Investigation of Root Spoilers on Horizontal Axis Wind

Transcription

Master of Science Thesis
Investigation of Root Spoilers on
Horizontal Axis Wind Turbines
Anantha Padmanabhan Kidambi Sekar
Version: Public
Investigation of Root Spoilers on
Horizontal Axis Wind Turbines
Master of Science Thesis
For obtaining the degree of Master of Science in Aerospace Engineering
at Delft University of Technology
Version: Public
Faculty of Aerospace Engineering
·
Delft University of Technology
Delft University of Technology
c Aerospace Engineering, Delft University of Technology
Copyright All rights reserved.
This thesis was carried out at the Energy Research Center of Netherlands.
c Energy Research Center of Netherlands,Wind Energy Technology
Copyright All rights reserved.
iv
M.Sc. Thesis
DELFT UNIVERSITY OF TECHNOLOGY
DEPARTMENT OF AEROSPACE ENGINEERING
The undersigned hereby certify that they have read and recommend to the Faculty of
Aerospace Engineering for acceptance the thesis entitled “Investigation of Root Spoilers
on Horizontal Axis Wind Turbines” by Anantha Padmanabhan Kidambi Sekar in
fulfillment of the requirements for the degree of Master of Science.
Version: Public
Supervisors:
Dr.Ir.C.J.Ferreira
Ir.W.A Timmer
Dr.A.Sciacchitano
Dr.Ir.J.G.Schepers
Dr.M.Caboni
Preface
For the purpose of flow analysis, the blade of a horizontal axis wind turbine (HAWT) can be
divided into the root, mid and the tip regions. The flow in the root region is very complex
and is affected due to the rotation of the blade and separation due to the presence of very
thick airfoil sections and the higher angle of attacks. Hence, a detailed knowledge of the flow
characteristics in the root region is needed to accurately create models that will represent
the flow characteristics in the root region. There exists an imbalance between the number of
studies conducted to analyse the root region of the HAWT blade when compared to the tip
region mainly because the contribution of the power production from the tip region of the
blade.
It is imperative that the physics behind the flow of the entire wind turbine blade is well known
and documented before an attempt can be made to model very high capacity wind turbines
(10MW). The thesis attempts to address these requirements by simulating the flow in the
root sections of the blades by means of experiments and also looking at the effects of spoilers
on the root sections of the blade.
As the aerodynamic forces in the root regions are not as high as the tip region, there is a lower
contribution to the aerodynamic and structural forces from the root regions. This study aims
to answer the question of the effect of lift enhancing devices (in this case, a spoiler) on the
flow characteristics and the power production of the blade.
The experiment design and the design of the blade was done by using the unsteady three
dimensional panel code (UMPM) from TU Delft and the lifting line free vortex code (AWSM)
from ECN. A model was then created based on the results of the simulation and the effects of
the spoilers on the performance output are tested with wind tunnel experiments at the open
jet facility of TU Delft (OJF). A parametric study was carried out to study the influence of
spoiler location on the performance. It is seen that the addition of spoilers to the existing
blade increases the performance of the wind turbine.
MSc. Thesis
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Preface
M.Sc. Thesis
Acknowledgement
Victoria Concordia Crescit.
I would like to start by thanking my supervisors Dr.C.J.Ferreira and Dr.G.Schepers. Without
your constant encouragement and support even when things were bleak, I never would have
been able to finish this thesis. I would like to thank ECN and everyone else involved for giving
me an opportunity to work on quite an interesting thesis topic and for guiding me when I
was stuck with a problem.
On a more personal note, I would like to express thank to all my friends I made over the past
two years and the old ones for creating beautiful memories. Finally a big thanks to my family
for being there when I needed you lot.
Ananth
MSc. Thesis
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Acknowledgement
M.Sc. Thesis
Table of Contents
Preface
vii
Acknowledgement
ix
List of Figures
xv
List of Tables
xix
Nomenclature
xxi
1 Introduction
1.1
1
Thesis Outline . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
2 Aim of the thesis
3
5
2.1
Research Questions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5
2.2
Thesis outline . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
5
3 Aerodynamics of Horizontal Axis Wind Turbine
7
3.1
Fundamentals of HAWTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
7
3.2
Modeling HAWTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
15
3.3
Approach . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
19
4 Experimental Technique: Particle Image Velocimetry
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Table of Contents
4.1
Working Principle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
21
4.2
Stereo Particle Induced Velocimetery . . . . . . . . . . . . . . . . . . . . . . . .
24
5 Initial Aerodynamic Investigation
31
5.1
Designing the Blade . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
31
5.2
Scaling based on Solidity Ratio . . . . . . . . . . . . . . . . . . . . . . . . . . .
31
5.3
Sectioning the Blade
. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
36
5.4
Tip Speed Ratio Scaling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
38
5.5
Effect of tip shapes on pressure distribution . . . . . . . . . . . . . . . . . . . .
39
5.6
Verification with Aero-Module . . . . . . . . . . . . . . . . . . . . . . . . . . .
41
5.7
Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
43
6 Blades and Experimental Set-Up
6.1
6.2
6.3
45
Blade Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
45
6.1.1
The OJF Tunnel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
47
Force Balance Experiment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
48
6.2.1
PIV Set Up . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
49
6.2.2
Spoilers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
50
Measurement Matrices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
51
7 Results
55
7.1
Loads Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
55
7.1.1
56
Thrust Coefficients . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
8 Uncertainty Analysis
65
8.1
Flow Uncertainities . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
65
8.2
Blade Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
67
8.3
Spoiler Position uncertainties . . . . . . . . . . . . . . . . . . . . . . . . . . . .
72
9 Recommendations and Future Work
73
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xiii
9.1
Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
73
9.2
Recommendations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
74
Bibliography
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xiv
Table of Contents
M.Sc. Thesis
List of Figures
1.1
Persian wind mill. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
1
1.2
Brush wind mill Shepherd (1990). . . . . . . . . . . . . . . . . . . . . . . . . .
2
1.3
An offshore wind farm in Denmark. . . . . . . . . . . . . . . . . . . . . . . . . .
3
3.1
Types of vortex generators. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
9
3.2
Hypothesis of flow around a Gurney flap. . . . . . . . . . . . . . . . . . . . . . .
10
3.3
Lift(Cl ) vs angle of attack(α) for different flap height as a function of chord c
Liebeck (1978). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
10
Drag (Cd ) vs angle of attack(α) for different flap height as a function of chord c
Liebeck (1978). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
10
Lift(Cl ) vs angle of attack(α) for different flap location as a function of chord c Li
et al. (2003). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
11
Drag(Cd ) vs angle of attack(α) for different flap location as a function of chord c
Li et al. (2003). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
11
3.7
Gurney flap mounting angle(Φ). Li et al. (2003). . . . . . . . . . . . . . . . . .
12
3.8
Lift(Cl ) vs angle of attack(α) for different flap mounting angle (Φ) Li et al. (2003). 12
3.9
Drag(Cd ) vs angle of attack(α) for different flap mounting angle (Φ) Li et al. (2003). 12
3.4
3.5
3.6
3.10 Airfoil geometries of FB3500-0050, FB3500-0875, FB353500-1750. . . . . . . . .
13
3.11 Visualization of streamlines around a S809 airfoil with a microtab of height1.1%c
at a location x/c = 0.95 Chow and Dam (2006). . . . . . . . . . . . . . . . . .
14
3.12 Effect of microtabs on aerodynamic properties Chow and Dam (2006). . . . . . .
14
3.13 Actuator disk with steam tube. . . . . . . . . . . . . . . . . . . . . . . . . . . .
15
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List of Figures
3.14 A)Schematic of blade elements(Top) B)Blade geometry for analysis [Manwell et al.
(2010)].(Bottom) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
16
3.15 A)Flow field model(Left) B)Wake geometry [Van Garrel (2003)](Right). . . . . .
17
3.16 Representation of airfoil section and wake. . . . . . . . . . . . . . . . . . . . . .
18
4.1
A sketch of a PIV-Set up. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
22
4.2
Cross-correlation map between two images taken at t and ∆t[F.Scarano (2013)].
23
4.3
Single-pass and multi pass techniques. . . . . . . . . . . . . . . . . . . . . . . .
24
4.4
Schematic for reconstruction of the three components of displacements. . . . . .
25
4.5
Angular displacement arrangement with Scheimpflug condition. . . . . . . . . . .
25
4.6
Sheimpflug condition. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
26
4.7
Perspective Effect of angle displacement arrangement.
. . . . . . . . . . . . . .
26
4.8
Reconstruction of Displacement components as give by Zhang (2013). . . . . . .
27
4.9
Control volume approach for determining integral forces as defined by Noca et al.
(1999) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
28
5.1
Solidity of a rotor blade. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
32
5.2
3 Blade ensemble generation in UMPM. . . . . . . . . . . . . . . . . . . . . . .
33
5.3
Pressure Distribution(Cp ) at different span wise locations of the wind turbine
blade(r/R) for three and two bladed turbine. . . . . . . . . . . . . . . . . . . . .
34
34
35
5.6
Axial force distribution along the span for three and two bladed turbine. . . . . .
35
5.7
blade(r/R) for 3 bladed, 2 bladed and a 20,25,30,35m model. . . . . . . . . . . .
37
37
38
5.4
5.5
5.8
5.9
M.Sc. Thesis
List of Figures
xvii
5.10 a) Flat tip b)Sweep back c)Sweep Forward plan view. . . . . . . . . . . . . . . .
39
5.11 Pressure Distribution(Cp ) at different span wise locations of the wind turbine
blade(r/R) for different tip shapes. . . . . . . . . . . . . . . . . . . . . . . . . .
40
40
41
5.14 Prandtl Correction Factor.
. . . . . . . . . . . . . . . . . . . . . . . . . . . . .
42
5.15 AWSM Correction Factor. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
43
6.1
Chord and Twist Distributions. . . . . . . . . . . . . . . . . . . . . . . . . . . .
45
6.2
Spanwise Reynolds Number distributions. . . . . . . . . . . . . . . . . . . . . .
46
6.3
Angle of attack distributions. . . . . . . . . . . . . . . . . . . . . . . . . . . . .
46
6.4
Wind tunnel model. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
47
6.5
Open Jet Facility. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
48
6.6
Convention for the force balance. . . . . . . . . . . . . . . . . . . . . . . . . . .
48
6.7
Stereoscopic Set Up. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
49
6.8
Drafting of the Spoiler1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
50
6.9
Drafting of the Spoiler 2.
. . . . . . . . . . . . . . . . . . . . . . . . . . . . .
50
6.10 Chord Wise Measurement Set Up [Akay et al. (2014)]. . . . . . . . . . . . . . .
52
7.1
Thrust coefficients (CT ) for various configurations for spoiler 1,2 as a function of
tip speed ratio (λ). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
57
7.2
Percentage change in thrust (CT ) Spoiler 1 as a function of tip speed ratio (λ).
58
7.3
Percentage change in thrust (CT ) Spoiler 2 as a function of tip speed ratio (λ).
59
7.4
Thrust coefficients(CT ) for various configurations for spoiler 1,2 as a function of
wind speed. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
61
7.5
Percentage change in Thrust(∆CT ) Spoiler 1 as a function of wind speed.
. . .
62
7.6
Percentage change in Thrust(∆CT ) Spoiler 1 as a function of wind speed.
. . .
63
MSc. Thesis
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8.1
List of Figures
Error plot for variation of thrust coefficients with respect to tip speed ratio for the
clean case. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
67
8.2
Axial force distribution over the span. . . . . . . . . . . . . . . . . . . . . . . .
68
8.3
blade(r/R) for linear and spline distributions for the root transition section. . . .
69
blade(r/R) for linear and spline distributions for the second transition section. . .
70
blade(r/R) for linear and spline distributions for the third transition section. . . .
71
8.4
8.5
M.Sc. Thesis
List of Tables
5.1
λ scaling.
. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
39
6.1
Spoiler 1 configurations. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
51
6.2
Spoiler 2 configurations. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
51
6.3
PIV Measurement Parameters . . . . . . . . . . . . . . . . . . . . . . . . . . . .
53
7.1
Spoiler 1,2 configurations. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .
55
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List of Tables
M.Sc. Thesis
Nomenclature
Abbreviations
AoA
BEM
CFD
ECN
FOV
HAWT
N-S
OJF
PIV
SPIV
TI
TSR
TTE
UMPM
Angle of Attack
Blade Element Method
Computational Fluid Dynamics
Energy research Centre of the Netherlands
Field Of View
Horizontal Axis Wind Turbine
Navier-Stokes
Open Jet Facility
Particle Induced Velocimetery
Stereo Particle Induced Velocimetery
Turbulence Intensity
Tip Speed Ratio
Thick Trailing Edge
TU Delft 3D unsteady panel code
Greek Symbols
α
λ
µ
Ω
ρ
Angle of Attack
Tip Speed Ratio
Kinematic Viscosity
Angular Velocity
Density
-[Degrees]
-[−]
2
-[m /s]
-[rad/s]
-[kg/m3
Latin Symbols
cD
cl
cP
cp
ct
f
Drag Coefficient
Lift Coefficient
Power Coefficient
Pressure Coefficient
Thrust Coefficient
Blade Rotational frequency
MSc. Thesis
-[−]
-[−]
-[−]
-[−]
-[−]
-[Hz]
xxii
p∞
Static Pressure
pdynamic Dynamic Pressure
ptotal Total Pressure
R
Universal Gas Constant
r
Blade Radius
T
Temperature
v∞
Free Stream Velocity
Nomenclature
-[P a]
-[P a]
-[P a]
-[J/(Kg.K)]
-[m]
-[K]
-[m/s]
M.Sc. Thesis
Chapter 1
Introduction
Wind has been a source of energy for mankind for a few centuries for purposes such as milling,
pumping water. Wind is also used to generate electricity by means of a wind turbine which
works on the principle of converting kinetic energy of the wind into electric energy. Wind
is a clean and renewable source of energy as it does not contribute to the environmentally
unfriendly phenomenon such as global warming, pollution as reported in Khan and Khan
(2013) . It is also a comparatively cheap form of energy which makes it an interesting option
for countries.
