CFD Assisted Design and Experimental Testing of a Wingsail

CFD assisted design and experimental

testing of a wingsail with high lift devices

Y. Noguchi, D.W. Atkins Mechanical

1 Introduction

Wingsails are solid aerofoil sections and create lift/thrust in the same manner as a conventional soft sail. During the late 70's and early 80's the cost

of crude oil rose sharply. As a result of this, the U.S. Maritime Administration (MARAD) commissioned a report entitled 'Wind propulsion for the ships of the American merchant marine' |1]. This report tested seven 'Wind Assist' devices and scored each on various aspects. Of the hardware alternatives examined, wingsails offer the greatest potential.

To optimise wing sail performance, the maximum thrust is required from a given wingsail area. To accomplish this, high lift devices at both leading and trailing edges may be used as are used in aircraft wings.

An aircraft wing is only required to create a lift in one direction, whereas a wingsail must be able to operate with the flow coming from any direction and therefore is required to be symmetrical. Another unique feature for wing sail is the optimisation of many operating incidences when the drag of the wing section combines in the forward thrust produced by the wingsail as shown in fig.l.

The aim of this project is to design a wingsail with high lift devices which will create the maximum thrust, for the majority of wingsail operating conditions, thus allowing maximum thrust.

188 Computational Methods and Experimental Measurements

2 Past Experiments

To test the viability of the symmetrical leading edge high lift device, a wind tunnel model was constructed and tested as shown in fig.2. A standard NACA 0010 section was used for the slat and a NACA 0018 for the main wing. The slat section chosen fulfilled the compromise between the required thinness and structural integrity. The wing section was chosen for a relatively large radius of curvature at the nose, well known stall characteristics and the increased efficiency derived from operating a flap with a thicker wing section, as shown in Abbott & Doenhoff [2|. The slat-chord:wing-chord ratio was chosen to be 1:5, the normal ratio for an aircraft slat/wing configuration.

This model showed that a leading edge symmetrical slat did indeed augment lift as show in fig.3. The difficulty was optimising the correct slat section, wing section combination and the slat-chord:wing-chord ratio.

3 Computational Fluid Dynamics (CFD) and Initial Tests

To construct and test a model takes 8 weeks. On the other hand using a

CFD to numerically change the geometrical profile of an aerofoil is easier and a computation takes 3 days. Therefore, a decision was made to use the CFD as a design tool.

A commercially available CFD package is used to avoid time consuming development for a purpose built CFD code. The package uses the Finite Element Method (FEM). FEM offers more geometric flexibility than other methods. A K-e turbulence model originally developed by Launder and Spalding [3] is also implemented. The package has no option to switch on/off the model at a specific point. Therefore, the boundary layer is treated as turbulent from the leading edge.

3.1 Mesh Tests

A mesh test was carried out to find an optimum mesh system. Two mesh shapes, a H-mesh and an O-mesh, were used in the tests, applied to a NACA 0009 aerofoil. Three mesh densities were used; coarse (8000 nodes), medium (12000 nodes) and dense (20000 nodes).

Of the two mesh types, the O-mesh was selected for having a very low ratio of distorted elements, 25% faster solution convergence and a very efficient mesh distribution. A difference of less than 2% on lift and drag was found between the 12000 & 20000 node solutions. The dense mesh increases computation time by 2 days and requires an additional 6 Mbytes of memory. It was therefore decided to use a medium density mesh.

Computational Methods and Experimental Measurements 189

3.2 Flat Plate Tests

The accuracy of the software's boundary layer prediction was tested next, using afla tplat e model. The flat plate was tested at a Reynolds number based on the chord length of IE* per metre. Only 6% of Cf and displacement/momentum thickness variation from the theory was found and thought to be satisfactory.

3.3 Aerofoil Section Tests

Software verification was then continued by testing a standard NACA 0018 section with an identical Re number 3.2E* to that used in the original NACA test f4]. Fig.4. shows that although lift and drag are both over estimated by the computational method, the lift and drag trends for the aerofoil section are predicted correctly.

3.4 Comparison with Previous Measurements Further testing was then performed, to compare the behaviour of the experimental slat model, with a computational model at the matching Reynolds number.

The computational model again predicted the correct lift and drag trend found from the experiment as shown in fig.5. However, the magnitude of the forces was again overestimated. This overestimation is thought to be due to discrepancies in the transition from laminar to turbulent boundary layer, turbulence treatment and the 3-dimensional effects.

3.5 Conclusion from Tests The results from this comprehensive validation exercise have given confidence that the CFD package correctly predicts the trend of the lift and drag forces for a multi element aerofoil, although not the magnitude of these forces. The package can now be used to predict the optimum multi-element aerofoil design combination for the wingsail.

