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

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.

i

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.

iii

TABLE OF CONTENTS SUMMARY ........................................................................................................................ i ACKNOWLEDGEMENT............................................................................................... iii TABLE OF CONTENTS ................................................................................................ iv LIST OF FIGURES ........................................................................................................ vii LIST OF TABLES ............................................................................................................ x LIST OF SYMBOLS ....................................................................................................... xi 1.

2.

3.

4.

INTRODUCTION..................................................................................................... 1 1.1.

OBJECTIVES ..................................................................................................... 2

1.2.

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

THEORY OF GROUND EFFECT AERODYNAMICS ...................................... 4 2.1

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

2.2

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

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

DESIGN ................................................................................................................... 20

iv

4.1. 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.

5.

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

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

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

C Lα vs. h

4.9.

Cm vs. α characteristic curve for wing-fuselage combination

vii

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

viii

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

C Di

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

Cmα

Slope of Cm vs. α curve

CLα

Slope of CL vs. α curve

Cm0

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 C m αt

Slope of Tail Moment Characteristic Curve

C m0 t

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

lt

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

St

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.

2.

Conceptual Design •

Define Mission Requirement

•

Literature Survey

•

First estimate of weight and size

•

Determine the configuration layout

Preliminary Design •

Calculations carry out using CFD

•

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

3.

•

Determine stability criteria of the craft

•

Finalizing the configuration layout

Fabrication •

Modular design for ease of transportation and modification

2

4.

1.2.

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

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-tochord (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:

CL =

L

CD =

D

CM = M

1 ρ ∞ V 2S 2 1 ρ∞ V 2S 2

1 ρ∞ V 2Sc 2

- (2.1)

- (2.2)

- (2.3)

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: 5

CL =

α h

CM = −

- (2.4)

α 3h

- (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: xp =

CM 1 =− CL 3

- (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 heightto-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 ), resulting in a reduction in induced S

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 C Di =

CL 2 πeAR

-(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 e ∝ C Di ∝ h

1 hence from equation 2.7, 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

∇u* = 0

- (3.1)

∂u* 1 2 * + (u * ∇)u* = −∇p* + ∇u * ∂t Re

- (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

CL

0.8 0.79 0.78 0.77 0.76 0.75 0.74 0.73 0.72 0.71 0

2000000

4000000

Our operating region

6000000

8000000

10000000 12000000

Re

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):

Re =

UL Inertia Force = Viscous 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:

UL ν 10x1 = 1.46x10−5 ≈ 6.85x105 Re =

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

13

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: 1.3 × 105 ≤ Re ≤ 4 × 105 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 1 0.8 CL

0.6 0.4 0.2 0 -4

-3

-2

-1

0

1

2

3

4

AOA in deg Re = 2.7e5

Re = 1.4e5

Re = 4e5

Fig. 3.7. CL vs. α for different Re Cd Vs AOA 0.06 0.05 Cd

0.04 0.03 0.02 0.01 0 -4

-3

-2

-1

0

1

2

3

4

AOA in Deg 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 Propulsion ( Prop, motor + speed

Mass / kg

% total Mass

0.40

20

Structural (fuselage, wings)

1.400

70

Electronics (servos, receiver, wires )

0.200

10

Total Mass

2.000

100

controller )

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 10 = a 330 = 0.03 ≈ 0 Ma =

- (4.1)

22

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

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

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

1.2 1

CL

0.8 In Ground Out Ground

0.6 0.4 0.2 0 0

1

2

3

4

5

AR

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 CL ∝ α

h

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

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

CL vs Angle of Attack 1

CL

0.8 0.6 0.4 0.2 -4

-2

0

2

4

6

CL = 0.1007α + 0.6139 CL = 0.0997α + 0.584 CL = 0.0903α + 0.5359

AOA in deg

Fig. 4.6. CL vs. α characteristic curve for wing-fuselage combination with various h CL vs ground clearance 0.64 0.62

