Thesis From Delf Univ

Master of Science Thesis

Boundary Layer Suction on a Horizontal Axis

Wind Turbine

An Aerodynamic Design of a Thick Airfoil for Application

Lex Zwang

November 12, 2009

Boundary Layer Suction on a Horizontal

Axis Wind Turbine

An Aerodynamic Design of a Thick Airfoil for Application

Master of Science Thesis

For obtaining the degree of Master of Science in Aerospace

Engineering at Delft University of Technology

Lex Zwang

November 12, 2009

Faculty of Aerospace Engineering Delft University of Technology

Actiflow B.V. Breda

Copyright © L. Zwang B.Sc.

All rights reserved.

Delft University Of Technology

Department Of

Aerodynamics and Wind Energy

The undersigned hereby certify that they have read and recommend to the Faculty of Aerospace

Engineering for acceptance a thesis entitled “Boundary Layer Suction on a Horizontal Axis Wind

Turbine” by Lex Zwang in partial fulfilment of the requirements for the degree of Master of Science.

Dated: November 12, 2009

Supervisors:

Prof.dr. G.J.W. van Bussel

ir. A. Barlas

ir. E. Terry

ir. A.W. Hulskamp

ir. W.A. Timmer

Abstract i

M.Sc. Thesis Lex Zwang

Abstract

In the past decades the size of wind turbines blades is considerably increased, there are even

prototypes designed and build with blade diameters up to 120 meters. This growth in blade length

and turbine size tends to make the blades a larger proportion of the production costs. However, the

blades themselves only represents 10-15% of total system costs, therefore reducing the Cost Of

Energy is limited. Nevertheless a reduction of COE is possible, an innovative blade design that

reduces loading and thereby affecting other major components (tower, drive train) will have a

snowball effect on the reduction of the costs. For advanced wind turbine designs an increasing

thickness of the airfoils is a major contributor to the restraining mass growth of the wind turbine,

though the increase in thickness will affect the aerodynamic performance of the blade. The objective

of this study is to design a thicker airfoil for a horizontal axis wind turbine with the application of

boundary layer suction. The extra thickness increases the stiffness of a wind turbine blade while the

desired aerodynamic performance is maintained by the application of boundary layer suction.

For this end a series of airfoils are produced with different thicknesses and originate from the DU-91-

W2-250 wind turbine airfoil. A script file then controls RFOIL-suc, to produce the new airfoils and

their polars, and Matlab to perform the validation of these airfoils by means of the selection criteria.

The possible candidate airfoils were fine-tuned by the script file, the camber of the airfoil is adjusted

and the airfoil goes through a similar selection process. The newly designed airfoil, the AF-0901, was

optimized while keeping equal aerodynamic performance in terms of maximum lift-to-drag ratio,

maximum lift coefficient and smooth stall control. A high maximum lift-to-drag ratio yields a high

power output, while minimizing the maximum lift coefficient and obtaining smooth stall control

reduces the operational loads on a wind turbine. The variation of the suction velocity is a possible

design option for engineers, this variation will be an assessment between additional power output or

reduced loads.

For the evaluation of the structural benefits of the newly designed airfoil, a structural 2-D

comparison was made between the ‘old’ DU-91-W2-250 and the ‘new’ AF-0901 airfoil. First a simple

beam theory and more accurate XFOIL thin walled shell stiffness data was used to compare both

airfoils. As a second comparison, a detailed structural design with actual lay-up data from the

reference wind turbine was implemented in PreComp. The new 20% thicker AF-0901 airfoil was then

optimized to match the original stiffness requirements. For the simple beam theory only the spar

caps were taken into account and for the shell structure one global material was assumed for the

different sections of the airfoil. These methods gave a reduction in materials of 17%. For a more in

depth analysis the program PreComp was applied, therefore a structural analysis with different

laminates and materials for the spar caps, webs and skins is possible. First, overall material used for

the airfoil was reduced. With this approach the mass density, the section mass per unit length, was

reduced with more than 28%. Secondly, the spar cap material was decreased as with the simple

beam theory. Reducing only the materials used for the spar caps gave a reduction in mass of almost

19%.

Acknowledgements iii


Acknowledgements

I wish to thank the following persons for their assistance and knowledge. First of all I would like to

thank the people at Actiflow for a great time during my graduation project and also had a

contribution to this thesis; Eric, Roy, Roland, Jeroen, Oscar and Vincent. My supervisors for their time

and assistance; ir. E. Terry from Actiflow, ir. A. Barlas and Prof.dr. G.J.W. van Bussel from Delft

University of Technology. Additionally, I also owe many thanks to ir. L.M.M. Boermans and ir. W.A.

Timmer for sharing their knowledge on airfoil design and ir. A.W. Hulskamp for gaining insight into

the structural analysis of a wind turbine blade.

Finally, thanks to thank all my friends and family for their support during my time as an aerospace

engineering student. In particular I like to thank my parents and my girlfriend for their unconditional

and never-ending support during my time in Delft.

Lex Zwang

Utrecht, November 12, 2009

https://phonebook.tudelft.nl/detail.html?uid=e9383b12bd44d776

https://phonebook.tudelft.nl/detail.html?uid=e9383b12bd44d776

Table of Contents v


Table of Contents

Abstract .................................................................................................................................................... i

Acknowledgements ................................................................................................................................. iii

List of Figures ........................................................................................................................................... ix

List of Symbols ......................................................................................................................................... xi

1 Introduction ..................................................................................................................................... 1

1.1 Introduction ............................................................................................................................. 1

1.2 Objectives of present study ..................................................................................................... 2

1.3 Outline of the report ............................................................................................................... 2

2 Theory overview .............................................................................................................................. 3

2.1 Structural design...................................................................................................................... 3

2.1.1 The main loads ................................................................................................................ 3

2.1.2 The structure of an airfoil ................................................................................................ 3

2.1.3 The simple beam theory .................................................................................................. 4

2.2 Airfoil design ............................................................................................................................ 6

2.3 Boundary layer control ............................................................................................................ 7

2.3.1 Boundary layer suction .................................................................................................... 7

2.3.2 Moving thickness backwards ........................................................................................... 7

3 Design Process ................................................................................................................................. 9

3.1 Design choices ......................................................................................................................... 9

3.1.1 Reference wind turbine ................................................................................................... 9

3.1.2 Suction distribution ....................................................................................................... 11

3.2 RFOIL-suc ............................................................................................................................... 13

3.3 Profile field ............................................................................................................................ 14

3.4 Selection criteria .................................................................................................................... 15

3.5 Flow diagram ......................................................................................................................... 16

3.6 Results from the profile field ................................................................................................. 21

3.7 Modify camber ...................................................................................................................... 33

3.8 Wind tunnel adjustments ...................................................................................................... 37

4 Structural benefits ......................................................................................................................... 39

4.1 Thin walled shell structure .................................................................................................... 39

4.2 The PreComp method ........................................................................................................... 41

vi Table of Contents

Lex Zwang M.Sc Thesis

4.2.1 Assumptions .................................................................................................................. 42

4.2.2 PreComp results ............................................................................................................ 43

5 Design issues ................................................................................................................................. 47

5.1 Modify suction distribution ................................................................................................... 47

5.1.1 Variation of suction length ............................................................................................ 47

5.1.2 Variation of suction velocity .......................................................................................... 48

5.1.3 Apply linear suction ....................................................................................................... 50

5.2 Suction variation due to gusts ............................................................................................... 52

5.3 Wind tunnel considerations .................................................................................................. 53

5.4 Comparison with similar profile ............................................................................................ 54

6 Conclusions and recommendations .............................................................................................. 57

6.1 Conclusions ............................................................................................................................ 57

6.2 Recommendations................................................................................................................. 58

6.2.1 Wind tunnel test for the AF-0901 airfoil ....................................................................... 58

6.2.2 Redesign of other airfoils to fit BLS ............................................................................... 58

6.2.3 Intensive study to optimize structural design ............................................................... 58

A Airfoil example files ....................................................................................................................... 59

A.1 Coordinate file ....................................................................................................................... 59

A.2 Suction distribution file ......................................................................................................... 59

A.3 Polar file from RFOIL .............................................................................................................. 60

B Visual Basic Script files .................................................................................................................. 61

B.1 CreateProfileField.vbs ........................................................................................................... 61

B.2 CreateSuction.vbs .................................................................................................................. 62

B.3 CreatePolar.vbs ..................................................................................................................... 63

C MATLAB files .................................................................................................................................. 67

C.1 CreateCorrectSuction.M ........................................................................................................ 67

C.2 FilterData.M........................................................................................................................... 68

C.3 MATLAB example data .......................................................................................................... 72

D Structural properties tables .......................................................................................................... 73

E Structural lay-up pictures .............................................................................................................. 75

F PreComp example input files ........................................................................................................ 79

F.1 Main Input File: “DU25.pci” .................................................................................................. 79

F.2 Airfoil Data File: “af-du25.inp” .............................................................................................. 79

F.3 Internal Structure Data File: “int-du25.inp” .......................................................................... 80

Table of Contents vii


F.4 Materials Data File: “materials.inp” ...................................................................................... 82

G PreComp example output files ...................................................................................................... 83

G.1 BModes Output File: “du25.out_bmd” ................................................................................. 83

H Pressure gradient in a closed rotating duct .................................................................................. 85

Bibliography ........................................................................................................................................... 87

List of Figures ix


List of Figures

Figure 1-1: Mass growth for commercial MW-scale blade design [21] .................................................................. 1

Figure 2-1: Common structural architecture for wind turbine blade cross section [21] .......................................... 4

Figure 2-2: Simplified structural model of a cross section [8] ................................................................................. 4

Figure 2-3: Development of the boundary layer affected by boundary layer suction ............................................. 7

Figure 3-1: DU-91-W2-250 profile normalized to its chord c ................................................................................ 10

Figure 3-2: Pressure along the length of the blade [1] .......................................................................................... 11

Figure 3-3: Suction area on an example profile .................................................................................................... 12

Figure 3-4: Suction velocity (v/U) vs chord wise position (x/c) .............................................................................. 13

Figure 3-5: Airfoil and wake panelling with vorticity () and source distributions (), with TE detail [6] ............. 13

Figure 3-6: Profile field of new created airfoils ..................................................................................................... 15

Figure 3-7: Polar curves for DU-91-W2-250 (Re = 8e6 and c/r = 0.09) .................................................................. 16

Figure 3-8: General flowchart design process ....................................................................................................... 17

Figure 3-9: CL/CD-max for v/U = 0 and clean configuration ...................................................................................... 18

Figure 3-10: Candidate airfoils for v/U = 0 (clean) ................................................................................................ 19

Figure 3-11: Definition of alternative selection criteria......................................................................................... 20

Figure 3-12: Polar examples of 1-5.air (Re = 8e6, c/r = 0.09 and clean) ................................................................ 20

Figure 3-13: CL,max and CL,design for clean conditions and v/U = 0 ........................................................................... 21

Figure 3-14: CL,max and CL,design for soiled conditions and v/U = 0 ........................................................................... 22

Figure 3-15: Result airfoils with thickness and position ........................................................................................ 23

Figure 3-16: Polars of the selected airfoils with v/U = 0.003 (clean; Re = 8e6; c/r = 0.09) .................................... 24

Figure 3-17: Polars of the selected airfoils with v/U = 0.003 (soiled; Re = 8e6; c/r = 0.09) .................................... 24

Figure 3-18: CL/CD-max for v/U = 0, clean configuration and Re = 8e6 ..................................................................... 25

Figure 3-19: CL/CD-max for v/U = 0.001, clean configuration and Re = 8e6 .............................................................. 25



Figure 3-22: CL/CD-max for v/U = 0, soiled conditions and Re = 8e6 ......................................................................... 27

Figure 3-23: CL/CD-max for v/U = 0.001, soiled conditions and Re = 8e6 .................................................................. 27



Figure 3-26: Candidate airfoils for v/U = 0 (clean) ................................................................................................ 29

Figure 3-27: Candidate airfoils for v/U = 0.001 (clean) ......................................................................................... 29



Figure 3-30: Candidate airfoils for v/U = 0 (soiled) ............................................................................................... 31

Figure 3-31: Candidate airfoils for v/U = 0.001 (soiled) ........................................................................................ 31



Figure 3-34: Camber distribution for example profile ........................................................................................... 33

Figure 3-35: Effect of moving the camber point for v/U = 0.003 (soiled; Re = 8e6; c/r = 0.09) .............................. 34

Figure 3-36: Effect of up scaling the camber value for v/U = 0.003 (soiled; Re = 8e6; c/r = 0.09) ......................... 35

Figure 3-37: Result of camber modification for v/U = 0.003 (soiled; Re = 8e6; c/r = 0.09) .................................... 36

Figure 3-38: Result of camber modification for v/U = 0.003 (clean; Re = 8e6; c/r = 0.09) ..................................... 36

Figure 3-39: Airfoil growth during the design process .......................................................................................... 37

Figure 3-40: Effect of wind tunnel modifications for v/U = 0.003 (clean; Re = 8e6; c/r = 0.09) ............................. 38

Figure 3-41: Effect of wind tunnel modifications for v/U = 0.003 (soiled; Re = 8e6; c/r = 0.09) ............................ 38

x List of Figures


Figure 4-1: Bending parameters from XFOIL ......................................................................................................... 39

Figure 4-2: Thickness ratio vs. bending moments ratio ......................................................................................... 41

Figure 4-3: Profile lay-out used for comparison .................................................................................................... 43

Figure 5-1: Polars of the AF-0901 airfoil for different suction lengths (soiled case; Re = 8e6; v/U = 0.003; c/r =

0.09) ...................................................................................................................................................................... 48

Figure 5-2: Effect of suction velocity (clean case; Re = 8e6; c/r = 0.09) ................................................................ 49

Figure 5-3: Effect of suction velocity (soiled case; Re = 8e6; c/r = 0.09) ................................................................ 50

Figure 5-4: Different suction distributions ............................................................................................................. 51

Figure 5-5: Effect of different suction distributions (soiled case; Re = 4e6; c/r = 0) .............................................. 52

Figure 5-6: Comparison of XFOIL, RFOIL and wind tunnel tests (clean case; Re = 3e6; M = 0.21; c/r = 0) ............ 54

Figure 5-7: Comparison of XFOIL, RFOIL and wind tunnel tests (soiled case; Re = 3e6; M = 0.22; c/r = 0) ........... 54

Figure 5-8: Comparison of airfoils with RFOIL (Re = 3e6; M = 0.22; c/r = 0) ......................................................... 55

Figure E-1: Blade profile lay-out [14] .................................................................................................................... 75

Figure E-2: Blade laminate lay-up [14] .................................................................................................................. 76

Figure E-3: Profile lay-out example [14] ................................................................................................................ 77

Figure E-4: Ply angle definition [2] ........................................................................................................................ 77

List of Symbols xi


List of Symbols

Abbreviations

AF ActiFlow COE Cost Of Energy DU Delft University HAWT Horizontal Axis Wind Turbine NACA National Advisory Committee for Aeronautics NREL National Renewable Energy Laboratory VBS Visual Basic Script

Greek Symbols

α Angle of attack ° θ Section pitch angle ° ρ Density kg/m2

σ Stress N/m2

φ Inflow angle ° Ω Rotational speed rad/s

Mathematical Symbols

A Area m2

a Axial induction factor - a Width m a’ Tangential induction factor - c Chord length m CD Drag coefficient - CL Lift coefficient - Cp Pressure coefficient - D Drag N d Height m E Young’s modulus Pa I Area moment of inertia m4

L Lift N MX Flap wise bending moment Nm MY Edgewise (lead-lag) bending moment Nm N Normal force N R Overall radius m r Local radius m t Layer thickness m t Spar cap thickness m t Wall thickness m U Air velocity at the blade m/s v Absolute suction velocity m/s Vwind Wind speed m/s

Introduction 1


1 Introduction

1.1 Introduction

In the past decades the size of wind turbine blades is considerably increased, there are even

prototypes designed and build with blade diameters up to 120 meters. This growth in blade length

and turbine size tends to make the blades a larger proportion of the production costs. However, the

blades themselves only represents 10-15% of total system costs, therefore reducing the Cost Of

Energy (COE) is limited [21]. Nevertheless a reduction of COE is possible, an innovative blade design

that reduces loading and thereby affecting other major components (tower, drive train) will have a

snowball effect on the reduction of the costs.

Following the laws of physics the mass of a wind turbine grows according to R3, the cubic law of mass

growth. The trend line of the blade mass in Figure 1-1 shows however a growth of R2.3, with the use

of advanced technologies the cubic law was beaten. The scatter of data is mainly due to different

design approaches of the manufactures. For example, the growth rate of the highlighted Vestas

blades is close to cubic law (R2.7) because it was already a lightweight design, the lower growth rate

than the cubic law is probably due to the use of thicker airfoils. On the other hand the growth rate of

the LM class is as low as R1.7, this is also due to thicker airfoils but especially the use of improved

materials here leads to the lower growth rate. For advanced wind turbine designs the increasing

thickness of the airfoils is thus a major contributor to the restraining mass growth. However, the

increase in thickness will negatively affect the aerodynamic performance of the blade.

Figure 1-1: Mass growth for commercial MW-scale blade design [21]


One of the possibilities to improve the aerodynamic performance of thicker airfoils is applying

boundary layer suction. In 2009 Actiflow BV, a spin-off company from the Technical University of

Delft, performed a feasibility study for the application of boundary layer control in wind turbine

design. This study showed that for airfoils with an increased thickness the application of boundary

layer suction results in the aerodynamic performance of relatively thinner airfoils. However, the

airfoils used here were not designed for use without boundary layer control and an aerodynamic

redesign of the blade would be beneficial for a successful applications of boundary layer suction in

wind turbine blade design [11].

1.2 Objectives of present study

During the research on boundary layer suction at Actiflow BV it became clear that drastically

changing the aerodynamic geometric design of the blade would have the largest potential in finding

successful applications for boundary layer suction in wind turbine blade design. The aim is now to

redesign an existing profile for use of boundary layer suction and thus focusing on the aerodynamic

design of a new wind turbine blade which uses boundary layer suction. The main objective of this

study can be summarized as follows:

“To design a thicker airfoil for a horizontal axis wind turbine with the application of boundary layer

suction, thereby increasing the stiffness of the wind turbine blade while maintaining the desired

aerodynamic performance.”