The earliest use of wind was the sail boat. Even though our ancestors were able to harness
the energy in the wind, they did not know how exactly it worked. The first windmills were
created for pumping water and grinding grains.
Figure 1.1: Persian wind mill.
Shepherd (1990) discusses the oldest design uncovered. This was the vertical axis system and
was built in Persia around 500-900 AD and is seen in Figure 1.1. This was used for grinding
MSc. Thesis
2
Introduction
grains. By the 11th century, these wind mills were being used extensively for food production.
These ideas were then bought to Europe by the trade merchants, the Dutch in particular who
modified the designs and used them for pumping water and draining lakes.
These discovery of electricity and the rapid industrialization of the world in the 19th and 20th
centuries lead to an increase in power demands and hence scientists started looking for new
methods of power generation. The first windmill to produce power was the Brush wind mill
which was built in 1888 in Ohio and is seen in Figure 1.2. This was a first of its kind as it
also included a gear box mechanism.
However, this windmill had some severe disadvantages as it was a high solidity rotor which
Figure 1.2: Brush wind mill Shepherd (1990).
was operating very slowly and was not very efficient. The first wind turbine which took into
account the aerodynamics of flow was built by Poul La Cour [Shepherd (1990)] who designed
a blade with a low solidity ratio and airfoil shapes. These were significantly faster than the
Brush wind mills and had a capacity of 25kW.
The development of wind turbine technology lay dormant until the ”oil crisis” in the 1970’s
when the US decided to research alternative sources of energy. Then spanned two decades of
work which laid the foundation for modern wind turbine technology. The formation of wind
energy research centers accelerated the growth of wind turbine technology with the blades
becoming more quieter, lighter and more efficient. Large scale power generation was possible
with the installation of wind farms like the one seen in Figure 1.3. Nowadays the average
capacity of a wind turbine is 4-5MW[Khan and Khan (2013)]. However as we look towards
the future, mega wind turbines capable of producing 10MW will be commonplace. Hence
the challenges in upscaling the turbines need to be tackled efficiently and innovative solutions
have to be researched. One of the challenges involved is controlling the flow over the blade.
This becomes especially difficult as future wind turbines will be very big with very thick root
sections and separation issues become significant. Hence flow control becomes imperative
M.Sc. Thesis
1.1 Thesis Outline
3
Figure 1.3: An offshore wind farm in Denmark.
when modeling and designing large wind turbines.
Most of previous research have been done by looking at the outboard sections/ tip parts of the
wind turbine blade as they have a stronger contribution to the power generation associated
with the wind turbine. Thus there have only been a few studies on the root sections of wind
turbines[Schubel and Crossley (2012)]. The root sections of the wind turbine produce in
comparison with the tip sections, lesser aerodynamic forces and moments which makes it an
area of study that has been neglected. Without knowledge of the complexity of the flow in
terms of three dimensionality and complexity, it is impossible to create aerodynamic models
of the flow over an entire blade section accurately. The effectiveness of flow control devices on
the root sections have not been studied as a result of lack of proper aerodynamic models. The
thesis focuses on understanding the flow on the root regions on wind turbines in the presence
of flow control devices and its impact on the performance output.
1.1
Thesis Outline
The flow in the root sections of a wind turbine blade is very complex as the flow is three
dimensional with rotational effects. The flow in the root sections have been studied with experiments and numerical methods as seen in Akay et al. (2013) where the flow around the root
section was modeled in order to study the 3D and rotational effects at the root. Even though
the root section of the turbine does not contribute massively to the aerodynamic loads, massive structural loads are present during the operation of the wind turbine. Without proper
knowledge of the flow characteristics in the root region, it is impossible to create a numerical
model capable of recreating the flow in the root section. Thus there are inaccuracies in the
numerical modeling of the flow in the root sections as there is very little information on the
root sections of the wind turbines.
There has also not been many studies on flow control devices on root sections of wind turbine
blades as the focus was on the part of the blade which contributes most to the power production ie. the tip sections. Even though the performance of passive and active flow control
MSc. Thesis
4
Introduction
devices such as gurney flaps, spoilers, vortex generators, plasma actuators, micro tabs have
been studied in detail for the tip sections, no such studies have been carried out for the root
sections.
It is essential that the physics behind the flow of the entire wind turbine blade is well known
and documented before an attempt can be made to model very high capacity wind turbines
(10+MW). The thesis attempts to address these requirements by simulating the flow in the
root sections of the blades by means of experiments and also looking at the effects of spoilers
on the root sections of the blade. From these experiments it is expected that a database
for flow control devices can be created using the results from the experiments and computational codes which can lead to the development of better numerical aerodynamic methods for
prediction of flow control device behavior and its impact on power generation.
M.Sc. Thesis
Chapter 2
Aim of the thesis
2.1
Research Questions
Study of existing literature did not show a lot of research on recreating the flow in the root
section of a wind turbine blade in a scaled model except for Akay et al. (2013). Hence, the
first step of the investigation is to obtain a model that can recreate the actual root flow
in the blade while not being entirely big when scaled down in order to maintain the three
dimensional effects that occur in the root regions. For this purpose, various numerical initial
aerodynamic investigation tools are used such as panel codes, blade element method (BEM)
codes and lifting line codes (AWSM ECN Aeromodule) are used. By changing the various
parameters of the wind turbine blade, an attempt is made to recreate the flow in the root
section to closely resemble an actual blade section in order to study the effectiveness of the
spoilers in an environment that resembles the real world conditions. Once a satisfactory
model is created, testing can be done in the wind tunnel. The model is studied at the open
jet facility(OJF) at the Delft University of Technology.
The flow around the blade is captured by using a high resolution three dimensional flow
measurement technique ”Stereo particle image velocimetry”(S-PIV), which will be able to
capture all the flow field around the blade sections with and without the spoilers. The flow
characteristics in the focused section are studied from the phase locked data obtained from
the experiments. The pressure data can be reconstructed from the velocity data which helps
in unraveling the aerodynamics of spoiler. The results of the experiments are used to answer
the motivation for carrying out the thesis.
2.2
Thesis outline
• Chapter 1 is an introduction to wind energy and its potential for power generation.
The chapter deals with the history and development of wind power over the last few
centuries while also discussing the potential challenges that have to be tackled in future
MSc. Thesis
6
Aim of the thesis
large scale wind turbines. The chapter also discusses a brief overview of the thesis in
terms of what is expected from the results.
• Chapter 2 is an overview of the research questions that could possibly be answered
over the course of the report.
• Chapter 3 is an introduction into the aerodynamics of horizontal axis wind turbines
and flow control devices. The chapter also deals with modeling HAWTS while listing
out the advantages and disadvantages of the possible methods which could potentially
be used for modeling.
• Chapter 4 consists of an introduction to stereo particle image velocimetry which is the
experimental method used.
• Chapter 5 deals with the initial aerodynamic investigations which are carried out using
the unsteady panel code, BEM code and the AWSM(lifting line+free vortex wake) code.
• Chapter 6 consists information regarding the wind turbine model and its blade characteristics and the experimental set up.
• Chapter 7 details the results from the experimental data from which conclusions are
drawn.
• Chapter 8 discusses the uncertainties associated with carrying out experiments and
its impact on the results.
• Chapter 9 details the conclusions of the thesis and also recommendations.
M.Sc. Thesis
Chapter 3
Aerodynamics of Horizontal Axis Wind
Turbine
3.1
Fundamentals of HAWTS
The aerodynamics of HAWTS are in most cases similar to the aerodynamics of fixed wing
aircraft’s. However the rotation of the blades introduces some effects which leads to complex
flow characteristics. An introduction to the aerodynamics of horizontal axis wind turbines
has been made in this section. Horizontal axis wind turbines are lift driven devices in which
the turbine rotational axis is parallel to the ground. The wind turbine converts the kinetic
energy present in the air into mechanical energy which is converted into electrical energy by
using generators. The maximum amount of power that can be theoretically extracted from
air can be calculated as:
P = 12 ρAV 3
Where, P is the power, A is the swept area of the blade, ρ is the density of air and V is
the velocity of air. However there exists a theoretical limit to the amount of energy that
can be extracted that is independent of the design of the wind turbine. Maximum power
indicates the entire kinetic energy of air is harvested but this cannot be achieved in real life.
The maximum power that can be extracted from a wind turbine is determined by the ”Betz
limit” that states that the efficiency of the turbine cannot exceed 59.3 % as stated by Burton
et al. (2001). However, in practice there are losses due to tip effects, wake losses, drive train
and blade shape. Modern wind turbines with 3 blades are capable of having an efficiency of
50% but this is dependent on the design of the blade and the operating conditions as stated
in Schubel and Crossley (2012).
An airfoil section under the influence of rotation and an incoming wind speed produces
thrust and torque forces. Both of these forces contribute to the thrust and torque output of
MSc. Thesis
8
Aerodynamics of Horizontal Axis Wind Turbine
the wind turbine. The sectional lift and drag forces acting on a blade are the functions of
the input relative velocity and the airfoil profile at that section and also the angle of attack
which contribute to the thrust and the torque. Each airfoil section has a different angle of
attack and effective velocity due to rotation. The root sections are thicker to withstand the
structural forces and the tip airfoils are thinner as they have a lesser structural loading and
high velocity.
The flow in the root region is very complex to model owing to the presence of high angles of
attack, three dimensional flow effects and Coriolis and centrifugal forces. These effects have
an impact on the boundary layer by making the flow faster and thinner and hence affect
the performance of the root region. This is widely known as stall delay which was discussed
in Schreck et al. (2007). The near wake is right behind the wind turbine and can be said
to extend two or three rotor diameters. This region is characterized by 3-D effects, nacelle
effects, tip vortices’s etc and is dependent on the design of the rotor as stated in Sanderse
(2009).
Being able to control the flow around a wind turbine blade holds a important position in
wind turbine aerodynamics as its application is capable of increasing the lift force with a
lower aerodynamic drag. This is of interest as the power generated by a wind turbine is
directly proportional to the square of the blade length but increasing the length leads to
higher mass of the blade which hold many structural challenges as discussed in Johnson and
Berg (2008). Hence active and passive flow control techniques are applied on a wind turbines.
Passive flow control has been used extensively on wind turbines for flow control purposes.
Passive flow control devices are quite useful in regulating the flow to optimum conditions for
maximum efficiency of the wind turbine [Schubel and Crossley (2012), Troldborg et al. (2013)].
Passive flow control devices
Vortex Generators
Investigation of passive flow control devices on wind turbines have been performed before
with some of the devices tested being vortex generators, gurney flaps etc. Vortex generators prevent transition of flow by mixing low momentum boundary layer flow with high
momentum free steam flow thus energizing the boundary layer. Vortex generators are small
devices which are attached at an angle to the suction side of the blade. They generate
longitudinal vortices which are responsible for mixing and regeneration of the boundary layer.
The effect of vortex generators have been tested extensively for different flow conditions
and positions in Troldborg et al. (2013). Properly designed vortex generators as seen in
Figure 3.1 are capable of producing coherent helical vortex structures that help in mixing
between the free stream and the boundary layer. Hence they are able to control flow
separation and increase the maximum lift coefficient Clmax but have a adverse effect on
the drag even at attached conditions as their effect on power generation is global. It
was observed that the effectiveness of the vortex generator was dependent on the position,
M.Sc. Thesis
3.1 Fundamentals of HAWTS
9
size, flow properties and the Reynolds number of the flow as seen in Johnson and Berg (2008).
Figure 3.1: Types of vortex generators.
Gurney Flaps
The gurney flap was introduced race car driver Dan Gurney in 1960 when he placed a right
angled sheet metal to the trailing edge of the rear wing of his car which has been explained
in Wang et al. (2008). He discovered that the presence of this device increased the maximum
down force thus enabling him to take corners faster and achieve higher speeds at straights.
More tests were performed by Liebeck (1978) on a Newman airfoil where it was stated that
the lift from airfoil was higher with a flap of height 1.25% of the chord placed at the trailing
edge of the pressure side. Liebeck (1978) also discusses scaling for optimum gurney flap height
and positioning for optimum lift to drag ratio.
For maximum efficiency if the flap, it must be placed at the trailing edge with a height scaled
with the height of the boundary layer at the local position. The gurney flap operates on the
basis of increasing the pressure on the pressure side, reducing pressure on the suction side
[Giguere et al. (1995)Nikoueeyan et al. (2014)] which is represented in Figure 3.2. The wake
downstream of the flap consists of a pair of counter rotating vortices which delay the flow
separation at the trailing edge suction side. This leads to an increase in suction and lift.
Liebeck (1978) discusses the effect of the flap height on the lift polars. It is seen that the
maximum lift coefficient(Cl ) increases with increase in the height of the flap along with an
increase in the zero lift angle of attack(α) as seen in Figure 3.3.
