WIG Aerodynamics (1)

AM90 Wing In Ground (WIG) Aircraft –

Aerodynamics

Submitted by

Ng Geok Hean

Department of Mechanical Engineering

In partial fulfilment of the

requirements for the Degree of

Bachelor of Engineering

National University of Singapore

Session 2004/2005

i

SUMMARY

This project, Wing in Ground (WIG) Aircraft – Aerodynamics, was initiated by

Wigetworks Pte Ltd, a local spin-off company aiming to revolutionize the marine

transport industry by marketing and being the lead manufacturer for the world’s first

commercialized WIG vehicle. The objective of this collaboration was to design, fabricate

and test fly a small scale WIG effect craft.

Based on the literature survey conducted, no research/published papers on a small scale

WIG craft were available because many existing WIG crafts were of a large scale, up to

the size of commercial jet aircraft. Even so, the amount of published data on existing

WIG was also limited as not much serious development had been done since the end of

the Cold War when the government of the former Soviet Union stopped its support for

WIG projects. Hence the motivation behind this project was to gain a better insight and

understanding of the aerodynamics of a small scale WIG and to obtain the aerodynamic

characteristics of the craft which could be used for future developments of such a vehicle.

One of the main challenges of this project was that it required multi- and interdisciplinary

skills. Therefore, this project was done as a team consisting of three other members

dealing with their respective areas: Structures, Propulsion and Stability, and Control. The

nature of this project also includes analyses, fabrication and field tests which involve the

integration of knowledge and skills from different specializations.

Another challenge of this project was the lack of technical data available for a small scale

WIG craft. Analyses and data for a small scale WIG had to be carried out through one

hundred over Computational Fluid Dynamics (CFD) runs to get the qualitative

relationships between the aerodynamic forces and moments with different dependent

variables. Good Computer Aided Design (CAD) modeling skills, proper mesh control

and understanding of numerical methods are also needed to model the physics of ground

effect aerodynamics and to ensure the predicted results are as accurate as possible.

Fifty over hours of flight tests were conducted both indoor and outdoor for validation of

the lift and drag predicted by CFD at different condition. On board instruments were

mounted onto the craft during the flight tests to quantify the test results for validation.

From the flight tests results, this craft was proven to have amphibious capabilities as it

was able to operate over both land and water surfaces, and it performed as expected from

the CFD results.

Finally, this project was selected for publication and was presented at the RSAF

Aerospace Technology Seminar 2005 in the Air Force School Auditorium.

ii

ACKNOWLEDGEMENT

The author will like to express his thanks and heartfelt gratitude to the following persons

who had contributed in this project:

Project Supervisor Assoc. Prof Gerard Leng for his guidance and advice during the

course of this project.

Team members, Mr. Benedict Ng Dyi En, Mr. Jonathan Quah Yong Seng and Mr. Toh

Boon Whye, for their time and effort put into this project.

Staffs from the dynamics lab, especially Mr. Ahmad Bin Kasa for frequently driving us to

the test site.

Mr. Favian Kang Hong An for allowing me to use his wind tunnel.

Mr. Anatoliy from Ukraine for sharing with us his technical expertise and experience in

WIG vehicles

Wigetworks Pte Ltd for supporting this project.

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TABLE OF CONTENTS

SUMMARY ........................................................................................................................ i

ACKNOWLEDGEMENT............................................................................................... iii

TABLE OF CONTENTS ................................................................................................ iv

LIST OF FIGURES ........................................................................................................ vii

LIST OF TABLES ............................................................................................................ x

LIST OF SYMBOLS ....................................................................................................... xi

1. INTRODUCTION..................................................................................................... 1

1.1. OBJECTIVES..................................................................................................... 2

1.2. ORGANIZATION OF THESIS ......................................................................... 3

2. THEORY OF GROUND EFFECT AERODYNAMICS ...................................... 4

2.1 CHORD DOMINATED GROUND EFFECT (CDGE) ..................................... 4

2.2 SPAN DOMINATED GROUND EFFECT (SDGE) ......................................... 6

3. PRELIMINARY CFD ANALYSIS ......................................................................... 8

3.1. CFD – SOME BASIC BACKGROUND............................................................ 8

3.2. THE NEED OF CFD .......................................................................................... 9

3.3. PREPROCESSING........................................................................................... 11

3.4. NUMERICAL SCHEMES ............................................................................... 14

3.4.1. SIMPLE .................................................................................................... 14

3.4.2. UPWIND SCHEME ................................................................................. 14

3.5. COMPARISONS OF RESULTS...................................................................... 15

3.6. CFD TRIALS CONDUCTED .......................................................................... 16

4. DESIGN ................................................................................................................... 20

iv

4.1. CONCEPTUAL DESIGN PHASE................................................................... 20

4.1.1 FIRST WEIGHT ESTIMATION ............................................................. 21

4.1.2. WING PLATFORM ................................................................................. 21

4.2. PRELIMINARY DESIGN PHASE.................................................................. 26

4.2.1. FUSELAGE DESIGN .............................................................................. 26

4.2.2. AERODYNAMIC CHARACTERISTICS OF A WIG. ........................... 27

4.3. CONFIGURATION LAYOUT ........................................................................ 34

4.3.1. PROPULSION SYSTEM INTEGRATION............................................. 34

4.3.2. POSITION OF CENTER OF GRAVITY................................................. 36

4.3.3. HORIZONTAL STABILIZER................................................................. 38

4.3.5. RESULTING LAYOUT........................................................................... 40

5. FLIGHT TESTS AND DISCUSSION .................................................................. 42

5.1. ONBOARD INSTRUMENTATION................................................................ 42

5.2. INDOOR FLIGHT TESTS............................................................................... 43

5.3. OUTDOOR FLIGHT TESTS........................................................................... 45

6. CONCLUSIONS ..................................................................................................... 49

7. RECOMMENDATIONS........................................................................................ 51

7.1. MORE STUDIES ON REVERSE DELTA WING.......................................... 51

7.2. FLOW OVER AIR-WATER INTERFACE..................................................... 51

7.3. OPTIMUM BLOWING PARAMETERS ........................................................ 52

REFERENCES................................................................................................................ 53

APPENDIX A – HISTORICAL DEVELOPMENT IN WIG ..................................... 56

APPENDIX B – FUNDAMENTAL FLUID MECHANICS ....................................... 59

v

APPENDIX C – PRESSURE CORRECTION METHOD ......................................... 61

APPENDIX D – TABULATIONS AND GRAPHS OF CFD RESULTS................... 63

APPENDIX E – DETAIL MASS BREAKDOWN OF CRAFT ................................. 67

APPENDIX F – DESIGN OF HORIZONTAL STABILIZER................................... 68

APPENDIX G – CALIBRATION OF AIRSPEED SENSOR AND FLIGHT TESTS

MEASUREMENTS ........................................................................................................ 70

APPENDIX H – HEIGHT MEASUREMENT............................................................. 73

vi

LIST OF FIGURES

1.1 WG-8 in flight (With courtesy of Wigetworks Pte Ltd.)

2.1 Contour plot of static pressure on an airfoil

2.2 Vortex strength of an aircraft in flight

3.1 Effect of Re on the Lift of a Gottingen 436 at 0 deg angle of attack and h/c = 0.05.

3.2. Geometry and Mesh for Overall Flow Domain

3.3. Mesh of WIG with fuselage and wing

3.4. Mesh across the mid section of WIG

3.5. Comparison between computational results and theoretical results at h/c = 0.1, Re

= 107

3.6 CL and CD vs. number of iterations when TOL is 10-5

3.7 CL vs. α for different Re

3.8 CD vs. α for different Re

4.1. Gottingen 436 airfoil

4.2. Effect of taper ratio with grey showing separated region

4.3 Velocity vector plot showing regions of separation (above) and cross section view

of wing with separation (below)

4.4 Graph of CL vs. AR in and out of ground effect as obtain through CFD

4.5 Velocity contour plot of WIG fuselage.

4.6 CL vs. α characteristic curve for wing-fuselage combination.

4.7 CL vs. h characteristic curve for wing-fuselage combination.

4.8 vs. LCα

h

4.9. Cm vs. α characteristic curve for wing-fuselage combination

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4.10 Static pressure plot along the upper surface of a wing. a. Out of ground effect. b.

