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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.
ii
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.
iii
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.
iv
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
v
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
vi
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
vii
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 LCα
vs. h
4.9. Cm vs. α characteristic curve for wing-fuselage combination
viii
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
ix
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.
x
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
xi
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
( )∇ i Divergence Operator
2∇ Laplacian Operator
Re Reynolds Number
U Reference Velocity
L Reference Length
ν Kinematics Viscosity
xii
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
1
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.)
2
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
3
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
4
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.
5
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 ∞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:
6
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
C 1xC 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 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.
7
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.
8
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):
9
* 0u∇ =i - (3.1)
2*
1( )t Re
** * * *u u u p u∂ + ∇ = −∇ + ∇
∂i i i - (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].
10
0.710.720.730.740.750.760.770.780.790.8
0 2000000 4000000 6000000 8000000 10000000 12000000
Re
CL
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.
Our operating region
11
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
12
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.
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.
Wall
Velocity Inlet
Outlet
Moving Wall
13
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
14
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
15
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
16
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
17
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,
18
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.
19
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
20
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.
21
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.
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
Table 4.1: First Estimation of mass breakdown of components
22
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)
23
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.
24
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.
25
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.
26
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
27
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 .
28
CL vs Angle of Attack
CL = 0.0997α + 0.584CL = 0.0903α + 0.5359
CL = 0.1007α + 0.6139
0.2
0.4
0.6
0.8
1
-4 -2 0 2 4 6
AOA in deg
CL
h=0.1ch=0.15ch = 0.085cLinear (h=0.1c)Linear (h=0.15c)Linear (h = 0.085c)
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 L LC C Cα
= α + - (4.1)
where LCα
is the gradient of the curve, 0LC is lift coefficient at 0 angle of attack and α is
expressed in degrees. Observe the value of LCα
for different h and they are plotted in
29
Fig. 4.8 and their relationship can be expressed by fitting a cubic curve onto the data.
Hence LCα
for different 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 0LC 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 LCα
and 0LC for given h from Fig. 4.8 and Fig. 4.7 respectively.
2. Calculate CL from LCα
and 0LC 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
0.0880.09
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
30
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
31
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.
32
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.
33
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.
34
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.
35
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.
Table 4.2: Second Estimation of mass breakdown of components
36
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
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.
Requirements Match
37
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.
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.
c
38
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.
39
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
t tm H LC V Cα α
≈ − - (4.7)
0 t tm H L 0 w tC V C ( i i )α
≈ 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. 0 tmC will
therefore determine the angle of incident of the tail, it. From calculation, the tail size
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
42
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.
43
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.
44
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:
45
L,indoor 212
212
LCV S
14.91.23 9.53 0.506
0.499
=ρ
=× × ×
=
Comparing the L,indoorC value obtain from the flight test to the CFD prediction presented
in Fig. 4.6, L,CFDC 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:
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
46
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
47
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
48
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
L,CFDC 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,
LCα
, 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:
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, L,outdoorC 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.
49
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
50
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.
51
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
52
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.
53
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
54
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.
55
APPENDICES
56
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
57
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.
58
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
59
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∇ =i - (B.1)
21( )tu u u p uδ + ∇ = − ∇ + ν∇
δ ρi i i - (B.2)
2
p
T k( ) T Tt c
uδ + ∇ = ∇δ ρ
i i - (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
= , * zzL
= , * ttL / V∞
=
* uuU∞
= , * vvU∞
= , * wwU∞
= , *2U∞ ∞
ρρ =ρ
} (B.4)−
60
Substituting equation B.4 into B.1 and B.2 yields the following non-dimensional form of
the incompressible N-S equations:
* 0u∇ =i - (3.1)
2*
1( )t Re
** * * *u u u p u∂ + ∇ = −∇ + ∇
∂i i i - (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 2Uρ and the
viscous force is result from the shear stress, u Uy L
∂τ = µ ≈ µ∂
. Hence by taking the ratio
between the two:
2Inertia U ULViscous U / L
ρ ρ= =µ µ
61
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 pc 0t x
∂ ∂+ =∂ ∂
- (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 1
tu u (p p )2 x
++ −
∆= − −∆
- (C.2)
Since n ni 1 i 1p p+ −= , then n 1 n
i iu u+ = 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:
n 1 n n n ni i i i 1 i 2
tu u ( 3p 4p p )2 x
+− −
∆= − − + −∆
- (C.3)
n ni i
tu ( 4p )2 x∆= − −∆
- (C.4)
62
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 p )x
++ + +
∆= − −∆
- (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.
63
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.
64
CL vs ground clearance
0.460.480.5
0.520.540.56
0 0.05 0.1 0.15 0.2 0.25
h/c
CL
CL vs Angle of Attack
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.
65
CL vs ground clearance
0.480.5
0.520.540.560.58
0 0.05 0.1 0.15 0.2 0.25
h/c
CL
CL vs Angle of Attack
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.
66
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
67
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
68
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:
t tm H LC V Cα α
≈ − = - 0.9673 - (4.5)
0 t tm H L 0 w tC V C ( i i )α
≈ 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 , tS and c are all fixed due to the geometry of the craft, hence the only unknowns are
tS and HV .
But by letting the tail geometry to be rectangular,
tS = btail x ctail - (F2)
Then ,
69
tail tailH
tail tail
0.42 b cV0.4 0.4
2.625b c
× ×=×
= ×
In equation 4.5, tLC
α is the slope of CL 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 8
AOA
CL
Fig. F1. CL of tail vs. angle of attack
From Fig. F1, tLC
α=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
70
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.
71
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
Air Speed Sensor
Tubing leading to
Pitot Static tube
72
Speed vs Time
02468
1012
0 2 4 6 8 10 12
Time (sec)
Spe
ed m
/s
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
Fig. G6. Measuring Angle of Attack.
6o
73
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.