The abrupt stall is due to the bursting of the leading edge separation bubbles instead of
gradual flow separation from the trailing edge. Looking at Figure 3.4, it is seen that after a
flap height of 3% of the chord, there is a substantial drag increase which is in agreement with
Liebeck (1978) who hypothesized that the flap height should be scaled with the boundary
layer. This was also validated by Li et al. (2003) who conducted experiments to study the
effects of flap height on the lift polars.
Li et al. (2003) discusses the effects of flap location on the lift and drag polar. Experiments
MSc. Thesis
10
Figure 3.2: Hypothesis of flow around a Gurney flap.
Figure 3.3: Lift(Cl ) vs angle of attack(α) for different flap height as a function of chord c Liebeck
(1978).
Figure 3.4: Drag (Cd ) vs angle of attack(α) for different flap height as a function of chord c
Liebeck (1978).
were conducted in a NACA 0012 airfoil with a gurney flap of height 1.5 % of the chord. The
flap was placed at 0,2,4,6 % of the chord away from the trailing edge and the effects on the
performance were studied.
M.Sc. Thesis
11
Figure 3.5: Lift(Cl ) vs angle of attack(α) for different flap location as a function of chord c Li
et al. (2003).
Figure 3.6: Drag(Cd ) vs angle of attack(α) for different flap location as a function of chord c Li
et al. (2003).
The variation of lift and drag due to the location of the gurney flap are seen in Figure 3.6 and
Figure 3.5. It is seen that the increase in lift decreased when the flap is moved away from
the trailing edge. Even though all the configurations performed better than a plain airfoil,
the maximum increase in lift of 17.4 % was seen only at the trailing edge. It is also seen that
when the flap is mounted closer to the trailing edge, it increases the apparent trailing edge
thickness of the airfoil which increases the drag. This is seen in the sharp increase in drag at
an angle of attack of 2 to 10 when the flap is mounted 4-6 % away from the trailing edge.
The angle at which the gurney flap is attached to the airfoil also has an impact on the
effectiveness of the flap. This was studied by Li et al. (2003) who studied the effect of
mounting angles on a 1.5 %c gurney flap on a NACA 0012 airfoil. Mounting angle is defined
as the angle between the chord line and the flap. This is seen in Figure 3.7.
MSc. Thesis
12
Figure 3.7: Gurney flap mounting angle(Φ). Li et al. (2003).
Figure 3.8: Lift(Cl ) vs angle of attack(α) for different flap mounting angle (Φ) Li et al. (2003).
Figure 3.9: Drag(Cd ) vs angle of attack(α) for different flap mounting angle (Φ) Li et al. (2003).
The variation of lift coefficient versus angle of attack for different mounting angles is seen
in Figure 3.8. The flap was mounted at 45, 60 and 90 degrees and it was seen that all
configurations performed better than the clean blade. The increase in lift was higher
for configurations with higher mounting angles. The drag polar are seen in Figure 3.9
and it is seen that the drag penalty can be reduced by reducing the mounting angle. It
can be concluded that more inclined gurney flaps work better as they reduce the drag penalty.
M.Sc. Thesis
13
Thick Trailing Edge Airfoils
Thick trailing edge(TTE) airfoils can be used to address the aerodynamic flow problems
associated with thick airfoils at subsonic conditions. A normal airfoil ends with a finite point
at the trailing edge where the upper and lower sides meet. But in the case of a wind turbine,
the root sections are larger as they have to bear the structural loads. Usually, the airfoils in
this section are pretty thick usually 35 % of the chord at the position of maximum thickness.
Making a sharp trailing edge leads to a sharp taper which creates a adverse pressure gradient
which makes the airfoil susceptible to stall as the boundary layer is very sensitive. Hence
making the trailing edge thicker reduces the possibility of flow seperation over the airfoil.
Example of a thick trailing edge airfoil is shown in Figure 3.10.
Figure 3.10: Airfoil geometries of FB3500-0050, FB3500-0875, FB353500-1750.
Also thick trailing edge airfoils are used on wind turbines as they are much more resistant to
fouling and contamination of blade leading edges[Standish and Van Dam (2003) Baker et al.
(2006)]. In addition to this, they also increase the lift coefficient of the airfoil and allows for
thicker airfoil sections which can take more structural loads when compared to normal airfoils
and also leads to a more linear lift polar with an increase in the lift to drag ratio (L/D). The
only constraint associated with trailing edge airfoils are increased base drag and more noise
production which can be minimized by proper design of the trailing edge.
Microtabs
Micro tabs are used to generate macro changes in flow properties by using micro structures
which act like add-ons to the existing blade geometry as seen in Figure 3.11 and discussed in
Chow and Dam (2006). These devices are placed at the trailing edge of the airfoil and are
controllable and work by changing the sectional camber and the trailing edge flow conditions
MSc. Thesis
14
[Chow and Van Dam (2007)]. The height of the device is in the order of the boundary layer
thickness. The micro tab increases the lift produced by the airfoil changing the sectional
camber thus affecting the circulation around the airfoil [Chow and Dam (2006)]. It is also
seen that the lift increase obtained by the micro tab is directly dependent on the solidity ratio
of the device Mayda et al. (2005) and also listed out at Figure 3.12.
Figure 3.11: Visualization of streamlines around a S809 airfoil with a microtab of height1.1%c
at a location x/c = 0.95 Chow and Dam (2006).
Figure 3.12: Effect of microtabs on aerodynamic properties Chow and Dam (2006).
M.Sc. Thesis
3.2 Modeling HAWTS
3.2
15
Modeling HAWTS
Several aerodynamic models can be used to investigate the aerodynamics of horizontal axis
wind turbines. The main investigation models are:
1. Actuator disk theory.
2. Blade element momentum theory(BEM).
3. Lifting Line and Vortex Wake modeling.
4. Panel Method.
5. Computational fluid dynamics(CFD).
All these models differ from each other in terms of complexity, computational time, ability
of the model to capture flow characteristics and so on. Hence a clear study of all the models
have to be done in order to choose a suitable simulation method.
Actuator disk theory
Figure 3.13: Actuator disk with steam tube.
This is a theoretical modeling tool which is used to model the near wake of a wind turbine
rotor. In this model the rotor is created as a surface which is perpendicular to the flow on
which the forces are distributed uniformly as referred in Glauert (1935).
Looking at the Figure 3.13, it can be seen that the actuator disk slows down the air in front
of the turbine while there is a loss of static pressure behind the rotor. The disk applies a
force on the flow [Sørensen and Kock (1995). The velocity and pressure of the air behind
the disk is lower than the velocity and pressure in front of the disk. At the position of the
disk, there is energy extraction from the flow in the form of force. Looking at the Figure 3.13
the pressure difference between the two sides of the disk are seen clearly. By applying the
conservation of mass, momentum and energy the rotor forces can be calculated in the form
of integrated values. The velocity of air at the actuator disk can be calculated as:
MSc. Thesis
16
Udisk = Uinf (1 − a)
Where a is the axial induction factor. The velocity at the wake can be written as.
UW ake = Uinf (1 − 2a)
It is seen that if the value of the axial induction factor goes beyond 0.5, then the velocity
in the wake becomes negative which is impossible. For the purpose of studying the effect of
spoilers on the flow on the blade and the near wake, this method is not suitable as it does
not provide the distribution of loads over the blade or the wake.
Blade Element Momentum(BEM) Methods
Figure 3.14: A)Schematic of blade elements(Top) B)Blade geometry for analysis [Manwell et al.
(2010)].(Bottom)
M.Sc. Thesis
3.2 Modeling HAWTS
17
BEM theory provides very fast results with reasonable accuracy and hence it is a widely
accepted technique. It makes use of the assumption that there is no aerodynamic interaction
between the blade elements and the forces on the blade are determined only by the lift
and drag characteristics of the airfoil shape.The lift and drag forces on the blade as seen in
Figure 3.14 can be calculated using BEM as these forces are responsible for the change of axial
and angular momentum of the air in the swept region of the blade [Burton et al. (2001)].
In this method, the blade is divided into several sections and the forces on the sections
are calculated by using two dimensional airfoil characteristics. Then the values of induced
velocities and the angle of attack are calculated. The process is iterative and is repeated till
convergence. The model is simple and easy to implement. The model is also unable to predict
3-D flow, unsteady effects and misalignment of the blades even though some accuracy can be
achieved by using tip correction factors of Prandtl and Glauert as discussed in Burton et al.
(2001). Hence BEM can be used to calculate the forces over the blade very fast but fails in
mapping the flow characteristics over the blade, dynamic stall characteristics and the near
wake .
Lifting Line and Vortex Wake modeling
Figure 3.15: A)Flow field model(Left) B)Wake geometry [Van Garrel (2003)](Right).
These methods are based on the assumptions that the flow is in compressible and inviscid.
They make use of vorticity transport and Biot-Savart equations to model the wake and its
influence on the turbine blades. The sectional lift and drag are calculated from the inflow
conditions. The aerodynamic wind turbine simulation module(AWSM) was created by Van
Garrel, [Van Garrel (2003)] is one of the codes based on a lifting line method with a free vortex
wake by making the assumption that the extension of the blade in the span wise direction is
more significant compared to the chord or thickness distribution of the wind turbine blade.
Hence the airfoils at various sections are replaced by a line passing through the quarter chord
positions of the blade as seen in Figure 3.15. The induced velocities on the blade can be
calculated using the Biot-Savart law.The shed vorticity at each time step is calculated using
the Kutta theorem. The shed vorticity, in the form of vortex rings, together forms the wake
of the rotor-blade. The model can provide the shape of the near wake of the turbine which is
a big advantage in the usage of this model.
MSc. Thesis
18
Panel Methods
The panel methods were first developed by Hess and Smith Erickson (1990) at ”The Douglas
Aircraft Company” in the development of an aircraft. Panel codes are able to accurately
analyze the 3-D flow around geometries which is not possible with other codes. These were
later replaced with CFD simulations due to the advancement of computers and because
of the reason that CFD codes are applicable to any flow problem. Panel methods can be
described as an extension of vortex method because they treat the wake in the same way as
that of vortex methods but do not need a table for Cl andCd data as seen in Erickson (1990).
They are able to model the geometry directly by solving the Laplace equation for inviscid
incompressible flows.
The blade surface is replaced by a number of discrete panels as seen in Figure 3.16. This
Figure 3.16: Representation of airfoil section and wake.
leads to a more accurate representation of the geometry of the blade.The various formulations
of panel codes have been explained in Johnson (1980) and Erickson (1990). Some of them
being the flow tangency at the surface, the distribution of panels, the implementation of
the Kutta condition, the discretization of the wake. However, the biggest disadvantage of
the panel code is that they exclude the effect of viscosity. Hence they are unable to handle
separated flow.
Computational Fluid Dynamics
The application of computational fluid dynamics(CFD) solvers on wind turbine blades is
widespread due to the increase in computational power. The exact flow characteristics and all
flow phenomenon that happen on the blade can be captured by solving the Navier-Stokes(NS) equations which describe the momentum, viscosity, pressure in the fluid. Mostly, the
equations are used in the form of the Reynolds Averaged Navier-Stokes (RANS) form. Flow
around a wind turbine is mostly incompressible with some compressibility effects happening
in the tip sections due to the high angular velocities present.
In a RANS solver, only the mean flow is solved around the mesh while the turbulence is
M.Sc. Thesis
3.3 Approach
19
recreated by using closure models. These closure models may range from simple algebraic
models such as the one equation models and two equation models like the K- ω models. CFD
solvers are also capable of calculating the flows and forces on a wind turbine blade with flow
control devices and add-ons. CFD solvers are able to capture all the flow phenomenon that
occur and help in obtaining a detailed understanding of the aerodynamics of a blade but are
computationally very intensive.
3.3
Approach
It is seen that the actuator disk model is not useful when it comes to calculating the forces
on the blade or the near wake structure as it is not capable of predicting these effects. The
blade element momentum method can be used to quickly calculate the distribution of the
forces over the blade while the shape of the near wake can be calculated by using the vortex
models. The panel code can be used to calculate the pressure distributions over the blade
sections as it uses the actual shape of the blade instead of approximating them unlike the
lifting line methods while also being capable of calculating the forces due to the presence of
flow control devices. CFD can be used to accurately determine the flow structure around the
blade to visualize the effects of the spoilers. But as time is a constraint, these will be carried
out only when time is available.
BEM code will be used in the preliminary design of the blade and the variation of the induced
velocities due to sectioning and other designs are simulated by using the lifting line code. The
pressure distributions around the different span sections of the blade are predicted by using
the unsteady panel method. By making use of these modeling tools it is now possible to
create an appropriate wind turbine model for testing.
MSc. Thesis
20
M.Sc. Thesis
Chapter 4
Experimental Technique: Particle Image
Velocimetry
4.1
Working Principle
Particle image velocimetry(PIV) is one of the most successful flow measurement techniques to
have emerged in the past decade. The technique behind PIV is measuring the displacement of
tracer particles which are injected in the flow between two small time intervals. The particles
have to be small enough so that they do not interfere with the fluid flow properties. The flow
is then illuminated by a thin light sheet from a pulsating light source usually a laser system.
An imaging device(camera) is then set up perpendicular to the measurement plane to capture
the light in two consecutive image frames. Processing these images gives the displacements
between the various tracer particles from which the velocity can be calculated using the time
difference between the two frames. PIV is capable of capturing any kind of flow as long as
the fluid itself is transparent to allow imaging of the tracer particles.
PIV Set Up
The typical PIV set-up for experiments has been shown in Figure 4.1. The main components
of the PIV system are the laser source, seeding particles and the imaging camera [Adrian
(2005)]. These are explained in the following section.