In ground effect.

4.11 Power Augmentation Ram System

4.12 PAR effects on a Wing. a. Separation prevented with PAR b. Velocity vector on

upper surface of the wing.

4.13 Thrust Characteristic for different propellers

4.14 Moment characteristic curves with different c.g position.

4.15 Pitching Moment Characteristic of WIG.

5.1 On board instrumentation for measuring airspeed

5.2 Screen shots from indoor flight tests

5.3 Sequential screen shots of WIG flipping during the encounter of a gust

5.4 Sequential screen shots of a successful outdoor flight test

A1 Various WIG concepts

D1 Aerodynamic characteristics of a wing-fuselage combination at 10m/s with

ground clearance h/c = 0.15.

D2 Aerodynamic characteristics of a wing with AR = 4 at 15m/s.

D3 Aerodynamic characteristics of a wing with AR = 5 at 12.5m/s.

D4 Aerodynamic characteristics of a wing with different AR

F1 CL of tail vs. angle of attack

G1 Calibration set up in a low speed wind tunnel

G2 Airspeed sensor setup for calibration.

G3 Airspeed sensor calibration curve

G4 Airspeed sensor readings for indoor flight test

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G5 Airspeed sensor readings for outdoor flight test

G6 Measuring Angle of Attack.

H1. a. Division of string segments. b. Under view of the string setup.

H2. a. Captured side view of string during flight. b. Height approximation using basic

trigonometry.

ix

LIST OF TABLES

4.1 First Estimation of mass breakdown of components

4.2 Second Estimation of mass breakdown of components

5.1. Average Speed calculation

E1 Mass breakdown of craft by components

G1 Calibration Results for airspeed sensor

x

LIST OF SYMBOLS

c Chord Length

h Height

h Height to Chord Ratio

CL Coefficient of Lift

CD Coefficient of Drag

iDC Coefficient of Induced Drag

Cm Coefficient of Moment

α Angle of Attack

xcp Center of Pressure

b Wing Span

S Projected Wing Area on ground plane

AR Aspect Ratio

e Span Efficiency

u* Dimensionless Velocity Vector

p* Dimensionless Pressure

T* Dimensionless Temperature

( )∇ Divergence Operator

2∇ Laplacian Operator

Re Reynolds Number

U Reference Velocity

L Reference Length

ν Kinematics Viscosity

xi

TOL Tolerance

xa/c Aerodynamic center

mC α Slope of Cm vs. α curve

LC α Slope of CL vs. α curve

m0C Intercept of Cm vs. α curve

VH Tail Volume Ratio

Cmwf Coefficient of Moment for Wing-Fuselage Combination

Cmt Coefficient of Moment for Tail combination

Cmwft Coefficient of Moment for Wing-Fuselage-Tail Combination

tmCα

Slope of Tail Moment Characteristic Curve

0 tmC Intercept of Tail Moment Characteristic Curve

ε0 Downwash Angle at 0 Angle of Attack

iw Wing Incident Angle

it Tail Incident Angle

n Time Level

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

tS Area of tail

T Thrust

xii

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. Pilots of huge airplane

like the 747 often experience the plane ‘bounces’ off the runway in the presence of

ground effect just before touch down. This phenomenon that resulted 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.

Recently, a local spin-off company, Wigetworks Pte Ltd, is trying to market a

commercialize Lippisch design WIG craft known as WG-8 (Fig 1). As the world first

commercialize WIG craft is going to be launch from Singapore, this project was initiated

by Wigetworks Pte Ltd to perform further studies and research to gain a better insight of

Ground Effect.

Fig 1.1 WG-8 in flight (With courtesy of Wigetworks Pte Ltd.)

1

1.1. OBJECTIVES

The aim of this project is to design and develop a small scale surface skimming craft

capable of traveling over land and water surfaces. Flight test will be carried out to

validate the results obtained through theoretical and computational methods of analyses.

The minimum design requirement for the craft is to maintain a straight and level flight

over land and water surfaces while carrying a minimum payload of sensors and

instrumentations. On the whole, this project is mainly divided into four stages:

Conceptual Design, Preliminary Design, Fabrication and finally Flight test and evaluation.

1. Conceptual Design

• Define Mission Requirement

• Literature Survey

• First estimate of weight and size

• Determine the configuration layout

2. Preliminary Design

• Calculations carry out using CFD

• Obtain better estimate of parameters e.g. Weight, Size, Cruising Speed and

ground clearance

• Determine stability criteria of the craft

• Finalizing the configuration layout

3. Fabrication

• Modular design for ease of transportation and modification

2

4. Flight Test and Evaluation

• Carrying out flight test both indoor and outdoor

• Evaluate results from flight test

• Minor modification and fine tuning for improvement of performances

• Validate flight tests results with computed data

1.2. ORGANIZATION OF THESIS

This thesis consists of a total of 7 chapters and is divided as follow:

Chapter 1 – Introduces the project, states the mission requirement and objectives.

Chapter 2 – Fundamentals of ground effect aerodynamics are covered here

Chapter 3 – Discussion of some basic aspect of CFD

Chapter 4 – Describe the design methodology develop in this project

Chapter 5 – Contains the results from the flight tests conducted and evaluation

Chapter 6 – Conclusion

Chapter 7 – Recommendation for future developments

3

2. THEORY 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, h . 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.

4

a b

Fig.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:

L2

LC 1 V S2 ∞

=ρ

- (2.1)

D2

DC 1 V S2 ∞

=ρ

- (2.2)

M2

MC 1 V Sc2 ∞

=ρ

- (2.3)

where is density of air, S is projected area on ground plane, V is free stream velocity

and c is the chord length.

∞ρ

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:

5

LChα

= - (2.4)

MC3hα

= − - (2.5)

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:

Mp

L

CxC 3

= = −1 - (2.6)

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 (2b

S), resulting in a reduction in induced

drag.

6

Fig.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

i

2L

DCCeAR

=π

-(2.7)

where e is known as the span efficiency and AR is the aspect ratio. In the presence of

ground effect, Rozhdestvensky [3] shows that 1eh

∝ hence from equation 2.7,

iDC ∝ h - (2.8)

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

height.

7

3. PRELIMINARY CFD ANALYSIS

In the study of aerodynamics, whether it is theoretical, experimental or computational, all

efforts are normally aimed at one objective: To determine the aerodynamic forces and

moments acting on a body moving through air. The main purpose of employing CFD

here is to predict and obtain these aerodynamic forces, Lift and Drag, and Moments,

acting on the craft so that the data can be use for design and analyses for later stage of the

project.

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 – SOME BASIC BACKGROUND

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):

8

* 0u∇ = - (3.1)

2*

1( )t R

** * *u u u p∂

+ ∇ = −∇ + ∇∂ e

*u - (3.2)

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. THE NEED OF CFD

Existing analytical solution for airfoils and wings that are developed were based on the

assumption of inviscid flow [2]. Those methods are fairly accurate if the operating

Reynolds’s number (Re) base on the free stream velocity and the chord length is very

high (in the order of 107 and above). From the Thin Airfoil Theory, the coefficient of lift

is proportional to the angle of attack and independent of the free stream velocity. This is

however not true for lower Re flow lesser than 4x106. By observing the relationship

between CL and Re from a series of CFD runs, the coefficient of lift is found to be highly

dependent on the Re for flow within this region as shown in Fig. 3.1. Similar

observations are also made by Hsiun and Chen [4].