1.3 Outline of the report

In chapter 1 an introduction into the subject of this study is given, along with a motivation for the

start of this project. Also the objective of this study is given in this chapter. An introduction into the

theory of the subject is presented in chapter 2, including an overview of the structural and

aerodynamic design of an airfoil and a small overview of boundary layer control is given. In chapter 3

the complete design process of the new airfoil will be discussed. Chapter 4 gives a description of the

structural benefits gained by this new airfoil. In chapter 5 the design issues belonging to this airfoil

are treated. The conclusions and recommendations that follow from this study are mentioned in

chapter 6. In appendix A some example files for the calculations are given, appendix B gives the used

visual basic script files and the Matlab files are shown in appendix C. Structural material data is given

in appendix D and accompanied pictures in appendix E. The input and output files for PreComp are

respectively given in appendix F en appendix G. The calculation of a closed rotating duct can be

found in appendix H.

Theory overview 3


2 Theory overview

In this chapter a small overview is given of the theory used for this thesis. The most important loads

on a wind turbine needed for a simple structural analysis are discussed in the first part. Section 2.2 is

a summary with highlights for the aerodynamic design of a wind turbine blade. The basics of

boundary layer control are given in the last section.

2.1 Structural design

2.1.1 The main loads

Wind turbines extract energy from the wind by slowing down the wind using a force in the upwind

direction, the thrust. This thrust is caused by a pressure jump over the rotor, induced by the flow

past the rotor blades. Besides a normal component to the flow, the thrust, there is also a tangential

component of the force in the rotational direction of the blades, which delivers the shaft torque.

Loads are transformed to a position close to the root as flap wise and edgewise (lead-lag) bending

moments together with a yaw and tilt rotor moment. The flap wise bending moment (MX) comes

mostly from the thrust and tends to deflect the blades out of the rotor plane in the downwind

direction (most important for tower clearance).The edgewise bending moment (MY) in the plane of

rotation stems from the tangential forces.

2.1.2 The structure of an airfoil

An example of a typical structural lay-out of an airfoil for a wind turbine blade is shown in Figure 2-1,

the main parts are the spar caps, the shear webs and the outer skins. The spar caps are the most

critical for preventing the blade from colliding with the tower as they carry the flap wise bending

loads. The spar caps are relatively thick laminates, up to 10 centimetres at inboard sections, with

primarily unidirectional fibres to bear the flap wise bending moment. The shear webs mainly pass on

shear loads and consist of much thinner laminates than the spar caps. The blade skins are particular

good for buckling resistance and have an outer layer which is weather resistant. They typically consist

of double-bias or tri-axial fibreglass layers with a balsa or foam core. The internal and external

structure of a wind turbine blade thus has several purposes, the most important structural reasons

are:

Providing stiffness (EI), especially against bending to prevent collision of the blade.

Providing strength for fatigue loading and extreme load cases.

Preventing buckling of the outer skins.

4 Theory overview


Figure 2-1: Common structural architecture for wind turbine blade cross section [21]

2.1.3 The simple beam theory

As mentioned above, stiffness is required to prevent collision of the blade with the tower of a wind

turbine. Modelling the blade as a beam, where the spar caps are the thick upper and lower flanges,

simplifies the structural calculations for our problem. Furthermore neglecting the edgewise moment

(MY) and the normal force N in comparison with the much larger flap wise bending moment is

acceptable. The stress from the bending moments MX and MY and the normal force N in the cross

section of an airfoil about the two principal axes is found with equation (2.1) [8].

, X Y

X Y

M M Nx y y x

I I A (2.1)

Figure 2-2: Simplified structural model of a cross section [8]

Moment of inertia about the flap wise axis for this simplified beam structure, where the spar caps

represent the thick upper and lower flanges of a beam, is [12]:

Theory overview 5


2 3 212* 2*

12X XI I Ad at a t d

(2.2)

With:

ĪX the moment of inertia about its centroidal axis

a the width of the spar cap

t the thickness of the spar cap (also t=b1-b2)

d the distance from the centreline to the middle of a spar cap

Taken into account that t << d and thus neglecting the term with t3, equation (2.2) simplifies to:

22*XI tad (2.3)

From the equation above it becomes clear that the stiffness (EI) of a blade modelled as a simple

beam is affected by the width (a) and thickness (t) of the spar caps, but most of all by the distance

from the centreline or the thickness of the blade. For the simplified beam in Figure 2-2, the edgewise

moment can be neglected with respect to the bending moment, adding that the stress due to the

normal force is small compared to the stress caused by the bending moments, equation (2.1) reduces

to:

, X

X

Mx y y

I (1.4)

It can now easily be seen that the highest stresses occur when y is at its largest, thus at y=d, this gives

together with equation (2.3) for the stress:

2,

2 2

X XM Mx y d

tad tad (2.5)

As an example we take two profiles with the same material lay-up, with one airfoil 20% thicker (dnew

= 1.2*dold) than the other airfoil. Furthermore the bending moment is assumed to be fixed for both

cases. Now we can compare both designs with the restriction that the stresses remain equal.

6 Theory overview


2 2

10.83

1.2 1.2

X X

old new

new old old

old new old

M M

tad tad

t d d

t d d

(2.6)

This simple example already shows the possible reduction in material, the spar cap thickness is

reduced to 83% of the original value by merely increasing the thickness of the airfoil.

2.2 Airfoil design

The reduction of the total cost of energy (COE) is the driving force behind the design of a modern

wind turbine. The design of airfoils not only focuses on maximizing the energy yield, an efficient

structural design is just as important. The objective to maximize energy yield means that from an

aerodynamic point of view airfoils producing the highest power coefficient are the best. This can only

be achieved with airfoils which have a high lift-to-drag ratio. The accompanying lift value at this lift-to

drag ratio, the design lift coefficient, determines the optimum chord length of the blade [16]. A

higher design lift will lead to slightly smaller blade chords for an optimal designed blade with the

same tip speed ratio.

For an efficient structural design the produced loads are the key parameter, the maximum lift

coefficient and the possible high dynamic overshoot are then to be considered. To select an airfoil

which meets these structural requirements the maximum static operating loads can be used. This is

the product of the maximum lift coefficient and its optimum chord (c* CL,max). For a modern pitch

controlled rotor, like the three bladed reference wind turbine, stall behaviour is not one of the

driving design issues anymore. Still, a gradual or smooth stall around and after the maximum lift

coefficient is advisable, this prevents high load fluctuations around stall. The optimum blade design is

thus a combination between aerodynamic and structural requirements:

A high lift-to-drag ratio to maximize energy yield.

A limited maximum lift coefficient to reduce aerodynamic loading on the blade.

These requirements for modern wind turbine airfoils vary along the span of the blade, toward the

root structural demands are higher than at the mid-span or tip locations, resulting in airfoils with

more than 25% thickness for additional stiffness. Airfoils can increase up to 40% relative thickness

inboards, at inboard locations also rotational effects play an important role. Thick airfoils provide

more structural stiffness and reduces weight, leading to a reduction of fatigue loads and costs.

However, the thickness causes increased pressure gradients over the aft part of the airfoil upper

surface, which in particular with leading edge contamination may lead to early turbulent separation

and reduction of the maximum lift coefficient [18].

Theory overview 7


Therefore, for example the DU-series airfoils have smaller upper surfaces and thicker lower surfaces

with ‘aft-loading’ to compensate for the reduction in lift due to the restricted upper surface

thickness. This is where boundary layer suction becomes beneficial, increasing the upper surface

without the loss of aerodynamic performance. The rotational effects may reduce roughness

sensitivity at the inboard sections, still contaminations of the blade nose cannot be avoided. The

degree of this soiling differs strongly, tests in the wind tunnel are not always similar and soiling is also

hard to predict and simulate.

2.3 Boundary layer control

2.3.1 Boundary layer suction

With boundary layer suction a small amount of air flowing around an airfoil is sucked through a

porous part of the skin. The driving force behind the airflow through this porous material is the

pressure difference between the outside and the inside, this principal is described in section 3.1.2.

Due to the adverse pressure gradient over the rear part of the upper surface the boundary layer

becomes thicker and more unstable, causing flow reversal and separation. By removing the air near

the wall the properties of the boundary layer close to the wall change, see Figure 2-3. Air from higher

in the boundary layer flows closer toward the surface of the airfoil, practically replacing the air that

has been sucked through the airfoil wall. This effect results in a higher average energy of the

boundary layer since high energy air replaces the low energy air that is removed by boundary layer

suction. This re-energizing of the boundary layer has a stabilizing effect, which postpones separation

and therefore increases lift en decreases drag [11].

Figure 2-3: Development of the boundary layer affected by boundary layer suction

2.3.2 Moving thickness backwards

Placing the maximum thickness of the airfoil far backwards and with that obtaining long runs of

laminar flow (and thus decreasing drag) was first seen at the well known NACA 6-series airfoils. This

was the first attempt to reduce skin-friction drag through laminarisation of the boundary layer. This

8 Theory overview


principle has led to very thick airfoils where the flow remained laminar all the way to the trailing

edge by proper shaping and adding a suction slot. For a long run of laminar flow at high Reynolds

numbers through favourable pressure gradients, airfoils will have relatively great maximum thickness

located far backwards. Thus increasing the danger for separation of the turbulent boundary layer

over the rear part of the airfoil due to the larger adverse pressure gradient. It is suggested by van

Ingen et al. [9] that laminarisation through shaping should be combined with passive or active

control of the boundary layer. The same article shows that laminarisation by shaping together with

turbulent boundary layer control, is an attractive scheme for performance improvement, for example

lower drag. Important properties for such laminar, or low drag, airfoils are:

Low minimal drag for a certain lift range (the ‘drag bucket’) due to long runs of attached flow

A rearward location of the pressure minimum which decreases the drag

Good stall characteristics of the ‘trailing edge stall’ type, separation begins at the trailing

edge and slowly moves to the leading edge

The pressure rise after the pressure minimum is strongly affected by the position of this pressure

minimum. A rearward pressure minimum will lead to an increased pressure rise behind this minimum

and thus an unfavourable pressure gradient. A large pressure rise will lead to separation of the

turbulent boundary layer and an increase in drag. Although suction could prevent this separation, a

compromise will have to be found. Suction of the boundary layer after this pressure minimum has

also practical advantages. There is no interference with the structural integrity of the airfoil, suction

takes place behind the load bearing box spar section with shear webs [7].

Design Process 9


3 Design Process

The complete design process of the new thicker airfoil is discussed in this chapter, first the starting

point is given in section 3.1. The developed tool to design and select possible airfoils is explained in

sections 3.3 to 3.5. Results from this design tool are discussed in section 3.6 and fine-tuning of the

airfoil is shown in the last two sections.

3.1 Design choices

3.1.1 Reference wind turbine

The first step in the design process is preparing a proper foundation for upcoming computations.

Hence we have to choose a baseline wind turbine which acts as a reference during the complete

process. The chosen turbine, commonly used in the wind energy industry, is based on the data of the

European UpWind reference wind turbine description [20]. This fictitious 5MW horizontal axis wind

turbine (HAWT) is originally based on models of a LMH 64.5 meter blade, which have been used with

various small modifications in projects such as DOWEC, MANGROVE, and the National Renewable

Energy Laboratory (NREL) baseline turbine. The general description of the wind turbine is given in

Table 3-1, although this wind turbine is not in production, it serves as a very useful baseline for this

study.

Table 3-1: Characteristics of the 5 MW UPWIND model

Wind regime IEC Cass 1A / Class 6 winds

Rotor orientation Clockwise rotation - Upwind

Control Variable speed - collective pitch

Cut in wind speed 4 m/s

Cut out wind speed 25 m/s

Rated power 5 MW

Number of blades 3

Rotor diameter 126 m

Hub diameter 3 m

Hub height 90 m

Rated rotor speed 12.1 rpm

Maximum tip speed 80 m/s

From this reference model the 25% thickness DU-91-W2-250 shown in Figure 3-1 is picked as the

baseline airfoil as input for the design process. At this mid-span location the aerodynamic properties

are less crucial than more outwards of the blade, and hence a small change or reduction in

aerodynamics will not critically effect the performance of the wind turbine. However the

10 Design Process


aerodynamic properties at this location are more dominating than at the inboard section, resulting in

a trade-off between aerodynamic and structural properties. This validates the application of

boundary layer suction for the mid-span location, increasing stiffness and still retaining equivalent

aerodynamic quality.

Figure 3-1: DU-91-W2-250 profile normalized to its chord c

Table 3-2: Aerodynamic properties reference wind turbine [20]

Element nr. Rotor radius

Twist Chord Pitch axis aft LE

Coord. Pitch axis

Thickness Airfoil

[m] [deg] [m] [*chord] [m] [%]

1 2.87 13.31 3.54 0.50 0.00 100.00 Cylinder 1

2 5.60 13.31 3.85 0.50 0.00 90.00 Cylinder 1

3 8.33 13.31 4.17 0.43 0.00 70.00 Cylinder 2

4 11.75 13.31 4.56 0.38 0.00 40.00 DU-00-W-401

5 15.85 11.48 4.65 0.38 0.00 35.00 DU-00-W-350

6 19.95 10.16 4.46 0.38 0.00 35.00 DU-00-W-350

7 24.05 9.01 4.25 0.38 0.00 30.00 DU-97-W-300

8 28.15 7.79 4.01 0.38 0.00 25.00 DU-91-W2-250

9 32.25 6.54 3.75 0.38 0.00 25.00 DU-91-W2-250

10 36.35 5.36 3.50 0.38 0.00 21.00 DU-93-W-210

11 40.45 4.19 3.26 0.38 0.00 21.00 DU-93-W-210

12 44.55 3.13 3.01 0.38 0.00 18.00 NACA-64618

13 48.65 2.32 2.76 0.38 0.00 18.00 NACA-64618

14 52.75 1.53 2.52 0.38 0.00 18.00 NACA-64618

15 56.17 0.86 2.31 0.38 0.00 18.00 NACA-64618

16 58.90 0.37 2.09 0.38 0.00 18.00 NACA-64618

17 61.63 0.11 1.42 0.38 0.00 18.00 NACA-64618

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1

-0.25

-0.2

-0.15

-0.1

-0.05

0

0.05

0.1

0.15

0.2

0.25

x/c

y/c

DU-91-W2-250

Thickness: t/c = 0.25 @ x/c = 0.325

Chord line

Lower surface

Trailing Edge (TE)

Upper surface

Leading Edge (LE)

Design Process 11


Another motive for choosing this airfoil is that suction is not possible along the entire length of the

blade. The pressure difference needed for suction is not sufficient anymore from 60-70% of the blade

span, more on this in section 3.1.2. Hence the most critical location where boundary layer suction is

feasible, is around mid-span. The accompanied DU-91-W2-250 profile, halfway the blade length, is

selected from the blade lay-out in Table 3-2. Common Reynolds numbers for such mid-span airfoils in

the higher mega Watt range are around 8 million, thus for the calculations yields Re = 8e6.

3.1.2 Suction distribution

Boundary layer suction is possible due to the pressure difference between the outside and the inside

of the porous material on the airfoil. The pressure difference needed for suction can be obtained by

the natural effect of the centrifugal force, acting on the air inside the blade. Creating an opening at

the tip and leaving the root closed creates an under pressure in the blade. The losses due to the

Coriolis force are minimal and can therefore be neglected [11]. To compute the pressure on the

inside, the blade is modelled as a closed rotating duct. At the position x/c = 0.7 along the chord of the

airfoil the pressure coefficient on the outside has an estimated average value of CP ≈ -0.7. With

known data of the reference turbine and the formulas in appendix H [1], the inside and outside

pressure can be plotted, see Figure 3-2.

Figure 3-2: Pressure along the length of the blade [1]

Figure 3-2 shows that suction is not possible along the whole blade, inside the blade a greater

negative pressure is needed to let the air flow through the porous material. Around R = 46m the

inside pressure is less negative than the outside pressure and air is blown out if the porous material

still is applied. Furthermore to overcome the time variant pressure fluctuations on the outside there

should be at least a 500Pa pressure difference at all times. Because of this safety margin, blowing out

air instead of suction due to pressure fluctuations will not be possible. The amount of sucked air can

be regulated by applying different porous materials and thicknesses. It has to be noted here that the

amount of suction applied is in reality not constant over the sucked area, this is because the outside

pressure will vary along the blade as well as along the suction area in chord wise direction.

12 Design Process


As mentioned before, suction is not feasible along the entire length of the blade, the same applies for

the chord wise position of the airfoil. To keep the structural integrity of the blade intact, only the last

30% of the airfoil’s chord can be used for suction of the boundary layer. This fact was confirmed after

several conversations with experts from the wind energy industry. Because of the small thickness at

the trailing edge, it is questionable if suction is possible over the last 5% of the chord. Therefore,

during the design process, suction will be applied from x/c = 0.7 to x/c = 0.95. After a new airfoil is

designed, the effect of shortening this suction length will be investigated, see section 5.1.1.

Figure 3-3: Suction area on an example profile

Figure 3-3 illustrates the position and the length of suction area described above. The complete

volume inside the airfoil beneath the suction area is used to accommodate the suction duct. The

shape of the suction distribution along the upper surface is shown in Figure 3-4, along the suction

area the distribution is formed as a block. This base suction (v/U = -0.001) can easily be imported and

scaled to larger suction velocities by RFOIL. RFOIL is a successor of XFOIL and is described in more

detail in section 3.2. The suction velocity plotted in Figure 3-4 is defined as the velocity of the sucked

air perpendicular to the airfoil (v) divided by the relative wind velocity at the blade (U). From results

of earlier studies the maximum suction velocity possible was already determined around v/U = -0.003

[1].

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1

-0.25

-0.2

-0.15

-0.1

-0.05

0

0.05

0.1

0.15

0.2

0.25

x/c

y/c

Suction Area/Length

Suction Duct

Design Process 13


Figure 3-4: Suction velocity (v/U) vs chord wise position (x/c)

3.2 RFOIL-suc

The required polars for the clean and soiled conditions are produced by RFOIL, which originates from

XFOIL. XFOIL is an inviscid linear-vorticity 2-D panel method (Figure 3-5) with a Karman-Tsien

compressibility correction, source distributions superimposed on the airfoil and wake permit

modelling of viscous layer influence on the potential flow. Both laminar and turbulent layers are

treated, with an e9-type amplification formulation determining the transition point [5]. Boundary

layer suction was already available in XFOIL at the Delft University of Technology and has now also

been implemented in RFOIL. RFOIL is a modified version of XFOIL with an improved prediction

around the maximum lift coefficient and capabilities of predicting the effects of rotation on airfoil

characteristics. These rotational effects scale with the local solidity (c/r), which is multiplied with a

factor of 2/3 before used as input for RFOIL. The local solidity was multiplied by this factor to match

the calculated results with the measurements [17]. RFOIL considerations are always a bit too

optimistic, resulting in a lower drag and a slightly higher (maximum) lift coefficient than in reality,

this is caused by underestimation of the boundary layer thickness [19].

