 0 2 4 6 8 10

Angle of Attack

Coef

ficie

nt o

f Mom

ent (

C M) Wing alone: AA = 8,

h/c = 0.1Wing + hull: AA=8,h/c=0.1Tail alone: h/c =0.1

Whole Aircraft

Wing alone: AA = 8,h/c =0.1Wing + hull: AA=8,h/c=0.1Tail alone: h/c=0.1

Whole Aircraft


- 22 -


4.6 Control Surfaces

The presence of control surfaces like a rudder and elevator on an aircraft is to

allow for motion control of the craft while moving, as well as for trimming the

aircraft for stability in reaction to environmental disturbances. The detailed

analysis of this section is handled by Mr Lee Qihui as well. The control

surfaces was simulated at varying angles of deflection while keeping other

parameters constant, and a graph of Moment against Deflection Angle can be

plotted in order to size the servos required for controlling the control surfaces.

Based on the results shown in fig. 4.5 and 4.6, a Hitec hs 55 servo and a

Hitec hs 85 servo is chosen to control the rudder and elevator respectively.

Figure 4.5: Graph of Moment versus Deflection Angle for Rudder

Graph 14: Moment generated by Rudder vs. Deflection Angle

y = 0.0254x

-0.8

-0.7

-0.6

-0.5

-0.4

-0.3

-0.2

-0.1

0

0.1

0.2

0.3

0.4

0.5

0.6

-10 -8 -6 -4 -2 0 2 4 6 8 10

Deflection Angle

Mom

ent g

ener

ated

due

to ru

dder

def

lect

ion

(Nm

)

Rudder Rudder



Figure 4.6: Graph of Moment versus Deflection Angle for Elevator

Graph 17: Moment generated by Elevator abt hinge vs. Elevator Deflection Angle

y = -0.0005x

-0.01

-0.008

-0.006

-0.004

-0.002

0

0.002

0.004

0.006

0.008

0.01

0.012

0.014

0.016

0.018

0.02

-22 -20 -18 -16 -14 -12 -10 -8 -6 -4 -2 0 2 4 6 8 10 12

Deflection Angle

Mom

ent g

ener

ated

(Nm

)

Elevator Elevator

4.7 Aerodynamic Characteristics

After obtaining the geometric and operating parameters for the whole aircraft,

a series of simulations can be done to determine the aerodynamic

characteristics of the prototype by varying the angle of attack and the

height/chord ratio. The objective is to derive a relationship between the

aerodynamic forces vs. angle of attack and height/chord ratio. Fig. 4.7 and

Fig. 4.8 show two different CL curves, one dependant on angle of attack and

another dependant on height/chord ratio. By relating these two curves, a

quantitative expression can be obtained to predict CL at any given angle of

attack and height/chord ratio.


- 24 -


Figure 4.7: Graph of Coefficient of Lift versus Angle of Attack

Graph of Coefficient of Lift vs Angle of Attack

y = 0.0710x + 0.151

y = 0.0632x + 0.142

y = 0.0545x + 0.134y = 0.0495x + 0.128

-0.4

-0.2

0

0.2

0.4

0.6

0.8

1

-8 -6 -4 -2 0 2 4 6 8 10

Angle of Attack (degrees)

Coef

ficie

nt o

f Lift

h/c = 0.075h/c = 0.100h/c = 0.150h/c = 0.200Linear (h/c = 0.075)Linear (h/c = 0.100)Linear (h/c = 0.150)Linear (h/c = 0.200)

Figure 4.8: Graph of Coefficient of Lift at α=0º versus Height/Chord Ratio

Graph of CL, 0 deg vs Height/Chord Ratio

y = 1.0492x2 - 0.465x + 0.1793

0.125

0.13

0.135

0.14

0.145

0.15

0.155

0 0.05 0.1 0.15 0.2 0.25

Height/Chord Ratio

C L, 0

deg

Fig. 4.7 shows a few graphs of CL against angle of attack at different

height/chord ratio, and they can generally be expressed as:

CL = CL, gradient α + CL, 0º - (4.2)



Where CL, gradient is the gradient of the curve, CL, 0º is the coefficient of lift at 0

angle of attack and α is the angle of attack expressed in degrees. The CL,

gradient for various height/chord ratios was plotted in a graph of CL, gradient

against height/chord ratio in fig. 4.9 and their relationship can be expressed

by fitting a cubic curve onto the data.

Figure 4.9: Graph of Coefficient of Lift’s Gradient versus Height/Chord Ratio

Graph of Coefficient of Lift's Gradient vs Height/Chord Ratio

CL, gradient = -8.8(h/c)3 + 4.7(h/c)2 - 0.931(h/c) + 0.1181

0

0.01

0.02

0.03

0.04

0.05

0.06

0.07

0.08

0 0.05 0.1 0.15 0.2 0.25

Height/Chord Ratio

Coe

ffici

ent o

f Lift

's G

radi

ent

Hence CL, gradient for any height/chord ratio can be found by:

CL, gradient = -8.8(h/c)3 + 4.7(h/c)2 - 0.931(h/c) + 0.1181 - (4.3)

The next unknown to be determined will be CL, 0º which can be obtained from

fig. 5.8 for any height/chord ratio and thus CL can be calculated from equation

5.2. To summarize the procedure, to calculate CL for any angle of attack and

height/chord ratio:




1) Determine CL, gradient and CL, 0º for the required height/chord ratio from

equation 4.3 and fig. 4.8 respectively.