9

0.710.720.730.740.750.760.770.780.790.8

0 2000000 4000000 6000000 8000000 10000000 12000000

Re

CL

Our operating region

Fig. 3.1. Effect of Reynolds number on the Lift of a Gottingen 436 at 0 deg angle of

attack and h/c = 0.05.

These dependency of CL on Re is due to the viscous effect of the fluid to become more

significant as Re decreases. This can be simply explained by looking at the physical

meaning of Re (See Appendix for derivation of Re):

UL Inertia ForceReViscous Force

= =

ν - (3.3)

From equation 3.3, as Re decreases, the viscous force will become more dominant.

Furthermore, from the momentum equation 3.2, the second term of the right hand side

represents the viscous term of the momentum equation. Note that the coefficient of the

viscous term is the inverse of Re. For very large Re, this viscous term can therefore be

neglected but not for values of small Re.

10

As the dimensions and operating speed of a small scale WIG is expected to be around the

order of 1m and 10m/s respectively, then the Re with air as the working fluid in room

temperature is:

5

5

ULRe

10x11.46x106.85x10

−

=ν

=

≈

Therefore the operating condition of the craft falls in the region where the lift is highly

dependent on Re. Hence classical methods of analysis will not be applicable here as

theoretical solution for flow at Re of this range is not available at this moment. Analysis

will then have to be carried out using CFD where the viscous effect of the flow will be

taken into account during computation.

3.3. PREPROCESSING

Before the solutions to the Navier-Stokes equation can be obtained, preprocessing work

has to be done. Preprocessing software, GAMBIT, is being use for Computer Aided

Design (CAD) modeling, mesh generation and implementation of boundary conditions.

Unstructured mesh is being use here due to its ability to adapt to more complex geometry.

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

11

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.

Wall

Velocity Inlet

Outlet

Moving Wall

Fig. 3.2. Geometry and Mesh for Overall Flow Domain

Although the craft is design to operate above the water surface as well, the physics

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

Base on the literature findings, the undulating surface effect is actually negligible [5]. 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. 3.2.

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.

12

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 3.3 and 3.4 shows two examples of the mesh

across the wing-fuselage combination of the craft.

Fig. 3.3. Mesh of WIG with fuselage and wing

Fig. 3.4. Mesh across the mid section of WIG

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3.4. 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:

3.4.1. 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, which is an iterative process, is develop to

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

continuity equation.

3.4.2. 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 zig zag

effect to be not reflected in the momentum equation. One proposed solution to take care

14

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.

Note: Please refer to the appendix for more details on the Upwind scheme and Reference

6 for the SIMPLE algorithm.

3.5. COMPARISONS OF RESULTS

Comparison is made between numerical results and Rozhdestvensk’s prediction of flow

over a flat plate in ground effect. The numerical results matched the theoretical solution

perfectly, hence validating the numerical scheme use.

Fig. 3.5. Comparison between computational results and theoretical results at h/c = 0.1,

Re = 107

15

To ensure proper convergence of the solutions, a study is made on the 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. 3.6, 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.

a b

Fig. 3.6. CL and CD vs. number of iterations when TOL is 10-5

3.6. CFD TRIALS CONDUCTED

With the numerical code validated, the scheme can now be applied to obtain a sets of

relationships needed to carry out design and analysis of a ground effect craft. But the

16

three important parameters, Lift, Drag and Moment are dependent on a number of

variables:

L = f (ρ, V, S, ν, α, h, c) - (3.4a)

D = f (ρ, V, S, ν, α, h, c) - (3.4b)

M = f (ρ, V, S, ν, α, h, c) - (3.4c)

Hence it will be very cumbersome and ineffective to run the CFD computations based on

all the variables above. Dimensional analysis is then needed to cut down the variables to

a few dimensionless parameters reduce the computational effort. 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 ) - (3.5a)

CD = f (Re, α, h ) - (3.5b)

CM = f (Re, α, h ) - (3.5c)

Therefore instead of six variables, only three variables needed for the computation to

obtain the characteristic of the WIG craft. However, from the analysis shown in Fig. 3.7

and Fig. 3.68, the Reynolds number seems to have almost no effect on the value of CL

and only a small effect on CD with various angle of attack. This may seems to be

contradicting at first as in section 3.2, the CL is said to be highly dependent on Re for the

operating region of this craft. But from the results shown, if the range of Re is kept small,

17

for example the same order of magnitude as shown here, the same set of values can be

used to predict the aerodynamic characteristics of another craft as long as the Re is not

too far off from the one being computed. Thus if the craft in this project is assumed to be

operating within 5m/s to 15m/s, the range of Re is given by:

5 51.3 10 Re 4 10× ≤ ≤ ×

Since the Re range is within the same order of magnitude, variables will now be further

cut down to two, height and angle of attack.

CL = f (α, h ) - (3.6a)

CD = f (α, h ) - (3.6b)

CM = f (α, h ) - (3.6c)

Henceforth all subsequent computation made and presented in this paper will be carried

out at Re = 2.7x105. The computational trials will begin by carrying out analyses on a

wing section 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.

18

CL vs Angle of Attack

0

0.2

0.4

0.6

0.8

1

-4 -3 -2 -1 0 1 2 3 4

AOA in deg

CL

Re = 2.7e5 Re = 1.4e5 Re = 4e5

Fig. 3.7. CL vs. α for different Re

Cd Vs AOA

00.010.020.030.040.050.06

-4 -3 -2 -1 0 1 2 3 4

AOA in Deg

Cd

Re = 2.7e5 Re = 1.4e5 Re =4e5

Fig. 3.8. CD vs. α for different Re

19

4. DESIGN

To start off with the design, the requirements for the craft have to be defined first and

they are as follow:

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

instrumentation.

2. Able to skim across both land and water surfaces

3. Operates only in ground effect.

4. Maintain a straight and level flight.

5. Speed limit not more than 20m/s

6. Ease of any repair or modification

7. 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.

With the requirements set, different phases of designs base on the design methodology

being develop in this project will be carry out in sequence and chronological order.

4.1. CONCEPTUAL DESIGN PHASE

In the conceptual design phase, the overall shape, dimensions and weight of the WIG

craft is determine so that a “rough sketch” of how the craft will look like can be

visualized.

20

To reduce the cost and fabrication time, off the shelves components like servos, electrical

engines and propellers are used. Therefore the size of the craft is also limited by the

availability and the constraints of these components.

4.1.1 FIRST WEIGHT ESTIMATION

Being a heavier than air vehicle, the craft cannot get off the ground unless it can produce

a lift greater than its own weight. An initial crude estimation of the takeoff gross weight

of the craft is done so that the desire wing size can be design to produce a lift force

sufficient to lift the craft. Since electric motor is use instead of IC Engine, there will be

no change in weight with respect to time due to fuel consumption. Hence from Table 4.1,

the first iteration will be base on a craft capable of lifting off with a maximum take off

weight of at least 2kg.

Table 4.1: First Estimation of mass breakdown of components

Components Mass / kg % total Mass

Propulsion ( Prop, motor + speed

controller )

0.40 20

Structural (fuselage, wings) 1.400 70

Electronics (servos, receiver, wires ) 0.200 10

Total Mass 2.000 100

4.1.2. WING PLATFORM

21

Here, the geometrical shape of the wing, (a) cross section airfoil, (b) wing sweep, (c)

taper ratio and (d) aspect ratio, will be taken into the consideration base on its design

requirement and operating region.