Figure 3-5: Airfoil and wake panelling with vorticity () and source distributions (), with TE detail [6]

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1-1

-0.9

-0.8

-0.7

-0.6

-0.5

-0.4

-0.3

-0.2

-0.1

0x 10

-3

x/c

v/U

Basic suction shape

14 Design Process


The advantage of a program like RFOIL for the design process lies in the fact that it can be automated

with help of script files. Script files are useful for command line-input programs which need the same

or almost similar input for many times. The commands for RFOIL of one profile are programmed in a

script file, and with help of looping any number of calculated profiles can now automated. The Visual

Basic Script (VBS) language developed by Microsoft is used for creating the script files, for more

information the reader is referred to the VBS internet site [13].

A disadvantage of the suction module in RFOIL is that the viscous boundary layer formulations

cannot handle a shape factor smaller than one. This happens when too much suction is applied,

especially at low angles of attack the suction causes the shape factor of the boundary layer to fall

below one and RFOIL-suc does not converges. However, in reality this is not possible, it means that

the suction prevents the build up of any more boundary layer and maximum profit of boundary layer

suction is reached.

3.3 Profile field

The base DU-91-W2-250 airfoil is altered in two different ways. First the position of the thickest point

is moved backwards step by step to ensure longer flow attachment, see section 2.3.2. XFOIL moves

the thickest point of the base airfoil (x/c = 0.325) backwards in small steps to the end point at x/c =

0.525, hereby not changing the camber of the airfoil. This end point is chosen as maximum to ensure

the new airfoils are still usable, the S-shape of the original airfoil deforms the shape contours.

Moving the thickest point too far backwards would give useless airfoils with too large surface

gradients at the trailing edge causing flow separation.

As a second step the thickness is gradually increased from 25% to 40% thickness. As explained in

section 2.1.3 the reason for this is clear, increasing the structural efficiency by increasing thickness of

the airfoil blade. The two modifications of the original airfoil produces a field of one hundred airfoils

with different thicknesses and shapes. This profile field is shown in Figure 3-6 where the airfoils are

labelled as “1stnumber-2ndnumber.AIR”. The 1st number represents the position of the thickest point,

with X = 1 being the original thickness point at x/c = 0.325 and X = 10 at x/c = 0.525. The second

number denotes the thickness of the airfoil, with Y = 1 again the original thickness of 25% and Y =10

the maximum thickness of 40%. On the horizontal axis the thickness is moved backward and

vertically the thickness is increased, the four corner points of the created profile field are shown with

the original airfoil in blue:

1-1.air: t/c = 0.25 at x/c = 0.325

1-10.air: t/c = 0.4 at x/c = 0.325

10-1.air: t/c = 0.25 at x/c = 0.525

10-10.air: t/c = 0.4 at x/c = 0.525

Design Process 15


Figure 3-6: Profile field of new created airfoils

3.4 Selection criteria

The goal of the design process is to develop a thicker airfoil with equal or better aerodynamic

properties. The most important parameter for an airfoil is the lift-to-drag ratio, a high lift-to-drag

ratio is needed for a high power coefficient to maximize energy yield. To ensure the same power

output the lift-to-drag ratio of the new airfoil cannot be reduced for the clean as well as the soiled

case. The lift coefficient at which the lift-to-drag ratio is at its maximum (L/Dmax) is the design lift

coefficient (CL,design), which determines then the chord length according to required lift.

The basic assumption is that the new airfoil can easily be inserted in the original model, meaning that

the chord length and thus the design lift cannot differ too much. Furthermore a margin between the

design lift with respect to the maximum lift coefficient (CL, max) is needed, usually a difference of ∆CL =

0.2 is sufficient. Also the maximum lift coefficient ought to be restrained to maximal its original

value, this is for limiting aerodynamic (static) loads on the blade. The above mentioned

characteristics are shown in Figure 3-7 and here again summarized:

Equal or better L/Dmax

Similar CL,design

Equal or lower CL, max

16 Design Process


Figure 3-7: Polar curves for DU-91-W2-250 (Re = 8e6 and c/r = 0.09)

3.5 Flow diagram

The general flowchart of the design process is given in Figure 3-8, the first basics for the start of the

design process were already mentioned. With the original DU-91-W2-250 profile a 10x10 field of

coordinate files is created, for an example profile coordinate file see appendix A.1. For each

coordinate file a matching suction distribution file is needed else RFOIL is not able to apply the

correct suction. RFOIL creates for an airfoil an unique set of points at which the suction velocity can

be defined. If for an airfoil the suction is not defined at those specific points, the applied suction

distribution is different from the intended distribution. To overcome this problem for each airfoil a

base suction distribution is created with the script file given in appendix B.2. The desired suction

velocity at the unique points are then written to the final suction files, thus ensuring the same base

suction distribution for each profile as shown in Figure 3-4. The MATLAB file shown in appendix C.1 is

used for this.

-10 -5 0 5 10 15 20-1

-0.5

0

0.5

1

1.5

2

X: 13.5Y: 1.663

CL - curve

CL

Clean configuration

Soiled configuration

-100 -50 0 50 100 150 200-1

-0.5

0

0.5

1

1.5

2

X: 114.6Y: 1.39

CL

CL/C

D

CL - C

L/C

D

X: 160.5Y: 1.03

CL/C

D-max,clean

CL/C

D-max,soiled

CL,design,clean

CL,design,soiled

CL,dif,soiled

CL,max

CL,dif,clean

Design Process 17


Figure 3-8: General flowchart design process

The required polar files for each airfoil, see appendix A.3, are produced by RFOIL-suc, again this

process is automated by a script file, see appendix B.3. The script file opens RFOIL-suc, loads the

airfoil, the required inputs for RFOIL-suc are given, the script file waits for RFOIL-suc to calculate the

polar and closes the program to start the process all over again for the next airfoil. Within this script

file the user can change the input according to the requirements:

The Reynolds number for the viscous flow is set to Re = 8e6

When soiled condition are modelled, the boundary layer is tripped on the upper surface at

5% from the leading edge, thus x/c = 0.05

If rotational effects are taken into account the local solidity (including a factor 2/3) is c/r =

0.09

The polar is calculated from an angle of attack of α = -10 degrees with steps of a half degree

to α = 20 degrees

To gain insight in all these polar files, the data will be sorted and filtered with help of MATLAB

(appendix C.2). First MATLAB reads in the data from the polar files and stores only the needed

columns of variables (α, CL and CD) up to CL,max. For clarity the polars of the different suction velocities

for each airfoil are combined to one polar with increasing suction. As mentioned before, if too much

18 Design Process


suction is applied RFOIL-suc will not converge. By combining the suction velocities step by step, each

suction velocity will have a complete polar without gaps. This means that the program stores only

one entry, the largest suction velocity, for each angle of attack.

With these filtered results the necessary data of each airfoil and for every suction velocity can be

calculated and stored in matrices. For example one matrix (10 by 10) contains the maximum lift-to-

drag ratio of all 100 airfoils for one suction velocity. Now the selection criteria like lift-to-drag ratio,

maximum lift coefficient and many more, can easily be compared for each case (clean, soiled or

different suction velocity). As an example the data for the maximum lift-to-drag ratio of the whole

profile field of 100 airfoils without suction applied is given in appendix C.3 and also graphically

illustrated below in Figure 3-9. The area under the coloured surface represents the profile field, with

on one axis the increasing thickness (from t/c = 0.25 to t/c = 0.4) and on the other axis the chord wise

position of the thickest point. For this example the surface itself consist of the 100 entries (CL/CD-max)

of the profile field. For the two points where the maximum lift-to-drag ratio is zero, RFOIL-suc could

not resolve the given airfoils.

Figure 3-9: CL/CD-max for v/U = 0 and clean configuration

The end result of this design tool is given in Figure 3-10, an overview of all airfoils which satisfy given

selection criteria, for each suction velocity an individual overview is generated. The right part of

Figure 3-10 gives a numerical listing of the available airfoils, the accompanying thickness and its

position is shown in the left part. The applied selection criteria are always given in the legend.

Design Process 19


Figure 3-10: Candidate airfoils for v/U = 0 (clean)

Usually the selection consist of checking if the maximum lift-to-drag ratio is big enough and whether

there is a sufficient safety margin between the design lift coefficient and the maximum lift

coefficient. As can be seen in Figure 3-10 also an alternative selection criteria can be used instead of

the before mentioned. The lift-to-drag ratio of the new airfoils is checked at the alternative lift

coefficient (CL,alt), this CL,alt is the lift coefficient with at least an 0.2 difference from the maximum lift

coefficient. When the lift-to-drag ratio at the alternative lift coefficient of a new airfoil is bigger than

the original, the airfoil is indicated as possible solution. The definition of this alternative approach

and an example of accompanying polars is plotted in Figure 3-11 and Figure 3-12.

Figure 3-11 shows the considerable larger lift-to-drag ratio at the alternative lift coefficient of the

new airfoil (‘1-5.air’). From Figure 3-12 it can be seen that indeed at the maximum lift-to-drag ratio of

the original airfoil, the new airfoil performs worse. However, for a reasonable alpha range (from α≈5

to α≈10 degrees) the lift-to-drag ratio of the new airfoil is substantially larger. Assuming the same

control strategy for both airfoils, a wind turbine equipped with the new airfoil would perform better

on this alpha range due to the higher lift-to-drag ratio. Important at this point is to take into account

the variation of the angle of attack during normal operation cycles.

The variation of this angle of attack for different wind velocities is given in Table 3-3, with the mean

angle of attack as well as the standard variation around this mean and the extremes. These values

were computed with BLADED for the reference model around mid-span, with normal wind shear and

turbulence conditions. For example at the rated wind speed, Vwind = 11.4 m/s, the average angle of

attack is α ≈ 4.6 degrees. At this angle of attack the new airfoil has a smaller lift-to-drag ratio than the

original airfoil, therefore the performance would be less and selecting this airfoil would be critical. It

is therefore important to look for which alpha range the new airfoils perform better, an overall larger

lift-to-drag ratio would of course be the ideal case.

0.325 0.375 0.425 0.475 0.525

0.25

0.275

0.3

0.325

0.35

0.375

0.4

Possible profiles (X-Y.air) with maximum thickness (t/cmax

) and position (x/c) for v/U = 0.000

x/c

t/c

1-1.air: t/c=0.25 @ x/c=0.325

1-10.air: t/c=0.4 @ x/c=0.325

10-1.air: t/c=0.25 @ x/c=0.525

10-10.air: t/c=0.4 @ x/c=0.525

Cl/Cdmax

; Cldif

(stall-design)

Cl/Cdalt

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

0

1

2

3

4

5

6

7

8

9

10

X

Y

20 Design Process


Figure 3-11: Definition of alternative selection criteria

Figure 3-12: Polar examples of 1-5.air (Re = 8e6, c/r = 0.09 and clean)

Table 3-3: Variation of angle of attack with the wind velocity for the reference wind turbine (BLADED)

Vwind [m/s] 8.0 11.4 18.0

Mean α *deg] 3.8 4.6 2.9

Maximum α *deg] 9.3 10.0 10.1

Minimum α*deg] -1.7 -0.8 -5.9

Standard deviation [deg] 1.8 1.9 2.3

0 2 4 6 8 10 12 14 16 18 200.4

0.6

0.8

1

1.2

1.4

1.6

1.8

2

CL - curve

CL

DU-91-W2-250

1-5.air

0 20 40 60 80 100 120 140 160 180 2000.4

0.6

0.8

1

1.2

1.4

1.6

1.8

2

X: 188Y: 1.487

CL - C

L/C

D

CL

CL/C

D

X: 187.9Y: 1.428

X: 161.2Y: 1.03

CL,max

CL,alt

CL 0.2

CL/C

D-max,alt

CL = C

L,dif < 0.2

0 2 4 6 8 10 12 14 16 18 200.4

0.6

0.8

1

1.2

1.4

1.6

1.8

2

CL

0 20 40 60 80 100 120 140 160 180 2000.4

0.6

0.8

1

1.2

1.4

1.6

1.8

2

CL

CL/C

D

0 0.02 0.04 0.06 0.08 0.1 0.12 0.140.4

0.6

0.8

1

1.2

1.4

1.6

1.8

2

CD

CL

DU-91-W2-250

1-5.air

0 2 4 6 8 10 12 14 16 18 200

50

100

150

200

CL/C

D

Design Process 21


3.6 Results from the profile field

Figure 3-18 until Figure 3-25 shown at the end of this section give a good overview of what happens

to the maximum lift-to-drag ratio across the profile field while increasing the suction velocity. The

thickness of an airfoil is found on the vertical axis, the position of the thickest point is represented by

the horizontal axis and the colour bar indicates the magnitude of the lift-to-drag ratio. The end result

of this design tool is given from Figure 3-26 to Figure 3-33, an overview of all airfoils which satisfy

given selection criteria, below follows an explanation of the results.

For the clean configurations shown in Figure 3-18 to Figure 3-21, increasing the thickness has a

positive effect on the maximum lift-to-drag ratio. The design lift coefficient increases with increasing

thickness leading to a higher L/D, but also the drag increases causing the lift-to-drag ratio to

decrease. The increase of maximum lift-to-drag ratio is mainly due to the design lift coefficient, by

increasing the thickness also the curvature of the airfoil is increased, and the lift is directly related to

the airfoils curvature. Moving the thickest point backwards and thereby decreasing the curvature of

the airfoil, as seen from the leading edge, decreases the maximum lift coefficient considerable.

Moreover by moving the thickest point backwards, the adverse pressure gradient after the negative

pressure peak is increased and can cause flow separation. These effects, increasing the thickness and

moving the thickest point backwards, on the design lift are shown in Figure 3-13 below.

Figure 3-13: CL,max and CL,design for clean conditions and v/U = 0

The drag grows with increasing thickness, but by moving the thickest point backwards the flow

remains longer attached to the airfoil and thus reducing the drag. This drag reduction was more

prominent for thinner airfoils in the lower part of the lift-to-drag ratio fields in Figure 3-18 up to

Figure 3-21, explaining the increase of the maximum lift-to-drag ratio in the lower right part of the

profile field. Placing the thickest point further backwards the lift is decreased and thus lowering L/D.

However, the drag is most of the time lower when moving the thickest point backwards, increasing

L/D again. For the more extreme airfoils, large thicknesses and more backwards, RFOIL could not

always produce polars with enough data. Therefore only a few data points were available to calculate

the lift-to-drag ratio, particular with small suction velocities the upper right part of the fields are

questionable.

22 Design Process


Figure 3-22 to Figure 3-25 at the end of this section shows the effect of increasing the thickness for

the soiled case, the lift-to-drag ratio decreases drastically. The increased thickness causes the

pressure gradients to rise over the aft part of the airfoil upper surface. These steep pressure

gradients combined with leading edge contamination may lead to early turbulent separation,

reducing the lift coefficient (see Figure 3-14 below) and thus the lift-to-drag ratio [18]. Moving the

thickest point backwards has the same effect as with the clean configurations. The lift coefficient is

decreased and also the drag is somewhat lower. However, when enough suction is applied this drag

reduction is more apparent. Again this effect is more clear for the thinner airfoils, increasing the

maximum lift-to-drag ratio in the lower right regime of the profile field, see Figure 3-22 to Figure

3-25.

Figure 3-14: CL,max and CL,design for soiled conditions and v/U = 0

The end result of the design tool is shown from Figure 3-26 to Figure 3-33 at the end of this section,

the figures give all the candidate airfoils for each suction velocity individual. For every suction

velocity the tool gives all the available airfoils. As mentioned before the lift-to-drag ratio at the

design lift coefficient should be greater than the original, this selection criteria is denoted as CL/CD,max.

Also a sufficient margin is needed between the design lift, CL,design, and the maximum lift coefficient

CL,max, this margin is shown as CL, diff in the legend with selection criteria. Another selection method

explained in section 3.4 can be used instead, this is mentioned as CL,alt. In the profile fields a

distinction is made between the two selection methods, it immediately becomes clear that with the

alternative selection criteria more airfoils become available.

As expected, for the clean conditions more candidate airfoils are available than for the soiled

conditions. The figures with the maximum lift-to-drag ratio already shown much better performance

for the clean conditions. From Figure 3-26 to Figure 3-29 the possible airfoils for the clean conditions

are shown, even some of the thickest airfoils satisfy the selection criteria, the maximum lift

coefficient was here added as an extra condition. This CL,max in the legend states that the maximum

lift coefficient of a new possible airfoil should be lower than the original airfoil, ensuring equal or

lower static operating loads (c*CL,max) with similar chord length. The clean field plots show that by

increasing the suction velocity more thicker airfoils become available. Because of the extra condition

for the maximum lift coefficient some airfoils disappear from the profile field with increasing suction

velocity. However, profiles found for a lower suction velocity still have a higher lift-to-drag ratio, now

the maximum lift coefficient rises above the original.

Design Process 23


Figure 3-30 illustrates the effect of contamination on the blade for wind turbines and thus the

potential for boundary layer suction for the wind energy industry. The original DU-91-W2-250 airfoil

was designed for low roughness sensitivity leading to a reduced upper surface, also the position of

the thickest point was optimized. Because of this, increasing the thickness and moving the thickest

point will seriously affect the performance of the airfoils for the soiled case, also shown in Figure

3-22 for the lift-to-drag ratio. Therefore, the only acceptable airfoil without suction according to the

selection criteria, is the original airfoil. By increasing the suction more candidate airfoil become

available, illustrated by Figure 3-30 to Figure 3-33. Comparing the clean and soiled cases it is needless

to state that the selection of a new airfoil is more critical for the soiled conditions, most of the airfoils

found for the soiled case will satisfy the same criteria for the clean conditions.

Checking the profile field of the soiled conditions for the maximum suction velocity possible, Figure

3-33, reveals the thickest airfoils possible with use of suction. The three thickest airfoils are now

chosen for further investigation:

The ‘1-4.air’ with a thickness of t/c = 0.300 at its original position of x/c = 0.326,

The ‘2-4.air’ with a thickness of t/c = 0.300 at a more backward position of x/c = 0.346 and

The ‘1-5.air’ with a thickness of t/c =0.316 at its original position of x/c = 0.326.

These selected airfoils are plotted below in Figure 3-15 together with the original DU-92-W2-250

(here ‘1-1.air’). The position of the thickest point of ‘1-5.air’ and ‘1-4.air’ is similar as that of the

original airfoil (x/c = 0.326), as can be seen the purple ‘1-5.air’ is slightly thicker than the red ‘1-4.air’.