CL

0.6

CL = 0.3641h-0.2136

0.58 0.56 0.54 0.52 0.5 0.07

0.09

0.11

0.13

0.15

0.17

0.19

h/c

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:

C L = C L α α + C L0

- (4.1)

where CLα is the gradient of the curve, CL0 is lift coefficient at 0 angle of attack and α is expressed in degrees. Observe the value of CLα 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 CLα for different h can be found by:

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

- (4.2)

The next unknown to be determined will be CL0 which can be obtained from Fig. 4.7 for different h and CL can be calculated from equation 4.1. To summarize the procedure, to calculate CL for different height and α: 1. Determine CLα and CL0 for given h from Fig. 4.8 and Fig. 4.7 respectively.

Gradient of CL

2. Calculate CL from CLα and CL0 as obtain from step 1 for given α.

Gradient of CL = -1.1897(h/c)3 - 1.3265(h/c)2 + 0.2001(h/c) + 0.0941 0.102 0.1 0.098 0.096 0.094 0.092 0.09 0.088 0.08 0.09 0.1 0.11 0.12 0.13 0.14 0.15

0.16

h/c

Fig. 4.8. CLα vs. h

29

CM vs AOA at Leading Edge

-3

-2

-1

-0.5 -0.7 0 -0.9 -1.1 -1.3 -1.5 -1.7 -1.9 -2.1 -2.3 -2.5

1

2

3

4

5

Cm = -0.2059α - 1.3804

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]: xa / c =

Cmα CLα

- (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 Propulsion ( Prop, motor + speed

Mass / kg

% total Mass

0.350

21.5

Structural (fuselage, wings)

1.100

67.5

Electronics (servos, receiver, wires )

0.180

11

Total Mass

1.630

100

controller )

Base on the second estimate of the weight of the craft, the required cruising speed is obtained from the Lift: L 1 2 ρ∞SC L

V= =

1 2

16 × 1.23 × 0.506 × 0.59

≈ 9.3m / s

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

3.00

Force N

2.50

Requirements Match

2.00

Drag 7"-4 blade 7"-2 blade 5.6"-4 blade 5.25"-2 blade

1.50 1.00 0.50 0.00 0

5

10

15

20

Speed

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]:

Cmα < 0 and

- (4.4)

Cm 0 > 0

- (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 0.2 0.1

Cm

0 -3

-2

-1

-0.1

0

1

2

3

4

5

-0.2 -0.3 -0.4 0c

0.333c

0.5c

AOA

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. 0.1 0.08

Cm

0.06 0.04

Wing-fuselage

0.02

Wing-fuselage-tail Tail

0 -4

-2

-0.02

0

2

4

6

-0.04 -0.06 AOA

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

Cmαt ≈ − VH CLαt

- (4.7)

Cm0 t ≈ VH CLαt (ε0 + i w − i t )

- (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. Cm0 t will therefore determine the angle of incident of the tail, it. From calculation, the tail size

39

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.

40

Plan View

Side View

Front View

Fig. 4.16. Resulting Layout of WIG

41

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

CL,indoor = =

L 2 1 2 ρV S

14.9 2 1 2 × 1.23 × 9.53 × 0.506

= 0.499

Comparing the CL,indoor value obtain from the flight test to the CFD prediction presented in Fig. 4.6, CL,CFD at 0o 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: CD,indoor = =

1 2

1 2

T ρV 2S

0.8835 ×1.23 × 9.532 × 0.506

= 0.0312

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: CL,outdoor = =

1 2

L 2 1 2 ρV S

14.9 ×1.23 × 8.52 × 0.506

= 0.85

Because it is flying at a lower ground clearance, Fig. 4.7 will be use to extrapolate CL,CFD at 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, CLα , at h = 0.05 by using equation 4.2. Finally, CL at h = 0.05 and α = 60 can be evaluated by equation 4.1:

CL,CFD = 0.1007α + 0.69 = 0.1007(6) + 0.69 = 1.2942

From the results of the indoor flight test, the actual CL value on a hard ground is 15% lesser than CFD prediction, hence CL,indoor = 1.1 vs. CL,outdoor = 0.85 there is a 23 % difference. When compared to the prediction by CFD, CL,outdoor 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.