Firstly, being a ground effect craft, (a) the airfoil chosen for the wing is the Gottingen

436 as it has a flat bottom surface which prevents suction effect as the wing approaches

the ground. Although base on the thin airfoil theory as discuss in section 3.1, a

symmetrical airfoil would seems to provide sufficient lift. This is however not true in the

real case. If a symmetrical airfoil is being used, the convergent and divergent area

between the airfoil and the ground plane will cause a drop in static pressure where the

cross section area is the minimum and hence creates a suction force which sucks the craft

towards the ground. [8]

Fig. 4.1. Gottingen 436 airfoil

Next, (b) the platform shape of the wing will be determined base on its operating region.

To decide whether if any swept angle is required will depends on whether the craft is

operating in subsonic, transonic or supersonic regime. This is done by looking at the

Mach Number. At sea level, the speed of sound, a, is approximately 300m/s. Hence the

Mach Number for our craft:

V 10Maa 330

0.03 0

= =

= ≈ - (4.1)

22

Since Ma ≈ 0, no swept is needed for the wing. The next factor will be (c) the taper

ratio, t

r

cc

, which is the ratio between the chord of the tip to the root of the wing. Wing

with different taper ratio will exhibit different flow phenomenon with flow separation

occurs at the root when taper ratio = 1 to separation at the tip when taper ratio = 0 [9]. The

effects can be summarized in Fig 4.2.

Fig. 4.2. Effect of taper ratio with grey showing separated region (Taken from Ref: 9)

From Fig. 4.2, it may seem beneficial to choose the 3rd case for our wing as separation

occurs only at the tip where the area is small hence the impact on the loss of lift is

minimum. But the 1st case is chosen instead for this project due to the advantage of its

ease of fabrication, maximum area hence more lift and also flow separation prevention

method is also incorporated in this project to prevent separation on the root which will be

discussed in the subsequent section.

23

Fig. 4.3. Velocity vector plot showing regions of separation (left) and cross section view

of wing with separation (right)

Flow visualization is carried out using CFD to confirm the region of separation. The blue

regions represents area with very flow velocity which indicates separation has occurs at

the root of the wing.

Finally, to decide on (d) the value of aspect ratio, CFD analyses are carried out for the

wing with different aspect ratio to obtain the its relationship with the amount of lift

generated. A comparison is made on the same wing in the absence on ground effect.

From Fig. 4.4, given the same aspect ratio, the wing in ground effect will have a 100%

higher coefficient of lift than without ground effect. This shows that a WIG can be made

smaller than an aircraft and yet generate more lift. This enables the WIG to carry more

payload than an aircraft as given its smaller size, the structural weight can be reduce.

Hence WIGs normally will have an aspect ratio much lesser than an aircraft.

24

0

0.2

0.4

0.6

0.8

1

1.2

0 1 2 3 4 5

AR

CL In Ground

Out Ground

Fig. 4.4. Graph of CL vs. AR in and out of ground effect as obtain through CFD

Another observation can be made from Fig. 4.4. The coefficient of lift will slowly reach

an optimum value at aspect ratio around 2.5 to 3 as beyond that, the increment of lift

seems to have plateau. As coefficient of lift is a measurement of lift efficiency, increasing

the aspect ratio further does not lead much gain in CL and by making the wing too large

may lead to structural penalty and the increment of wetted area will lead to an increase in

skin profile drag.

To summarized, the following is selected for the wing geometry:

a. Gottingen 436 airfoil section

b. Straight Wing.

c. No tapering

d. Aspect Ratio of 2.5.

25

4.2. PRELIMINARY DESIGN PHASE

With the geometry of the wing decided during the conceptual phase, the preliminary

phase will then involve extensive Computational Fluid Dynamics (CFD) analysis being

carried out to determine the aerodynamic characteristics of the wing-fuselage

combination of the WIG craft.

4.2.1. FUSELAGE DESIGN

Like most aircrafts and boats, the fuselage/hull is where all the components and payloads

are housed. But being a craft which operates in both water and air, the design of a WIG’s

fuselage involves taking into the account of the aero-hydrodynamic effect of the craft.

Hence this section is done in a joint effort with Mr. Toh Boon Whye who is overseeing

the hydrodynamics and propulsions of the WIG. To minimize the hydrodynamic drag, the

bottom surface of the fuselage which is in contact with water is design according to naval

architecture technology and tow tank tests are carried out by Mr. Toh for measurements

of the hydrodynamic drag. For more details please refer to Mr. Toh’s paper.

On the other hand, while airborne, the WIG will behave almost like an aircraft. To

minimize the aerodynamic drag, the fuselage is made as streamline as possible according

to the physic of low speed aerodynamics. Similar to a low speed aerofoil, the nose of the

fuselage has to be made as round as possible and the trailing edge as thin as possible to

allow air to flow around it smoothly without much abruption. Fig. 4.6 shows the velocity

26

contour plot generated by CFD. No separation of flow can be seen occurring on the

surface of the fuselage and minimum wake is observed at the trailing edge.

Fig. 4.5. Velocity contour plot of WIG fuselage.

4.2.2. AERODYNAMIC CHARACTERISTICS OF A WIG.

The aerodynamic characteristics of a WIG are obtained by running a series of simulation

of a wing with various angle of attack and height. The lift, drag and moment are then

obtain by integration of the pressure force and shear force acting on the wing. Recall in

section 2.1, Rozhdestvensky prediction LC hα∝ and in section 3.6, the aerodynamic

forces are proven to be function of α and h , therefore the purpose here is to derive a

relationship between the aerodynamic forces vs. ground clearance and height. Fig. 4.6

and Fig. 4.7 shows two different CL curves, one dependant on α another dependant on h .

Attempts will now be made to relate these two curves and obtain a quantitative

expression for calculation of lift with different α and h .

27


CL = 0.0997α + 0.584CL = 0.0903α + 0.5359

h=0.1ch=0.15ch = 0.085c

0.2

0.4

0.6

0.8

1

-4 -2 0 2 4 6

AOA in deg

CL

Linear (h=0.1c)Linear (h=0.15c)Linear (h = 0.085c)

CL = 0.1007α + 0.6139

Fig. 4.6. CL vs. α characteristic curve for wing-fuselage combination with various h

CL vs ground clearance

CL = 0.3641h-0.2136

0.5

0.520.54

0.56

0.58

0.60.62

0.64

0.07 0.09 0.11 0.13 0.15 0.17 0.19

h/c

CL

Fig. 4.7. CL vs. h characteristic curve for wing-fuselage combination at α = 0o.

Fig. 4.6 shows plots of a few lift characteristic curve which can generally be expressed by:

0L LC C Cα

= α + L - (4.1)

where is the gradient of the curve, is lift coefficient at 0 angle of attack and α is

expressed in degrees. Observe the value of

LCα 0LC

LCα

for different h and they are plotted in

28

Fig. 4.8 and their relationship can be expressed by fitting a cubic curve onto the data.