2) Calculate CL from CL, gradient and CL, 0º obtained from step 1 for the

required angle of attack using equation 4.2.



5 Propulsion System

In order to size the propulsion system, the first step would be to determine the

maximum drag that the prototype will experience. Since the thrust from the

propulsion system has to be at least equal to the maximum drag, thus the

appropriate motors and propellers can be determined through both their

technical specifications as well as actual experimental testing.

5.1 Determining Maximum Drag Intuitively, the maximum drag experienced will be due to water resistance

while it is attempting to take off, however due to the complex interaction

between the buoyancy force and the lift force with respect to time, the

maximum drag cannot be easily determined. With the lack of a tow tank

facility to determine the drag through experiments on the actual prototype, the

only viable way to determine the drag would be through CFD simulations. By

simulating the submerged portion of the fuselage (when in rest position) at the

cruising speed, the drag obtained is definitely higher than the maximum drag

experienced by the prototype. This is due to the wetted surface area

decreasing as the lift force generated increases when the velocity increases.

Thus the propulsion system sized to this simulated value will definitely be able

to power the plane adequately. Due to the complexity involved in simulating

two-phase flows, the following assumptions were made in order to simplify the

simulation to a single phase flow:

1) Drag due to air is negligible compared to the drag due to water.

2) Parasite drag due to the undulating effect is negligible compared to the

drag due to water.



3) Only the submerged portion of the fuselage at rest will be simulated.

4) Maximum drag experienced is below the amount of drag experienced

at cruising speed with the initial submerged fuselage portion.

The cruising speed of the prototype can be calculated from the following

equation:

V2 = L / (½ ρ∞ S CL) - (5.1)

With the estimated payload at 2 kg, the amount of lift force needed would be

19.62 N, with the coefficient of lift at 0.4 according to fig. 4.7, thus the cruising

velocity of the prototype would be 12.5 m/s.

In order to determine the portion of the fuselage under the water, a simple

experiment was conducted to find the waterline on the fuselage when it was

loaded down with weights at 2kg. The waterline is then transferred onto the

CAD model, and the volume above the waterline is removed. This modified

CAD model, shown in fig. 5.1 below, is then used in the simulation to

determine the water drag, using the cruising speed of 12.5 m/s as its velocity.


Figure 5.1: Submerged portion of the fuselage at rest

The drag obtained from the simulation gives a force of 32.9 N. The amount of

power that the motor(s) and propellers need to supply can be calculated by

the following equation:

Pinstant = Fv - (5.2)

Therefore the amount of power which the propulsion system is required to

provide is 411.6 W.

5.2 Motors Selection

With the maximum drag and power required known, it is possible to begin the

motor(s) selection for the propulsion system. Several criteria were set in order

to make the selection process easier, and they are as follow:

1) Propeller diameter should be as small as possible.




2) Affordable and easily available motors.

3) High power/weight ratio.

Based on the first criteria above, a dual motor configuration was decided on,

as a single motor configuration would require a propeller diameter of at least

14” to provide over 400 W of power. With the motors being mounted above

the fuselage, the motors have to be mounted at least half the propeller’s

diameter above the fuselage. Thus a larger propeller diameter will create a

higher centre of gravity for the prototype, as well as a greater nose down

moment which can possibly affect the longitudinal stability. With a dual motor

configuration, each motor thus has to provide 205.8 W of power.

From second and third criteria, the suitable range of motors was short listed

into 3 models of brushless motors produced by AXI as they are among the

most affordable brushless motors that are easily available in Singapore

compared to brands like Hacker. The short listed motors are the Axi 2814/12,

Axi 2820/12 and Axi 2826/12. Based on 3 cells lithium polymer batteries, the

amount of power that these motors can generate with different propellers is

shown in table 5.1.



Table 5.1: Comparison of power generated with different propellers

Motor Mass (g) Propeller Power (W) Power/Mass Ratio (W/g)

2814/12 106 8,5 x 6" 215 2.03

106 9,5 x 5" 247 2.33

106 10 x 6" 273 2.58

2820/12 151 10.5 x 7" 202 1.33

151 11 x 7" 221 1.46

151 12 x 6.5" 223 1.48

151 11 x 8" 229 1.52

2826/12 181 12 x 8" 213 1.18

Based on the technical specifications of the motors and propellers, the most

suitable motor and propeller combination would be an Axi 2814/12 motor

coupled with an 8.5 x 6” propeller. However, due to the limited types of

propellers available in Singapore stores, the most similar propeller available

was 9 x 6”. Given that a propeller size of 10 x 6” and 8.5 x 6” generates 273

W and 215 W of power with the Axi 2814/12 respectively, it can be safely

assumed that a 9 x 6” propeller will be able to generate more than 215 W of

power. Thus the required power of 205.8 W per motor is satisfied.