The green ‘2-4.air’ and the red ‘1-4.air’ have the same thickness (t/c = 0.300), only the thickest point

for the green ‘2-4.air’ airfoil is more backwards (x/c =0.346). The polars of these airfoils are shown in

Figure 3-16 and Figure 3-17.

Figure 3-15: Result airfoils with thickness and position

0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1

-0.25

-0.2

-0.15

-0.1

-0.05

0

0.05

0.1

0.15

0.2

0.25

x/c

y/c

1-4.air: t/c = 0.300 @ x/c = 0.326

2-4.air: t/c = 0.300 @ x/c = 0.346

1-5.air: t/c = 0.316 @ x/c = 0.326

1-1.air: t/c = 0.250 @ x/c = 0.326

24 Design Process


For the soiled case the performance is quite acceptable, over the operational range of the angle of

attack (see Table 3-3) the lift-to-drag ratio is higher. However, for higher angles of attack the lift-to-

drag ratio of the candidate airfoils drops below the original. Also due to the sharp peak in the lift

curves, the stall is more severe than the build in smooth stall control of the original airfoil. Figure

3-17 also shows the negative effect on the performance due to increasing the thickness (from ‘1-

4.air’ to ‘1-5.air’) and placing the thickest point further backwards (from ‘1-4.air’ to ‘2-4.air’) for the

soiled conditions. For the clean case this is not as apparent, increasing the thickness even improves

the performance of the airfoil (from ‘1-4.air’ to ‘1-5.air’), see Figure 3-16.

Figure 3-16: Polars of the selected airfoils with v/U = 0.003 (clean; Re = 8e6; c/r = 0.09)

Figure 3-17: Polars of the selected airfoils with v/U = 0.003 (soiled; Re = 8e6; c/r = 0.09)

0 2 4 6 8 10 12 14 16 18 200

0.5

1

1.5

2

CL

0 50 100 150 200 2500

0.5

1

1.5

2

CL

CL/C

D

0 0.02 0.04 0.06 0.08 0.1 0.12 0.140

0.5

1

1.5

2

CD

CL

DU-91-W2-250

1-4.air (clean)

2-4.air (clean)

1-5.air (clean)

0 2 4 6 8 10 12 14 16 18 200

50

100

150

200

250

CL/C

D

0 2 4 6 8 10 12 14 16 18 200

0.5

1

1.5

2

CL

0 50 100 150 200 2500

0.5

1

1.5

2

CL

CL/C

D

0 0.02 0.04 0.06 0.08 0.1 0.12 0.140

0.5

1

1.5

2

CD

CL

DU-91-W2-250

1-4.air (soiled)

2-4.air (soiled)

1-5.air (soiled)

0 2 4 6 8 10 12 14 16 18 200

50

100

150

200

250

CL/C

D

Design Process 25


Profile fields with maximum lift-to-drag ratio:

Figure 3-18: CL/CD-max for v/U = 0, clean configuration and Re = 8e6

Figure 3-19: CL/CD-max for v/U = 0.001, clean configuration and Re = 8e6

26 Design Process




Design Process 27


Figure 3-22: CL/CD-max for v/U = 0, soiled conditions and Re = 8e6

Figure 3-23: CL/CD-max for v/U = 0.001, soiled conditions and Re = 8e6

28 Design Process




Design Process 29


Profile fields with candidate airfoils:

Figure 3-26: Candidate airfoils for v/U = 0 (clean)

Figure 3-27: Candidate airfoils for v/U = 0.001 (clean)

0.325 0.375 0.425 0.475 0.525

0.25

0.275

0.3

0.325

0.35

0.375

0.4



x/c

t/c

1-1.air: t/c=0.25 @ x/c=0.325

1-10.air: t/c=0.4 @ x/c=0.325

10-1.air: t/c=0.25 @ x/c=0.525

10-10.air: t/c=0.4 @ x/c=0.525

Cl/Cdmax

; Cldif

(stall-design); Clmax

Cl/Cdalt

; Clmax 0 1 2 3 4 5 6 7 8 9 10

0

1

2

3

4

5

6

7

8

9

10

X

Y

0.325 0.375 0.425 0.475 0.525

0.25

0.275

0.3

0.325

0.35

0.375

0.4



x/c

t/c

1-1.air: t/c=0.25 @ x/c=0.325

1-10.air: t/c=0.4 @ x/c=0.325

10-1.air: t/c=0.25 @ x/c=0.525

10-10.air: t/c=0.4 @ x/c=0.525

Cl/Cdmax

; Cldif


Cl/Cdalt

; Clmax 0 1 2 3 4 5 6 7 8 9 10

0

1

2

3

4

5

6

7

8

9

10

X

Y

30 Design Process




0.325 0.375 0.425 0.475 0.525

0.25

0.275

0.3

0.325

0.35

0.375

0.4Possible profiles (X-Y.air) with maximum thickness (t/c

max) and position (x/c) for v/U = 0.002

x/c

t/c

1-1.air: t/c=0.25 @ x/c=0.325

1-10.air: t/c=0.4 @ x/c=0.325

10-1.air: t/c=0.25 @ x/c=0.525

10-10.air: t/c=0.4 @ x/c=0.525

Cl/Cdmax

; Cldif


Cl/Cdalt

; Clmax 0 1 2 3 4 5 6 7 8 9 10

0

1

2

3

4

5

6

7

8

9

10

X

Y

0.325 0.375 0.425 0.475 0.525

0.25

0.275

0.3

0.325

0.35

0.375

0.4



x/c

t/c

1-1.air: t/c=0.25 @ x/c=0.325

1-10.air: t/c=0.4 @ x/c=0.325

10-1.air: t/c=0.25 @ x/c=0.525

10-10.air: t/c=0.4 @ x/c=0.525

Cl/Cdmax

; Cldif


Cl/Cdalt

; Clmax 0 1 2 3 4 5 6 7 8 9 10

0

1

2

3

4

5

6

7

8

9

10

X

Y

Design Process 31


Figure 3-30: Candidate airfoils for v/U = 0 (soiled)

Figure 3-31: Candidate airfoils for v/U = 0.001 (soiled)

0.325 0.375 0.425 0.475 0.525

0.25

0.275

0.3

0.325

0.35

0.375

0.4



x/c

t/c

1-1.air: t/c=0.25 @ x/c=0.325

1-10.air: t/c=0.4 @ x/c=0.325

10-1.air: t/c=0.25 @ x/c=0.525

10-10.air: t/c=0.4 @ x/c=0.525

Cl/Cdmax

; Cldif

(stall-design)0 1 2 3 4 5 6 7 8 9 10

0

1

2

3

4

5

6

7

8

9

10

X

Y

0.325 0.375 0.425 0.475 0.525

0.25

0.275

0.3

0.325

0.35

0.375

0.4



x/c

t/c

1-1.air: t/c=0.25 @ x/c=0.325

1-10.air: t/c=0.4 @ x/c=0.325

10-1.air: t/c=0.25 @ x/c=0.525

10-10.air: t/c=0.4 @ x/c=0.525

Cl/Cdmax

; Cldif

(stall-design)

Cl/Cdalt

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

0

1

2

3

4

5

6

7

8

9

10

X

Y

32 Design Process




0.325 0.375 0.425 0.475 0.525

0.25

0.275

0.3

0.325

0.35

0.375

0.4



x/c

t/c

1-1.air: t/c=0.25 @ x/c=0.325

1-10.air: t/c=0.4 @ x/c=0.325

10-1.air: t/c=0.25 @ x/c=0.525

10-10.air: t/c=0.4 @ x/c=0.525

Cl/Cdmax

; Cldif

(stall-design)

Cl/Cdalt

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

0

1

2

3

4

5

6

7

8

9

10

X

Y

0.325 0.375 0.425 0.475 0.525

0.25

0.275

0.3

0.325

0.35

0.375

0.4



x/c

t/c

1-1.air: t/c=0.25 @ x/c=0.325

1-10.air: t/c=0.4 @ x/c=0.325

10-1.air: t/c=0.25 @ x/c=0.525

10-10.air: t/c=0.4 @ x/c=0.525

Cl/Cdmax

; Cldif

(stall-design)

Cl/Cdalt

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

0

1

2

3

4

5

6

7

8

9

10

X

Y

Design Process 33


3.7 Modify camber

Up to now the camber of the airfoil remained unchanged, however the camber has a substantial

effect on the performance of the airfoils. There are numerous ways of defining and changing the

camber of an airfoil, for this report the camber line is defined as a maximum camber value at a

position along the chord of the airfoil. The camber of the original DU-91-W2-250 airfoil is then 0.027

at x/c = 0.78 and for each airfoil of the profile field the camber has this same value. In Figure 3-34 an

example camber distribution is plotted by XFOIL, the lower graph shows the actual airfoil and the

upper graph shows the symmetrical version of the airfoil in red and the camber as a purple line.

Figure 3-34: Camber distribution for example profile

The camber is now changed by moving and scaling the camber point to determine the effect of

altering the camber. This procedure is done for all three candidate airfoils from the previous section.

Here the purpose of camber variation is to fine-tune the airfoils so they will perform better. Better

performance can mean a higher lift-to-drag ratio, smoother stall after the maximum lift coefficient,

lowering the maximum lift coefficient to reduce static operation loads and lengthening the alpha

34 Design Process


range where the lift-to-drag ratio is larger than the original. Evaluation and selection of the airfoils is

now done by checking and comparing the individual polars.

The effect of the camber modifications is similar for each of the candidate airfoils, the outcome of

moving the camber point is shown for the ‘1-4.air’ airfoil in Figure 3-35. The original airfoil is of

course without suction and for the new airfoil the suction velocity is v/U = 0.003. For the original

airfoil the camber point lies at x/c = 0.78, moving this point forward lowers the CL,max, the CL-α curve

is slightly effected and the overall lift-to-drag ratio is decreased. Moreover the range of angle of

attack where the lift-to-drag ratio is greater than the original is reduced. Placing the camber point

more backwards has the opposite effect, thus improving performance but with an increase of the

maximum lift coefficient and thus increasing static operational loads.

Figure 3-35: Effect of moving the camber point for v/U = 0.003 (soiled; Re = 8e6; c/r = 0.09)

Figure 3-36 illustrates the effect of scaling the camber of an airfoil, as expected not only the

maximum lift coefficient is increased the whole lift curve moves upwards. Furthermore the range

where the lift-to-drag ratio is improved increases due to scaling of the camber value. However, the

lift coefficient is also greatly affected by moving the CL-α curve, thus scaling the camber value has to

be done with great care.

0 2 4 6 8 10 12 14 16 18 200

0.5

1

1.5

2

CL

0 50 100 150 200 2500

0.5

1

1.5

2

CL

CL/C

D

0 0.02 0.04 0.06 0.08 0.1 0.12 0.140

0.5

1

1.5

2

CD

CL

DU-91-W2-250

1-4.air: x/c = 0.66 scale = 100%

1-4.air: x/c = 0.70 scale = 100%

1-4.air: x/c = 0.74 scale = 100%

1-4.air: x/c = 0.78 scale = 100%

1-4.air: x/c = 0.82 scale = 100%

0 2 4 6 8 10 12 14 16 18 200

50

100

150

200

250

CL/C

D

Design Process 35


Figure 3-36: Effect of up scaling the camber value for v/U = 0.003 (soiled; Re = 8e6; c/r = 0.09)

During the design process of altering the camber of the airfoil, the ‘2-4.air’ airfoil was quickly

discarded and the choice was between the ‘1-4.air’ and the thicker ‘1-5.air’ airfoil. This because the

effect of the camber changes was similar for the three airfoils and there the ‘1-4.air’ performed

better than the ‘2-4.air’ and both have equal thicknesses. As seen in Figure 3-16 and Figure 3-17 the

lift-to-drag ratio in general of the ‘1-4.air’ is higher than that of the ‘2-4.air’ and the range where it

performs better than the original DU-91-W2-250 is also larger. Both are critical selection criteria for

choosing the new airfoil.

After the iterative design process of changing the camber of the candidate airfoils the end result for

the two remaining candidate airfoils is plotted in Figure 3-37 and Figure 3-38. As mentioned before,

altering the camber is done by scaling the camber value and moving its position along the chord.

During this process the above mentioned effects of both modifications were taken into account. As

can be seen from Figure 3-37 for the soiled case, the camber changes were more effective for the ‘1-

5.air’ than for the ‘1-4.air’. The lift-to-drag ratio of the ‘1-5.air’ is improved, just as the range where it

is larger than the original DU-91-W2-250 airfoil. For the ‘1-4.air’ the lift-to-drag ratio is sometimes

even lower than without camber changes, however it is larger than the original for a greater range of

angles of attack, although it is just one degree. For the clean conditions shown in Figure 3-38, the

camber alterations did not produce major effects, only the CL,max of the ‘1-5.air’ is slightly larger.

Important for choosing the final airfoil is the performance within the normal operating range of

angles of attack. These are listed in Table 3-3 with a maximum of around 10 degrees, -5 degrees for a

minimum and an average angle of attack between 3 and 5 degrees with a standard deviation around

2 degrees. The overall performance of the ‘1-4.air’ is better than that of the ‘1-5.air’, for the soiled

conditions the lift-to-drag ratio is larger and the drop-off after the maximum lift coefficient is smaller

thus smoother stall conditions for the ‘1-4.air’. Furthermore the angle of attack range with improved

lift-to-drag ratio is larger for the ‘1-4.air’. The ‘drop-off’ of the lift-to-drag ratio for the ‘1-5.air’ is just

0 2 4 6 8 10 12 14 16 18 200

0.5

1

1.5

2

CL

0 50 100 150 200 2500

0.5

1

1.5

2

CL

CL/C

D

0 0.02 0.04 0.06 0.08 0.1 0.12 0.140

0.5

1

1.5

2

2.5

CD

CL

DU-91-W2-250

1-5: x/c = 0.78 scale = 100%

1-5: x/c = 0.78 scale = 150%

1-5: x/c = 0.78 scale = 200%

0 2 4 6 8 10 12 14 16 18 200

50

100

150

200

250

CL/C

D

36 Design Process


after the maximum normal operational angle of attack of 10 degrees, making it slightly critical. For

the clean conditions, Figure 3-38, the lift-to-drag ratio of the ‘1-5.air’ is somewhat better than the ‘1-

4.air’. However, the maximum lift coefficient of the ‘1-5.air’ is slightly higher and thus increasing

static operational loads. Although the ‘1-5.air’ is somewhat thicker, the ‘1-4.air’is chosen as the final

airfoil because of before mentioned properties. The better performance over the normal range of

angles of attack for the soiled case is one of the decisive factors for taking the ‘1-4.air’ airfoil.

Figure 3-37: Result of camber modification for v/U = 0.003 (soiled; Re = 8e6; c/r = 0.09)

Figure 3-38: Result of camber modification for v/U = 0.003 (clean; Re = 8e6; c/r = 0.09)

-10 -5 0 5 10 15 20-1

-0.5

0

0.5

1

1.5

2

CL

-100 -50 0 50 100 150 200 250-1

-0.5

0

0.5

1

1.5

2

CL

CL/C

D

0 0.02 0.04 0.06 0.08 0.1 0.12 0.14-1

-0.5

0

0.5

1

1.5

2

CD

CL

DU-91-W2-250

1-4.air

1-4.air: x/c = 82 scale = 90%

1-5.air

1-5.air: x/c = 82 scale = 100%

-10 -5 0 5 10 15 20-100

-50

0

50

100

150

200

250

CL/C

D

-10 -5 0 5 10 15 20-1

0

1

2

CL

-100 -50 0 50 100 150 200 250-1

0

1

2

CL

CL/C

D

0 0.02 0.04 0.06 0.08 0.1 0.12 0.14-1

0

1

2

CD

CL

DU-91-W2-250

1-4.air

1-4.air: x/c = 82 scale = 90%

1-5.air

1-5.air: x/c = 82 scale = 100%

-10 -5 0 5 10 15 20-100

0

100

200

CL/C

D

Design Process 37


3.8 Wind tunnel adjustments

The design of a new airfoil is just one part the complete process of boundary layer suction

development for a wind turbine. As a next step the airfoil should be tested in the wind tunnel,

particularly to validate the applied design tools like RFOIL-suc. During the preparation by Actiflow for

these tests and after a close investigation of the airfoil, it became clear that some alterations were

necessary. Straightening the rear part of the upper surface is just a minor change which was applied,

because this part was already nearly straight for the last 40%. The effect of this is thus negligible, also

because suction is applied over the most part of the rear upper surface. The biggest advantage of a

straight rear upper surface is the simplicity for manufacturing. Moreover the suction is employed

over this part of the airfoil making the use of straight sheets possible instead of curved sheets.

Figure 3-39: Airfoil growth during the design process

Another modification to the airfoils is done by rotating the lower surface of the airfoil. Fixing the

nose of the airfoil and rotating the lower part creates a larger trailing edge, much like truncated

airfoils. This larger trailing edge is needed to fit the measuring instruments for the wind tunnel tests

in the rear part of the airfoil. In Figure 3-39 the new airfoils with these modifications are plotted. The

final airfoil has thus a straight rear part of the upper surface and a rotated lower surface. This final

airfoil will from now be labelled as ‘AF-0901’, the Actiflow (AF) airfoil designed in 2009 (09) and the

first one (01) of that year. In Figure 3-40 and Figure 3-41 the final AF-0901 airfoil is compared with

the original DU-91-W2-250 and the end result of the camber modifications from section 3.7. Mainly

due to the rotation of the lower surface the lift-to-drag ratio is slightly reduced, the maximum lift

coefficient is increased, but the range of angles of attack with better lift-to-drag than the original DU-

91-W2-250 is increased.