Hence for different LCα

h can be found by:

LCα

= -1.1897( h )3 - 1.3265( h )2 + 0.2001( h ) + 0.0941 - (4.2)

The next unknown to be determined will be which can be obtained from Fig. 4.7 for

different

0LC

h and CL can be calculated from equation 4.1. To summarize the procedure, to

calculate CL for different height and α:

1. Determine and for given LCα 0LC h from Fig. 4.8 and Fig. 4.7 respectively.

2. Calculate CL from and as obtain from step 1 for given α. LCα 0LC

Gradient of CL = -1.1897(h/c)3 - 1.3265(h/c)2 + 0.2001(h/c) + 0.0941

0.0880.09

0.0920.0940.0960.098

0.10.102

0.08 0.09 0.1 0.11 0.12 0.13 0.14 0.15 0.16

h/c

Gra

dien

t of C

L

Fig. 4.8. LCα

vs. h

29

CM vs AOA at Leading Edge

Cm = -0.2059α - 1.3804

-2.5-2.3-2.1-1.9-1.7-1.5-1.3-1.1-0.9-0.7-0.5

-3 -2 -1 0 1 2 3 4 5

Fig. 4.9. Cm vs. α characteristic curve for wing-fuselage combination at h 0.1=

Secondly, in an aircraft, there is a position where the moment is constant with varying

angle of attack. This position is known as the aerodynamic center (a/c) and it is found to

be located at the quarter position of the chord. This point can be obtained mathematically

by the ratio of the slope of the Moment characteristic curve and Lift characteristic curve

with respect to angle of attack [10]:

ma / c

L

CxC

α

α

= - (4.3)

But from the characteristic curves that were obtained using CFD, as shown in Fig. 4.6

and Fig. 4.9, the a/c was found to be located at 31% of the chord for h 0.1= . The results

of the thin airfoil theory presented in section 2.1 shows that the a/c of a flat plate in

ground effect is at one-third of the chord. On the other hand, an aircraft has its a/c is

located slightly aft of the chord. Hence from these two results, one can deduce that when

30

a WIG fly out of ground effect, the a/c will start to shift in front and this will lead to some

implication on the longitudinal stability of the craft.

Secondly, recall in section 2.1, the Lift force is said to increase with decreasing ground

clearance and increasing angle of attackand is presented in Fig. 4.7 and Fig. 4.6, one

could conclude that if the WIG can operate at very high angle of attack and at very small

ground clearance, then maximum lift can be achieved which maximize the potential of a

WIG. Is this really true? Unfortunately, in reality, we do not get something out of nothing.

Let’s look at Fig. 4.9 which shows the difference between the static pressure obtain by

CFD on the upper surface between two similar wings, one in ground effect and another

out of ground effect, at 3 degrees angle of attack. Comparing the pressure plot between

the two figures, the one on the left which is in ground effect has a much higher adverse

gradient than the one on the right which is in the absence of ground effect. So, another

observation could also be seen for a WIG: low stall angle. Do note data for angle of

attack beyond 5 degrees angle of attack are not presented here. This is because stalling is

observed to occur at 5 degrees and beyond due to the separation at the wing root as

shown in Fig. 4.3. The cause of it is mainly due to the high adverse pressure gradient on

the upper surface of the wing in the presence of ground effect.

31

a b

Fig. 4.10. Static pressure plot along the upper surface of a wing. a. Out of ground effect.

b. In ground effect.

This problem however, can be overcome. In aerodynamics, different methods have been

proposed to prevent or delay stall by either passive or active method. One such active

method is by blowing of air across the upper surface of the wing to increase the

momentum of the air so that it could overcome the high adverse pressure gradient. This

can be done by using one of the design features of a WIG by placing engines or

propellers placed ahead of the leading edge of the wing known as Power Augmentation

Ram Effect or PAR [11] which is employed to overcome the large hydrodynamic drag

during the initial take off phase. But by allowing part of the slip stream from the propeller

to flow through the upper surface of the wing not only prevents separation, but the higher

velocity on the upper surface will create a larger suction force therefore increasing the

total lift of the wing. Simulations results have shown that the total lift can be increased up

to 20% with PAR. This small increment is due to the size of the prop which is relatively

much smaller than the wing area hence only part of the wing is exposed to the slipstream

from the propeller. Fig. 4.11b shows the area of the wing that the PAR has an effect on.

32

Fig. 4.11. Power Augmentation Ram System

a b

Fig. 4.12 PAR effects on a Wing. a. Separation prevented with PAR b. Velocity vector

on upper surface of the wing.

But, like all other blowing methods propose by aerodynamicists, this method will work

only if the velocity of the slipstream from the propeller is higher than the velocity of the

craft is cruising. Hence at high speed, this method is practically ineffective and therefore

the craft will be design to cruise at angle of attack lesser than 3 degrees to prevent stalling

from occurring.

33

4.3. CONFIGURATION LAYOUT

The design of various parts of the WIG will be look into in this section. As most of these

parts are coupled with other disciplines, the discussion will focus on how they are

integrate with the knowledge of aerodynamics. Reference will have to be made for more

details for the design of the respective parts.

4.3.1. PROPULSION SYSTEM INTEGRATION

With the majority of the airframe components determined, the next item will be the

propulsion system to drive the craft on air.

The propulsion system integration is done with Mr. Toh Boon Whye who is in charge of

the propulsion system in this project. To determine how much thrust is needed to propel

the craft forward will then depend on the drag force acting on the craft. In cruise

condition, the thrust produce must be equal to the drag of the craft at that cruising speed.

Hence the selection of the propulsion system will depend on the cruising speed of the

craft. At this stage, a better weight estimate of the craft will now be available since most

of the design work is done.

34

Table 4.2: Second Estimation of mass breakdown of components

Components Mass / kg % total Mass

Propulsion ( Prop, motor + speed

controller )

0.350 21.5

Structural (fuselage, wings) 1.100 67.5

Electronics (servos, receiver, wires ) 0.180 11

Total Mass 1.630 100

Base on the second estimate of the weight of the craft, the required cruising speed is

obtained from the Lift:

1L2

12

LVSC

161.23 0.506 0.59

9.3m / s

∞

=ρ

=× × ×

≈

Where CL is obtain from Fig. 4.6, S is the projected area of the craft on the ground plane,

and ρ∞ is density of air.

From the thrust analyses conducted by Mr. Toh Boon Whye, the thrust characteristic of

the different propellers is plotted in Fig. 4.11 with the drag force predicted by CFD.

The intersection of between the thrust characteristic and the drag curve represents the

cruising speed of the craft. Hence of the four types of propellers available, only the two 7

inch diameter propeller are found match the thrust requirement for the craft. The four

blades 7 inch diameter is selected finally as it produces slightly more thrust than its two

blades counterpart.

35

0.00

0.50

1.00

1.50

2.00

2.50

3.00

0 5 10 15 20

Speed

Forc

e N

Drag7"-4 blade7"-2 blade5.6"-4 blade5.25"-2 blade

Requirements Match

Fig. 4.13. Thrust Characteristic for different propellers.

4.3.2. POSITION OF CENTER OF GRAVITY

Similar to an aircraft, the c.g position of the WIG plays an important role in achieving

longitudinal stability. The analyses for stability of the craft are done with Mr. Quah Yong

Seng, Jonathan who is in charge of the flight control system design. To achieve

longitudinal stability calls for the following two conditions to be met [12]:

0Cm <α and - (4.4)

0C 0m > - (4.5)

Mathematically, equation 4.4 and 4.5 means the WIG moment characteristic curve must

intercept at the positive y-axis and has a negative gradient.

36

Fig. 4.14 shows the moment characteristic curves of the craft taken at different c.g

position. It shows that a wing alone design is normally not stable especially if the airfoil

used is positively cambered as regardless which position the c.g is placed, it will never

meet the above two requirements. Typically, a convenient position for the c.g. is chosen

to be near or at the aerodynamic center which in this case at 33% of the chord. This is

done in particular to enable the horizontal stabilizer to be easily design to suit conditions

4.4 and 4.5 and will be discussed in further details in the next section.

c

Cm Vs AOA

-0.4

-0.3

-0.2

-0.1

0

0.1

0.2

-3 -2 -1 0 1 2 3 4 5

AOA

Cm

0c 0.333c 0.5c

Fig. 4.14. Moment characteristic curves with different c.g position.

37

4.3.3. HORIZONTAL STABILIZER

The horizontal stabilizer is used to provide longitudinal trim and stability of the craft. For

an aircraft, it can be either mounted behind the main wing which is the conventional way

or in front of the main wing and is known as the canard. Here, the conventional design

will be chosen and therefore the horizontal stabilizer will be mounted at the tail of the

craft.

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

height stability, the horizontal stabilizer is normally mounted high out of ground effect.