Seeding Particles
Microscopic particles are used to seed the flow for the purpose of visualization. The most
important aspect of the seeding particles are that it should not interfere with the flow properties. Hence the particles have to be buoyant and small with respect to the fluid [Raffel
MSc. Thesis
22
Experimental Technique: Particle Image Velocimetry
Figure 4.1: A sketch of a PIV-Set up.
et al. (2002)]. The size of the seeding particles are around 5 to 200 µm. The particles have
to be highly reflective in order to yield good particle images. The ability of the particle to
scatter and reflect light is highly dependent upon the diameter and the refractive index of the
particle.
Light Source
Illumination is usually provided by means of a laser sheet. The main requirement of a PIV
imaging system is that the pulse duration δt should be as small as possible. The laser
sheet is used to define the area of interrogation. Laser is used as it is capable of delivering
a thin bright sheet of light that is coherent and monochromatic without any diffusion as
discussed in Stamhuis (2006). The laser sheet is aligned with the flow direction to maintain
the same particles in the interrogation window when the pictures are taken. Both pulsating
and continuous laser sheets can be used for illumination. Pulsed lasers can provide higher
levels of illuminations in very short intervals and hence are more suitable for high speed
experiments.
Imaging
A camera is set up with its optical axis perpendicular to the illuminated plane.The quality of
the lenses have to be good enough to capture the light levels and the particles. charge coupled
device(CCD) or complementary metal-oxide semiconductor(CMOS) devices are capable of
delivering images real time and are used to analyze the images before experiments [Willert
and Gharib (1991)]. The efficiency of the imaging system is dependent on the focal length, f
number, spatial resolution, recording medium and the image magnification factor [F.Scarano
(2013)]. The magnification is dependent on the image and object diameter.
M.Sc. Thesis
4.1 Working Principle
23
PIV Image Post Processing
The images obtained from the camera has to be processed in order to obtain the velocity
fields. To achieve this, the images are divided into smaller regions called interrogation zones
or windows. Then cross correlation analysis is carried out on these windows. Cross correlation is the measure of similarity of two series (in this case the images). By comparing
the two images, a correlation peak can be calculated. This peak corresponds to the particle
motion between the two images which is a measure of displacement. The velocity can then be
calculated by dividing the displacement with the time interval between the two laser pulses.
This is shown in Figure 4.2. Good cross-correlation demands that the particles remain in the
Figure 4.2: Cross-correlation map between two images taken at t and ∆t[F.Scarano (2013)].
interrogation window during both time frames. If the correlation size is too small, then there
is an error induced due to wrong velocity vectors while larger values of correlation windows
will lead to a lower resolution of the flow characteristics.
Overlapping can be done to increase the quality of the data coming from processing. Overlapping allows for two interrogation windows to overlap. As the particles near the edges of
the interrogation windows do not usually have pairs, overlapping allows them to recover the
lost particles thus helping in determining the displacement of that particle. Adrian (2005)
define usual overlap as about 50 %.
A small interrogation window leads to particle loss while a large window is not able to map
the flow details. To further increase the accuracy of the velocity vectors, multi pass technique is used which is performed by analyzing the flow with multiple passes with the size of
the interrogation window decreasing every pass. The windows are displaced by the average
displacement of the particle during the passes. By performing multi pass operations, the
calculation of large displacement vectors are made possible.
Figure 4.3 shows the difference between single and multi pass techniques. The multi pass
technique is able to resolve the flow field more clearly than the single pass operation. The
effect of noise can be removed by using filtering techniques which are explained in Zhang
(2013).
MSc. Thesis
24
Figure 4.3: Single-pass and multi pass techniques.
4.2
Stereo Particle Induced Velocimetery
Stereo particle induced velocimetery(S-PIV) is capable of capturing the out of plane
displacements. This is achieved by using by using two cameras to measure the velocity in
the third plane instead of a single one. Both cameras are capable of capturing sufficient
information to calculate the out of plane motion of the particles. The two views are then
combined by using various algorithms to recreate the three dimensional field (3 component
velocities). [Prasad and Adrian (1993)].
Principle
A stereo-PIV setup makes use of two cameras to capture the out of plane displacements in
the measurement plane as explained previously. The schematic for reconstruction of the three
components of displacements is shown in Figure 4.4. The point O is a known point on the
measurement plane which is visible to both of the cameras. The x and y axes are along the
measurement plane while the z axis is perpendicular to the measurement plane. The cameras
are placed at the points L1 and L2. The angles α1 and α2 refer to the angle between the
camera viewing ray and the light sheet normal direction. β1 and β2 refer to the angle created
with the Y − Z plane. The values of the displacement vector δα and δβ are neglected as the
viewing distance is larger than the displacement vector. The displacement vectors can then
be calculated by the following equations as given in Zhang (2013).:
dx =
dx2 ·tanα1 −dx1 ·tanα2
,
tanα1 −tanα2
dy =
dy2 ·tanβ1 −dy1 ·tanβ2
,
tanβ1 −tanβ2 )
dz =
dy2 −dy1
tanβ1 −tanβ2 .
M.Sc. Thesis
4.2 Stereo Particle Induced Velocimetery
25
Figure 4.4: Schematic for reconstruction of the three components of displacements.
If the viewing axes become collinear in either of the 2-D projections, then the numerator
approaches zero. Hence the angles of β1 and β2 become very small so the equation of dy can
be rewritten as:
dy =
(dy1 +dy2 )
2
+
(dx2 +dx1 )
2
tanβ2 −tanβ1
· ( tanα
).
1 −tanα2
Set Up
The stereo PIV system can be set up in three different ways: the lens translation method,
the general angle displacement method and the angle displacement with the Scheimpflug
condition [Prasad and Jensen (1995)]. The 4.5 shows two cameras that are non parallel with
Figure 4.5: Angular displacement arrangement with Scheimpflug condition.
MSc. Thesis
26
their axes intersecting at the image plane. The plane perpendicular to the image plane and
the cameras are set at a particular angle. The larger this angle, the higher is the accuracy
of the out of plane vector but might lead to image distortion. This set up also makes use
of the Scheimpflug condition that can overcome the disadvantage of a normal displacement
configuration and provide focus for the entire measurement window.
As the cameras are tilted, the object plane and the lens plane are no longer parallel. Hence
Figure 4.6: Sheimpflug condition.
the image plane has to be rotated by an angle so that the image plane, the lens plane and
the image plane lie on the same point. This is called as Scheimpflug condition. This is shown
in Figure 4.6. The inclination angle is usually around 30-50 degrees as seen in Prasad and
Adrian (1993). Higher angles increase the accuracy of out of plane displacements but affect
the magnification over the measurement region and leads to distortion which can be corrected
using adapters.
Image Reconstruction
Figure 4.7: Perspective Effect of angle displacement arrangement.
The rectangular grid on the measurement plane becomes trapezoidal when looked at with
the two cameras as seen in Figure 4.7. Hence a calibration procedure is needed to determine
the mapping function between the measurement plane and the image plane. This calibration
helps to avoid and correct the distortions of the measurements. Calibration is done by placing
a target in the object plane. The target is a plate with dots placed on a Cartesian grid and is
captured at several locations along the thickness of the light sheet. It can also be seen that
the accuracy of the out-of-plane displacement is inversely proportional to the viewing angle.
σδz
σδx
=
1
tanα .
M.Sc. Thesis
27
Where σδz and σδx are the errors of the out of plane and in plane components. The process
Figure 4.8: Reconstruction of Displacement components as give by Zhang (2013).
of reconstructing the displacement components are described in Figure 4.8. The higher the
order of the mapping function, the more accurate the reconstruction of the displacements will
be as higher order reconstruction methods take into account the image non linearities and
distortion effects [Prasad and Adrian (1993)].
Pressure and Forces from PIV
The pressure field can be calculated by integrating the 2-D steady N-S equations if the flow is
nearly two dimensional, steady and incompressible as explained in Imaichi and Ohmi (1983).
2
∂2U
),
∂Y 2
2
∂2V
).
∂Y 2
∂U
U ∂X
+V
∂U
∂Y
=
−1 ∂p
ρ ∂X
∂ U
+ ν( ∂X
2 +
∂V
U ∂X
+V
∂V
∂Y
=
−1 ∂p
ρ ∂Y
∂ V
+ ν( ∂X
2 +
A more accurate representation of the pressure and forces can be obtained by using the control
volume approach as suggested by Noca et al. (1999) where closed form expressions are defined
for the evaluation of time dependent forces on a body for incompressible, viscous rotational
flow. Forces on the blade originate due to the pressure distribution and the shear stresses over
the blade. These can be calculated by measuring the change in momentum inside the control
MSc. Thesis
28
Figure 4.9: Control volume approach for determining integral forces as defined by Noca et al.
(1999) .
volume. The force inside by the flow on the airfoil can be calculated using the following
equation from Campo et al. (2014).:
d
F~f low−>airf oil = −
dt
ZZZ
ZZ
ZZ
ZZ
¯0
~
~r · ~n)V
~r ds−
ρVr dν −
p~nds +
τ ~nds −
ρ(V
v
S
S
s
ZZZ
ZZZ
~
~
~ × (Ω
~ × ~r))dv
2ρ(Ω × Vr )dv −
ρ(Ω
v
v
If a stationary frame of reference is used, there exists a time derivative in the equation. Hence
a moving reference frame is chosen to remove the time derivative from the equation because
the flow characteristics will not be changing with respect to the moving frame. Coriolis and
centrifugal forces are added to the equation due to changing the frame of reference. By
solving the resulting equation the forces can be obtained directly from the PIV data. The
pressure data can also be calculated the same way. The pressure data is calculated from the
momentum equation in the differential form using the relative velocity field and its derivatives
as seen in,
5p =
d
~
dt (ρVr )
~ r · 5V
~r + 2ρ(Ω
~ × (Ω
~ × ~r) + µ∆V
~r .
− ρV
Inertial forces, Reynolds stresses and the assumption that the flow is in compressible is
considered when deriving the above equation. From the pressure gradient data, it is possible
to calculate the pressure at the points by making use of several space marching algorithms
algorithms. But there happens a propagation of errors with the marching. To minimize
this error, the pressure data is calculated from the equation using a forcing function g(u, v)
derived from the momentum equations as seen in equation 4.2.
M.Sc. Thesis
29
52 p ≈ Dρ = g(u, v) → p = D−1 g(u, v).
The forcing function is defined as,
pi+1,j −2pi,j +pi−1,j
∆x2
+
pi,j+1 −2pi,j +pi,j−1
∆y 2
= Dρ
i,j=1...n,1...m
Conclusions
Stereo-PIV can be performed on the OJF rotor to study the origins of the hypothesized
increase in performance. Stereo PIV is capable of capturing the out of plane velocities as
well as the in plane velocities from which the vorticity distribution around the blade can
be determined as well as the pressure distributions. Loads data can be obtained from the
pressure distributions around the blade to visualize the sectional increase in lift due to the
presence of spoilers so that comparison with the clean case can be performed.
MSc. Thesis
30
M.Sc. Thesis
Chapter 5
Initial Aerodynamic Investigation
5.1
Designing the Blade
The idea behind the experiment was to take a full scale wind turbine blade and then sectioning
by removing the outboard section of the blade. This is done so that when the smaller model
is scaled, the chord values at the root sections of the blade are larger than when the entire full
blade gets scaled and hence larger Reynolds numbers whcih are more closer to representing
the real flow are present in the spans of the root region (r/R=0.3). It was expected that by
cutting off the blade at a particular span section, the inboard flow characteristics of the blade
will not be affected by any span-wise circulation and flow issues. Other tip designs were also
taken into account for the study to maintain the same flow properties in the inboard sections
of the blade such as winglets, tapered tips.
These designs are first simulated in the panel code after which the results are compared with
the results from the lifting line code. The Nordex N80 turbine was taken as the base model
for the experiment. The model can then be sectioned and then scaled and converted to a two
bladed model by changing the scaling using the solidity ratio and the tip speed ratio. The
pressure distributions along the span of the blade and the induced velocities are obtained from
the panel code and the lifting line code and are used as a measure of comparison between
the different models. This chapter explains the method of scaling the model based on various
parameters and studying the pressure distributions and the induced velocities.
5.2
Scaling based on Solidity Ratio
The original design of the blade consisted of three blades of 40m radius each. However this
poses a problem during the measurement campaigns as it is harder to conduct phase locked
PIV studies on 3 bladed rotors and that the rotor head at the OJF can only support two
blades. Hence the model has to be scaled from a three bladed model to a two bladed model.
MSc. Thesis
32
This was achieved by keeping the solidity ratio of the rotor constant. The solidity ratio is the
ratio of the total rotor platform area to the total swept area of the blade.
Figure 5.1: Solidity of a rotor blade.
For a three bladed rotor, the solidity ratio can be written as as,
S=
3·a
A ,
For a two bladed rotor, the solidity ratio becomes,
S=
2·a
A ,
Where a is the area of the blade and A is the total swept area. From this equation, it is seen
that to maintain the same solidity ratio when downsizing from 3 to 2 blades, the area of the
blade has to go up by a factor of 1.5. Hence the chord of the new blade is 1.5 times higher
than the 3 bladed turbine.