5.3 Experimental Verification of Motor’s Technical Specifications

An experiment was designed to verify the amount of power which the motor

can generate. The motor coupled with the 9 x 6” propeller was secured to a

stand, and an ammeter was attached in the circuit in-between the motor and

the battery. The motor was powered up to full throttle, and held steady for a

few second before the current reading was taken. The current drawn was

measured to be 29.7 A. With a battery voltage of 11.1 V, the power drawn is

therefore given by the following equation:


P = VI - (5.3)

The power drawn is found to be 329.7 W. Based on the motor and propeller

combination’s listed efficiency of 75% (see appendix D), the power output of

the motor with the 9 x 6” propeller is 247.3 W. Thus a dual motor

configuration will be able to produce a maximum power of 494.6 W.

5.4 Determining the Thrust of the Motor

An experiment was designed in order to determine the amount of thrust

provided by the motor and propeller combination (see appendix E). By varying

the throttle percentage, the amount of force generated by the motor was

noted down, and a graph of Thrust versus Throttle Percentage can be plotted,

as shown in fig. 5.2 below.

Figure 5.2: Graph of Thrust (N) versus Throttle (%)

Graph of Thrust (N) vs Throttle (%)

y = 0.216x

0

5

10

15

20

25

0 20 40 60 80 100

Throttle (%)

Thr

ust (

N)

120



6 Fabrication and Integration of the Prototype

After the parameters for the prototype is fully defined, the fabrication process,

which is a joint effort between me and Mr Lee Qihui, started in full swing.

Different materials are chosen to fabricate different components of the

prototype according to their individual requirements. For the fuselage, it was

duplicated from Mr Toh Boon Whye’s prototype using fibreglass. The usage of

fibreglass allows the duplication process to have high accuracy, as well as

having good toughness properties which will allow it to withstand the wear and

tear during hard ground testing.

For the wings and horizontal stabilizers, their skeleton frames are constructed

from light weight balsa wood which are cut to specific shapes, and then joined

together using aliphatic resin glue as seen in fig 6.1 below. The skeleton

frames are then covered using a type light weight heat shrink wrap known as

Ora Cover. The wings and tail is then connected to the fuselage through the

use of mechanical fasteners like bolts and nuts to allow for easy assembly

and disassembly, thus making the transportation of the prototype of test sites

easier. The modular design of the prototype will also allow easier modification

of design as only the affected component needs to be changed.

Figure 6.1: Photographs of the prototype’s skeleton structure



6.1 Integration of the Prototype

After all the components have been fabricated or purchased, they have to be

integrated seamlessly into a working prototype. Care has to be taken to make

sure that the center of gravity of the prototype is in the ideal range of position,

and this involves proper distribution of the various components. An

experiment was conducted by Mr Lee Qihui to locate the center of gravity

after the various components were integrated, and the CG position was found

to be within the ideal range of position. Details about the experiment can be

found in Mr Lee’s thesis. The fig. 6.2 below shows the schematic layout of the

prototype in plan view, and fig. 6.3 shows the photographs of the FS8 which

the prototype was based on, and the fabricated prototype side by side.

Figure 6.2: Schematic layout of the prototype in plan view



Figure 6.3: Photographs of the FS8 and the prototype




7 Results and Analysis of the Flight Test Results

Many hours of flight tests were done to validate the design methodology and

the various results predicted by the CFD simulations. The flight tests were

carried out in a few different locations having different environmental

conditions in order to investigate the aerodynamic characteristics. In order to

validate the various geometric and operating parameters obtained from CFD

simulations, the method of comparing the theoretical CL and CD of the aircraft

with the experimental CL and CD is utilized. The experimental CL and CD have

to be calculated based on the velocity and the thrust of the prototype in flight

respectively using the following equations:

CL = L / (0.5 ρ∞v2S) - (7.1)

CD = D / (0.5 ρ∞v2S) - (7.2)

The velocity of the prototype in flight will be calculated by observing the

distance the prototype covered while flying, divided by the amount of time it is

in the air from the video footages. The lift force in equation 7.1 is assumed to

be the weight of the prototype, and the drag force in equation 7.2 is assumed

to be the thrust of the motors at cruise velocity. To prove the validity of the

propulsion system, the prototype simply has to be able to take off from the

water starting from rest, since it was sized to be able to overcome the water

resistance.



7.1 Indoor Flight Test

One of the test sites was the Multi-Purpose Sports Hall 2 (MPSH 2). It was

chosen for its large spacious area, while being sheltered from environmental

conditions like wind which can significantly affect the flying characteristics of

the prototype. The craft is able to take off smoothly and able to sustain a

straight level flight. The height readings captured by the sonar sensor during

the flight (see appendix F) shows that the flying height is at the targeted

height/chord ratio of 0.1. Based on the video footages, the velocity of the

prototype in air has been tabulated in table 7.1.