More detailed analyses of longitudinal and height stability is carried out by Mr. Quah

Yong Seng, Jonathan. 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 taken as a wing in the absence of ground effect.

-0.06

-0.04

-0.02

0

0.02

0.04

0.06

0.08

0.1

-4 -2 0 2 4 6

AOA

Cm

Wing-fuselageWing-fuselage-tailTail

Fig. 4.15. Pitching Moment Characteristic of WIG.

38

The total moment acting on the WIG is the sum of the moment contribution about the c.g

from the wing-fuselage combination and the tail. When expressed in dimensionless form,

the moment equation taken with respect from the c.g is given as:

Cmwf + Cmt = Cmwft - (4.6)

The blue curve represents the desired condition for stability. The moment characteristic

of the wing-fuselage combination is obtained from Fig. 4.15 with the thrust taken into

account. The tail moment characteristic curves can be obtain from equation 4.6 and is

presented as the red curve.

To design the tail, the slope and the intercept of the tail moment characteristic curves can

be written as

tm HC V Cα

≈ −tLα

t )

- (4.7)

0 t tm H L 0 wC V C ( i iα

≈ ε + − - (4.8)

where VH is the tail volume ratio which is proportional to the tail area, ε0 is downwash

angle at zero angle of attack, iw is the wing angle of incidence and in this case 3 degrees,

and finally it is the tail angle of incidence.

The size of the tail will therefore be determined by the slope of the curve. will

therefore determine the angle of incident of the tail, i

0 tmC

t. From calculation, the tail size

39

40

requires will have a span of 0.4 meters with a chord of 0.2 meters mounted at an incident

angle of 0.65 degrees.

Detailed working of obtaining the tail size and its angle if incident can be found in the

Appendix F.

4.3.5. RESULTING LAYOUT

All aspect of chapter four is to achieve the final layout of the craft beginning with the

weight estimation, determining the wing size, fuselage design, propulsion system

integration and finally the control system. From this point onwards, fabrications are done

base on the final layout drawings shown in Fig. 4.16. The final weight of the craft with all

individual components integrated is given in Appendix E.

41

Plan View

Side View Front View

Fig. 4.16. Resulting Layout of WIG

5. FLIGHT TESTS AND DISCUSSION

The objective of the flight test is to validate the design and the results from the

calculations obtained using numerical or theoretical methods. A pitot-tatic tube is

mounted next to the nose of the craft to measure the airspeed which will therefore use to

validate the value of CL and CD from CFD analyses. Two different flight tests are carried

out, indoor and outdoor.

5.1. ONBOARD INSTRUMENTATION

A telemetry system manufactured by the German R/C manufacturer, Robbe, is used in

this project for airspeed measurement. Reasons why this telemetry set is chosen is

because it is small and compact which can be easily fitted onto the craft and its velocity

range is within this project’s requirement. The purpose of the pitot-static tube is to

measure the difference between the stagnation free stream pressure and static air pressure

to obtain the velocity. Hence it must be mounted ahead of the craft at the nose such that it

is away from the slip stream of the propeller.

42

a b

Fig. 5.1. On board instrumentation for measuring airspeed. a. Pitot static tube mounted on

the nose. b. Airspeed sensor.

Calibration of the speed sensor is done using a low speed wind tunnel located in the Fluid

Mechanics lab in WS2. The procedure and calibration curve of the tests is described in

detail in the Appendix G.

5.2. INDOOR FLIGHT TESTS

Indoor flight tests are carried out in the Multi-Purpose Sports Hall (MPSH). Being a

closed environment, the MPSH is being sheltered from environmental factors like

weather, especially wind. Results from the indoor flight tests are excellent and tallies

quite well with calculations. The craft is able to take off smoothly and able to sustain a

straight level flight with the ability to trimmed itself.

43

a) Front View b) Rear View

Fig. 5.2. Screen shots from indoor flight tests

The average cruising speed of the craft is obtain through several runs by taking the

average time taken for the craft to travel from one end of the MPSH to the other which is

100 meters. The instantaneous velocity of the craft is obtained by the speed sensors

mounted on board the craft (Results presented in Appendix G).

Table 5.1. Average Speed calculation

Speed calculation base on video Trial Distance Time Taken Average Speed

1 100 11 9.09 2 100 10 10.00 3 100 9 11.11 4 100 11 9.09 5 100 12 8.33 Average 9.53

From both methods, the average speed of the craft is found to be between 8m/s to 10m/s.

This tallies with the designed condition as stated in section 5 that for the given weight of

this craft, the designed cruising speed for the craft to stay on air is 10m/s. Using this to

obtain the Lift coefficient:

44

L,indoor 212

212

LCV S

14.91.23 9.53 0.506

0.499

=ρ

=× × ×

=

Comparing the value obtain from the flight test to the CFD prediction presented

in Fig. 4.6, at 0

L,indoorC

L,CFDC o angle of attack is 0.59, there is only a 15 % difference. This

difference is attributed to the imperfection which occurs during fabrication which is not

taken into account in the simulation. Since the measured result from the indoor flight test

is a more accurate prediction of the WIG performance, this result will then be use for

comparison with the results obtain from the outdoor flight tests.

The value of CD,CFD = 0.0272 is expected to be an unpredicted value of CD,indoor as the

surface roughness and the parasite drags of other components on the WIG is ignored

during the computation. According to the flight tests, during cruise D = T and T is taken

from Fig. 4.13:

D,indoor 212

212

TCV S

0.88351.23 9.53 0.506

0.0312

=ρ

=× × ×

=

Comparisons between the two values shows that CD,CFD has under predict the actual drag

by 12.8 %.

5.3. OUTDOOR FLIGHT TESTS

45

Outdoor flight tests are much more challenging than indoor tests. Unpredictable results

are sometime obtained during outdoor flight tests due to environmental turbulences. The

outdoor tests are carried out at a pond located in West Coast Park. Environmental

disturbances encountered during the outdoor tests are strong gust of wind and ripples in

the water created by a fountain at the center of the pond. Tests will have to be carried out

ideally at a time when the fountain is off and a less windy day. A strong gust of wind

with a speed of about 5m/s heading in the opposite direction of the craft will cause a

sudden increment in lift hence a sudden nose up moment which results in the craft

flipping over. (See Fig. 5.3)

a b

c d

Fig. 5.3. Sequential screen shots of WIG flipping during the encounter of a gust

46

The measured velocity of craft is also not as accurate due to cross wind which will cause

a misalignment between the pitot tube and the velocity vector as pitot-static tube is

relatively not sensitive to yaw effects. Nevertheless, under less windy condition, the craft

is still able to skim above the water surface but the ground clearance is much lesser as

compared to flight tests conducted indoor. This is due to the undulating effect of the

water surface which is not taken into consideration during the design phase. Also,

because of the large hydrodynamic drag, the craft has to fly at 60 angle of attack to

compensate for the loss of lift while flying at a lower speed measuring from 7m/s to

8.5m/s.

a b

c b Fig. 5.4. Sequential screen shots of a successful outdoor flight test

47

From the observation of the flight test, the predicted CL can be calculated:

L,outdoor 212

212

LCV S

14.91.23 8.5 0.506

0.85

=ρ

=× × ×

=

Because it is flying at a lower ground clearance, Fig. 4.7 will be use to extrapolate

t L,CFDC a h 0.05= and 0 degrees angle of attack and is found to be 0.69. To calculate CL

at α = 60, another extrapolation is needed to evaluate the gradient of the CL vs. α curve,

, at LCα

h 0.05= by using equation 4.2. Finally, CL at h 0.05= and α = 60 can be

evaluated by equation 4.1:

L,CFDC 0.1007 0.690.1007(6) 0.691.2942

= α +

= +=

From the results of the indoor flight test, the actual CL value on a hard ground is 15%

lesser than CFD prediction, hence L,indoorC 1.1= vs. L,outdoorC 0.85= there is a 23 %

difference. When compared to the prediction by CFD, is only 35% of what was

predicted. Therefore a conclusion can be made here that the free surface effect for a WIG

on this scale has a significant impact and hence cannot be ignore, unlike for a large scale

WIG.