M.Sc. Thesis
5.2 Scaling based on Solidity Ratio
33
Validation from UMPM
In order to verify the proper scaling of the blade, simulations are done using the panel
code(UMPM) from TU Delft. This code is a 3D, multibody, freewake panel code created in
matlab. More information regarding the capabilities of the code can be found in K.R.Dixon
(2008). The input to the code consists of a column matrix where the twist, chord and span
are specified. The number of blades are specified in the ”ensemble.m” file. The simulation
parameters are better listed out and explained in the appendix. 50 chordwise and 52 spanwise
panels are used for the geometry generation. Figure 5.2 shows the geometry generation
Figure 5.2: 3 Blade ensemble generation in UMPM.
along with the ensemble geometry generation. The scaling of solidity ratio is validated from
the pressure distributions at different span sections of the 3 bladed and 2 bladed turbine
blades as obtained from the simulations of the unsteady panel method(UMPM) as seen in
Figure 5.3, Figure 5.4, Figure 5.5. The simulations were done for the design tip speed ratio
of 6.8 which corresponds to a input velocity of 10 m/s and a rotational speed of 16.403 rpm.
It is seen that the pressure distributions at the different span sections are a little different
between the two and three bladed turbine. Looking at Figure 5.3,which shows the pressure
distributions at the inner sections(r/R=0.07 to 0.185), it is seen that the distributions are
skewed. This is because of the fact that as the code is inviscid, the code has problems solving
the flow around the almost cylindrical root sections. It is seen that at the thicker airfoil
sections with the blunt trailing edge, the kutta condition is not enforced properly which can
be seen clearly from the trailing edge pressure distributions which do not meet at a single
point. However the outboard pressure distributions are correct because of the relatively
thinner airfoils and sharp trailing edges. The maximum Cp for the two bladed turbine is
lower than the three bladed turbine. As the two bladed turbine has a longer chord because
of scaling, the flow over the blade leading edge takes a longer time to accelerate because of
the smaller gradient change over the airfoil chord. However this difference is negligible as the
shape of the pressure distribution still remains the same thus leading to the smaller peaks of
Cp .
MSc. Thesis
34
Cp Section 1 r/R=0.038
6
6
3 Bladed
2 Bladed
2
2
0
0
-2
-2
0
0.5
3 Bladed
2 Bladed
4
Cp
4
1
0
Normalized Chord
6
2
2
0
0
-2
-2
0
0.5
1
3 Bladed
2 Bladed
4
Cp
Cp
6
3 Bladed
2 Bladed
4
0.5
Normalized Chord
1
0
Normalized Chord
0.5
1
Normalized Chord
Figure 5.3: Pressure Distribution(Cp ) at different span wise locations of the wind turbine
blade(r/R) for three and two bladed turbine.
6
6
3 Bladed
2 Bladed
2
2
0
0
-2
-2
0
0.5
3 Bladed
2 Bladed
4
Cp
4
1
0
Normalized Chord
6
2
2
0
0
-2
-2
0
0.5
Normalized Chord
1
3 Bladed
2 Bladed
4
Cp
Cp
6
3 Bladed
2 Bladed
4
0.5
Normalized Chord
1
0
0.5
1
Normalized Chord
M.Sc. Thesis
5.2 Scaling based on Solidity Ratio
35
6
6
3 Bladed
2 Bladed
2
2
0
0
-2
-2
0
3 Bladed
2 Bladed
4
Cp
4
0.5
1
0
0.5
Normalized Chord
Cp Section 11 r/R=0. 476
6
3 Bladed
2 Bladed
4
Cp
1
Normalized Chord
2
0
-2
0
0.5
1
Normalized Chord
3500
3blade
2 blade
3000
2500
Fx
2000
1500
1000
500
0
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0.9
1
r/R
Figure 5.6: Axial force distribution along the span for three and two bladed turbine.
MSc. Thesis
36
Also a good comparison would be the distribution of the forces along the span of the blade.
As the pressure distributions are plotted only at specific sections over the blade, analysis of
forces over the blade would also be helpful in verifying the scaling.
Figure 5.6 shows the distribution of axial forces along the blade span. This force is positive
along the direction of the wind. This force is made dimensionless by using the density of air,
the input velocity and the chord length at the specified span location which can be written
as,
Fx =
Fx
1
ρv 2 cspan
2
It is seen that the axial force distribution for the two bladed turbine is lower than that of
the three bladed turbine. This is explained from the previous plots of pressure distribution
where it is seen that for a two bladed turbine the maximum pressure coefficient is lower than
that of the three bladed turbine. Because of this, the two bladed turbine produces lower lift
per section that the three bladed turbine and hence the lower axial force distribution.
5.3
Sectioning the Blade
Now that test model has been converted from a three bladed turbine to a two bladed turbine,
the effects of beheading (such as cross flow across the tip) in the outer sections can be visualized by performing simulations by means of the panel code.
The 2 bladed configuration was obtained by maintaining the same solidity ratio as that of the
three bladed turbine while increasing the chord. The blade was then sectioned at different
span lengths. This was all done by maintaining the same operating conditions that is the
same wind speed and the tip speed ratio( ≈ 6.8) which are optimum operating conditions.
The pressure distribution from the various simulations are displayed in Figure 5.7, Figure 5.8,
Figure 5.9. The pressure coefficient is plotted against the normalized chord radius at different span sections which is normalized using the span of the longest section(40m).The plots
of 3blade refer to the full size 40m rotor with three blades. It is seen that sectioning the
blade till 25m has almost no effect on the pressure distributions on the inner sections of the
blade which are under observation. However the pressure distributions start to vary when
the blade is sectioned even further seen at a radial position of 0.4R. When the blade is made
smaller and smaller the circulation around the tip of the blade starts to have an impact on
the pressure distributions in the inner sections. If the blade is sectioned too close to the root,
there is a big change in the pressure plots.
As explained in the previous section, the pressure distribution at the innermost
spans(r/R=0.07 to 0.185) are incorrect due to the inability of the panel code to handle
cylindrical sections while the rest of the pressure distributions are more or less the same at
the root regions(r/R ≈ 0.3) of the wind turbine blade.
It can then be concluded that the blade can be sectioned from an original radius of 40m to a
new radius of 25m without any significant changes in the flow in the root section. Now the
model can be scaled accordingly to fit the dimensions of the OJF rotor.
M.Sc. Thesis
5.3 Sectioning the Blade
37
6
2 Bladed
3 Bladed
2 blade cut 20m
2blade cut 25m
2blade cut 30m
2blade cut 35m
Cp
4
2
0
-2
0
6
6
4
4
Cp
Cp
2
0
-2
-2
0.5
1
2
0
0
0.5
Normalized Chord
1
0
Normalized Chord
0.5
1
Normalized Chord
blade(r/R) for 3 bladed, 2 bladed and a 20,25,30,35m model.
6
4
4
Cp
6
2
2
0
0
-2
-2
0
0.5
1
0
6
6
4
4
2
0
-2
-2
0.5
Normalized Chord
1
2
0
0
0.5
Normalized Chord
Cp
Cp
Normalized Chord
1
0
0.5
1
Normalized Chord
MSc. Thesis
38
6
4
4
Cp
6
2
2
0
0
-2
-2
0
0.5
1
0
Normalized Chord
0.5
1
Normalized Chord
6
Cp
4
2
0
-2
0
0.5
1
Normalized Chord
5.4
Tip Speed Ratio Scaling
When down scaling a model, it is desirable to keep the Reynolds number and tip speed ratios
constant. For Reynolds number conservation during scaling, We can write that,
uOJF = Rscaling · uReal
and
2
ΩOJF = Rscaling
· ΩReal
This would mean, for a scaling ratio of 29, the tunnel speed has to be increased by a factor
of 29 times and the rotational speed 292 which is impossible to achieve. Hence only tip speed
ratio is scaled to maintain the same angle of attack distribution.
The tip speed ratio is the ratio of the rotational speed of the blade tip to the speed of the
incoming wind and can be defined as,
λ=
ΩR
V ,
M.Sc. Thesis
5.5 Effect of tip shapes on pressure distribution
Parameter
Vinf (m/s)
R (m)
ω (RPM)
λ
Full Scale Model
10
40
16.403
6.87
39
Wind Tunnel Model
8
0.850
617
6.87
Table 5.1: λ scaling.
Where λ is the tip speed ratio, Ω is the rotational speed in rad/s , R is the radius of the blade
and V is the incoming wind speed in m/s. Tip speed ratio is an important consideration as
it determines the angle of attack of each airfoil section. Ideally the airfoil sections have to be
located in such a way that they are operating at the maximum lit to drag ratio. So in order
to maintain the same angle of attack distribution after scaling the model, the tip speed ratio
has to be also scaled. As λ and V are constant, only the rotational speed increases when the
model is downscaled. This is shown in the table above.
This rotational speed is achievable in the OJF Rotor without any significant vibrations and
is hence chosen.
5.5
Effect of tip shapes on pressure distribution
In order to mitigate the effects of cross circulation on the blade tip, a number of tip shapes
as discussed in Hoerner and Borst (1985) were implemented in the design and simulated with
the panel method. Wing tips are effective in altering the flow and forces over the wing. The
actual location of the tip vortex and the parasitic drag is dependent on the design of the wing
tips. By altering the position of the tip vortex on the blades, it is possible to alter the induced
velocities and the pressure distributions over the blade span. The tip shape of the blade is
altered by implementing a swept back and a swept forward tip shape as seen in Figure 5.10
and the effect of this design on the pressure distributions are shown.
Figure 5.10: a) Flat tip b)Sweep back c)Sweep Forward plan view.
MSc. Thesis
40
Cp Section
1
6
Flat Tip
Sweep Forward
Sweep Backward
Cp
4
0.5
2
0
-2
0
0
0.5
1
0
6
6
4
4
2
0
-2
-2
0.5
1
2
0
0
0.5
Normalized Chord
Cp
Cp
Normalized Chord
1
0
Normalized Chord
0.5
1
Normalized Chord
blade(r/R) for different tip shapes.
6
4
4
Cp
6
2
2
0
0
-2
-2
0
0.5
1
0
6
6
4
4
2
0
-2
-2
0.5
Normalized Chord
1
1
Flat Tip
Sweep Forward
Sweep Backward
2
0
0
0.5
Normalized Chord
Cp
Cp
Normalized Chord
0
0.5
1
Normalized Chord
M.Sc. Thesis
5.6 Verification with Aero-Module
41
Cp Section 10r/R=0. 476
6
4
4
Cp
6
2
2
0
0
-2
-2
0
0.5
1
Flat Tip
Sweep Forward
Sweep Backward
0
Normalized Chord
0.5
1
Normalized Chord
Figure 5.11,Figure 5.12, Figure 5.13 show the pressure distributions obtained from UMPM
at various spanwise sections of the blade for the three tip shapes. It is seen that the pressure
distributions in the root region is not affected at all by the design of the tip. Even though
some differences are seen in the more outboard regions of the blade due to changes in the tip
geometry, it is mostly irrelevant in this case. Hence the flap tip shape which is the result of
simply sectioning the blade is adopted for the blade.
5.6
Verification with Aero-Module
Blade Element Momentum Method
In order to visualize the effects of sectioning the blade at the tip, simulations using the ECN
AeroModule was carried out in addition to the simulations from the panel code. The BEM
part of the simulation was done to see the effect of sectioning the blade on the Prandtl loss
factor. The input to the code consists of airfoil data in terms of 2D airfoil polars obtained
from experiments and other geometric parameters. Using the parameters of wind turbine
chord, span, twist, airfoil polars and t/c ratios, a BEM simulation was performed on the full
scale blade and the sectioned models.
26 spanwise elements are used to provide input regarding the blade. Prandtl root and tip
correction factors are taken into account for the simulations to account for the number of
blades. The Prandtl factor accounts for the annular averaged values of the axial and tangential
induction to the local induction at each element. More information regarding the input to
the ”AEROMODULE” code is given in appendix.
The results of the BEM simulation, in particular the Prandtl correction factors are shown
in the Figure 5.14 for the full blade and the sectioned blades. The span section has been
normalized to be between 0 and 1 by using the radius of the said blade either full or sectioned.
It is seen that sectioning the blade has no effect on the tip loss factor in the mid section of
the blade(r/R≈0.35 to r/R≈0.65). This can be explained by the derivation of the prandtl
correction factor which is defined as:
MSc. Thesis
42
Prandtl Tip Loss Factor
1
25m
30m
35m
40m
0.9
TIp Loss Factor
0.8
0.7
0.6
0.5
0.4
0.3
0.2
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0.9
1
Span Section(r/R)
Figure 5.14: Prandtl Correction Factor.
F = π2 cos−1 (e−π
(R−r)
d
).
Where, d is the distance between the vortex sheets, r is the radius at which F is calculated
and R is the radius of the rotor. As the blade becomes shorter due to the sectioning, the
distance between the vortex sheets start to decrease when the blades are rotating at the same
angular frequency. Hence the induced velocities at the tip for the actual blade is smaller when
compared to the induced velocities of the sectioned blade. The same reasoning can be applied
to the root correction factor.
Lifting line/Free Vortex Wake Method Method
In order to visualize the effects of sectioning the blade at the tip, the lifting line free vortex
wake code AWSM was used. The lifting line part of the simulation was done to see the effect
of sectioning the blade on the induced velocities. The input to the code consists of airfoil
data in terms of polars and other geometric parameters. It should be noted that in this case,
the distribution of chord over the blade is considered negligible when compared to the span.
More information is presented regarding the input parameters are provided in the appendix.
26 spanwise elements are used to provide input regarding the blade.