Table 7.1: Velocity of the Prototype in Air (land takeoff)

Trial Distance (m) Time Taken (s) Average Speed (m/s)

1 30 2.9 10.3

2 36 3.3 10.9

Average 10.6

With the mass of the prototype at 2.027 kg (see appendix G for detailed

breakdown), thus the experimental CL value can be calculated to be:

CL, Indoor = L / (0.5 ρ∞v2S)

= 19.88 / (0.5 x 1.23 x 10.62 x 0.51)

= 0.56

From fig. 7.1 below, it can be seen that the angle of attack the prototype was

flying at was higher than the targeted angle of 4 degree. It can be observed

from the photos that when it was flying at a height/chord ratio of 0.1, the angle

of attack is about 7.5 degree (see appendix H). From fig 7.2, the CL, theoretical

when the angle of attack is 7.5 degree is 0.62.


Figure 7.1: Screenshots of the prototype flying in the MPSH 2

Figure 7.2: Graph of CL versus Angle of Attack

Graph 3: Coefficient of Lift (CL) vs. Angle of Attack(For Anhedral Angle = 8, h/c = 0.1, Wing angle of incidence = 4)

y = 0.0559x + 0.0846

y = 0.0696x + 0.1007

-0.4

-0.2

0

0.2

0.4

0.6

0.8

1

-7 -6 -5 -4 -3 -2 -1 0 1 2 3 4 5 6 7 8 9 10

Angle of Attack

Coef

ficie

nt o

f Lift

(CL)

Wing Alone: AA =8, h/c = 0.1Wing + Hull AA = 8, h/c =0.1Wing Alone (AA = 8, h/c=0.1)Wing + Hull (AA = 8, h/c = 0.1)

Comparing the CL, Indoor value obtained from the flight test with the CL, theoretical

predicted by the CFD simulations, it is observed that the CL, Indoor is lower than

CL, theoretical by only 9.0 %. This shows that the results from the indoor flight test

tallies quite well with the results predicted by the CFD simulations. The slight

difference in CL can probably be attributed to the imperfection introduced in

the fabrication process.

During the test flight, it was noted by the pilot that a throttle percentage of

about 10 % was needed to keep the plane in level flight. From fig. 5.2, the

amount of thrust is about 2.16 N, thus the experimental CD value can be

calculated to be:



CD, Indoor = D / (0.5 ρ∞v2S)

= 2.16 / (0.5 x 1.23 x 10.62 x 0.51)

= 0.06

From fig. 7.3 below, the value of CD, theoretical when the angle of attack is 7.5

degree is 0.05. Comparing the values of the experimental CD, Indoor with CD,

theoretical, it was observed that the CD, theoretical was lower than the CD, Indoor by

18.4 %. The theoretical CD was already expected to under predict the actual

CD, as many of the drag contributing factors was not modeled into the CFD

simulations. Factors like surface roughness and the parasite drag of the

motors and propellers were ignored during the simulations in order to speed

up the computational time.

Figure 7.3: Graph of CD versus Angle of Attack

Graph 2: Coefficient of Drag (CD) vs. Angle of Attack(For Anhedral Angle = 8, h/c = 0.1 Wing angle of incidence = 4)

y = 0.0004x2 + 0.0004x + 0.0096

y = 0.0006x2 + 0.0004x + 0.0151

0

0.005

0.01

0.015

0.02

0.025

0.03

0.035

0.04

0.045

0.05

0.055

0.06

0.065

-7 -5 -3 -1 1 3 5 7 9

Angle of Attack

Coe

ffic

ient

of D

rag

(CD)








7.2 Outdoor Flight Test

Outdoor flight tests have been carried out in 2 locations, mainly at the football

field in NUS and at a 35 m long swimming pool. Unfortunately the flight tests

at the football field were not successful, due to 2 main reasons. Firstly, there

were frequent gusts of wind that causes the prototype to lift off suddenly and

resulting in crashes. Secondly, the field was quite uneven, leading to frequent

unintended veering of direction during the takeoff. Thus only the results from

the water flight test will be presented here.

The experimental CL, outdoor from the water test flight would be compared with

CL, theoretical and CL, indoor to see the effect of the undulating water surface on the

CL. Due to the limited test area, the elevator was deflected fully in order to fly

the prototype at a higher angle of attack, and thus resulting in a shorter take

off distance due to the higher CL. Based on the video footages, the velocity of

the prototype in air has been tabulated in table 7.2.

Table 7.2: Velocity of the Prototype in Air (water takeoff)

Trial Distance (m) Time Taken (s) Average Speed (m/s)

1 17 1.9 8.5

2 10.5 1.2 8.8

Average 8.7

With the mass of the prototype at 2.111 kg due to the additional pontoons,

thus the experimental CL value can be calculated to be:

CL, outdoor = L / (0.5 ρ∞v2S)

= 20.71 / (0.5 x 1.23 x 8.72 x 0.51)

= 0.87


From the sonar sensor attached to the prototype, it was noted that it was only

flying at the height/chord ratio of 0.037. The angle of attack which the

prototype was flying at is about 9 degree (see appendix H). Fig. 7.4 shows the

screenshots of the prototype flying over water.