L,outdoorC

48

6. CONCLUSIONS

A study on a small scale WIG craft has been done over the past nine months of the

academic year. Although there is a lack of available technical data for design purposes,

understanding the philosophy behind numerical methods has enable the proper use of

CFD to obtain accurate prediction of the aerodynamic forces acting on the craft.

As the forces are dependant on a large number of variables, it will therefore be ineffective

and computationally costly to conduct CFD runs base on all the variables. Thus

dimensional analyses are needed carried out to reduce the number of variables to only

two: height and angle of attack. Although theoretical methods have been proposed to

predict the aerodynamic forces in ground effect, these methods are only limited to simple

shapes (e.g. flat plate) and intensive mathematical operation need to be applied, making it

very tedious. CFD runs are then carried out to obtain a series of aerodynamic data so that

empirical relationships between the aerodynamic forces, the angle of attack and the

ground clearance can be derived. These relationships can therefore be used for future

development and design of WIG. Although these relationships are based on data that

were computed based on a constant Reynolds Number at 2.7 x 105, results have shown

that if the operating range of Reynolds number is kept small, in this case 1 x 105 < Re < 4

x 105, the variation of the force coefficients are not that significant, therefore the

relationships will be valid.

Despite that WIG is much more effective in generating lift of up to 100 percent more than

an airplane, this project has shown that WIG has its limitation. Although through

theoretical analyses show that by flying the WIG at high angle of attack and at very small

49

ground clearance can achieve very high amount of lift, CFD results have shown that this

is not possible as the high ramming pressure below the wing surface results in a build up

of high adverse on the upper surface of the wing and this promotes flow separation on the

upper surface of the wing hence decreasing the stall angle. Therefore, the WIG is limited

to flying at a small angle of attack in order to prevent stalling from occurring.

From the series flight tests conducted, the results shows that there is a better prediction on

the CL for the indoor tests by CFD than for the outdoor tests. CFD manage to predict 85%

of the total lift for indoor but only 65% for outdoor. This is due to the free surface effect

of the water which was not taken into account during the computation. This shows that

for a small scale WIG, it is rather sensitive to the wave and water surface effect, unlike a

larger WIG which is relatively insensitive.

Overall, the requirements for this project have been met. A small scale WIG craft with

amphibious capability has been successfully developed. From the flight tests conducted

during the course of this project have shown that the craft is able to maintain a straight,

level flight and also has the binding to ground effect which prevents it from lifting off

like an aircraft.

50

7. RECOMMENDATIONS

To facilitate any further developments in this project, the followings are recommended:

7.1. MORE STUDIES ON REVERSE DELTA WING

The Lippisch’s reverse delta wing platform is said to be very insensitive to the changes in

the moment acting on the wing with respect to ground clearance. Therefore the reverse

delta wing is said to have inherent stability since the position of the pitch center will not

vary as much as it would on the rectangular wing platform [13], making the vehicle more

stable. However, no published data on reverse delta wing is available and the

aerodynamics of a reverse delta wing is unknown at the moment. More studies can be

made on the reverse delta wing by conducting wind tunnel studies to obtain quantitative

measurements on the pressure distribution, lift, drag and moment acting on the wing.

Flow visualization can also be carried out using water tunnel.

7.2. FLOW OVER AIR-WATER INTERFACE

When a WIG skims above the water surfaces, free surfaces effect induced by the

aircushion can be observed as ripples and wakes behind the craft. The physics between

the air-water interface is very complex and modeling it using CFD requires large

computational power. Although from literature findings [5], these undulating effects will

51

cause insignificant changes to the lift force, this is however not true from observation

during the flight tests for a small scale WIG. Hence for future research works, the free

surface effect of the water will have to be taken into account. To model the physics of

such flow problem, it is recommended that one will have to look into better numerical

schemes that can not only model the physics accurately but also efficiently in order to

reduce computational costs.

7.3. OPTIMUM BLOWING PARAMETERS

In this project, although the PAR is proven to be an effective way improving the

aerodynamic efficiency and preventing separation, it is however not fully optimized. The

amount of lift augmented by PAR depends on the amount of slipstream flowing across

the upper and lower surface of the wing. CFD studies conducted in this project have

shown that for a fixed distance between the wing and the PAR, the increment of lift

varies with the angle of the PAR as well as the free stream velocity. Therefore from the

observed results, by carefully controlling the amount of air flowing across the upper and

lower surface of the wing, one can optimized the total lift force. However, as too many

parameters are involved here, a trial and error process of getting the optimized blowing

parameters will be impractical. A study on quantifying the relationship between the lift

generated and the PAR parameters is therefore needed to improve the efficiency and

design of a WIG.

52

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

53

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.

54

APPENDICES

55

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

56

Soviet Union’s Ekranoplan (See Fig. A.1b). The world’s first commercialized WIG

vehicle is base on the Lippisch concept (See Fig. 1.1).

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.

57

a) Alexeyver’s KM-1 b) Lippish’s X-114

c) Hybrid Airship d) Aerotrain

e) Boeing’s Pelican

Fig. A.1. Various WIG concepts

58

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:

0u∇ = - (B.1)

21( )tu u u pδ

+ ∇ = − ∇ + ν∇δ ρ

u - (B.2)

2

p

T k( ) Tt c

uδ+ ∇ = ∇

δ ρT - (B.3)

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

* xxL

= , * yyL

= , * zz , L

= * ttL / V∞

= } (B.4)−

* uuU∞

= , * vvU∞

= , * wwU∞

= , *2U∞ ∞

ρρ =

ρ

59

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

the incompressible N-S equations:

* 0u∇ = - (3.1)

2*

1( )t R

** * *u u u p∂

+ ∇ = −∇ + ∇∂ e

*u - (3.2)

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 and the

viscous force is result from the shear stress,

2Uρ

uy L

U∂τ = μ ≈ μ

∂. Hence by taking the ratio

between the two:

2Inertia U ULViscous U / L

ρ ρ= =

μ μ

60

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:

u pct x

∂ ∂+ =

∂ ∂0 - (C.1)

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

on the time derivative yields:

n 1 n n ni i i 1 i

tu u (p p2 x

++ −

Δ= − −

Δ 1)

− i

- (C.2)

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

in this case.

n ni 1 i 1p p+ = n 1 n

iu u+ =

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

yields:

n 1 n n n ni i i i 1 i

tu u ( 3p 4p p2 x

+− −

Δ= − − + −

Δ 2 ) - (C.3)

ni

tu ( 42 xΔ

= − −Δ

nip ) - (C.4)

61

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:

n 1 n n ni 1/ 2 i 1/ 2 i 1 i

tu u (p px

++ + +

Δ= − −

Δ) - (C.5)

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

second order upwind scheme is preferred in this project.

62

APPENDIX D – TABULATIONS AND GRAPHS OF CFD RESULTS

0.20.30.40.50.60.70.80.9

-4 -2 0 2 4

AOA in deg

CL

00.010.020.030.040.050.06

-4 -2 0 2 4

AOA in Deg

Cd

a b

-0.12

-0.11

-0.1

-0.09

-0.08

-0.07

-0.06-4 -3 -2 -1 0 1 2 3 4

AOA in degrees

Cm

12

13

14

15

16

17

18

19

20

-4 -3 -2 -1 0 1 2 3 4

AOA in Deg

CL/

Cd

c d

Fig. D1. Aerodynamic characteristics of a wing-fuselage combination at 10m/s with

ground clearance h/c = 0.15.