4 Design Procedure/Comparisons

The CFD package is utilized fully to test the variables. They are;

i) Slat section ii) Wing section iii) Flap section iv) Slat:Wing:Flap section ratio v) Slat:Wing:Flap chord ratio

190 Computational Methods and Experimental Measurements

To achieve this, the following tests were run, using the experimental Reynolds number of IE*.

Wing:Flap section thickness ratio was first tested, to discover the optimum ratio. The ratio of the wing:flap chord lengths were then tested. Following that the slat section thickness was tested and then finally the ratio of slat chord to wing chord.

5 Final Model Design

The result of this testing provided the optimum configuration of slat:wing:flap chord ratio and slat:wing chord thickness/chord (t/c) values for the wingsail. Shown in the fig.6.

Fig.7. shows the lift and drag coefficients of the original NACA 0018 section, with leading edge device only, trailing edge device only and both high lift devices. It shows an increase in lift by 130% when compared with NACA 0018, 95% over leading edge device only and 43% over trailing edge device only. It is known that the Cl achieved with the section is overestimated, it is estimated that an actual lift coefficient of Cl = 2.4 will be achieved.

6 Concluding Remarks

The project aim, to optimise the design of a wingsail with symmetrical high lift devices was fulfilled. The large number of possible design configurations lead to the introduction of a commercially available CFD package. Initial tests of the CFD package showed the capability to predict the trend, but not the magnitude of aerofoil performance. The final design improved lift by 130% when compared with the original NACA 0018, 95% over a leading edge device only and 43% over trailing edge device only. Showing that even a commercially available CFD package can be an effective and time saving tool for wingsail design.

Computational Methods and Experimental Measurements 191

T- THRUST

U-BOAT SPEED L - SAIL LIFT

D - SAIL DRAG w H - HEELING FORCE V - APPARENT WINDSPEED

W - TRUE WINDSPEED R - RESULTANT SAIL

FORCE p- APPARENT WIND ANGLE

Figure 1

PIVOT POINT OVERLAP

SYMMETRICAL SLAT DESIGN

STANDARD SLAT DESIGN

Figure 2

192 Computational Methods and Experimental Measurements

SYMMETRIC,M SLAT TEST 1 O Cl&Cd re. ALPHA /\ -7 1 .£. V. / Cd (WING + SL AT) 0.6 1 ^ \ Cl (WING+SLAT) -0.5 0.8 Cd (WING) X -0.4 -=- ^x 0.6 0 Cl (WING) X/^ I 8 0.4 0.1 0.2 0 0^ -0.1

-U.!•<>'< _ -\J.£r« o. 0^ 4 8 12 16 20 24 28

ALPHA

Figure 3

CFD vs NACA - AEROFOIL CHARS. NACA 0018 SECTION AT Re 3.2E6 1Q S\ ff- .O u.u 1.6 Cl (CFD) . " 0.5 1.4- Cd (C FD) " w * •

1.2 -0.4 Cl (NACN * * *"*

O Cd (fvJACA) * -0.3 S 0.8 * u " * * * 0.6 + n -0.2 0.4 - " . % •> " -0.1 0.2- * * n* * c ° 0-a+ a a a ft 0 4 8 12 16 20 24 28

ALPHA

Figure 4

Computational Methods and Experimental Measurements 193

Cl vs INCIDENCE (CFD vs EXPERIMENTAL) AT VARIOUS SLAT ANGLES (DEG) ^.3 21 DEG (CFD) 2 » * * 16 DEG (CFD) 9

1.5 21 DEG (EXP)

0 1 16 DEG (EXP) . * a„ c- " A 0.5 s s * " *

"HIf

5 10 15 20 25 3C ALPHA

Figure 5

FINAL AEROFOIL SECTION DESIGN

THICKNESS/CHORD (Vc) RATIO 15:18:15 %

WING CHORD RATIO 3:12:8

Figure 6

194 Computational Methods and Experimental Measurements

Cl/Cd vs INCIDENCE (CFD) 15 DEG INCIDENCE (Re 1E6)

3.5-t

PI AIM WING WING & FLAP WING & SLAT SLAT/WING/FLAP ANGLE

Figure 7

7 Acknowledgements

This project is supported by Engineering Physical Science Research Council, U.K.

8 References

111 L. Bergeson, C.K. Greenwald Sail assist developments 1979-84

Journal Wind Eng. Industrial Aerodynamics Vol.19 1985

[21 I.H. Abbott, A.E. Von Doenhoff Theory of Wing Sections Dover

[3] B.E. Launder, D.B. Spalding Mathematical Models of Turbulence London, Academic Press, 1972

[41 The Characteristics of 78 related airfoil sections from tests in

the variable density windtunnel NACA Report No.460 19th Annual Report 1933

[5| J.O. Scherer Aerodynamics of High Performance Wingsails

Marine Technology VI1. Pt.3 July 1974