0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1

-0.25

-0.2

-0.15

-0.1

-0.05

0

0.05

0.1

0.15

0.2

0.25

x/c

y/c

1-4.air: camber (x/c = 0.82 scale = 90%)

1-4.air: camber and straigth upper surface

1-4.air: camber, straigth upper surface and rotated lower surface

DU-91-W2-250

38 Design Process


Figure 3-40: Effect of wind tunnel modifications for v/U = 0.003 (clean; Re = 8e6; c/r = 0.09)

Figure 3-41: Effect of wind tunnel modifications for v/U = 0.003 (soiled; Re = 8e6; c/r = 0.09)

-10 -5 0 5 10 15 20-1

-0.5

0

0.5

1

1.5

2

CL

-100 -50 0 50 100 150 200 250-1

-0.5

0

0.5

1

1.5

2

CL

CL/C

D

0 0.02 0.04 0.06 0.08 0.1 0.12 0.14-1

-0.5

0

0.5

1

1.5

2

CD

CL

DU-91-W2-250

1-4.air: x/c = 0.82 scale = 90%

AF-0901

-10 -5 0 5 10 15 20-100

-50

0

50

100

150

200

250

CL/C

D

-10 -5 0 5 10 15 20-1

-0.5

0

0.5

1

1.5

2

CL

-100 -50 0 50 100 150-1

-0.5

0

0.5

1

1.5

2

CL

CL/C

D

0 0.02 0.04 0.06 0.08 0.1 0.12 0.14-1

-0.5

0

0.5

1

1.5

2

CD

CL

DU-91-W2-250

1-4.air: x/c = 0.82 scale = 90%

AF-0901

-10 -5 0 5 10 15 20-100

-50

0

50

100

150

200

250

CL/C

D

Structural benefits 39


4 Structural benefits

For the evaluation of the structural benefits of the newly designed airfoil, a structural 2-D

comparison is made between the ‘old’ DU-91-W2-250 and the ‘new’ AF-0901 airfoil. First a simple

beam theory with accurate XFOIL thin walled shell stiffness data is given to compare both airfoils. As

a second comparison, a detailed structural design from the reference wind turbine is implemented in

PreComp. The new thicker AF-0901 airfoil will then be optimized to match the original stiffness

requirements. In section 2.1.3 the airfoil is modelled as a simple beam and in section 4.1 as a thin

walled structure consisting of one material for comparison, thus without webs and spar caps.

However, in reality various materials with different properties are used for every sandwich section of

the layup. Therefore, in section 4.2 another more in depth structural analysis is made.

4.1 Thin walled shell structure

In section 2.1.3 the simple beam theory was used to show the advantage of increasing the thickness

of the blade. The airfoil was modelled as a simple beam, the flanges of the beam then represent the

spar caps, anything else was dismissed. The airfoil is now modelled as a thin walled structure, with a

material of thickness t along its contour line. When an airfoil is seen as a thin walled structure, XFOIL

can be used for a structural analysis. For such a thin walled profile XFOIL gives the moment if inertia

divided by this thickness (skin Ixx/t), see Figure 4-1.

Figure 4-1: Bending parameters from XFOIL



With the same reasoning as before equation (1.4) now becomes:

,/ /

X X

X Xold new

M Mx y y y

I t t I t t

(4.1)

Assuming again an equal bending moment and rearranging for the thickness ratio gives:

/

/

Xnew old new

old X oldnew

I tt y

t I t y (4.2)

As said before the value of the moment of inertia divided by the thickness are provided by XFOIL.

Also the maximum value for y can be delivered by XFOIL (max Y-Yc), at this value of y the stresses are

at their maximum. With the data from XFOIL the thickness ratio now becomes:

2

2

1.504 10 0.14790.8286 ~ 0.83

2.208 10 0.1216

new

old

t

t

(4.3)

Next the influence of the edgewise moment on the thickness ratio is investigated. Again from XFOIL

the exact data is extracted for the moments of inertia and distances. Still neglecting the normal force

from equation (2.1), the relation for the thickness ratio is given in equation (4.4) and plotted in Figure

4-2.

/ /

/ /

X Y

X Ynew new

old X Y

X Y old

M My x

I t t I t tt

t M My x

I t t I t t

(4.4)

The exact bending moments are not know, thus the thickness ratio is plotted against the ratio

between the edgewise and flap wise bending moment. As can be seen when the edgewise moment is

neglected (My = 0 and thus the moment ratio equals zero) the thickness ratio is around 0.83.



Furthermore the thickness ratio decreases some when the edgewise moment is taken into account.

However when the edgewise moment is small compared to the flap wise moment (My/ Mx around

zero) this is a marginal difference and thus validating the assumption of neglecting the edgewise

moment. For the 5MW reference wind turbine at conditions below rated the flap wise bending

moment is an order of magnitude larger than the edgewise bending moment, and above rated 3 to 5

times bigger than the edgewise bending moment [10].

Figure 4-2: Thickness ratio vs. bending moments ratio

4.2 The PreComp method

PreComp [15] (Pre-processor for computing Composite blade structural properties) was developed

by the NREL and computes the stiffness and inertial properties of a composite blade. These

properties are almost always needed by designers for fast evaluation of different composite layouts

or as inputs for aeroelastic codes (FAST, ADAMS, BLADED) used in the wind energy industry. A

complete blade is processed by PreComp with a novel approach that integrates a modified classic

laminate theory with a shear-flow approach [2]. The advantage of PreComp over other structural

computing tools lies in its efficiency, there it is not based on the time consuming finite element

method [4].

Because PreComp is not based on finite element methods it cannot give a detailed load-displacement

or load-stress distribution. However, by simplification PreComp directly calculates the structural

properties and usually within a second. As inputs PreComp requires the description of the external

blade (chord, twist, airfoil shape, webs, etc) and the internal layup of the composite laminates

(laminate schedule, orientation, properties, etc). Common underlying assumptions are listed below:

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 10.74

0.75

0.76

0.77

0.78

0.79

0.8

0.81

0.82

0.83

My/M

x

t ne

w/t

old



Each blade is a thin-walled, closed, multi cellar section (constant shear flow around cells).

No hoop stresses in any wall of a section.

Straight blade and webs must be normal to the chord.

Each blade section is free to warp out of its plane.

Transverse shearing is negligible and the blade section is rigid in its own blade.

4.2.1 Assumptions

The reference station for the profile used for comparison is chosen at a radius station of R = 29.0

meters, this station falls within the range where the DU-91-W2-250 airfoil is used, see Table 3-2 [20].

Figure E-1, taken from [14], shows the stations available for comparison, structural data as materials

and different lay-ups are summarized in Table D-1 and Table D-2. Detailed aerodynamic and

structural properties from the reference wind turbine are listed below in Table 4-1 and Table 4-2.

Table 4-1: Aerodynamic properties (detailed: N=131) [20]

Element nr. Rotor radius

Twist Chord Pitch axis aft LE

Coord. Pitch axis

Thickness Airfoil

[m] [deg] [m] [*chord] [m] [%]

26 28.99 7.53 3.95 0.38 0.00 25.51 DU-91-W2-250

Table 4-2: Structural properties (detailed) [20]

Element nr.

Rotor radius

Flap bending stiffness (EI)

Flap bending stiffness *1.2 (corrected)

Lead-lag bending stiffness (EI)

Torsion stiffness (GJ)

[m] [Nm2] [Nm2] [Nm2] [Nm2]

24 29.70 8.76E+08 1.05E+09 3.14E+09 1.20E+08

Figure E-2 shows the total lay-up of the blade, the station for comparison here is K, thus with R =29.0

meters. For the complete blade, the lay-up starts with a TRIAX-1 layer at the outside and ends with a

TRAIX-3 layer on the inside. Dependent on the chord wise position on the airfoil, the layer between is

of the type SKINFOAM, UD (spar caps) or UD_TE more towards the trailing edge. Again in Table D-1

and Table D-2 the characteristics of these layers can be found; materials used, their properties,

number of layers and thicknesses. With linear interpolation between the two known points (R =

14.0m and R = 47.0m) the number of layers for UD_TE at station K (R = 29.0m) used, is around 16

layers. An example for the structure of a profile is shown in Figure E-3. The final lay-out used for

comparison is shown in Figure 4-3, with the following assumptions taken into account:

Upper and lower surface lay-up are similar.

Profile is divided into five sectors along the chord.

The chord length is 3.95 meters.



The pitch axis lies in the middle of the spar caps at x/c = 0.38.

The width of the spar caps (UD layer) is 0.75/3.95 = 0.19 ≈0 .20 (is fraction of chord), so the

UD layer starts at x/c = 0.28 and ends at x/c = 0.48.

The width of the TE_UD layer is 0.60/3.93 = 0.15 (is fraction of chord), starts at x/c = 0.80 and

ends at x/c = 0.95.

The shear webs consists of 3 layers (2*R4545, 50mm WEBPS, 2*R4545) as in Figure E-3 and

the thickness of the layer R4545 is t = 0.94mm per layer.

The trailing edge sandwich panel of sector 5 is modelled without the skin foam layer, this

foam layer would also not fit in reality due to the small thickness at the trailing edge.

The blunt trailing edge is modelled in PreComp as a third web with the same structure as the

trailing edge sandwich panel in sector 5 (6*UD45R).

Figure 4-3: Profile lay-out used for comparison

4.2.2 PreComp results

First all available data for the original airfoil DU-91-W2-250, with a thickness of 25 percent, is used

for PreComp. The main input file for PreComp can be found in F.1, this file calls the files with data of

the airfoil (F.2), the internal structure (F.3) and materials used (F.4). With this PreComp computes the

stiffness and inertial properties of the airfoil, see G.1. It should be noted here that at least two

airfoils should be given as input, using two identical airfoils without any twist applied PreComp gives

the same results for both airfoils. Also the length of the blade is of no importance for this case and

thus giving a 2-D comparison between different airfoils.



With PreComp only the most important parameters are used for comparison here, these are the flap

wise, edgewise (lead-lag) and torsion stiffness. The first results of the original airfoil used with data

known from the reference wind turbine (Table 4-2) are collected in Table 4-3. The computed values

from PreComp have the same order as the data from the original turbine and are even a good

approximation. To compare the new airfoil design an identical internal structure and the same

materials are used. From Table 4-3 it is clear that the new airfoil (AF-0901) is much more stiffer than

its old airfoil (DU-91-W2-250). However the new airfoil is also a bit heavier, here the mass density is

the section mass per unit length. As stated before the chord is equal for each airfoil and so with the

mass number, the original DU-91-W2-250 airfoil is taken as reference, the airfoils can easily be

compared.

Table 4-3: Collected results PreComp

Profile Mass density

Mass number

Flap bending stiffness

Lead-lag bending stiffness

Torsion stiffness

[kg/m] [%] [Nm2] [Nm2] [Nm2]

REFERENCE 0.1050E+10 0.3140E+10 0.1200E+09

DU-91-W2-250 0.4089E+03 100.0 0.1028E+10 0.3299E+10 0.1093E+09

AF-0901 0.4175E+03 102.1 0.1541E+10 0.3511E+10 0.1587E+09

AF-0901 -20% 0.2764E+03 67.6 0.1251E+10 0.2341E+10 0.1139E+09

AF-0901 -20% (webs: 4 plies)

0.3128E+03 76.5 0.1253E+10 0.3709E+10 0.1139E+09

AF-0901 -25% 0.2642E+03 64.6 0.1190E+10 0.2253E+10 0.1136E+09

AF-0901 -25% (webs: 4 plies)

0.3005E+03 73.5 0.1192E+10 0.3624E+10 0.1136E+09

AF-0901 -25%UD (only spar caps )

0.3573E+03 87.4 0.1224E+10 0.3242E+10 0.1594E+09

AF-0901 -35%UD (only spar caps )

0.3321E+03 81.2 0.1084E+10 0.3105E+10 0.1594E+09

As a next step for optimization the material used is decreased for the new airfoil, because the aim is

to have the same structural properties (equal stiffness). The number of plies used (see Internal

Structure Data File: “int-du25.inp”) for each laminate is now decreased with 20 percent. It has to be

noted that the number of plies should be an integer and that it is actual not possible to have for

example 1.6 ply. However a first good approximation of the structural benefits now becomes

possible. The AF-0901 version minus 20 percent per ply is also shown in Table 4-3, the airfoil is now

lighter, flap and torsion stiffness is acceptable, only the lead-lag bending stiffness is decreased

considerable.

To increase the bending stiffness to an appropriate level, one method is to add stiffness to the shear

webs. With the same reasoning as for the spar caps, which mostly effect the flap wise bending

stiffness, the shear webs contribute to the lead-lag bending stiffness. The original shear webs have 2

plies of R4545 at each side of the WEBPS, this is increases to 4 plies at each side to improve the lead-

lag bending stiffness. The lead-lag bending stiffness is because of this increased, but also the mass



number is now slightly higher (AF-0901 -20% with 4 plies, Table 4-3). This sequence is repeated for

the AF-0901 with a reduction of 25 percent for the number of all plies and layers. Table 4-3 (AF-0901

-25% with 4 plies) shows that the mass number is decreased to 73.5, meaning 26.5 percent less mass

for the new designed airfoil.

Reducing only the materials of the spar caps should resemble the results from the thin walled

structure more. Table 4-3 also gives these results from PreComp, when 35% of the spar caps material

is left out the mass number is reduced to 81.2% of its original value. So by only reducing the material

of the spar caps, the weight is reduced by 18.8%. As a final note it should be mentioned that the

effect of the material reduction on the buckling resistance is not taken into account. By removing

material from the spar caps or other sections the stiffness is lowered. However, when too much

material is removed, buckling of the spar caps or airfoil skin could occur.

Design issues 47


5 Design issues

Now the final design of the AF-0901 airfoil is known, in this chapter important issues to take into

account while using this airfoil for a new blade are discussed. The first part focuses on different

suction distributions, then in section 5.2 the effect of gusts or variation of the angle of attack is

mentioned. In section 5.3 another comparison with the original DU-91-W2-250 is given to discuss the

upcoming wind tunnel test. In the last section of this chapter, section 5.3, a comparison is made with

the DU-97-W-300 airfoil which has a similar thickness as the AF-090.

5.1 Modify suction distribution

5.1.1 Variation of suction length

Figure 5-1 shows the influence of the length of the suction distribution, changing the length effects

the sucked volume of air and thus the performance of the airfoil. To keep the structural integrity of

the airfoil intact, only over the last 30 percent of the chord suction is applied. Furthermore, the last 5

percent is not used at all because of the limited volume for that part of the airfoil. The original

applied sucked area along the upper surface is from x/c = 0.70 to x/c = 0.95, thus the blue line in

Figure 5-1. During the design process the maximum possible suction area along the upper surface is

used.

This length is shortened in two ways, first suction is applied from the original starting point of x/c =

0.70 up to x/c = 0.85. As expected this results in an overall lower lift-to-drag ratio and stall

postponement is less than with the original length. By shortening the length of the sucked area the

volume of sucked air is reduced and thus the effects of boundary layer suction are reduced. The red

line in Figure 5-1, the second variation of suction area is now from x/c = 0.80 to x/c = 0.95, also shows

these reduced effects due to shortening of the suction length.

In addition this example shows that it is favourable to apply suction as soon as possible. In other

words if it necessary to shorten the length of the sucked area it is better to apply suction closer to

the nose. For both cases (x/c = 0.70 to x/c = 0.85 and x/c = 0.80 to x/c = 0.95) the length of the

sucked are is equal, also the suction distribution is similar and thus the sucked volume of air is the

same. Keeping this in mind Figure 5-1 than shows that with suction applied closer to the nose (x/c =

0.70 to 0.85) a higher lift-to-drag ratio is obtained and stall is longer postponed. If flow separation

occurs before the starting point suction has less effect, thus moving the suction starting point

forwards reduces this drawback.

48 Design issues


Figure 5-1: Polars of the AF-0901 airfoil for different suction lengths (soiled case; Re = 8e6; v/U = 0.003; c/r = 0.09)

5.1.2 Variation of suction velocity

The polars for different suction velocities for the clean and soiled conditions are plotted in Figure 5-2

and Figure 5-3. As mentioned before, for the clean case the performance of the new AF-0901 airfoil

is equivalent to that of the original DU-91-W2-250 airfoil, even without suction (see the red line in

Figure 5-2). Performance is here specified in terms of lift-to-drag ratio, smooth stall and maximum

and design lift coefficient. For the new AF-0901 up to an angle of attack of 5 degrees the lift-to-drag

ratio lags behind that of the original airfoil and applying suction improves the performance for this

range by increasing this lift-to-drag ratio, see Figure 5-2. Increasing the suction velocity even more

does not have a great effect for the lift-to-drag ratio for the same alpha range, it does increases the

lift-to-drag ratio up to α = 13 degrees.

By increasing the suction velocity the maximum lift coefficient is also increased and rises above the

original, thereby increasing static operating loads. Furthermore the smooth stall behaviour of the

DU-91-W2-250 airfoil is in less extent present, the AF-0901 has a greater dip after the maximum lift

coefficient. This does not necessary has to be a negative property, indeed a higher maximum lift

coefficient increases the static loads. However, the decrease of the lift coefficient after its maximum

for the new AF-0901 lowers the loads on the blade again, whereas the lift coefficient of the original

airfoil remains nearly constant after its maximum due to its smooth stall. Besides this, the AF-0901

airfoil has a much higher lift-to-drag ratio in the region before the maximum lift coefficient leading to

a higher energy yield. Summarizing for the clean conditions, the AF-0901 with boundary layer suction

(v/U = 0.003) has up to α = 5 degrees a similar lift-to-drag ratio and maximum lift coefficient, after

this angle of attack the lift-to-drag ratio improves. The lift coefficient rises above the original and

after its maximum it drops again below that of the original.

-10 -5 0 5 10 15 20-1

-0.5

0

0.5

1

1.5

2

CL

-100 -50 0 50 100 150-1

-0.5

0

0.5

1

1.5

2

CL

CL/C

D

0 0.02 0.04 0.06 0.08 0.1 0.12 0.14-1

-0.5

0

0.5

1

1.5

2

CD

CL

AF-0901 (x/c = 0.70 to 0.95)

AF-0901 (x/c = 0.70 to 0.85)

AF-0901 (x/c = 0.80 to 0.95)

-10 -5 0 5 10 15 20-100

-50

0

50

100

150

CL/C

D

Design issues 49


For the soiled case shown in Figure 5-3 the suction velocity is more critical. The range of alpha where

the lift-to-drag ratio is greater than the original is reduced when the suction velocity is lowered from

0.3 to 0.2 percent. The maximum lift coefficient is also more effected, the drop of its maximum (for

0.3 to 0.2 percent) is substantially larger than for the clean case. Lowering the original suction

velocity of v/U = 0.003 now becomes interesting, a lower maximum lift coefficient will lower the

static operational loads. On the other hand reducing the suction velocity to v/U = -0.002 creates a

critical region with a lower lift-to-drag ratio. Important here is keeping in mind the normal operating

alpha range of the wind turbine as given in Table 3-3. Choosing the suction velocity will then be a

trade-off between a greater range of superior lift-to-drag ratio for extra power production and a

lower maximum lift coefficient to reduce loads. A small reduction in maximum power coefficient is

seen as less important than the increase in expected maximum operational loads [16].