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 twodimensional 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: ∇u = 0

- (B.1)

δu 1 + (u∇)u = − ∇p + ν∇ 2u δt ρ

- (B.2)

δT k 2 + (u∇)T = ∇T δt ρc p

- (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 x* =

u* =

x * y * z * t , y = ,z = , t = L L L L / V∞ u v w ρ , v* = , w* = , ρ* = ρ∞ U ∞2 U∞ U∞ U∞

}

−(B.4)

59

Substituting equation B.4 into B.1 and B.2 yields the following non-dimensional form of the incompressible N-S equations: ∇u* = 0

- (3.1)

∂u* 1 2 * + (u * ∇)u* = −∇p* + ∇u * ∂t Re

- (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 ρU 2 and the viscous force is result from the shear stress, τ = μ

∂u U ≈ μ . Hence by taking the ratio L ∂y

between the two: Inertia ρU 2 ρUL = = Viscous μU / L 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 nonuniform pressure field will appear to be uniform when if the usual central difference case is applied. Consider a simplified one dimensional convection equation: ∂u ∂p +c =0 ∂t ∂x

- (C.1)

After applying the central difference scheme on the pressure field and the explicit Euler on the time derivative yields: u in +1 = u in −

Δt (pin+1 − pin−1 ) 2Δx

- (C.2)

Since pin+1 = pin−1 , then u in +1 = u in 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: u in +1 = u in − = u in −

Δt (−3pin + 4pin−1 − pin− 2 ) 2Δx

Δt (−4pin ) 2Δx

- (C.3)

- (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: u in++1/1 2 = u in+1/ 2 −

Δt n (pi +1 − pin ) Δx

- (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 0.06

-4

-2

0.05 0.04 Cd

CL

0.9 0.8 0.7 0.6 0.5 0.4 0.3 0.2

0.03 0.02 0.01 0

0

2

-4

4

-2

0

a

b

-0.06 -3

-2

-1

4

AOA in Deg

AOA in deg

-4

2

20

0

1

2

3

4

19

-0.07

18

-0.09

17 CL/Cd

Cm

-0.08

16 15

-0.1

14 13

-0.11 -0.12

12 -4

-3

AOA in degrees

-2

-1

0

1

2

3

4

AOA in Deg

c

d

Fig. D1. Aerodynamic characteristics of a wing-fuselage combination at 10m/s with ground clearance h/c = 0.15.