Figure 7.4: Screenshots of the prototype flying over water

Due to the difference in height/chord ratio, Fig. 4.7 will be use to extrapolate

C L, theorectical at height/chord ratio of 0.037 at 0 degrees angle of attack and it is

found to be 0.164. The gradient of the C L, theorectical against angle of attack

curve at a height/chord ratio of 0.037 is given by equation 4.3, and it is found

to be 0.0896. Thus CL, theorectical at height/chord ratio of 0.037 and at an angle

of attack of 9 degree can be evaluated by equation 4.2:

CL, theorectical = 0.0896α + 0.164

= 0.0896(9) + 0.164

= 0.97

From the results of the indoor flight test, the CL, Indoor value on a hard ground is

9.0 % lesser than the CFD simulation’s prediction, thus the CL, Indoor for a

height/chord ratio of 0.037 and an angle of attack of 9 degree is approximately

0.88. It can be observed that CL, outdoor is lower than CL, indoor by only a 1.1 %

difference. When compared to the CL, theorectical by CFD simulation, CL, outdoor is

only 10.1 % lower than predicted, and the error can be attributed to the same




reasons mentioned above. Although it was noted in the previous year WIG

prototype’s analysis that the free surface effect for a WIG of this scale has a

significant impact and hence cannot be ignored, but it was observed

differently in this case. It might be due to the fact that although it is an outdoor

test, the pool that was utilized is located in a sheltered area, thus there is

minimum waves that might cause inaccuracy in the experiment.



8 Conclusion and Recommendations

The study of a small scale WIG aircraft with an inverted delta wing

configuration has been conducted over a period of 9 months. Although the

prototype is based on an actual WIG aircraft, only a few basic parameters like

the dimension of the wings are known. Many of the important parameters like

the anhedral angle and angle of incidence which determines the aerodynamic

characteristics of the aircraft still have to be obtained through the careful

application of CFD simulations.

As the aerodynamic forces are dependant on a large number of variables, it is

inefficient and too time consuming to investigate all of the variables through

simulation. Thus dimensional analysis was performed to reduce the number of

variables to only three: height/chord ratio, anhedral angle and angle of attack.

CFD simulations are then carried out to obtain various aerodynamic data that

can be used to derive the empirical relationships between the aerodynamic

forces, the anhedral angle, the angle of attack and the height/chord ratio. With

these relationships, the aerodynamic characteristics of the aircraft are thus

known, and the prototype can be fabricated.

In order to size the propulsion system, the maximum amount of drag

experienced by the prototype must be known. Although it is intuitive that it

occurs during a takeoff from the water surface, however due to the complex

interaction between the buoyancy force and the lift force with respect to time,

the maximum drag cannot be easily determined. By simulating the submerged

portion of the fuselage (when in rest position) at the cruising speed, the drag



obtained is definitely higher than the maximum drag experienced by the

prototype. This is due to the wetted surface area decreasing as the lift force

generated increases when the velocity increases. Thus the propulsion system

sized to this simulated value will definitely be able to power the plane

adequately.

Once all the components were fabricated and integrated together, test flights

of the prototype was conducted in a few test sites. From the series flight tests

conducted, it was observed from the indoor tests that the lift and drag

coefficient predicted by the CFD simulations was quite accurate. The results

show that the experimental CL was only lower than the predicted value by 9.0

%, and the CD was higher than the predicted value by 18.4 %. The

discrepancies are possibly due to the imperfect fabrication of the prototype

and the simplification of the simulation models for faster computation time. For

the outdoor test flights on the water, it was observed that the CFD simulations

predicted the results quite accurately again. The experimental CL obtained

from the outdoor test flights was very near the indoor test’s CL, with only a

very small difference of 1.1 %. This shows that the flight characteristics of the

prototype is almost the same for both land and water testing, and thus it can

be concluded that the undulating water surface has insignificant

consequences as assumed in the simulations.

Overall, the objectives for this project have been achieved. A small scale

inverted delta wing configuration WIG craft with amphibious capability has



been successfully developed. From the flight tests conducted at the various

test sites, the prototype is shown to be able to maintain a straight, level flight.

8.1 Recommendations

Due to the limited time and manpower, several aspects of the project were

simplified in an attempt to save time. Those aspects can be the focus for

further studies and research on this topic of inverted delta wing WIG aircraft,

and their significance on the performance of the aircraft can be investigated.

Firstly, more research can be done on the interaction between the buoyancy

force and the lift force with respect to time. By quantifying this relationship, it

will be possible to find out the amount of wetted surface area of the fuselage

at any instance of time during takeoff as long as the velocity is known. Thus

the propulsion system of the aircraft can be more appropriately sized. By

knowing the amount of drag at any given instant of time will also allow an

automatic height control system to control the throttle settings for the whole

flight.

Secondly, there was limited investigation on the turning manoeuvres of the

prototype. All of the experiments were done to investigate straight level flight.

By researching more on the turning and banking capabilities of the prototype

will allow more insight to be gained on the significance of ground effect on

such manoeuvres.



Thirdly, the test flights for this project were done mostly in sheltered

environments. Further research can be done to investigate the effects of

environment turbulences like cross wind and choppy water on the inverted

delta wing WIG aircraft.