In case of extreme conditions the maximum lift coefficient can be reduced even more, this can be

done by shutting off the boundary layer suction system. For example by closing a valve at the tip the

passive system is turned off, air cannot flow from inboard towards the tip section and therefore no

under pressure is possible. The polars of the new AF-0901 airfoil without suction is shown by the red

lines in Figure 5-2 and Figure 5-3, notice the large difference in maximum lift coefficient in

comparison to the green lines (v/U = -0.003).

Figure 5-2: Effect of suction velocity (clean case; Re = 8e6; c/r = 0.09)

-10 -5 0 5 10 15 20-1

-0.5

0

0.5

1

1.5

2

CL

-100 -50 0 50 100 150 200 250-1

-0.5

0

0.5

1

1.5

2

CL

CL/C

D

0 0.02 0.04 0.06 0.08 0.1 0.12 0.14-1

-0.5

0

0.5

1

1.5

2

CD

CL

DU-91-W2-250

AF-0901 with v/U = 0

AF-0901 with v/U = -0.001

AF-0901 with v/U = -0.002

AF-0901 with v/U = -0.003

-10 -5 0 5 10 15 20-100

-50

0

50

100

150

200

250

CL/C

D

50 Design issues


Figure 5-3: Effect of suction velocity (soiled case; Re = 8e6; c/r = 0.09)

5.1.3 Apply linear suction

As said before the air from the boundary layer is passively sucked into the blade and released again

at the tip. Because of this passive system the pressure along the length of the blade will vary,

however at a specific blade position, along the chord of an airfoil, the inside pressure is constant.

There the pressure along the upper surface changes along the chord also the pressure difference will

vary along the chord. After a negative pressure peak the pressure rises towards the trailing edge

decreasing the under pressure along the upper surface. For simplicity reasons the suction material

and thickness along the chord is assumed to remain equal. The applied block suction distribution up

to now will not be possible in reality, a linear increasing suction velocity is therefore more realistic.

-10 -5 0 5 10 15 20-1

-0.5

0

0.5

1

1.5

2

CL

-100 -50 0 50 100 150 200 250-1

-0.5

0

0.5

1

1.5

2

CL

CL/C

D

0 0.02 0.04 0.06 0.08 0.1 0.12 0.14-1

-0.5

0

0.5

1

1.5

2

CD

CL

DU-91-W2-250

AF-0901 with v/U = 0

AF-0901 with v/U = -0.001

AF-0901 with v/U = -0.002

AF-0901 with v/U = -0.003

-10 -5 0 5 10 15 20-100

-50

0

50

100

150

200

250

CL/C

D

Design issues 51


Figure 5-4: Different suction distributions

Figure 5-4 shows the original block distribution and three other applied linear distributions. Two of

which have an average suction velocity equal to the original, only starting at a lower value and

ending at a higher one. The third (purple line) starts at zero and rises to a suction velocity of v/U = -

0.003, lowering the average suction velocity over the whole length and thus the total sucked volume

of air. Because of this lower average suction velocity, the performance for this case will be less than

for the other cases. This is illustrated in Figure 5-5 with a lower lift coefficient and lower lift-to-drag

ratio for the purple line (v/U = 0 to -0.003). Although for the other two cases the sucked volume of

air is the same as for the original block distribution, the overall lift-to-drag ratio and lift coefficient is

reduced. In addition the linear suction distribution with a higher starting velocity (the red line in

Figure 5-4 and Figure 5-5) has a slightly better performance in terms of lift-to-drag ratio. The

reasoning behind this effect is analogous to the effect of moving the suction distribution more

forward, preventing a thicker boundary layer is better than curing one.

0 0.2 0.4 0.6 0.8 1 1.2 1.4-5

-4.5

-4

-3.5

-3

-2.5

-2

-1.5

-1

-0.5

0x 10

-3

x/c

v/U

Block suction (v/U = -0.003 constant)

Linear suction (v/U = -0.001 to -0.005)



52 Design issues


Figure 5-5: Effect of different suction distributions (soiled case; Re = 4e6; c/r = 0)

5.2 Suction variation due to gusts

Due to for example gusts the angle of attack of an airfoil could almost instantly increase

considerable, furthermore during this process the rotor has little time to react and the rotational

speed remains constant. As a result of these two events the relative wind velocity and thus the

suction velocity (v/U) is affected. The relation between the angle of attack (α), the rotational speed

(Ω) and the relative wind velocity (Urel) is given in equation (5.1) [3]. The angle of relative wind (φ)

consists of the angle of attack and the section pitch angle (θ), a presentable sectional pitch angle

here is θ = 7.8 degrees, see Table 3-2.

cos cosrel

r

U

(5.1)

As an example case, the angle of attack is increased from α = 5 degrees to α = 15 degrees. Assuming a

constant rotational speed (Ωr = 1), the relation for the relative wind velocity than becomes:

-10 -5 0 5 10 15 20-1

-0.5

0

0.5

1

1.5

2

CL

-60 -40 -20 0 20 40 60 80 100 120 140-1

-0.5

0

0.5

1

1.5

2

CL

CL/C

D

0 0.02 0.04 0.06 0.08 0.1 0.12 0.14 0.16-1

-0.5

0

0.5

1

1.5

2

CD

CL

AF-0901 (v/U = 0.003 constant)

AF-0901 (v/U = 0.001 to 0.005)

AF-0901 (v/U = 0.002 to 0.004)

AF-0901 (v/U = 0.000 to 0.003)

-10 -5 0 5 10 15 20-100

-50

0

50

100

150

CL/C

D

Design issues 53


5

15

5

15

1cos 7.8 5

1cos 7.8 15

cos 7.8 15

cos 7.8 5

U

U

U

U

(5.2)

The new suction velocity is than easily calculated by:

5

15 5 15

cos 7.8 150.003 0.003 0.945 0.00284

cos 7.8 5

Uv v

U U U

(5.3)

If the angle of attack increases from α = 5 degrees to α = 10 degrees the suction velocity becomes v/U

= 0.0029. Therefore the suction velocity is hardly affected by gusts, only a reduction of 5 percent for

extreme cases.

5.3 Wind tunnel considerations

In Figure 5-6 and Figure 5-7 a comparison is given between the original airfoil characteristics

calculated with XFOIL, RFOIL (without rotation) and measured results from wind tunnel tests. Clearly

shown is the improved prediction of RFOIL over XFOIL near and beyond the maximum lift coefficient

compared to the test results. Still RFOIL calculations are a bit too optimistic resulting in a lower drag

and a slightly higher (maximum) lift coefficient for the calculated results. For the clean case the

maximum lift coefficient is well predicted, on the other hand for the rough conditions the maximum

lift coefficient is over predicted. Furthermore it is known from experts that RFOIL usually produces

10-15 percent less drag, for thicker airfoils (t/c > 0.3) the drag underestimation is closer to 15

percent. This is mainly caused by underestimation of the boundary layer thickness by RFOIL [19]. For

the new AF-0901 airfoil similar results are expected from wind tunnel tests.

54 Design issues


Figure 5-6: Comparison of XFOIL, RFOIL and wind tunnel tests (clean case; Re = 3e6; M = 0.21; c/r = 0)

Figure 5-7: Comparison of XFOIL, RFOIL and wind tunnel tests (soiled case; Re = 3e6; M = 0.22; c/r = 0)

5.4 Comparison with similar profile

In section 5.3 the new airfoil is compared with the original DU-91-W2-250 airfoil, in this section the

AF-0901 will be equated with the DU-97-W300 airfoil, which has an equal thickness of t/c =0.3. The

clean and soiled cases for both airfoils are shown in Figure 5-8 and the stiffness properties are given

in Table 5-1. These structural properties were calculated with PreComp and with the original lay-up

as discussed in section 4.2, the stiffness is calculated with the same lay-up for each airfoil. Taking a

-10 -5 0 5 10 15 20 25-1

-0.5

0

0.5

1

1.5

2

CL

-100 -50 0 50 100 150-1

-0.5

0

0.5

1

1.5

2

CL

CL/C

D

0 0.05 0.1 0.15 0.2 0.25 0.3 0.35-1

-0.5

0

0.5

1

1.5

2

CD

CL

DU-91-W2-250 test results

DU-91-W2-250 (XFOIL)

DU-91-W2-250 (RFOIL)

AF-0901 (RFOIL)

-10 -5 0 5 10 15 20 25-100

-50

0

50

100

150

CL/C

D

-10 -5 0 5 10 15 20 25-1

-0.5

0

0.5

1

1.5

2

CL

-100 -50 0 50 100 150-1

-0.5

0

0.5

1

1.5

2

CL

CL/C

D

0 0.05 0.1 0.15 0.2 0.25 0.3 0.35-1

-0.5

0

0.5

1

1.5

2

CD

CL

DU-91-W2-250 test results

DU-91-W2-250 (XFOIL)

DU-91-W2-250 (RFOIL)

AF-0901 (RFOIL)

-10 -5 0 5 10 15 20 25-100

-50

0

50

100

150

CL/C

D

Design issues 55


closer look at Figure 5-8 and Table 5-1 the following can be concluded for a comparison of the AF-

0901 and DU-97-W-300 airfoils:

Up to an angle of attack of α =10 degrees the lift-to-drag ratio for the clean condition is much

higher for the AF-0901 airfoil.

For soiled conditions the lift-to-drag ratio for both airfoils is comparable up to α =5 degrees,

after this point the lift-to-drag ratio of the AF-0901 drops considerable. This better

performance of the DU-97-W-300 airfoil is mainly due to its reduced upper surface thickness.

For the soiled conditions the maximum lift coefficient is higher for the DU-97-W-300 airfoil,

with similar chord lengths this could lead to increased operational loads.

For the clean conditions the AF-0901 airfoil has a slightly higher maximum lift coefficient.

The smooth stall characteristics are comparable for both airfoils, the drop of the lift

coefficient is more significant for the clean case.

Calculated stiffness properties are almost similar for both airfoils, with a small advantage for

the AF-0901 airfoil.

Table 5-1: Structural properties calculated with PreComp

Profile Mass density

Mass number

Flap bending stiffness

Lead-lag bending stiffness

Torsion stiffness

[kg/m] [%] [Nm2] [Nm2] [Nm2]

REFERENCE 0.1050E+10 0.3140E+10 0.1200E+09

DU-91-W2-250 0.4089E+03 100.0 0.1028E+10 0.3299E+10 0.1093E+09

AF-0901 0.4175E+03 102.1 0.1541E+10 0.3511E+10 0.1587E+09

DU-97-W-300 0.4178E+03 102.2 0.1447E+10 0.3504E+10 0.1512E+09

Figure 5-8: Comparison of airfoils with RFOIL (Re = 3e6; M = 0.22; c/r = 0)

-10 -5 0 5 10 15 20 25-1

-0.5

0

0.5

1

1.5

2

CL

-100 -50 0 50 100 150-1

-0.5

0

0.5

1

1.5

2

CL

CL/C

D

0 0.05 0.1 0.15 0.2 0.25 0.3 0.35-1

-0.5

0

0.5

1

1.5

2

CD

CL

AF-0901 (clean)

AF-0901 (soiled)

DU-97-W-300 (clean)

DU-97-W-300 (soiled)

-10 -5 0 5 10 15 20 25-100

-50

0

50

100

150

CL/C

D

Conclusions and recommendations 57


6 Conclusions and recommendations

6.1 Conclusions

Increasing the thickness of the blade results in a greater stiffness for a wind turbine blade, while

keeping the stresses due to bending constant a reduction of material is then possible. However, due

to the thickness increase the aerodynamic properties of the airfoil are reduced. An aerodynamic

redesign with the use of boundary layer suction of an airfoil is for this reason performed. With the

application of boundary layer suction the thickness of the original airfoil was increased from t/c =

0.25 to t/c = 0.30, an increase in thickness of 20%. The newly designed airfoil, the AF-0901, was

optimized while keeping similar aerodynamic performance in terms of maximum lift-to-drag ratio,

maximum lift coefficient and smooth stall control. A high maximum lift-to-drag ratio yields a high

power output, while minimizing the maximum lift coefficient and obtaining smooth stall control

reduces the operational loads on a wind turbine.

The variation of the suction velocity is a possible design option for engineers, this variation will be an

assessment between additional power output or reduced loads. Increasing the suction velocity

affects the maximum lift-to-drag ratio and thereby increasing the power output. However, the

maximum lift coefficient is also increased, which increases the operational loads and reduces the

smoother stall after the maximum lift coefficient.

It is shown that the increased stiffness reduces the use of material considerably. The airfoil was

optimized with respect to stiffness and maximum stress requirements, the possible buckling of the

skin or spar caps due to the reduction of material use was not taken into account. An airfoil can for

example be modelled as a simple beam, with the spar caps as the thick upper and lower flanges, or

as a thin walled shell structure. For the simple beam theory only the spar caps were taken into

account and for the shell structure one global material was assumed for the different sections of the

airfoil. These methods gave a reduction in material of 17%.

For a more in depth analysis the program PreComp was applied, because of this a structural analysis

with different laminates and materials for the spar caps, webs and skins is possible. First, overall

material used for the airfoil was reduced, with this approach the mass density, the section mass per

unit length, was reduced with more than 28%. Secondly, the spar cap material was lowered similar to

the simple beam theory, reducing only the materials used for the spar caps gave a reduction of

almost 19%.


6.2 Recommendations

6.2.1 Wind tunnel test for the AF-0901 airfoil

For the redesign of the original DU-91-W2-250 airfoil RFOIL-suc was used, the program RFOIL-suc is a

combination of the already validated codes RFOIL and XFOIL (with suction). The need of validation of

RFOIL-suc is apparent, just like the results produced by RFOIL-suc. As with most airfoils it is important

to test the AF-0901 airfoil in the wind tunnel during fixed and controllable conditions to verify the

current design. By testing the airfoil in de wind tunnel the design is checked and in addition to this

the application of boundary layer suction can be validated. The preparations for the wind tunnel test

of the AF-0901 are already made by Atiflow BV and the process is in an advanced phase.

6.2.2 Redesign of other airfoils to fit BLS

In this study one airfoil around mid span of the reference UpWind 5MW wind turbine was optimized

for the application of boundary layer suction. To fully understand the effect of boundary layer

suction for the wind turbine, also the airfoils at the other stations should be redesigned to fit

boundary layer suction. With the complete redesign of the blade the performance of the new blade

can be analyzed and the wind turbine can be further optimized.

6.2.3 Intensive study to optimize structural design

The structural benefits of the AF-0901 airfoil compared with the original airfoil were examined, to

fully benefit from the boundary layer suction application the complete blade should be analyzed. An

intensive study of the structure of the new thicker blade is needed to optimize the use of materials

and thereby reducing the costs. Currently a study of the internal structure of the UpWind 5MW wind

turbine is in progress at Actiflow BV. Instead of reducing the blade costs by optimizing the internal

structure, the study could also focus on scaling the original wind turbine blade to increase the power

output without the standard additional costs for the blades. Furthermore a practical boundary layer

suction application is needed in the future to apply on a actual wind turbine so the complete system

including weather resistance can be tested .

Appendix 59


A Airfoil example files

A.1 Coordinate file

1-1.air (DU 91-W2-250)

0.0 0.0 2.00 2.00 6.00

1.000000 0.003247

0.992034 0.005528

0.980658 0.008716

0.966567 0.012514

...

...

0.000000 0.000000

...

...

0.939895 0.006296

0.953314 0.004622

0.966976 0.002574

0.980244 0.000330

0.991806 -0.001748

1.000000 -0.003239

A.2 Suction distribution file

Base suction distribution 1-1.suc v/U=-.001 x/c=0.7-0.95

1.000001 0.000000

0.991323 0.000000

0.979244 0.000000

0.964411 0.000000

0.947401 -0.001000

0.929197 -0.001000

0.910549 -0.001000

0.891742 -0.001000

0.872911 -0.001000

0.854163 -0.001000

0.835456 -0.001000

0.816629 -0.001000

0.797844 -0.001000

0.779249 -0.001000

0.760752 -0.001000

0.742249 -0.001000

0.723695 -0.001000

0.705193 -0.001000

0.686803 0.000000

0.668549 0.000000

0.650394 0.000000

...

0.000000 0.000000

...

0.935261 0.000000

0.947761 0.000000

0.960297 0.000000

0.972619 0.000000

0.983944 0.000000

0.993225 0.000000

1.000001 0.000000

60 Appendix


A.3 Polar file from RFOIL

RFOIL Version 1.0

Calculated polar for: 1-1.air (DU 91-W2-250)

1 Reynolds number fixed

xtrf = 1.000 (suction) 1.000 (pressure)

Rot. Parameters: f0 = 1.000 c/r = 0.090

Mach = 0.000 Re = 8.000 e 6 Ncrit = 9.000

alpha CL CD Re(CL) CM S_xtr P_xtr CDp

------- -------- --------- --------- -------- ------- ------- --------

0.000 0.4610 0.00572 8.00000 -0.1349 0.4334 0.4247 0.00179

0.500 0.5252 0.00575 8.00000 -0.1367 0.4280 0.4324 0.00182

1.000 0.5895 0.00576 8.00000 -0.1384 0.4237 0.4383 0.00185

1.500 0.6534 0.00579 8.00000 -0.1400 0.4200 0.4452 0.00188

2.000 0.7171 0.00585 8.00000 -0.1416 0.4149 0.4511 0.00193

2.500 0.7803 0.00593 8.00000 -0.1431 0.4090 0.4567 0.00200

3.000 0.8438 0.00596 8.00000 -0.1446 0.4037 0.4628 0.00206

3.500 0.9067 0.00606 8.00000 -0.1460 0.3975 0.4687 0.00214

4.000 0.9688 0.00620 8.00000 -0.1472 0.3849 0.4729 0.00225

4.500 1.0302 0.00639 8.00000 -0.1484 0.3680 0.4782 0.00239

5.000 1.0834 0.00735 8.00000 -0.1483 0.2865 0.4835 0.00294

5.500 1.1364 0.00823 8.00000 -0.1481 0.2245 0.4869 0.00350

6.000 1.1895 0.00901 8.00000 -0.1479 0.1765 0.4926 0.00404

6.500 1.2418 0.00978 8.00000 -0.1475 0.1354 0.4967 0.00460

7.000 1.2912 0.01065 8.00000 -0.1466 0.0946 0.5004 0.00524

7.500 1.3414 0.01136 8.00000 -0.1458 0.0704 0.5034 0.00580

8.000 1.3905 0.01204 8.00000 -0.1448 0.0536 0.5090 0.00639

8.500 1.4384 0.01269 8.00000 -0.1436 0.0414 0.5128 0.00697

9.000 1.4835 0.01337 8.00000 -0.1419 0.0333 0.5162 0.00759

9.500 1.5223 0.01407 8.00000 -0.1391 0.0275 0.5198 0.00826

10.000 1.5471 0.01490 8.00000 -0.1337 0.0236 0.5241 0.00909

10.500 1.5759 0.01593 8.00000 -0.1296 0.0210 0.5285 0.01013

11.000 1.6008 0.01724 8.00000 -0.1255 0.0185 0.5317 0.01146

11.500 1.6231 0.01885 8.00000 -0.1217 0.0170 0.5346 0.01310

12.000 1.6400 0.02097 8.00000 -0.1179 0.0158 0.5390 0.01528

12.500 1.6546 0.02355 8.00000 -0.1148 0.0145 0.5427 0.01792

13.000 1.6643 0.02671 8.00000 -0.1121 0.0139 0.5465 0.02118

13.500 1.6673 0.03064 8.00000 -0.1099 0.0134 0.5492 0.02520

14.000 1.6643 0.03512 8.00000 -0.1080 0.0129 0.5515 0.02979

14.500 1.6522 0.04049 8.00000 -0.1067 0.0124 0.5559 0.03529

15.000 1.6418 0.04601 8.00000 -0.1064 0.0120 0.5589 0.04095

15.500 1.6362 0.05144 8.00000 -0.1069 0.0114 0.5616 0.04649

16.000 1.6362 0.05681 8.00000 -0.1082 0.0111 0.5644 0.05198

16.500 1.6430 0.06168 8.00000 -0.1097 0.0109 0.5669 0.05693

17.000 1.6457 0.06749 8.00000 -0.1122 0.0107 0.5709 0.06285

17.500 1.6525 0.07297 8.00000 -0.1146 0.0104 0.5744 0.06843

18.000 1.6603 0.07846 8.00000 -0.1171 0.0102 0.5776 0.07400

18.500 1.6670 0.08412 8.00000 -0.1198 0.0099 0.5805 0.07974

19.000 1.6736 0.08993 8.00000 -0.1225 0.0095 0.5835 0.08563

19.500 1.6849 0.09510 8.00000 -0.1250 0.0091 0.5875 0.09087

20.000 1.6901 0.10128 8.00000 -0.1282 0.0085 0.5911 0.09715

Appendix 61


B Visual Basic Script files

B.1 CreateProfileField.vbs

Batch file: Cscript.exe //NoLogo NewAirfoils.vbs | xfoil_new.exe

' *******************************************************************************

' First we need to declare START and END values for the thickness (y=t/c)

' and position of thickest point (x=x/c at y/c_max),

' plus the number of points and thus the stepsize.