The distribution of the induced velocities along the span wise sections of the blade are
obtained from the lifting line code and are shown in the form of the AWSM factor which
is the ratio of the induced velocity on the blade to the averaged induced velocity over the
entire azimuth. In Figure 5.15, the span sections have been normalized between 0 and 1 by
M.Sc. Thesis
5.7 Results
43
AWSM Tip Loss Factor
1.5
40m
25m
30m
35m
1
0.5
0
F
-0.5
-1
-1.5
-2
-2.5
-3
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0.9
1
Span Section(r/R)
Figure 5.15: AWSM Correction Factor.
using the length of that particular blade as the scale. It is seen that sectioning the blade has
no effect on the tip loss factor in the mid section of the blade(r/R≈0.35 to r/R≈0.65).
It can be seen that the correction factors remain constant for different blade lengths which is
an indication that there are no cross flow effects across the tip region of the blade in the mid
span. The higher values of the AWSM correction factor near the blade root is due to the
rotational effects which contribute to a higher lift in the inner sections. Is is seen that AWSM
tip loss factor seems to increase as we move towards the tip. This is because when the blade
geometry is replaced by a single line, the induced velocity becomes infinite according to the
Biot-Savart law. It can be concluded that the effect of cutting off the blade is not present
on the pressure distributions at different span sections but is seen in the induced velocity
distributions.
5.7
Results
From the results of the simulations, some conclusions can be drawn.
• It is seen from the pressure distributions obtained from the panel code, that it is possible
to scale from a three bladed wind turbine model to a two blade turbine model just by
modifying the chord distribution over the span wise direction while keeping the solidity
ratio constant.
MSc. Thesis
44
• Tip speed ratio can be applied to obtained an optimum angle of attack distribution over
the blade span when scaling is applied. This can also be discerned from the pressure
distributions obtained at the various spanwise positions from the panel code.
• The tip speed ratio scaling and the solidity scaling are verified with the ECN AEROModule code and the blade element method where the induced velocity distributions
over the blade surface remain constant over the mid span (r/R ≈0.3 to r/R≈0.7).
• Tip shape has more or less no influence on the induced velocities over the blade span.
The analysis of data from the panel code indicates that it is enough to maintain a flat
tip instead of incorporating a swept back or swept forward tip design into the blade.
M.Sc. Thesis
Chapter 6
Blades and Experimental Set-Up
The objective of the experiment was to test the effectiveness and influence of the spoilers
on the wind turbine loads and also capture a detailed measurement and analysis of the flow
around the blade with the flow control devices on and off. This is done by using a loads setup to record the blade performance and stereoscopic particle induced velocimetery. S-PIV is
done to capture the details of the flow around the blade.
In this chapter, the design of the blade is discussed along with the specifics of the wind tunnel,
the PIV set up, the varying spoiler configurations and the test matrix configuration setups
are discussed.
6.1
Blade Design
Figure 6.1: Chord and Twist Distributions.
The wind turbine used for the experiment is a 2 bladed scaled and modified model of the
Nordex N80 wind turbine with a design optimal tip speed ratio of 6.8. The rotor radius is
85cm from root to tip with a maximum twist of 9 degrees at the root and a minimum twist
of 2.3 degrees at the tip. The maximum chord is 17.74 cm which occurs at a radial position
MSc. Thesis
46
of 31 cm.
The twist and the chord distributions of the blade are displayed in Figure 6.1. Detailed
information can be found in ??. The root section is circular which is transitioned into a
modified DU405 thick trailing edge airfoil with maximum thickness at 40 percent of the
chord. The Reynolds number distribution along the span wise direction of the blade is seen
in Figure 6.3. The maximum Reynolds number at the tip at a tip speed of 6.8 is ≈ 300,000.
The Nacelle is about 38cm in diameter and is made in such a way that its influence on the
flow is negligible. A DC brushless motor, a hall encoder, a gear box and a optical trigger
are built into the nacelle.The gearbox to the motor is coupled with a gear ratio of 1:5. The
nacelle is placed on a 3m tall tower.
Figure 6.2: Spanwise Reynolds Number distributions.
Figure 6.3: Angle of attack distributions.
The wind tunnel model is shown in Figure 6.4. The nacelle has a diameter of 38cm and the
tower diameter is 20cm. The distance from the tip of the blade to the axis of rotation is
104cm. As stated in the previous chapter, the blade has been designed to have a thicker
root section so that the three dimensional effects that occur in the root region are more
pronounced. The angle of attack distribution of the blade is shown in Figure 6.4.
M.Sc. Thesis
6.1 Blade Design
47
Figure 6.4: Wind tunnel model.
6.1.1
The OJF Tunnel
The experiments are carried out at the open jet facility(OJF) at Delft University of Technology. The tunnel is closed loop open jet wind tunnel with an octagonal nozzle and a contraction
ratio of 4:1. The test section is 6m in width, 6.5m in height and 11m length in the inflow
direction with a maximum achievable speed of 35 m/s. The flow in the tunnel is driven by a
large fan rated at 500kW. A schematic of the tunnel is seen in Figure 6.5. The flow is uniform
with a maximum turbulent intensity of 0.5 %. The test section is maintained at a constant
temperature by using a radiator which is rated at 350kW. The flow is calculated to be steady
upto 3m from the exit which corresponds to little bit more than 3 turbine diameters of the
current model.
MSc. Thesis
48
Figure 6.5: Open Jet Facility.
6.2
Force Balance Experiment
Figure 6.6: Convention for the force balance.
To obtain the variation of the loads with respect to the tip speed ratio of the wind turbine,
force balance experiments are done. The forces on the blades are obtained by using an external
six component balance. The forces are obtained by running the blade at different rotational
speeds and measuring the axial component of the forces induced. The balance consists of six
Wheatstone bridges that give a voltage output that corresponds to the loads that act on the
blades. The balance system is capable of measuring upto 130% of the load range of ± 250N
in the axial direction, ± 500N in the y-direction and ± 500N in the vertical direction. The
conventions of the coordinate axes are shown in Figure 6.6.
Only the axial force is measured during the course of the experiment as the torque sensors
were misaligned and hence not working properly. The loads experiments are carried out
by studying the variation of the loads for a varying tip speed ratio and also for a range of
velocities while maintaining the tip speed ratio. The blades are mounted and the coefficient
of thrust is calculated for varying tip speed ratios by either changing the rotational speed of
the turbine or the free stream velocity. The forces and loads due to the tower and the nacelle
are calculated by dismounting the blades from the set-up and the running the rotor for the
same conditions that is for different wind tunnel settings. These values are then subtracted
from the values from the blades on configuration to calculate the forces exerted by the blades.
M.Sc. Thesis
6.2 Force Balance Experiment
6.2.1
49
PIV Set Up
The stereoscopic set up is seen in Figure 6.7. The entire stereo system which consists of
the two cameras and laser is mounted on a traversing system. This system is used in order
to scan the flow field at different span wise positions of the blade. The traverse system is
capable of moving in two directions with a range of 130cm in x-direction and 35cm in the
y-direction. This ensures that when the traverse system is moved, the cameras and the lasers
move along with it making sure that the plane of measurement is always in the field of view
of the cameras and the lasers.
Figure 6.7: Stereoscopic Set Up.
The field of interest is illuminated by a low speed Quantel Evergreen Nd:YAG laser system
with a pulse energy of 200mJ/pulse at a maximum frequency of 15Hz. The wavelength is
about 527nm. The laser light is focused into a 3mm thick sheet of width 25cm at the field
of view by a combination of spherical lens of focal lengths f=-40mm and f=-50mm and a
cylindrical lens of focal length f=75mm.
The cameras used to capture the region of interest are two LaVision Imager Pro LX 16
Mpx cameras. These cameras have a resolution of 4870x3246 P x2 with a pixel pitch of
7.4 µ m/px . The required field of view(FOV) size of 340*250mm is obtained by using a
f=180mm lens and a f number setting of 5.4. The magnification factor is calculated as
0.103. To prevent peak locking, the focusing plane is offset with respect to the laser plane.
Schiempflug adapters are attached to the cameras to avoid defocussing and to obtain an
uniformly focused image plane.
The flow is seeded using a SAFEX smoke generator with SAFEX mix which provides
particles of size 1 µ m. The smoke is made up of diethyl glycol mixture which is turned
into smoke by letting it drip over a electrically heated hot plate. Constant levels of
seeding is achieved by using a control unit placed in the tunnel control room. The pulse
MSc. Thesis
50
for the phase locking system is provided by a ”opto-coupler” TCST 2103 together with
a disc perforated at a set azimuthal position which rotates with the turbine shaft. This
system is capable of providing 1 or n pulses per revolution( in this experiment 1). This
pulse sent to the programmable transfer unit(PTU) which triggers the cameras and the lasers.
6.2.2
Spoilers
Two spoilers of different heights were chosen for the experiment. The spoilers were LM Tshaped spoilers with a height of 10mm and 20mm respectively and were sheet metal formed.
The angle between the ”T-shape” is 55 degrees. The spoiler is placed on the root pressure
side and is fixed to the blade by using a double sided tape with its front pointing towards
the leading edge. The reference point of placing the spoiler is the rounded edge. The length
of each spoiler is about 50mm and 4 of them were machined so that a spoiler length span of
100mm (50mm*2) can be achieved on each blade. The 3D drawing of the spoilers are shown
in Figure 6.8 and Figure 6.9.
Figure 6.8: Drafting of the Spoiler1.
Figure 6.9: Drafting of the Spoiler 2.
M.Sc. Thesis
6.3 Measurement Matrices
Configuration
1
2
3
4
5
6
Spoiler
1
1
1
1
1
1
51
Spanwise Position(m)
0.01-0.11
0.01-0.11
0.01-0.11
0.07-0.17
0.07-0.17
0.07-0.17
Chordwise Position(x/c)
0.75
0.85
0.95
0.75
0.85
0.95
Table 6.1: Spoiler 1 configurations.
Configuration
1
2
3
4
5
6
Spoiler
2
2
2
2
2
2
0.01-0.11
0.01-0.11
0.01-0.11
0.07-0.17
0.07-0.17
0.07-0.17
0.75
0.85
0.95
0.75
0.85
0.95
Table 6.2: Spoiler 2 configurations.
Configurations
For parameter studies, the spoilers are placed in various span wise and chord wise positions
to study its impact on the thrust and power generation. Three chord wise and two span wise
positions are chosen for both the spoilers, 12 in total. The configuration names and their
parameters are given in Table 6.1 and Table 6.2. Two spanwise locations are chosen in order
to study the rotational effects on the performance.
6.3
Measurement Matrices
The chord wise measurement setup which is used in the experiment to measure the velocity
field around the blades is shown in Figure 6.10. The laser is placed below the blades and
is made vertical by using a 90 degree mirror. The laser sheet is about 25cm wide and 3mm
thick at the field of view and. The cameras are placed perpendicular to the laser sheet and
are focused by using the Schiempflug adapters. The blade is kept at the same position and
the cameras and lasers are moved along the blade span by using the traverse system. The
measurements are made at Θ=0 azimuth angle which corresponds to a 3 o’clock position
when looking from upwind to downwind direction.
Separate measurements are made for the suction and pressure sides to because the blade
casts a shadow when the entire blade is imaged in one view. The suction and pressure side
measurements are then stitched together to obtain the distribution of velocities around the
blade.
Using the velocities, the pressure distributions around the blades can be reconstructed by
solving the incompressible Navier-Stokes equations from which the loads can be calculated.
MSc. Thesis
52
Figure 6.10: Chord Wise Measurement Set Up [Akay et al. (2014)].
The measurements are done for 13 radial positions of the blade with 10 of them in the root
region.
Test Conditions
For the loads experiments, the clean blade and the blades with the spoiler on configurations
are tested for a range of velocities and tip speed ratios. The tip speed ratio was varied between
3.5 to 8.9 by keeping the rotational speed of the blades constant and varying the velocities.
A test case of constant tip speed ratio was also carried out by varying both the rotational
speed and the inlet velocity.
The PIV runs were carried out for the design tip speed ratio for both the clean and one spoiler
configuration(spoiler 1 configuration1). The parameters of the PIV set-up are shown in table
6.3.
M.Sc. Thesis
6.3 Measurement Matrices
Parameter
Field of View
F-Number
Seperation Time
Window Size
Focal Length
Magnification Factor
Laser Sheet Width
Laser Sheet Thickness
Ensemble Size
Overlap
Processing
53
Value
350*280 mm
4.8
100 µ s
64x64 to 16x16
180 mm
0.103
30cm
3mm
100 image pairs
50 %
Stereo cross-correlation
Table 6.3: PIV Measurement Parameters
MSc. Thesis
54
M.Sc. Thesis
Chapter 7
Results
This chapter deals with the results obtained from the experimental campaign. The first
few sections are devoted to analysis of the data from the loads experiments from which the
influence of the spoilers can be deciphered from the plots of thrust coefficients (CT ). It was
also noticed after the experiments that the data from the PIV experiments were not good
enough to recreate the velocity fields as the separation time was set too high.
7.1
Loads Data
The results from the loads experiments for different spoiler configurations are displayed in
the form of variation of thrust coefficients with respect to velocity and tip speed ratios. The
force measurements were obtained by means of torque and thrust strain gauges placed in
the front and rear of the OJF rotor nacelle with an excitation voltage of 5V. The signals are
then amplified with an amplification factor of 100.
The recording frequency was 10,000 Hz and the recording was done for 10 seconds for all the
cases. The values of strain gauge displacements are then averaged over the azimuth and the
number of rotations to calculate the thrust values. The different configurations are listed out
again in Table 7.1 for easier reading.