References 1. K.V. Rozhdestvensky, Aerodynamics of a Lifting System in Extreme

Ground Effect, 1st ed., Springer-Verlag, 2000, pp 63-67.

2. J.D. Anderson Jr., Fundamentals of Aerodynamics, 3rd ed., McGraw-Hill,

2001.

3. K.V. Rozhdestvensky, Aerodynamics of a Lifting System in Extreme

Ground Effect, 1st ed., Springer-Verlag, 2000, pp 263 – 280.

4. Chin-Min Hsiun, Cha’o-Kuang Chen, Aerodynamic characteristics of a two

dimensional airfoil with ground effect, J. Aircraft v33 (2), 1996, pp 386-392

5. Knud Benedict , Nikolai Kornev , Michael Meyer, Jost Ebert, Complex

mathematical model of the WIG motion including the take-off mode, Ocean

Engineering 29 (2002), pp 315–357.

6. J.D. Anderson Jr., Computational Fluid Dynamics: The Basics with

Application, 1st ed., McGraw-Hill, 1995.

7. Bruce R. Munson, Donald F. Young, Theodore H. Okiishi, Fundamentals of

Fluid Mechanics, 4th Edition, John Wiley & Sons, 2002.

8. M.R. Ahmed. S.D. Sharma, An investigation on the aerodynamics of a

symmetrical airfoil in ground effect, Experimental Thermal and Fluid Science,

In Press, 2004.

9. J.D. Anderson Jr., Aircraft Performances and Design, 1st Edition, Mcgraw

Hill, 1999.

10. H.H. Chun, C.H Chang, Longitudinal stability and dynamic motions of a

small passenger WIG craft, Ocean Engineering 29, 2002, pp 1145-1162.



11. V. Bebyakin Ed., EKRANOPLANS: Peculiarity of the theory and design,

Saint Peterburg, "Sudostroeniye", 2000.

12. Robert C. Nelson, Flight Stability and Automatic Control, 2nd ed.,

McGraw-Hill, 1998.

13. Bill Husa, WIG Configuration development from component matrix,

Aerospace Design and Engineering, Orion Technologies, 2000.

14. Ron Laurenzo, A long wait for big WIGs, Aerospace America AIAA, June

2003, pp 36-40.

15. D.E. Calkins, Feasibility Study of a Hybrid Airship Operating in Ground

Effect, J. Aircraft Vol.14, No.8, August 1977, pp 809 – 815.

16. Ng Geok Hean, AM90 Wing In Ground (WIG) Aircraft – Aerodynamics, 2004/2005. 17. Toh Boon Whye, Propulsion System for a Wing-in-ground effect model, 2004/2005.



APPENDICES



Appendix A: Historical Development in WIG The phenomenon of ground effect was observed as early as the Wright

Brothers’ Wright Flyer I which flew in the presence of ground effect. During

World War II, war planes which were low on fuel flew in ground effect in to fly

back to base in order to make use of the increase in efficiency when operating

in ground effect.

Despite the early discovery of the phenomenon of ground effect before the

cold war, the main advances in ground effect technology took place during the

1960s in the Soviet Union by a Russian engineer, Rostislav E. Alexeyver, and

his Hydrofoil Design and Construction Bureau. Alexeyver and his company

designed and built a number of very successful WIG vehicles known to the

Soviet Union as Ekranoplans. One of Alexeyver’s projects includes the most

famous and the largest of all the ekranoplans, KM, also known to the west as

the Caspian Sea Monster (See Fig. A.1a). Its dimension was documented to

have reached a wing span of 40m, a length of 100m, with a maximum take off

weight to reach 540 tons and had a cruising speed of over 400km/h. The end

of the cold war saw the end of the development of WIG vehicle in the Soviet

Union.

Several European countries were involved in developing ground effect

vehicles. In particular, Dr. Alexander Lippisch, the famous German aircraft

designer and widely known for his invention of delta wing aircrafts, made

significant contribution in the development of WIG vehicles. WIG vehicles,

based on the reverse delta wing which was pioneered by Lippisch, still exist

today and is said to be a much better design to the Soviet Union’s Ekranoplan



(See Fig. A.1b). The world’s first commercialized WIG vehicle is base on the

Lippisch concept.

The most recent development in WIG is perhaps Boeing’s own WIG project

named Pelican [14]. With a wing span of 152m and a fuselage of length 109m,

the Pelican will be the largest aircraft ever build in the world and also the first

non-Russian large WIG.

Being built as a military transport vehicle, the Pelican is designed to carry a

payload of more than 1400 tonnes. Cruising at 6m above water at 480km/h

and powered by four turboprop engines, the Pelican if necessary can also fly

at 20 000feet in the air..

Other interesting WIG concepts proposed includes the Hybrid ground effect

airship by Calkins [15] for the purpose of transoceanic cargo transportation and

the Aerotrain by the Tohoku University Institute of Fluid Science in Sendai.