' *******************************************************************************

dblThicknessStart = .25 ' t/c start value (Y=t/c)

dblThicknessEnd = .4 ' t/c end value

dblThickPositionStart = .325 ' x/c at t/c(max) start value (X=x/c)

dblThickPositionEnd = .525 ' x/c at t/c(max) end value

intPoints = 9

' number of points MINUS 1

'(so 10 points --> intPoints=9; first point plus number of steps=intPoints)

' stepsize for thickness

dblThickStepsize = (dblThicknessEnd - dblThicknessStart) / intPoints

' stepsize for thickness point

dblThickPosStepsize = (dblThickPositionEnd - dblThickPositionStart ) / intPoints

' *******************************************************************************

l = 1

m = 1

' *******************************************************************************

' First FOR loop is for going through the X-values (x/c) and for each X-value a new

' loop is created to run through the Y-values (t/c). During the second loop XFOIL

' will create a new airfoil file (X-Y.air) with the current X- and Y-value.

' *******************************************************************************

For i = dblThickPositionStart to dblThickPositionEnd step dblThickPosStepsize

' loop for x/c

dblThickPosition = i

For j = dblThicknessStart to dblThicknessEnd step dblThickStepsize

'loop for t/c

dblThickness = j

'************************************************************************

' Xfoil inputs

'************************************************************************

Wscript.Stdout.WriteLine "load du25.air" ' load base irfoil

Wscript.Sleep 10

Wscript.Stdout.WriteLine "gdes" ' enter GDES menu

Wscript.Sleep 10

Wscript.Stdout.WriteLine "tset" ' set new thickness and camber

Wscript.Sleep 10

Wscript.Stdout.WriteLine dblThickness ' new thickness

Wscript.Sleep 10

Wscript.Stdout.WriteLine "" ' ENTER: no new camber

Wscript.Sleep 10

Wscript.Stdout.WriteLine "exec" ' set current airfoil

Wscript.Sleep 10

Wscript.Stdout.WriteLine "high" ' move camber and thickness highpoints

Wscript.Sleep 10

62 Appendix


Wscript.Stdout.WriteLine dblThickPosition ' new thickness highpoint

Wscript.Sleep 10

Wscript.Stdout.WriteLine "" ' ENTER: no new camber highpoint

Wscript.Sleep 10

Wscript.Stdout.WriteLine "exec" ' set current airfoil

Wscript.Sleep 10

Wscript.Stdout.WriteLine "" ' ENTER: go to OPER menu

Wscript.Sleep 10

Wscript.Stdout.WriteLine "name" ' name airfoil

Wscript.Sleep 10

Wscript.Stdout.WriteLine l & "-" & m & ".air (DU 91-W2-250)" ' name

Wscript.Sleep 10

Wscript.Stdout.WriteLine "isav" ' write airfoil to ISES coordinate file

Wscript.Sleep 10

Wscript.Stdout.WriteLine l & "-" & m & ".air" ' write as XY.air

Wscript.Sleep 10

'***********************************************************************

m=m+1

Next

m=1

l=l+1

Next

Wscript.Stdout.WriteLine "quit" ' quit XFOIL

Wscript.Sleep 10

' *******************************************************************************

B.2 CreateSuction.vbs

Batch file: Cscript.exe //NoLogo CreateSuction.vbs

' Create explicit variables

Option Explicit

Dim objShell ' run/quit xfoil

Dim x,y 'integers

' ********************************************************************************

' First FOR loop is for going through the X-values (x/c) and for each X-value a new

' loop is created to run through the Y-values (t/c). --> x-y.air

' ********************************************************************************

For x = 1 to 10 step 1 ' loop for x/c

For y = 1 to 10 step 1 'loop for t/c

' Create Object Shell to run and quit Xfoil

Set objShell = CreateObject("WScript.Shell")

objShell.Run "rfoilsuc", 3, false ' Open Xfoilsuc

Wscript.Sleep 600

'========================================================================

' Rfoil inputs

'========================================================================

objShell.SendKeys "iloa" & "{ENTER}" ' load airfoil

Wscript.Sleep 100

objShell.SendKeys x & "-" & y & ".air" & "{ENTER}" ' load airfoil

Wscript.Sleep 100

objShell.SendKeys "oper" & "{ENTER}" ' enter OPER menu

Wscript.Sleep 100

objShell.SendKeys "vdes" & "{ENTER}" ' enter VDES menu

Wscript.Sleep 100

objShell.SendKeys "3e6" & "{ENTER}" ' set Reynolds number

Wscript.Sleep 100

objShell.SendKeys "load" & "{ENTER}" ' load suction distribution

Wscript.Sleep 100

objShell.SendKeys "base.suc" & "{ENTER}" ' enter RFoilSuc base suction

Wscript.Sleep 100

Appendix 63


objShell.SendKeys "exec" & "{ENTER}" ' set active Vsuc from buffer Vsuc

Wscript.Sleep 100

objShell.SendKeys "save" & "{ENTER}" ' save current Vsuc to file

Wscript.Sleep 100

objShell.SendKeys x & "-" & y & ".bsuc" & "{ENTER}" ' (x-y.suc)

Wscript.Sleep 100

objShell.SendKeys x & "-" & y & ".bsuc" & "{ENTER}" ' (x-y.suc)

Wscript.Sleep 100

objShell.SendKeys "{ENTER}"

Wscript.Sleep 100


Wscript.Sleep 100


Wscript.Sleep 100

objShell.SendKeys "quit" & "{ENTER}" 'quit RFOIL

Wscript.Sleep 100

objShell.SendKeys "%y"

WScript.Sleep 900

Next

Next

' End script

WScript.Quit

B.3 CreatePolar.vbs

Batch file: Cscript.exe //NoLogo Createpolar.vbs

' Create explicit variables

Option Explicit

Dim objShell, objExec, objWMIService, objProcess, colProcess ' run/quit xfoil

Dim strComputer, strProcessKill 'quit routine xfoil

Dim x, y, s 'integers

' String variables needed for Quit routine

strComputer = "."

strProcessKill = "'rfoilsuc.exe'"

For x = 1 to 10 step 1 ' loop for x/c

For y = 1 to 10 step 1 ' loop for t/c

For s = 0 to 6 step 1 ' suction scale factor

' Create Object Shell to run and quit Xfoil

Set objShell = CreateObject("WScript.Shell")

objShell.Run "rfoilsuc", 3, false ' Open Xfoilsuc

Wscript.Sleep 900

'============================================================================

' Rfoil inputs

'============================================================================

' Note 1:

' Maximize Rfoil window for input of VBScript otherwise graphic window is the

' active window and no input is possible anymore. When window is maximized the

' 'input' window will be the active window.

' Note 2:

' When Rfoil is calculating the window is not active and the script has to

' "sleep" untill Rfoil is finished or the commands are lost. This is different

' from Xfoil, where the script commands will wait for Xfoil to finish, the

' commands are queued and then past to Xfoil (This doesn't work for Rfoil!!!).

objShell.SendKeys "%{ENTER}" ' maximize window

WScript.Sleep 1000

64 Appendix


objShell.SendKeys "iloa" & "{ENTER}" ' load airfoil

Wscript.Sleep 1000

objShell.SendKeys x & "-" & y & "-b.air" & "{ENTER}" ' airfoil name

Wscript.Sleep 1000

objShell.SendKeys "oper" & "{ENTER}" ' enter OPER menu

Wscript.Sleep 1000

objShell.SendKeys "vdes" & "{ENTER}" ' enter VDES menu

Wscript.Sleep 1000

objShell.SendKeys "8e6" & "{ENTER}" ' set Reynolds number

Wscript.Sleep 1000

objShell.SendKeys "load" & "{ENTER}" ' load suction distribution

Wscript.Sleep 1000

objShell.SendKeys x & "-" & y & ".suc" & "{ENTER}" ' enter base suction

Wscript.Sleep 1000


Wscript.Sleep 1000

objShell.SendKeys "scal" & "{ENTER}" ' scale buffer Vsuc

Wscript.Sleep 1000

objShell.SendKeys s & "{ENTER}" ' scale factor

Wscript.Sleep 1000


Wscript.Sleep 1000

objShell.SendKeys "{ENTER}" ' ENTER: return to OPER menu

Wscript.Sleep 1000

objShell.SendKeys "vpar" & "{ENTER}" ' enter VPAR menu

Wscript.Sleep 1000

objShell.SendKeys "xtr" & "{ENTER}" ' change trip positions Xtr/c

Wscript.Sleep 1000

objShell.SendKeys "0.05" & "{ENTER}" ' enter suction side Xtr/c

Wscript.Sleep 1000

objShell.SendKeys "1" & "{ENTER}" ' enter pressure side Xtr/c

Wscript.Sleep 1000

objShell.SendKeys "cr" & "{ENTER}" ' change c/r ratio

Wscript.Sleep 1000

objShell.SendKeys "0.09" & "{ENTER}" ' set c/r ratio

Wscript.Sleep 1000

objShell.SendKeys "{ENTER}" ' ENTER: return to OPER menu

Wscript.Sleep 1000

objShell.SendKeys "pacc" & "{ENTER}" ' start auto point accumulation

Wscript.Sleep 1000

objShell.SendKeys x & "-" & y & "-s" & s & ".plr" & "{ENTER}"

' enter polar name

Wscript.Sleep 1000

objShell.SendKeys "{ENTER}" ' ENTER: no dump file

Wscript.Sleep 1000

objShell.SendKeys "aseq" & "{ENTER}" ' calculate sequence of alpha's

Wscript.Sleep 1000

objShell.SendKeys "-10" & "{ENTER}" ' enter first alpha

Wscript.Sleep 1000

objShell.SendKeys "20" & "{ENTER}" ' enter last alpha

Wscript.Sleep 1000

objShell.SendKeys "0.5" & "{ENTER}" ' enter alpha increment

Wscript.Sleep 1000

' Check status Xfoil (0=running)

Set objExec = objShell.Exec("rfoilsuc")

' Display status Xfoil in CMD window

WScript.Echo "Status rfoilsucblind = " & objExec.Status

Wscript.Sleep 1000

' Quit Xfoil

If objExec.Status = 0 Then

' Search for Xfoil with strProcessKill

Set objWMIService = GetObject("winmgmts:" &

"{impersonationLevel=impersonate}!\\" & strComputer & "\root\cimv2")

Set colProcess = objWMIService.ExecQuery ("Select * from Win32_Process Where

Name = " & strProcessKill )

' Terminate Xfoil

Appendix 65


For Each objProcess in colProcess

objProcess.Terminate()

Next

' Display Xfoil killed in CMD window

WSCript.Echo "Just killed process " & strProcessKill & "-" & x & "-" & y & "-"

& s

End If

Wscript.Sleep 900

Next

Next

Next

' End script

WScript.Quit

Appendix 67


C MATLAB files

C.1 CreateCorrectSuction.M

for x = 1:10 for y = 1:10 % sucfilename = x-y.bsuc sucfilename = [num2str(x,15) '-' num2str(y,15) '.bsuc']

% read textfilename for positon and suction velocity [pos vel] = textread(sucfilename,'%f %f','headerlines',1);

% number of entries in pos length_pos = length(pos);

% size SUC = [2,length(pos)] suc1 = pos'; % 1st row is position suc2 = zeros(1,length_pos); % 2nd row is suction velocity SUC = [suc1; suc2];

% set end of loop (only first half of positions-> % upper half of airfoil!!) i_end = length_pos/2;

% if position is between suction_start and suction_end % set velocity to -0.001 % Suction from x/c=.70 to x/c=.95 suc_end = 0.95; suc_start = 0.70; for i = 1 : i_end if SUC(1,i)>suc_start && SUC(1,i)<suc_end SUC(2,i) = -0.001; end end

% set suction area!! suc_area = '7095'; writefilename = [num2str(x,15) '-' num2str(y,15) '.suc'] suctionname = ['Base suction distribution ' num2str(x,15) '-'... num2str(y,15) '.suc ' 'v/U=-.001 x/c=' num2str(suc_start,15)... '-' num2str(suc_end,15)]

% First write suctionname to first line of writefilename fid = fopen(writefilename,'wt'); fprintf(fid,'%s\n',suctionname); fclose(fid);

% Append suction distribution (matrix SUC) to writefilename fid = fopen(writefilename,'at'); fprintf(fid,'%f %f\n',SUC); fclose(fid); end end

68 Appendix


C.2 FilterData.M

%========================================================================== % CreateMatFiles.M % Read in polar files to MATLAB, creat variables (Alpha CL CD) and filtered % variables (only store up to CL_max) %========================================================================== for x = 1 : 10 % loop for x/c for y = 1 : 10 % loop for t/c for s = 0 : 1: 6 % loop for suction scale factor

% clear existing polar variables clear Alpha CL CD Re CM Top_Xtr ... Bot_Xtr Dp Alpha_new CL_new CD_new

% create textfilename (x-y.plr) to read file textfilename = [num2str(x,15) '-' num2str(y,15)... '-s' num2str(s,15) '.splr']

% try, catch, end: for error handling try

% read x-y.plr file with 8 variables and 13 text lines [Alpha CL CD Re CM Top_Xtr Bot_Xtr Dp] = ... textread(textfilename,'%f %f %f %f %f %f %f %f',... 'headerlines',13);

% filter Cl values, only store for increasing Cl q = length(CL); CL_fil(1,1) = CL(1); CD_fil(1,1) = CD(1); Alpha_fil(1,1) = Alpha(1); for r = 2:q if CL(r) > CL(r-1) CL_fil(r,1) = CL(r); CD_fil(r,1) = CD(r); Alpha_fil(r,1) = Alpha(r); else break end end

% create savefilename to save to .mat file (x-y.mat) savefilename = [num2str(x,15) '-' num2str(y,15) '-v' ... num2str(s,15) '-soiled']

% save desired variables (Alpha, CL, CD) save (savefilename, 'Alpha', 'CL', 'CD',... 'Alpha_fil', 'CL_fil', 'CD_fil') % OUTPUT: X-Y-vS-soiled-fil.mat % [Alpha, CL, CD (_fil)] % X = 1:10, Y=1:10, S=1:6 catch end end end end

Appendix 69


%========================================================================== % CreatTotalPolarsMat(-fil).M % Combine polars of different suction velocities, find for each alpha % biggest v/U possible and store only that point. This is done for each % step of v/U, thus creating only one polar for each v/U. % Note: has to run separately for unfiltered results if desired %========================================================================== clear all for x = 1 : 10 % loop for x/c for y = 1 : 10 % loop for t/c for v = 0:1:6 % loop for v/U

% create empty vectors Alpha_tot = []; CL_tot = []; CD_tot = []; for s = 0 : 1: v % clear existing polar variables clear Alpha CL CD Alpha_fil CL_fil CD_fil

% create textfilename to read .M-file (Alpha,CL,CD) openfilename = [num2str(x,15) '-' num2str(y,15) '-v'... num2str(s,15) '-soiled-fil']

% try, catch, end: for error handling try load(openfilename) % store current alpha, Cl and Cd Alpha_tot = [Alpha_tot;Alpha_fil]; CL_tot = [CL_tot;CL_fil]; CD_tot = [CD_tot;CD_fil]; catch end end

% create empty vectors ALPHA =[]; CL = []; CD = []; for a = 0: 0.5: 20 % find indices for alpha index = find(Alpha_tot == a);

% find index with highest suction velocity index_max = max(index);

% set Cl and Cd for alpha with highest suction velocity ALPHA_NEW = Alpha_tot(index_max); CL_NEW = CL_tot(index_max); CD_NEW = CD_tot(index_max);

% store current alpha, Cl and Cd ALPHA =[ALPHA;ALPHA_NEW]; CL = [CL;CL_NEW]; CD = [CD;CD_NEW] ; end savefilename = [num2str(x,15) '-' num2str(y,15)... '-TOTAL-v' num2str(v,15) '-soiled-fil'] save (savefilename, 'ALPHA', 'CL', 'CD')

70 Appendix


% OUTPUT: X-Y-TOTAL-vS-soiled-fil.mat % [ALPHA, CL, CD] % X = 1:10, Y=1:10, S=1:6 end end end