Configuration
1
2
3
4
5
6
Spoiler
1,2
1,2
1,2
1,2
1,2
1,2
0.01-0.11
0.01-0.11
0.01-0.11
0.07-0.17
0.07-0.17
0.07-0.17
0.75
0.85
0.95
0.75
0.85
0.95
Table 7.1: Spoiler 1,2 configurations.
MSc. Thesis
56
7.1.1
Results
Thrust Coefficients
The thrust coefficients are plotted as functions of velocity and tip speed ratio for all the
configurations for both the spoilers in Figure 7.1. The influence of spoilers is clearly seen
from the plots of thrust where most of the spoiler on configurations show a higher thrust and
power values when compared to the clean configuration. It is seen that in most of the cases,
the spoiler on configurations work better than the clean blade in terms of thrust coefficients.
A quick initial glance at the plots indicates that the first spoiler might be performing better
than spoiler 2. The figures Figure 7.2 and Figure 7.3 show the percentage change in thrust
as a function of tip speed ratio defined by,
∆Ct =
Ct Spoiler−Ct Clean
Ct Clean
From these plots, it is seen clearly that the spoiler 1 works better than spoiler 2 for the entire
range of tip speed ratios. This is supported by the findings of Li et al. (2003) where it is
recommended that the gurney flap should be submerged in the boundary layer for obtaining
the optimum lift to drag ratio. Spoiler 2 with its taller height is not submerged fully in the
boundary layer at its location and hence is not operating at optimum. At the optimum tip
speed ratio, the configuration 1 works the best for spoiler 1 and the configuration 2 works the
best for the spoiler 1 with a thrust increase of 7 % for spoiler 1 and 4 % for the spoiler 2.
The variation of thrust with respect to the spoiler position is then discussed. It is seen that
when the spoiler is moved towards the trailing edge, the increase in thrust at the design tip
speed is higher in accordance with Wang et al. (2008) where it is stated that the increment in
lift coefficient is reduced when the location of the flap is moved away from the trailing edge.
The decrease in thrust at higher tip speeds is probably due to the bursting of the leading
edge bubbles at higher angles of attack as discussed in the paper.
Also it is seen that as a whole, the effectiveness of the spoilers when placed right at the
root(r/R=0.01m-0.11m) is higher than placing it relatively outward(r/R=0.07m to 0.17m).
This can be due to a variety of reasons. The second set of configurations(4,5,6) are placed
more radially outboard(r/R=0.21) when compared to the first set(r/R=0.1). The effects due
to rotation is more pronounced in at the more inboard regions which translates to higher lift
generation when compared to the outboard regions.
This effect along with the presence of the spoilers lead to a higher thrust and power values. The second possible explanation are the Reynolds number effects. Higher Reynolds
numbers present more outboard region of the blade might affect the performance of the spoilers. Another possible effect is the radial outflow due to the centrifugal and Coriolis forces l
forces which makes the boundary layer thinner. It is seen from Figure 7.3, that the second
spoiler when placed more outboard from the root section affect the performance of the blade
adversely.
M.Sc. Thesis
7.1 Loads Data
57
Spoiler 1
1.1
Clean
0.75 x/c & 0.01-0.11m
0.85 x/c & 0.01-0.11m
0.95 x/c & 0.01-0.11m
0.75 x/c & 0.07-0.17m
0.85 x/c & 0.07-0.17m
0.95 x/c & 0.07-0.17m
1
0.9
0.8
CT
0.7
0.6
0.5
0.4
0.3
0.2
0.1
3
3.5
4
4.5
5
5.5
6
6.5
7
7.5
6
6.5
7
7.5
λ
Spoiler 2
1.1
Clean
0.75 x/c & 0.01-0.11m
0.85 x/c & 0.01-0.11m
0.95 x/c & 0.01-0.11m
0.75 x/c & 0.07-0.17m
0.85 x/c & 0.07-0.17m
0.95 x/c & 0.07-0.17m
1
0.9
0.8
CT
0.7
0.6
0.5
0.4
0.3
0.2
0.1
3
3.5
4
4.5
5
5.5
λ
Figure 7.1: Thrust coefficients (CT ) for various configurations for spoiler 1,2 as a function of tip
speed ratio (λ).
MSc. Thesis
58
Results
Explanation
The percentage increase of thrust for the spoiler on configurations with respect to the clean
configuration is shown in Figure 7.2 and Figure 7.3 as a function of tip speed ratio.
Spoiler-1
S1 0.01 to 0.11m &0.75 x/c
10
∆CT
∆CT
10
0
-10
8
4
λ
S1 0.01 to 0.11m &0.95 x/c
10
∆CT
∆CT
6
0
-10
6
8
0
8
4
λ
S1 0.07 to 0.17m &0.85 x/c
10
∆CT
∆CT
6
λ
S1 0.07 to 0.17m &0.75 x/c
-10
4
10
0
-10
4
10
S1 0.01 to 0.11m &0.85 x/c
0
-10
6
8
λ
S1 0.07 to 0.17m &0.95 x/c
0
-10
4
6
λ
8
4
6
8
λ
Figure 7.2: Percentage change in thrust (CT ) Spoiler 1 as a function of tip speed ratio (λ).
Figure 7.2 shows the percentage variation in thrust coefficient for various spoiler-1 configuration with respect to the clean case as a function of tip speed ratio.
The inboard spoiler configurations(0.01m to 0.11m) show the most promise with the highest
increase in thrust at lower wind speeds with a maximum of 10 % at λ ≈ 6 for configuration
1. For the same conditions at the outboard sections(config 4,5,6), the increase in thrust is
almost negligible. Comparison of the spanwise locations of the spoiler for the same chordwise location indicates that the inboard spoilers perform better than the outboard ones. It
is also seen clearly that the percentage change in thrust is positive till a certain tip speed
after which it starts to decrease and even become negative for some configurations which is
more prominent in the outboard configurations. The reason behind this is being that, after a
certain tip speed ratio, the root section starts to operate in stalled conditions. This behavior
can be seen in all the configurations.
M.Sc. Thesis
7.1 Loads Data
59
Spoiler-2
Figure 7.3 shows the percentage variation in thrust coefficient for various spoiler-2 configuration with respect to the clean case as a function of free stream velocity.
As stated earlier, the height of spoiler-2 is larger than that of the first spoiler and the spoiler
is not fully immersed in the boundary layer. Hence the second spoiler does not operate as
efficiently as spoiler 1. Similar to spoiler 1, the inboard spoiler configurations show the most
promise with the highest increase in thrust at lower wind speeds with a maximum of 5
% at λ = 6.8(Design TSR) for configuration 1. For the same conditions at the outboard
sections(config 4), the increase in thrust is almost negative. Comparison of the spanwise
locations of the spoiler for the same chordwise location indicates that the inboard spoilers
perform better than the outboard ones. It is also seen clearly that the percentage change in
thrust is positive till a certain velocity after which it starts to decrease and even become negative for some configurations. The reason behind this is being that, after a certain velocity,
the root section starts to operate in stalled conditions. This behavior can be seen in all the
configurations.
S2 0.01 to 0.11m &0.75 x/c
10
∆CT
∆CT
10
0
-10
8
4
λ
S2 0.01 to 0.11m &0.95 x/c
10
∆CT
∆CT
6
0
-10
6
8
λ
S2 0.07 to 0.17m &0.75 x/c
0
8
4
λ
S2 0.07 to 0.17m &0.85 x/c
10
∆CT
∆CT
6
-10
4
10
0
-10
4
10
S2 0.01 to 0.11m &0.85 x/c
0
-10
6
8
λ
S2 0.07 to 0.17m &0.95 x/c
0
-10
4
6
λ
8
4
6
8
λ
Figure 7.3: Percentage change in thrust (CT ) Spoiler 2 as a function of tip speed ratio (λ).
MSc. Thesis
60
Results
Reynolds Number Effects
This section discusses the influence of Reynolds number on the thrust coefficients for the clean
and the spoiler on configurations.
The variation of thrust with respect to velocity for constant tip speed ratios are displayed
in Figure 7.4. The figure Figure 7.5 show the percentage change in thrust as a function of
tip speed ratio. Looking at the figure, the root configurations(0.11 to 0.11m) perform better
than the radially outward configurations. Also it is seen that the spoiler 1 performs better
than the second spoiler in most of the cases. This is because of the reason that the spoiler 1
is more immersed in the boundary layer than the second spoiler.
It is also seen that for configurations 3 and 6 (0.95 x/c), there is a negative change in thrust.
This is because of the mounting angle of the spoilers.When placed at 0.95 x/c, the mounting
angle which is the angle between the chord line and the spoiler line is very small as studied
in Li et al. (2003).
According to Li et al. (2003), the smaller the mounting angle, the smaller there is an advantage
due to the spoiler. At smaller mounting angles, the lift increase due to the spoiler is not enough
to offset the drag penalty associated with it. The more inclined the spoiler is, the more the
increase in lift to drag ratio. The negative change in the value of thrust for the more outboard
spoilers specifically configurations 4,5,6 is a combination of this effect and Reynolds number
effects.
The Figure 7.5 shows the variation of thrust for different spoiler cases with respect to the
velocity at the same tip speed ratio. This can be construed as Reynolds number effects.
Reynolds number here is defined as,
Re =
λ·UInf ·c
.
ν
As λ is constant along with the chord c and the kinematic viscosity ν, only the change in
velocity leads to change in the Reynolds number and hence can be interpreted as Reynolds
number effects. It is seen that with increasing velocity the change in thrust is positive for
spoiler 1 till a particular velocity after which the change is negative. This is possibly because
of the fact that the boundary layer gets thinner with increase Reynolds number that it leads
to earlier onset of transition.
As stated earlier, the configurations more outboard have a less enhancing effect on the
thrust and also negative past 6m/s. This is possibly separated flow because of the high
inlet velocities and rotational speeds where the root section stalls. It can be said that the
spoilers no matter what size and shape have a negative effect on the thrust when placed in
stalled conditions.The smaller spoiler works better than the larger one because of the reasons
explained in the previous chapter.
M.Sc. Thesis
7.1 Loads Data
61
Spoiler 1
1
Clean
0.75 x/c & 0.01-0.11m
0.85 x/c & 0.01-0.11m
0.95 x/c & 0.01-0.11m
0.75 x/c & 0.07-0.17m
0.85 x/c & 0.07-0.17m
0.95 x/c & 0.07-0.17m
0.9
CT
0.8
0.7
0.6
0.5
0.4
4
4.5
5
5.5
6
6.5
1
7.5
8
Clean
0.75 x/c & 0.01-0.11m
0.85 x/c & 0.01-0.11m
0.95 x/c & 0.01-0.11m
0.75 x/c & 0.07-0.17m
0.85 x/c & 0.07-0.17m
0.95 x/c & 0.07-0.17m
0.9
0.8
CT
7
Velocity(m/s)
Spoiler 2
0.7
0.6
0.5
0.4
4
4.5
5
5.5
6
6.5
7
7.5
8
Velocity (m/s)
Figure 7.4: Thrust coefficients(CT ) for various configurations for spoiler 1,2 as a function of
wind speed.
MSc. Thesis
62
Results
Spoiler-1
4
6
V(m/s)
S1 0.01 to 0.1m &0.95 x/c
0
-10
10
4
6
-10
6
8
4
6
8
V(m/s)
S1 0.07 to 0.17m &0.75 x/c
0
-10
V(m/s)
S1 0.07 to 0.17m &0.85 x/c
4
0
10
8
0
S1 0.01 to 0.11m &0.85 x/c
-10
8
∆CT
∆CT
10
10
∆CT
0
-10
∆CT
S1 0.01 to 0.11m &0.75 x/c
10
∆CT
∆CT
10
4
6
8
V(m/s)
S1 0.07 to 0.17m &0.95 x/c
0
-10
V(m/s)
4
6
8
V(m/s)
Figure 7.5: Percentage change in Thrust(∆CT ) Spoiler 1 as a function of wind speed.
Figure 7.5 shows the percentage variation in thrust coefficient for various spoiler-1 configuration with respect to the clean case as a function of free stream velocity. The tip speed ratio
is maintained at the design tip speed ratio. The y-axis ∆cT is defined as,
∆CT =
Ct Spoiler−Ct Clean
Ct Clean
The inboard spoiler configurations show the most promise with the highest increase in thrust
at lower wind speeds with a maximum of 8 % at v = 4m/s for configuration 1. For the
same conditions at the outboard sections(config 4), the increase in thrust is almost negligible.
Comparison of the spanwise locations of the spoiler for the same chordwise location indicates
that the inboard spoilers perform better than the outboard ones. It is also seen clearly that
the percentage change in thrust is positive till a certain velocity after which it starts to
decrease and even become negative for some configurations. The reason behind this is being
that, after a certain velocity, the root section starts to operate in stalled conditions. This
behavior can be seen in all the configurations.
M.Sc. Thesis
7.1 Loads Data
63
Spoiler-2
4
6
V(m/s)
S2 0.01 to 0.11m &0.95 x/c
0
-10
10
4
6
4
6
V(m/s)
8
4
6
8
V(m/s)
S2 0.07 to 0.17m &0.75 x/c
0
-10
V(m/s)
S2 0.07 to 0.17m &0.85 x/c
0
-10
10
8
S2 0.01 to 0.11m &0.85 x/c
0
-10
8
∆CT
∆CT
10
10
∆CT
0
-10
∆CT
S2 0.01 to 0.11m &0.75 x/c
10
∆CT
∆CT
10
4
6
8
V(m/s)
S2 0.07 to 0.17m &0.95 x/c
0
-10
4
6
8
V(m/s)
Figure 7.6: Percentage change in Thrust(∆CT ) Spoiler 1 as a function of wind speed.