63


0.460.480.5

0.520.540.56

0 0.05 0.1 0.15 0.2 0.25

h/c

CL


00.20.40.60.8

1

-2 -1 0 1 2 3 4

AOA in deg

CL

a b

Cd Vs Ground Clearance

0.010.012

0.0140.016

0.018

0 0.1 0.2 0.3

Ground Clearance

Cd

Cd Vs AOA

0

0.01

0.02

0.03

0.04

-2 -1 0 1 2 3 4

AOA in Deg

Cd

c d

2527293133353739414345

0 0.05 0.1 0.15 0.2 0.25

Ground clearance

CL/

Cd

Cd Vs AOA

0

0.01

0.02

0.03

0.04

-2 0 2 4 6

AOA in Deg

Cd Cd

e f

Cm Vs Ground Clearance

-0.12

-0.115

-0.11

-0.105

-0.10 0.1 0.2 0.3

Ground Clearance

Cm

Cm Vs AOA

-0.2

-0.15

-0.1

-0.05

0-2 -1 0 1 2 3 4

AOA

Cm

g h Fig. D2. Aerodynamic characteristics of a wing with AR = 4 at 15m/s.

64


0.480.5

0.520.540.560.58

0 0.05 0.1 0.15 0.2 0.25

h/c

CL


00.20.40.60.8

1

-2 0 2 4 6

AOA in deg

CL

a b

Cd Vs Ground Clearance

0.01

0.012

0.014

0.016

0.018

0 0.05 0.1 0.15 0.2 0.25

Ground Clearance

Cd

Cd Vs AOA

0

0.01

0.02

0.03

0.04

-2 0 2 4 6

AOA in Deg

Cd

c d

303234363840424446

0 0.05 0.1 0.15 0.2 0.25

ground clearance

2022242628303234363840

-2 -1 0 1 2 3 4 5

AOA in degrees

CL/

Cd

e f Cm Vs Ground Clearance

-0.094-0.092-0.09

-0.088-0.086-0.084-0.082

0 0.1 0.2 0.3

Ground Clearance

Cm

-0.15

-0.13

-0.11

-0.09

-0.07

-0.05-2 0 2 4 6

AOA

Cm

g h Fig. D3. Aerodynamic characteristics of a wing with AR = 5 at 12.5m/s.

65

0

0.2

0.4

0.6

0.8

1

1.2

0 1 2 3 4 5

AR

CL

Re = 10e5Out GroundRe = 10e7

a

0.025

0.027

0.029

0.031

0.033

0.035

0.037

0.039

0 1 2 3 4 5

AR

Cd

Re = 10e5

Out Ground

Re = 10e7

b

05

10152025303540

0 1 2 3 4 5

AR

Cl/C

d V = 10m/sOut Ground

Re = 10e7

c

Fig. D4. Aerodynamic characteristics of a wing with different AR

66

APPENDIX E – DETAIL MASS BREAKDOWN OF CRAFT

Table E1: Mass breakdown of craft by components

Component Actual Weight (kg)

Percentage Weight

(%)

Motors, servos, servo mount, receiver 0.142 9.5

Propellers 0.036 2.4

Batteries 0.113 7.6

PAR 0.197 13.2

Vertical Fin with motor connected 0.170 11.4

Horizontal tail 0.063 4.2

Fuselage 0.332 22.4

Wings 0.437 29.3

Total 1.49 100

Table E1: Mass breakdown of craft by components

67

APPENDIX F – DESIGN OF HORIZONTAL STABILIZER

From section 4.3.3, the tail can be design by referring to the tail moment characteristic

curves. The slope of the curve and the intercept can be express by the following:

tm HC V Cα

≈ −tLα

t )

= - 0.9673 - (4.5)

0 t tm H L 0 wC V C ( i iα

≈ ε + − = 0.037781 - (4.6)

In equation 4.5 and 4.6, VH is known as the tail volume ratio which can be expressed in

the following terms:

t tH

l SVSc

= - (F1)

where

tl = distance between the C.G and a/c of tail = 0.42m

tS = Area of tail

S = Area of Wing = 0.4m2

c = reference chord length of wing = 0.4m

tl , and tS c are all fixed due to the geometry of the craft, hence the only unknowns are

and tS HV .

But by letting the tail geometry to be rectangular,

tS = btail x ctail - (F2)

Then ,

68

tail tailH

tail tail

0.42 b cV0.4 0.4

2.625b c

× ×=

×= ×

In equation 4.5, is the slope of CtLC

αL of tail vs. angle of attack which is obtain by CFD

0.3

0.35

0.4

0.45

0.5

0.55

0.6

0.65

0.7

0.75

0.8

0 2 4 6

AOA

CL

8

Fig. F1. CL of tail vs. angle of attack

From Fig. F1, =4.606. Hence from equation 4.5, tLC

α

btail x ctail = 0.08, if we choose the chord of the tail to be 0.2m, the span will be

btail = 0.4m

and from equation 4.6, iw = 30

Hence it =0.650

69

APPENDIX G – CALIBRATION OF AIRSPEED SENSOR AND

FLIGHT TESTS MEASUREMENTS

Calibration of the airspeed sensor is carried out in a low speed wind tunnel located in the

Fluid Mechanics lab shown in Fig. G1.

Fig. G1. Calibration set up in a low speed wind tunnel

Calibration is done by mounting the Pitot - static tube at the center of the test section,

connected to the airspeed sensor through tubing which is located outside the wind tunnel.

(See Fig. G2). Readings are taken at an interval of approximately 2m/s from 0m/s to

20m/s. Three readings are taken from each interval and the mean value is used for

calibration. Results from the calibration are shown in Table G1 and the curve on Fig. G3.

70

Tubing leading to

Pitot Static tube

Air Speed Sensor

Fig. G2. Airspeed sensor setup for calibration.

y = 15.591x + 0.9585

0

5

10

15

20

25

0 0.2 0.4 0.6 0.8 1 1.2 1.4

Velocity

Volta

ge

Fig. G3. Airspeed sensor calibration curve

Table. G1. Calibration Results for airspeed sensor

Data Reading Free stream velocity (m/s) 1st Reading (V) 2nd Reading (V) 3rd Reading (V) Average

0 0.02 0.02 0.02 0.02 2.12 0.05 0.04 0.05 0.0466674.31 0.25 0.2 0.225 0.225 7.67 0.4 0.475 0.45 0.441667

10.02 0.525 0.5 0.55 0.525 12.52 0.675 0.7 0.675 0.68333314.05 0.825 0.775 0.85 0.81666717.1 1 1 1.05 1.016667

18.79 1.225 1.25 1.2 1.225

71

Speed vs Time

02468

1012

0 2 4 6 8 10 1

Time (sec)

Spe

ed m

/s

2

Fig. G4. Airspeed sensor readings for indoor flight test

Speed vs Time

0

2

4

6

8

10

0 5 10 15

Time (sec)

Spee

d m

/s

Fig. G5. Airspeed sensor readings for outdoor flight test

6o

Fig. G6. Measuring Angle of Attack.

72

APPENDIX H – HEIGHT MEASUREMENT

The measurement of the cruising height of the WIG is carried out by Mr. Jonathan Quan

Yong Seng. It is done using a simple piece of rope (Attached to the bottom of the craft)

that was divided into separate segments. The string attachment is shown below. A side

view of the craft was taken to measure the cruising height.

a b

Fig. H1. a. Division of string segments. b. Under view of the string setup.

A close up photo of the craft was captured while performing a flight at the MPSH. The

photos were captured using a 1/1000sec shutter speed. From the marked divisions on the

string, basic trigonometry can be use to approximate the altitude of the vehicle.

a. b Fig. H2. a. Captured side view of string during flight. b. Height approximation using

basic trigonometry.

73

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WIG Aerodynamics (1)