%========================================================================== % CreateResultsMatrixTOTALMat.M % Construct matrices from 10x10 airfoils (x-y): % - Cl/Cd_max with Alpha and Cl % - Cl_max with Alpha % - Cl_dif (= Cl_max - Cl_design) % - Alpha_dif (= Alpha_max - Alpha_design) %========================================================================== clear all clcd_max(10,10)=0; cl_clcdmax(10,10)=0; alpha_clcdmax(10,10)=0; cl_max(10,10)=0; cl_dif(10,10)=0; alpha_clmax(10,10)=0; alpha_dif(10,10)=0; cl_design(10,10)=0; alpha_design(10,10)=0; clcd_design(10,10)=0; for s = 0:6 % Suction velocity for n = 1:10 % length(x) = n for m = 1:10 % length(y) = m

% clear existing polar variables clear ALPHA CL CD

% create textfilename (x-y-c) to load file textfilename = [num2str(n,15) '-' num2str(m,15) '-TOTAL-v'... num2str(s,15) '-soiled-FIL']

% Error handling try % load x-y-c.mat = l-m-c.mat load(textfilename);

% calculate Cl/Cd clcd = (CL./CD);

% locate Cl/Cd_max and its index number [clcdmax,clcdmaxindex] = max(clcd);

% store Cl/Cd_max for this airfoil clcd_max(m,n)=clcdmax;

% store Cl at Cl/Cd_max for this airfoil cl_clcdmax(m,n) = CL(clcdmaxindex);

% store alpha at Cl/Cd_max for this airfoil alpha_clcdmax(m,n) = ALPHA(clcdmaxindex);

Appendix 71


% locate Cl_max with index number [clmax,clmaxindex] = max(CL);

% store Cl_max for this airfoil cl_max(m,n) = clmax;

% store Cl difference (Cl_max - Cl_design) cl_dif(m,n) = clmax - cl_clcdmax(m,n);

% store alpha at Cl_max for this airfoil alpha_clmax(m,n) = ALPHA(clmaxindex);

% store alpha difference alpha_dif(m,n) = alpha_clmax(m,n) - alpha_clcdmax(m,n);

% alternative cl_design = cl_alt: % Cl at which Cl/Cd > Cl/Cl_original AND difference with its Cl_max > 0.2 % Cl_dif = Cl_max - Cl_alt = 0.2 % Cl_alt = Cl_max - 0.2 % CL/CD(Cl_alt) > CL/CD_original -> new criteria

% find index CL for cl_des with Cl_alt = Cl_max - 0.2 CL_alt = clmax - 0.2; clindex = find(CL_alt>CL(1:clmaxindex),1,'last'); % store alpha, CL, CL/CD at CL_alt cl_alt(m,n) = CL(clindex); alpha_alt(m,n) = ALPHA(clindex); clcd_alt(m,n) = clcd(clindex); catch end end end savefilename = ['ResultMatrix-TOTAL-v' num2str(s,15) '-soiled-fil'] save(savefilename, 'clcd_max', 'cl_clcdmax', 'alpha_clcdmax',... 'cl_max', 'cl_dif', 'alpha_clmax', 'alpha_dif', 'cl_alt',... 'alpha_alt', 'clcd_alt') % OUTPUT: ResultsMatrix-TOTAL-vS-soiled-fil.mat % Matrices of 10*10 of all data for each suction velocity % X = 1:10, Y=1:10, S=1:6 end

72 Appendix


C.3 MATLAB example data

clcd_max =

Columns 1 through 5

161.2207 151.7559 148.8401 149.1667 128.0426

168.2389 165.3907 154.9745 146.1102 161.9959

176.7338 170.7478 162.9937 163.8294 164.2830

182.5134 179.7234 176.0752 171.6393 159.8957

188.0278 186.9041 179.3478 171.2809 164.1402

191.1794 189.6565 181.2650 173.7659 175.7437

193.3095 191.4145 183.0591 175.9728 175.8045

191.3734 186.9722 178.9560 172.2978 166.8831

184.0446 178.2220 174.2066 164.3876 145.6906

173.4516 160.8226 155.0182 131.6646 89.6871

Columns 6 through 10

152.7335 172.1963 0 112.3294 158.3161

131.2391 170.4619 170.8333 112.6139 109.1388

148.3164 143.9871 172.3326 144.1538 93.6770

175.9506 143.9683 158.6437 156.8980 0

177.9099 154.2240 136.4428 147.5092 49.2602

179.6741 165.5422 127.6557 68.6340 31.6851

173.7559 149.9078 92.7907 36.0241 24.3335

157.5816 90.8169 43.3645 25.9863 19.8167

90.4745 46.2875 42.9615 20.1333 14.6240

47.4159 43.7048 23.7803 17.1489 11.8126

Appendix 73


D Structural properties tables

Table D-1: Lay-up description[14]

Layer ID Material name Label Radius Nr of layers Thickness per layer Total thickness

[m] [-] [mm] [mm]

TRIAX-1 UD45R 2.0 3 0.94 2.82

TRIAX-1 UD45R 63.5 3 0.94 2.82

UD UD_OB K 29.0 172 0.47 80.94

UD_TE UD_OB U 14.0 29 0.47 13.60

UD_TE UD_OB V 47.0 1 0.47 0.47

SKINFOAM SKINFOAM X 7.5 40.00

SKINFOAM SKINFOAM Y 43.0 40.00

TRIAX-3 UD45R 2.0 3 0.94 2.82

TRIAX-3 UD45R 63.5 3 0.94 2.82

Table D-2: Material properties [14]

Nr. Material name E11 E22 G12 nu12 density

[MPa] [MPa] [MPa] [-] [kg/m3]

1 UD_OB 38887 9000 3600 0.249 1869

2 UD45R 24800 11500 4861 0.416 1826

3 R4545 11700 11700 9770 0.501 1782

4 SKINFOAM 256 256 22 0.300 200

5 WEBPS 25 25 12 0.300 45

Appendix 75


E Structural lay-up pictures

Figure E-1: Blade profile lay-out [14]

76 Appendix


Figure E-2: Blade laminate lay-up [14]

Appendix 77


Figure E-3: Profile lay-out example [14]

Figure E-4: Ply angle definition [2]

Appendix 79


F PreComp example input files

F.1 Main Input File: “DU25.pci”

***************** main input file for PreComp ***************************** DU25 Composite Blade Section Properties General information ----------------------------------------------- 1 Bl_length : blade length (m) 2 N_sections : no of blade sections (-) 5 N_materials : no of materials listed in the materials table (material.inp) 3 Out_format : output file (1: general format, 2: BModes-format, 3: both) f TabDelim (true: tab-delimited table; false: space-delimited table) Blade-sections-specific data -------------------------------------- Sec span l.e. chord aerodynamic af_shape int str layup location position length twist file file Span_loc Le_loc Chord Tw_aero Af_shape_file Int_str_file (-) (-) (m) (degrees) (-) (-) 0.0000 0.38 3.950 0.00 'af-du25.inp' 'int-du25.inp' 1.0000 0.38 3.950 0.00 'af-du25.inp' 'int-du25.inp' Webs (spars) data -------------------------------------------------- 3 Nweb : number of webs (-) ! enter 0 if the blade has no webs 1 Ib_sp_stn : blade station number where inner-most end of webs is located (-) 2 Ob_sp_stn : blade station number where outer-most end of webs is located (-) Web_num Inb_end_ch_loc Oub_end_ch_loc (fraction of chord length) 1 0.28 0.28 2 0.48 0.48 3 1.00 1.00

F.2 Airfoil Data File: “af-du25.inp”

54 N_af_nodes :no of airfoil nodes, counted clockwise starting with leading edge (see users' manual, fig xx) Xnode Ynode !! chord-normalized coordinated of the airfoil nodes 0.00E+00 0.00E+00 !! the first node, a leading-edge node, must be (0,0) 0.00043 0.00553 0.00281 0.01348 0.00773 0.02130 0.01466 0.02925 0.02371 0.03760 0.03537 0.04659 0.05045 0.05645 0.07015 0.06741 0.09594 0.07949 0.12904 0.09226 0.16931 0.10467 0.21497 0.11538 0.26374 0.12332 0.31393 0.12791 0.36525 0.12868 0.41936 0.12564 0.47662 0.11945 0.53667 0.11050 0.59893 0.09918

80 Appendix


0.66299 0.08594 0.72862 0.07119 0.79504 0.05546 0.86107 0.03927 0.92291 0.02374 0.97374 0.01061 1.00000 0.00325 1.00000 -0.00325 0.97432 0.00134 0.93259 0.00703 0.88960 0.00879 0.84525 0.00612 0.79916 -0.00105 0.75024 -0.01301 0.69665 -0.02979 0.63780 -0.05035 0.57811 -0.07102 0.52124 -0.08880 0.46710 -0.10289 0.41466 -0.11311 0.36305 -0.11960 0.31253 -0.12247 0.26354 -0.12161 0.21654 -0.11711 0.17270 -0.10939 0.13356 -0.09917 0.10056 -0.08767 0.07415 -0.07610 0.05368 -0.06514 0.03795 -0.05483 0.02586 -0.04504 0.01654 -0.03576 0.00943 -0.02696 0.00437 -0.01845

F.3 Internal Structure Data File: “int-du25.inp”

Composite laminae lay-up inside the blade section *************************** TOP SURFACE **************************** 5 N_scts(1): no of sectors on top surface normalized chord location of nodes defining airfoil sectors boundaries (xsec_node) 0.0 0.28 0.48 0.80 0.95 1.00 .................................................................. Sect_num no of laminae (N_laminas) 1 3 lamina num of thickness fibers_direction composite_material ID number plies of ply (m) (deg) (-) lam_num N_plies Tply Tht_lam Mat_id 1 3 0.000940 45 2 (UD45R) 2 1 0.040 0 4 (SKINFOAM) 3 3 0.000940 45 2 (UD45R) .................................................................. Sect_num no of laminae 2 3 lamina num of thickness fibers_direction composite_material ID number plies of ply (m) (deg) (-) lam_num N_plies Tply Tht_lam Mat_id 1 3 0.000940 45 2 (UD45R) 2 172 0.00047 0 1 (UD_OB) 3 3 0.000940 45 2 (UD45R) .................................................................. Sect_num no of laminae 3 3 lamina num of thickness fibers_direction composite_material ID number plies of ply (m) (deg) (-)

Appendix 81


lam_num N_plies Tply Tht_lam Mat_id 1 3 0.000940 45 2 (UD45R) 2 1 0.040 0 4 (SKINFOAM) 3 3 0.000940 45 2 (UD45R) .................................................................. Sect_num no of laminae 4 3 lamina num of thickness fibers_direction composite_material ID number plies of ply (m) (deg) (-) lam_num N_plies Tply Tht_lam Mat_id 1 3 0.000940 45 2 (UD45R) 2 16.3 0.00047 0 1 (UD_OB) 3 3 0.000940 45 2 (UD45R) .................................................................. Sect_num no of laminae 5 1 lamina num of thickness fibers_direction composite_material ID number plies of ply (m) (deg) (-) lam_num N_plies Tply Tht_lam Mat_id 1 6 0.000940 45 2 (UD45R) *************************** BOTTOM SURFACE **************************** 5 N_scts(2): no of sectors on bottom surfaces normalized chord location of nodes defining airfoil sectors boundaries (xsec_node) 0.0 0.28 0.48 0.80 0.95 1.00 .................................................................. Sect_num no of laminae (N_laminas) 1 3 lamina num of thickness fibers_direction composite_material ID number plies of ply (m) (deg) (-) lam_num N_plies Tply Tht_lam Mat_id 1 3 0.000940 45 2 (UD45R) 2 1 0.040 0 4 (SKINFOAM) 3 3 0.000940 45 2 (UD45R) .................................................................. Sect_num no of laminae 2 3 lamina num of thickness fibers_direction composite_material ID number plies of ply (m) (deg) (-) lam_num N_plies Tply Tht_lam Mat_id 1 3 0.000940 45 2 (UD45R) 2 172 0.00047 0 1 (UD_OB) 3 3 0.000940 45 2 (UD45R) .................................................................. Sect_num no of laminae 3 3 lamina num of thickness fibers_direction composite_material ID number plies of ply (m) (deg) (-) lam_num N_plies Tply Tht_lam Mat_id 1 3 0.000940 45 2 (UD45R) 2 1 0.040 0 4 (SKINFOAM) 3 3 0.000940 45 2 (UD45R)) .................................................................. Sect_num no of laminae 4 3 lamina num of thickness fibers_direction composite_material ID number plies of ply (m) (deg) (-) lam_num N_plies Tply Tht_lam Mat_id 1 3 0.000940 45 2 (UD45R) 2 16.3 0.00047 0 1 (UD_OB) 3 3 0.000940 45 2 (UD45R) .................................................................. Sect_num no of laminae 5 1 lamina num of thickness fibers_direction composite_material ID number plies of ply (m) (deg) (-) lam_num N_plies Tply Tht_lam Mat_id 1 6 0.000940 45 2 (UD45R)

82 Appendix


********************************************************************** Laminae schedule for webs (input required only if webs exist at this section): web_num no of laminae (N_weblams) 1 3 lamina num of thickness fibers_direction composite_material ID number plies of ply (m) (deg) (-) wlam_num N_Plies w_tply Tht_Wlam Wmat_Id 1 2 0.00153 45 3 (R4545) 2 1 0.050 0 5 (WEBPS) 3 2 0.00153 45 3 (R4545) web_num no of laminae 2 3 lamina num of thickness fibers_direction composite_material ID number plies of ply (m) (deg) (-) wlam_num N_Plies w_tply Tht_Wlam Wmat_Id 1 2 0.00153 45 3 (R4545) 2 1 0.050 0 5 (WEBPS) 3 2 0.00153 45 3 (R4545) web_num no of laminae 3 1 lamina num of thickness fibers_direction composite_material ID number plies of ply (m) (deg) (-) wlam_num N_Plies w_tply Tht_Wlam Wmat_Id 1 6 0.000940 45 2 (UD45R)

F.4 Materials Data File: “materials.inp”

Mat_Id E1 E2 G12 Nu12 Density Mat_Name (-) (Pa) (Pa) (Pa) (-) (Kg/m^3) (-) 1 38.9e+9 9.0e+9 3.6e+9 0.25 1869.0 (UD_OB) 2 24.8e+9 11.5e+9 4.9e+9 0.42 1826.0 (UD45R) 3 11.7e+9 11.7e+9 9.8e+9 0.50 1782.0 (R4545) 4 25.6e+7 25.6e+7 2.2e+7 0.30 200.0 (SKINFOAM) 5 2.5e+7 2.5e+7 1.2e+7 0.30 45.0 (WEBPS)

Appendix 83


G PreComp example output files

G.1 BModes Output File: “du25.out_bmd”

================================================================================== Results generated by PreComp (v1.00.01, 03-Aug-2006) on 07-Oct-2009 at 14:04:58. DU25 Composite Blade Section Properties ================================================================================== blade length (meters) = 1.00 span_loc str_tw tw_iner mass_den flp_iner edge_iner flp_stff (-) (deg) (deg) (kg/m) (kg-m) (kg-m) (Nm^2) 0.0000 2.011 1.712 0.4089E+03 0.5839E+02 0.3137E+03 0.1028E+10 1.0000 2.011 1.712 0.4089E+03 0.5839E+02 0.3137E+03 0.1028E+10 edge_stff tor_stff axial_stff cg_offst sc_offst tc_offst (Nm^2) (Nm^2) (N) (m) (m) (m) 0.3299E+10 0.1093E+09 0.6314E+10 0.283 0.158 0.225 0.3299E+10 0.1093E+09 0.6314E+10 0.283 0.158 0.225

Appendix 85


H Pressure gradient in a closed rotating duct

The targeted duct is a closed rotating rectangular straight duct. One end of the duct is in the centre

of rotation, the other end, further referred as tip, is rotating in a circle with radius r, the length of the

duct. The tip speed, Vtip, is calculated by:

tipV r (H.1)

with ω the angular velocity in rad/s. The volume of air enclosed by the duct is:

Volume length width height r width height (H.2)

And the mass of this volume is therefore:

m r width height (H.3)

The centre of this mass lies at 0.5r. The centrifugal force, Fc, is therefore:

2 0.5cF m r (H.4)

The centrifugal force will cause a pressure gradient between the two ends of the duct, interest is in

calculating this pressure gradient and thereby the maximum under pressure that can be generated in

the duct.

cp width height F (H.5)

With (width ∗ height) being the cross section of the duct. Substituting above equation leads to:

86 Appendix


cFpwidth height

(H.6)

2 0.5r width height r

pwidth height

(H.7)

2

0.5p r (H.8)

or

20.5 tipp V (H.9)

With help Bernoulli equation

21

2V gh p constant (H.10)

the inside pressure becomes

2 21 1

2 2R gR r gr

(H.11)

The outside pressure at the suction area is

21

2rel PV C (H.12)

Bibliography

[1] J. Bink, "Basic Pressure Gradient Calculation in a Closed Rotating Duct," Actiflow B.V., Breda,

2009.

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Technology, Delft, 2008.

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Composite Wind Turbine Blade Properties," 2009.

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[6] M. Drela, "XFOIL: An Analysis and Design System for Low Reynolds Number Airfoils.," 1989.

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speed flight employing turbulent boundary layer suction over the rear part," Delft University of

Technology, Delft, 1985.

[10] P.J. van Langen, "Aeroelastic analysis of the UpWind reference wind turbine," ECN, Petten, ECN-

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[11] J. Lousberg, "Boundary Layer Control in Wind Turbine desing: A Preliminary Study of the

Potential for Application," Actiflow B.V., Master's Thesis 2008.

[12] J.L. Meriam and L.G. Kraige, Enginering Mechanics: Statics. USA: John Wiley & Sons, Inc, 1998.

[13] Microsoft Corporation. (2009, Mar.) VBScript User's Guide. [Online].

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[14] R. Nijssen, G.D. de Winkel, and J.M. Peeringa, "UpWind WMC5MW laminate data," WMC,

Wieringerwerf, 2007.

[15] PreComp. (2007, March) NWTC Design Codes. [Online].

http://wind.nrel.gov/designcodes/preprocessors/precomp/

[16] R.P.J.O.M. van Rooij, "Aerodynamic design for a pitch regulated rotor," Delft University of


Technology, Delft, WE-01172, 2001.

[17] R.P.J.O.M. van Rooij, "Modification of the boundary layer calculation in RFOIL for improved

airfoil stall prediction," Delft University of Technology, Delft, IW-96087R, 1996.

[18] R.P.J.O.M. van Rooij and W.A. Timmer, "Roughness sensitivity considerations for thick rotor

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