Figure 7.6 shows the percentage variation in thrust coefficient for various spoiler-2 configuration with respect to the clean case as a function of free stream velocity.
As stated earlier, the height of spoiler-2 is larger than that of the first spoiler and the spoiler
is not fully immersed in the boundary layer. Hence the second spoiler does not operate as
efficiently as spoiler 1. Similar to spoiler 1, the inboard spoiler configurations show the most
promise with the highest increase in thrust at lower wind speeds with a maximum of 7 %
at v = 7m/s for configuration 1. For the same conditions at the outboard sections(config
4), the increase in thrust is almost negligible. Comparison of the spanwise locations of the
spoiler for the same chordwise location indicates that the inboard spoilers perform better
than the outboard ones. It is also seen clearly that the percentage change in thrust is positive
till a certain velocity after which it starts to decrease and even become negative for some
configurations. The reason behind this is being that, after a certain velocity, the root section
starts to operate in stalled conditions. This behavior can be seen in all the configurations.
MSc. Thesis
64
Results
M.Sc. Thesis
Chapter 8
Uncertainty Analysis
There exists some uncertainties in the results because of the fact that measurements in the
wind tunnel are always affected by experimental errors which might be systematic or random
in nature. Systematic errors are occur consistently throughout the experiments and random
errors are statistical fluctuations in the measured data and vary every observation. Determining the source of errors during the experimental campaign is hard because the set up is
complex. This chapter discusses about the possible sources of experimental errors and its
impact on the results. The uncertainties might be of the form of wind tunnel uncertainties,
uncertainties in the positions of the spoilers.
8.1
Flow Uncertainities
The flow in the wind tunnel is under the influence of external parameters such as temperature
of air, shear layer expansion in the test section, size of the model in comparison to the test
section. The blockage ratio of the wind tunnel model was 5 % so blockage effects can be
considered as negligible as the flow is allowed to expand in the open test section. As the
experiments were done throughout the day, there are some changes in the ambient temperature
that affects the flow properties in the test section. Hence the test section was maintained at
a constant temperature by means of a heat exchanger.
P = ρRT
This can be re-written as,
ρ=
ρlower =
MSc. Thesis
P
RTlower
=
P
RT
101325
287×18.4+273
kg
= 1.211 m
3
66
ρU pper =
P
RTHigher
=
101325
287×21.3+273
kg
= 1.1996 m
3
The average velocity is the velocity set as tunnel input speed which is 8 m
s.
ρAverage =
P
RTAverage
=
101325
287×19.85+273
kg
= 1.2055 m
3
Conservation laws can be applied to calculate the difference in the velocities that occur due
to temperature differences.
ρAverage · uAverage = ρLower · uLower = ρHigher · uHigher
The lower and upper limits of velocity can be calculated.
uLower = 7.964m/s
uLower = 8.045m/s
Velocity u = 8m/s corresponds to a tip speed ratio of 6.8645. Including the effects of changes
in the temperature of the test section, it is seen that the tip speed ratio varies between 6.8956
and 6.8261. Study of older experiments at the OJF seem to indicate that the region of steady
flow reduces at a rate of 16.7 cm for every meter traveled downstream. At a distance of 1.5D
from the tunnel exit, the blade still lies in the steady region of the expanding jet. This is
within the shear layer and hence the unsteady characteristics in the flow are negligible. In
this region, the turbulence intensity is lower than 0.5 %.
The rotational speed of the rotor is kept at the desired value throughout the course of the
experiment by means of a motor and a RPM controlling mechanism. Though the motor
can keep the frequency of the rotor constant, it could not account for the instances when
oscillations occurred mainly when changing the tunnel speeds or changing the rotational
speed of the rotor. At the design tip speed ratio of 6.8645, the frequency was seen to be
oscillating between 10.267 to 10.312 Hz while the ideal frequency had to be 10.28 Hz. This
leads to a variation in tip speed ratio which can be calculated as follows,
λ=
λlower =
λHigher =
ωr
uinf
=
2πf r
uinf
2π×10.267×0.85
8
2π×10.312×0.85
8
= 6.850
= 6.880
M.Sc. Thesis
8.2 Blade Design
67
The variation of of the tip speed ratio from the design bound is 0.0645 on the lower bound
and 0.03 on the upper bound. This is a change of 0.215 % in the lower bound and 0.226 % in
the upper bound. Even though this value is pretty negligible, the uncertainty must be taken
into account. An error plot that shows information about the uncertainties that exist in the
calculation of the thrust forces are shown in the Figure 8.1.
1.1
Clean
1
0.9
0.8
CT
0.7
0.6
0.5
0.4
0.3
0.2
0.1
3
3.5
4
4.5
5
5.5
6
6.5
7
7.5
λ
Figure 8.1: Error plot for variation of thrust coefficients with respect to tip speed ratio for the
clean case.
8.2
Blade Design
The blades are created by 3 dimensional modeling tools where each span wise airfoil section
is imported and then blended together to form a three dimensional model. The tool used
to create the model was PTC Creo. It was seen that the interpolation between the sections
during the blending process was linear while this is not the case in the real wind turbine
blade. The blade was then machined in a CNC lathe. It was also seen that the two blades
weighed differently. Even though the difference in weight is pretty limited, there happens to
be a significant change in the centrifugal forces during rotations at tip speed scaled higher
speeds which was 617 RPM.
The blade was hand finished with blade matte paint which was sprayed. Spraying might have
lead to some imperfections on the blade surface which acts as surface roughness strips. The
middle span regions were not studied in the PIV sections as it was not an area of interest but
the shape of the transition might have affected the lift and drag of the airfoil sections thus
directly impacting the thrust and torque coefficients.
The pressure distributions at the transition region for the real blade and the wind tunnel blade
are shown in Figure 8.3 which shows the transition in the region of the thicker airfoil Figure 8.4
and Figure 8.5 which show the transition region in the more outboard airfoil sections. It is
MSc. Thesis
68
see that for the linear transition sections, the pressure peak is lower and the sections have a
higher pressure gradient which makes separation issues a possibility especially for the thicker
airfoil sections.This phenomenon is seen reoccurring in all the transition regions.
12
Spline
Linear
10
Fx
8
6
4
2
0
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0.9
1
r/R
Figure 8.2: Axial force distribution over the span.
The axial force distributions normalized with respect to the chord length is also shown in
Figure 8.2. This force is positive along the direction of the wind. This force is made dimensionless by using the density of air, the input velocity and the chord length at the specified
span location which can be written as,
Fx =
Fx
1
ρv 2 cspan
2
It can be seen that because of the higher adverse pressure gradient in the transition sections
of the actual wind tunnel model, the axial force which is produced by the wind tunnel model
is a little lower than the axial force of the real blade more at the root sections(r/R=0.3). This
has an impact on the performance results from the experiments.
M.Sc. Thesis
8.2 Blade Design
69
Cp
Airfoil Shape
0.5
6
Linear
actual
Linear
actual
2
y
Cp
4
0
0
-2
-0.5
0
0.2
0.4
0.6
0.8
1
0
c/cMax
Cp
6
0.2
0.4
0.6
0.8
1
x
Airfoil Shape
0.5
Linear
actual
Linear
actual
2
y
Cp
4
0
0
-2
-0.5
0
0.2
0.4
0.6
0.8
1
0
c/cMax
Cp
6
0.2
0.4
0.6
0.8
1
x
Airfoil Shape
0.5
Linear
actual
Linear
actual
2
y
Cp
4
0
0
-2
-0.5
0
0.2
0.4
0.6
c/cMax
0.8
1
0
0.2
0.4
0.6
0.8
1
x
blade(r/R) for linear and spline distributions for the root transition section.
MSc. Thesis
70
Cp
Airfoil Shape
0.5
6
Linear
actual
Linear
actual
2
y
Cp
4
0
0
-2
-0.5
0
0.2
0.4
0.6
0.8
1
0
c/cMax
Cp
6
0.2
0.4
0.6
0.8
1
x
Airfoil Shape
0.5
Linear
actual
Linear
actual
2
y
Cp
4
0
0
-2
-0.5
0
0.2
0.4
0.6
0.8
1
0
c/cMax
Cp
6
0.2
0.4
0.6
0.8
1
x
Airfoil Shape
0.5
Linear
actual
Linear
actual
2
y
Cp
4
0
0
-2
-0.5
0
0.2
0.4
0.6
c/cMax
0.8
1
0
0.2
0.4
0.6
0.8
1
x
blade(r/R) for linear and spline distributions for the second transition section.
M.Sc. Thesis
8.2 Blade Design
71
Cp
Airfoil Shape
0.5
6
Linear
actual
Linear
actual
2
y
Cp
4
0
0
-2
-0.5
0
0.2
0.4
0.6
0.8
1
0
c/cMax
Cp
6
0.2
0.4
0.6
0.8
1
x
Airfoil Shape
0.5
Linear
actual
Linear
actual
2
y
Cp
4
0
0
-2
-0.5
0
0.2
0.4
0.6
c/cMax
0.8
1
0
0.2
0.4
0.6
0.8
1
x
blade(r/R) for linear and spline distributions for the third transition section.
MSc. Thesis
72
8.3
Spoiler Position uncertainties
The way of mounting the spoilers onto the blade sections could also introduce an element
of uncertainty into the experiment. As seen before, the spoilers were sheet metal formed
and its orientation with respect to its position to the blade was discussed. It was seen that
the spoilers were placed on the pressure side and was fixed to the blade root by means of
double sided tape. The reference point of the spoiler was the rounded edge. The spoilers
were placed at 0.75,0.85,0.95% of the chord. This distance was measured from the trailing
edge with a fold able tape measure. The dimensional accuracy of the tape measure was type
2 which corresponds to a deviation of ±0.6 mm. The double sided tape which was used to fix
the spoilers onto the blade surface had a thickness of 0.19mm. Even though great care was
taken while fixing the spoilers to the respective configurations, the spoiler positions are still
susceptible to human and parallax errors. A good way to avoid these errors would be the use
of ensemble averaging where many sets of data collection are done for the same test conditions.
M.Sc. Thesis
Chapter 9
Recommendations and Future Work
A wind tunnel model was created which was able to recreate the actual root flow in a wind
turbine. Experiments are performed for different spoiler types, positions and configuration
and the performance data was recorded.
This chapter summarizes the key findings and observations while providing recommendations
for future thesis work.
9.1
Conclusions
From the simulations and the experiments, the following conclusions can be drawn.
• It is seen from the pressure distributions and the axial force distributions obtained from
the panel code, that it is possible to scale from a three bladed wind turbine model to
a two blade turbine model just by modifying the chord distribution over the span wise
direction while keeping the solidity ratio constant. This has to be done because the
OJF rotor could support only two blades.
• Tip speed ratio can be applied to obtained an optimum angle of attack distribution
over the blade span when scaling is applied. As it is impossible to maintain Reynolds
number conservation during scaling, tip speed ratio is conserved in order to maintain
the same angle of attack distribution over the blade span.
• The tip speed ratio scaling and the solidity scaling are verified with the ECN AEROModule code and the blade element method where the induced velocity distributions
over the blade surface remain constant over the mid span (r/R=0.3 to r/R=0.7).
• Sectioning the original blade from 40m to 25m has no effect on the pressure distributions
on the inner sections of the blade. It is seen that by beheading the blade beyond 25m,
the pressure distributions are influenced possibly due to cross flow issues.
MSc. Thesis
74
Recommendations and Future Work
• Tip shape has more or less no influence on the induced velocities over the blade span.
The analysis of data from the panel code indicates that it is enough to maintain a flat
tip instead of incorporating a swept back or swept forward tip design into the blade.
• It is seen that by applying the scaling parameters discussed in the previous points,
a wind turbine blade which recreates the root flow can be created. The maximum
Reynolds number at the tip is ≈ 300,000 at optimal tip speed ratio. This blade can be
then tested at the open jet facility where performance measurements are carried out.
• Initial performance results for the spoiler on and off configurations indicate that the
spoilers when placed in certain configurations have a positive performance on the rotor
thrust.
• The smaller spoiler works better than the second spoiler. This is because that the first
spoiler is completely immersed in the boundary layer at the root part of the flow.
• The influence of Reynolds number is clearly seen in the thrust plots where it is seen that
the spoilers cause a decrease in thrust at higher velocities usually after uinf = 7m/s.
• The performance of the spoilers are under the influence of boundary layer, rotational
and Reynolds number effects.
• Spoilers have an adverse effect on the performance at higher inlet velocities and rotational speeds.
• The experiment was subject to uncertainties in the form of flow parameters, blade design
and uncertainties in the position of the spoilers.
9.2
Recommendations
The experiment performed on the wind turbine model created to maximize the root effects
that occur on an actual wind turbine with a high enough Reynolds number at the tip at the
design tip speed ratio. Further studies can be performed to improve the design of the blade.
As time was a constraint during the course of the experiment, no CFD calculations could be
performed to verify the design of the blade. This can be done to study in detail the blade
design and also perform some simulations with the spoilers.
As the torque sensor was not working during the course of the experiments, the rotor head
at the OJF should be sent for maintenance so that the slip rings in the rotor can be repaired
to make better thrust and torque measurements. Better PIV studies have to be performed to
study the source of the change in performance originating due to the spoilers.
M.Sc. Thesis
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M.Sc. Thesis

Investigation of Root Spoilers on Horizontal Axis Wind

Transcription

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