Appendix B: Fundamental Fluid Mechanics The physical aspects of any fluid flow are governed by the 3 fundamental

principles of mechanics:

1) Conservation of Mass

2) Conservation of Momentum

3) Conservation of Energy

When expressed in terms mathematical equations, the governing equations

for fluid (the Navier-Stoke’s equations) takes the form of the respective partial

differential equations. When the condition of incompressible flow is applied,

the following sets of incompressible Navier-Stoke’s equation are obtained:

Equation 3.1 is known as the continuity equation, equation 3.2 is the

momentum equation and equation 3.3 is the energy equation. If only the

continuity and momentum equations are solved, the flow variables and

coordinates can be non dimensionalized by

Substituting equation B.4 into B.1 and B.2 yields the following non-

dimensional form of the incompressible N-S equations:



Reynolds number is qualitatively defined as the ratio of inertia force over

viscous force and can be easily proven by the following.

Considering that the inertia force will follow the magnitude of the order ρU2

and the viscous force is result from the shear stress,

Hence by taking the ratio between the two:



Appendix C: Pressure Correction Method In the process of discretizating the N-S equations, it is common to define the

pressure and velocity components on the same mesh points. The drawback of

this is that a highly non uniform pressure field will appear to be uniform when

if the usual central difference case is applied. Consider a simplified one

dimensional convection equation:

After applying the central difference scheme on the pressure field and the

explicit Euler on the time derivative yields:

Since , which is not true as the pressure variation is not

reflected in this case.

Now, let’s consider applying the second order upwind scheme on the

pressure field which yields:

Thus the pressure variation is now reflected.



Alternatively, the staggered mesh is use which the pressure and velocity are

not define on the same node as shown below.

Applying the central difference scheme on the pressure field:

The use of the staggered mesh however is only limited to structured mesh,

hence the second order upwind scheme is preferred in this project.



Appendix D: AXI 2814/12 Motor Specifications



Appendix E: Motor Thrust Experiment For the motor thrust experiment, the lever concept was utilised. The

experimental setup is shown in the photograph below. The motor is mounted

on one end of the lever, and a modified camera tripod was used as the

fulcrum. On the other end of the lever, it was supported and weighted down

on a weighing scale.

When the motor is powered up, the reading on the scale will adjust according

to the amount of moment generated by the thrust of the motor. Thus by noting

down the changes in the mass reading of the scale, the thrust of the motor

can be determined.



Appendix F: Graphs of Height Readings (cm) versus Time (s)

Height Readings From On-Board Sensor(Hard Ground Testing)

0

2

4

6

8

10

12

0 2 4 6 8 10 12 14

Time into flight(s)

Hei

ght r

eadi

ngs

(cm

)

Height readings

Height Readings From On-Board Sensor(Water Testing)

0

2

4

6

0 2 4 6 8 10 12

Time into flight(s)

Hei

ght r

eadi

ngs

(cm

)

Height readings




Appendix G: Detailed Mass Breakdown of the Components

Components Mass (kg)

Fuselage + Top Cover 0.496

Horizontal Tail + Servo 0.271

Battery + ESC (x 02) 0.206 x 2

Motor Mount 0.085

Motor + Propeller (x 02) 0.144 x 2

Wings 0.406

Total: 2.027


Appendix H: Determination of Experimental Angle of Attack

Angle B is 15 degrees when it is measured directly from the photo above;

however angle B with respect to the plane is a known angle of 4 degree.

Angle A is 13 degree as measured directly from the photo and by using angle

B’s conversion ratio, angle A is 3.47 degree. Thus it can be assumed that the

prototype is flying at an angle of attack of 7.47 degree.

The angle of attack is measured to be 9 degrees directly from the photo

above.



Appendix I: Tabulations and Graphs of the CFD Simulations

Graph of CL versus Anhedral Angle

Graph 1: Coefficient of Lift (CL) vs. Anhedral Angle - In Extreme Ground Effect(For various h/c, Wing Angle of Incidence = 0 , Angle of Attack = 0)

0

0.01

0.02

0.03

0.04

0.05

0.06

0.07

5 6 7 8 9 10 11

Anhedral Angle

Coe

ffic

ient

of L

ift (C

L) h/c = 0.01

h/c = 0.015

h/c = 0.02

h/c = 0.025

h/c = 0.01

h/c = 0.015

h/c = 0.02

h/c = 0.025

Graph of CD versus Anhedral Angle

Graph 3: Coefficient of Drag (CD) vs. Anhedral Angle - In Extreme Ground Effect(For various h/c, Wing Angle of Incidence = 0 , Angle of Attack = 0)

0.0158

0.0159

0.016

0.0161

0.0162

0.0163

0.0164

0.0165

0.0166

0.0167

0.0168

0.0169

5 6 7 8 9 10 11

Anhedral Angle

Coe

ffic

ient

of D

rag

(CD) h/c = 0.01

h/c = 0.015

h/c = 0.02

h/c = 0.025

h/c = 0.01

h/c = 0.015

h/c = 0.02

h/c = 0.025


Note: No visible trend can be observed for the graph of CD versus Anhedral Angle, but the values of CD for all values of anhedral angle were fairly constant as they are all within the range of 0.0160 to 0.0168.