63

CL vs Angle of Attack

0.56

1

0.54

0.8

0.52

0.6

CL

CL

CL vs ground clearance

0.5

0.4

0.48

0.2

0.46

0 0

0.05

0.1

0.15

0.2

0.25

-2

-1

0

h/c

a

b

Cd Vs Ground Clearance

Cd Vs AOA

0.018

0.04

0.016

0.03 Cd

0.014

Cd

1

2

3

4

2

3

4

AOA in deg

0.02

0.012

0.01

0.01

0

0

0.1

0.2

0.3

-2

-1

0

Ground Clearance

1 AOA in Deg

c

d Cd Vs AOA

45 43 41

0.04 0.03

37 35

Cd

CL/Cd

39

33

0.02

31

0.01

29

0

27 25 0

0.05

0.1

0.15

0.2

-2

0.25

4

f

Cm Vs Ground Clearance

Cm Vs AOA

0.1

0.2

0.3

-2 Cm

-0.105

2

e

-0.1 0

0

-0.11 -0.115

6

AOA in Deg

Ground clearance

Cm

Cd

0 -1 0 -0.05

1

2

3

4

-0.1 -0.15

-0.12

-0.2

Ground Clearance

AOA

g h Fig. D2. Aerodynamic characteristics of a wing with AR = 4 at 15m/s. 64

CL vs Angle of Attack

0.58

1

0.56

0.8

0.54

0.6

CL

CL

CL vs ground clearance

0.52

0.4

0.5

0.2

0.48

0 0

0.05

0.1

0.15

0.2

0.25

-2

0

2

4

6

4

6

AOA in deg

a

b

Cd Vs Ground Clearance

Cd Vs AOA

0.018

0.04

0.016

0.03

0.014

Cd

Cd

h/c

0.02

0.012

0.01

0.01

0

0

0.05

0.1

0.15

0.2

0.25

-2

0

Ground Clearance

2 AOA in Deg

c

d

46

40 38 36 34 32 30 28 26 24 22 20

44 42 CL/Cd

40 38 36 34 32 30 0

0.05

0.1

0.15

0.2

0.25

-2

-1

0

g r o und clear ance

1

2

3

4

5

AOA in degrees

e

f

-0.082 -0.084 0 -0.086 -0.088 -0.09 -0.092 -0.094

-2

-0.05

0

2

4

6

-0.07

0.1

0.2

0.3

Cm

Cm

Cm Vs Ground Clearance

-0.09 -0.11 -0.13 -0.15

Ground Clearance

AOA

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

65

1.2 1

CL

0.8

Re = 10e5

0.6

Out Ground

0.4

Re = 10e7

0.2 0 0

1

2

3

4

5

AR

a 0.039 0.037

Cd

0.035

Re = 10e5

0.033

Out Ground

0.031

Re = 10e7

0.029 0.027 0.025 0

1

2

3

4

5

AR

b

Cl/Cd

40 35 30 25 20 15 10

V = 10m/s Out Ground Re = 10e7

5 0 0

1

2

3

4

5

AR

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 Percentage Weight Component Actual Weight (kg)

(%)

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:

Cmαt ≈ − VH CLαt = - 0.9673 Cm0 t ≈ VH CLαt (ε0 + i w − i t ) = 0.037781

- (4.5) - (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: VH =

l t St Sc

- (F1)

where l t = distance between the C.G and a/c of tail = 0.42m St = Area of tail S = Area of Wing = 0.4m2

c = reference chord length of wing = 0.4m

l t , St and c are all fixed due to the geometry of the craft, hence the only unknowns are St and VH . But by letting the tail geometry to be rectangular, St = btail x ctail

- (F2)

Then , 68

0.42 × b tail × c tail 0.4 × 0.4 = 2.625b tail × c tail VH =

In equation 4.5, CLαt is the slope of CL of tail vs. angle of attack which is obtain by CFD

0.8 0.75 0.7 0.65 CL

0.6 0.55 0.5 0.45 0.4 0.35 0.3 0

2

4

6

8

AOA

Fig. F1. CL of tail vs. angle of attack From Fig. F1, C Lαt =4.606. Hence from equation 4.5, 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. 25 y = 15.591x + 0.9585 Voltage

20 15 10 5 0 0

0.2

0.4

0.6

0.8

1

1.2

1.4

Velocity

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) 0 0.02 0.02 0.02 2.12 0.05 0.04 0.05 4.31 0.25 0.2 0.225 7.67 0.4 0.475 0.45 10.02 0.525 0.5 0.55 12.52 0.675 0.7 0.675 14.05 0.825 0.775 0.85 17.1 1 1 1.05 18.79 1.225 1.25 1.2

Average 0.02 0.046667 0.225 0.441667 0.525 0.683333 0.816667 1.016667 1.225

71

Speed vs Time 12 Speed m/s

10 8 6 4 2 0 0

2

4

6

8

10

12

Time (sec)

Fig. G4. Airspeed sensor readings for indoor flight test Speed vs Time

Speed m/s

10 8 6 4 2 0 0

5

10

15

Time (sec)

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.

b a. Fig. H2. a. Captured side view of string during flight. b. Height approximation using basic trigonometry. 73