- 63 -


Graph of Lift/Drag Ratio versus Anhedral Angle

Graph 4: Lift/Drag Ratio vs. Anhedral Angle - In Extreme Ground Effect(For various h/c, Wing Angle of Incidence = 0 , Angle of Attack = 0)

0

0.5

1

1.5

2

2.5

3

3.5

4

4.5

5 6 7 8 9 10 11

Anhedral Angle

Lift/

Dra

g R

atio

h/c = 0.01h/c = 0.015h/c = 0.02h/c = 0.025h/c = 0.01h/c = 0.015h/c = 0.02h/c=0.025

Graph of CL versus Height/Chord Ratio

Graph 7: Coefficient of Lift (CL) vs.Dimensionless Height (h/c)(For Anhedral Angle = 8, Wing angle of incidence = 4)

y = -0.6362x + 0.3636

-0.4

-0.2

0

0.2

0.4

0.6

0.8

0 0.05 0.1 0.15 0.2 0.25 0.3

Dimensionless Height (h/c)

Coe

ffic

ient

of L

ift (C

L)

Wing Alone: AA =8, AoA=3

Wing Alone (AA = 8, AoA=3)


- 64 -


Graph of CD versus Height/Chord Ratio

Graph 6: Coefficient of Drag (CD) vs. Dimensionless Height (h/c)(For Anhedral Angle = 8, Wing angle of incidence = 4)

y = 0.002x + 0.0213

0

0.005

0.01

0.015

0.02

0.025

0.03

0 0.05 0.1 0.15 0.2 0.25 0.3

Dimensionless Height (h/c)

Coe

ffic

ient

of D

rag

(CD)

Wing alone: AA = 8, AoA=3

Wing alone: AA = 8, AoA=3

Graph of Lift/Drag ratio versus Height/Chord Ratio

Graph 5: Lift/Drag Ratio vs. Dimensionless Height (h/c)(For Anhedral Angle = 8, Wing angle of incidence = 4)

8

9

10

11

12

13

14

15

16

17

0 0.05 0.1 0.15 0.2 0.25 0.3Dimensionless Height (h/c)

Lift/

Dra

g R

atio

Wing alone: AA = 8, AoA=3Wing alone: AA = 8, AoA=3



Graph of CL versus Angle of Attack

Graph 3: Coefficient of Lift (CL) vs. Angle of Attack(For Anhedral Angle = 8, h/c = 0.1, Wing angle of incidence = 4)

y = 0.0559x + 0.0846

y = 0.0696x + 0.1007

-0.4

-0.2

0

0.2

0.4

0.6

0.8

1

-7 -6 -5 -4 -3 -2 -1 0 1 2 3 4 5 6 7 8 9 10

Angle of Attack

Coe

ffic

ient

of L

ift (C

L)

Wing Alone: AA =8, h/c = 0.1

Wing + Hull AA = 8, h/c =0.1

Wing Alone (AA = 8, h/c=0.1)

Wing + Hull (AA = 8, h/c = 0.1)

Graph of CD versus Angle of Attack

Graph 2: Coefficient of Drag (CD) vs. Angle of Attack(For Anhedral Angle = 8, h/c = 0.1 Wing angle of incidence = 4)

y = 0.0004x2 + 0.0004x + 0.0096

y = 0.0006x2 + 0.0004x + 0.0151

0

0.005

0.01

0.015

0.02

0.025

0.03

0.035

0.04

0.045

0.05

0.055

0.06

0.065

-7 -5 -3 -1 1 3 5 7 9

Angle of Attack

Coe

ffic

ient

of D

rag

(CD)





Graph of Lift/Drag Ratio versus Angle of Attack



Graph 1: Lift/Drag Ratio vs. Angle of Attack(For Anhedral Angle = 8, h/c = 0.1 Wing angle of incidence = 4)

y = -0.0365x2 + 0.366x + 13.71

y = -0.0214x2 + 0.3458x + 11.062

8

9

10

11

12

13

14

15

16

17

-6 -5 -4 -3 -2 -1 0 1 2 3 4 5 6 7 8 9

Angle of Attack

Lift

/Dra

g R

atio





Graph of CM versus Angle of Attack

Coefficient of Moment (CM) vs. Angle of Attack(For Anhedral Angle = 8, h/c = 0.1 Wing angle of incidence = 4)

y = 0.0082x - 0.0178

y = 0.0118x - 0.0229

y = -0.0273x + 0.0846

y = -0.0156x + 0.063

-0.6

-0.5

-0.4

-0.3

-0.2

-0.1

0

0.1

0.2

0.3

0.4

0.5

-8 -6 -4 -2 0 2 4 6 8 10

Angle of Attack

Coe

ffici

ent o

f Mom

ent (

C M)



Tail alone: h/c =0.1

Whole Aircraft



Tail alone: h/c=0.1

Whole Aircraft


inverted delta wig effect aircraft 2

Documents

working wig aircraft

ground wig effect

cfd simulations

actual flight tests

flight manoeuvrability

water flight tests

propulsion system

software simulations